- Email: [email protected]
Nuclear-Powered Payload Safety
CHAPTER
Nuclear-Powered Payload Safety
6
Firooz A. Allahdadi, Sayavur I. Bakhtiyarov, Gregory D. Wyss, Gary F. Polansky, Joseph A. Sholtis, Curt D. Botts
CHAPTER OUTLINE 6.1 6.2 6.3 6.4 6.5 6.6 6.7 6.8 6.9
Introduction to Space Nuclear Systems ........................................................... SNPS Launch History and Accidents ................................................................ Launch Abort Environments Affecting SNPSs .................................................... Containment Design........................................................................................ Risk Assessment for Nuclear Missions ............................................................ International Protocols and U.S. Environmental Review .................................... Nuclear Mission Launch Approval ................................................................... Nuclear Mission Launch Integration ................................................................ Symbols and Acronyms...................................................................................
256 268 275 335 338 348 355 357 361
It would be nice if the risks of doing space exploration could be borne only by the people like me who believe in doing them e just like it would be nice if people driving cars only put their own lives at risk. Unfortunately, it’s hard to see how to structure human life that way. If you put a big sculpture of a foot on top of the Ferry building in San Francisco, there’s a finite chance that it could fall over and injure a passerby e possibly even a passerby that thinks it’s really ugly. Should we forbid all public sculpture? After all, we don’t need it to survive (at least not in any obvious way). How can we justify even a slight chance that someone could die under that gargantuan heel? The value of human life is high, but it is not infinite. If we try and act as though it is infinite, it is in danger of becoming worthless. But there’s nothing to get upset about this time, right? After all the good guys won, didn’t they? Cassini’s in flight, on its way to Saturn. except that what’s going to happen the next time they try and build a space probe using an RTG? Is this another useful technology that’s going to be moved to the list of things that are politically untenable? Heads they win, tails they win. (Brenner, 1997)
This chapter introduces the concepts of Space Nuclear Power Systems (SNPSs). It briefly describes the history and nature of these ingenious machines. For more than five decades, larger and more powerful rockets have inspired awe Safety Design for Space Operations. http://dx.doi.org/10.1016/B978-0-08-096921-3.00006-4 Copyright Ó 2013 Elsevier Ltd. All rights reserved.
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among those privileged to observe their launches. But the very factor that makes the firing of a large rocket so spectacular e the consumption of tons of fuel in just seconds e also suggests practical limits for chemical propulsion. And the limits apply not only to propulsion. Any application that requires high power for more than the few minutes of a chemical rocket’s firing, or power levels greater than can be provided by collecting the Sun’s diffuse radiation energy, is a candidate for nuclear power. The reason for nuclear power in space is simple and compelling: an increase by a factor of as much as one million in the energy that is contained in a mass of nuclear fuel compared with the same mass of chemical fuel. Since the start of the space age, a range of nuclear power supply options has been developed by the United States to support the civilian and military space programs. Tomorrow’s space exploration and exploitation activities will require even larger quantities of reliable, long-lived power (Angelo, 1985).
6.1 Introduction to Space Nuclear Systems Space nuclear systems are vital to the exploration of space and have seen moderate changes in the basic design and fundamental operating schemes. Nuclear-based systems can supply electricity, heat, and propulsion for missions that are well beyond the capacities of solar power, fuel cells, and traditional chemical methods. Nuclear-based systems consist of (1) radioisotope thermoelectric generators (RTGs); (2) fission-based nuclear reactor power systems; (3) nuclear thermal propulsion (NTP); (4) nuclear electric propulsion (NEP), and other advanced nuclear technologies. Since the “Seebeck effect” was discovered in 1826, the basic concept of reliable space nuclear power has grown in efficiency and capacity. German physicist Thomas Johann Seebeck (Duckworth, 1960) published results of experiments conducted in 1822 in which he observed that electrical current is present in a series circuit of two dissimilar metals, provided the junctions of the two metals are at different temperatures (Figure 6.1). This thermoelectric effect increases as the temperature differential increases. The “Seebeck effect” is a result of two phenomena: charge carrier diffusion and phonon drag. Charge carriers (electrons, ions) will diffuse when one end of a conductor is at a different temperature than the other. Hot (or cold) carriers will diffuse from the hot (or cold) end to the cold (or hot) end since there is a lower density of hot (or cold) carriers at the cold (or hot) end of the conductor. The phonon drag phenomenon is an increase in the effective mass of conduction electrons due to contact with the crystal lattice. When an electron interacts with atoms in the lattice its charge alters or polarizes the adjacent lattice. This effect results in a decrease in the electron mobility, and consequently, conductivity decrease. As the magnitude of the thermopower increases with phonon drag, it is advantageous for direct energy conversion applications. Maintaining both connections at the same temperature, and periodically opening and closing one connection, causes a temperature
Thermoelectric effects.
FIGURE 6.1
6.1 Introduction to Space Nuclear Systems
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dependent alternating current (AC) voltage to be generated. This effect will still be observed even when the wires do not touch. Hence, one would conclude that no diffusion is needed for this phenomenon. The Seebeck effect is one of three reversible phenomena illustrating analogous processes associated to thermoelectricity, conductivity, and temperature. French scientist Jean-Charles Peltier in 1834 described the second closely related phenomena (known as the “Peltier effect”): the temperature at either junction changed proportionally to the voltage between the metal conductors. German scientist Heinrich Lenz in 1839 expanded on Seebeck’s and Peltier’s discoveries and showed that heat transfer at the junctions depends on the direction of the current in the circuit (“SeebeckePeltier effect”). In 1851, British physicist William Thomson (Kelvin) described the rate of heat created (or absorbed) in a current-carrying conductive material subjected to a temperature gradient (“Thomson effect”). A simple and most popular electrical tool based on measuring the SeebeckePeltiereThompson (SPT) effects is the thermocouple, which converts the thermal potential difference into electric potential difference. Several thermocouples connected in series are called a thermopile, which can be used to increase the output voltage. SPT effects are also the basic work principle of thermoelectric generators such as RTGs used to generate power from heat differentials. The United States has designed, tested and launched RTGs for more than three decades. The RTG systems have been used with great success in several extraterrestrial missions and space exploration experiments. Those missions would not have been possible without these important power sources. However, these RTG systems are limited in producing power up to tenths of kilowatts, making them less ideal for use in maintaining equipment operations and sustaining human life during long duration space missions beyond the Moon and Mars. In 1958 a U.S. project, Orion (Figure 6.2), was initiated by General Atomics (Shipps, 1964). The purpose of the Orion was to launch a spacecraft (~1000 tons) INTERMEDIATE PLATFORM 1st STAGE SHOCK ABS.
2nd STAGE SHOCK ABS.
UPPER MODULE SECTION (BODY)
PUSHER
PAYLOAD SECTION PROPELLANT MAGAZINES PROPULSION MODULE
FIGURE 6.2 Orion nuclear spacecraft.
6.1 Introduction to Space Nuclear Systems
FIGURE 6.3 Nuclear Engine for Rocket Vehicle Applications (NERVA).
utilizing a series of small nuclear explosions against a “pusher plate” to provide thrust. The project was terminated in 1963 when the Atmospheric Test Ban Treaty made it illegitimate, as the radioactive fallout could have been a main problem. The concept, nonetheless, is still living as one of the more efficient methods for producing propulsion energy. From 1959 to 1973 a U.S. multi-mission nuclear rocket program, “Nuclear Engine for Rocket Vehicle Applications” (NERVA), was conducted to substitute a nuclear power for chemical rockets at certain launch phases. Electric propulsion provides remarkable advantages over chemical propulsion. It enables new classes of Solar System exploration missions with multiple targets, eliminates or reduces launch windows required for gravity assists, reduces cruise time to distant targets, and reduces mission cost because smaller launch vehicles may be used (Scolese, 2002). The proposed NERVA system (Figure 6.3) used graphite-core reactors heating hydrogen and ejecting it through a nozzle delivering a minimum of 75,000 lbf thrust. A total of 20 reactor tests and two engine tests were conducted (Robbins & Finger, 1991) in Nevada and delivered thrust more than half that of the space shuttle main engines. Since the cancellation of the program, “nuclear rockets” have been proposed for space propulsion, but not for the initial launch phase. According to the U.S. National Aeronautics and Space Administration’s (NASA’s) 2006 Solar System Exploration Roadmap (NASA, 2006), the lunar and Mars explorations eventually will be manned missions in around 2030 to answer the following fundamental questions: • • •
How did the Sun’s family of planets and minor bodies originate? How did the Solar System evolve to its current diverse state? What are the characteristics of the Solar System that led to the origin of life?
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• •
How did life begin and evolve on Earth and has it evolved elsewhere in the Solar System? What are the hazards and resources in the Solar System environment that will affect the extension of human presence in space?
A scientific input to NASA in developing and achieving these goals and priorities is offered by the NASA Advisory Council (NAC) and by other science advisory groups, such as the Outer Planets Assessment Group (OPAG), the Venus Exploration Analysis Group (VEXAG), the Mars Exploration Program Analysis Group (MEPAG), and the Lunar Exploration Analysis Group (LEAG). In the Solar System Exploration (SSE) Roadmap (NASA, 2006) the proposed missions could be facilitated by Radioisotope Power Systems (RPSs) identified as one of the highest priority technologies. The following goals are established by the MEPAG for Mars exploration: • • • •
Determining if life ever arose on Mars. Understanding the process and history of climate on Mars. Determining the evolution of the surface and interior of Mars. Preparing for human exploration.
In November of 2011 NASA’s Mars Science Laboratory (MSL) launched “Curiosity,” a nuclear powered rover to Mars using an Atlas V 541 Launch Vehicle. The MSL mission will explore Mars habitability (i.e., capability of the Mars environment to sustain a life). Presently, NASA has developed two new types of RPSs. These are the MultiMission Radioisotope Thermoelectric Generator (MMRTG), with static power conversion, and the Stirling Radioisotope Generator (SRG), utilizing dynamic conversion. The MMRTG includes a significant amount (~4.8 kg) of ceramic plutonium dioxide (PuO2) as a heat source to supply continuous power at approximately 110 watts on the Martian surface. In the MMRTG design the leadetelluride (PbeTe) thermoelectric sources are protected from exposure to the Martian atmosphere, which produces helium due to radioactive decay of the PuO2 as is vented. In the U.S., space missions that involve radioactive materials require presidential launch approval per Presidential Directive/National Security Council Directive 25 (PD/NSC-25, 1977). The Interagency Nuclear Safety Review Panel (INSRP) was formed to conduct an independent safety review and evaluation of the nuclear risk associated with the MSL mission. The INSRP process was established in the 1960s to review safety and assurance issues related to the nuclear space launches sponsored by NASA, the Department of Defense (DOD) and the Department of Energy (DOE) (former Atomic Energy Commission e AEC). The U.S. Environmental Protection Agency (EPA) and U.S. Nuclear Regulatory Commission (NRC) also participate in the INSRP review process. The Advanced Stirling Radioisotope Generator (ASRG) is a next-generation RPSs system. A Stirling engine is used to convert thermal energy to electricity with higher efficiency (~29%) than thermocouples in RTGs. The ASRG utilizes
6.1 Introduction to Space Nuclear Systems
moving parts, produces more electricity than MMRTGs, and requires only onefourth of the plutonium-238 (238Pu) required for RTG operation. The introduction of the SRG design shifted the focus of space agencies to development of a flightready and reliable power system since it can minimize a demand for 238Pu for long-duration space missions (NRC-RPS Committee, 2009). However, the long term reliability of the Stirling design, especially since moving parts are required is, as yet, unproven. Many hours of bench testing have been accomplished and early space missions should reveal the trustworthiness of this more efficient power generator. In the Stirling engine, the conversion is based on oscillation of helium gas at one end of the generator, which is heated by radioactive decay of 238Pu, while the other end is cooled by a heat sink. This oscillating helium gas pushes a piston in a linear alternator that generates AC electricity. Then the AC is converted to direct current (DC) electronically, and the current is fed to the power management system of the spacecraft (NRC-RPS Committee, 2009). During some missions the orbiters and landers will be constantly exposed to the high radiation, which must be diminished. RPSs with static conversion utilize hundreds of thermocouples, which provide a built-in redundancy. That is, failure of a few thermocouples would have a minimal impact on the overall power system performance. In contrast to the RPS, the SRGs are more sensitive to radiation. Therefore, the control systems require a considerable shielding. Electromagnetic radiation (EMR) from the SRG could interfere with scientific instrumentations, while EMR shielding to reduce it would impact system mass. The power systems such as solar arrays are not feasible for the demanding harsh extraterrestrial environment and due to the lack of sunlight and meaningful storage facilities such as batteries. For example, sunlight at Jupiter and Pluto is 96% and 99.94% less intense, respectively, than at Earth. The Juno mission to Jupiter will be powered by solar arrays, but it will be in a highly elliptical polar orbit. It will not orbit near the Jovian equatorial plane where the most intense portions of the belts are located. Thus, it will spend little time in the belts themselves (NRC-RPS Committee, 2009). Currently, RPSs are considered the only existing reliable power source that can work unrestrained in the extraterrestrial environment for the extended periods of time needed in order to accomplish many missions (NRCRPS Committee, 2009). RPSs are compact, rugged and very reliable spacecraft power systems that provide a dependable and long-lived power in space environment. RPSs are not nuclear reactors and they do not use nuclear fission or fusion processes to generate energy. RPSs produce heat through the natural radioactive decay of 238Pu. The generated heat flows through thermocouples to a heat sink generating DC electricity via the conversion process. All of the RPSs flown to date have been RTGs. Since 1961 RPSs have supported 26 NASA and DOD missions and have an outstanding safety and reliability record (see Table 6.3, p. 15). The design of an RPS incorporates many safety features to support continuous operation in the severe space environment. However, the energy conversion efficiency of an RPS is quite low (~6%). The RPS technologies require environmental documentation and safety analysis before approval for use and launch, and
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special accommodations that can impact mission architecture and mission design (Balint, 2007). They must be designed for all mission phases, including launch, cruise, entry, descent, landing and in-situ operations. It is noteworthy to mention that the conditions between these mission phases are different. For example, the heat transfer mechanisms to reject the excess heat created by the radioisotope decay of the plutonium fuel are different for the operations in planetary environments, where heat is rejected through convection, conduction and radiation. During the cruise phase RPSs can be encapsulated inside an aeroshell, while the excess heat is removed by a fluid loop and rejected to space through external radiators. However, for orbiting missions, the RPSs can be exposed directly to the space environment, and heat rejected through radiation directly to space, but the system must still address the pre-launch and launch environments on Earth (Balint, 2007). Plutonium-238 with an 88 year half-life is considered the only practical and convenient isotope for energizing those RPSs for space exploration. The crucial role of RPSs and the need for a supply of 238Pu for U.S. space exploration programs have been proved during space missions such as Viking, Voyager, Cassini and Pluto New Horizons. Unfortunately, 238Pu does not exist in nature. It is created by irradiating Neptunium-237 (237Np) targets in a nuclear reactor; however, 238Pu is unsuitable for use in nuclear detonation weapons. For space power applications, 238Pu is preferable over other isotopes (NRC-RPS Committee, 2009) since it: • • • • •
generates heat for adequate time (radioactive decay half-life of sufficient length); produces a type and quantity of emissions by radioactive decay of fuel that allow it to be handled safely; provides high specific power (heat/mass) and high power density (heat/volume); is a non-corrosive, water-insoluble, and chemically stable fuel form with good engineering properties at high temperatures; can be produced in sufficient quantity at an affordable cost.
Because 238Pu emits alpha particles, RPSs can cause only a biological vulnerability if the 238Pu is somehow released into the environment (then ingested through the food or inhaled through the atmosphere). RPSs are fueled with 238Pu in the form of a ceramic oxide (238PuO2) that has a high melting point and very low solubility to minimize fuel vaporization and transport in the atmosphere, and minimize fuel retention within the human body (NRC-RPS Committee, 2009). The RPSs are designed with multiple fuel containment systems to prevent a fuel release and dispersal of 238Pu into the biosphere during a space mission. (NRC-RPS Committee, 2009). Protective systems are described in the section on “Nuclear Mission Launch Approval”. Only two countries in the world e the U.S. and Russia e have produced 238Pu in the past. Hence, a short supply of the 238Pu may have a significant impact on several pending missions. There are three mission classes for the SSE Program: Discovery e Scout for Mars (small), New Frontiers (medium), and Flagship (large). The projected net balance (supply minus demand) and need for 238Pu to conduct future space operations is shown in Figure 6.4 and Table 6.1 (NRC-RPS Committee, 2009). Due
6.1 Introduction to Space Nuclear Systems
120
Pu Balance = Supply - Demand
100 Pu demand (status quo)
Kilograms of Pu-238
80 60
Pu demand (best case)
40
Pu balance (best case)
20 0 –20
Pu balance (status quo)
The Problem
–40 2008
2010
2012
2014
2016 2018 2020 Calendar year
2022
2024
2026
2028
FIGURE 6.4 Projected
238
Pu demand and net balance (NRC-RPS Committee, 2009).
Table 6.1 NASA’s demand for 238Pu in RPS-involved space missions 238
Pu (kg) 3.5 1.8 24.6 3.5 5.3 14 14 1.8–5.3 3.5 14 5.3–6.2 14 105–110 5.3–5.5
Mission
Launch year
Watts
Mars Science 2011 100 Laboratory Discovery 2014 250 12/Scout Outer Planets 2017 600–850 Flagship 1 Discovery 14 2020 500 New Frontiers 4 2021 800 Pressurized 2022 2000 Rover 1 ATHLETE Rover 2024 2000 New Frontiers 5 2026 250–800 Discovery 16 2026 500 Pressurized 2026 2000 Rover 2 Outer Planets 2027 700-850 Flagship 2 Pressurized 2028 2000 Rover 3 Total demand for Pu (kg) 2009–2028 Annual demand (20-year average in kg/year).
Type of RPS MMRTG ASRG MMRTG ASRG ASRG High Performance SRG High Performance SRG ASRG ASRG High Performance SRG ASRG High Performance SRG
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to low cost caps (~$425M FY06) the Discovery and the Scout class missions are not budgeted to utilize RPSs. Even with the protective containment systems the RPSs may create some risk that 238 Pu could be released into the biosphere. To evaluate this risk, the U.S. has established a flight safety review and launch approval process for RPS-powered missions established by Presidential Directive/National Security Council memorandum-25 (PD/NSC-25, 1977). During this process the DOE prepares a detailed Safety Analysis Report (SAR) to evaluate radiological risk for each mission. The Final Safety Analysis Report (FSAR) is reviewed by the INSRP. The INSRP coordinators conduct an independent safety review and evaluation of the nuclear risk for each RPS involved mission. The INSRP provides technical oversight in six distinct technical Working Groups (WGs): Launch Abort, Re-entry, Power Systems, Meteorology, Biomedical and Environmental Effects, and Risk Integration and Uncertainty. The INSRP then prepares a mission specific Safety Evaluation Report (SER), which identifies, characterizes, and quantifies probable accident scenarios, including the probabilities that 238Pu will be released. NASA uses the FSAR and the SER to request launch approval from the Office of Science and Technology Policy (OSTP), which is within the Executive Office of the President. After reviewing the SER and evaluating the potential risk, OSTP then recommends to the President whether to authorize or deny the requested launch. The President may choose to defer the decision to the Vice President or his Chief Scientist at OSTP. The launch approval process on average takes 3 years (NRC-RPS Committee, 2009). In near future space missions may require building human stations on the Moon and Mars for long-term human attendance and equipment operations. The required electrical power for human stations might be tens of kilowatts for several years. Nowadays nuclear fission power is considered as another promising alternative for outer planet exploration where solar power is not realistic. Nuclear fission power reactors can effectively and safely supply the power required for long-period human and robotic missions to explore the outer planets and targets beyond the Solar System. Nuclear fission reactors can be used to power operating systems of a wide range of magnitudes. Nuclear fission reactors can generate safe, efficient and consistent electric power during the space missions and operations on other planets. According to NASA reports, reactor systems can provide tens of kilowatts of electrical power for electric propulsion and/or operation of science instruments (Scolese, 2002). A nuclear fission reactor can provide significantly more power than ~0.1 kW electricity provided by RTGs. However, in order to reduce mission cost and the amount of fuel, it is enviable to build fission reactor power systems with longer operating lives (more than 20 years) to supply ~40 kWe power for future stations on the Moon and Mars (Schriener & El-Genk, 2011). Several reactor power systems have been designed and tested to provide electric power up to 40 kW for 5e10 years. The Affordable Fission Surface Power System (AFSPS) concept was developed by NASA and DOE for providing electrical power to space stations on Moon and Mars surfaces Mason et al. (2008). This proposed system utilizes a compact fission reactor that operates at
6.1 Introduction to Space Nuclear Systems
temperatures below 1000 K. It is coupled to six pairs of Free Piston Stirling Engines (FPSEs) to generate 40 kW of electric power for 8 full power years (FPY). The FPSE transfers the thermal power to produce AC electricity with a predicted thermal efficiency of 22.85%. The highly enriched (93 wt% 235U) UO2 pellets in the core’s fuel rods are doped with Gd2O3 thermal neutron poison and a clad with 316 stainless steel. The Gd2O3 guarantees that the reactor is safe should it be exposed to moisture. The first, and thus far only, U.S. nuclear fission space power system, SNAP10A, was launched in the 1960s. Project Prometheus, which was NASA’s most recent attempt to develop space nuclear power reactors, selected a nuclear electric propulsion reactor concept that was scalable from 20 kWe to 300 kWe (NRC-RPS Committee, 2009). However, the performance and reliability of space nuclear power reactor systems using current technology remains unproven, especially for missions with long lifetimes (NRC, 2009). Both radioisotope power (via RTG) and fission-based reactor systems are highly reliable technologies with no moving parts. Hence, they can make it possible to provide safer and more dependable thrust generation for space missions than conventional chemical power and propulsion systems. The Sectored Compact Reactor (SCoRe-S) power system concept was developed to provide 100 kW electric power for up to 6 years El-Genk et al. (2005); Hatton & El-Genk, (2006, 2009). The SCoRe-S fission heat source that operates at relatively high temperatures (1150e1250 K) is cooled by circulating liquid lithium El-Genk et al. (2005). The electrical power to the load and to the electromagnetic (EM) pumps for circulating the liquid lithium is provided by the power system which uses SiGe thermoelectric energy conversion elements. The SCoRe-S core (Figure 6.5) is divided into six equal triangular sectors separated by flat, liquid
Rx HPs graphite support disk(0.2 cm)
66.1 cm
Top section of Rx vessel Enclosing gas plenum of UN pins
Control drums actuators
Openspaces for shield cooling
FIGURE 6.5 Sectored Compact Reactor (SCoRe-S).
Rx control actuators Rx lithium heat pipes in multifoils insulation
24.9
Tungsten γ-shield
92.24
15°
59.98
BeO radial
B4C reflector clad in sectors MA-ODS steel
Natural LiH neutrons shield in SS honeycomb
All dimensions in cm
Graphite conduction disk (2.0)
Depleted LiH in SS honeycomb
W-LiH shadow shield clad InMA-ODS steel (0.05)
56 cm
Reactor Li heat pipes In MFT insulation (25 10 μm-thick foils, 100 μm gap)
Steel encasement (0.5 mm) 74.48
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metal heat pipe dividers, are thermally and neutronically coupled, but hydraulically independent El-Genk et al. (2005); El-Genk, (2008). Each sector is thermal-hydraulically coupled to a separate pair of circulating liquid lithium primary and secondary loops with thermoelectric (TE) energy conversion modules and heat rejection radiator panels. The fastest method to drive spacecraft to the target with a lesser amount of propellant is to utilize Nuclear Thermal Propulsion (NTP) systems. A broad development of NTP systems was performed by the U.S. and the Russian Federation in 2007. They evaluated NTP space programs for advanced unmanned missions to the outer planets. Within NASA’s new Vision for Space Exploration Program, Marshall Space Flight Center (MSFC) and Los Alamos National Laboratory (LANL) initiated a development of a new NTP engines. The design process interfaces with the ROCket Engine Transient Simulation (ROCETS) system tool which has been used extensively in the liquid rocket engine industry for many years. ROCETS produces an integrated engine and reactor system design for the specific mission and material limits, at the same time maximizing system thrust to weight ratio Amiri et al. (2007). Today, nuclear fission power is considered as the only alternative for extra-solar exploration. Nuclear fission power reactors can effectively and safely supply the needed energy for long-period human and robotic missions to explore the outer planets and targets beyond the Solar System. Besides propulsion, nuclear fission reactors can be used to power operating systems of a wide range of magnitudes. They can generate safe, efficient and consistent electric power during space missions and operations on other planets. According to NASA, these types of systems can provide tens of kilowatts of electrical power for electric propulsion and operation of science instruments (Scolese, 2002). A nuclear fission power system can provide significantly more power than the ~0.1 kW electrical that RTG systems are currently limited to. Only a single U.S. nuclear fission power system has been launched occurring on April 3, 1965. This System for Nuclear Auxiliary Power (SNAP)-10A functioned for 43 days before it failed, but demonstrated successful operation of the contingency safing plan, placing it into a 4000 year orbit.
Basic Concept e Have Heat, Will Travel If you need constant and reliable power for a long period of time there are several ways to provide that terrestrially; but for long space voyages the criticality of the space vehicle’s power source is a primary concern. For spacecraft missions in the vicinity of the inner planets (Mercury, Venus, Earth, Mars) the standard for power supply has traditionally been based on solar panels, especially if the mission does not call for landing on the planet surface. As seen in the Mars Exploration Rover (MER) missions utilizing the Spirit and Opportunity rovers, solar power is subject to degradation as panels become covered by atmospheric dust (Mazumder, 2003). As related in Figure 6.6, the distance a planet is located in relation to the Sun
6.1 Introduction to Space Nuclear Systems
affects the available solar irradiance (Honsberg, 2010). The total solar radiation, emitted by the Sun is given by Psun ¼ sT 4 Asun ¼ sT 4 4pR2sun ¼ 4:5 1026 Watt
(1)
where: Psun is the power emitted at the Sun’s surface (in Watts) as determined by StefaneBoltzmann’s blackbody equation; s is the StefaneBoltzmann constant ¼ 5.670400 10e8 Jse1me2Ke4 T is 6000 degrees Kelvin Rsun is the radius of the Sun in meters (6.9599 108 m) and the solar radiation intensity at a given distance, D, in meters is given by: H0 ¼
R2sun Hsun D2
(2)
The mean solar irradiance in watts per square meter is provided in Table 6.2 for the planetary bodies in our Solar System.
FIGURE 6.6 Solar System scaled to football field.
Table 6.2 Solar irradiance at planetary distances Planetary body
Distance (3109 m)
Mean solar irradiance (W/m2)
Mercury Venus Earth Mars Jupiter Saturn Uranus Neptune Pluto
58 108 150 228 778 1427 2871 4497 5906
9145 2619 1370 590 51 15 4 2 1
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Noticeable is the drop in mean solar irradiance once the orbit of Jupiter is reached. This significant reduction in available power density beckons the need to provide power for spacecraft applications through other means. Using Seebeck’s discovery, engineers developed SNPSs since extremely reliable heating sources are necessary to guarantee long-duration mission power.
Small to Large Systems Depending upon the needs of the space vehicle mission, the heat source can range from small pellets providing environmental heating only, to larger pellets providing primary heating to thermoelectric conversion units, to large nuclear reactor components providing electrical power conversion more akin to Earthbound systems.
Early to Today Development Historical development of SNPS A good source is Angelo (1985), covering: Polonium Plutonium Uranium LWRHUs, small RTGs, SRGs, MMRTGs, space nuclear reactors, etc.
6.2 SNPS Launch History and Accidents The history of launching spacecraft that have employed radioisotope systems for heat or to generate power began nearly as soon as mankind began to launch rockets. Only two countries however have launched these types of systems.
U.S. Space Nuclear Systems Launch History and Accidents Since 1961, the U.S. has launched 46 nuclear power systems and hundreds of radioisotope heater units (RHUs) on 30 space missions intended for navigation, meteorology, communications, and exploration of the Sun, Moon, Mars, Jupiter, Saturn, and the remainder of our outer Solar System and beyond (Table 6.3). One of these space missions, SNAPSHOT in 1965, employed a 500 We nuclear reactor power system, the remainder employed RPSs, specifically RTGs, and/or RHUs. The objectives of these missions could not have been accomplished without onboard nuclear power systems (Shotlis, 1994), (NRC-RPS Committee, 2009). Currently, there are eight nuclear-powered space missions/spacecraft that are still active; they are: Lincoln Experimental Satellites 8 and 9 (LES 8 and 9), Voyager 1 and 2, Cassini, Mars Exploration Rovers A and B (Spirit and
Table 6.3 U.S. Mission Spacecraft using nuclear power systems, launched 1961–2006 Nuclear system (Num.-nom. pwr)
Mission type
Launch
TRANSIT 4A
Navigational
Jun 1961
TRANSIT 4B
Navigational
Nov 1961
TRANSIT 5BN-1
Navigational
Sep 1963
TRANSIT 5BN-2
Navigational
Dec 1963
TRANSIT 5BN-3
Navigational
Apr 1964
SNAPSHOT
Experimental
Apr 1965
NIMBUS B-1
Meteorological
May 1968
SNAP-19B2 (2-40 We RTGs)
NIMBUS III
Meteorological
Apr 1969
APOLLO 12
Lunar Exploration Lunar Exploration
Nov 1969
SNAP-19B3 (2-40 We RTGs) SNAP-27 (1-70 We RTG) SNAP-27 (1-70 We RTG)
APOLLO 13
Apr 1970
SNAP-3B7 (1-2.7 We RTG) SNAP-3B8 (1-2.7 We RTG) SNAP-9A (1-25 We RTG) SNAP-9A (1-25 We RTG) SNAP-9A (1-25 We RTG) SNAP-10A (1-500 We Reactor)
Status Successfully achieved orbit; ops terminated 1966 Successfully achieved orbit; ops terminated 1967 Successfully achieved orbit; ops terminated 1970 Successfully achieved orbit; ops terminated 1971 Failed to achieve orbit; SNAP-9A RTG burned up on re-entry as then designed/intended Successfully achieved orbit; voltage regulator on S/C bus failed after 43 days of ops; SNAP-10A reactor shutdown automatically & permanently in 3000þ year orbit Launch vehicle intentionally destroyed during early launch phase; SNAP-19B2 RTGs retrieved intact from Santa Barbara Channel; fuel used on subsequent mission Successfully achieved orbit; ops terminated 1979 Successfully placed on Moon; ops terminated 1980 Mission aborted en route to Moon; SNAP-27 survived re-entry, as designed; under 7000þft of water in S. Pacific Tonga Trench (Continued)
6.2 SNPS Launch History and Accidents
Mission
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Mission
Mission type
Launch
APOLLO 14
Lunar Exploration Lunar Exploration Solar System Exploration Lunar Exploration Navigational
Jan 1971
APOLLO 15 PIONEER 10 APOLLO 16 TRANSIT/ TRIAD-01-1X APOLLO 17
Jul 1971 Mar 1972 Mar 1972 Sep 1972
LES 8
Lunar Exploration Solar System Exploration Mars Exploration Mars Exploration Communications
Mar 1976
LES 9
Communications
Mar 1976
VOYAGER 2
Solar System Exploration Solar System Exploration
Aug 1977
PIONEER 11 VIKING 1 VIKING 2
VOYAGER 1
Dec 1972 Apr 1973 Aug 1975 Sep 1975
Sep 1977
Nuclear system (Num.-nom. pwr) SNAP-27 (1-70 We RTG) SNAP-27 (1-70 We RTG) SNAP-19 (4-40 We RTGs) SNAP-27 (1-70 We RTG) TRANSIT-RTG (1-30 We RTG) SNAP-27 (1-70 We RTG) SNAP-19 (4-40 We RTGs) SNAP-19 (2-40 We RTGs) SNAP-19 (2-40 We RTGs) MHW-RTG (2-150 We RTGs) MHW-RTG (2-150 We RTGS) MHW-RTG (3-150 We RTGs) MHW-RTG (3-150 We RTGs)
Status Successfully placed on Moon; ops terminated 1980 Successfully placed on Moon; ops terminated 1980 Successfully placed on interplanetary trajectory; ops terminated 2008 Successfully placed on Moon; ops terminated 1980 Successfully achieved orbit; ops terminated 1977 Successfully placed on Moon; ops terminated 1980 Successfully placed on interplanetary trajectory; ops terminated 2008 Successfully placed on Mars surface; ops terminated 1980 Successfully placed on Mars surface; ops terminated 1980 Successfully achieved orbit; still operational Successfully achieved orbit; still operational Successfully placed on interplanetary trajectory; still operational near heliosheath Successfully placed on interplanetary trajectory; still operational near heliosheath
CHAPTER 6 Nuclear-Powered Payload Safety
Table 6.3 U.S. Mission Spacecraft using nuclear power systems, launched 1961–2006dContinued
Jovian Exploration
Oct 1989
ULYSSES
Solar Polar Exploration Mars Rover Exploration Saturnian Exploration
Oct 1990
Mars Rover Exploration Mars Rover Exploration Pluto & Kuiper Belt Exploration
Jun 2003
MARS PATHFINDER CASSINI/ HUYGENS
MER-A MER-B PLUTO NEW HORIZONS
Dec 1996 Oct 1997
Jul 2003 Jan 2006
GPHS-RTG (2-275 We RTGs) & LWRHU (120-1 Wt RHUs) GPHS-RTG (1-275 We RTG) LWRHU (3-1 Wt RHUs) GPHS-RTG (3-275 We RTGs) & LWRHU (117-1 Wt RHUs) LWRHU (8-1 Wt RHUs) LWRHU (8-1 Wt RHUs) GPHS-RTG (1-275 We RTG
Successfully placed in orbit around Jupiter; intentionally de-orbited into atmosphere of Jupiter in 2004 to ensure future protection of Europa
Successfully placed on solar polar orbit; ops terminated 2010 Successfully placed on Mars surface; ops ceased 1997 Successfully placed in orbit around Saturn; still operational
Successfully placed on Mars surface; still operational Successfully placed on Mars surface; still operational En-route to Pluto w/flyby in 2015; still operational
6.2 SNPS Launch History and Accidents
GALILEO
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Opportunity), and Pluto New Horizons. In the immediate future, the MSL mission/spacecraft with a rover (Curiosity) powered completely by the MultiMission RTG (MMRTG) is scheduled for launch in late 2011. Nuclear power systems are expected to continue to enable and support further robotic exploration of our Solar System and beyond in the future. Moreover, nuclear systems for power and propulsion will be required to enable future human missions to the Moon, Mars, and asteroids as we expand our presence in space. Table 6.3 provides a listing of all known U.S. launches of payloads that utilized some form of radioisotope. Launch, orbital and final trajectory insertion, and space flight involve risks of failure. Some failures can present severe accident environments to onboard systems, including nuclear systems. In general, the most critical mission phases are launch and ascent, when large quantities of propellants are present. Accidents during launch and ascent can threaten an onboard nuclear system from blast overpressure, small shrapnel and large fragment impacts, Earth surface impact, propellant fires, and falling debris. In addition to accidents during launch and ascent, malfunctions during orbital or final trajectory insertion can result in an inadvertent re-entry, with its associated re-entry heating, ablation, and thermo-mechanical shock environments, followed by Earth surface impact. For nuclear systems that require active cooling (e.g., space reactors once operated in space), internal malfunctions during operation in space that can result in overpower or under-cooling conditions can also threaten the nuclear system’s integrity. Overall, the probability of catastrophic launch vehicle and upper stage accidents is in the range of a few percent. Consequently, space nuclear systems must be designed to respond to such accidents without presenting undue risk to the Earth’s population or environment (Shotlis, 1994). Four U.S. space missions with nuclear systems onboard have experienced failures; three were caused by the launch vehicle or upper stage. In each case, built-in safety features performed as designed and there were no adverse radiological consequences. A chronological discussion of these events follows (Shotlis, 1994). The first failure occurred on April 21, 1964, when the TRANSIT 5BN-3 navigational satellite failed to achieve Earth orbit because a computer malfunction prematurely shut down an upper stage booster. The satellite and its SNAP-9A RTG power supply re-entered the Earth’s atmosphere and burned up completely e as early RTGs were designed to do e at an altitude of ~50 km. Approximately 20,000 Curies of 238Pu were released into the upper atmosphere and dispersed worldwide. Although this accident did not pose a health threat to any member of the population, it did involve an uncontrolled release of radioactivity into the biosphere. Consequently, the design requirement for RPSs and RHUs under accidental re-entry was changed from complete breakup and dispersal at high altitude to survival intact, that is, with fuel containment and confinement preserved through re-entry. In May 1965, 43 days after the successful launch and deployment of the SNAPSHOT experimental spacecraft, powered by a SNAP-10A reactor, a voltage
6.2 SNPS Launch History and Accidents
regulator failure on the spacecraft bus caused automatic and permanent (i.e., irreversible) shutdown of the reactor. No deleterious consequences were involved as a result of this failure. The reactor currently remains in a 3000þ year Earth orbit. When it does re-enter the Earth’s atmosphere in the distant future, the reactor will contain the same insignificant amount of radioactivity as it had when it was launched. That is, the fuel will again be essentially radioactively cold and clean. On May 18, 1968, approximately one minute into the launch of the SNAP19B2 powered NIMBUS B-1 meteorological satellite, the Range Safety Officer at Vandenberg Air Force Base, CA, destroyed the launch vehicle by command destruct action, so its errant launch trajectory would not place the public in danger. Although the launch vehicle, upper stage, and spacecraft were completely destroyed, the two on-board SNAP-19B2 RTGs survived intact, with no release of their radioactive 238Pu fuel. The RTGs were retrieved from the Santa Barbara channel, and their nuclear fuel was used on a subsequent space mission. Lastly, in April 1970, the APOLLO 13 mission was aborted en-route to the Moon because of an explosion of an oxygen tank in the APOLLO 13 Service Module. A SNAP-27 RTG was on the APOLLO 13 Lunar Lander to power a lunar surface experiment package. Because the Lunar Lander returned to Earth with the Crew Re-entry Module, it and the SNAP-27 RTG experienced re-entry into and through the Earth’s atmosphere. The SNAP-27 RTG survived re-entry intact, with no release of its radioactive 238Pu fuel, and sank to a depth in excess of 7000 feet at the bottom of the Tonga trench in the South Pacific, where it remains. Although failures/accidents have occurred, they were anticipated and specifically accounted for in the design of the onboard space nuclear systems, to prevent harmful radiological consequences. The historical record, documented in Table 6.3, including the failures/accidents discussed above, serves to provide confidence that the space nuclear safety program in the U.S. has been and continues to be effective as well as vital.
Soviet/Russian Space Nuclear Systems Launch History and Accidents The Soviet Union/Russia has launched 40 space nuclear systems in support of 40 space missions during the period September 3, 1965, and November 16, 1996. Five of those systems/missions involved RTGs, two involved RHUs, while the remaining 33 involved space reactors (Bennet, 1989; Ponomarev-Stepnoi et al., 2000), and (Sholtis et al., 2008). The first two Soviet nuclear-powered space missions, Cosmos-84 and Cosmos90, launched on September 3, 1965 and September 18, 1965, respectively, each used a 20 We RTG fueled with polonium-210. Two other RTG-powered LUNOKHOD spacecraft were launched on Cosmos-300 and Cosmos-305, on
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September 23, 1969 and October 22, 1969, respectively. Both of these LUNOKHOD spacecraft failed to achieve Earth orbit and the onboard RTGs, each fueled with polonium-210, re-entered the Earth’s atmosphere and burned up, as designed, on September 27, 1969 and October 24, 1969, respectively. Lastly, among the Soviet/Russian RTG-powered space missions, was the MARS 96 mission, launched on November 16, 1996, with a 0.2 We plutonium-238 fueled RTG. The MARS 96 spacecraft failed to achieve Earth escape velocity, and reentered the Earth’s atmosphere on November 17, 1996, falling into the South Pacific off the coast of Chile and Bolivia. This RTG was designed to survive reentry, and no radioactivity was detected from the re-entry or impact of the MARS 96 onboard RTG. Two Soviet space missions, Luna 17 and Luna 21, launched on November 10, 1970, and January 8, 1973, respectively, each used a 1.0 kWt RHU, fueled with polonium-210. These RHUs apparently served to provide heat for the batteries and electronics of the Soviet MOONCAR-1 and MOONCAR-2 lunar vehicles (Sholtis, Marshall, Bennett, Brown, Usov and Dawson, 2008). The remaining 33 Soviet space nuclear systems/missions used space reactors. Thirty-one of these systems were 3.0 kWe reactor systems known as Buk (bul, meaning “beech” in Russian) reactors. These Buk space reactors, which had a uranium-235 fuel loading of 30 kg, produced 100 kWt, and used thermoelectrics to convert this heat into 3.0 kWe. The total Buk space reactor power system mass was 930 kg. The Soviet Buk space reactor powered their Radar Ocean Reconnaissance Satellites (RORSATs), launched during the period October 3, 1970, and March 14, 1988 (Sholtis, Marshall, Bennett, Brown, Usov & Dawson, 2008). Two other Soviet reactor-powered space missions, Cosmos-1818 and Cosmos1867, launched on February 1, 1987, and July 10, 1987, respectively, each used a different space reactor known as TOPAZ. The Soviet TOPAZ space reactor incorporates in-core thermionics to convert the 150 kWt generated by the reactor directly into 5.0 kWe. The uranium-235 fuel loading of the TOPAZ reactor is 11.5 kg, and the total mass of the TOPAZ space reactor power system is 980 kg (Sholtis, Marshall, Bennett, Brown, Usov & Dawson, 2008). Three RORSAT missions (Cosmos-954, Cosmos-1402, and Cosmos-1900), each with a Buk space reactor onboard, experienced accidents. The first accident involved Cosmos-954, which was launched on September 18, 1977. The Cosmos-954 RORSAT operated in low Earth orbit for 43 days. At the end of its operational mission, the Buk reactor successfully separated from the spacecraft, but failed to boost to its intended higher, long-lived orbit. Consequently, the reactor re-entered the Earth’s atmosphere and impacted the Earth’s surface near Great Slave Lake in the Northwest Territories of Canada on January 24, 1978. The Buk reactor was designed to burnup on re-entry, but re-entry burnup of the reactor was not complete. Approximately 65 kg of radioactive debris and nuclear
6.3 Launch Abort Environments Affecting SNPSs
fuel was recovered from the crash site. It is very fortunate, indeed, that this impact occurred in an unpopulated area. The second accident involved Cosmos-1402, which was launched on August 30, 1982. During this RORSAT mission, an anomaly occurred and the spacecraft automatically separated into three parts. The Buk reactor is believed to have re-entered over the South Atlantic, approximately 1600 km east of Brazil on February 7, 1983. No radioactivity was detected from this re-entry or impact. This suggests that the design of the Buk space reactor was modified following the Cosmos-954 accident, to ensure its intact re-entry. The last Soviet space nuclear accident involved Cosmos-1900, which was launched on December 12, 1987. On May 13, 1988, the Soviet Union reported that it had lost communications with the Cosmos-1900 RORSAT in April, 1988, and that as a result the onboard Buk reactor could not be boosted to a higher, long-lived orbit via a ground command signal. At the same time, the Soviet Union reported that there was an automatic separation and boost capability built into the spacecraft. Subsequently, on September 30, 1988, the Buk reactor automatically separated from the spacecraft and boosted to a higher (~720 km) orbit. This event, although it does involve a malfunction, does not represent an accident, in the strict sense, since the situation was resolved by an onboard capability designed to automatically operate for such a malfunction.
6.3 Launch Abort Environments Affecting SNPSs A Space Nuclear Power System designed and tested for a specific mission (e.g., deep space, planetary) has the additional constraint of surviving transport from Earth to its mission location. This requires that it has to remain undamaged not only during the normal elevated vibration and shock environments of launch, but it also must be robust enough in its containment design to survive a worst case launch vehicle accident. These types of accidents normally include extreme shock, vibration, blast, fragmentation, overpressure, and thermal environments which expose the payload to their damaging effects. The following sections introduce the reader to the hazardous environment that could exist around the nuclear source during a launch abort accident. By evaluating the potential high temperature and shock conditions provided to the source in such a scenario, the design requirements for protecting the nuclear material can be established. Methodologies to handle the additional difficulty of predicting the potential for a launch failure concludes this section.
Pre-Launch and Launch Abort Scenarios and Issues The safety design and testing of nuclear material used in space power systems is predicated on survival of its containment in one of the most destructive environments
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FIGURE 6.7 Delta II GPS-IIR launch abort, January 17, 1977.
imaginable: the pressure and heat of an explosion. While the likelihood of a rocket to “spontaneously combust” is usually low, especially in more mature vehicles, the surprise attack has been demonstrated, as was the case on January 17, 1997. Figure 6.7 depicts the 241st launch of a Delta II rocket (the 35th Delta 7925 model) 12.5 seconds into flight. Critical to the design of a containment system for nuclear material is a thorough understanding of the potential explosive environment through historical and state-of-the-art characterization of launch vehicle fires and explosions. Characterization begins with the type of launch vehicle used to transport the space nuclear power system. A brief description of typical rockets and their explosive potentials leads to discussions on the physics of rocket propellant explosions, the blast overpressure hazard, fragments induced by these explosions, and the possible thermal fireball environment subjected upon the payload. Just how likely is it that a launch failure will occur? A description of probabilistic studies of these events is included.
Launch Vehicle Types While there have been a plethora of new launch vehicles arriving on the scene, the basic propulsive technology applied to these rockets dates back to the 1960s era. It seems the methods for propelling spacecraft into orbit over the last eight decades has remained consistent and provides for two fundamental elements e liquid propellant and solid propellant. What follows is a brief description of these propellant types, their applications in launch vehicles, and some typical quantities found on current rockets. In 1926, when Robert Goddard (Figure 6.8) launched his rockets, the technology to propel objects through the air was already well understood by the Chinese who in the 13th century had successfully lofted destructive missiles at their enemies using
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.8 Robert Goddard’s liquid fueled rocket, circa 1926.
what are today know as fireworks (Figure 6.9). While the Chinese propulsion element of choice consisted mainly of gunpowder (or solid propellant), Goddard, after improving on their designs, turned to liquid propellant. Both have their advantages and disadvantages, some of which are shown in Table 6.4. All of the rockets launched today use one or both of these forms of propellant in stages to reach their destination. New designs may combine both elements and are called hybrids. A combined form of gel propellant looks promising for bringing together benefits of both types. Table 6.5 lists some of the current launch vehicle inventory and the types and amounts of propellant utilized by each. This data will be useful in calculating the explosive environment later in this chapter.
Threat environment A very important aspect of the propellants used is their ability to impart energy to other parts of the launch vehicle during uncontrolled burning or explosive events,
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FIGURE 6.9 Chinese fire arrows.
particularly hazardous systems such as RTGs, which could in turn be demolished causing release of harmful radioactive material. The section on “Containment Design” describes the vulnerabilities of the nuclear source material such as clad melting temperatures, overpressure limits, and effects of vaporization. Predicting the environment that this material must potentially survive is the objective of the rest of this section.
Propellant Fires Regardless of the propellant combination (solid, liquid, gel), if the flame directed out the nozzle to provide thrust escapes the confines of the combustion chamber or exit, due to leaking or burn-through, the result will likely be a propellant fire with subsequent explosion once lower explosive limits are reached (Figure 6.10). Hypergolic propellants (e.g., hydrazine fuel and nitrogen tetroxide oxidizer) are designed to combust, without an ignition source, upon contact. Cracking or delamination of solid propellant can provide pathways for flame to travel to the outer sidewalls of the booster causing torch-like burn-through, which can impinge on
Table 6.4 Propellant pros and cons
Solid propellant
Liquid propellant
Advantages
Disadvantages
Storable Non-toxic (in storage) Simplified systems Throttleable Higher Isp
No stop/start capability Explosive impacts Lower Isp Toxic (N2H4/N2O4) Cryo storage (LOX/LH2) More complex systems
Table 6.5 Launch vehicle properties
Vehicle Delta 7425
Stg 0
4 Ground Lit GEMs
1
Delta III
Delta IV (medium)
Solid quantity lb
S
HTPB
25,942 ea
L
STAR48 6 Gnd Lit, 3 Air Lit GEMs
1 2 3 0
Solid prop type
L S S
L S S
Liquid ox quantity lb [gal]
Liquid prop fuel
Liquid fuel quantity lb [gal]
LOX
146,070 [15,940] 8712 [720]
RP-1
66,504 [9597] 5250 [700]
N2O4 HTPB HTPB
LOX N2O4 HTPB HTPB
A-50
4405 25,942 ea
L
STAR48 6 Gnd Lit, 3 Air Lit GEMs
Liquid prop oxidizer
146,070 [15,940] 8712 [720]
RP-1
146,070 [15,940] 31,179 [3402] n/a
RP-1
A-50
66,504 [9597] 5250 [700]
4405 37,500 ea
1
L
LOX
2
L
LOX cat bed
1
L
LOX
2
L
LOX
371,321 [40,520] 38,571 [4209]
LH2 N 2H 4 LH2 LH2
66,504 [9597] 5788 [10,101] 160 [19] 68,679 [119,859] 6429 [11,220] (Continued)
6.3 Launch Abort Environments Affecting SNPSs
Delta 7925
2 3 0
Solid/ Liquid
279
280
Vehicle Delta IV (medium þ)
Delta IV (heavy)
Atlas IIA
Atlas IIAS
Stg 0
2 Gnd Lit/ 2 Gnd þ 2 Air Lit GEMs
Solid prop type
Solid quantity lb
S
HTPB
41,990 non TVC 42,500 TVC ea
Liquid prop oxidizer
Liquid ox quantity lb [gal]
Liquid prop fuel
Liquid fuel quantity lb [gal]
371,321 [40,520] 38,571 [4209]
LH2
1,113,963 [121,560] 38,571 [4209]
LH2
242,144 [25,489] 31,350 [3290] n/a
RP-1
68,679 [119,859] 6429 [11,220] 206,037 [359,577] 6429 [11,220] 107,616 [15,826] 5900 [10,000] 170 [20]
242,144 [25,489] 31,350 [3290] n/a
RP-1
289,289 [31,568]
RP-1
1
L
LOX
2
L
LOX
L
LOX
2
L
LOX
1
L
LOX LOX cat bed
1
3 core boosters
2
Centaur
L
0
2 Gnd Lit 2 Air Lit Castor IV-As
S
1 2
Atlas III
Solid/ Liquid
1
Centaur
HTPB
LH2
LH2
LH2 N2H4
22,300 ea
L
LOX
L
LOX cat bed
L
LOX
LH2 N2H4
107,616 [15,826] 5900 [10,000] 170 [20] 112,072 [16,172]
CHAPTER 6 Nuclear-Powered Payload Safety
Table 6.5 Launch vehicle propertiesdContinued
Atlas V
Titan IVB
Centaur III
L
0 1
0 to 5 SRBs
S L
2
Centaur III
2
Centaur III (stretched)
31,616 [3450]
LH2
5990 [10,454]
LOX
452,906 [49,422] 31,350 [3290] n/a 38,077 [4155] n/a 1,358,824 [117,140] 28,324 [2442] 5554 [479]
RP-1
Jet-A
174,194 [25,136] 5900 [10,000] 170 [20] 6923 [12,082] 170 [20] 181,176 [26,801] 3776 [559]
Jet-A
741 [110]
A-50
117,240 [12,540] 27,570 [2950] 123 [14.6]
102,000 ea
LOX cat bed L
LOX cat bed
1
L
H2O2
2
L
H2O2
3 0 1
2 Gnd Lit SRMUs
2
STS
HTPB
LOX
L S L
H2O2 HTPB
IUS
S
N2O4
3
Centaur
L
0
2 Gnd Lit SRBs
S
PBAN
Jet-A
cat bed
222,760 [18,410] 49,630 [4100] n/a
N 2H 4
LOX cat bed
38,260 [4000] n/a
LH2 N 2H 4
n/a
n/a
N 2H 4 APU
N2O4 HTPB
LH2 N 2H 4
695,427 ea
L
3
LH2 N 2H 4
SRM-1 21,404 SRM-2 6002
1,287,188 ea
A-50
7940 [13,400] 170 [20] 67 [8]
6.3 Launch Abort Environments Affecting SNPSs
Beal BA-2
2
(Continued)
281
282
Vehicle
Pegasus Proton D-1-e
Zenit 3SL
H-2A
Stg
Solid/ Liquid
Solid prop type
Solid quantity lb
Liquid prop oxidizer
Liquid ox quantity lb [gal]
Liquid prop fuel
Liquid fuel quantity lb [gal]
LOX
1,361,936 [143,351] 1549 [128]
LH2
227,641 [385,265] 946 [129]
ET
L
Orb 1 2 1
L S S L
2
L
N 2O 4
3
L
N 2O 4
N2O 4 HTPB HTPB
N2O 4
4
DM-5L
L
LOX
4
or Breeze
L
N2O 4
1
L
LOX
2
L
LOX LOX
3
DM
L
0 0
2 Gnd Lit SRB-A 2 Solid Boosters (opt.) 1-2 Liq Boosters (opt.)
S S
0
L
HTPB HTPB
MMH
26,790 6670 659,014 [54,464] 276,893 [22,884] 82,561 [6823] 23,369 [2460] 22,120 [1828] 511,044 [55,766] 129,278 [14,107] 23,369 [2460]
UDMH
187,039 [20,410]
LH2
UDMH UDMH SYNTIN UDMH RP-1 RP-1 SYNTIN
244,986 [37,119] 67,207 [10,183] 20,039 [3036] 9480 [1335] 11,060 [1676] 196,556 [28,363] 49,722 [7175] 9480 [1335]
133,400 ea 28,900 ea LOX
31,701 [55,325]
CHAPTER 6 Nuclear-Powered Payload Safety
Table 6.5 Launch vehicle propertiesdContinued
Ariane IV
Long March 3B
L
LOX
2
L
LOX
0 0
0-4 Gnd Lit SRBs 0-4 Liq Boosters (opt.)
S L
CTPB
N2O4
L
N2O4
2
L
N2O4
3
L
LOX
2 Gnd Lit SRBs
2 0
Aestus 4 Liquid boosters
1 2 3
S L
HTPB
LH2 LH2
31,957 [55,771] 6100 [10,646]
20,900 ea
1
0 1
188,543 [20,574] 30,500 [3328]
54,148 [6435] 323,630 [26,746] 47,876 [3957] 19,697 [2149]
UH25
287,280 [31,349] 14,384 [1189] 82,908 total ea 379,260 total 108,045 total 39,690 total
LH2
UH25 UH25 LH2
31,852 [4550] 190,370 [27,196] 28,162 [4023] 4103 [7161]
506,000 ea LOX
L L
N2O4 N2O4
L L L
N2O4 N2O4 LOX
MMH UDMH
54,720 [95,497] 7016 [957] unavailable
UDMH UDMH LH2
unavailable unavailable unavailable
6.3 Launch Abort Environments Affecting SNPSs
Ariane V
1
283
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CHAPTER 6 Nuclear-Powered Payload Safety
FIGURE 6.10 Titan I explosion, December 12, 1959.
secondary combustible systems (e.g., the Challenger accident). Should the combustible materials reach conditions for explosion, a fireball will be created. If the leaking propellant settles in a particular area on the ground and ignites (spontaneously or via a source) a pool fire is the result. Some historical launch pad explosions are shown in Figures 6.10e6.15. On- or near-pad accidents (altitude 170 ft) are assumed to be of such severity (major structural failures) that all of the onboard liquid propellants (except possibly the spacecraft’s) are released and are thus available for bulk pad fires (Marietta, 1996). During the propellant loading phase (pre-launch), the total propellant weight would be consistent with what had been currently loaded into the launch vehicle. For tanks being topped off, the mission specific weight would be directly proportional to the elapsed time since the start of loading, assuming a constant pumping rate. In addition, if destruct action occurred (either by automatic or commanded action) after the solid rockets have been ignited, solid propellant ground fires will be present, although considerably dispersed by the fragmentation process.
6.3 Launch Abort Environments Affecting SNPSs
(A)
(B)
FIGURE 6.11 (A) Juno II going horizontal and (B) exploding, July 16, 1959.
FIGURE 6.12 Titan I fireball sequence, December 12, 1959.
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FIGURE 6.13 Titan I fireball sequence, August 14, 1959.
FIGURE 6.14 Vanguard fireball sequence, December 6, 1957.
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.15 Convair Atlas.
Increasing flight speed and altitude (170 ft h 3800 ft) will usually disperse the liquid propellants. Consequently, the reduced concentration of mixed propellant that falls back will greatly reduce any fireball formation at the ground. Airborne reactions will form more of a combustion cloud (like the Titan 34D-9, see Figure 6.16) than a ground fireball. Solid propellant ground fires resulting from fragmentation of the boosters will be more dispersed than for on- or nearpad accidents. During the period from land clear to shortly after payload fairing jettison, high flight dynamic pressure or rapidly decreasing ambient pressure will further increase the liquid propellant dispersion and reduce the intensity of airborne combustion. The flight path is completely over water so there is no ground fire hazard from either the liquid or solid propellants. Once at high altitude (~475,000 ft), the only thermal environment hazard to the payload’s nuclear material is from accidental launch vehicle re-entry. Aeroloading and heating, with subsequent breakup during fall back from this altitude regime, are covered in the re-entry section.
Liquid propellant fireballs A conceptual model of a fireball is shown in Figure 6.17. Combusting propellants embedded within the fireball are the source of the combustion products (Dobranich, 1997). The physics of fireball formation is thoroughly described in Sandia Report SAND97-1585 (Dobranich, 1997), which explains their Fireball
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CHAPTER 6 Nuclear-Powered Payload Safety
FIGURE 6.16 Titan 34D-9 explosion at Tþ9 seconds.
Integrated Code Package simulating responses of physical and chemical processes occurring in a fireball and is presented here:
Fireball temperature Based on a single control volume, the time-dependent temperature of the fireball is determined by solving the following energy balance equation (assuming a constantpressure combustion process): d nf h f ¼ RLþE (3) dt where nf is the quantity of combustion products and entrained air that comprise the fireball (mol), hf is the molar enthalpy of the fireball (J/mol), t is time (sec), R is the reactant enthalpy inflow (W), L is the fireball energy loss rate (W), and E is the enthalpy inflow from air entrainment (W). Note that both nf and hf are dependent
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.17 Conceptual fireball model.
variables. The molar gas quantity in the fireball is: nf ¼ np þ na
(4)
where np is the quantity of combustion products (mol), and na is the quantity of entrained air (mol). The reactant enthalpy inflow term, R, is given by: R ¼ n_ r hr
(5)
where n_ r is the molar combustion rate of propellant reactants (mols/s), which is assumed known, and hr is the specified enthalpy of the reactants (J/mol). Likewise, the air entrainment enthalpy inflow term is given by:
E ¼ n_ a ha
(6)
where n_ a is the molar rate of air entrainment (mol/s), and ha is the enthalpy of the ambient air (J/mol). Equation (6) can be expanded using the chain rule and rearranged to provide dhf ¼ n_ r hr n_ p hf L þ n_ a ha hf nf dt
(7)
dnp is the molar product production rate (mol/s). dt The product production rate, n_ p , is expressed in terms of the reactant combustion rate using: dnr ¼ yn_ r (8) n_ p ¼ y dt where n_ p ¼
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where y is the molar quantity of combustion products produced per mole of reactant, and n_ r is the specified reactant combustion rate (mol/s). This equation is required for the solution of eq. (8). Based on equilibrium chemistry, the fraction y is determined using: y ¼
np Wr ¼ nr Wp
(9)
where np is the quantity of products (mol), nr is the quantity of reactants (mol), and Wr and Wp are the mean molecular weights of the reactants and products, respectively (g/mol). The mean molecular weights are given by: Wr ¼
Nr X
y jr Wrj
(10)
yip Wpi
(11)
j¼1
and Wp ¼
Np X i¼1
where Nr is the number of reactant species, Np is the number of product species, yrj and yip are the mole fractions of reactant species j and product species I, respectively, and Wrj and Wpi are the molecular weights of the reactant and product species, respectively (g/mol). Equation (7) can now be rearranged to provide: dhf ¼ n_ r hr yhf L þ n_ a ha hf nf dt
(12)
This equation is based on the following definition of enthalpy: h ¼ hF þ Dh
(13)
where hF is the enthalpy of formation (also known as the heat of formation) at the standard state (298.15 K, 1 atm), and Dh is the change in enthalpy from the standard state to some other state, including sensible heat and heats of transition when appropriate. The loss term in eq. (7), L, is given by: 4 þ hAl AAl ðT TAl Þ qAl L ¼ ε s A T 4 Ta4 þ εAl ε s AAl T 4 TAl
(14)
where ε is the effective emissivity of the fireball, s is the StefaneBoltzmann constant (5.67 10e12 W/cm2,K4), A is the surface area of the spherical fireball (cm2), T is the temperature of the fireball (K), Ta is the ambient environment temperature (K), εAl is the surface emissivity of the aluminum structures within the fireball, TAl is the surface temperature of the aluminum structures (K), hAl is the heat transfer coefficient for convective heat transfer between the aluminum structures and
6.3 Launch Abort Environments Affecting SNPSs
the fireball (W/cm2,K), and qAl is the heat added to the fireball from aluminum combustion (W). Calculation of the aluminum surface temperature and the combustion source term are not described in this chapter. Qualitatively, eq. (3) indicates that the time rate of change of the fireball energy equals the rate of energy inflow from reactants and air entrainment minus the rate of fireball energy loss. The loss term in eq. (14) includes four contributions: (1) thermal radiation from the fireball to the ambient environment; (2) thermal radiation from the fireball interior (treated as gray semitransparent participating medium) to the aluminum structures (the rocket) within the fireball; (3) convective heat transfer from the fireball interior to the aluminum structures, and (4) heat added to the fireball from combustion of aluminum structures. Following liftoff, only the first loss term contribution is included because the structures are assumed to no longer be immersed in the rising fireball. Using the calculated fireball enthalpy and product mole fractions, the fireball temperature is determined using: hf ¼ ð1 ya Þ
Np X
yip hip þ ya ha
(15)
i¼1
where ya is the molar air fraction defined as the molar quantity of entrained air in the fireball divided by the molar quantity of air and all combustion products in the fireball, yip is the mole fraction of product species i, hip is the enthalpy of product species i (J/mol), and ha is the enthalpy of the air entrained into the fireball (J/mol). The molar air fraction is defined as: na ya ¼ (16) nf Because hip and ha are functions of fireball temperature T, eq. (15) must be solved iteratively for temperature. This is accomplished using Newton iteration, recognizing that the derivative of enthalpy with respect to temperature at constant pressure is equal to cp, the constant-pressure specific heat (J/mol,K). Enthalpies are provided as fifth-order polynomials in temperature and account for heat of formation, sensible heat, phase transitions, and corrections for non-ideal gas behavior. Specific heats are provided as fourth-order polynomials in temperature.
Empirical formulae Roberts (1982) used the empirical data of Hasegawa and Sato (1977) to correlate the measured radiation flux q received by a detector at a distance L (m) from the center of a fireball with a hydrocarbon fuel mass (AICE, 1994), mf (kg): q ¼ 828 m0:771 L2 f
(17)
Heat flux density thermochemical model predictions are shown along with upper bound experimental data, plotted versus dimensionless time, t/tb, in Figure 6.18. Temperature within the fireball can be estimated from the heat flux data by assuming
291
CHAPTER 6 Nuclear-Powered Payload Safety
440
5000
3064 396
352 4500 308
264
MO D
220
P EL
2898
C DI RE TI
P EX
DA
FIREBALL STEM LIFTOFF (1.5 × tb)
176
2803
O
.
N FO RL OX/L H 8 LO EX X/RP-1 (U.B .) P. D ATA FOR H YPERGOLS (U.B.) TA
4000
HEAT FLUX (kW/m2)
TH ER MO CH EM IC AL
FIREBALL SPHERE LIFTOFF (tb)
HEAT FLUX (BTU/sec/ft2)
2985
TEMPERATURE (K)
292
132 3500 88
44
0
0 0
0.5 1 1.5 NONDIMENSIONAL TIME (t/tb)
3000
2
FIGURE 6.18 Liquid propellant heat flux versus non-dimensional time.
that radiation effects dominate and the emissivity is 1.0 (blackbody). Consider the initial radiant flux of 2000 kW/m2. This corresponds to a temperature of 2433 K via the StefaneBoltzmann relation: Eb ¼ s T 4
(18)
where Eb is the blackbody emissive power, or heat flux (kW/m ), s is the Stefane Boltzman constant, and T is the fireball temperature (K). The maximum fireball temperatures are higher than this value, but rapidly drop down within milliseconds. 2
Fireball rise time Failure of the propellant containment system can lead to a subsequent explosion that can be located near the ground, during early flight, or in the air, during later flight. In
6.3 Launch Abort Environments Affecting SNPSs
surface fireballs associated with propellant explosions, as the pressure of the detonation products decreases to ambient, the density of the gas is considerably below ambient density and the resultant buoyant force causes the gases to begin to rise. The bulk of the propellant is then entrained in the fireball and rapidly burned (High, 2006). A hemispherical shape is maintained during most of this period until the fireball growth rate is exceeded by the buoyancy and a spherical shape develops. After the spherical shape is completely formed, the fireball lifts from the ground, and surrounding air is drawn into the fireball by a combination of convective forces and a vortex motion to continue mass addition. A stem is created with propellant spilled on the ground to produce a mushroom-type cloud. The hot fireball continues to change into an oblate spheroid and finally a torus, as seen in Figure 6.19. Combustion between the fuel-rich gases and entrained air dilutes and cools the gases. The radiative losses also contribute to a rapid cooling process. The dynamic behavior of the fireball can be described as a spherical thermal composed of hot fireball gases on top and entrained air below. The development and rise of this thermal are
FIGURE 6.19 Fireball rise model.
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CHAPTER 6 Nuclear-Powered Payload Safety
controlled by thermal radiation, addition and combustion of propellant, and air entrainment. Following completion of the combustion stage, the fireball lifts from the ground, rapidly entraining air. The molar rate of air entrainment, based on an assumed entrainment coefficient, is determined using: n_ a ¼
auAp RT
(19)
where a is an entrainment coefficient, u is the rise velocity of the fireball (cm/s), A is the surface area of the fireball (cm2), p is the pressure of the fireball (atm), which is assumed to equal the ambient pressure, R is the universal gas constant (82 cm3,atm/ mol,K), and T is the fireball temperature (K). The quantity of entrained air is added to the existing product species in the calculation of fireball temperature via eq. (15). Also, during the entrainment stage, the reactant enthalpy inflow term is zero. The air entrainment model is based on the “macro-scale” process of a bubble (the fireball) rising through a medium (air), which is applicable after fireball liftoff. Such a model is common in bubble flow and atmospheric simulations. However, the model is also used before fireball liftoff, when air entrainment is a “micro-scale” process in which turbulent eddies near the boundaries of the fireball bring in ambient air. Because these turbulent eddies are not captured in the fireball model, there is no reasonable way to predict this entrainment. The micro-scale entrainment is expected to be much smaller than the macro-scale entrainment. Different entrainment coefficients may be used during the combustion and entrainment stages to reflect the different processes (e.g., a ¼ 0.025 during the combustion stage and 0.25 during the entrainment stage). It is also possible to introduce entrained air by specifying it as a reactant; thus micro-scale entrainment calculated by some other means (such as a computational fluid dynamics code) can be incorporated in the fireball simulation if desired. The rise velocity of the fireball is determined from solution of the following momentum balance: Cd d r A P u2 ðrVuÞ ¼ gV ra r 2 a dt
(20)
where r is the density of the fireball (g/cm3), V is the volume of the fireball (cm3), g is the acceleration due to gravity (981 cm/s2), Cd is the drag coefficient, ra is the density of the ambient air (g/cm3), and AP is the projected area of the fireball (cm2). The first term to the right of the equal sign is the buoyant force due to the density difference of the ambient air and the fireball, and the second term is the drag force. Application of the chain rule and rearrangement yields the following equation for numerical solution: du ¼ c2 c1 u c3 u2þb dt
(21)
6.3 Launch Abort Environments Affecting SNPSs
with
" ln 1 dðrVÞ d lnðrVÞ ¼ z c1 ¼ rV dt dt r c2 ¼ g a 1 r
kþ1 # nf W f k nf W f Dt
(22) (23)
and c3 ¼
1þb b bþ2 p að2Þb1 ra r ma nf Wf
(24)
where b is a constant that depends on the Reynold’s number, nf is the quantity of all fireball constituents (mol), which includes combustion products and entrained air, Wf is the mean molecular weight of the fireball mixture (g/mol), the superscript k indicates the time step number, Dt is the selected time step (sec), r is the fireball density (g/cm3), a is another constant that depends on the Reynold’s number, ra is the density of the ambient air (g/cm3), ma is the dynamic viscosity of the ambient air (g/cm-sec), and r is the fireball radius (cm). The product nfWf, which is the fireball mass, is calculated as: nf Wf ¼ np Wp þ na Wa
(25)
where Wa is the molecular weight of air (28.97 g/mol). The drag coefficient, Cd, is expressed as: Cd ¼ a Reb
(26)
where Re is the Reynold’s number of the rising fireball given as
Re ¼
ra u2r ma
(27)
where u is the rise velocity (cm/s) and r is the fireball radius (cm). The constants a and b are both functions of the Reynold’s number, and for flow around a sphere are provided in Table 6.6, based on simple curve fits to graphical data. Although the fireball will deviate from a spherical shape, the errors introduced are considered insignificant for our purposes. Table 6.6 Drag coefficient constants Range
a
b
Re 1.9 500 > Re > 1.9 Re 500
24 18.5 0.44
1 0.6 0
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CHAPTER 6 Nuclear-Powered Payload Safety
The height of the center of the fireball is determined by integrating the rise velocity over time with z equal to z0 as the initial condition. Thus: Zt z ¼ z0 þ u dt (28) 0
A hemispherical fireball is formed initially. The fireball grows in the shape of a truncated sphere as it rises, attaining a full spherical shape when it lifts from the ground. This growth is depicted schematically in Figure 6.20. The volume and surface area of the fireball before liftoff are given by: 4 3 1 2 pr pl 3rc þ l2 3 6
V ¼
(29)
and A ¼ 4pr 2 2prl
(30)
l ¼ rz
(31)
with and rc ¼
pffiffiffiffiffiffiffiffiffiffiffiffiffiffi r 2 z2
(32)
where V is the volume (cm3), r is the radius of the fireball (cm), rc is the radius of the circle formed by the intersection of the fireball truncated sphere and the ground (cm), and z is the height of the fireball center (cm). Equations (29)e(32) are solved to determine r and A given V and z. Following liftoff, the volume and surface area of a full sphere are used.
z
r z rC
l
FIGURE 6.20 Fireball formation geometry.
ground level
6.3 Launch Abort Environments Affecting SNPSs
The volume of the fireball is based on the ideal gas law: V ¼
nf RT p
(33)
The quantity of all fireball constituents, nf, is found by integrating the rate of fireball growth, n_ f , over time. Thus: Ztend
Ztend ðyn_ r þ n_ a Þdt
n_ f dt ¼
nf ¼ 0
(34)
0
where the fraction y (eq. (9)) is defined as the molar ratio of combustion products to reactants (determined by the equilibrium combustion solution), and n_ a is the molar rate of air entrainment (mol/s). The effective emissivity of the fireball, accounting for all particles within the fireball and assuming non-absorbing fireball gases and diffuse particle surfaces, is given by:
P ε ¼ 1e
Lb
Ns
i¼1
εip C i Aip
(35)
where Lb is an effective thickness of the fireball (cm), Ns is the number of particle sizes, εip is the emissivity of particle size i, Ci is the number concentration of particles of size i (cme3), and Aip is the effective projected area of particle size i (cme3), which is given by p 2 (36) Aip ¼ ci dei 4 where ci is the dynamic shape factor for particles of size i, and dei is the diameter of the volume-equivalent sphere for particles of size i (cm). The effective fireball thickness, Lb, is usually referred to as the mean beam length, which for a sphere is given as: Lb ¼ 1:3r
(37)
where r is the fireball radius (cm). Before the fireball attains a full spherical shape, the mean beam length is approximated as: Lb ¼ 3:9
V A
(38)
Equation (35) is easily modified to approximately account for absorbing gases (CO2 and water vapor) within the fireball. Thus: PN s i i i L ε C A b p i¼1 p ε ¼ 1 1 εg e (39)
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CHAPTER 6 Nuclear-Powered Payload Safety
where εg is the effective gas emissivity based on the temperature, pressure, and mean beam length of the fireball product gases. The calculation of the gas emissivity is based on an engineering approach assuming that water vapor and carbon dioxide are the dominant contributors to emissivity. Thus: εg ¼ εH2 O fH2 O þ εCO2 fCO2 Dε
(40)
where the H2O and CO2 subscripts refer to water vapor and carbon dioxide gas, respectively, and f and Dε are correction factors. The emissivities for the two dominant gases along with the correction factors are based on approximate curve fits to experimental data and are given as: log10 εH2 O ¼ ð0:88424 þ 0:65524T Þ
þ ð0:49502 þ 0:294787T Þlog10 Lb yH2 O p 2 ð0:061316 þ 0:05272T Þ log10 Lb yH2 O p ¼ 1 þ 0:897 0:2 log10 Lb yH2 O p p 1 þ yH 2 O 1
f H2 O log10 εCO2 ¼
(41) (42)
i
h
1:34561 þ 0:0758127T 0:122506ðT Þ2 i h þ 0:35278 þ 0:072651T þ 8:4779 103 ðT Þ2 log10 Lb yCO2 p i h 2 þ 0:074333 þ 0:0262193T 9:8478 103 ðT Þ2 log10 Lb yCO2 p (43)
fCO2
¼ 1 0:768622 0:195377 log10 Lb yCO2 p 0:69448 log10 Lb yCO2 p
Dε ¼ sin
*
yH2 O p 4:2 108 h2 þ 1:24 104 h þ 0:0361 yH2 O þ yCO2 h ¼ Lb p yH2 O þ yCO2
(44) (45) (46)
where T is equal to T divided by 1000; T is the fireball temperature (K), Lb is the mean beam length (cm), y is the mole fraction, and p is the fireball pressure (atm). The data points for the curve fits were selected to emphasize the high temperatures expected in a fireball. Equation (39) for fireball emissivity couples the aerosol physics-governing equations, which are used to determine the agglomerated particle concentrations, to the fireball temperature and size equations, and to the structure and particle heat transfer equations. Both the fireball energy equation and the fireball rise velocity equation are solved using a RungeeKuttaeFehlberg approach. An adaptive time step algorithm is incorporated to ensure both equations are solved accurately. This algorithm uses the difference between a fourth- and fifth-order solution to control the truncation error
6.3 Launch Abort Environments Affecting SNPSs
as the solutions are advanced in time. Thus different combustion scenarios can be simulated without concern for time step selection. It is necessary to define gas mixture properties for viscosity and conductivity in terms of the fireball constituent properties. The mixing model of Wilke as modified by Herning and Zipperer is used for this averaging process: m ¼
Np X k¼1
"
, yk m k
1=2 # Np X Wj yj Wk j¼1
(47)
where m is the dynamic viscosity (g/cm,sec), Np is the number of gaseous product species in the fireball at the current time, yk is the mole fraction for species k, and Wk, is the molecular weight for species k. As suggested in a prior report (Reid, Prausnitz and Poling, 1987), the same equation is used for conductivity with m replaced with k (W/cm,K). The final equations presented here for fireball physics describe the calculation of the turbulent dissipation term, which follows the work of Turner. This term is required in the aerosol physics solution of agglomeration due to turbulent diffusion. Although particle agglomeration due to Brownian motion is expected to dominate, the enhancement of agglomeration due to turbulence within the fireball should not be neglected. The local turbulent energy dissipation rate is approximated as the product of kinematic viscosity and the square of the vorticity: m .2 (48) εT ¼ V y r where εT is the turbulent energy dissipation rate (cm2/s3), m is the dynamic viscosity . of the fireball gas mixture (g/cm,sec), r is the fireball density (g/cm3), and y is the gas velocity vector (cm/sec). The flow field within the fireball is assumed to be approximated by a spherical vortex, for which the stream function is: ur 2 sin2 q r2 (49) 1 2 J ¼ ry 4 where u is the vortex (fireball) rise velocity, r is the radial position within the vortex, q is the angle describing the azimuthal position within the fireball, and ry is the radius of the vortex. The radial and angular components of the velocity vector, yr and yq are given by: 1 vJ u r2 (50) yr ¼ 2 ¼ cos q 1 2 ry r sinq vq 2 and yq ¼
1 vJ u 2r 2 ¼ sin q 2 ry 1 rsinq vr 2
(51)
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CHAPTER 6 Nuclear-Powered Payload Safety
Now the vorticity cross product can be evaluated as: 5ur sinq . V y ¼ 2ry2
(52)
The turbulent dissipation rate, assuming isotropic turbulence, can be expressed as: εT ¼
25m u2 r 2 sin2 q 4rry4
(53)
Because the fireball is modeled as a single uniform control volume, the spatial dependence of turbulent dissipation must be removed. A volumetrically averaged dissipation can be derived as: 3 εT ¼ 4pry2
Z2p Zry Zp
6px εT r sinq dq dr df ¼ 4pry3
Zry Zp r 4 sin3 q dq dr
2
0
0
0
0
(54)
0
where x ¼
25 u2 m 4 ry4 r
(55)
Now the average turbulent dissipation rate is: εT ¼
5 u2 m 2 ry2 r
(56)
where the radius of the vortex is taken as the radius of the fireball. This quantity is evaluated based on the current state of the fireball and used for calculation of agglomeration due to turbulent diffusion. To determine the products of combustion and their relative amount, a calculation of the combustion thermodynamics is required. The thermodynamic properties of the fireball are necessary to determine the temperature and size of the fireball, and to assess the vaporization and agglomeration of the plutonium-bearing particles. To simplify the combustion thermodynamics, two assumptions are made in the Fireball Code Package: • •
All combustion product gases are considered ideal although corrections are made to enthalpy for non-ideal behavior. An equilibrium chemical solution describes the situation adequately, which is probably reasonable for the high temperatures and reaction rates involved.
Chemical equilibrium is based on the minimization of either Gibb’s free energy or Helmholtz free energy. The Fireball Code Package utilizes the minimization of Gibb’s free energy, which is a function of pressure, temperature, and chemical
6.3 Launch Abort Environments Affecting SNPSs
composition. For a mixture of Np product species, the chemical contribution to the Gibb’s free energy for the mixture is given by: g ¼
Np X
m j nj
(57)
j¼1
where mj is the chemical potential for species j. If pressure and temperature are constant then: dg ¼
Np X
mj dnj
(58)
j¼1
The condition for equilibrium is the minimization of free energy and is applied to this equation subject to mass balance and non-negativity constraints. Of principal interest for fireball simulations is the behavior of plutonium-bearing particles. The discipline of aerosol physics deals primarily with the agglomeration (or forming into round masses) of aerosol particles, of the same or different size, in a gaseous suspension. Generally, particles of diameter less than 0.01 cm (100 mm) are considered to behave as aerosols. Larger particles, often referred to as “rocks”, are assumed not to agglomerate. The objective of aerosol physics is to determine how the fireball environment modifies the size distribution of plutonium-bearing aerosol articles.
Empirical formulae Previous studies (Mansfield, 1969), (Willoughby, Wilton & Mansfield, 1968), (Gayle & Bransford, 1965) have shown that the size, duration, and temperature of fireballs of various propellant mixtures have nearly the same quantitative behavior. Also, the empirical models used to characterize fireballs are generally well established and accepted. These estimates are based on experimental data from tests with up to 25,000 lb of liquid propellant and summarized in (Reinhart, 1999). In the studies, the fireball size and duration relationships were estimated from accumulated test and incident data, and analyzed empirically to obtain the following relationships: The maximum diameter of the fireball is given by: D ¼ 9:56 W 0:325
(59)
where D is the diameter of the fireball (ft), and W is the combined (fuel and oxidizer) propellant weight (lb). The standard error in D (i.e., the standard deviation of the data from the fitted equation), is about 30%. The fireball liftoff time is given by: tb ¼ 0:572W 0:167
(60)
where tb (secs) is the point in the fireball development when buoyancy and entrainment become dominant.
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CHAPTER 6 Nuclear-Powered Payload Safety
The duration of visible radiation (in seconds) is given by: tT ¼ 0:197W 0:349
(61)
with a standard error estimated to be 84%. The stem liftoff time defines the point when the heat flux drops abruptly, ceasing to be a measurable threat, is given by: tc ¼ 1:5tb
(62)
Pool Fires The accidental release of hundreds of gallons of flammable commodity such as hydrazine may not lead to an explosion; rather a pooling effect may occur leading to a large surface area at the bottom of the launch structure awaiting a potential ignition source to set the liquid alight. Unfortunately, hypergolic propellants like hydrazine can be easily ignited by the heat of reaction when they come in contact with organic material, especially rust, which can be plentiful in the high saline environment of beach-side launch pads. The major hazard from a burning pool fire at the base of the launch vehicle is that it will provide an ignition source for subsequent leaks of fuel along the stages of the rocket leading upward to the nuclear source in the payload. Secondarily, radiant heat effects, while probably not extremely threatening to a distant payload, can provide damaging consequences to a clad RHU that has been ejected from the payload during an explosion or tip-over event. However, the aeroshell protecting the clad is designed for the intense heating of re-entry and it is unlikely that the radiant temperature of the pool fire could significantly damage the RHU.
Solid Propellant Fire Hazard Threats similar to those provided by liquid propellant vehicles are introduced to the nuclear material onboard the payload by launch vehicles utilizing solid propellant. However, while a liquid explosion can provide a relatively short duration fire environment in an airborne explosion, a solid rocket motor can fragment as a result of either mechanical failure or purposeful destruct action to halt an erroneous flight trajectory. These fragments can return to the ground either burning, if conditions are ripe to maintain the burning process, or unburned, in varying sizes. If the size of a fragment (i.e., greater than critical diameter), and impact velocity are sufficient, these secondary “bombs” can explosively burn forming large craters, blast overpressures and intense heat. If the distribution of these solid propellant fragments is such that one or more is near a nuclear energy source (e.g., exposed or clad pellet, RHU, or RTG), there is a danger of vaporizing the nuclear material.
Solid propellant composition The thermal environment exhibited by burning solid propellant is determined in large part by its chemical makeup. Other factors affecting variability in the thermal
6.3 Launch Abort Environments Affecting SNPSs
environment include the geometrical configuration/size of the burning fragments, the amount of fracturing due to mechanical damage, and ambient conditions. Solid propellant stores the fuel and oxidizer together in solid form. The fuel is typically powdered aluminum, and the oxidizer is ammonium perchlorate (Wertz & Larson, 1991). A synthetic rubber such as polybutadiene holds the fuel and oxidizer powders together. Table 6.7 shows the typical type and quantity of some solid propellants used on launch vehicles and also lists the ingredients of several popular propellants.
Burn rate The surface recession rate r, called the burning rate, is an empirically determined function of the propellant composition and certain conditions within the combustion chamber. These conditions include propellant initial temperature, combustion pressure, and the velocity of the gaseous combustion products over the surface of the solid (Hill and Peterson, 1970). This function is approximated by: r ¼ a pn
(63)
where a and n are constants for a particular propellant mix and p is the pressure within the combustion chamber. Values of the constants for several propellants are also shown in Table 6.7.
Combustion process The qualitative description of the combustion process which a solid propellant chunk will exhibit during ambient-atmosphere burning on the ground is presented here from Price et al. (1979): Because of slow aluminum combustion, the combustion zone has the macroscopic structure shown in Figure 6.21. In a thin region (roughly one centimeter thick) adjoining the propellant surface, the oxidizer and polymeric binder decompose and interact. This is followed by an extended Region A in which the aluminum droplets are burning, and the gaseous products are near equilibrium composition (i.e., equilibrium for the amount of aluminum burned at that point). Region B corresponds to a distance from the burning surface great enough so that all of the aluminum has burned (observed to be about 1 meter). Along the sides of the combustion plume there are regions where the environmental air has mixed with propellant products. Because the propellant products are highly fuel rich, air admixture leads to further reaction and heat release, concurrent with dilution and cooling with atmospheric nitrogen. The important aspect of this macroscopic description of the combustion zone is the recognition of the difference in reaction rate-controlling processes in the different regions. Near the propellant surface, the dominant rates involve binder and oxidizer decomposition and mixing. In Region A, the rate is limited by the aluminum droplet combustion. In Region B, combustion is constant. In Regions C and D, the rate is limited by air mixing. The principal scale effect in going to large propellant samples is the decreasing importance of edge effects (Regions C and D).
303
304
Vehicle use
Propellant (mfr.)
Ingredients (% by weight)
Titan IV SRMU
QDT (Alliant)
50% 19% 19% 09%
02% 01% Titan IIIC
UTP-3001 (United Technology Center)
44.34% 23.17% 16.12% 10.86% 05.26% 00.25%
Ammonium perchlorate, 200 m (oxidizer) Ammonium perchlorate, 20 m (oxidizer) Aluminum powder (fuel) Hydroxyl-terminated polybutadiene (HTPB) (binder) Dioctyl sebecate (DOS) (plasticizer) Other (curative, catalyst, bonding agents) Ammonium perchlorate, 200 m (oxidizer) Ammonium perchlorate, 8 m (oxidizer) Aluminum (fuel) Polybutadiene acrylonitrile (PBAN) (binder) Other (curative, plasticizer) Iron oxide (burn rate modifier)
Burn rate constant, a in/sec
Burn rate constant, n
0.0401 500 > p > 1300 psia T ¼ 80 F
0.30
0.0694 500 > p > 1200 psia T ¼ 70 F
0.25
CHAPTER 6 Nuclear-Powered Payload Safety
Table 6.7 Typical solid propellant ingredients
Space Shuttle SRB
TP-H-1148 (Thiokol)
69.9% 16.0% 14.0% 00.1%
Delta II GEM
QDL (Alliant)
69.000%
Minuteman III
ANB-3066 (Aerojet)
73.00% 15.00% 08.56%
03.44%
0.04424 200 > p > 1700 psia T ¼ 77 F
0.33
0.2353 500 > p > 1500 psia T ¼ 80 F
0.35
0.0597 400 > p > 1000 psia T ¼ 80 F
0.27
6.3 Launch Abort Environments Affecting SNPSs
19.000% 06.362% 03.602% 02.036%
Ammonium perchlorate (oxidizer) Aluminum (fuel) Polybutadiene acrylonitrile (PBAN) (binder, curative) Iron oxide (burn rate modifier) Ammonium perchlorate (oxidizer) Aluminum powder (fuel) HTPB (binder) DOS (plasticizer) Other (curative, accelerator, catalyst, bonding agent) Ammonium perchlorate (oxidizer) Aluminum (fuel) Carboxy-terminated polybutadiene (CTPB) (binder) Polybutene (plasticizer, crosslinker)
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CHAPTER 6 Nuclear-Powered Payload Safety
REGION C: Oxidizer, binder, aluminum, and air in equilibrium, stoichiometric mixture
ION D: REGION Mixing region,, with unreacted aluminum m present
REGION B: Oxidizer binder, bin Oxidizer, aluminum reacted, no air admixture REGION A:
1 m (observed)
1 cm
Oxidizer and binder reacted, aluminum oxidation proceeding
PROPELLANT
FIGURE 6.21 General structure of solid propellant plume.
Region A remains the same thickness for all sample sizes that are large enough for complete aluminum combustion before appreciable air admixture. The macroscopic structure of the combustion zone is substantially determined by microscopic details, as is the nature of the two-phase flow seen by an object immersed in the fire. Looking at the combustion sequence on a more microscopic scale, the oxidizer and binder on the propellant surface decompose to gases, while the aluminum particles (10 to 30 microns) accumulate and adhere to form larger agglomerate droplets up to 299 microns in diameter. The oxidizer and binder gases mix and react near the burning surface. The aluminum agglomerate droplets ignite and burn on the surface, then move out into the gas flow. There they burn so slowly that they produce an extended combustion zone roughly 1 meter in thickness. The final product flow is roughly 30% aluminum oxide, which is in liquid droplet form. Starting with aluminum agglomerate droplets in the 20 to 200 micron diameter range, the droplets burn by two paths that produce distinctive product oxide droplets. Fine (<2 micron) smoke droplets are formed in a flame envelope around each aluminum droplet. Oxide also forms and accumulates on the aluminum. This in turn forms a distinctive oxide lobe that eventually becomes a residual droplet when the aluminum is consumed. Thus an increasingly dense oxide cloud develops as the burning aluminum droplets and smoke move away from the propellant droplets, and larger ones being formed later as the larger aluminum droplets burn out. A series of chemical equilibrium calculations were made for UTP 3001 propellant at 1 atm at the Air Force Weapons Laboratory, with percent aluminum reacted as
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.22 Gaseous reaction products as a function of aluminum burned.
a parameter. The results are shown in Figures 6.22 and 6.23. In these calculations, the temperature of the unburned aluminum was assumed as shown in the figure. Of the three temperatures referred to, 2533 K is most plausible for Region A, while lower values are appropriate near the burning surface, which corresponds to near 0% aluminum reacted. From these figures it can be seen that the aluminum combustion brings the temperature in Region A up from 2350 K at about a centimeter above the burning surface, to 2990 K at the start of Region B where aluminum combustion is complete. In the process, H2O and CO2 are reduced to H2 and CO. The equilibrium properties shown at 100% aluminum burned are applicable to Region B of the combustion zone. This is a major part of the fire environment of a large piece of propellant, encompassing that region more than 1 meter or so from the burning surface, out to those regions where air admixture has become important. For a large propellant sample (e.g., 1 meter across the burning surface), the temperature calculated by
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3000
2900 Calculated Equilibrium Temperature, K
308
2800
2700 300 K
Assumed Temperature of 833 K Unbumed 2633 K Aluminum
2600
2500
2400
2300 0
20
40
60
80
100
Aluminum Burned, %
FIGURE 6.23 Calculated combustion zone temperature.
this means is probably more accurate than can be measured directly. The calculation for less than 100% aluminum burned tells nothing about the droplet size distribution at each point in the combustion zone or the spatial variation of temperature within the combustion zone. Table 6.8 provides a summary of the reaction zone characteristics for UTP 3001 propellant. The extended region of the combustion zone of the propellant in which aluminum droplets burn is a region where chemical heat release continues locally around the aluminum droplets and flows to the surrounding media, with corresponding temperature gradients. The most intense heat release is in the high temperature zone, which is a detached flame envelope around each aluminum droplet. In this flame envelope, the most energetic step is the formation of Al2O3 in liquid form, primarily on the surface of existing smoke droplets in that envelope. It is the resulting energy of that reaction, and the corresponding high temperature of those smoke droplets, that makes the combustion so luminous. At 1 atm the Al2O3 temperature is about 3800 K. In a convective environment the flame envelope may be drawn out into a tail. From a practical viewpoint, it should be stressed that the high temperature
6.3 Launch Abort Environments Affecting SNPSs
Table 6.8 Estimates of reaction zone characteristics for UTP 3001 Propellant burn rate (measured) Propellant density (measured) Propellant mass burn rate Combustion zone thickness (Region A, observed) Burning time of large agglomerates Temperature after complete burning of Al (calculated) Gas density in the combustion zone Mass of Al2O3 per unit volume in the product flow1 Flow velocity2 1 2
0.127 cm/sec 1.76 g/cm3 0.233 g/s,cm3 1.0 meter 0.07 sec 2990 K 0.78 10–4 g/cm3 0.31 10–4 g/cm3 15–30 m/sec
Based on chemical equilibrium data actually applicable to Region B. Based on several methods of determination.
around the burning agglomerates, while the source of a large part of the radiant energy and heat, contains very little mass, and only small droplets that are not expected to be deposited extensively on objects in the flow. Inside the flame zone, the aluminum droplet is necessarily at a temperature below the aluminum boiling point (2740 K) and above the Al2O3 melting point (2318 K). The reason is that the flame cannot support the high vaporization rate of the boiling point, while the combustion would be arrested, or drastically reduced, if the oxide were frozen (because it would encapsulate the agglomerate). The aluminum droplet, at a temperature that lies between 2318 and 2740 K, with 2500 K taken as a best estimate, and, in aggregate, representing a substantial mass, readily impinges on immersed objects, reacting exothermally after being retained on the object. This effect (impingement and retention on an immersed object) would be more pronounced near the burning surface of the propellant where the aluminum concentration is high. The third temperature zone of the aluminum combustion is all that volume outside the droplet flame envelope, where the temperature is something like the flame temperature of the propellant (with due regard for the percent aluminum consumed at the particular location). This temperature ranges from 2350 K near the propellant surface to 2990 K further out where the aluminum is 100% consumed. In summary, most of the fire environment consists of the following: •
• •
The bulk of the volume (gas, smoke, and residual oxide droplets) is at a temperature ranging from 2350 K (0% Al burned) to 3000 K (rounded from 2990 K, for 100% Al burned). Only the residual oxide droplets are large enough for impingement. The unreacted aluminum in burning droplets is at about 2500 K. Droplets are large enough for impingement. The flame envelope around the aluminum droplets, containing a very small mass of hot gases and Al2O3 smoke droplets, is at about 3800 K. It is highly visible because of its high temperature.
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Spatial distribution of temperature is as follows: Near the propellant burning surface (lower Region A) there is a heavy flow of burning aluminum (droplet temperature about 2500 K), with a correspondingly large flame envelope radiation (temperature about 3800 K); and with a correspondingly low smoke density in the rest of the volume (temperature range of 2400 to 2500 K). As the material flows further from the surface (outer Region A), the aluminum droplets decrease in size and total mass while remaining at about 2500 K; the total volume and radiation of the flame envelope (temperature of 3800 K) decreases; the population of both smoke and residual oxide droplets increases and the temperature rises towards 3000 K if the air admixture does not intervene unduly. Beyond about 1 meter (Region B), the aluminum droplets and their flame envelopes disappear, leaving only a homogenous flow of gas, oxide smoke, and residual oxide droplets at about 3000 K. The radiation field in the combustion zone showed spectral lines of several gaseous species and a higher level of thermal radiation from the Al2O3 smoke cloud. Measurements of the smoke cloud 15 cm from the burning surface indicated a temperature of about 2800 K, an emissivity of 0.03, and an absorption coefficient of about 0.1/cm. Intensity-emissivity measurements are wavelength-dependent, apparently due to the presence of a small amount of smoke that is in the agglomerate flame envelope, which is much hotter than the bulk volume.
Experimental measurement Immersion heat probes, including bare thermocouples, immersion calorimeters and water-cooled calorimeter tubes, were used to investigate the fire environment. The environment experienced by these types of devices would be similar to a test object in the fire such as an RTG. Bare thermocouples experienced direct impingement with droplets in the flow, gave very erratic readings, and survived only briefly. Some measurements made in the wake of deflectors survived longer and gave better records; however, it was not clear how the temperature there could be related to the free stream temperature, even if the thermocouple survived. In high-temperature flows where immersion probes deteriorate, measurements of the heat-up rate of the probe can be used to estimate the plume temperature when an equation for the heat transfer between the probe and the flow can be written. When only convection and radiation were considered, an inordinately high plume temperature was inferred. This demonstrated that the heat transfer from deposition and reaction of condensed material on the probe was a major component of the heat transferred to the probe. In tests at Los Alamos National Laboratory to determine the response of Light Weight Radioisotope Heater Units (LWRHUs) when exposed to solid propellant fires (Tate & Land, 1985), thermocouples measured temperatures of 2333 K out to distances of at least 1.8 meters for UTP 3001 propellant. In tests at Lawrence Livermore National Laboratory to characterize the thermal environment of burning Minuteman III Stage 3 propellant (ANB-3066) (Diaz, 1993, 1994), spectral radiometry and pyrometry methods were used to determine plume temperatures. Temperatures along the centerline of the plume were observed by using graphite sight tubes, with the inboard end cut at an angle. The sight tube was inserted
6.3 Launch Abort Environments Affecting SNPSs
into the plume with the slant cut facing in the direction of flow of the plume. The sensor (for spectrometry or pyrometry methods) and sight tube were mounted, one on each end, to an alignment tube. Heat fluxes to objects engulfed in the plume were estimated by imbedding thermocouples in solid metal objects and placing them in the plume. Inverse modeling calculations were then completed using a heat transfer model to duplicate the observed temperature history of the calorimeters. Two devices were used. One consisted of a 2.5 cm diameter by 15.25 cm long molybdenum rod with two tungstene rhenium thermocouples embedded on the centerline at 3.18 cm and 8.25 cm from the exposed end of the rod. An additional thermocouple was placed on the wall at 8.25 cm from the exposed end. The rod was wrapped with an insulating ceramic blanket and placed in a holder that exposed only one end of the rod. The second calorimeter was a 5 cm diameter by 5 cm long stainless-steel cylinder, with one end machined into a hemisphere. The other end was drilled to the center point of the hemisphere and a tungstenerhenium thermocouple was placed in the center. The temperature, emissivity, and heat flux profiles obtained from these experiments are shown in Figure 6.24. The temperature is seen to vary from 2273 K at the burning surface to a peak at 2573 K about 32 cm off the initial surface. The heat flux of about 20 cal/ cm2$sec was measured at a standoff of 25 cm. There was a high level of confidence in the experimental methods; however, the temperatures were observed to vary by 10% due to the turbulence and thermal characteristics of the propellant fire. ANB3066 has an adiabatic flame temperature of 3003 K. In tests at Sandia National Laboratories to characterize the thermal environment from an unpressurized burning bed of an AP, Al, and HTPB composite solid propellant Gill et al. (1994), measurements were made with arrays of slug calorimeters located adjacent to the plume and with cylindrical calorimeters inserted into the plume. The locations of the slug calorimeters were chosen so the plume was viewed from several vantage points. The analytic model viewed the plume as a well-stirred reactor, which could be characterized by a thermal loss that is a fixed percentage of the total heat released in the reactor volume, and by the reactor temperature. From this the plume was found to have an effective temperature near 2073 K and to radiate 2600
50
0.5
Temperature 30
2200 Heat Flux 2000
20
1800
10
1600 0
10
20
30
40
50
60
70
80
90
0 100
Standoff from Initial Surface of Propellant, cm
FIGURE 6.24 Temperature, heat flux, and emissivity profiles for ANB-3066 burning in air.
0.4 0.3 0.2 0.1 0.0
Emissivity
40
2400
Heat Flux, cal/cm2 s
Temperature, C
Emissivity
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CHAPTER 6 Nuclear-Powered Payload Safety
34% of the total heat released. To simulate a body immersed within the plume, cylindrical calorimeters were positioned horizontally above the propellant burning surface at two elevations, at 5 cm and 20 cm standoffs. A thin-walled (0.64 cm) and thick-walled (3.2 cm) version, both with a 10.2 cm outside diameter, were positioned 11 cm apart at each standoff. The calorimeters were divided into thermally isolated quadrants to provide a measure of circumferential variation of the heat flux. Each quadrant of the thick-walled unit and the top and side quadrants of the thin-walled unit had one thermocouple. The bottom of the thin-walled calorimeter had three thermocouples. They were oriented so that each quadrant preferentially viewed the top, bottom, and lateral heat fluxes. The heat flux versus time from each segment indicated a slow increase in flux to a maximum level. The thin calorimeters then melted while the thick calorimeters remained intact. The maximum heat flux levels are shown in Figure 6.25. The average heat flux into a calorimeter is 84.7 W/cm2, with a range of 8%. The average incident heat flux (W/cm2) to a specific quadrant of these calorimeters is given by: q ¼ 84:7 33:5x3 þ 11:5x4 x3 ¼ þ1 for facing up x3 ¼ –1 for facing down x3 ¼ 0 for being vertical
(64)
x4 ¼ þ1 for facing in x4 ¼ –1 for facing out x4 ¼ 0 for being horizontal
36
52
60
79
77
110
107 11 cm typ.
Thin walled calorimeters
60
Thick walled calorimeters 77
67
92
113
93
68
20 cm standoff
133 5 cm standoff
PROPELLANT BURNING SURFACE
FIGURE 6.25 Heat flux (W/cm2) into quadrants of cylindrical calorimeters immersed in the plume.
6.3 Launch Abort Environments Affecting SNPSs
The Chemical Propulsion Information Agency (CPIA) estimated the adiabatic temperature for Titan IV SRMU propellant burning at ambient pressure at 3097 K using the NASA Lewis e Chemical Equilibrium and Transport Properties of Chemical Systems Program (Gordon, McBride & Coveny, 1990).
Explosions Pre-launch to early flight phase explosions Even before the launch vehicle begins its escape from the gravity well there are scenarios in which onboard nuclear source materials can be damaged. The final preparations for integration of the nuclear source system and the launch vehicle/ spacecraft involve hazardous operations which, in failure situations, can lead to adverse environments similar to launch failure scenarios. Drops during crane lifts, electrical connection failures causing fires, electrostatic discharge to solid propellants, and other processing failures can lead to the destruction of the entire launch pad or vehicle processing facility. This was woefully exemplified on August 22, 2003, during final processing of Brazil’s Veı´culo Lanc¸ador de Sate´lites (Satellite Launch Vehicle) (VLS) at Alcaˆntara (Figure 6.26). Twenty-one people working at the launch site were killed when one of the rocket’s four first stage solid rocket motors ignited accidentally.
FIGURE 6.26 VLS-1 pre-launch accident, 2003.
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Probably the most hazardous threat to nuclear material onboard the payload occurs in a scenario in which the vehicle launches nominally, then shortly after liftoff executes an immediate rotation impacting the ground nose first. In this position, large amounts of unburned solid propellant can be placed either on top of or beside the nuclear source. Impact velocities sufficient to cause secondary explosions and subsequent burning of solid propellant sets up an environment conducive to potential removal of the nuclear material’s protective shielding followed by vaporization of the material. This scenario was played out (sans nuclear payload) on February 15, 1996, when the Chinese Long March 3B rocket veered off-course immediately after launch, crashing in the nearby village only 22 seconds later, killing an estimated 100 people and destroying 80 houses in Xichang (Figure 6.27).
Ascent phase explosions During later flight, as the vehicle increases velocity, an explosion below the payload would tend to throw the nuclear material forward of the flight path and since this segment would have different aerodynamic characteristics its ballistic trajectory should keep its impact point clear of the returning solid fragments. Figure 6.28 shows the ballistic instantaneous impact points for two times. The early breakup of the vehicle at Tþ20 seconds shows that debris tends to impact in one area as the flight is nearly straight up and the programmed turn
FIGURE 6.27 Chinese Long March 3B launch accident, February 1996.
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.28 Early versus later flight impacts: at Tþ20 seconds (left) and Tþ40 seconds (right).
down-range has not yet been initiated. Later flight is depicted in the second figure at Tþ40 seconds revealing the debris field has elongated as more impacts occur over a range of points due to varying ballistic coefficients of pieces. Intact impact of the spacecraft itself with the smaller quantities of liquid propellants minimizes the possibilities of overpressures that would harm the nuclear source. There is some concern that solid “kick” motors attached to the spacecraft could impact with sufficient velocity to damage the nuclear material or impose a thermal environment upon burning on the ground to vaporize the source.
Fallback secondary explosions Figure 6.29 depicts the bombardment of Launch Complex 17 as debris and solid propellant fragments fall during the Delta II explosion on January 16, 1997. These fragments can impact onto hard surfaces (e.g., concrete or steel, Figure 6.30), soft surfaces (e.g., sand or grassy soil), or water. Included in the falling debris, as well as the solid propellant, are any nuclear sources either maintained in their original configuration or freed by the explosive events. These can impact prior to or following a propellant fragment. The response of solid propellant to ground impact may range from possible ignition and burning to explosive burning or even a detonation. It can be characterized knowing the size of the fragment, its impact velocity, and the type of surface that it impacts. The consequence of this response can then be defined knowing the spatial relationship between the propellant fragment impact point and the nuclear material during the short time interval immediately following the solid propellant ground impact. For a solid motor fragment to sustain detonation it must be of supercritical size and it must be subjected to a shock stimulus sufficient to initiate detonation.
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FIGURE 6.29 Secondary impacts, Delta II explosion, January 16, 1997.
FIGURE 6.30 Impact Crater Delta II explosion, January 16, 1997.
Determining this critical diameter, at which detonation could occur, was one of the objectives of solid propellant testing.
Detonation response The mechanisms of propellant detonation are dependent upon the variables of size and impact velocity. Detonation is a process in which an explosive reacts to products very close behind a steady, high-pressure shock wave that is supported by the explosive reaction. Detonation is characterized by high rates of propagation (4e9 km/s), extremely high pressures (10e40 GPa), and a short time to the start of reaction (typically 1 ms) (Lee & Green, 1988). In a detonation, 10 cm of material is consumed
6.3 Launch Abort Environments Affecting SNPSs
fully in slightly more than 10 ms. Conversely, deflagration refers to a range of reaction rates involving burning. The term burning is usually applied to a slow deflagration reaction that does not generate significant overpressure and whose time scale to consume 10 cm of material is greater than 1 second. If the material has a relatively large surface area, then a high-velocity deflagration can involve measurable or damaging overpressure and a time scale of 100 ms to 10 ms to consume the material. In this case, the reaction is propagated by conduction, convection, and radiation, and not by a shock wave. The term “explosive burn” is used for high-rate deflagration with significant overpressure to distinguish it from a high yield event like detonation. There are several responses that the propellant can exhibit given these varying parameters. The shock-to-detonation (SDT) response is a process in which a shock wave entering an explosive material develops into a detonation (Figure 6.31). Each type of explosive has a characteristic distance of travel for the transition of a shock wave of given amplitude into a detonation. The deflagration-to-detonation transition (DDT) is a process in which a deflagration reaction evolves into a detonation. The term is somewhat misleading in that gas pressure created by the deflagration compresses unburned, porous explosive material to cause a reaction that eventually leads to an SDT process. The initial deflagration does not simply burn faster and faster until a charge detonates; rather, its role is to drive the compaction wave. Figure 6.32 shows a porous bed of explosive that is ignited. The burning explosive compresses unburned material to form a compaction wave and a reaction within the wave by the intensity of compaction. The intensity of the compaction wave finally grows to that of a shock wave, and a detonation develops. XDT is an initiation process that originally referred to an “unknown” (i.e., X) detonation transition. XDT now commonly means a delayed detonation in which impact damages and creates high porosity in an explosive and then causes a subsequent compaction wave that ultimately leads to a detonation (Figure 6.33). XDT is closely related to DDT in that a compaction wave is the route to detonation. However, the compaction wave in XDT is mechanically, rather than combustion, driven.
Unshocked, unreacted
Shock front
explosive
Detonation
Detonation
products
front
Entering Shock Wave Shocked, partially reacted explosive
Reaction zone
FIGURE 6.31 Shock-to-detonation transition (SDT)eshock wave entering explosive develops to detonation.
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Porous
Compaction
Detonation
wave
products
Container
explosive
Combustion of porous explosive
Compacted Combustion
explosive
Detonation front
Partially reacted explosive
products
FIGURE 6.32 Deflagration-to-detonation transition (DDT) e initial deflagration drives a compaction wave, which grows to the intensity of a shock wave and detonation occurs.
Undamaged
Impact compressed
Recompressed
explosive
explosive Compaction
explosive
wave
Fractured and rarified Impact surface
explosive
FIGURE 6.33 Delayed detonation transition (XDT) e impact damages an explosive and creates high porosity. The continuing impact then causes a compaction wave that leads to detonation.
Detonation hazard The hazards associated with ground impact induced solid propellant detonation, whether arising from SDT, DDT, or XDT, can be very severe if appropriate conditions exist to facilitate the detonation transition. Because of the severe blast overpressure and fragment environments potentially generated by solid propellant detonation, the special conditions necessary to cause SDT, DDT, or XDT should be examined in detail. Detonations typically produce: • • •
Impact crater dimensions two to three times greater than those expected for typical propellant impacts. Air shock overpressures near the crater as high as 2000 psi. Pressures within the crater region as high as detonation pressures (i.e., 20 to 30 kbar, or greater).
6.3 Launch Abort Environments Affecting SNPSs
The capability of the propellant segment or fragment to achieve sufficient impact velocity to undergo detonation must also be examined. In the case of the Cassini payload launched on a Titan IV, the flight profile made conditions necessary to achieve SRMU propellant detonation very unlikely. The launch profile limited maximum SRMU segment and propellant fragment impact velocities to less than about 600 ft/sec, with impacts most likely to occur on soft sandy soil. For mission elapsed times greater than approximately 23 seconds, fragments would impact in the Atlantic Ocean. These velocity limitations bound the upper limit of ground impact pressures to the 2e3 kbar range for soft sandy soil impact and 7e9 kbar for hard surface impact (e.g., steel or concrete). As shown in Figure 6.34, this threshold is well below the rather conservative impact velocity of 1000 ft/sec (and nominal pressure of 20e30 kbar) required for SRMU propellant detonation. The probability that solid propellant segments or fragments will explode on impact depends on their amount and impact velocity. For Class 1.3 solid propellant in general, Figure 6.35 shows these probabilities given propellant weight and impact speed.
Critical diameter In order to sustain the detonation shock wave, the propellant geometry must be such that the segment or fragment is supercritical. Below this critical size, the shock wave will lose strength and diminish below supersonic speed which will likely result in deflagration only. Burning propellant does have its hazards however, especially in providing a high thermal environment to exposed RHUs. Above critical size, chemical and physical interactions will “feed” the shock wave, maintaining state dynamics for continuation of the shock’s travel leading to detonation.
LAND
800 100% Yield Possible for Intact Motors/Segments
Minuteman Stage 2 hard surface impact
600
Significant Fracture/Explosive Burn of Bare Propellant
shotgun & plate tests hard surface impact
400 Fracture/Explosive Burn Threshold
impact motors or segments Ttan III 34D-9
LAND
200
WATER
Impact Velocity, ft/sec
WATER
Detonation Threshold (conservative)
lg chunks,2 - 4 x critical dia, hard surface impact
1000
Ignition Threshold
0 0
5
Titan IV Cassini SRMU Propellant Impact Velocity
15 10 Mission Elapsed Time (from T-0), seconds
20
23
25
FIGURE 6.34 Comparison of Cassini Titan IV SRMU segment and propellant fragment maximum ground impact velocities with impact thresholds.
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100% 90% 80%
40%
1, 00 0
lb
lb 00 2,0
10,0
50%
lb
60%
4,00 0
lb
00 l b
70% 20,0 00
Probability of Explosion
320
30% 20% 10% 0% 0
100
200
300
500 600 400 700 Impact Velocity (ft/sec)
800
900
1000
FIGURE 6.35 Probability of solid propellant (1.3) explosion given mass and impact.
There is a unique and explicit relationship between the critical diameter of a material and the critical dimensions of a charge of any shape with uniform cross-section cast from that material (Elwell, Irwin & Vail, 1967). This relationship is the critical geometry, sc, which is defined by: 4A (65) sc ¼ P c where A is the area of the cross section, P is its total perimeter, and the subscript c refers to the critical conditions. For a right circular cylinder, eq. (65) becomes: sc ¼ dc
(66)
where dc is the critical diameter. If the critical diameter is known, eqs. (65) and (66) define the critical ratio of cross-sectional area to perimeter for any shape. Since area and perimeter can be expressed in terms of the dimensions of the cross section, the critical value of the dimensions (i.e., the critical size) of any shape is thereby defined. Testing of various propellant sample shapes were used to verify and refine eq. (66) through statistical analysis of data. The empirical form of the equation is: sc ¼ ð0:92 0:03Þdc
(67)
Solid propellant experimental testing Most of the knowledge concerning the explosive behavior of hydrocarbon rubber binderealuminum fueleammonium perchlorate oxidizer (HC/Al/AP) propellant was gained in the 1960s from the results of Project SOPHY (Elwell, Irwin & Vail, 1967) conducted at Edwards Air Force Base, CA (Merril, Nichols & Lee, 1999).
6.3 Launch Abort Environments Affecting SNPSs
The SOPHY program looked at only 84% total solids (i.e., ingredients besides binder) propellant. More modern propellants have higher total solids loadings (86e90%) that produce higher thrust performance. Little is known about the explosive behavior of these propellants. Solid rocket motors for the Titan IV (SRMU) have 88% total solids and the Space Shuttle SRBs have 86% total solids (Table 6.7). The rocket propulsion industry had seemed to assume that explosive potential and impact velocity thresholds for violent propellant response would be similar to that found in the SOPHY program. Table 6.9 shows the range of critical diameters derived from experimentation. These data, and incidents like the 1986 Titan 34D-9 explosion at Vandenberg Air Force Base, CA, indicated that large booster propellants may be more explosive than previously thought and highlighted the need to improve our understanding of booster propellant explosive behavior. Technology now enables us to study this behavior more economically. Polyvinyl difluoride (PVF2) piezoelectric gauges and advanced recording instrumentation now more accurately measure and record times of arrival coupled with shock magnitude measurements and shock magnitude time histories. During the 1960s, impact experiments were carried out by rocket sled impact. Hardware and range maintenance and refurbishment costs involved high capital outlays and use of substantial manpower. Experimental methods for substantially sized propellant impacts have evolved from ramming a rocket motor or propellant sample on a rocket sled into a barrier, to putting impact barrier material on a rocket sled so that a static, highly instrumented test sample could more easily provide high speed data than in the moving sample impact process, to explosively driven steel plate impactors into instrumented propellant samples resting on relatively inexpensive, replaceable wooden stands. Advanced techniques in combination with computer code simulation could make very large propellant sample testing unnecessary. The Lawrence Livermore National Laboratory (LLNL) has developed a reactive response hydrodynamics computer code called CALE/PERMS (C-language based Arbitrary Lagrangian-Eulerian/Propellant Energetic Response to Mechanical Stimuli), which is moving in that direction. In 1996 the Propellant Impact Risk Assessment Team (PIRAT) was assembled to specifically investigate the explosive risk posed by a Titan IV-SRMU launch. Selecting propellant sample sizes based on critical diameters reported by the French (Brunet & Salvetat, 1988), and the Super HIPPO experiments in 1992 (Merrill, 1992), a test program was developed with the intent to have impact samples in
Table 6.9 Propellant testing data Propellant type 84% HC/Al/AP 88% HTPB/Al/AP 90% HTPB/Al/AP 90% HTPB/Al/AP
Experimental critical diameter 1778 mm (70 in) 200–250 mm (8–10 in) 83 mm (3.27 in) 305 mm (12 in)
Year tested 1967 SOPHY (Elwell, Irwin & Vail, 1967) 1988 French Study (Brunet & Salvetat, 1988) 1988 French Study (Brunet & Salvetat, 1988) 1992 Super HIPPO (Merrill, 1992)
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excess of critical diameter or at least close to critical diameter so that substantial energetic response to impacts could be observed. With support of CALE/PERMS it was projected that even data from samples smaller than critical diameter could provide useful inputs for computer code modification. Cylindrical diameters of 203 mm (8 in), 254 mm (10 in), 305 mm (12 in) and 559 mm (22 in) were selected. Using PVF2 gauges, detonation velocity probes, foil gauges, crystal pins, high speed data recorders, overpressure sensors, and high speed film and video recorders, test articles were initiated using the setup shown in Figure 6.36 with some variations for each diameter tested. After getting encouragement that critical diameters for 88% solid HTPB propellants might be in the range of 200 to 250 mm, it was found that the first booster propellant formulation tested has a critical diameter larger than 559 mm and, perhaps, twice as large as 559 mm. Although the propellant sizes are large in normal testing of explosive materials, they are definitely subscale compared with solid propellant booster motors that can contain more than 500 metric tons of propellant. Larger size samples might be employed in the future since investigation of characteristics above critical diameter is desired. Solid rocket booster propellant in explosive events seems to provide high atmospheric overpressure yields without being involved in true detonations. If detonations are steady state supersonic events
Nonelectric cap
1 1/2" C4 Sphere 20" Primacord, typ.
42" Primacord 2 EBWs
Nonelectric cap, typ. 21" Plywood template 5" × 21.5" dia. C4 Aluminum light shield
Instrumentation Crystal Pins 6" apart 2 columns PVF2 gauges @ center, radius, and mid-radius 4 sets
37" × 22" dia. propellant segment
7" × 22" dia. propellant segments
2’ and 4’ velocity gauges, 180° apart 3" × 22" dia. paraffin
FIGURE 6.36 Twenty-two inch critical diameter test apparatus.
6.3 Launch Abort Environments Affecting SNPSs
in the energetic material and deflagrations are explosive events characterized by subsonic shock in the energetic materials, perhaps a new definition for an explosive process may need to be developed (Merril, Nichols & Lee, 1999).
Additional computer modeling It appears that detonations in Class 1.3 composite propellants are unconfirmed (some physicists have stated that such detonations are impossible) rapid deflagration is entirely possible and has been observed experimentally. Class 1.3 presents less of an explosive hazard than its more volatile counterpart graded Class 1.1 (e.g., Trident missile propellant). There is data from flyer plate experiments at Lawrence Livermore National Laboratory (LLNL) that show that “violent” explosions were induced in 1.3 propellant from high speed impacts. The test data indicates these explosions were not from shock induced detonation (i.e., there was no SDT); instead the explosion was a rapid conflagration induced in test samples (Williams, 1987). The Solid Rocket Motor Ground Impact (SRMGI) model is an analytical/engineering code developed by RDA Nance, Crawford et al. (1990) to provide estimates of the effects of solid motor impacts. Previous work for the Air Force and NASA in analyzing the Titan 34D-9 failure, particularly ground impact explosions and blowout fragment ranges, led to the development of this code. It was subsequently expanded to include the aft segment of the Shuttle boosters, and Delta II Graphite Epoxy Motors (GEMs). Major subroutines determine the theoretical Sandia penetration depth (Sandia Labs developed penetration scale), gas generating rate void pressure, fragment accelerations and motion, venting, and final fragment velocities. Key assumptions involve percent increase in burn area and, for the stage impact, the cavity formation time, initial cavity size, and blowout fraction. Maximum fuel burn rates, soil burn area characteristics, and blowout fraction can be parametrically varied. The summary of critical velocities versus soil number and burn rate, as predicted by the modified SRMGI code, are presented in the semi-log plot in Figure 6.37. This figure shows the smaller GEMs require the highest speed (> 400 ft/sec) and hardest surfaces (S < 1.2) to produce runaway explosive. The Titan upper segments could go critical on somewhat softer surfaces at 300e400 ft/sec but require 500e600 ft/sec for soil numbers up to 2.4. Above S ¼ 2.4 the SRMGI code indicates that impacts of the Titan segments should not produce explosive deflagration even if burn rates are as high as 500 in/sec. Both the GEM and the Titan will not go critical if burn rates are less than 8e15 in/sec. The STS aft segments appear to present the greatest hazard. As indicated in the figure, the STS’s large size produced runaway condition for soil impacts from S z 1.4 up to S z 15 to 20 and for maximum fuel burn rates which are only 5e6 in/sec. For hard surfaces (S < 1.4) when the aft stage does not penetrate more than 6 ft (half the stage diameter, the minimum depth assumed in the code for this segment, impacting end on, to create a cavity) then it is assumed conditions would not be created for a runaway deflagration. The code, with all of its engineering approximations and assumptions, only indicates that conditions for runaway deflagration have been created.
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Runaway explosion regions
800
700 Impact velocity (ft/s)
GEM 18 Burn rate = 14 30
TITAN 8 1215 40 500
500
600 STS AFT SEGMENT (min. Br 5 in/s)
500
400
300
200 20
Muck
10
Bay mud
8
Top soil
6
Moist clay, fine sand
Alluvial soil
Packed soil
Hard clay
Dense wet sand
4 2 SANDIA Soil No.
1
Shale
Broken rock
0.8
Soft concrete, asphalt
0.6
Hard concrete, granite
324
FIGURE 6.37 SRMGI code runs for space booster SRMs.
Blast Overpressure Hazard A characteristic feature of explosions is blast. Gas explosions are characterized by rapid combustion in which high-temperature combustion products expand and affect their surroundings. In this fashion, the heat of combustion of a fueleair mixture (chemical energy) is partially converted into expansion (mechanical energy). Mechanical energy is transmitted by the explosion process into the surrounding atmosphere in the form of a blast wave (Lee & Green, 1988). If the combustion process within an explosion is relatively slow, then expansion is slow, and the blast consists of a low-amplitude pressure wave that is characterized by a gradual increase in gas dynamic state variables (e.g., pressure, density, and particle velocity) (Figure 6.38). If, on the other hand, combustion is rapid, the blast is characterized by a sudden increase in the gas dynamic state variables: a shock. Due to extensive military investigations on the destructive potential of high explosives, relating the explosive power of an accidental explosion to an equivalent TNT charge is usually the comparison of choice.
6.3 Launch Abort Environments Affecting SNPSs
Pressure
SHOCK LOW AMPLITUDE PRESSURE WAVE
t
FIGURE 6.38 Blast wave shapes.
TNT Equivalency The use of TNT equivalency methods for blast prediction purposes is quite simple. The available combustion energy is converted into an equivalent charge weight of TNT with the following formula: WTNT ¼ ae
Wf Hf ¼ am Wf HTNT
(68)
where Wf is the weight of the fuel involved (kg), WTNT is the equivalent weight of TNT or yield (kg), Hf is the heat of combustion of the fuel in question (J/kg), HTNT is the TNT blast energy (J/kg), ae is the TNT equivalency based on energy, and am is the TNT equivalency based on mass. If the equivalent weight of TNT is known, the blast characteristics, in terms of peak side-on overpressure of the blast wave, can be derived for varying distances from the explosion. Concerns about this peak overpressure caused by the explosion of either liquid or solid propellant are focused in the near-field where the nuclear source material resides. Far-field effects of explosive blast are more concerned with effects of pressure increases on facilities, equipment, window glass breakage and the safety of personnel in the immediate area of the launch pad or at farther distances where pressure waves can focus.
Blast scaling The upper half of Figure 6.39 represents how a spherical explosive charge of diameter d, produces a blast wave of side-on peak overpressure P, and positive-phase duration tþ, at a distance R from the charge center. Experimental observations show that an explosive charge of diameter Kd, produces a blast wave of identical side-on peak overpressure p, and positive-phase duration Ktþ, at a distance KR from the charge center (lower half of Figure 6.40). Consequently,
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R
d
I
P t+
KR
Kd KI
P Kt+
FIGURE 6.39 Blast wave scaling.
charge size can be used as a scaling parameter for blast. Charge size, however, is not a customary unit for expressing the power of an explosive charge; charge weight is more appropriate. Therefore, the cube root of the charge weight, which is proportional to the charge size, is used as a scaling parameter. If the distance to the charge and the duration of the wave are scaled with the cube root of the charge weight, the distribution of the blast parameters in a field can be graphically represented, independent of charge weight. This technique, which is common practice for high-explosive blast data, is called the Hopkinson scaling law (Hopkinson, 1915). The compilation of experimental data for TNT surface bursts are presented in Figures 6.40 and 6.41. Early tests of liquid oxygen and RP-1 fuel (kerosene) were conducted under project PYRO in the 1960s to substantiate this curve. Project PYRO, conducted in the late 1960s, consisted of a comprehensive program to determine the blast and thermal characteristics of the three liquid propellant combinations in most common use at the time in military missiles and space vehicles: liquid oxygen with RP-1, liquid oxygen with liquid hydrogen, and nitrogen tetroxide with Aerozine 50 (50% UDMH/50% N2H4) (Willoughby, Wilton & Mansfield, 1968). During the course of the program some 270 tests were conducted with these propellant combinations on weight scales ranging from 200 to 100,000 lb. This basic explosive test program was supplemented by analytical and statistical studies, laboratory-scale experimental studies, simulation tests with inert propellant combinations, and a series of high-explosive tests for calibration and evaluation purposes. Several test configurations were used. The Confinement by Ground Surface (CBGS) condition best simulated the
6.3 Launch Abort Environments Affecting SNPSs
25,000 LB DATA LARGE POOL DIAMETER DATA ALL OTHER DATA
400
200
100 80
PEAK OVERPRESSURE (psi)
60
40
20
10 8 6
4
2 Note: The curve is not fit to the data, but rather the TNT equivalent blast curve when the far field equivalent yield is extrapolated to the near field.
1
1
2
4
6 8 10 20 40 60 80 100 SCALED DISTANCE (ft/lb½) Source: Willoughby et al. (1968)
FIGURE 6.40 Project PYRO test: Confined by Ground Surface (CBGS) case.
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25,000 LB DATA 400
ALL OTHER DATA
200
100 80 60 PEAK OVERPRESSURE (psi)
328
40
20
10 8 6
4
2 Note: The curve is not fit to the data, but rather the TNT equivalent blast curve when the far field equivalent yield is extrapolated to the near field.
1
1
2
4
6
8
10
SCALED DISTANCE (ft/lb½)
FIGURE 6.41 Project PYRO test: Confinement by Missile (CBM) case.
20
40
60
80 100
Source: Willoughby et al. (1968)
6.3 Launch Abort Environments Affecting SNPSs
sequential ground impact of separate fuel and oxidizer propellant tanks. This condition could occur by over pressurization of tanks or by the fallback or topple-over of the entire launch vehicle. Mixing interaction between the two propellants involved consideration of two configurations for the CBGS. Examination of the various length-to-diameter (L/D, or L-over-D) ratios and propellant flow velocity combinations revealed that the two propellant masses could have led to two cases of study: when both propellants were vertical, and when both propellants were horizontal. Figure 6.42 represents the Confinement by the Missile (CBM) configuration in which case an internal failure is assumed to occur, and one propellant falls down into the other. Cause of this type of failure could be the rupture of a common bulkhead between fuel and oxidizer tanks due to over pressurization or fragments from
FIGURE 6.42 CBM case setup.
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a primary explosion (e.g., engine blowup) penetrating the bulkhead. To simulate this failure, a test setup as shown in Figure 6.42 was used with a small C-4 charge used to propel the breaker ram through the tempered glass diaphragm separating the oxidizer from the fuel. Duration of the CBM case was limited to the time that the propellants remained confined by the walls of the vehicle. This time was determined by the strength of the tankage, the rate of vaporization of the cryogenic material, the initial pressure in the tanks, and the initial ullage space. The mixing mode in this case is determined by the interaction of the upper propellant and lower propellant at the time of opening. The position and velocity distributions of the lower propellant at the moment of first contact are fairly well specified, since it was assumed to be in its original configuration and to have zero velocity. Those of the upper propellant, however, may have a large range of values, depending on how full the tanks are and how large an opening is created between them. If a relatively small opening is produced, the L/D ratio of the top propellant will be effectively very much greater than that of the bottom one, and its total weight (entering into the mixing at an early enough stage to be a matter of concern) will be much smaller than that for the bottom one. In addition, the top propellant will have an initial velocity given by the fluid head in the top tank and the pressure differential between the two tanks. As the opening between the tanks becomes larger, the L/D ratio and effective weight of the upper propellant more nearly approaches the values they originally had in the missile. In addition, at the time the opening between the tanks is the full cross section of the original tank, the velocity would reach the value given by the acceleration of gravity through the distance of the ullage space in the lower tank, provided there was no pressure difference between the two tanks. Where the opening between the tanks is significantly less than the full cross section of the vehicle, the different initial L/D values for each propellant can be treated by assuming that the L/D ratio refers to the overall vehicle geometry and by defining a Do/Dt ratio, where Do is the opening diameter and Dt the vehicle diameter. The CBGS scenario produces a hemispherical blast while the CBM case generates a spherical shape. When using these graphs, a scaled distance must be determined, equal to K ¼
R W
=
330
1 3
(69)
where R is the distance of interest in feet and W is the combined liquid propellant weight in pounds. Adjustment factors are applied to W to reflect the fraction of terminal yield and the geometry of the explosion wave (i.e., spherical or hemispherical). Empirical data on TNT equivalency of liquid propellant combinations were used to develop Table 6.10 from Air Force Manual 91-201 (AFSC, 2011). The net explosive weight of liquid propellant is the total quantity of fuel (e.g., liquid hydrogen, kerosene (RP-1)) and oxidizer (e.g., liquid oxygen, nitrogen tetroxide) that would mix and explode. Assume a prelaunch explosion of a Delta II first stage occurs at a distance of 69.3 ft from the nuclear power source installed in the spacecraft. Predict the overpressure
6.3 Launch Abort Environments Affecting SNPSs
Table 6.10 Liquid propellant TNT equivalence Propellant combination
TNT equivalence
LO2/LH2 B5H9 þ an oxidizer LO2/LH2 plus LO2/RP-1
60%1 60% Sum of (60% for LO2/LH2) þ (20% for LO2/RP-1) 20% up to 500,000 lb plus 10% over 500,000 lb 20% up to 500,000 lb plus 10% over 500,000 lb 20% up to 500,000 lb plus 10% over 500,000 lb 10% 10% 10% 10% 10% plus equivalence of the solid propellant 100% 100%
LO2/RP-1 LO2/NH3 B5H9 þ a fuel IRFNA and aniline IRFNA and UDMH IRFNA and UDMH plus JP-4 N2O4/UDMH plus N2H4 N2O4/UDMH plus N2H4 plus solid propellants Tetranitromethane (alone or in combination) Nitromethane (alone or in combination)
Recent testing at white sands may lower this value substantially.
environment at this distance assuming the propellants leak to the flame trench of the launch pad forming a pool that is subsequently ignited in a hemispherical explosion. The Delta II first stage propellant load of LO2 and RP-1 equals 212,574 lb from Table 6.5. Using a TNT equivalence rating of 20% for the LO2 and RP-1 combination from Table 6.10, a scaled distance can be calculated using equation (6) as follows: =
W
1 3
¼
69:3 feet ð212; 574 lb 20%Þ
=
R
1 3
ft
¼ 2:0
=
K ¼
1 3
lb
(70)
Looking at Figure 6.40 for the hemispherical case (CBGS) the overpressure at a scaled distance of 2.0 ft is approximately 300 psi. If the destruct system had inadvertently activated, opening the LO2 and RP-1 tanks to afford mixing and ignition at the tanks’ interface, a spherical explosion could be assumed and a factor of 0.5 would need to be applied to the total propellant weight, W in calculating the scaled distance. 69:3 =
ð212; 574 0:5 20%Þ
1 3
¼ 2:5
ft =
K ¼
1 3
lb
(71)
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10 00 STATIC OVERPRESSURE (psi)
332
WSTF PYRO Deep Hole 90%
100
50%
10%
10
1 0.10
10.0
1.0 SCALED DISTANCE
100.0
(ft/lb1/3)
FIGURE 6.43 Static overpressure for MMH/NTO propellant combination.
which relates to an overpressure of approximately 240 psi from Figure 6.40. Experimental data for more recent testing done with nitrogen tetroxide (N2O4) and monomethyl hydrazine (MMH) are shown along with PYRO deep hole results in Figure 6.43.
Impulse pressure The impulse, I, is the area under the positive portion of the overpressureetime response curve (Figure 6.39). This is the more damaging portion of the blast wave, therefore the impulse predicted will be used along with peak overpressure to characterize potential adverse effects of explosive commodities.
Solid propellant yield models Over the years, great effort has been expended on debris risk model input parameters to improve, understand, clarify, or justify the values used for assessing launch hazards. One of the most controversial issues remaining is the estimation of solid propellant impact yield. There are many models and methods available to estimate the explosive yield produced upon impact of the solid propellant at variable velocities and configurations, many of which have been used for Titan SRMUs, Atlas SRBs, and Delta II GEMs. Solid Rocket Motors have consistently driven explosive debris risks to the general public and personnel for launch area and downrange overflight accident scenarios. Solid propellant also tends to drive risks to payloads requiring SNPSs (e.g., Cassini, MERs, PNH, MSL). The Space Shuttle SRB in particular provides the highest launch area explosive debris risk value of all the other launch vehicles for one reason: the aft segment does not include a linear shaped charge (LSC) to split the case as is done for the forward segments. This leaves a contained mass of approximately 260,000 lb of 1.3 HTPB
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.44 STAR-48 destruct test at White Sands Missile Center.
propellant descending to impact should destruct action be necessary during ascent. The short lived Ares replacement vehicle of NASA’s Constellation Program provided for extension of the standard LSC to incorporate the aft segment. However, how do I know my yield model is accurate? Considerable effort has been expended culling available historical data to validate our models, but the truth remains: there is very little controlled testing data. The only true test I know of was the 1964 Titan IIIC segment (~82K lbm 1.3 HTPB) strapped to a sled and slapped end-on into a concrete wall at 667 ft/sec. Models tend to over-predict the yield calculated from that resultant test data. Other yield “data” tends to come from investigations of real incidents when crater dimensions are used to back-calculate appropriate yield. Some small scale data exists from critical diameter testing but it is questionable if that data adequately relates to fallback yield. Nonetheless, all available data has been previously scrutinized, weighed, and factored into the methods of yield prediction. Why not get more data? Since 1964, when the sled test was accomplished, constraints to large scale high explosive testing have increased considerably (environmental limitations, high cost). A STAR-48 full scale destruct test, accomplished at WSMC for the MER nuclear risk assessment was an excellent demonstration of the difficulties and the rewards of this type of challenging data acquisition program (Figure 6.44).
Solid propellant computer modeling Special computer models have been developed to predict the blast environment provided when a solid propellant chunk impacts the ground. The PISCES-2DELK continuum mechanics model, developed by FOILS Engineering, is a hydrodynamics code which was used to envision the impact of a General Purpose Heat Source (GPHS) on sand or concrete followed by a direct impact or near-field impact of a large propellant chunk (Mukunka, 1996). It encompasses the cratering process, the response of the propellant to it, the subsequent fracturing and burning of the
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FIGURE 6.45 GPHS impact modeling via PISCES-2 DELK at 0 seconds.
propellant, and finally derivation of the flow conditions (pressure, density, etc.) radially away from the impact point. Figures 6.45e6.48 depict predictions of the response and deformation of the GPHS to these events.
Orbital Re-Entry Failures that can occur once the spacecraft has achieved orbital velocity can end the mission abruptly or over a longer delay ending in re-entry through Earth’s atmosphere, heating, and eventual breakup and impact.
FIGURE 6.46 GPHS impacting modeling via PISCES-2 DELK at 3.0 ms.
6.4 Containment Design
FIGURE 6.47 GPHS impacting modeling via PISCES-2 DELK at 10.0 ms.
FIGURE 6.48 GPHS impact modeling via PISCES-2 DELK.
6.4 Containment Design The safety issues for radioisotope heat sources and fission reactor based heat generation are fundamentally different and must be considered separately. For radioisotope
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systems, the fuel material is hazardous at launch, and containment of the fuel material, to the greatest degree possible, must be achieved for all accident environments. Radioisotope systems also produce heat continuously and this energy must be managed though all mission phases. For fission reactor systems, the nuclear fuel is relatively nonhazardous at launch and the hazard increases with reactor operating time as the fission product inventory builds up. For reactor systems, the safety approach focuses on avoiding inadvertent criticality ensuring that any fission product inventory is adequately decayed before accidental re-entry. Reactor systems produce heat on demand, therefore simplifying thermal management though the various missions phases. However, reactor systems are inherently more complex devices than radioisotope systems, leading to a wide variety of failure modes that must be considered. The safety approach for radioisotope systems has been well developed through the long history of the use in space applications. Radioisotope systems are designed to contain the fuel though the variety of insults than can occur in launch vehicle accidents and inadvertent re-entry.
Small Heater Containment Design A small application of radioisotope material in spacecraft systems involves placing Light Weight Radioisotope Heater Units (LWRHUs) in locations throughout the space vehicle that can benefit from a constant source of heat to maintain thermal control. The Mars Pathfinder Sojourner utilized three of these to keep its Warm
FIGURE 6.49 LWRHU layers.
6.4 Containment Design
Electronics Box (WEB) above critical temperatures during the cold Martian night. The Cassini spacecraft has 117 LWRHUs located in various places for thermal management (Furlong & Wahlquist, 1999). LWRHUs consist of 2.68 grams of PuO2 in a fuel pellet that produce 1 watt of heat output. Each fuel pellet is encapsulated in a platinum rhodium cladding and encased in a multilayer graphite containment to protect the pellet in the event of an accident (Figure 6.49).
General Purpose Heat Source Module The workhorse of the Radioisotope Thermoelectric Generator (RTG) is the General Purpose Heat Source (GPHS) (Figure 6.50).
RTG/MMRTG structure Containment is at the GPHS level, but is not as necessary at the RTG level. This is actually a benefit since in most accident cases the RTG can be broken apart leaving the individual GPHS modules to fall back to land with impact. Since they are designed for the re-entry heating, as individual “boxes” they are lighter and their areal density can decrease the impact speed (Figure 6.51).
Space Reactor System Containment Design The safety approach for reactor based systems is less developed due to the extremely limited application of this technology to date and the fact that no programs in recent history have gone to flight. Nonetheless, a number of programs in the past 20 years
Cap Aeroshell Cap
Fueled Floating Clad Membrane Fuel Pellet
Graphite Impact Shell
Lock Member
Carbon Bonded Carbon Fiber Sleeve Aeroshell Lock Screw Individual GPHS Module
FIGURE 6.50 General purpose heat source stack.
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FIGURE 6.51 Radioisotope Thermoelectric Generator (RTG).
have proposed the use of space reactor systems (Angelo, 1985) and the safety of such systems had been studied extensively (Ponomarev-Stepnoi, Talyzin & Usov, 2000). At launch, the uranium fuel of a reactor system is relatively non-hazardous as compared to the plutonium fuel for radioisotope systems. As the reactor operates, the fission product inventory builds up increasing the hazard. The safety of reactor systems, therefore, depends primarily on preventing operation in the biosphere and ensuring that there is sufficient decay time after reactor operation before it can return to the biosphere. The first element of ensuring safe operation of space reactor systems is the prevention of inadvertent criticality. Inadvertent criticality must be prevented for both normal credible accident conditions. The primary accident scenarios of concern for preventing criticality are compaction and water immersion. Compaction is a concern in impact accidents where the nuclear fuel could potentially be arranged to a critical configuration. Water immersion is a concern as flooding of the reactor core and reflection provided by water immersion can also lead to a critical configuration. In some cases, the reactor can be designed to be inherently insensitive to compaction and water immersion. If such design is not feasible, then active safety measure such as removable neutron poisons can be employed to address inadvertent criticality concerns. Additional operational measures are employed to prevent a radiologically hot reactor from entering the biosphere. Prelaunch, the reactor can only be operated for low power testing such that negligible radioactivity is produced. After launch, the reactor can only be operated when the system reaches a “sufficiently high orbit.” A sufficiently high orbit is defined as that which the orbital lifetime is adequate to allow for decay of the fission products before re-entry. For reactor disposal in Earth orbit, such disposal must be in a sufficiently high orbit and re-entry of a radiologically hot reactor is precluded.
6.5 Risk Assessment for Nuclear Missions The use of radioactive materials has enabled humankind to pursue scientific research and gain insights into our universe that would have been impossible or impractical by other means. However, these same materials that have proven
6.5 Risk Assessment for Nuclear Missions
so beneficial in outer space can pose a significant hazard to human health and the environment if an accident causes them to be dispersed. For this reason, a rigorous risk assessment process is performed for all launches involving radioactive material. The purpose of a risk assessment is twofold: first, during the design of the mission and the spacecraft, risk assessment enables mission decision makers to identify, assess, and manage possible safety concerns (and, in this context, especially radiological safety concerns) related to the mission. Then, once the mission and spacecraft design have been completed, the results from the risk assessment are presented to the administrative and political decision makers who participate in the launch approval process as evidence to document the safety of the mission. One way to view the risk assessment is as the means of adjudicating the competition between the extreme environments that might be encountered during a launch accident and the robust containment structures that are designed to prevent the radioactive materials from reaching the environment. The risk assessment examines the scenarios that have the potential to cause a release of radioactive material and estimates the likelihood that a release will occur, the characteristics of the release that might be expected, and the effects that the release might have on the public and the environment. Thus, a risk assessment is a broad and far-reaching analysis that is performed by a team with many different skills and tools. So what is risk? Risk can be viewed as the potential for realization of unwanted or adverse consequences. Risk estimates are based on the likelihood and consequences of an adverse event, considering uncertainty. Thus, a risk analyst must provide a rigorous and complete understanding of which adverse consequences can occur, the mechanisms by which such consequences can be produced (scenarios), the likelihood of occurrence for each scenario, the specific magnitude of consequences that would be expected should the consequence occur, and the degree of certainty involved in each of these estimates. For activities related to the use of radioactive materials in outer space, the principal adverse consequences of an inadvertent release of radioactive materials are the potential for environmental contamination and human health effects (especially latent cancer fatalities). It should be noted that spaceflight presents many risks other than those involved with the release of radioactive material. Examples include the use of toxic and flammable materials, asphyxiation in confined spaces, and other common industrial hazards. These risks must also be identified and managed to ensure the overall safety of persons and the environment. However, the focus of this section is specific to the special risks associated with the use of radioactive materials. The basic process for assessing risks for the launch of radioactive materials requires the analyst to: • •
Analyze the launch vehicle and mission to determine the frequency of events that might threaten the radioactive material. Determine the distribution of the potentially threatening environments in terms of their conditional likelihood, magnitude, time of occurrence, and spatial characteristics.
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• • • •
Determine the expected response of the structures intended to contain and isolate the radioactive material to each of the potential threatening environments. Determine the amount and characteristics of the radioactive release, if any. Determine how any released material will be transported through the environment cause contamination or to cause human exposure. Estimate the health effects of these potential radioactive exposures.
To be successful this process must involve the collaboration of experts from a wide variety of technical disciplines working together to understand complex nonlinear phenomena and provide straightforward insights for decision makers.
Characterizing the Insult Environment Spacecraft designers understand the importance of preventing the release of radioactive materials to the environment, so they design structures into the spacecraft to contain and isolate the radioactive material. The job of the risk analyst is to understand the conditions under which the radioactive material might escape these structures. The chief mechanism for escape is the introduction of energy into these structures that is sufficient either to breach the structure or to change the material form of the radioactive material inside to enable it to escape through some other pre-existing pathway. An example of the former is a crack in cladding material, while the latter may occur if heat causes the material to vaporize and escape through a vent. The introduction of energy to radioactive material or its protective structures is often called an “insult.” Energy insults can take the form of atmospheric overpressure, explosive shock, mechanical force, impact by fragments, impact with the ground, or thermal energy from burning propellants or the heat of re-entry. Historically, other sources of energy such as electric power or chemical reaction have usually not been present in sufficient magnitude to challenge the containment of the radioactive material except to the degree that they contribute to the creation of the other energy sources. The potential sources and magnitudes of these insult environments have been described earlier in this chapter (see section on “Launch Abort Environments Affecting SNPSs”). The risk analyst looks for mechanisms that can apply sufficient energy to damage the radioactive material or its protective structures and cause the material to be released to the environment. Understanding the threshold for “sufficient” energy for any individual insult type is extraordinarily challenging. For example, extensive testing and computational simulation programs may be required to characterize the behavior of the protective structures, with the objective of providing risk analysts with the most accurate possible understanding of how these materials would perform in a real accident should it occur. Limitations on testing and computational resources mean that such studies often involve the application of only a single insult type in any experiment or simulation. However, a real accident involving a launch vehicle or spacecraft produces an extremely chaotic environment in which multiple insults of various types might be created in rapid succession. The superposition of results from
6.5 Risk Assessment for Nuclear Missions
separate effects tests into a comprehensive picture of material behavior in this chaotic environment remains a significant challenge for risk analysts. In addition to understanding the type and magnitude of the insult environments that may cause the release of radioactive material from its protective structures, the risk analyst must also understand how those insult environments might come to exist. The analyst must use the list of insult environments which have the potential to result in radioactive releases when examining all aspects of the spacecraft, the launch vehicle, and the mission. It may be possible for accidents to create such environments not only during the launch itself, but also during the preparation of the spacecraft and launch vehicle for the launch event, and during any possible re-entry of the spacecraft into the Earth’s atmosphere. After many years of studying launch vehicle and spacecraft accidents that have occurred, and studying postulated accident scenarios from risk assessment studies for all phases of previous missions, a rich historical database has been developed to relate which insult types should be studied in which mission phases. This list serves as a generic starting point for consideration of insult environments because it incorporates information from a broad cross section of past experience. The potential for a launch vehicle to explode, with its resultant overpressure, possible explosive shock, fragment field, ground impact, and thermal insults must be considered for any mission phase when the radioactive material is installed in a fueled launch vehicle, but the characteristics of these insults (e.g., magnitude and spatial distribution) may vary dramatically depending on when the accident occurs and which events or malfunctions trigger the accident. The risk analyst must not stop, however, with the list of generic insult environments and their causes. Each space mission is a first-ever event using one-of-a-kind hardware that has been designed for a special purpose but never before completely operated as an integrated unit. Therefore, it is possible that some of the generic insult environments may not apply in the “usual” mission phases, and it is likely that unique new accident conditions may exist with the potential to cause insults to the radioactive material or its protective structures. For example, the Pluto New Horizons mission, launched in 2006, marked the first time a U.S. mission used a large solid rocket upper stage in close proximity to an RTG. This mission feature meant that previous assumptions regarding the mission time over which accidents might produce fragments of solid rocket propellant and solid rocket motor cases with the opportunity to impact the RTG had to be changed. In addition, the proximity of the solid rocket motor to the RTG meant that the impact of large high velocity propellant fragments on the RTG was far more likely for this mission than for any previous mission. Thus, the risk assessment team paid special attention to this unusual source of insults in order to ensure that it was appropriately captured in the mission risk profile. Another example of a mission-specific insult environment occurs when radioactive material is contained deep inside of a spacecraft. When radioactive material such as an RTG is highly exposed, the overpressure produced by the deflagration of liquid fuels may tear the RTG away from the spacecraft and allow it to fly
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free. In such situations, the RTG may land well clear of any burning launch vehicle debris, thus reducing the potential for mechanical insults such as the impact of launch vehicle debris or thermal insults from burning solid rocket propellant. When the radioactive material is contained deep inside of the spacecraft, there is significantly more potential for these insults, and also the potential for insults from the structure of the spacecraft itself. These examples demonstrate how the consideration of mission-specific insult environments is critical to be accurate depiction of mission risk.
Accident Likelihood In order to assess mission risks, the risk analyst must understand the insult environment “hazard curve” e that is, the relative likelihood with which insults of each type and magnitude will occur. To accomplish this, a mission accident assessment brings together information about the likelihood of various observed and postulated accident-initiating events with mission characteristics and physical response data so that the risk analyst can understand the circumstances under which each insult type and magnitude will occur. The process begins by identifying scenarios that could be important to mission risk. In general, analysts look for potential situations that might impart kinetic, thermal, or chemical energy to the radioactive material. For example, prior to launch, accidents during vehicle assembly might result in kinetic energy insults from accidental launch vehicle collapse or dropping materials from great heights, or thermal insults form accidental fires or failure to adequately cool RTGs. During launch, vehicle malfunctions may cause the launch vehicle to explode or crash, or they may cause the launch vehicle to fly off-course requiring actuation of the vehicle destruct systems. The vehicle destruct systems may also malfunction directly to cause the launch vehicle either to be destroyed at an unintended time or to fail to be destroyed when destruction is needed. These chaotic environments can result in many different sources of kinetic, thermal and chemical insults to the radioactive materials. Different types of insult environments occur once the spacecraft reaches suborbital or orbital altitudes, as the insult of re-entry heating becomes possible but reduced inventories of rocket fuels reduce the likelihood and severity of explosive insults. The situations under which these insults might occur are identified using systematic probabilistic risk assessment tools such as failure modes and effects analysis, hazards and operability analysis, and fault tree analysis. Tools like event sequence diagrams and event tree analysis are used to systematically explore how each type of accident might progress to specific insult-producing outcomes, or conversely, be mitigated before insults are produced. For example, if a scenario is identified that could produce a thermal insult to an RTG due to failure of cooling systems, the event sequence diagram or event tree would incorporate options for re-establishing cooling via alternative means or removing the RTG from the spacecraft before adverse temperatures could be reached.
6.5 Risk Assessment for Nuclear Missions
The likelihood of each accident initiating event is highly time-dependent not only between mission phases but even within a mission phase. For example, the likelihood of catastrophic failure for solid- and liquid-fueled rocket engines is often much higher during the first several seconds after ignition than at any other time during the mission. These failures, in turn, occur at relatively low altitudes near the launch pad where the falling materials may impact hard surfaces such as concrete and steel structures. Thus, the analyst must carefully account for the time profile of each failure type in order to produce an accurate assessment of mission risks. The frequency of specific accident initiating events is estimated by considering a wide variety of statistical and non-statistical evidence. Detailed examination of past space flight failures provides evidence regarding the overall frequency of such accidents. However, the historical database covers a wide variety of dissimilar launch vehicles and spacecraft, and some of these historical failures may not be relevant to the mission being assessed. Evidence also comes from the probabilistic risk and reliability studies that are performed for the launch vehicle, the spacecraft, and the ground support systems, as well as component and subsystem test data. In many cases, however, these studies and measured data may not be fully applicable to the mission being assessed. For example, a tank failure pressure may have been measured assuming a full load of fuel while the profile for the mission being assessed calls for this tank to be only 70% full. Thus, it is an inexact science to estimate the likelihood of each accident initiating event for each mission’s specific one-of-a-kind characteristics from the sea of partially-applicable data. Analysts accomplish this task using statistical techniques when possible (especially Bayesian analyses), but often must make use of expert judgment techniques because of the disparate nature of the underlying data. A thorough risk mission assessment examines a large number of accidents derived from a large number of possible initiating events. However, the insult environments for the radioactive material are often similar for accidents from many different initiating events, so analysts may group the accident scenarios so that accidents that produce similar event sequences and insult combinations can be analyzed together for their radioactive release potential. The result of this analysis can be thought of as a hazard curve for each insult type. It includes groups of accidents that are expected to produce similar combinations of insults, and a probability distribution for each group of accidents as a function of mission time. It also includes statistical distributions to represent the intensity and location of each insult type (e.g., fragment size, velocity and angular distribution). For U.S. launches, this information is documented in a Launch Vehicle Databook (Rumerman, 1999) that is unique to each mission.
Radioactive Release Characterization A group of accidents often presents the radioactive material and its protective structure with an environment that includes multiple sequential or even concurrent insults with physics that are coupled and nonlinear. The potential for radioactive release
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must be estimated for these complex and chaotic environments on the basis of separate effects data from experimental and computational sources. Where possible, limited sets of combined environment and multiple insult evaluations are performed to help analysts understand how best to use the data from the more numerous and complete separate effects sources. Otherwise, the separate effects data are applied to the material sequentially to predict the effects on the radioactive material and its protective structures. The objective of the analysis is to evaluate whether a particular sequence of insults will cause radioactive material to be released to the environment, and if so, to predict the characteristics of the release that are important to its potential transport through the environment and ability to cause human health effects. The key characteristics that are estimated are the mass of radioactive material released, its particle size distribution, its chemical composition, and the thermal environment of the release. The mass of radioactive material released depends on the performance of both the protective structure and the material itself. Breaches in the protective structures enable material to escape to the environment, and the released mass depends strongly on the size of the breach. But the released mass also depends on the physical form of the radioactive material. For example, material that has been crushed into a powder by a kinetic energy insult can escape through a breach more easily than material that is in the form of a pellet, extruded metal, or large chunks. Also, liquid or vaporized material can escape through pathways that are inaccessible to solids. The risk analysis makes use of models to predict how much material escapes under specific insult conditions, and these models are tied to both experimental and computational results. The scatter in the experimental data for such models can be great, and it is incumbent upon the risk analyst to understand and account for this source of uncertainty. Two other important characteristics of the released radioactive material are its particle size distribution and its chemical form. The particle size distribution is important because small particles (e.g., less than 10 microns in diameter) are easily transported by wind and are readily inhaled by humans, while large particles (e.g., greater than 100 microns in diameter) fall out of the atmosphere quickly and are not easily inhaled. The chemical form is also important for predicting the biological uptake of certain radioactive materials. The size distribution and chemical form of the material can be significantly altered by the insult environment. For example, while the initial form of the material may be a ceramic oxide pellet with very benign distribution and uptake characteristics, a kinetic insult may crush the ceramic to produce very fine particles, and a severe fire environment may vaporize the material and change its chemical form significantly. A final important characteristic of the radioactive release is its thermal characteristics. Materials that are released in a fire environment may be lofted high into the atmosphere and, thus, transported far greater distances in the environment than comparable releases in an ambient temperature environment. Each of these release characteristics can have a significant effect on the behavior of the material as it is transported through the environment and as it affects exposed
6.5 Risk Assessment for Nuclear Missions
humans. Since the insult environment is complex and nonlinear, and since there is significant statistical uncertainty in the release models, the radioactive release problem is not a candidate for closed form solution or even simple approximation. Instead, release predictions can be computed using a tool like a Monte Carlo discrete event simulation tool that explicitly considers the known uncertainties to estimate releases. The tool might select the mission elapsed time at which the accident occurs and position the spacecraft at the appropriate point. It might follow the spacecraft and radioactive material as it flies off-course and/or falls to the ground. It might generate a statistical distribution of fragments for any explosive environment and determine which fragments impact the radioactive material e either in the air of on the ground. It might also apply the appropriate overpressure insult and thermal environment to the material and compute the characteristics of any radioactive releases that may occur. This process can be repeated for a large number of Monte Carlo trials until the distribution of possible releases is sufficiently characterized. This same process is repeated for each group of accidents. This computational method produces far more possible releases than can be assessed in the radioactive material transport or health effects estimation analyses. A frequent characteristic of the set of computed releases is that an overwhelming majority of the releases are very small with characteristics that would be expected to cause minimal land contamination or human health effects. Large releases very rare, but since they may be orders of magnitude larger than typical releases, they can have a profound effect on statistics such as the mean release and mean consequences. Furthermore, these very rare trials often involve a combination of insults that are at or beyond the bounds of the experimental dataset, and are, thus, relatively uncertain. The sensitivity of the mean to computational artifacts and uncertain data is a fact that is difficult to effectively communicate to decision makers.
Radioactive Material Transport The released radioactive material is transported through the environment and may be deposited far from the original accident site. The objective of this part of the mission risk assessment is to understand how each release of radioactive material is transported through the environment, to estimate the exposure of humans to the radioactive material, and to predict the area and magnitude of land contamination. These results depend not only on the characteristics of the release described previously, but also on the weather conditions into which the release occurred and the population distribution around the accident site. The weather conditions into which the release occurs are critical. A release on a day when winds would blow the radioactive plume out to sea will produce minimal land contamination and human health effects. However, the same release on a day with winds blowing toward a population centre may produce very different results. The principal tools used in the radioactive material transport analysis are plume rise and atmospheric transport models that explicitly account for population. The analysis is
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based on the assumption that the distribution of possible weather conditions at the time of the potential accident can be appropriately estimated from the weather conditions from recent years for dates near the potential accident date (as related to the projected launch date). Typical analyses use weather data from the previous several years for the dates in question, resulting in 150e250 days of weather measured conditions as potential conditions into which the radioactive material might be released. Additional data are required to support the plume rise and transport models, such as model parameters related to the degree to which a plume spreads or the speed with which particles of a particular size are deposited from the atmosphere on the ground. There is significant uncertainty regarding which values of particular model parameters produce the best agreement with experimentally observed behavior, so the results of this analysis must reflect these uncertainties for decision makers. A complete analysis of radioactive material transport would evaluate all radioactive releases into all weather conditions using a variety of model parameters. Such an analysis is not possible within current computational capabilities, so a statistical subset radioactive releases and weather conditions must be analyzed. Methods for generating this statistical subset have included pseudo-Monte Carlo sampling and importance sampling. Since the data contain a few rare but large releases and a few weather trials that yield unusually large consequences, the potential for random combination of these rare events in a Monte Carlo analysis can make the mean and other statistics very sensitive to the presence or absence of specific random data pairings. Overcoming these sampling artifacts will require a significant increase in computational power to be applied to these analyses.
Health Effects Estimation A key metric in the assessment of risks for space missions involving radioactive material is the potential for released material to cause human health effects such as latent cancer or early fatality. To compute this metric, risk analysts use methods sanctioned by the International Commission on Radiological Protection (ICRP) (ICRP, 1999, 2005). These methods consider human exposure to radiation on an organ-by-organ basis for exposures related to radioactive materials that get into the body through pathways such as ingestion or inhalation, as well as exposure from radioactive materials outside of the body through pathways such as the “shine” from radioactive materials in the air or on the ground. The methods also consider mechanisms by which materials can be changed from one exposure pathway to another, such as when the material is dissolved in water or when material is resuspended in air after having been previously deposited on the ground. Since specific radionuclides are preferentially deposited in particular organs, the analysis is very material-specific. As stated previously, the vast majority of radioactive releases computed in typical mission risk assessments are very small. This leads to radiological doses to humans that are very small, and are increasingly small as one moves to greater distances from the accident site. However, these greater distances also hold the possibility for a much larger population to be exposed to these very small
6.5 Risk Assessment for Nuclear Missions
radiological doses e often an exceedingly small percentage of normal background radiation. The ICRP method holds that all nonzero radiological doses hold the potential to cause latent cancer, so these very small computable (but likely not measurable) radiological doses are applied to very large numbers of people to produce health effects, and these health effects from extremely small doses often constitute a large fraction of the overall mission risk. An alternative to the ICRP model holds that a radiation dose threshold exists below which the potential for latent cancer is negligible (a de minimis dose). Thus, persons who receive extremely low doses of radiation are not counted in the mission health effects. However, there is no scientific consensus as to whether a de minimis threshold exists and, if it exists, what its value should be. Furthermore, political decision makers are not of the same mind on this subject. For this reason, risk analysts may present results based on numerous hypothetical de minimis threshold values so that one can understand how this technical uncertainty affects the overall mission risk results.
Risk Integration The analysis method described in this paper leads to the definition of several possible accident scenarios, each with many potential radioactive releases, each of which could occur under a variety of weather conditions leading to different land contamination and human exposure potentials, each of which might lead to different of health effect estimates (depending on the value of the de minimis dose selected). These results are then integrated and summarized to provide a predictive risk profile. To accomplish this, the population of accidents, releases, and consequences is sorted and integrated into a complementary cumulative distribution function (CCDF). An example CCDF is shown in Figure 6.52. A point on a CCDF is defined as an ordered pair consisting of a probability (the ordinate) and a consequence value (the abscissa), and is interpreted as the probability that the consequences will exceed this value (thus, it is sometimes referred to as an “exceedence frequency” curve). A CCDF can be generated from a data set as the complement of the integral of the probability density function to a specific point. There are several benefits to displaying risk as a CCDF. First, it visually reinforces the fact that large consequences are very unlikely to occur. Also, when comparing among alternatives, comparing the CCDFs quickly shows differences between the risk profiles that cannot be seen by comparing single values such as the mean risk (e.g., one alternative has a lower likelihood of large consequences but a higher likelihood of accidents involving a small number of consequences). Finally, a CCDF directly answers typical risk-related questions, such as, “What is the probability that at least one person will die as a result of this mission?” and “How bad could a one-in-a-million accident be for this mission?” The answers to these questions are simply points on the CCDF. The integrated risk results may be composed of several CCDFs (e.g., one each for health effects and land area contamination, or one to represent the risks
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Mission Risk for the Pluto New Horizons Mission 1.E-02
Probability of Exceedence
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1.E-03
1.E-04
1.E-05
1.E-06 1.E-05
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FIGURE 6.52 Risk is represented as a complementary cumulative distribution function.
associated with each mission phase, or one for each analyzed de minimis threshold). In addition, mean risk values (computed as loss expectation values, or the integration of probability times consequence) and other single-value estimates (e.g., the probability of at least one heath effect) can be computed and presented to decision makers if requested, although the accompanying discussion must place these in their proper perspective.
6.6 International Protocols and U.S. Environmental Review Several international treaties, agreements, conventions, and principles adopted to date have some relevance to the use of space nuclear systems. Similarly, the environmental process formally established in the U.S. is applicable to the launch and use of space nuclear systems. These sections address these international protocols and the U.S. environmental process as they relate to the use of space nuclear systems.
International Protocols, Principles and Guidelines Applicable to the Launch and Use of Space Nuclear Systems The international community, primarily through the United Nations (UN), has taken an increasingly active role in establishing agreements on safety and environmental issues relating to space activities as well as on nuclear power considerations (Sholtis, Marshall, Bennett, Brown, Usov & Dawson, 2008). During the 1960s through the
6.6 International Protocols and U.S. Environmental Review
1980s, five international treaties were established to address space safety and environmental issues; namely: 1. Treaty on Principles Governing the Activities of States in the Exploration and Use of Outer Space, Including the Moon and Other Celestial Bodies (known as the Outer Space Treaty), adopted October 10, 1967; 2. Agreement on the Rescue and Return of Astronauts and the Return of Objects Launched into Outer Space (known as the Rescue Agreement), adopted December 3, 1968; 3. Convention on International Liability for Damage Caused by Space Objects (known as the Liability Convention), adopted October 9, 1973; 4. Convention on Registration of Objects Launched into Outer Space (known as the Registration Convention), adopted September 15, 1976; and 5. Agreement Governing the Activities of States on the Moon and Other Celestial Bodies (known as the Moon Treaty), which was not signed by the U.S., adopted July 11, 1984. The first four treaties govern activities in space, whether nuclear or non-nuclear, while the Moon Treaty does mention emplacement of radioactive materials on the Moon. In the late 1980s, in the wake of the Chernobyl nuclear power reactor accident in 1986, the UN established two conventions related to nuclear power; namely: 1. Convention on Early Notification of a Nuclear Accident, adopted October 27, 1987; and 2. Convention on Assistance in the Case of a Nuclear Accident or Radiological Emergency, adopted February 26, 1987. Space nuclear systems have received international attention ever since the unplanned re-entry of the former Soviet Union’s Cosmos 954, Radar Ocean Reconnaissance Satellite (RORSAT), with an onboard radioactively hot space reactor, into the Earth’s atmosphere and subsequent impact of the reactor debris onto the Great Slave Lake area in the Northwest Territories of Canada on January 24, 1978. As a direct result of this accident, in 1978 the UN Committee on the Peaceful Uses of Outer Space (COPUOS) established a Working Group on Nuclear Power Sources (NPS) in Space principally under its Scientific and Technical Subcommittee, but also under its Legal Subcommittee (Bennett, Sholtis & Rashkow, 1987). This Working Group on NPS in Space, which has convened periodically since the late 1970s, worked to formulate a Preamble and eleven Principles related to the safe use of NPS in space. That Preamble and eleven Principles were adopted by the UN General Assembly in June 1992 (UN Legal Subcommittee, 1992). A summary of the Preamble and Principles follows:
Preamble The Preamble recognizes that for some missions nuclear power systems are essential and affirms that the Principles only apply to nuclear power systems “. devoted to
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generation of electric power onboard space objects for non-propulsive purposes, which have characteristics generally comparable to those of systems used and missions performed at the time of the adoption of the Principles .” In other words, the Principles do not apply to nuclear propulsion systems or to new space nuclear systems. The Preamble also recognizes “. that this set of Principles will require future revision in view of emerging nuclear-power applications and of evolving international recommendations on radiological protection”.
Principle 1 Applicability of international law e basically states that the use of nuclear power systems will be carried out in accordance with international law.
Principle 2 Use of terms e defines a number of terms, in particular “. the terms ‘foreseeable’ and ‘all possible’ describe a class of events or circumstances whose overall probability of occurrence is such that it is considered to encompass only credible possibilities for purposes of safety analysis.” In addition, the definition of the term “general concept of defense-in-depth” allows flexibility in achieving this goal by allowing consideration of “. the use of design features and mission operations in place of or in addition to active systems, to prevent or mitigate the consequences of system malfunctions. Redundant safety systems are not necessarily required for each individual component to achieve this purpose. Given the special requirements of space use and of varied missions, no particular set of systems or features can be specified as essential to achieve this objective”.
Principle 3 Guidelines and criteria for safe use e this principle sets forth general goals for radiation protection and nuclear safety followed by specific design safety criteria for nuclear reactors and for radioisotope generators.
Principle 4 Safety assessment e requires a “thorough and comprehensive” safety assessment which is to be made publicly available prior to each launch.
Principle 5 Notification of re-entry e requires a timely notification of the re-entry of radioactive materials into the Earth’s atmosphere and provides a format for such notification.
Principle 6 Consultations e requires States providing information under Principle 5 to respond promptly to requests for further information or consultations sought by other States.
Principle 7 Assistance to States e requires States with tracking capabilities to provide information to the Secretary-General of the United Nations and to any concerned
6.6 International Protocols and U.S. Environmental Review
State, and requires the launching State to promptly offer assistance. After re-entry, other States and international organizations with relevant technical capabilities should also provide assistance to the extent possible when requested by an affected State.
Principle 8 Responsibility e States shall bear international responsibility for their use of space nuclear power systems.
Principle 9 Liability and compensation e holds the launching State and the State procuring such a launch internationally liable for any damage, including restoration “. to the condition which would have existed if the damage had not occurred”. Compensation includes “. reimbursement of the duly substantiated expenses for search, recovery and clean-up operations, including expenses for assistance received from third parties.”
Principle 10 Settlement of disputes e disputes “. shall be resolved through negotiations or other established procedures for the peaceful settlement of disputes, in accordance with the Charter of the United Nations.”
Principle 11 Review and revision e requires that “These Principles shall be reopened for revision by the Committee on the Peaceful Uses of Outer Space no later than two years after their adoption.” Principle 3 is perhaps the most important and most controversial of these Principles. During the adoption of the Principles, the U.S. delegation formally expressed reservations about the technical validity of these Principles to the United Nations. For example, in Section 1.3 of Principle 3, dose limits were established for accidents. Although the dose limits did not apply to “low probability accidents with potentially serious consequences” the U.S. delegation pointed out that Principle 3 should address risk (probability of an adverse consequence) rather than numerical dose limits (Lange, 1991a). In particular, the U.S. delegation made the point that “. this modification, by taking into account the probabilistic concept of risk, which is a central feature of a thorough safety assessment, relates the recommendation directly to the well-proven [space nuclear power system] practices of the United States” (Lange, 1991b). “In November 1990 the International Commission on Radiological Protection published new recommendations in the form of ICRP-60, which supersede the approach taken in Principle 3 when it was developed earlier that year” (Smith, 1992). The International Atomic Energy Agency (IAEA) independently supported the U.S. position by stating that “The sole use of the individual-related dose limits, rather than the complete ICRP system of radiation protection (including sourcerelated constraint), is, in the Agency’s view, inappropriate and does not conform
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with the aims of the ICRP recommendations. Secondly, as the ICRP has recently issued new recommendations on dose limitation . It might, therefore, be problematic to issue guidelines and criteria of safe use of [nuclear power systems] in outer space that would be outdated from their inception” (IAEA, 1991). The United States also proposed clarifying language in 1991 with the statement that “We believe that this clarification removes any doubts as to the intent behind the application of the term ‘defense-in-depth’. As was clear at the time that the Legal Subcommittee reached consensus on this principle, the Subcommittee did not intend to apply the terrestrial standards as such to space systems.” In 1991 the U.S. delegation proposed changes to the wording of Section 3.2 (relating to re-entry of radioisotope sources) “. to take into account the fact that the probability of accidental re-entry from a hyperbolic or highly elliptical orbit can be reduced to a very low value by mission design and operations” and to recognize “. the fact that the practical design objective for RTG containment systems is localization rather than zero release under all circumstances, and that there are practical limits from a costversus-risk standpoint on ‘complete’ clearing of radioactivity by a recovery operation” (Lange, 1991b). At a meeting of the Special Political Committee on October 28, 1992, the U.S. representative stated “The United States did not block the consensus recommendation of the Committee to forward the principles to the General Assembly, nor will the United States oppose their adoption here. On some points, however, it remains our view that the principles related to safe use of nuclear power sources in outer space do not yet contain the clarity and technical validity appropriate to guide safe use of nuclear power sources in outer space. The United States has an approach on these points which it considers to be technically clearer and more valid and has a history of demonstrated safe and successful application of nuclear power sources. We will continue to apply that approach” (Hodgkins, 1992). Although the UN Principles are non-binding, the United States and other nations are working together to establish generally acceptable international safety and environmental guidance relating to space nuclear power. International cooperation in space enterprises and the voice of the international community is expected to increase during the twenty-first century (Sholtis, Marshall, Bennett, Brown, Usov and Dawson, 2008).
The U.S. Environmental Process for the Launch and Use of Space Nuclear Systems The National Environmental Policy Act, or NEPA, was enacted within the U.S. in 1970 (NEPA, 1970 et seq). It requires all U.S. federal agencies to formally consider and document the consideration of planned programs, projects, actions, and activities that could have a significant impact on the quality of the human environment. Planning and pursuing a space mission which would be launched and operated with nuclear materials or systems onboard could potentially trigger the NEPA process (Sholtis et al., 2008).
6.6 International Protocols and U.S. Environmental Review
NEPA implementing guidance and procedures provide detailed guidelines for programs and projects to ensure NEPA requirements are met. The NEPA implementing guidance and procedures require that planned federal programs, projects, actions, and activities with potentially significant impacts complete an environmental analysis sufficient to support and document a formal decision on the matter. Moreover, NEPA requires that this environmental analysis be subjected to public comment, and that any decision in the matter must be made and documented after public comment and responses to those comments have been addressed. Lastly, NEPA requires that a formal, documented decision in the matter must be completed before any substantial expenditure of funds and any irrevocable environmental impacts occur. This required NEPA environmental analysis can be in one of two forms: (1) An Environmental Assessment (EA); or (2) an Environmental Impact Statement (EIS). An EIS involves much more in-depth analyses than an EA, but both must be prepared in Draft form, for public comment, and subsequently in Final form, where all public comments are explicitly addressed. In summary, the EA is a concise public document that provides sufficient information and analyses to determine if a more involved EIS needs to be prepared. Figure 6.53 provides an overview of the NEPA compliance flow process for a planned action involving nuclear materials and shows the trigger points for development of an EIS as opposed to an EA. If, after a Draft EA is completed, subjected to public comment, and updated as a Final EA, it is determined that there are no significant environmental impacts, then a Finding Of No Significant Impact (FONSI) is prepared and published in the Federal Register; NEPA compliance is complete at this point. If, on the other hand, a determination is made that significant environmental impacts do potentially exist, then an EIS is prepared e first as a Draft EIS, and subsequently after public comment, as a Final EIS. An EIS can be done without an EA preceding it if it is reasonably certain that an EIS will be required. See Figure 6.53. An EIS is a detailed and rigorous document, which provides a full discussion of potentially significant environmental impacts, based on the best information available at that time. It presents a discussion of the purpose of the proposed action and details the need for the action. The EIS provides a general understanding of how the action will be implemented, and what potential environmental impacts, both positive and negative, could result. It is also designed to inform decision makers and the public of a reasonable range of alternatives that are compatible with the purpose and need of the action. These alternatives must be compared with the proposed action in terms of their potential environmental impacts. In addition to alternatives, the EIS must also consider the no-action alternative. The EIS undergoes two public comment periods. The first comment period follows a Notice of Intent (NOI) to prepare an EIS which is published in the Federal Register, with a request for comments on the scope of the project. The NOI is also sent to any person or group who has indicated an interest in the project. The second public comment opportunity is a review of the Draft EIS (DEIS). Copies of the DEIS
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START
Sufficient Nuclear Material or Controversy?
NO
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Environmental Impact Statement (EIS)
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END
FIGURE 6.53 NEPA compliance process.
are sent to anyone who responded to the NOI and to any others who have indicated an interest in the project. Additionally, a Notice of Availability (NOA) is published in the Federal Register. Also, a DEIS is provided to any interested or involved federal and state agencies for review. Anyone can provide comments on the DEIS. Comments received are considered by the sponsoring agency, and agency responses are published in the Final EIS (FEIS). At the conclusion of the EIS process a Record of Decision (ROD) is prepared and filed, documenting the agency’s decision on whether to go forward with the action, an alternative, or no action. An EIS for space nuclear missions must provide a detailed analysis of the impact of a normal launch on the air, water, plants, etc. around the launch area. It must also discuss a reasonable range of postulated accidents and their potential impacts. Alternatives considered can include alternate launch vehicles, alternate power sources, and alternate flight trajectories. The launch radiological impact analysis for credible accidents at the launch site, downrange, and inadvertent re-entry from failure to achieve Earth orbit is based on best available information with respect to launch
6.7 Nuclear Mission Launch Approval
vehicle accidents, accident probabilities, the resulting potential accident environments, the nuclear system responses to those environments, and an assessment of the potential health impacts. Postulated re-entry accidents following Earth orbit insertion are also considered. The analysis is a combined effort of the launch vehicle owner, the launching organization (NASA or DOD), the Department of Energy, and other involved government agencies and supporting contractors. Although the analysis performed for an EIS is similar to those performed for the U.S. launch approval process for space nuclear missions, it is separate and distinct from it due to the different timing, purpose, and specifics of each process. For missions like MARS PATHFINDER that use Radioisotope Heater Units (RHUs) with no Radioisotope Power Systems (RPSs), either an EA or an EIS is required. The decision on which will satisfy NEPA compliance requirements is based on a number of factors, most importantly whether it is likely that the process will end with a FONSI. If the EA is completed and it is determined that there is a potential for significant environmental impacts, an EIS will be prepared. As already mentioned, the NEPA process differs significantly in several ways, and is wholly separate from the U.S. launch approval process for space nuclear missions. Because the NEPA process is completed early in the program cycle, the agency may have to use preliminary data and analyses instead of the later, more specific data or more refined analyses used for the launch approval process. In addition, the NEPA process differs in that it involves the public and a large number of federal, state, and local agencies. The U.S. launch approval process for space nuclear systems is not subject to NEPA. Instead, it involves an internal government process, with independent review and evaluation, ultimately leading to a launch decision by the White House; see section “Nuclear Mission Launch Approval” below.
6.7 Nuclear Mission Launch Approval The effects of a release of radioactive materials into the environment from an accident involving a space mission are not limited to quantifiable consequences such as land contamination and potential cancer fatalities. Such a release could also be a serious political matter for the country sponsoring the mission. In order to ensure that both the technical and political risks are appropriately weighed when such missions are authorized, the launch approval process for missions involving the use of radioactive materials should involve both independent scientific reviews of mission risks and participation by political leaders in the launch approval process. The United States formally established such a process in the form of a Presidential Directive in 1977 (PD/NSC-25, 1977). This directive establishes a launch approval process that requires approval from key political leaders following an independent assessment of mission risks. In order to facilitate this process, the directive establishes an Interagency Nuclear Safety Review Panel (INSRP) for each mission involving the use of radioactive
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materials. The INSRP serves two functions: first, to provide expert independent review of all nuclear-related safety and risk assessments performed for the mission, and second, to provide an independent estimate of the nuclear-related risks for use by decision makers in the launch approval process. The INSRP for each mission is an ad hoc panel formed at the request of the agency sponsoring the mission. The panel and its scientific and technical support personnel use the methods described previously in this chapter (see Section “Risk Assessment for Nuclear Missions”) to review information provided by the mission project team such as the mission characteristics, risk assessment methods, physical and structural assessment methods, and the supporting data on which the nuclear risk assessment will be based. A formal but robust technical dialogue between the INSRP and the mission project team ensues to ensure that the INSRP has ample opportunity to probe the methods, data and results of the mission project team, to perform independent analyses, and to ensure that safety questions and issues are resolved. The safety review and risk assessment process results in two independent evaluations of mission risk: a Final Safety Analysis Report from the mission project team, and a Safety Evaluation Report (SER) from the INSRP. Both of these documents are presented to decision makers for use in a multistep mission launch approval process that culminates in the Office of Science and Technology Policy (OSTP) in the Executive Office of the President of the United States. Launch approval authority rests with the President, although he generally authorizes his Science Advisor as the decision making authority for launch approval. The launch approval process generally takes at least 6 months. Both the risk assessments and the launch approval are generally valid only for a limited period of time (e.g., 1 month) in part because the assessments are based on specific weather conditions that are unique to this time interval. The analyses, reviews, and launch approval are all based on the assumption that the launch vehicle and spacecraft will meet specific quality standards. However, the launch vehicle and spacecraft will not be completed until mere hours before launch. Thus, the launch approval decision comes with conditions attached e namely that the launch vehicle will be prepared in such a way that mission personnel can certify that the necessary quality standards have been met.
Launch Vehicle Certification for Flight Besides the payload safety review requirements for a spacecraft that utilizes a SNPS, considerations have to be made for the launch vehicle chosen to provide “the ride” to the destination. The most valuable validation of the selected launch vehicle’s strengths and weaknesses comes from the program’s launch vehicle databook. This probabilistic risk assessment (PRA) based book provides detailed information about each of the launch vehicle subsystem failure modes and reliability values. The effects of these failure modes are then examined further in the Safety Analysis Report (SAR) by developing release scenarios and ultimately calculating radioisotope exposure risk. The many existing safety requirements and regulations that exist for launching
6.8 Nuclear Mission Launch Integration
a “normal” payload apply and mission-unique “special” requirements are usually listed in these. For the most part, however, these regulations lead a Range User, preparing to bring a radioisotope system to the launch area for assembly, test, and finally launch to one place e the aforementioned PD/NSC-25 requirement to empanel an INSRP and begin a thorough review of the program’s nuclear payload mission. As an example, NASA’s Procedural Requirements (NPR) document for General Safety Program Requirements (NPR 8715.3, latest version) provides details on how NASA complies with the necessary requirements of PD/NSC-25. NASA Policy Directive 8610.7 (latest version) provides requirements for selection of a launch vehicle to transport varying classes of payloads into orbit/space. Increased risks (based on hazard, cost, national interest) required increased scrutiny of launch vehicle design and greater flight experience to mitigate those risks.
6.8 Nuclear Mission Launch Integration The hazardous nature of the radioisotope-laden payload introduces unique challenges to the normal spacecraft processing and integration process. The rigorous safety review process that exists for “normal” payloads is followed for nuclear missions as well; however, one simple principle tends to apply e introduce the radioisotope system at the launch vehicle processing facility as late in the prelaunch schedule as possible.
Design Influence This basic assumption does not come without a price however, since design, fabrication, test, and integration of often specialized attach mechanisms to accommodate a late arrival requires ingenuity, schedule availability, and valued spacecraft mass allocation. The investment is rewarded though, especially in minimizing overall “exposure” time of launch processing workers required to conduct tasks in proximity to the nuclear system’s appointed spacecraft real estate. Radioisotope systems can be mounted on booms to provide distance from sensitive instrumentation. This also allows more efficient radiative thermal control by having heat bearing surfaces exposed to the space environment. Simple radioisotope heater units tend to be “buried” in equipment such as the Warm Electronics Box (WEB) on the MER Rovers Spirit and Opportunity. Thermal cooling loops are utilized to transfer the decay heat as necessary to maintain internal component temperatures within critical functional ranges.
Potential detrimental effects of destruct The normal design of Range Safety Systems (RSSs) are driven by the need to (a) halt thrusting rocket engines (solid or liquid) in order to control a failing vehicle’s terminal impact point; and (b) disperse propellants as necessary to minimize explosive yield events upon impact with the ground. Since the main idea is to destroy the pressurization capability of a thrusting vehicle, there is very little concern for the consequences to other, mission-necessary subsystems (e.g., telemetry, solar power, command
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computers) since once a need arises that warrants activation of the destruct systems, their utility becomes rapidly less important in the ensuing explosive consequence. It is a very different story for payloads that maintain a radioisotope component. One can argue that any final data that can be capture via telemetry downlink is invaluable to resolve the “What happened” investigation that will follow a launch anomaly. Conversely, the vulnerability of an onboard SNPS must be considered so that an action meant to mitigate risks to the public, personnel, and critical resources from falling debris, toxic effluent exposure, and distant focusing overpressure hazards does not consequently expose the radioisotope containments systems to extreme environments. Range safety destruct system ordnance components provide penetration energy to defeat tankage and break motor casings. However, what occurs at the “non-business end” of a conical shaped charge (CSC) can be hazardous to spacecraft mounted RTGs. Very small metallic casing pieces of the CSC can translate from the back plane of the device at near-rifle velocity and the proximity of spacecraft systems leave them vulnerable to impacts. Heavier mounting bosses and brackets can also be expelled at slower velocities but the higher “beta” value makes them also very lethal. The additional concern for evaluation of side effects from flight termination activation is very similar in the world of human spaceflight. Much has been learned about the chaotic nature of vehicle breakup and explosion effects to space vehicle crew which translates well to the nuclear payload sensitivity issue. Difficulties in understanding the near-field effects of liquid and solid propellant explosions can greatly increase uncertainties in modeling these abort scenarios to calculate overall risks.
Launch Site Processing Issues After a spacecraft has completed assembly and test, normally at the owners offsite facility (e.g., JPL, APL, Boeing, Lockheed), it is shipped in whole or in part to the launch area. For nuclear payloads the utilization of radioisotope materials adds constraints to the already hectic processing schedule. Special permitting, licenses, and security needs can accompany these payloads. Additional personnel hazards are included since many tasks require workers to approach the radioisotope hazard zones in order to carry out their unassociated subsystem work. Normally used protective equipment may be newly vulnerable to the inherently high temperatures that exist on the surfaces of the radioisotope systems. One example is the standard Self-Contained Atmospheric Protection Ensemble (SCAPE) worn by propellant handlers. These suits are not designed for use around the high temperature system so special procedures are used and care has to be taken when required to be near the radioisotope sources. Radiation zones are set up to indicate exposure limitations and special personnel access and monitoring procedures are used. Dosimetry badges can provide tracking and warning of overexposure. One of the greatest advantages of transferring the radioisotope components to the spacecraft late in the flow is that many of the most hazardous processes have been completed. Launch vehicle electrical connections have been completed and tested, ordnance devices have been installed and verified, propellant loading has been tested
6.8 Nuclear Mission Launch Integration
so tanks have been through a cycle validating integrity. Catastrophic accidents, such as the Brazil explosion, which could be extremely detrimental to the radioisotope system, are much less likely to occur.
Launch Day Concerns As the final countdown approaches, missions utilizing radioisotope materials will have enhanced requirements to monitor early failures. Tracking the position of the launch vehicle along its trajectory is accomplished via human observation, metric optics (i.e., cinetheodolites), onboard guidance and navigation (inertial measurement unit) data transmitted to the ground, and radar information that is fed into a tracking solution. Except for the visual methods (human observation and cinetheodolites) the tracking methods tend to provide a point estimate of the vehicle’s position in three dimensional space. This information is compared to flight constraints (destruct lines) set to mitigate risk to the public and personnel. Mission Flight Control Officers (MFCOs) monitoring the launch vehicle behavior initiate the RSS if prescribed criteria are met that include shutdown of liquid rocket engines and, as necessary, activation of onboard destruct ordnance. The area within a calculated risk contour in which debris would possibly fall during a launch anomaly, termed the Flight Caution Area (FCA) is cleared of all public and personnel. Only a few Launch Essential Personnel are allowed to remain inside this area and only if the predicted risk is within acceptable limits and the purpose is “essential” to support the launch. Unique to nuclear payload missions, observing the launch vehicle longitudinal orientation becomes very important. During a typical launch, an accident that might deliver debris impacts to the ground around the launch pad certainly could happen, but the limits of that area are predictable and cleared of people to mitigate any risk. For a nuclear mission, one of the most devastating accident scenarios that can lead to a maximum release of radioisotope material involves a full stack intact impact (FSII), especially in a nose down configuration. The payload would impact the ground first, delivering extreme forces to the containment systems, followed by the remaining structure of the rocket and subsequent explosions and fires. While the containment devices are built robustly, there can be a low probability of failure and the additional exposure to high temperature solid propellant fire can lead to vaporization of the material. To mitigate this risk it is important that an accident condition is quickly recognized and MFCO action is taken to initiate the vehicle destruct system preventing the FSII (Botts, 2003). The destructed case also can provide risk of radioisotope release and subsequent vaporization but to a much less extent. The Range Safety Advisory System (RSAS) is used to recognize the potential of anomalous flight behavior that may drive the launch vehicle to a FSII condition. During the early phase of launch ascent (T-0 to approximately Tþ50 seconds) MFCOs monitor a display (Figure 6.54) that provides a visual depiction of the longitudinal attitude as the vehicle lifts, initiates a pitchover maneuver downrange, and
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Atlas
RSAS UNCLASSIFIED
Data Loss
Analog
Guidance
Total Rotation
Mission Time
0
Vehicle Status
0.0
45
Vehicle Attitude Centaur Cleared Tower
Atlas Angle From Nominal
90
0.0
Alarm Status
Acceleration
Tone Alarm Active
135
Tone Alarm override
180
Atlas Rotation
0.0
Tone Alarm Test
0
45
Main Engine L
90
+6
+6
135
+4
180
+2
Chamber Pressures
+4
1
4 +6 +4 +2
Main Engine R
–2 –4 –6 –2
L
R
–4
2
3
–6
+2
5 –6 –4 –2
+2 +4 +6 –2 –4 –6
FIGURE 6.54 Range Safety Advisory System display.
progresses from over land flight to ocean overflight. A deviation of the real-time longitudinal vector that reaches a delta value of 45 degrees from the planned nominal direction indicates that the failure in progress has reached a point that is unrecoverable. The RSAS triggers an audible alarm and visual indication that destruct action is necessary. MFCOs visually verify the alarm condition via camera observation and human observer confirmation and follow quickly with RSS activation. Additional cues can be provided to the MFCO by observing the launch vehicle rocket nozzle gimbal angles (also shown). A failure of the guidance system or mechanical actuators might be indicated if the nozzle thrustline indicator (white dot) moves to an extreme angle and remains there. This tells the MFCO to expect a pitch/yaw response from the vehicle that can lead to a FSII condition. Core and
6.9 Symbols and Acronyms
attached booster rocket chamber pressures indicate any loss of thrust conditions that can cue the MFCO to an expected pitch/yaw response. Based on thrust to weight (T/ W) capability of the vehicle, a loss of thrust to critical rocket motors may indicate a fallback is imminent. Once the vehicle and payload have traveled downrange sufficiently to confirm only ocean impacts would occur in a flight failure, the RSAS can be shut down and more traditional long range tracking becomes primary.
Radiological Contingency Concerns Anomalies can occur during the period from final integration of the SNPS through initiation of transfer to an Earth escape trajectory and, if the mission includes, through possible return to Earth (to gain gravitational boost) until the nuclear payload is bound for the extraterrestrial destination. Planning for these accidents and the radiological consequences is a large part of extra effort that takes place. Impacts near the launch pad and launch range would be initially handled by the Launch Emergency Operations Center (LEOC) and the Radiation Contingency Center (RADCC). The LEOC would act to cordon the area preventing entry so no one can approach any of the harmful toxic, explosive or, in this case, radioactive materials. Off-site detection, warning, emergency response, and recovery would be managed via the RADCC with representatives from range, program, local, state, and federal emergency agencies that have established procedures to mitigate risks to the general public. RADCC personnel can also initiate U.S. state department responses to alert other nations of possible or pending re-entry failures. Efforts following the initial response to mitigate risk to humans would focus on locating, assessing, containing, and transporting any SNPS components in the cordon area. Longer term cleanup operations may be necessary for very extreme accident cases.
6.9 Symbols and Acronyms a Dh ε εAl εg εT 3 T q ma ms r ra s sc
entrainment coefficient change in enthalpy effective emissivity of fireball surface emissivity of aluminum structure effective gas emissivity turbulent energy dissipation rate volumetrically-averaged dissipation rate vortex azimuthal position within fireball dynamic viscosity of ambient air microsecond fireball density density of ambient air Stefan–Boltzmann constant ¼ 5.67 10–12 critical geometry
• Watts • • • cm2/sec3 cm2/sec3 radians g/cm,sec g/cm,sec g/cm3 g/cm3 W/cm2,K4 cm
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! y yq yr c J V
a A A AFSPS Al2O3 AP Ap APU ASRG atm b C cal CALE CBGS CBM CCDF Cd Ci cm COPUOS cp CPIA CTPB D dc de DEIS DELK Dt E Eb EIS EM
gas velocity vector angular component of vortex velocity vector radial component of vortex velocity vector dynamic shape factor of particle stream function del operator for cross product
cm/sec cm/sec cm/sec • • •
constant based on Reynold’s number surface area of spherical fireball cross-sectional area of propellant Affordable Fission Surface Power System aluminum oxide projected area of fireball effective projected area of particle Auxiliary Power Unit Advanced Stirling Radioisotope Generator atmosphere constant based on Reynold’s number number concentration of particles calorie C-language based Arbitrary Lagrangian–Eulerian Confined By Ground Surface Confined By Missile Complementary Cumulative Distribution Function drag coefficient Curie centimeter Committee On the Peaceful Uses of Outer Space [UN] constant-pressure specific heat Chemical Propulsion Information Agency carboxy-terminated polybutadiene maximum fireball diameter critical diameter diameter of volume-equivalent sphere for particles Draft EIS [PISCES-2] dimensional Eulerian–Lagrangian Coupled maximum fireball diameter enthalpy inflow from air entrainment blackbody emissive power (heat flux) Environmental Impact Statement electromagnetic
• cm2 cm2 • • cm2 Cm–3 • • atm • Cm–3 cal • • • • • Ci cm • J/mol,K • • feet cm cm • • feet Watts kW/m2 • •
6.9 Symbols and Acronyms
f FCA FEIS FONSI FPSE FPY FSAR FSII ft FTS FY g Gd2O3 GEM gnd GPa GPHS ha H2 H2O 2 ha hAl hf hF hp hr HTNT HTPB IAEA ICRP Isp IUS J K kbar Kd Ktþ kW kWe kWt L LANL
emissivity correction factor Flight Caution Area Final Environmental Impact Statement Finding of No Significant Impact Free Piston Stirling Engine full power year Final Safety Analysis Report Full-Stack Intact Impact feet Flight Termination System fiscal year acceleration due to gravity ¼ 981 gadolinium oxide Graphite Epoxy Motor ground gigapascal (109 pascal) General Purpose Heat Source enthalpy of ambient air hydrogen hydrogen peroxide enthalpy of entrained air heat transfer coefficient between Al structure and fireball molar enthalpy of fireball enthalpy of formation enthalpy of product species specified enthalpy of the reactants TNT blast energy hydroxy-terminated polybutadiene International Atomic Energy Agency International Commission on Radiological Protection specific impulse Inertial Upper Stage joule (kg$m2/sec2) Kelvin temperature kilobar explosive charge diameter blast wave positive-phase duration kiloWatts kiloWatts electric kiloWatts thermal fireball energy loss Los Alamos National Laboratory, New Mexico
• • • • • yr • • ft • yr cm/sec2 • • • GPa • J/mol • • J/mol W/cm2,K J/mol Watts J/mol J/mol J/kg • • • sec • J K kbar msec kW kW kW Watts •
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Lb lbf lbm lb LEOC LH2 LLNL LO2 LSC LWRHU m MEPAG MER mf MFCO MHW MMH mol ms MSFC N2H 4 N2O4 na n_ a n_ p n_ r NAC NASA NEPA NERVA nf NOA NOI 237 Np np Np Nr Ns NSC NTO NTP
effective thickness of fireball (mean beam length) pounds force pounds mass pounds Launch Emergency Operations Center liquid hydrogen Lawrence Livermore National Laboratory, California liquid oxygen Linear-Shaped Charge Light Weight Radioisotope Heater Unit meter Mars Exploration Program Analysis Group Mars Exploration Rover (Spirit and Opportunity) mass of fuel Mission Flight Control Officer multi-hundred Watt monomethyl hydrazine mole millisecond Marshall Space Flight Center, Alabama anhydrous hydrazine nitrogen tetroxide quantity of entrained air molar rate of air entrainment molar product production rate molar combustion rate of propellant reactants NASA Advisory Council National Aeronautics and Space Administration, USA National Environmental Protection Act, USA Nuclear Engine for Rocket Vehicle Applications quantity of combustion products and entrained air of fireball Notice of Availability Notice of Intent neptunium quantity of combustion products number of product species number of reactant species number of particle sizes National Security Council nitrogen tetroxide Nuclear Thermal Propulsion
cm lbf lbm lb • • • • • • m • • kg • • • mol ms • • • moles moles/sec moles/sec moles/sec • • • • moles • • • moles • • • • • •
6.9 Symbols and Acronyms
OMS OSTP p p PBAN Pb–Te PD PERMS PISCES-2DELK PIRAT PNH psia Pu 238 Pu 238 PuO2 PVF2 pwr q qAl r r R R RADCC rc Re RHU ROCETS RORSAT
RPS RSAS RSS ry s SCoRe-S SDT sec SER
Orbital Maneuvering System [Executive] Office of Science and Technology Policy pressure perimeter of propellant polybutadiene acrylonitrile lead telluride Presidential Directive Propellant Energetic Response to Mechanical Stimuli Physics Int’l multi-Spatial Code for Engineering and Science Propellant Impact Risk Assessment Team Pluto New Horizons pounds per square inch absolute plutonium plutonium-238 isotope plutonium-238 oxide polyvinylflouride power thermal radiation flux heat added to fireball from aluminum combustion fireball radius burning rate reactant enthalpy inflow universal gas constant ¼ 82 Radiological Control Center, NASA Kennedy Space Center truncated fireball radius Reynold’s number Radioisotope Heater Unit ROCket Engine Transient Simulation Radar Ocean Reconnaissance Satellite (U.S. name for Russian UPracm>fn9k SPutojl Altjco9k; Managed Satellite Activities) Radioisotope Power System Range Safety Advisory System Range Safety System vortex radius second Sectored Compact Reactor-Small Shock to Detonation Transition Second Safety Evaluation Report
• • atm cm • • • •
• • psia • • • • kW/m2 Watts cm in/sec Watts cm3,atm/mol,K • cm • • • •
• • • cm s • • S •
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SiGe SNPS SOPHY SPT SRB SRG SRM SRMGI SRMU SSME Stg STS t T T* Ta TAl tb tc Tc TE Th tT TVC 2DELK u 235
U UDMH UN UNCOPUOS UO2 U.S. V VLS W Wa Wf WG Wk Wp Wr WSMC
silicon germanium Space Nuclear Power System Solid Propellant Hazards Program Seebeck–Peltier–Thompson Solid Rocket Booster Stirling Radioisotope Generator Solid Rocket Motor Solid Rocket Motor Ground Impact [model] Solid Rocket Motor Upgrade (Titan IV) Space Shuttle Main Engine Stage Space Transportation System (Space Shuttle) time temperature T / 1000 ambient temp surface temperature of aluminum structure fireball liftoff time stem liftoff time cold temperature thermoelectric hot temperature duration of visible radiation Thrust Vector Control [PISCES]-2 dimensional Eulerian–Lagrangian Coupled fireball rise velocity uranium-235 isotope unsymmetrical dimethylhydrazine United Nations United Nations COPUOS uranium oxide United States fireball volume Veı´culo Lanc¸ador de Sate´lites (Brazilian Launch Vehicle) weight of fuel and oxidizer combined molecular weight of air ¼ 28.97 mean molecular weight of fireball mixture Working Group species k molecular weight mean molecular weight of products mean molecular weight of reactants White Sands Missile Center, New Mexico
• • • • • • • • • • • • seconds K K K K sec sec • • • sec • • cm/s • • • • • • cm3 • lb g/mol g/mol • g/mol g/mol g/mol •
6.9 Symbols and Acronyms
wt WTNT XDT y ya yp Yr z
weight TNT equivalent weight Unknown Detonation Transition molar quant. of comb. products produced per mole of reactant molar air fraction (na/nf) mole fraction of product species year height of fireball center
• • • • • • cm
References AFSC. (2011). Explosives Safety Standards. Kirtland Air Force Base, NM, USA: Air Force Safety Center. Retrieved February 2013, from http://www.af.mil/shared/media/epubs/ AFMAN91-201.pdf. AICE. (1994). Guidelines for Evaluating the Characteristics of Vapor Cloud Explosions, Flash Fires, and BLEVES. American Institute of Chemical Engineers. NY: Center for Chemical Process Safety. Angelo, J. A. (1985). Space Nuclear Power (Original Edition ed.). Malabar, FL: Orbit Book Company, Inc. Bennet, G. L. (1989). A Look at the Soviet Space Nuclear Power Program. Proceedings of the 24th Intersociety Energy Conversion Engineering Conference, IECEC-89. 21187e1194, Washington, DC: IEEE. Bennett, G. L., Sholtis, J. J., & Rashkow, B. C. (1987). United Nations Deliberations on the Use of Nuclear Power Sources in Space:1978e1987. Proceedings of the 5th Symposium on Space Nuclear Power Systems, 8, 45e57, Albuquerque, NM. Botts, C. (2003). Assessment of MFCO Response for MER Mission. Patrick Air Force Base, FL: Technical Report, 45th Space Wing Safety Office. Brenner, J. M. (1997, November 17). In Defense of Cassini. Retrieved November 24, 2010, from Obsidian Rook - Mirthless Doomfiles: www.obsidianrook.com/Mirthless/web/ cassiniova.html. Brunet, J., & Salvetat, B. (1988). n.t. Proceedings of the Joint International Symposium on Compatibility of Plastics and Other Materials with Explosives, Propellants, Pyrotechnics and Processing of Explosives, Propellant and Ingredients. New Orleans, LA: JISCP. Diaz, J. C. (1993). Thermal Environment of a Solid Rocket Propellant in Ambient Atmospheric Conditions. Proceedings of the JANNAF Propulsion Systems Hazards Subcommittee Meeting 389. Diaz, J. C. (1994). Minuteman III e Stage 3 Propellant Fire Characterization. Lawrence Livermore National Laboratory. Dobranich, D., Powers, D. A., & Harper, F. T. (1997). The Fireball Integrated Code Package. Sandia National Laboratories. Duckworth, H. E. (1960). Electricity and Magnetism. New York: Holt, Rinehart and Winston. Elwell, R. B., Irwin, O. R., & Vail, J. R. (1967). Project SOPHY e Solid Propellant Hazards Program. Sacremento, CA: Aerojet General.
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Gayle, J. B., & Bransford, J. W. (1965). Size and Duration of Fireballs from Propellant Explosions. NASA. Gill, W., et al. (1994). A Preliminary Model for Heat Transfer from an Uncontained Solid Rocket Motor Burn. JANNAF Propulsion Systems Hazards Subcommittee Meeting. JANNAF. Gordon, S., McBride, B., & Coveny, L. (1990). Chemical Equilibrium and Transport Properties of Chemical Systems Program. NASA Lewis Research Center. Hasegawa, K., & Sato, K. (1977). Study on the Fireball Following Steam Explosion of n-Pentane. Proceedings of the Second International Symposium on Loss Prevention and Safety Promotion in the Process Industries, 297e304, Heidelberg, Germany. High, R. W. (2006). The Saturn Fireball. Houston, TX: NASA Manned Spacecraft Center. Hill, P. G., & Peterson, C. R. (1970). Mechanics and Thermodynamics of Propulsion (Third printing ed.). In H. W. Emmons, & S. S. Penner (Eds.). Reading, MA, USA: AddisonWesley Publishing Company. Hodgkins, K. (1992, October 28). U.S. Adviser to the 47th Session of the United Nations General Assembly. Special Political Committee on Item #72, International Cooperation in the Peaceful Uses of Outer Space. NY: United States Mission to the United Nations. Honsberg, C. (2010, 12 8). Solar Radiation in Space. Retrieved 2010, from PV Education: www.pveducation.org/pvcdrom/properties-of-sunlight/solar-radiation-in-space. Hopkinson, B. (1915). British Ordnance Board Minutes. IAEA. (1991). International Atomic Energy Agency Statement to the Scientific and Technical SubCommittee. Committee on the Peaceful Uses of Outer Space. NY: United Nations. ICRP (1999). Protection of the Public in Situations of Prolonged Radiation Exposure. ICRP Publication 82. Ann. ICRP 29 (1-2). ICRP (2005). Protecting People against Radiation Exposure in the Event of a Radiological Attack. ICRP Publication 96. Ann. ICRP 35 (1). Lange, R. (1991a). Statement by Robert Lange, U.S. Adviser at the 28th Session of the Scientific and Technical Subcommittee of the UNCOPUOS on Agenda Item 7, Nuclear Power Sources in Outer Space. United Nations Committee on the Peaceful Uses of Outer Space. NY: United Nations. Lange, R. (1991b). Statement by Robert Lange, U.S. Adviser at the 30th Session of the Legal Subcommittee of the UNCOPUOS to the Working Group on Nuclear Power Sources in Outer Space. United States Mission to the United Nations. NY: United Nations. LEAG. (n.d.). Lunar Exploration Analysis Group. Retrieved September 24, 2006, from www. lpi.usra.edu/leag/ Lee, E., & Green, L. (1988). Detonation of Explosives in Accidents. Lawrence Livermore National Laboratory. Mansfield, J. A. (1969). Heat Transfer Hazards of Liquid Rocket Propellant Explosions. URS Systems Corporation. Marietta, M. (1996). Cassini Titan IV RTG Safety Databook. Mazumder, M.e. (2003). Solar Panel Obscuration in the Dusty Atmosphere of Mars. ESAIEEE Joint Meeting 11. MEPAG. (n.d.). Mars Exploration Program Analysis Group. Retrieved September 24, 2006, from http://mepag.jpl.nasa.gov/ Merril, C., Nichols, A., & Lee, E. (1999). Explosvie Response of HTPB Propellants to Impact and Shock. CPIA.
6.9 Symbols and Acronyms
Merrill, C. I. (1992). Large Class 1.3 Rocket Motor Detonation Behavior. Proceedings of the Department of Defense Explosive Safety Board Conference. DDESB. Mukunka, M. (1996). RTG Response to Solid Propellant Fallback Fragment Environments. Orbital Sciences. Nance, W., Crawford, R., et al. (1990). Fuel Impact Explosion Study. Logicon RDA. NASA. (2006). Solar System Exploration Roadmap for NASA’s Science Mission Directorate. SSE Roadmap Tema, Solar System Exploration. Washington, DC: NASA Science Missions Directorate, Planetary Science Division. NEPA. (1970 et seq). Title 42, U.S. Code, Part 4321. National Environmental Protection Agency. NPR 8715.3. (Latest Version). NASA General Safety Program Requirements. Office of Safety and Mission Assurance. Washington, DC: NASA HQ. National Research Council e Radioisotope Power Systems (NRC-RPS) Committee. (2009). Radioisotope Power Systems: An Imperative for Maintaining U.S. Leadership in Space Exploration (ISBN-13: 978-0-309-13857-4; ISBN-10: 0-309-13857-4 ed.). Washington, DC: The National Academies Press. Retrieved February 2013, from. http://www.nap.edu/ openbook.php?record_id=12653. OPAG. (n.d.). Outer Planets Assessment Group. Retrieved September 24, 2006, from www. lpi.usra.edu/opag/ PD/NSC-25. (1977). Scientific or Technological Experiments with Possible Large-Scale Adverse Environmental Effects and Launch of Nuclear Systems into Space. Presidential Directive/National Security Council Memorandum. Ponomarev-Stepnoi, N. N., Talyzin, V. M., & Usov, V. A. (2000, December 13). Russian Space Nuclear Power and Nuclear Thermal Propulsion Systems. Nuclear News, 43(13), 13. Price, E. W., et al. (1979). The Fire Environment of a Solid Propellant Burning in Air. Air Force Weapons Laboratory, AFWL. Reid, R. C., Prausnitz, J. M., & Poling, B. E. (1987). The Properties of Gases and Liquids (4th ed.). McGraw Hill Book Company. Reinhart, L. (1999). Liquid and Solid Propellant Fires. Delta II Databook, Jet Propulsion Laboratory. Robbins, W. H., & Finger, H. B. (1991). An Historical Perspective of the NERVA Nuclear Rocket Engine Technology Program, NASA. Roberts, A. F. (1981e1982). Thermal Radiation Hazards from Releases of LPG from Pressurized Storage. Fire Safety Journal, 4, 197e212. Rumerman, J. A. (1999). NASA Historical Data Book. Washington, DC: NASA History Office. Scolese, C. (2002). NASA’s Nuclear Systems Initiative. Presentation by Nuclear Energy Research Advisory Committee. Shipps, P. R. (1964). Nuclear Pulse Space Vehicle Study. Marshall Space Flight Center. Sholtis, J. A., Marshall, A. C., Bennett, G. L., Brown, N. W., Usov, V. A., & Dawson, S. M. (2008). Safety Implementation, Chapter 12. In A. C. Marshall, & A. C. Marshall (Eds.), Space Nuclear Safety. Malabar, FL, U.S.: Krieger Publishing Company. Shotlis, E. A. (1994). U.S. Space Nuclear Safety: Past, Present, and Furture. New Your, NY: American Intitute of Physics (AIP) Press. Smith, P. G. (1992). Statement by Peter G. Smith, U.S. Representative to the 29th Session of the Scientific and Technical Subcommittee of the UNCOPUOS. NY: UN Committee on the Peaceful Uses of Outer Space. Tate, R. E., & Land, C. C. (1985). Environmental Safety Analysis Tests on the Light Weight Radioisotope Heater Unit (LWRHU). Los Alamos National Laboratory.
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UN Legal Subcommittee. (1992). UN Report of the Legal Subcommittee on the Work of its 31st Session, Chairman’s Text, Principles Relevant to the Use of Nuclear Power Sources in Outer Space. NY: UN. VEXAG. (n.d.). Venus Exploration Analysis Group. Retrieved September 24, 2006, from www.lpi.usra.edu/vexag/ Wertz, J. R., & Larson, W. J. (Eds.). (1991). Space Mission Analysis and Design. Netherlands: Kluwer Academic Publishers. Williams, M. L. (1987). Historical Perspective on Solid Rocket Structural Integrity. Air Force Institute of Technology. WPAFB, OH: AFIT. Willoughby, A. B., Wilton, C., & Mansfield, J. (1968). Liquid Propellant Explosive Hazards. Edwards, CA: AFRPL.
Nuclear-Powered Payload Safety
6
Firooz A. Allahdadi, Sayavur I. Bakhtiyarov, Gregory D. Wyss, Gary F. Polansky, Joseph A. Sholtis, Curt D. Botts
CHAPTER OUTLINE 6.1 6.2 6.3 6.4 6.5 6.6 6.7 6.8 6.9
Introduction to Space Nuclear Systems ........................................................... SNPS Launch History and Accidents ................................................................ Launch Abort Environments Affecting SNPSs .................................................... Containment Design........................................................................................ Risk Assessment for Nuclear Missions ............................................................ International Protocols and U.S. Environmental Review .................................... Nuclear Mission Launch Approval ................................................................... Nuclear Mission Launch Integration ................................................................ Symbols and Acronyms...................................................................................
256 268 275 335 338 348 355 357 361
It would be nice if the risks of doing space exploration could be borne only by the people like me who believe in doing them e just like it would be nice if people driving cars only put their own lives at risk. Unfortunately, it’s hard to see how to structure human life that way. If you put a big sculpture of a foot on top of the Ferry building in San Francisco, there’s a finite chance that it could fall over and injure a passerby e possibly even a passerby that thinks it’s really ugly. Should we forbid all public sculpture? After all, we don’t need it to survive (at least not in any obvious way). How can we justify even a slight chance that someone could die under that gargantuan heel? The value of human life is high, but it is not infinite. If we try and act as though it is infinite, it is in danger of becoming worthless. But there’s nothing to get upset about this time, right? After all the good guys won, didn’t they? Cassini’s in flight, on its way to Saturn. except that what’s going to happen the next time they try and build a space probe using an RTG? Is this another useful technology that’s going to be moved to the list of things that are politically untenable? Heads they win, tails they win. (Brenner, 1997)
This chapter introduces the concepts of Space Nuclear Power Systems (SNPSs). It briefly describes the history and nature of these ingenious machines. For more than five decades, larger and more powerful rockets have inspired awe Safety Design for Space Operations. http://dx.doi.org/10.1016/B978-0-08-096921-3.00006-4 Copyright Ó 2013 Elsevier Ltd. All rights reserved.
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among those privileged to observe their launches. But the very factor that makes the firing of a large rocket so spectacular e the consumption of tons of fuel in just seconds e also suggests practical limits for chemical propulsion. And the limits apply not only to propulsion. Any application that requires high power for more than the few minutes of a chemical rocket’s firing, or power levels greater than can be provided by collecting the Sun’s diffuse radiation energy, is a candidate for nuclear power. The reason for nuclear power in space is simple and compelling: an increase by a factor of as much as one million in the energy that is contained in a mass of nuclear fuel compared with the same mass of chemical fuel. Since the start of the space age, a range of nuclear power supply options has been developed by the United States to support the civilian and military space programs. Tomorrow’s space exploration and exploitation activities will require even larger quantities of reliable, long-lived power (Angelo, 1985).
6.1 Introduction to Space Nuclear Systems Space nuclear systems are vital to the exploration of space and have seen moderate changes in the basic design and fundamental operating schemes. Nuclear-based systems can supply electricity, heat, and propulsion for missions that are well beyond the capacities of solar power, fuel cells, and traditional chemical methods. Nuclear-based systems consist of (1) radioisotope thermoelectric generators (RTGs); (2) fission-based nuclear reactor power systems; (3) nuclear thermal propulsion (NTP); (4) nuclear electric propulsion (NEP), and other advanced nuclear technologies. Since the “Seebeck effect” was discovered in 1826, the basic concept of reliable space nuclear power has grown in efficiency and capacity. German physicist Thomas Johann Seebeck (Duckworth, 1960) published results of experiments conducted in 1822 in which he observed that electrical current is present in a series circuit of two dissimilar metals, provided the junctions of the two metals are at different temperatures (Figure 6.1). This thermoelectric effect increases as the temperature differential increases. The “Seebeck effect” is a result of two phenomena: charge carrier diffusion and phonon drag. Charge carriers (electrons, ions) will diffuse when one end of a conductor is at a different temperature than the other. Hot (or cold) carriers will diffuse from the hot (or cold) end to the cold (or hot) end since there is a lower density of hot (or cold) carriers at the cold (or hot) end of the conductor. The phonon drag phenomenon is an increase in the effective mass of conduction electrons due to contact with the crystal lattice. When an electron interacts with atoms in the lattice its charge alters or polarizes the adjacent lattice. This effect results in a decrease in the electron mobility, and consequently, conductivity decrease. As the magnitude of the thermopower increases with phonon drag, it is advantageous for direct energy conversion applications. Maintaining both connections at the same temperature, and periodically opening and closing one connection, causes a temperature
Thermoelectric effects.
FIGURE 6.1
6.1 Introduction to Space Nuclear Systems
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dependent alternating current (AC) voltage to be generated. This effect will still be observed even when the wires do not touch. Hence, one would conclude that no diffusion is needed for this phenomenon. The Seebeck effect is one of three reversible phenomena illustrating analogous processes associated to thermoelectricity, conductivity, and temperature. French scientist Jean-Charles Peltier in 1834 described the second closely related phenomena (known as the “Peltier effect”): the temperature at either junction changed proportionally to the voltage between the metal conductors. German scientist Heinrich Lenz in 1839 expanded on Seebeck’s and Peltier’s discoveries and showed that heat transfer at the junctions depends on the direction of the current in the circuit (“SeebeckePeltier effect”). In 1851, British physicist William Thomson (Kelvin) described the rate of heat created (or absorbed) in a current-carrying conductive material subjected to a temperature gradient (“Thomson effect”). A simple and most popular electrical tool based on measuring the SeebeckePeltiereThompson (SPT) effects is the thermocouple, which converts the thermal potential difference into electric potential difference. Several thermocouples connected in series are called a thermopile, which can be used to increase the output voltage. SPT effects are also the basic work principle of thermoelectric generators such as RTGs used to generate power from heat differentials. The United States has designed, tested and launched RTGs for more than three decades. The RTG systems have been used with great success in several extraterrestrial missions and space exploration experiments. Those missions would not have been possible without these important power sources. However, these RTG systems are limited in producing power up to tenths of kilowatts, making them less ideal for use in maintaining equipment operations and sustaining human life during long duration space missions beyond the Moon and Mars. In 1958 a U.S. project, Orion (Figure 6.2), was initiated by General Atomics (Shipps, 1964). The purpose of the Orion was to launch a spacecraft (~1000 tons) INTERMEDIATE PLATFORM 1st STAGE SHOCK ABS.
2nd STAGE SHOCK ABS.
UPPER MODULE SECTION (BODY)
PUSHER
PAYLOAD SECTION PROPELLANT MAGAZINES PROPULSION MODULE
FIGURE 6.2 Orion nuclear spacecraft.
6.1 Introduction to Space Nuclear Systems
FIGURE 6.3 Nuclear Engine for Rocket Vehicle Applications (NERVA).
utilizing a series of small nuclear explosions against a “pusher plate” to provide thrust. The project was terminated in 1963 when the Atmospheric Test Ban Treaty made it illegitimate, as the radioactive fallout could have been a main problem. The concept, nonetheless, is still living as one of the more efficient methods for producing propulsion energy. From 1959 to 1973 a U.S. multi-mission nuclear rocket program, “Nuclear Engine for Rocket Vehicle Applications” (NERVA), was conducted to substitute a nuclear power for chemical rockets at certain launch phases. Electric propulsion provides remarkable advantages over chemical propulsion. It enables new classes of Solar System exploration missions with multiple targets, eliminates or reduces launch windows required for gravity assists, reduces cruise time to distant targets, and reduces mission cost because smaller launch vehicles may be used (Scolese, 2002). The proposed NERVA system (Figure 6.3) used graphite-core reactors heating hydrogen and ejecting it through a nozzle delivering a minimum of 75,000 lbf thrust. A total of 20 reactor tests and two engine tests were conducted (Robbins & Finger, 1991) in Nevada and delivered thrust more than half that of the space shuttle main engines. Since the cancellation of the program, “nuclear rockets” have been proposed for space propulsion, but not for the initial launch phase. According to the U.S. National Aeronautics and Space Administration’s (NASA’s) 2006 Solar System Exploration Roadmap (NASA, 2006), the lunar and Mars explorations eventually will be manned missions in around 2030 to answer the following fundamental questions: • • •
How did the Sun’s family of planets and minor bodies originate? How did the Solar System evolve to its current diverse state? What are the characteristics of the Solar System that led to the origin of life?
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• •
How did life begin and evolve on Earth and has it evolved elsewhere in the Solar System? What are the hazards and resources in the Solar System environment that will affect the extension of human presence in space?
A scientific input to NASA in developing and achieving these goals and priorities is offered by the NASA Advisory Council (NAC) and by other science advisory groups, such as the Outer Planets Assessment Group (OPAG), the Venus Exploration Analysis Group (VEXAG), the Mars Exploration Program Analysis Group (MEPAG), and the Lunar Exploration Analysis Group (LEAG). In the Solar System Exploration (SSE) Roadmap (NASA, 2006) the proposed missions could be facilitated by Radioisotope Power Systems (RPSs) identified as one of the highest priority technologies. The following goals are established by the MEPAG for Mars exploration: • • • •
Determining if life ever arose on Mars. Understanding the process and history of climate on Mars. Determining the evolution of the surface and interior of Mars. Preparing for human exploration.
In November of 2011 NASA’s Mars Science Laboratory (MSL) launched “Curiosity,” a nuclear powered rover to Mars using an Atlas V 541 Launch Vehicle. The MSL mission will explore Mars habitability (i.e., capability of the Mars environment to sustain a life). Presently, NASA has developed two new types of RPSs. These are the MultiMission Radioisotope Thermoelectric Generator (MMRTG), with static power conversion, and the Stirling Radioisotope Generator (SRG), utilizing dynamic conversion. The MMRTG includes a significant amount (~4.8 kg) of ceramic plutonium dioxide (PuO2) as a heat source to supply continuous power at approximately 110 watts on the Martian surface. In the MMRTG design the leadetelluride (PbeTe) thermoelectric sources are protected from exposure to the Martian atmosphere, which produces helium due to radioactive decay of the PuO2 as is vented. In the U.S., space missions that involve radioactive materials require presidential launch approval per Presidential Directive/National Security Council Directive 25 (PD/NSC-25, 1977). The Interagency Nuclear Safety Review Panel (INSRP) was formed to conduct an independent safety review and evaluation of the nuclear risk associated with the MSL mission. The INSRP process was established in the 1960s to review safety and assurance issues related to the nuclear space launches sponsored by NASA, the Department of Defense (DOD) and the Department of Energy (DOE) (former Atomic Energy Commission e AEC). The U.S. Environmental Protection Agency (EPA) and U.S. Nuclear Regulatory Commission (NRC) also participate in the INSRP review process. The Advanced Stirling Radioisotope Generator (ASRG) is a next-generation RPSs system. A Stirling engine is used to convert thermal energy to electricity with higher efficiency (~29%) than thermocouples in RTGs. The ASRG utilizes
6.1 Introduction to Space Nuclear Systems
moving parts, produces more electricity than MMRTGs, and requires only onefourth of the plutonium-238 (238Pu) required for RTG operation. The introduction of the SRG design shifted the focus of space agencies to development of a flightready and reliable power system since it can minimize a demand for 238Pu for long-duration space missions (NRC-RPS Committee, 2009). However, the long term reliability of the Stirling design, especially since moving parts are required is, as yet, unproven. Many hours of bench testing have been accomplished and early space missions should reveal the trustworthiness of this more efficient power generator. In the Stirling engine, the conversion is based on oscillation of helium gas at one end of the generator, which is heated by radioactive decay of 238Pu, while the other end is cooled by a heat sink. This oscillating helium gas pushes a piston in a linear alternator that generates AC electricity. Then the AC is converted to direct current (DC) electronically, and the current is fed to the power management system of the spacecraft (NRC-RPS Committee, 2009). During some missions the orbiters and landers will be constantly exposed to the high radiation, which must be diminished. RPSs with static conversion utilize hundreds of thermocouples, which provide a built-in redundancy. That is, failure of a few thermocouples would have a minimal impact on the overall power system performance. In contrast to the RPS, the SRGs are more sensitive to radiation. Therefore, the control systems require a considerable shielding. Electromagnetic radiation (EMR) from the SRG could interfere with scientific instrumentations, while EMR shielding to reduce it would impact system mass. The power systems such as solar arrays are not feasible for the demanding harsh extraterrestrial environment and due to the lack of sunlight and meaningful storage facilities such as batteries. For example, sunlight at Jupiter and Pluto is 96% and 99.94% less intense, respectively, than at Earth. The Juno mission to Jupiter will be powered by solar arrays, but it will be in a highly elliptical polar orbit. It will not orbit near the Jovian equatorial plane where the most intense portions of the belts are located. Thus, it will spend little time in the belts themselves (NRC-RPS Committee, 2009). Currently, RPSs are considered the only existing reliable power source that can work unrestrained in the extraterrestrial environment for the extended periods of time needed in order to accomplish many missions (NRCRPS Committee, 2009). RPSs are compact, rugged and very reliable spacecraft power systems that provide a dependable and long-lived power in space environment. RPSs are not nuclear reactors and they do not use nuclear fission or fusion processes to generate energy. RPSs produce heat through the natural radioactive decay of 238Pu. The generated heat flows through thermocouples to a heat sink generating DC electricity via the conversion process. All of the RPSs flown to date have been RTGs. Since 1961 RPSs have supported 26 NASA and DOD missions and have an outstanding safety and reliability record (see Table 6.3, p. 15). The design of an RPS incorporates many safety features to support continuous operation in the severe space environment. However, the energy conversion efficiency of an RPS is quite low (~6%). The RPS technologies require environmental documentation and safety analysis before approval for use and launch, and
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special accommodations that can impact mission architecture and mission design (Balint, 2007). They must be designed for all mission phases, including launch, cruise, entry, descent, landing and in-situ operations. It is noteworthy to mention that the conditions between these mission phases are different. For example, the heat transfer mechanisms to reject the excess heat created by the radioisotope decay of the plutonium fuel are different for the operations in planetary environments, where heat is rejected through convection, conduction and radiation. During the cruise phase RPSs can be encapsulated inside an aeroshell, while the excess heat is removed by a fluid loop and rejected to space through external radiators. However, for orbiting missions, the RPSs can be exposed directly to the space environment, and heat rejected through radiation directly to space, but the system must still address the pre-launch and launch environments on Earth (Balint, 2007). Plutonium-238 with an 88 year half-life is considered the only practical and convenient isotope for energizing those RPSs for space exploration. The crucial role of RPSs and the need for a supply of 238Pu for U.S. space exploration programs have been proved during space missions such as Viking, Voyager, Cassini and Pluto New Horizons. Unfortunately, 238Pu does not exist in nature. It is created by irradiating Neptunium-237 (237Np) targets in a nuclear reactor; however, 238Pu is unsuitable for use in nuclear detonation weapons. For space power applications, 238Pu is preferable over other isotopes (NRC-RPS Committee, 2009) since it: • • • • •
generates heat for adequate time (radioactive decay half-life of sufficient length); produces a type and quantity of emissions by radioactive decay of fuel that allow it to be handled safely; provides high specific power (heat/mass) and high power density (heat/volume); is a non-corrosive, water-insoluble, and chemically stable fuel form with good engineering properties at high temperatures; can be produced in sufficient quantity at an affordable cost.
Because 238Pu emits alpha particles, RPSs can cause only a biological vulnerability if the 238Pu is somehow released into the environment (then ingested through the food or inhaled through the atmosphere). RPSs are fueled with 238Pu in the form of a ceramic oxide (238PuO2) that has a high melting point and very low solubility to minimize fuel vaporization and transport in the atmosphere, and minimize fuel retention within the human body (NRC-RPS Committee, 2009). The RPSs are designed with multiple fuel containment systems to prevent a fuel release and dispersal of 238Pu into the biosphere during a space mission. (NRC-RPS Committee, 2009). Protective systems are described in the section on “Nuclear Mission Launch Approval”. Only two countries in the world e the U.S. and Russia e have produced 238Pu in the past. Hence, a short supply of the 238Pu may have a significant impact on several pending missions. There are three mission classes for the SSE Program: Discovery e Scout for Mars (small), New Frontiers (medium), and Flagship (large). The projected net balance (supply minus demand) and need for 238Pu to conduct future space operations is shown in Figure 6.4 and Table 6.1 (NRC-RPS Committee, 2009). Due
6.1 Introduction to Space Nuclear Systems
120
Pu Balance = Supply - Demand
100 Pu demand (status quo)
Kilograms of Pu-238
80 60
Pu demand (best case)
40
Pu balance (best case)
20 0 –20
Pu balance (status quo)
The Problem
–40 2008
2010
2012
2014
2016 2018 2020 Calendar year
2022
2024
2026
2028
FIGURE 6.4 Projected
238
Pu demand and net balance (NRC-RPS Committee, 2009).
Table 6.1 NASA’s demand for 238Pu in RPS-involved space missions 238
Pu (kg) 3.5 1.8 24.6 3.5 5.3 14 14 1.8–5.3 3.5 14 5.3–6.2 14 105–110 5.3–5.5
Mission
Launch year
Watts
Mars Science 2011 100 Laboratory Discovery 2014 250 12/Scout Outer Planets 2017 600–850 Flagship 1 Discovery 14 2020 500 New Frontiers 4 2021 800 Pressurized 2022 2000 Rover 1 ATHLETE Rover 2024 2000 New Frontiers 5 2026 250–800 Discovery 16 2026 500 Pressurized 2026 2000 Rover 2 Outer Planets 2027 700-850 Flagship 2 Pressurized 2028 2000 Rover 3 Total demand for Pu (kg) 2009–2028 Annual demand (20-year average in kg/year).
Type of RPS MMRTG ASRG MMRTG ASRG ASRG High Performance SRG High Performance SRG ASRG ASRG High Performance SRG ASRG High Performance SRG
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to low cost caps (~$425M FY06) the Discovery and the Scout class missions are not budgeted to utilize RPSs. Even with the protective containment systems the RPSs may create some risk that 238 Pu could be released into the biosphere. To evaluate this risk, the U.S. has established a flight safety review and launch approval process for RPS-powered missions established by Presidential Directive/National Security Council memorandum-25 (PD/NSC-25, 1977). During this process the DOE prepares a detailed Safety Analysis Report (SAR) to evaluate radiological risk for each mission. The Final Safety Analysis Report (FSAR) is reviewed by the INSRP. The INSRP coordinators conduct an independent safety review and evaluation of the nuclear risk for each RPS involved mission. The INSRP provides technical oversight in six distinct technical Working Groups (WGs): Launch Abort, Re-entry, Power Systems, Meteorology, Biomedical and Environmental Effects, and Risk Integration and Uncertainty. The INSRP then prepares a mission specific Safety Evaluation Report (SER), which identifies, characterizes, and quantifies probable accident scenarios, including the probabilities that 238Pu will be released. NASA uses the FSAR and the SER to request launch approval from the Office of Science and Technology Policy (OSTP), which is within the Executive Office of the President. After reviewing the SER and evaluating the potential risk, OSTP then recommends to the President whether to authorize or deny the requested launch. The President may choose to defer the decision to the Vice President or his Chief Scientist at OSTP. The launch approval process on average takes 3 years (NRC-RPS Committee, 2009). In near future space missions may require building human stations on the Moon and Mars for long-term human attendance and equipment operations. The required electrical power for human stations might be tens of kilowatts for several years. Nowadays nuclear fission power is considered as another promising alternative for outer planet exploration where solar power is not realistic. Nuclear fission power reactors can effectively and safely supply the power required for long-period human and robotic missions to explore the outer planets and targets beyond the Solar System. Nuclear fission reactors can be used to power operating systems of a wide range of magnitudes. Nuclear fission reactors can generate safe, efficient and consistent electric power during the space missions and operations on other planets. According to NASA reports, reactor systems can provide tens of kilowatts of electrical power for electric propulsion and/or operation of science instruments (Scolese, 2002). A nuclear fission reactor can provide significantly more power than ~0.1 kW electricity provided by RTGs. However, in order to reduce mission cost and the amount of fuel, it is enviable to build fission reactor power systems with longer operating lives (more than 20 years) to supply ~40 kWe power for future stations on the Moon and Mars (Schriener & El-Genk, 2011). Several reactor power systems have been designed and tested to provide electric power up to 40 kW for 5e10 years. The Affordable Fission Surface Power System (AFSPS) concept was developed by NASA and DOE for providing electrical power to space stations on Moon and Mars surfaces Mason et al. (2008). This proposed system utilizes a compact fission reactor that operates at
6.1 Introduction to Space Nuclear Systems
temperatures below 1000 K. It is coupled to six pairs of Free Piston Stirling Engines (FPSEs) to generate 40 kW of electric power for 8 full power years (FPY). The FPSE transfers the thermal power to produce AC electricity with a predicted thermal efficiency of 22.85%. The highly enriched (93 wt% 235U) UO2 pellets in the core’s fuel rods are doped with Gd2O3 thermal neutron poison and a clad with 316 stainless steel. The Gd2O3 guarantees that the reactor is safe should it be exposed to moisture. The first, and thus far only, U.S. nuclear fission space power system, SNAP10A, was launched in the 1960s. Project Prometheus, which was NASA’s most recent attempt to develop space nuclear power reactors, selected a nuclear electric propulsion reactor concept that was scalable from 20 kWe to 300 kWe (NRC-RPS Committee, 2009). However, the performance and reliability of space nuclear power reactor systems using current technology remains unproven, especially for missions with long lifetimes (NRC, 2009). Both radioisotope power (via RTG) and fission-based reactor systems are highly reliable technologies with no moving parts. Hence, they can make it possible to provide safer and more dependable thrust generation for space missions than conventional chemical power and propulsion systems. The Sectored Compact Reactor (SCoRe-S) power system concept was developed to provide 100 kW electric power for up to 6 years El-Genk et al. (2005); Hatton & El-Genk, (2006, 2009). The SCoRe-S fission heat source that operates at relatively high temperatures (1150e1250 K) is cooled by circulating liquid lithium El-Genk et al. (2005). The electrical power to the load and to the electromagnetic (EM) pumps for circulating the liquid lithium is provided by the power system which uses SiGe thermoelectric energy conversion elements. The SCoRe-S core (Figure 6.5) is divided into six equal triangular sectors separated by flat, liquid
Rx HPs graphite support disk(0.2 cm)
66.1 cm
Top section of Rx vessel Enclosing gas plenum of UN pins
Control drums actuators
Openspaces for shield cooling
FIGURE 6.5 Sectored Compact Reactor (SCoRe-S).
Rx control actuators Rx lithium heat pipes in multifoils insulation
24.9
Tungsten γ-shield
92.24
15°
59.98
BeO radial
B4C reflector clad in sectors MA-ODS steel
Natural LiH neutrons shield in SS honeycomb
All dimensions in cm
Graphite conduction disk (2.0)
Depleted LiH in SS honeycomb
W-LiH shadow shield clad InMA-ODS steel (0.05)
56 cm
Reactor Li heat pipes In MFT insulation (25 10 μm-thick foils, 100 μm gap)
Steel encasement (0.5 mm) 74.48
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metal heat pipe dividers, are thermally and neutronically coupled, but hydraulically independent El-Genk et al. (2005); El-Genk, (2008). Each sector is thermal-hydraulically coupled to a separate pair of circulating liquid lithium primary and secondary loops with thermoelectric (TE) energy conversion modules and heat rejection radiator panels. The fastest method to drive spacecraft to the target with a lesser amount of propellant is to utilize Nuclear Thermal Propulsion (NTP) systems. A broad development of NTP systems was performed by the U.S. and the Russian Federation in 2007. They evaluated NTP space programs for advanced unmanned missions to the outer planets. Within NASA’s new Vision for Space Exploration Program, Marshall Space Flight Center (MSFC) and Los Alamos National Laboratory (LANL) initiated a development of a new NTP engines. The design process interfaces with the ROCket Engine Transient Simulation (ROCETS) system tool which has been used extensively in the liquid rocket engine industry for many years. ROCETS produces an integrated engine and reactor system design for the specific mission and material limits, at the same time maximizing system thrust to weight ratio Amiri et al. (2007). Today, nuclear fission power is considered as the only alternative for extra-solar exploration. Nuclear fission power reactors can effectively and safely supply the needed energy for long-period human and robotic missions to explore the outer planets and targets beyond the Solar System. Besides propulsion, nuclear fission reactors can be used to power operating systems of a wide range of magnitudes. They can generate safe, efficient and consistent electric power during space missions and operations on other planets. According to NASA, these types of systems can provide tens of kilowatts of electrical power for electric propulsion and operation of science instruments (Scolese, 2002). A nuclear fission power system can provide significantly more power than the ~0.1 kW electrical that RTG systems are currently limited to. Only a single U.S. nuclear fission power system has been launched occurring on April 3, 1965. This System for Nuclear Auxiliary Power (SNAP)-10A functioned for 43 days before it failed, but demonstrated successful operation of the contingency safing plan, placing it into a 4000 year orbit.
Basic Concept e Have Heat, Will Travel If you need constant and reliable power for a long period of time there are several ways to provide that terrestrially; but for long space voyages the criticality of the space vehicle’s power source is a primary concern. For spacecraft missions in the vicinity of the inner planets (Mercury, Venus, Earth, Mars) the standard for power supply has traditionally been based on solar panels, especially if the mission does not call for landing on the planet surface. As seen in the Mars Exploration Rover (MER) missions utilizing the Spirit and Opportunity rovers, solar power is subject to degradation as panels become covered by atmospheric dust (Mazumder, 2003). As related in Figure 6.6, the distance a planet is located in relation to the Sun
6.1 Introduction to Space Nuclear Systems
affects the available solar irradiance (Honsberg, 2010). The total solar radiation, emitted by the Sun is given by Psun ¼ sT 4 Asun ¼ sT 4 4pR2sun ¼ 4:5 1026 Watt
(1)
where: Psun is the power emitted at the Sun’s surface (in Watts) as determined by StefaneBoltzmann’s blackbody equation; s is the StefaneBoltzmann constant ¼ 5.670400 10e8 Jse1me2Ke4 T is 6000 degrees Kelvin Rsun is the radius of the Sun in meters (6.9599 108 m) and the solar radiation intensity at a given distance, D, in meters is given by: H0 ¼
R2sun Hsun D2
(2)
The mean solar irradiance in watts per square meter is provided in Table 6.2 for the planetary bodies in our Solar System.
FIGURE 6.6 Solar System scaled to football field.
Table 6.2 Solar irradiance at planetary distances Planetary body
Distance (3109 m)
Mean solar irradiance (W/m2)
Mercury Venus Earth Mars Jupiter Saturn Uranus Neptune Pluto
58 108 150 228 778 1427 2871 4497 5906
9145 2619 1370 590 51 15 4 2 1
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Noticeable is the drop in mean solar irradiance once the orbit of Jupiter is reached. This significant reduction in available power density beckons the need to provide power for spacecraft applications through other means. Using Seebeck’s discovery, engineers developed SNPSs since extremely reliable heating sources are necessary to guarantee long-duration mission power.
Small to Large Systems Depending upon the needs of the space vehicle mission, the heat source can range from small pellets providing environmental heating only, to larger pellets providing primary heating to thermoelectric conversion units, to large nuclear reactor components providing electrical power conversion more akin to Earthbound systems.
Early to Today Development Historical development of SNPS A good source is Angelo (1985), covering: Polonium Plutonium Uranium LWRHUs, small RTGs, SRGs, MMRTGs, space nuclear reactors, etc.
6.2 SNPS Launch History and Accidents The history of launching spacecraft that have employed radioisotope systems for heat or to generate power began nearly as soon as mankind began to launch rockets. Only two countries however have launched these types of systems.
U.S. Space Nuclear Systems Launch History and Accidents Since 1961, the U.S. has launched 46 nuclear power systems and hundreds of radioisotope heater units (RHUs) on 30 space missions intended for navigation, meteorology, communications, and exploration of the Sun, Moon, Mars, Jupiter, Saturn, and the remainder of our outer Solar System and beyond (Table 6.3). One of these space missions, SNAPSHOT in 1965, employed a 500 We nuclear reactor power system, the remainder employed RPSs, specifically RTGs, and/or RHUs. The objectives of these missions could not have been accomplished without onboard nuclear power systems (Shotlis, 1994), (NRC-RPS Committee, 2009). Currently, there are eight nuclear-powered space missions/spacecraft that are still active; they are: Lincoln Experimental Satellites 8 and 9 (LES 8 and 9), Voyager 1 and 2, Cassini, Mars Exploration Rovers A and B (Spirit and
Table 6.3 U.S. Mission Spacecraft using nuclear power systems, launched 1961–2006 Nuclear system (Num.-nom. pwr)
Mission type
Launch
TRANSIT 4A
Navigational
Jun 1961
TRANSIT 4B
Navigational
Nov 1961
TRANSIT 5BN-1
Navigational
Sep 1963
TRANSIT 5BN-2
Navigational
Dec 1963
TRANSIT 5BN-3
Navigational
Apr 1964
SNAPSHOT
Experimental
Apr 1965
NIMBUS B-1
Meteorological
May 1968
SNAP-19B2 (2-40 We RTGs)
NIMBUS III
Meteorological
Apr 1969
APOLLO 12
Lunar Exploration Lunar Exploration
Nov 1969
SNAP-19B3 (2-40 We RTGs) SNAP-27 (1-70 We RTG) SNAP-27 (1-70 We RTG)
APOLLO 13
Apr 1970
SNAP-3B7 (1-2.7 We RTG) SNAP-3B8 (1-2.7 We RTG) SNAP-9A (1-25 We RTG) SNAP-9A (1-25 We RTG) SNAP-9A (1-25 We RTG) SNAP-10A (1-500 We Reactor)
Status Successfully achieved orbit; ops terminated 1966 Successfully achieved orbit; ops terminated 1967 Successfully achieved orbit; ops terminated 1970 Successfully achieved orbit; ops terminated 1971 Failed to achieve orbit; SNAP-9A RTG burned up on re-entry as then designed/intended Successfully achieved orbit; voltage regulator on S/C bus failed after 43 days of ops; SNAP-10A reactor shutdown automatically & permanently in 3000þ year orbit Launch vehicle intentionally destroyed during early launch phase; SNAP-19B2 RTGs retrieved intact from Santa Barbara Channel; fuel used on subsequent mission Successfully achieved orbit; ops terminated 1979 Successfully placed on Moon; ops terminated 1980 Mission aborted en route to Moon; SNAP-27 survived re-entry, as designed; under 7000þft of water in S. Pacific Tonga Trench (Continued)
6.2 SNPS Launch History and Accidents
Mission
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270
Mission
Mission type
Launch
APOLLO 14
Lunar Exploration Lunar Exploration Solar System Exploration Lunar Exploration Navigational
Jan 1971
APOLLO 15 PIONEER 10 APOLLO 16 TRANSIT/ TRIAD-01-1X APOLLO 17
Jul 1971 Mar 1972 Mar 1972 Sep 1972
LES 8
Lunar Exploration Solar System Exploration Mars Exploration Mars Exploration Communications
Mar 1976
LES 9
Communications
Mar 1976
VOYAGER 2
Solar System Exploration Solar System Exploration
Aug 1977
PIONEER 11 VIKING 1 VIKING 2
VOYAGER 1
Dec 1972 Apr 1973 Aug 1975 Sep 1975
Sep 1977
Nuclear system (Num.-nom. pwr) SNAP-27 (1-70 We RTG) SNAP-27 (1-70 We RTG) SNAP-19 (4-40 We RTGs) SNAP-27 (1-70 We RTG) TRANSIT-RTG (1-30 We RTG) SNAP-27 (1-70 We RTG) SNAP-19 (4-40 We RTGs) SNAP-19 (2-40 We RTGs) SNAP-19 (2-40 We RTGs) MHW-RTG (2-150 We RTGs) MHW-RTG (2-150 We RTGS) MHW-RTG (3-150 We RTGs) MHW-RTG (3-150 We RTGs)
Status Successfully placed on Moon; ops terminated 1980 Successfully placed on Moon; ops terminated 1980 Successfully placed on interplanetary trajectory; ops terminated 2008 Successfully placed on Moon; ops terminated 1980 Successfully achieved orbit; ops terminated 1977 Successfully placed on Moon; ops terminated 1980 Successfully placed on interplanetary trajectory; ops terminated 2008 Successfully placed on Mars surface; ops terminated 1980 Successfully placed on Mars surface; ops terminated 1980 Successfully achieved orbit; still operational Successfully achieved orbit; still operational Successfully placed on interplanetary trajectory; still operational near heliosheath Successfully placed on interplanetary trajectory; still operational near heliosheath
CHAPTER 6 Nuclear-Powered Payload Safety
Table 6.3 U.S. Mission Spacecraft using nuclear power systems, launched 1961–2006dContinued
Jovian Exploration
Oct 1989
ULYSSES
Solar Polar Exploration Mars Rover Exploration Saturnian Exploration
Oct 1990
Mars Rover Exploration Mars Rover Exploration Pluto & Kuiper Belt Exploration
Jun 2003
MARS PATHFINDER CASSINI/ HUYGENS
MER-A MER-B PLUTO NEW HORIZONS
Dec 1996 Oct 1997
Jul 2003 Jan 2006
GPHS-RTG (2-275 We RTGs) & LWRHU (120-1 Wt RHUs) GPHS-RTG (1-275 We RTG) LWRHU (3-1 Wt RHUs) GPHS-RTG (3-275 We RTGs) & LWRHU (117-1 Wt RHUs) LWRHU (8-1 Wt RHUs) LWRHU (8-1 Wt RHUs) GPHS-RTG (1-275 We RTG
Successfully placed in orbit around Jupiter; intentionally de-orbited into atmosphere of Jupiter in 2004 to ensure future protection of Europa
Successfully placed on solar polar orbit; ops terminated 2010 Successfully placed on Mars surface; ops ceased 1997 Successfully placed in orbit around Saturn; still operational
Successfully placed on Mars surface; still operational Successfully placed on Mars surface; still operational En-route to Pluto w/flyby in 2015; still operational
6.2 SNPS Launch History and Accidents
GALILEO
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Opportunity), and Pluto New Horizons. In the immediate future, the MSL mission/spacecraft with a rover (Curiosity) powered completely by the MultiMission RTG (MMRTG) is scheduled for launch in late 2011. Nuclear power systems are expected to continue to enable and support further robotic exploration of our Solar System and beyond in the future. Moreover, nuclear systems for power and propulsion will be required to enable future human missions to the Moon, Mars, and asteroids as we expand our presence in space. Table 6.3 provides a listing of all known U.S. launches of payloads that utilized some form of radioisotope. Launch, orbital and final trajectory insertion, and space flight involve risks of failure. Some failures can present severe accident environments to onboard systems, including nuclear systems. In general, the most critical mission phases are launch and ascent, when large quantities of propellants are present. Accidents during launch and ascent can threaten an onboard nuclear system from blast overpressure, small shrapnel and large fragment impacts, Earth surface impact, propellant fires, and falling debris. In addition to accidents during launch and ascent, malfunctions during orbital or final trajectory insertion can result in an inadvertent re-entry, with its associated re-entry heating, ablation, and thermo-mechanical shock environments, followed by Earth surface impact. For nuclear systems that require active cooling (e.g., space reactors once operated in space), internal malfunctions during operation in space that can result in overpower or under-cooling conditions can also threaten the nuclear system’s integrity. Overall, the probability of catastrophic launch vehicle and upper stage accidents is in the range of a few percent. Consequently, space nuclear systems must be designed to respond to such accidents without presenting undue risk to the Earth’s population or environment (Shotlis, 1994). Four U.S. space missions with nuclear systems onboard have experienced failures; three were caused by the launch vehicle or upper stage. In each case, built-in safety features performed as designed and there were no adverse radiological consequences. A chronological discussion of these events follows (Shotlis, 1994). The first failure occurred on April 21, 1964, when the TRANSIT 5BN-3 navigational satellite failed to achieve Earth orbit because a computer malfunction prematurely shut down an upper stage booster. The satellite and its SNAP-9A RTG power supply re-entered the Earth’s atmosphere and burned up completely e as early RTGs were designed to do e at an altitude of ~50 km. Approximately 20,000 Curies of 238Pu were released into the upper atmosphere and dispersed worldwide. Although this accident did not pose a health threat to any member of the population, it did involve an uncontrolled release of radioactivity into the biosphere. Consequently, the design requirement for RPSs and RHUs under accidental re-entry was changed from complete breakup and dispersal at high altitude to survival intact, that is, with fuel containment and confinement preserved through re-entry. In May 1965, 43 days after the successful launch and deployment of the SNAPSHOT experimental spacecraft, powered by a SNAP-10A reactor, a voltage
6.2 SNPS Launch History and Accidents
regulator failure on the spacecraft bus caused automatic and permanent (i.e., irreversible) shutdown of the reactor. No deleterious consequences were involved as a result of this failure. The reactor currently remains in a 3000þ year Earth orbit. When it does re-enter the Earth’s atmosphere in the distant future, the reactor will contain the same insignificant amount of radioactivity as it had when it was launched. That is, the fuel will again be essentially radioactively cold and clean. On May 18, 1968, approximately one minute into the launch of the SNAP19B2 powered NIMBUS B-1 meteorological satellite, the Range Safety Officer at Vandenberg Air Force Base, CA, destroyed the launch vehicle by command destruct action, so its errant launch trajectory would not place the public in danger. Although the launch vehicle, upper stage, and spacecraft were completely destroyed, the two on-board SNAP-19B2 RTGs survived intact, with no release of their radioactive 238Pu fuel. The RTGs were retrieved from the Santa Barbara channel, and their nuclear fuel was used on a subsequent space mission. Lastly, in April 1970, the APOLLO 13 mission was aborted en-route to the Moon because of an explosion of an oxygen tank in the APOLLO 13 Service Module. A SNAP-27 RTG was on the APOLLO 13 Lunar Lander to power a lunar surface experiment package. Because the Lunar Lander returned to Earth with the Crew Re-entry Module, it and the SNAP-27 RTG experienced re-entry into and through the Earth’s atmosphere. The SNAP-27 RTG survived re-entry intact, with no release of its radioactive 238Pu fuel, and sank to a depth in excess of 7000 feet at the bottom of the Tonga trench in the South Pacific, where it remains. Although failures/accidents have occurred, they were anticipated and specifically accounted for in the design of the onboard space nuclear systems, to prevent harmful radiological consequences. The historical record, documented in Table 6.3, including the failures/accidents discussed above, serves to provide confidence that the space nuclear safety program in the U.S. has been and continues to be effective as well as vital.
Soviet/Russian Space Nuclear Systems Launch History and Accidents The Soviet Union/Russia has launched 40 space nuclear systems in support of 40 space missions during the period September 3, 1965, and November 16, 1996. Five of those systems/missions involved RTGs, two involved RHUs, while the remaining 33 involved space reactors (Bennet, 1989; Ponomarev-Stepnoi et al., 2000), and (Sholtis et al., 2008). The first two Soviet nuclear-powered space missions, Cosmos-84 and Cosmos90, launched on September 3, 1965 and September 18, 1965, respectively, each used a 20 We RTG fueled with polonium-210. Two other RTG-powered LUNOKHOD spacecraft were launched on Cosmos-300 and Cosmos-305, on
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September 23, 1969 and October 22, 1969, respectively. Both of these LUNOKHOD spacecraft failed to achieve Earth orbit and the onboard RTGs, each fueled with polonium-210, re-entered the Earth’s atmosphere and burned up, as designed, on September 27, 1969 and October 24, 1969, respectively. Lastly, among the Soviet/Russian RTG-powered space missions, was the MARS 96 mission, launched on November 16, 1996, with a 0.2 We plutonium-238 fueled RTG. The MARS 96 spacecraft failed to achieve Earth escape velocity, and reentered the Earth’s atmosphere on November 17, 1996, falling into the South Pacific off the coast of Chile and Bolivia. This RTG was designed to survive reentry, and no radioactivity was detected from the re-entry or impact of the MARS 96 onboard RTG. Two Soviet space missions, Luna 17 and Luna 21, launched on November 10, 1970, and January 8, 1973, respectively, each used a 1.0 kWt RHU, fueled with polonium-210. These RHUs apparently served to provide heat for the batteries and electronics of the Soviet MOONCAR-1 and MOONCAR-2 lunar vehicles (Sholtis, Marshall, Bennett, Brown, Usov and Dawson, 2008). The remaining 33 Soviet space nuclear systems/missions used space reactors. Thirty-one of these systems were 3.0 kWe reactor systems known as Buk (bul, meaning “beech” in Russian) reactors. These Buk space reactors, which had a uranium-235 fuel loading of 30 kg, produced 100 kWt, and used thermoelectrics to convert this heat into 3.0 kWe. The total Buk space reactor power system mass was 930 kg. The Soviet Buk space reactor powered their Radar Ocean Reconnaissance Satellites (RORSATs), launched during the period October 3, 1970, and March 14, 1988 (Sholtis, Marshall, Bennett, Brown, Usov & Dawson, 2008). Two other Soviet reactor-powered space missions, Cosmos-1818 and Cosmos1867, launched on February 1, 1987, and July 10, 1987, respectively, each used a different space reactor known as TOPAZ. The Soviet TOPAZ space reactor incorporates in-core thermionics to convert the 150 kWt generated by the reactor directly into 5.0 kWe. The uranium-235 fuel loading of the TOPAZ reactor is 11.5 kg, and the total mass of the TOPAZ space reactor power system is 980 kg (Sholtis, Marshall, Bennett, Brown, Usov & Dawson, 2008). Three RORSAT missions (Cosmos-954, Cosmos-1402, and Cosmos-1900), each with a Buk space reactor onboard, experienced accidents. The first accident involved Cosmos-954, which was launched on September 18, 1977. The Cosmos-954 RORSAT operated in low Earth orbit for 43 days. At the end of its operational mission, the Buk reactor successfully separated from the spacecraft, but failed to boost to its intended higher, long-lived orbit. Consequently, the reactor re-entered the Earth’s atmosphere and impacted the Earth’s surface near Great Slave Lake in the Northwest Territories of Canada on January 24, 1978. The Buk reactor was designed to burnup on re-entry, but re-entry burnup of the reactor was not complete. Approximately 65 kg of radioactive debris and nuclear
6.3 Launch Abort Environments Affecting SNPSs
fuel was recovered from the crash site. It is very fortunate, indeed, that this impact occurred in an unpopulated area. The second accident involved Cosmos-1402, which was launched on August 30, 1982. During this RORSAT mission, an anomaly occurred and the spacecraft automatically separated into three parts. The Buk reactor is believed to have re-entered over the South Atlantic, approximately 1600 km east of Brazil on February 7, 1983. No radioactivity was detected from this re-entry or impact. This suggests that the design of the Buk space reactor was modified following the Cosmos-954 accident, to ensure its intact re-entry. The last Soviet space nuclear accident involved Cosmos-1900, which was launched on December 12, 1987. On May 13, 1988, the Soviet Union reported that it had lost communications with the Cosmos-1900 RORSAT in April, 1988, and that as a result the onboard Buk reactor could not be boosted to a higher, long-lived orbit via a ground command signal. At the same time, the Soviet Union reported that there was an automatic separation and boost capability built into the spacecraft. Subsequently, on September 30, 1988, the Buk reactor automatically separated from the spacecraft and boosted to a higher (~720 km) orbit. This event, although it does involve a malfunction, does not represent an accident, in the strict sense, since the situation was resolved by an onboard capability designed to automatically operate for such a malfunction.
6.3 Launch Abort Environments Affecting SNPSs A Space Nuclear Power System designed and tested for a specific mission (e.g., deep space, planetary) has the additional constraint of surviving transport from Earth to its mission location. This requires that it has to remain undamaged not only during the normal elevated vibration and shock environments of launch, but it also must be robust enough in its containment design to survive a worst case launch vehicle accident. These types of accidents normally include extreme shock, vibration, blast, fragmentation, overpressure, and thermal environments which expose the payload to their damaging effects. The following sections introduce the reader to the hazardous environment that could exist around the nuclear source during a launch abort accident. By evaluating the potential high temperature and shock conditions provided to the source in such a scenario, the design requirements for protecting the nuclear material can be established. Methodologies to handle the additional difficulty of predicting the potential for a launch failure concludes this section.
Pre-Launch and Launch Abort Scenarios and Issues The safety design and testing of nuclear material used in space power systems is predicated on survival of its containment in one of the most destructive environments
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FIGURE 6.7 Delta II GPS-IIR launch abort, January 17, 1977.
imaginable: the pressure and heat of an explosion. While the likelihood of a rocket to “spontaneously combust” is usually low, especially in more mature vehicles, the surprise attack has been demonstrated, as was the case on January 17, 1997. Figure 6.7 depicts the 241st launch of a Delta II rocket (the 35th Delta 7925 model) 12.5 seconds into flight. Critical to the design of a containment system for nuclear material is a thorough understanding of the potential explosive environment through historical and state-of-the-art characterization of launch vehicle fires and explosions. Characterization begins with the type of launch vehicle used to transport the space nuclear power system. A brief description of typical rockets and their explosive potentials leads to discussions on the physics of rocket propellant explosions, the blast overpressure hazard, fragments induced by these explosions, and the possible thermal fireball environment subjected upon the payload. Just how likely is it that a launch failure will occur? A description of probabilistic studies of these events is included.
Launch Vehicle Types While there have been a plethora of new launch vehicles arriving on the scene, the basic propulsive technology applied to these rockets dates back to the 1960s era. It seems the methods for propelling spacecraft into orbit over the last eight decades has remained consistent and provides for two fundamental elements e liquid propellant and solid propellant. What follows is a brief description of these propellant types, their applications in launch vehicles, and some typical quantities found on current rockets. In 1926, when Robert Goddard (Figure 6.8) launched his rockets, the technology to propel objects through the air was already well understood by the Chinese who in the 13th century had successfully lofted destructive missiles at their enemies using
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.8 Robert Goddard’s liquid fueled rocket, circa 1926.
what are today know as fireworks (Figure 6.9). While the Chinese propulsion element of choice consisted mainly of gunpowder (or solid propellant), Goddard, after improving on their designs, turned to liquid propellant. Both have their advantages and disadvantages, some of which are shown in Table 6.4. All of the rockets launched today use one or both of these forms of propellant in stages to reach their destination. New designs may combine both elements and are called hybrids. A combined form of gel propellant looks promising for bringing together benefits of both types. Table 6.5 lists some of the current launch vehicle inventory and the types and amounts of propellant utilized by each. This data will be useful in calculating the explosive environment later in this chapter.
Threat environment A very important aspect of the propellants used is their ability to impart energy to other parts of the launch vehicle during uncontrolled burning or explosive events,
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FIGURE 6.9 Chinese fire arrows.
particularly hazardous systems such as RTGs, which could in turn be demolished causing release of harmful radioactive material. The section on “Containment Design” describes the vulnerabilities of the nuclear source material such as clad melting temperatures, overpressure limits, and effects of vaporization. Predicting the environment that this material must potentially survive is the objective of the rest of this section.
Propellant Fires Regardless of the propellant combination (solid, liquid, gel), if the flame directed out the nozzle to provide thrust escapes the confines of the combustion chamber or exit, due to leaking or burn-through, the result will likely be a propellant fire with subsequent explosion once lower explosive limits are reached (Figure 6.10). Hypergolic propellants (e.g., hydrazine fuel and nitrogen tetroxide oxidizer) are designed to combust, without an ignition source, upon contact. Cracking or delamination of solid propellant can provide pathways for flame to travel to the outer sidewalls of the booster causing torch-like burn-through, which can impinge on
Table 6.4 Propellant pros and cons
Solid propellant
Liquid propellant
Advantages
Disadvantages
Storable Non-toxic (in storage) Simplified systems Throttleable Higher Isp
No stop/start capability Explosive impacts Lower Isp Toxic (N2H4/N2O4) Cryo storage (LOX/LH2) More complex systems
Table 6.5 Launch vehicle properties
Vehicle Delta 7425
Stg 0
4 Ground Lit GEMs
1
Delta III
Delta IV (medium)
Solid quantity lb
S
HTPB
25,942 ea
L
STAR48 6 Gnd Lit, 3 Air Lit GEMs
1 2 3 0
Solid prop type
L S S
L S S
Liquid ox quantity lb [gal]
Liquid prop fuel
Liquid fuel quantity lb [gal]
LOX
146,070 [15,940] 8712 [720]
RP-1
66,504 [9597] 5250 [700]
N2O4 HTPB HTPB
LOX N2O4 HTPB HTPB
A-50
4405 25,942 ea
L
STAR48 6 Gnd Lit, 3 Air Lit GEMs
Liquid prop oxidizer
146,070 [15,940] 8712 [720]
RP-1
146,070 [15,940] 31,179 [3402] n/a
RP-1
A-50
66,504 [9597] 5250 [700]
4405 37,500 ea
1
L
LOX
2
L
LOX cat bed
1
L
LOX
2
L
LOX
371,321 [40,520] 38,571 [4209]
LH2 N 2H 4 LH2 LH2
66,504 [9597] 5788 [10,101] 160 [19] 68,679 [119,859] 6429 [11,220] (Continued)
6.3 Launch Abort Environments Affecting SNPSs
Delta 7925
2 3 0
Solid/ Liquid
279
280
Vehicle Delta IV (medium þ)
Delta IV (heavy)
Atlas IIA
Atlas IIAS
Stg 0
2 Gnd Lit/ 2 Gnd þ 2 Air Lit GEMs
Solid prop type
Solid quantity lb
S
HTPB
41,990 non TVC 42,500 TVC ea
Liquid prop oxidizer
Liquid ox quantity lb [gal]
Liquid prop fuel
Liquid fuel quantity lb [gal]
371,321 [40,520] 38,571 [4209]
LH2
1,113,963 [121,560] 38,571 [4209]
LH2
242,144 [25,489] 31,350 [3290] n/a
RP-1
68,679 [119,859] 6429 [11,220] 206,037 [359,577] 6429 [11,220] 107,616 [15,826] 5900 [10,000] 170 [20]
242,144 [25,489] 31,350 [3290] n/a
RP-1
289,289 [31,568]
RP-1
1
L
LOX
2
L
LOX
L
LOX
2
L
LOX
1
L
LOX LOX cat bed
1
3 core boosters
2
Centaur
L
0
2 Gnd Lit 2 Air Lit Castor IV-As
S
1 2
Atlas III
Solid/ Liquid
1
Centaur
HTPB
LH2
LH2
LH2 N2H4
22,300 ea
L
LOX
L
LOX cat bed
L
LOX
LH2 N2H4
107,616 [15,826] 5900 [10,000] 170 [20] 112,072 [16,172]
CHAPTER 6 Nuclear-Powered Payload Safety
Table 6.5 Launch vehicle propertiesdContinued
Atlas V
Titan IVB
Centaur III
L
0 1
0 to 5 SRBs
S L
2
Centaur III
2
Centaur III (stretched)
31,616 [3450]
LH2
5990 [10,454]
LOX
452,906 [49,422] 31,350 [3290] n/a 38,077 [4155] n/a 1,358,824 [117,140] 28,324 [2442] 5554 [479]
RP-1
Jet-A
174,194 [25,136] 5900 [10,000] 170 [20] 6923 [12,082] 170 [20] 181,176 [26,801] 3776 [559]
Jet-A
741 [110]
A-50
117,240 [12,540] 27,570 [2950] 123 [14.6]
102,000 ea
LOX cat bed L
LOX cat bed
1
L
H2O2
2
L
H2O2
3 0 1
2 Gnd Lit SRMUs
2
STS
HTPB
LOX
L S L
H2O2 HTPB
IUS
S
N2O4
3
Centaur
L
0
2 Gnd Lit SRBs
S
PBAN
Jet-A
cat bed
222,760 [18,410] 49,630 [4100] n/a
N 2H 4
LOX cat bed
38,260 [4000] n/a
LH2 N 2H 4
n/a
n/a
N 2H 4 APU
N2O4 HTPB
LH2 N 2H 4
695,427 ea
L
3
LH2 N 2H 4
SRM-1 21,404 SRM-2 6002
1,287,188 ea
A-50
7940 [13,400] 170 [20] 67 [8]
6.3 Launch Abort Environments Affecting SNPSs
Beal BA-2
2
(Continued)
281
282
Vehicle
Pegasus Proton D-1-e
Zenit 3SL
H-2A
Stg
Solid/ Liquid
Solid prop type
Solid quantity lb
Liquid prop oxidizer
Liquid ox quantity lb [gal]
Liquid prop fuel
Liquid fuel quantity lb [gal]
LOX
1,361,936 [143,351] 1549 [128]
LH2
227,641 [385,265] 946 [129]
ET
L
Orb 1 2 1
L S S L
2
L
N 2O 4
3
L
N 2O 4
N2O 4 HTPB HTPB
N2O 4
4
DM-5L
L
LOX
4
or Breeze
L
N2O 4
1
L
LOX
2
L
LOX LOX
3
DM
L
0 0
2 Gnd Lit SRB-A 2 Solid Boosters (opt.) 1-2 Liq Boosters (opt.)
S S
0
L
HTPB HTPB
MMH
26,790 6670 659,014 [54,464] 276,893 [22,884] 82,561 [6823] 23,369 [2460] 22,120 [1828] 511,044 [55,766] 129,278 [14,107] 23,369 [2460]
UDMH
187,039 [20,410]
LH2
UDMH UDMH SYNTIN UDMH RP-1 RP-1 SYNTIN
244,986 [37,119] 67,207 [10,183] 20,039 [3036] 9480 [1335] 11,060 [1676] 196,556 [28,363] 49,722 [7175] 9480 [1335]
133,400 ea 28,900 ea LOX
31,701 [55,325]
CHAPTER 6 Nuclear-Powered Payload Safety
Table 6.5 Launch vehicle propertiesdContinued
Ariane IV
Long March 3B
L
LOX
2
L
LOX
0 0
0-4 Gnd Lit SRBs 0-4 Liq Boosters (opt.)
S L
CTPB
N2O4
L
N2O4
2
L
N2O4
3
L
LOX
2 Gnd Lit SRBs
2 0
Aestus 4 Liquid boosters
1 2 3
S L
HTPB
LH2 LH2
31,957 [55,771] 6100 [10,646]
20,900 ea
1
0 1
188,543 [20,574] 30,500 [3328]
54,148 [6435] 323,630 [26,746] 47,876 [3957] 19,697 [2149]
UH25
287,280 [31,349] 14,384 [1189] 82,908 total ea 379,260 total 108,045 total 39,690 total
LH2
UH25 UH25 LH2
31,852 [4550] 190,370 [27,196] 28,162 [4023] 4103 [7161]
506,000 ea LOX
L L
N2O4 N2O4
L L L
N2O4 N2O4 LOX
MMH UDMH
54,720 [95,497] 7016 [957] unavailable
UDMH UDMH LH2
unavailable unavailable unavailable
6.3 Launch Abort Environments Affecting SNPSs
Ariane V
1
283
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CHAPTER 6 Nuclear-Powered Payload Safety
FIGURE 6.10 Titan I explosion, December 12, 1959.
secondary combustible systems (e.g., the Challenger accident). Should the combustible materials reach conditions for explosion, a fireball will be created. If the leaking propellant settles in a particular area on the ground and ignites (spontaneously or via a source) a pool fire is the result. Some historical launch pad explosions are shown in Figures 6.10e6.15. On- or near-pad accidents (altitude 170 ft) are assumed to be of such severity (major structural failures) that all of the onboard liquid propellants (except possibly the spacecraft’s) are released and are thus available for bulk pad fires (Marietta, 1996). During the propellant loading phase (pre-launch), the total propellant weight would be consistent with what had been currently loaded into the launch vehicle. For tanks being topped off, the mission specific weight would be directly proportional to the elapsed time since the start of loading, assuming a constant pumping rate. In addition, if destruct action occurred (either by automatic or commanded action) after the solid rockets have been ignited, solid propellant ground fires will be present, although considerably dispersed by the fragmentation process.
6.3 Launch Abort Environments Affecting SNPSs
(A)
(B)
FIGURE 6.11 (A) Juno II going horizontal and (B) exploding, July 16, 1959.
FIGURE 6.12 Titan I fireball sequence, December 12, 1959.
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FIGURE 6.13 Titan I fireball sequence, August 14, 1959.
FIGURE 6.14 Vanguard fireball sequence, December 6, 1957.
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.15 Convair Atlas.
Increasing flight speed and altitude (170 ft h 3800 ft) will usually disperse the liquid propellants. Consequently, the reduced concentration of mixed propellant that falls back will greatly reduce any fireball formation at the ground. Airborne reactions will form more of a combustion cloud (like the Titan 34D-9, see Figure 6.16) than a ground fireball. Solid propellant ground fires resulting from fragmentation of the boosters will be more dispersed than for on- or nearpad accidents. During the period from land clear to shortly after payload fairing jettison, high flight dynamic pressure or rapidly decreasing ambient pressure will further increase the liquid propellant dispersion and reduce the intensity of airborne combustion. The flight path is completely over water so there is no ground fire hazard from either the liquid or solid propellants. Once at high altitude (~475,000 ft), the only thermal environment hazard to the payload’s nuclear material is from accidental launch vehicle re-entry. Aeroloading and heating, with subsequent breakup during fall back from this altitude regime, are covered in the re-entry section.
Liquid propellant fireballs A conceptual model of a fireball is shown in Figure 6.17. Combusting propellants embedded within the fireball are the source of the combustion products (Dobranich, 1997). The physics of fireball formation is thoroughly described in Sandia Report SAND97-1585 (Dobranich, 1997), which explains their Fireball
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CHAPTER 6 Nuclear-Powered Payload Safety
FIGURE 6.16 Titan 34D-9 explosion at Tþ9 seconds.
Integrated Code Package simulating responses of physical and chemical processes occurring in a fireball and is presented here:
Fireball temperature Based on a single control volume, the time-dependent temperature of the fireball is determined by solving the following energy balance equation (assuming a constantpressure combustion process): d nf h f ¼ RLþE (3) dt where nf is the quantity of combustion products and entrained air that comprise the fireball (mol), hf is the molar enthalpy of the fireball (J/mol), t is time (sec), R is the reactant enthalpy inflow (W), L is the fireball energy loss rate (W), and E is the enthalpy inflow from air entrainment (W). Note that both nf and hf are dependent
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.17 Conceptual fireball model.
variables. The molar gas quantity in the fireball is: nf ¼ np þ na
(4)
where np is the quantity of combustion products (mol), and na is the quantity of entrained air (mol). The reactant enthalpy inflow term, R, is given by: R ¼ n_ r hr
(5)
where n_ r is the molar combustion rate of propellant reactants (mols/s), which is assumed known, and hr is the specified enthalpy of the reactants (J/mol). Likewise, the air entrainment enthalpy inflow term is given by:
E ¼ n_ a ha
(6)
where n_ a is the molar rate of air entrainment (mol/s), and ha is the enthalpy of the ambient air (J/mol). Equation (6) can be expanded using the chain rule and rearranged to provide dhf ¼ n_ r hr n_ p hf L þ n_ a ha hf nf dt
(7)
dnp is the molar product production rate (mol/s). dt The product production rate, n_ p , is expressed in terms of the reactant combustion rate using: dnr ¼ yn_ r (8) n_ p ¼ y dt where n_ p ¼
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where y is the molar quantity of combustion products produced per mole of reactant, and n_ r is the specified reactant combustion rate (mol/s). This equation is required for the solution of eq. (8). Based on equilibrium chemistry, the fraction y is determined using: y ¼
np Wr ¼ nr Wp
(9)
where np is the quantity of products (mol), nr is the quantity of reactants (mol), and Wr and Wp are the mean molecular weights of the reactants and products, respectively (g/mol). The mean molecular weights are given by: Wr ¼
Nr X
y jr Wrj
(10)
yip Wpi
(11)
j¼1
and Wp ¼
Np X i¼1
where Nr is the number of reactant species, Np is the number of product species, yrj and yip are the mole fractions of reactant species j and product species I, respectively, and Wrj and Wpi are the molecular weights of the reactant and product species, respectively (g/mol). Equation (7) can now be rearranged to provide: dhf ¼ n_ r hr yhf L þ n_ a ha hf nf dt
(12)
This equation is based on the following definition of enthalpy: h ¼ hF þ Dh
(13)
where hF is the enthalpy of formation (also known as the heat of formation) at the standard state (298.15 K, 1 atm), and Dh is the change in enthalpy from the standard state to some other state, including sensible heat and heats of transition when appropriate. The loss term in eq. (7), L, is given by: 4 þ hAl AAl ðT TAl Þ qAl L ¼ ε s A T 4 Ta4 þ εAl ε s AAl T 4 TAl
(14)
where ε is the effective emissivity of the fireball, s is the StefaneBoltzmann constant (5.67 10e12 W/cm2,K4), A is the surface area of the spherical fireball (cm2), T is the temperature of the fireball (K), Ta is the ambient environment temperature (K), εAl is the surface emissivity of the aluminum structures within the fireball, TAl is the surface temperature of the aluminum structures (K), hAl is the heat transfer coefficient for convective heat transfer between the aluminum structures and
6.3 Launch Abort Environments Affecting SNPSs
the fireball (W/cm2,K), and qAl is the heat added to the fireball from aluminum combustion (W). Calculation of the aluminum surface temperature and the combustion source term are not described in this chapter. Qualitatively, eq. (3) indicates that the time rate of change of the fireball energy equals the rate of energy inflow from reactants and air entrainment minus the rate of fireball energy loss. The loss term in eq. (14) includes four contributions: (1) thermal radiation from the fireball to the ambient environment; (2) thermal radiation from the fireball interior (treated as gray semitransparent participating medium) to the aluminum structures (the rocket) within the fireball; (3) convective heat transfer from the fireball interior to the aluminum structures, and (4) heat added to the fireball from combustion of aluminum structures. Following liftoff, only the first loss term contribution is included because the structures are assumed to no longer be immersed in the rising fireball. Using the calculated fireball enthalpy and product mole fractions, the fireball temperature is determined using: hf ¼ ð1 ya Þ
Np X
yip hip þ ya ha
(15)
i¼1
where ya is the molar air fraction defined as the molar quantity of entrained air in the fireball divided by the molar quantity of air and all combustion products in the fireball, yip is the mole fraction of product species i, hip is the enthalpy of product species i (J/mol), and ha is the enthalpy of the air entrained into the fireball (J/mol). The molar air fraction is defined as: na ya ¼ (16) nf Because hip and ha are functions of fireball temperature T, eq. (15) must be solved iteratively for temperature. This is accomplished using Newton iteration, recognizing that the derivative of enthalpy with respect to temperature at constant pressure is equal to cp, the constant-pressure specific heat (J/mol,K). Enthalpies are provided as fifth-order polynomials in temperature and account for heat of formation, sensible heat, phase transitions, and corrections for non-ideal gas behavior. Specific heats are provided as fourth-order polynomials in temperature.
Empirical formulae Roberts (1982) used the empirical data of Hasegawa and Sato (1977) to correlate the measured radiation flux q received by a detector at a distance L (m) from the center of a fireball with a hydrocarbon fuel mass (AICE, 1994), mf (kg): q ¼ 828 m0:771 L2 f
(17)
Heat flux density thermochemical model predictions are shown along with upper bound experimental data, plotted versus dimensionless time, t/tb, in Figure 6.18. Temperature within the fireball can be estimated from the heat flux data by assuming
291
CHAPTER 6 Nuclear-Powered Payload Safety
440
5000
3064 396
352 4500 308
264
MO D
220
P EL
2898
C DI RE TI
P EX
DA
FIREBALL STEM LIFTOFF (1.5 × tb)
176
2803
O
.
N FO RL OX/L H 8 LO EX X/RP-1 (U.B .) P. D ATA FOR H YPERGOLS (U.B.) TA
4000
HEAT FLUX (kW/m2)
TH ER MO CH EM IC AL
FIREBALL SPHERE LIFTOFF (tb)
HEAT FLUX (BTU/sec/ft2)
2985
TEMPERATURE (K)
292
132 3500 88
44
0
0 0
0.5 1 1.5 NONDIMENSIONAL TIME (t/tb)
3000
2
FIGURE 6.18 Liquid propellant heat flux versus non-dimensional time.
that radiation effects dominate and the emissivity is 1.0 (blackbody). Consider the initial radiant flux of 2000 kW/m2. This corresponds to a temperature of 2433 K via the StefaneBoltzmann relation: Eb ¼ s T 4
(18)
where Eb is the blackbody emissive power, or heat flux (kW/m ), s is the Stefane Boltzman constant, and T is the fireball temperature (K). The maximum fireball temperatures are higher than this value, but rapidly drop down within milliseconds. 2
Fireball rise time Failure of the propellant containment system can lead to a subsequent explosion that can be located near the ground, during early flight, or in the air, during later flight. In
6.3 Launch Abort Environments Affecting SNPSs
surface fireballs associated with propellant explosions, as the pressure of the detonation products decreases to ambient, the density of the gas is considerably below ambient density and the resultant buoyant force causes the gases to begin to rise. The bulk of the propellant is then entrained in the fireball and rapidly burned (High, 2006). A hemispherical shape is maintained during most of this period until the fireball growth rate is exceeded by the buoyancy and a spherical shape develops. After the spherical shape is completely formed, the fireball lifts from the ground, and surrounding air is drawn into the fireball by a combination of convective forces and a vortex motion to continue mass addition. A stem is created with propellant spilled on the ground to produce a mushroom-type cloud. The hot fireball continues to change into an oblate spheroid and finally a torus, as seen in Figure 6.19. Combustion between the fuel-rich gases and entrained air dilutes and cools the gases. The radiative losses also contribute to a rapid cooling process. The dynamic behavior of the fireball can be described as a spherical thermal composed of hot fireball gases on top and entrained air below. The development and rise of this thermal are
FIGURE 6.19 Fireball rise model.
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CHAPTER 6 Nuclear-Powered Payload Safety
controlled by thermal radiation, addition and combustion of propellant, and air entrainment. Following completion of the combustion stage, the fireball lifts from the ground, rapidly entraining air. The molar rate of air entrainment, based on an assumed entrainment coefficient, is determined using: n_ a ¼
auAp RT
(19)
where a is an entrainment coefficient, u is the rise velocity of the fireball (cm/s), A is the surface area of the fireball (cm2), p is the pressure of the fireball (atm), which is assumed to equal the ambient pressure, R is the universal gas constant (82 cm3,atm/ mol,K), and T is the fireball temperature (K). The quantity of entrained air is added to the existing product species in the calculation of fireball temperature via eq. (15). Also, during the entrainment stage, the reactant enthalpy inflow term is zero. The air entrainment model is based on the “macro-scale” process of a bubble (the fireball) rising through a medium (air), which is applicable after fireball liftoff. Such a model is common in bubble flow and atmospheric simulations. However, the model is also used before fireball liftoff, when air entrainment is a “micro-scale” process in which turbulent eddies near the boundaries of the fireball bring in ambient air. Because these turbulent eddies are not captured in the fireball model, there is no reasonable way to predict this entrainment. The micro-scale entrainment is expected to be much smaller than the macro-scale entrainment. Different entrainment coefficients may be used during the combustion and entrainment stages to reflect the different processes (e.g., a ¼ 0.025 during the combustion stage and 0.25 during the entrainment stage). It is also possible to introduce entrained air by specifying it as a reactant; thus micro-scale entrainment calculated by some other means (such as a computational fluid dynamics code) can be incorporated in the fireball simulation if desired. The rise velocity of the fireball is determined from solution of the following momentum balance: Cd d r A P u2 ðrVuÞ ¼ gV ra r 2 a dt
(20)
where r is the density of the fireball (g/cm3), V is the volume of the fireball (cm3), g is the acceleration due to gravity (981 cm/s2), Cd is the drag coefficient, ra is the density of the ambient air (g/cm3), and AP is the projected area of the fireball (cm2). The first term to the right of the equal sign is the buoyant force due to the density difference of the ambient air and the fireball, and the second term is the drag force. Application of the chain rule and rearrangement yields the following equation for numerical solution: du ¼ c2 c1 u c3 u2þb dt
(21)
6.3 Launch Abort Environments Affecting SNPSs
with
" ln 1 dðrVÞ d lnðrVÞ ¼ z c1 ¼ rV dt dt r c2 ¼ g a 1 r
kþ1 # nf W f k nf W f Dt
(22) (23)
and c3 ¼
1þb b bþ2 p að2Þb1 ra r ma nf Wf
(24)
where b is a constant that depends on the Reynold’s number, nf is the quantity of all fireball constituents (mol), which includes combustion products and entrained air, Wf is the mean molecular weight of the fireball mixture (g/mol), the superscript k indicates the time step number, Dt is the selected time step (sec), r is the fireball density (g/cm3), a is another constant that depends on the Reynold’s number, ra is the density of the ambient air (g/cm3), ma is the dynamic viscosity of the ambient air (g/cm-sec), and r is the fireball radius (cm). The product nfWf, which is the fireball mass, is calculated as: nf Wf ¼ np Wp þ na Wa
(25)
where Wa is the molecular weight of air (28.97 g/mol). The drag coefficient, Cd, is expressed as: Cd ¼ a Reb
(26)
where Re is the Reynold’s number of the rising fireball given as
Re ¼
ra u2r ma
(27)
where u is the rise velocity (cm/s) and r is the fireball radius (cm). The constants a and b are both functions of the Reynold’s number, and for flow around a sphere are provided in Table 6.6, based on simple curve fits to graphical data. Although the fireball will deviate from a spherical shape, the errors introduced are considered insignificant for our purposes. Table 6.6 Drag coefficient constants Range
a
b
Re 1.9 500 > Re > 1.9 Re 500
24 18.5 0.44
1 0.6 0
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CHAPTER 6 Nuclear-Powered Payload Safety
The height of the center of the fireball is determined by integrating the rise velocity over time with z equal to z0 as the initial condition. Thus: Zt z ¼ z0 þ u dt (28) 0
A hemispherical fireball is formed initially. The fireball grows in the shape of a truncated sphere as it rises, attaining a full spherical shape when it lifts from the ground. This growth is depicted schematically in Figure 6.20. The volume and surface area of the fireball before liftoff are given by: 4 3 1 2 pr pl 3rc þ l2 3 6
V ¼
(29)
and A ¼ 4pr 2 2prl
(30)
l ¼ rz
(31)
with and rc ¼
pffiffiffiffiffiffiffiffiffiffiffiffiffiffi r 2 z2
(32)
where V is the volume (cm3), r is the radius of the fireball (cm), rc is the radius of the circle formed by the intersection of the fireball truncated sphere and the ground (cm), and z is the height of the fireball center (cm). Equations (29)e(32) are solved to determine r and A given V and z. Following liftoff, the volume and surface area of a full sphere are used.
z
r z rC
l
FIGURE 6.20 Fireball formation geometry.
ground level
6.3 Launch Abort Environments Affecting SNPSs
The volume of the fireball is based on the ideal gas law: V ¼
nf RT p
(33)
The quantity of all fireball constituents, nf, is found by integrating the rate of fireball growth, n_ f , over time. Thus: Ztend
Ztend ðyn_ r þ n_ a Þdt
n_ f dt ¼
nf ¼ 0
(34)
0
where the fraction y (eq. (9)) is defined as the molar ratio of combustion products to reactants (determined by the equilibrium combustion solution), and n_ a is the molar rate of air entrainment (mol/s). The effective emissivity of the fireball, accounting for all particles within the fireball and assuming non-absorbing fireball gases and diffuse particle surfaces, is given by:
P ε ¼ 1e
Lb
Ns
i¼1
εip C i Aip
(35)
where Lb is an effective thickness of the fireball (cm), Ns is the number of particle sizes, εip is the emissivity of particle size i, Ci is the number concentration of particles of size i (cme3), and Aip is the effective projected area of particle size i (cme3), which is given by p 2 (36) Aip ¼ ci dei 4 where ci is the dynamic shape factor for particles of size i, and dei is the diameter of the volume-equivalent sphere for particles of size i (cm). The effective fireball thickness, Lb, is usually referred to as the mean beam length, which for a sphere is given as: Lb ¼ 1:3r
(37)
where r is the fireball radius (cm). Before the fireball attains a full spherical shape, the mean beam length is approximated as: Lb ¼ 3:9
V A
(38)
Equation (35) is easily modified to approximately account for absorbing gases (CO2 and water vapor) within the fireball. Thus: PN s i i i L ε C A b p i¼1 p ε ¼ 1 1 εg e (39)
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CHAPTER 6 Nuclear-Powered Payload Safety
where εg is the effective gas emissivity based on the temperature, pressure, and mean beam length of the fireball product gases. The calculation of the gas emissivity is based on an engineering approach assuming that water vapor and carbon dioxide are the dominant contributors to emissivity. Thus: εg ¼ εH2 O fH2 O þ εCO2 fCO2 Dε
(40)
where the H2O and CO2 subscripts refer to water vapor and carbon dioxide gas, respectively, and f and Dε are correction factors. The emissivities for the two dominant gases along with the correction factors are based on approximate curve fits to experimental data and are given as: log10 εH2 O ¼ ð0:88424 þ 0:65524T Þ
þ ð0:49502 þ 0:294787T Þlog10 Lb yH2 O p 2 ð0:061316 þ 0:05272T Þ log10 Lb yH2 O p ¼ 1 þ 0:897 0:2 log10 Lb yH2 O p p 1 þ yH 2 O 1
f H2 O log10 εCO2 ¼
(41) (42)
i
h
1:34561 þ 0:0758127T 0:122506ðT Þ2 i h þ 0:35278 þ 0:072651T þ 8:4779 103 ðT Þ2 log10 Lb yCO2 p i h 2 þ 0:074333 þ 0:0262193T 9:8478 103 ðT Þ2 log10 Lb yCO2 p (43)
fCO2
¼ 1 0:768622 0:195377 log10 Lb yCO2 p 0:69448 log10 Lb yCO2 p
Dε ¼ sin
*
yH2 O p 4:2 108 h2 þ 1:24 104 h þ 0:0361 yH2 O þ yCO2 h ¼ Lb p yH2 O þ yCO2
(44) (45) (46)
where T is equal to T divided by 1000; T is the fireball temperature (K), Lb is the mean beam length (cm), y is the mole fraction, and p is the fireball pressure (atm). The data points for the curve fits were selected to emphasize the high temperatures expected in a fireball. Equation (39) for fireball emissivity couples the aerosol physics-governing equations, which are used to determine the agglomerated particle concentrations, to the fireball temperature and size equations, and to the structure and particle heat transfer equations. Both the fireball energy equation and the fireball rise velocity equation are solved using a RungeeKuttaeFehlberg approach. An adaptive time step algorithm is incorporated to ensure both equations are solved accurately. This algorithm uses the difference between a fourth- and fifth-order solution to control the truncation error
6.3 Launch Abort Environments Affecting SNPSs
as the solutions are advanced in time. Thus different combustion scenarios can be simulated without concern for time step selection. It is necessary to define gas mixture properties for viscosity and conductivity in terms of the fireball constituent properties. The mixing model of Wilke as modified by Herning and Zipperer is used for this averaging process: m ¼
Np X k¼1
"
, yk m k
1=2 # Np X Wj yj Wk j¼1
(47)
where m is the dynamic viscosity (g/cm,sec), Np is the number of gaseous product species in the fireball at the current time, yk is the mole fraction for species k, and Wk, is the molecular weight for species k. As suggested in a prior report (Reid, Prausnitz and Poling, 1987), the same equation is used for conductivity with m replaced with k (W/cm,K). The final equations presented here for fireball physics describe the calculation of the turbulent dissipation term, which follows the work of Turner. This term is required in the aerosol physics solution of agglomeration due to turbulent diffusion. Although particle agglomeration due to Brownian motion is expected to dominate, the enhancement of agglomeration due to turbulence within the fireball should not be neglected. The local turbulent energy dissipation rate is approximated as the product of kinematic viscosity and the square of the vorticity: m .2 (48) εT ¼ V y r where εT is the turbulent energy dissipation rate (cm2/s3), m is the dynamic viscosity . of the fireball gas mixture (g/cm,sec), r is the fireball density (g/cm3), and y is the gas velocity vector (cm/sec). The flow field within the fireball is assumed to be approximated by a spherical vortex, for which the stream function is: ur 2 sin2 q r2 (49) 1 2 J ¼ ry 4 where u is the vortex (fireball) rise velocity, r is the radial position within the vortex, q is the angle describing the azimuthal position within the fireball, and ry is the radius of the vortex. The radial and angular components of the velocity vector, yr and yq are given by: 1 vJ u r2 (50) yr ¼ 2 ¼ cos q 1 2 ry r sinq vq 2 and yq ¼
1 vJ u 2r 2 ¼ sin q 2 ry 1 rsinq vr 2
(51)
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CHAPTER 6 Nuclear-Powered Payload Safety
Now the vorticity cross product can be evaluated as: 5ur sinq . V y ¼ 2ry2
(52)
The turbulent dissipation rate, assuming isotropic turbulence, can be expressed as: εT ¼
25m u2 r 2 sin2 q 4rry4
(53)
Because the fireball is modeled as a single uniform control volume, the spatial dependence of turbulent dissipation must be removed. A volumetrically averaged dissipation can be derived as: 3 εT ¼ 4pry2
Z2p Zry Zp
6px εT r sinq dq dr df ¼ 4pry3
Zry Zp r 4 sin3 q dq dr
2
0
0
0
0
(54)
0
where x ¼
25 u2 m 4 ry4 r
(55)
Now the average turbulent dissipation rate is: εT ¼
5 u2 m 2 ry2 r
(56)
where the radius of the vortex is taken as the radius of the fireball. This quantity is evaluated based on the current state of the fireball and used for calculation of agglomeration due to turbulent diffusion. To determine the products of combustion and their relative amount, a calculation of the combustion thermodynamics is required. The thermodynamic properties of the fireball are necessary to determine the temperature and size of the fireball, and to assess the vaporization and agglomeration of the plutonium-bearing particles. To simplify the combustion thermodynamics, two assumptions are made in the Fireball Code Package: • •
All combustion product gases are considered ideal although corrections are made to enthalpy for non-ideal behavior. An equilibrium chemical solution describes the situation adequately, which is probably reasonable for the high temperatures and reaction rates involved.
Chemical equilibrium is based on the minimization of either Gibb’s free energy or Helmholtz free energy. The Fireball Code Package utilizes the minimization of Gibb’s free energy, which is a function of pressure, temperature, and chemical
6.3 Launch Abort Environments Affecting SNPSs
composition. For a mixture of Np product species, the chemical contribution to the Gibb’s free energy for the mixture is given by: g ¼
Np X
m j nj
(57)
j¼1
where mj is the chemical potential for species j. If pressure and temperature are constant then: dg ¼
Np X
mj dnj
(58)
j¼1
The condition for equilibrium is the minimization of free energy and is applied to this equation subject to mass balance and non-negativity constraints. Of principal interest for fireball simulations is the behavior of plutonium-bearing particles. The discipline of aerosol physics deals primarily with the agglomeration (or forming into round masses) of aerosol particles, of the same or different size, in a gaseous suspension. Generally, particles of diameter less than 0.01 cm (100 mm) are considered to behave as aerosols. Larger particles, often referred to as “rocks”, are assumed not to agglomerate. The objective of aerosol physics is to determine how the fireball environment modifies the size distribution of plutonium-bearing aerosol articles.
Empirical formulae Previous studies (Mansfield, 1969), (Willoughby, Wilton & Mansfield, 1968), (Gayle & Bransford, 1965) have shown that the size, duration, and temperature of fireballs of various propellant mixtures have nearly the same quantitative behavior. Also, the empirical models used to characterize fireballs are generally well established and accepted. These estimates are based on experimental data from tests with up to 25,000 lb of liquid propellant and summarized in (Reinhart, 1999). In the studies, the fireball size and duration relationships were estimated from accumulated test and incident data, and analyzed empirically to obtain the following relationships: The maximum diameter of the fireball is given by: D ¼ 9:56 W 0:325
(59)
where D is the diameter of the fireball (ft), and W is the combined (fuel and oxidizer) propellant weight (lb). The standard error in D (i.e., the standard deviation of the data from the fitted equation), is about 30%. The fireball liftoff time is given by: tb ¼ 0:572W 0:167
(60)
where tb (secs) is the point in the fireball development when buoyancy and entrainment become dominant.
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CHAPTER 6 Nuclear-Powered Payload Safety
The duration of visible radiation (in seconds) is given by: tT ¼ 0:197W 0:349
(61)
with a standard error estimated to be 84%. The stem liftoff time defines the point when the heat flux drops abruptly, ceasing to be a measurable threat, is given by: tc ¼ 1:5tb
(62)
Pool Fires The accidental release of hundreds of gallons of flammable commodity such as hydrazine may not lead to an explosion; rather a pooling effect may occur leading to a large surface area at the bottom of the launch structure awaiting a potential ignition source to set the liquid alight. Unfortunately, hypergolic propellants like hydrazine can be easily ignited by the heat of reaction when they come in contact with organic material, especially rust, which can be plentiful in the high saline environment of beach-side launch pads. The major hazard from a burning pool fire at the base of the launch vehicle is that it will provide an ignition source for subsequent leaks of fuel along the stages of the rocket leading upward to the nuclear source in the payload. Secondarily, radiant heat effects, while probably not extremely threatening to a distant payload, can provide damaging consequences to a clad RHU that has been ejected from the payload during an explosion or tip-over event. However, the aeroshell protecting the clad is designed for the intense heating of re-entry and it is unlikely that the radiant temperature of the pool fire could significantly damage the RHU.
Solid Propellant Fire Hazard Threats similar to those provided by liquid propellant vehicles are introduced to the nuclear material onboard the payload by launch vehicles utilizing solid propellant. However, while a liquid explosion can provide a relatively short duration fire environment in an airborne explosion, a solid rocket motor can fragment as a result of either mechanical failure or purposeful destruct action to halt an erroneous flight trajectory. These fragments can return to the ground either burning, if conditions are ripe to maintain the burning process, or unburned, in varying sizes. If the size of a fragment (i.e., greater than critical diameter), and impact velocity are sufficient, these secondary “bombs” can explosively burn forming large craters, blast overpressures and intense heat. If the distribution of these solid propellant fragments is such that one or more is near a nuclear energy source (e.g., exposed or clad pellet, RHU, or RTG), there is a danger of vaporizing the nuclear material.
Solid propellant composition The thermal environment exhibited by burning solid propellant is determined in large part by its chemical makeup. Other factors affecting variability in the thermal
6.3 Launch Abort Environments Affecting SNPSs
environment include the geometrical configuration/size of the burning fragments, the amount of fracturing due to mechanical damage, and ambient conditions. Solid propellant stores the fuel and oxidizer together in solid form. The fuel is typically powdered aluminum, and the oxidizer is ammonium perchlorate (Wertz & Larson, 1991). A synthetic rubber such as polybutadiene holds the fuel and oxidizer powders together. Table 6.7 shows the typical type and quantity of some solid propellants used on launch vehicles and also lists the ingredients of several popular propellants.
Burn rate The surface recession rate r, called the burning rate, is an empirically determined function of the propellant composition and certain conditions within the combustion chamber. These conditions include propellant initial temperature, combustion pressure, and the velocity of the gaseous combustion products over the surface of the solid (Hill and Peterson, 1970). This function is approximated by: r ¼ a pn
(63)
where a and n are constants for a particular propellant mix and p is the pressure within the combustion chamber. Values of the constants for several propellants are also shown in Table 6.7.
Combustion process The qualitative description of the combustion process which a solid propellant chunk will exhibit during ambient-atmosphere burning on the ground is presented here from Price et al. (1979): Because of slow aluminum combustion, the combustion zone has the macroscopic structure shown in Figure 6.21. In a thin region (roughly one centimeter thick) adjoining the propellant surface, the oxidizer and polymeric binder decompose and interact. This is followed by an extended Region A in which the aluminum droplets are burning, and the gaseous products are near equilibrium composition (i.e., equilibrium for the amount of aluminum burned at that point). Region B corresponds to a distance from the burning surface great enough so that all of the aluminum has burned (observed to be about 1 meter). Along the sides of the combustion plume there are regions where the environmental air has mixed with propellant products. Because the propellant products are highly fuel rich, air admixture leads to further reaction and heat release, concurrent with dilution and cooling with atmospheric nitrogen. The important aspect of this macroscopic description of the combustion zone is the recognition of the difference in reaction rate-controlling processes in the different regions. Near the propellant surface, the dominant rates involve binder and oxidizer decomposition and mixing. In Region A, the rate is limited by the aluminum droplet combustion. In Region B, combustion is constant. In Regions C and D, the rate is limited by air mixing. The principal scale effect in going to large propellant samples is the decreasing importance of edge effects (Regions C and D).
303
304
Vehicle use
Propellant (mfr.)
Ingredients (% by weight)
Titan IV SRMU
QDT (Alliant)
50% 19% 19% 09%
02% 01% Titan IIIC
UTP-3001 (United Technology Center)
44.34% 23.17% 16.12% 10.86% 05.26% 00.25%
Ammonium perchlorate, 200 m (oxidizer) Ammonium perchlorate, 20 m (oxidizer) Aluminum powder (fuel) Hydroxyl-terminated polybutadiene (HTPB) (binder) Dioctyl sebecate (DOS) (plasticizer) Other (curative, catalyst, bonding agents) Ammonium perchlorate, 200 m (oxidizer) Ammonium perchlorate, 8 m (oxidizer) Aluminum (fuel) Polybutadiene acrylonitrile (PBAN) (binder) Other (curative, plasticizer) Iron oxide (burn rate modifier)
Burn rate constant, a in/sec
Burn rate constant, n
0.0401 500 > p > 1300 psia T ¼ 80 F
0.30
0.0694 500 > p > 1200 psia T ¼ 70 F
0.25
CHAPTER 6 Nuclear-Powered Payload Safety
Table 6.7 Typical solid propellant ingredients
Space Shuttle SRB
TP-H-1148 (Thiokol)
69.9% 16.0% 14.0% 00.1%
Delta II GEM
QDL (Alliant)
69.000%
Minuteman III
ANB-3066 (Aerojet)
73.00% 15.00% 08.56%
03.44%
0.04424 200 > p > 1700 psia T ¼ 77 F
0.33
0.2353 500 > p > 1500 psia T ¼ 80 F
0.35
0.0597 400 > p > 1000 psia T ¼ 80 F
0.27
6.3 Launch Abort Environments Affecting SNPSs
19.000% 06.362% 03.602% 02.036%
Ammonium perchlorate (oxidizer) Aluminum (fuel) Polybutadiene acrylonitrile (PBAN) (binder, curative) Iron oxide (burn rate modifier) Ammonium perchlorate (oxidizer) Aluminum powder (fuel) HTPB (binder) DOS (plasticizer) Other (curative, accelerator, catalyst, bonding agent) Ammonium perchlorate (oxidizer) Aluminum (fuel) Carboxy-terminated polybutadiene (CTPB) (binder) Polybutene (plasticizer, crosslinker)
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CHAPTER 6 Nuclear-Powered Payload Safety
REGION C: Oxidizer, binder, aluminum, and air in equilibrium, stoichiometric mixture
ION D: REGION Mixing region,, with unreacted aluminum m present
REGION B: Oxidizer binder, bin Oxidizer, aluminum reacted, no air admixture REGION A:
1 m (observed)
1 cm
Oxidizer and binder reacted, aluminum oxidation proceeding
PROPELLANT
FIGURE 6.21 General structure of solid propellant plume.
Region A remains the same thickness for all sample sizes that are large enough for complete aluminum combustion before appreciable air admixture. The macroscopic structure of the combustion zone is substantially determined by microscopic details, as is the nature of the two-phase flow seen by an object immersed in the fire. Looking at the combustion sequence on a more microscopic scale, the oxidizer and binder on the propellant surface decompose to gases, while the aluminum particles (10 to 30 microns) accumulate and adhere to form larger agglomerate droplets up to 299 microns in diameter. The oxidizer and binder gases mix and react near the burning surface. The aluminum agglomerate droplets ignite and burn on the surface, then move out into the gas flow. There they burn so slowly that they produce an extended combustion zone roughly 1 meter in thickness. The final product flow is roughly 30% aluminum oxide, which is in liquid droplet form. Starting with aluminum agglomerate droplets in the 20 to 200 micron diameter range, the droplets burn by two paths that produce distinctive product oxide droplets. Fine (<2 micron) smoke droplets are formed in a flame envelope around each aluminum droplet. Oxide also forms and accumulates on the aluminum. This in turn forms a distinctive oxide lobe that eventually becomes a residual droplet when the aluminum is consumed. Thus an increasingly dense oxide cloud develops as the burning aluminum droplets and smoke move away from the propellant droplets, and larger ones being formed later as the larger aluminum droplets burn out. A series of chemical equilibrium calculations were made for UTP 3001 propellant at 1 atm at the Air Force Weapons Laboratory, with percent aluminum reacted as
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.22 Gaseous reaction products as a function of aluminum burned.
a parameter. The results are shown in Figures 6.22 and 6.23. In these calculations, the temperature of the unburned aluminum was assumed as shown in the figure. Of the three temperatures referred to, 2533 K is most plausible for Region A, while lower values are appropriate near the burning surface, which corresponds to near 0% aluminum reacted. From these figures it can be seen that the aluminum combustion brings the temperature in Region A up from 2350 K at about a centimeter above the burning surface, to 2990 K at the start of Region B where aluminum combustion is complete. In the process, H2O and CO2 are reduced to H2 and CO. The equilibrium properties shown at 100% aluminum burned are applicable to Region B of the combustion zone. This is a major part of the fire environment of a large piece of propellant, encompassing that region more than 1 meter or so from the burning surface, out to those regions where air admixture has become important. For a large propellant sample (e.g., 1 meter across the burning surface), the temperature calculated by
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3000
2900 Calculated Equilibrium Temperature, K
308
2800
2700 300 K
Assumed Temperature of 833 K Unbumed 2633 K Aluminum
2600
2500
2400
2300 0
20
40
60
80
100
Aluminum Burned, %
FIGURE 6.23 Calculated combustion zone temperature.
this means is probably more accurate than can be measured directly. The calculation for less than 100% aluminum burned tells nothing about the droplet size distribution at each point in the combustion zone or the spatial variation of temperature within the combustion zone. Table 6.8 provides a summary of the reaction zone characteristics for UTP 3001 propellant. The extended region of the combustion zone of the propellant in which aluminum droplets burn is a region where chemical heat release continues locally around the aluminum droplets and flows to the surrounding media, with corresponding temperature gradients. The most intense heat release is in the high temperature zone, which is a detached flame envelope around each aluminum droplet. In this flame envelope, the most energetic step is the formation of Al2O3 in liquid form, primarily on the surface of existing smoke droplets in that envelope. It is the resulting energy of that reaction, and the corresponding high temperature of those smoke droplets, that makes the combustion so luminous. At 1 atm the Al2O3 temperature is about 3800 K. In a convective environment the flame envelope may be drawn out into a tail. From a practical viewpoint, it should be stressed that the high temperature
6.3 Launch Abort Environments Affecting SNPSs
Table 6.8 Estimates of reaction zone characteristics for UTP 3001 Propellant burn rate (measured) Propellant density (measured) Propellant mass burn rate Combustion zone thickness (Region A, observed) Burning time of large agglomerates Temperature after complete burning of Al (calculated) Gas density in the combustion zone Mass of Al2O3 per unit volume in the product flow1 Flow velocity2 1 2
0.127 cm/sec 1.76 g/cm3 0.233 g/s,cm3 1.0 meter 0.07 sec 2990 K 0.78 10–4 g/cm3 0.31 10–4 g/cm3 15–30 m/sec
Based on chemical equilibrium data actually applicable to Region B. Based on several methods of determination.
around the burning agglomerates, while the source of a large part of the radiant energy and heat, contains very little mass, and only small droplets that are not expected to be deposited extensively on objects in the flow. Inside the flame zone, the aluminum droplet is necessarily at a temperature below the aluminum boiling point (2740 K) and above the Al2O3 melting point (2318 K). The reason is that the flame cannot support the high vaporization rate of the boiling point, while the combustion would be arrested, or drastically reduced, if the oxide were frozen (because it would encapsulate the agglomerate). The aluminum droplet, at a temperature that lies between 2318 and 2740 K, with 2500 K taken as a best estimate, and, in aggregate, representing a substantial mass, readily impinges on immersed objects, reacting exothermally after being retained on the object. This effect (impingement and retention on an immersed object) would be more pronounced near the burning surface of the propellant where the aluminum concentration is high. The third temperature zone of the aluminum combustion is all that volume outside the droplet flame envelope, where the temperature is something like the flame temperature of the propellant (with due regard for the percent aluminum consumed at the particular location). This temperature ranges from 2350 K near the propellant surface to 2990 K further out where the aluminum is 100% consumed. In summary, most of the fire environment consists of the following: •
• •
The bulk of the volume (gas, smoke, and residual oxide droplets) is at a temperature ranging from 2350 K (0% Al burned) to 3000 K (rounded from 2990 K, for 100% Al burned). Only the residual oxide droplets are large enough for impingement. The unreacted aluminum in burning droplets is at about 2500 K. Droplets are large enough for impingement. The flame envelope around the aluminum droplets, containing a very small mass of hot gases and Al2O3 smoke droplets, is at about 3800 K. It is highly visible because of its high temperature.
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Spatial distribution of temperature is as follows: Near the propellant burning surface (lower Region A) there is a heavy flow of burning aluminum (droplet temperature about 2500 K), with a correspondingly large flame envelope radiation (temperature about 3800 K); and with a correspondingly low smoke density in the rest of the volume (temperature range of 2400 to 2500 K). As the material flows further from the surface (outer Region A), the aluminum droplets decrease in size and total mass while remaining at about 2500 K; the total volume and radiation of the flame envelope (temperature of 3800 K) decreases; the population of both smoke and residual oxide droplets increases and the temperature rises towards 3000 K if the air admixture does not intervene unduly. Beyond about 1 meter (Region B), the aluminum droplets and their flame envelopes disappear, leaving only a homogenous flow of gas, oxide smoke, and residual oxide droplets at about 3000 K. The radiation field in the combustion zone showed spectral lines of several gaseous species and a higher level of thermal radiation from the Al2O3 smoke cloud. Measurements of the smoke cloud 15 cm from the burning surface indicated a temperature of about 2800 K, an emissivity of 0.03, and an absorption coefficient of about 0.1/cm. Intensity-emissivity measurements are wavelength-dependent, apparently due to the presence of a small amount of smoke that is in the agglomerate flame envelope, which is much hotter than the bulk volume.
Experimental measurement Immersion heat probes, including bare thermocouples, immersion calorimeters and water-cooled calorimeter tubes, were used to investigate the fire environment. The environment experienced by these types of devices would be similar to a test object in the fire such as an RTG. Bare thermocouples experienced direct impingement with droplets in the flow, gave very erratic readings, and survived only briefly. Some measurements made in the wake of deflectors survived longer and gave better records; however, it was not clear how the temperature there could be related to the free stream temperature, even if the thermocouple survived. In high-temperature flows where immersion probes deteriorate, measurements of the heat-up rate of the probe can be used to estimate the plume temperature when an equation for the heat transfer between the probe and the flow can be written. When only convection and radiation were considered, an inordinately high plume temperature was inferred. This demonstrated that the heat transfer from deposition and reaction of condensed material on the probe was a major component of the heat transferred to the probe. In tests at Los Alamos National Laboratory to determine the response of Light Weight Radioisotope Heater Units (LWRHUs) when exposed to solid propellant fires (Tate & Land, 1985), thermocouples measured temperatures of 2333 K out to distances of at least 1.8 meters for UTP 3001 propellant. In tests at Lawrence Livermore National Laboratory to characterize the thermal environment of burning Minuteman III Stage 3 propellant (ANB-3066) (Diaz, 1993, 1994), spectral radiometry and pyrometry methods were used to determine plume temperatures. Temperatures along the centerline of the plume were observed by using graphite sight tubes, with the inboard end cut at an angle. The sight tube was inserted
6.3 Launch Abort Environments Affecting SNPSs
into the plume with the slant cut facing in the direction of flow of the plume. The sensor (for spectrometry or pyrometry methods) and sight tube were mounted, one on each end, to an alignment tube. Heat fluxes to objects engulfed in the plume were estimated by imbedding thermocouples in solid metal objects and placing them in the plume. Inverse modeling calculations were then completed using a heat transfer model to duplicate the observed temperature history of the calorimeters. Two devices were used. One consisted of a 2.5 cm diameter by 15.25 cm long molybdenum rod with two tungstene rhenium thermocouples embedded on the centerline at 3.18 cm and 8.25 cm from the exposed end of the rod. An additional thermocouple was placed on the wall at 8.25 cm from the exposed end. The rod was wrapped with an insulating ceramic blanket and placed in a holder that exposed only one end of the rod. The second calorimeter was a 5 cm diameter by 5 cm long stainless-steel cylinder, with one end machined into a hemisphere. The other end was drilled to the center point of the hemisphere and a tungstenerhenium thermocouple was placed in the center. The temperature, emissivity, and heat flux profiles obtained from these experiments are shown in Figure 6.24. The temperature is seen to vary from 2273 K at the burning surface to a peak at 2573 K about 32 cm off the initial surface. The heat flux of about 20 cal/ cm2$sec was measured at a standoff of 25 cm. There was a high level of confidence in the experimental methods; however, the temperatures were observed to vary by 10% due to the turbulence and thermal characteristics of the propellant fire. ANB3066 has an adiabatic flame temperature of 3003 K. In tests at Sandia National Laboratories to characterize the thermal environment from an unpressurized burning bed of an AP, Al, and HTPB composite solid propellant Gill et al. (1994), measurements were made with arrays of slug calorimeters located adjacent to the plume and with cylindrical calorimeters inserted into the plume. The locations of the slug calorimeters were chosen so the plume was viewed from several vantage points. The analytic model viewed the plume as a well-stirred reactor, which could be characterized by a thermal loss that is a fixed percentage of the total heat released in the reactor volume, and by the reactor temperature. From this the plume was found to have an effective temperature near 2073 K and to radiate 2600
50
0.5
Temperature 30
2200 Heat Flux 2000
20
1800
10
1600 0
10
20
30
40
50
60
70
80
90
0 100
Standoff from Initial Surface of Propellant, cm
FIGURE 6.24 Temperature, heat flux, and emissivity profiles for ANB-3066 burning in air.
0.4 0.3 0.2 0.1 0.0
Emissivity
40
2400
Heat Flux, cal/cm2 s
Temperature, C
Emissivity
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CHAPTER 6 Nuclear-Powered Payload Safety
34% of the total heat released. To simulate a body immersed within the plume, cylindrical calorimeters were positioned horizontally above the propellant burning surface at two elevations, at 5 cm and 20 cm standoffs. A thin-walled (0.64 cm) and thick-walled (3.2 cm) version, both with a 10.2 cm outside diameter, were positioned 11 cm apart at each standoff. The calorimeters were divided into thermally isolated quadrants to provide a measure of circumferential variation of the heat flux. Each quadrant of the thick-walled unit and the top and side quadrants of the thin-walled unit had one thermocouple. The bottom of the thin-walled calorimeter had three thermocouples. They were oriented so that each quadrant preferentially viewed the top, bottom, and lateral heat fluxes. The heat flux versus time from each segment indicated a slow increase in flux to a maximum level. The thin calorimeters then melted while the thick calorimeters remained intact. The maximum heat flux levels are shown in Figure 6.25. The average heat flux into a calorimeter is 84.7 W/cm2, with a range of 8%. The average incident heat flux (W/cm2) to a specific quadrant of these calorimeters is given by: q ¼ 84:7 33:5x3 þ 11:5x4 x3 ¼ þ1 for facing up x3 ¼ –1 for facing down x3 ¼ 0 for being vertical
(64)
x4 ¼ þ1 for facing in x4 ¼ –1 for facing out x4 ¼ 0 for being horizontal
36
52
60
79
77
110
107 11 cm typ.
Thin walled calorimeters
60
Thick walled calorimeters 77
67
92
113
93
68
20 cm standoff
133 5 cm standoff
PROPELLANT BURNING SURFACE
FIGURE 6.25 Heat flux (W/cm2) into quadrants of cylindrical calorimeters immersed in the plume.
6.3 Launch Abort Environments Affecting SNPSs
The Chemical Propulsion Information Agency (CPIA) estimated the adiabatic temperature for Titan IV SRMU propellant burning at ambient pressure at 3097 K using the NASA Lewis e Chemical Equilibrium and Transport Properties of Chemical Systems Program (Gordon, McBride & Coveny, 1990).
Explosions Pre-launch to early flight phase explosions Even before the launch vehicle begins its escape from the gravity well there are scenarios in which onboard nuclear source materials can be damaged. The final preparations for integration of the nuclear source system and the launch vehicle/ spacecraft involve hazardous operations which, in failure situations, can lead to adverse environments similar to launch failure scenarios. Drops during crane lifts, electrical connection failures causing fires, electrostatic discharge to solid propellants, and other processing failures can lead to the destruction of the entire launch pad or vehicle processing facility. This was woefully exemplified on August 22, 2003, during final processing of Brazil’s Veı´culo Lanc¸ador de Sate´lites (Satellite Launch Vehicle) (VLS) at Alcaˆntara (Figure 6.26). Twenty-one people working at the launch site were killed when one of the rocket’s four first stage solid rocket motors ignited accidentally.
FIGURE 6.26 VLS-1 pre-launch accident, 2003.
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Probably the most hazardous threat to nuclear material onboard the payload occurs in a scenario in which the vehicle launches nominally, then shortly after liftoff executes an immediate rotation impacting the ground nose first. In this position, large amounts of unburned solid propellant can be placed either on top of or beside the nuclear source. Impact velocities sufficient to cause secondary explosions and subsequent burning of solid propellant sets up an environment conducive to potential removal of the nuclear material’s protective shielding followed by vaporization of the material. This scenario was played out (sans nuclear payload) on February 15, 1996, when the Chinese Long March 3B rocket veered off-course immediately after launch, crashing in the nearby village only 22 seconds later, killing an estimated 100 people and destroying 80 houses in Xichang (Figure 6.27).
Ascent phase explosions During later flight, as the vehicle increases velocity, an explosion below the payload would tend to throw the nuclear material forward of the flight path and since this segment would have different aerodynamic characteristics its ballistic trajectory should keep its impact point clear of the returning solid fragments. Figure 6.28 shows the ballistic instantaneous impact points for two times. The early breakup of the vehicle at Tþ20 seconds shows that debris tends to impact in one area as the flight is nearly straight up and the programmed turn
FIGURE 6.27 Chinese Long March 3B launch accident, February 1996.
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.28 Early versus later flight impacts: at Tþ20 seconds (left) and Tþ40 seconds (right).
down-range has not yet been initiated. Later flight is depicted in the second figure at Tþ40 seconds revealing the debris field has elongated as more impacts occur over a range of points due to varying ballistic coefficients of pieces. Intact impact of the spacecraft itself with the smaller quantities of liquid propellants minimizes the possibilities of overpressures that would harm the nuclear source. There is some concern that solid “kick” motors attached to the spacecraft could impact with sufficient velocity to damage the nuclear material or impose a thermal environment upon burning on the ground to vaporize the source.
Fallback secondary explosions Figure 6.29 depicts the bombardment of Launch Complex 17 as debris and solid propellant fragments fall during the Delta II explosion on January 16, 1997. These fragments can impact onto hard surfaces (e.g., concrete or steel, Figure 6.30), soft surfaces (e.g., sand or grassy soil), or water. Included in the falling debris, as well as the solid propellant, are any nuclear sources either maintained in their original configuration or freed by the explosive events. These can impact prior to or following a propellant fragment. The response of solid propellant to ground impact may range from possible ignition and burning to explosive burning or even a detonation. It can be characterized knowing the size of the fragment, its impact velocity, and the type of surface that it impacts. The consequence of this response can then be defined knowing the spatial relationship between the propellant fragment impact point and the nuclear material during the short time interval immediately following the solid propellant ground impact. For a solid motor fragment to sustain detonation it must be of supercritical size and it must be subjected to a shock stimulus sufficient to initiate detonation.
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FIGURE 6.29 Secondary impacts, Delta II explosion, January 16, 1997.
FIGURE 6.30 Impact Crater Delta II explosion, January 16, 1997.
Determining this critical diameter, at which detonation could occur, was one of the objectives of solid propellant testing.
Detonation response The mechanisms of propellant detonation are dependent upon the variables of size and impact velocity. Detonation is a process in which an explosive reacts to products very close behind a steady, high-pressure shock wave that is supported by the explosive reaction. Detonation is characterized by high rates of propagation (4e9 km/s), extremely high pressures (10e40 GPa), and a short time to the start of reaction (typically 1 ms) (Lee & Green, 1988). In a detonation, 10 cm of material is consumed
6.3 Launch Abort Environments Affecting SNPSs
fully in slightly more than 10 ms. Conversely, deflagration refers to a range of reaction rates involving burning. The term burning is usually applied to a slow deflagration reaction that does not generate significant overpressure and whose time scale to consume 10 cm of material is greater than 1 second. If the material has a relatively large surface area, then a high-velocity deflagration can involve measurable or damaging overpressure and a time scale of 100 ms to 10 ms to consume the material. In this case, the reaction is propagated by conduction, convection, and radiation, and not by a shock wave. The term “explosive burn” is used for high-rate deflagration with significant overpressure to distinguish it from a high yield event like detonation. There are several responses that the propellant can exhibit given these varying parameters. The shock-to-detonation (SDT) response is a process in which a shock wave entering an explosive material develops into a detonation (Figure 6.31). Each type of explosive has a characteristic distance of travel for the transition of a shock wave of given amplitude into a detonation. The deflagration-to-detonation transition (DDT) is a process in which a deflagration reaction evolves into a detonation. The term is somewhat misleading in that gas pressure created by the deflagration compresses unburned, porous explosive material to cause a reaction that eventually leads to an SDT process. The initial deflagration does not simply burn faster and faster until a charge detonates; rather, its role is to drive the compaction wave. Figure 6.32 shows a porous bed of explosive that is ignited. The burning explosive compresses unburned material to form a compaction wave and a reaction within the wave by the intensity of compaction. The intensity of the compaction wave finally grows to that of a shock wave, and a detonation develops. XDT is an initiation process that originally referred to an “unknown” (i.e., X) detonation transition. XDT now commonly means a delayed detonation in which impact damages and creates high porosity in an explosive and then causes a subsequent compaction wave that ultimately leads to a detonation (Figure 6.33). XDT is closely related to DDT in that a compaction wave is the route to detonation. However, the compaction wave in XDT is mechanically, rather than combustion, driven.
Unshocked, unreacted
Shock front
explosive
Detonation
Detonation
products
front
Entering Shock Wave Shocked, partially reacted explosive
Reaction zone
FIGURE 6.31 Shock-to-detonation transition (SDT)eshock wave entering explosive develops to detonation.
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Porous
Compaction
Detonation
wave
products
Container
explosive
Combustion of porous explosive
Compacted Combustion
explosive
Detonation front
Partially reacted explosive
products
FIGURE 6.32 Deflagration-to-detonation transition (DDT) e initial deflagration drives a compaction wave, which grows to the intensity of a shock wave and detonation occurs.
Undamaged
Impact compressed
Recompressed
explosive
explosive Compaction
explosive
wave
Fractured and rarified Impact surface
explosive
FIGURE 6.33 Delayed detonation transition (XDT) e impact damages an explosive and creates high porosity. The continuing impact then causes a compaction wave that leads to detonation.
Detonation hazard The hazards associated with ground impact induced solid propellant detonation, whether arising from SDT, DDT, or XDT, can be very severe if appropriate conditions exist to facilitate the detonation transition. Because of the severe blast overpressure and fragment environments potentially generated by solid propellant detonation, the special conditions necessary to cause SDT, DDT, or XDT should be examined in detail. Detonations typically produce: • • •
Impact crater dimensions two to three times greater than those expected for typical propellant impacts. Air shock overpressures near the crater as high as 2000 psi. Pressures within the crater region as high as detonation pressures (i.e., 20 to 30 kbar, or greater).
6.3 Launch Abort Environments Affecting SNPSs
The capability of the propellant segment or fragment to achieve sufficient impact velocity to undergo detonation must also be examined. In the case of the Cassini payload launched on a Titan IV, the flight profile made conditions necessary to achieve SRMU propellant detonation very unlikely. The launch profile limited maximum SRMU segment and propellant fragment impact velocities to less than about 600 ft/sec, with impacts most likely to occur on soft sandy soil. For mission elapsed times greater than approximately 23 seconds, fragments would impact in the Atlantic Ocean. These velocity limitations bound the upper limit of ground impact pressures to the 2e3 kbar range for soft sandy soil impact and 7e9 kbar for hard surface impact (e.g., steel or concrete). As shown in Figure 6.34, this threshold is well below the rather conservative impact velocity of 1000 ft/sec (and nominal pressure of 20e30 kbar) required for SRMU propellant detonation. The probability that solid propellant segments or fragments will explode on impact depends on their amount and impact velocity. For Class 1.3 solid propellant in general, Figure 6.35 shows these probabilities given propellant weight and impact speed.
Critical diameter In order to sustain the detonation shock wave, the propellant geometry must be such that the segment or fragment is supercritical. Below this critical size, the shock wave will lose strength and diminish below supersonic speed which will likely result in deflagration only. Burning propellant does have its hazards however, especially in providing a high thermal environment to exposed RHUs. Above critical size, chemical and physical interactions will “feed” the shock wave, maintaining state dynamics for continuation of the shock’s travel leading to detonation.
LAND
800 100% Yield Possible for Intact Motors/Segments
Minuteman Stage 2 hard surface impact
600
Significant Fracture/Explosive Burn of Bare Propellant
shotgun & plate tests hard surface impact
400 Fracture/Explosive Burn Threshold
impact motors or segments Ttan III 34D-9
LAND
200
WATER
Impact Velocity, ft/sec
WATER
Detonation Threshold (conservative)
lg chunks,2 - 4 x critical dia, hard surface impact
1000
Ignition Threshold
0 0
5
Titan IV Cassini SRMU Propellant Impact Velocity
15 10 Mission Elapsed Time (from T-0), seconds
20
23
25
FIGURE 6.34 Comparison of Cassini Titan IV SRMU segment and propellant fragment maximum ground impact velocities with impact thresholds.
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100% 90% 80%
40%
1, 00 0
lb
lb 00 2,0
10,0
50%
lb
60%
4,00 0
lb
00 l b
70% 20,0 00
Probability of Explosion
320
30% 20% 10% 0% 0
100
200
300
500 600 400 700 Impact Velocity (ft/sec)
800
900
1000
FIGURE 6.35 Probability of solid propellant (1.3) explosion given mass and impact.
There is a unique and explicit relationship between the critical diameter of a material and the critical dimensions of a charge of any shape with uniform cross-section cast from that material (Elwell, Irwin & Vail, 1967). This relationship is the critical geometry, sc, which is defined by: 4A (65) sc ¼ P c where A is the area of the cross section, P is its total perimeter, and the subscript c refers to the critical conditions. For a right circular cylinder, eq. (65) becomes: sc ¼ dc
(66)
where dc is the critical diameter. If the critical diameter is known, eqs. (65) and (66) define the critical ratio of cross-sectional area to perimeter for any shape. Since area and perimeter can be expressed in terms of the dimensions of the cross section, the critical value of the dimensions (i.e., the critical size) of any shape is thereby defined. Testing of various propellant sample shapes were used to verify and refine eq. (66) through statistical analysis of data. The empirical form of the equation is: sc ¼ ð0:92 0:03Þdc
(67)
Solid propellant experimental testing Most of the knowledge concerning the explosive behavior of hydrocarbon rubber binderealuminum fueleammonium perchlorate oxidizer (HC/Al/AP) propellant was gained in the 1960s from the results of Project SOPHY (Elwell, Irwin & Vail, 1967) conducted at Edwards Air Force Base, CA (Merril, Nichols & Lee, 1999).
6.3 Launch Abort Environments Affecting SNPSs
The SOPHY program looked at only 84% total solids (i.e., ingredients besides binder) propellant. More modern propellants have higher total solids loadings (86e90%) that produce higher thrust performance. Little is known about the explosive behavior of these propellants. Solid rocket motors for the Titan IV (SRMU) have 88% total solids and the Space Shuttle SRBs have 86% total solids (Table 6.7). The rocket propulsion industry had seemed to assume that explosive potential and impact velocity thresholds for violent propellant response would be similar to that found in the SOPHY program. Table 6.9 shows the range of critical diameters derived from experimentation. These data, and incidents like the 1986 Titan 34D-9 explosion at Vandenberg Air Force Base, CA, indicated that large booster propellants may be more explosive than previously thought and highlighted the need to improve our understanding of booster propellant explosive behavior. Technology now enables us to study this behavior more economically. Polyvinyl difluoride (PVF2) piezoelectric gauges and advanced recording instrumentation now more accurately measure and record times of arrival coupled with shock magnitude measurements and shock magnitude time histories. During the 1960s, impact experiments were carried out by rocket sled impact. Hardware and range maintenance and refurbishment costs involved high capital outlays and use of substantial manpower. Experimental methods for substantially sized propellant impacts have evolved from ramming a rocket motor or propellant sample on a rocket sled into a barrier, to putting impact barrier material on a rocket sled so that a static, highly instrumented test sample could more easily provide high speed data than in the moving sample impact process, to explosively driven steel plate impactors into instrumented propellant samples resting on relatively inexpensive, replaceable wooden stands. Advanced techniques in combination with computer code simulation could make very large propellant sample testing unnecessary. The Lawrence Livermore National Laboratory (LLNL) has developed a reactive response hydrodynamics computer code called CALE/PERMS (C-language based Arbitrary Lagrangian-Eulerian/Propellant Energetic Response to Mechanical Stimuli), which is moving in that direction. In 1996 the Propellant Impact Risk Assessment Team (PIRAT) was assembled to specifically investigate the explosive risk posed by a Titan IV-SRMU launch. Selecting propellant sample sizes based on critical diameters reported by the French (Brunet & Salvetat, 1988), and the Super HIPPO experiments in 1992 (Merrill, 1992), a test program was developed with the intent to have impact samples in
Table 6.9 Propellant testing data Propellant type 84% HC/Al/AP 88% HTPB/Al/AP 90% HTPB/Al/AP 90% HTPB/Al/AP
Experimental critical diameter 1778 mm (70 in) 200–250 mm (8–10 in) 83 mm (3.27 in) 305 mm (12 in)
Year tested 1967 SOPHY (Elwell, Irwin & Vail, 1967) 1988 French Study (Brunet & Salvetat, 1988) 1988 French Study (Brunet & Salvetat, 1988) 1992 Super HIPPO (Merrill, 1992)
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excess of critical diameter or at least close to critical diameter so that substantial energetic response to impacts could be observed. With support of CALE/PERMS it was projected that even data from samples smaller than critical diameter could provide useful inputs for computer code modification. Cylindrical diameters of 203 mm (8 in), 254 mm (10 in), 305 mm (12 in) and 559 mm (22 in) were selected. Using PVF2 gauges, detonation velocity probes, foil gauges, crystal pins, high speed data recorders, overpressure sensors, and high speed film and video recorders, test articles were initiated using the setup shown in Figure 6.36 with some variations for each diameter tested. After getting encouragement that critical diameters for 88% solid HTPB propellants might be in the range of 200 to 250 mm, it was found that the first booster propellant formulation tested has a critical diameter larger than 559 mm and, perhaps, twice as large as 559 mm. Although the propellant sizes are large in normal testing of explosive materials, they are definitely subscale compared with solid propellant booster motors that can contain more than 500 metric tons of propellant. Larger size samples might be employed in the future since investigation of characteristics above critical diameter is desired. Solid rocket booster propellant in explosive events seems to provide high atmospheric overpressure yields without being involved in true detonations. If detonations are steady state supersonic events
Nonelectric cap
1 1/2" C4 Sphere 20" Primacord, typ.
42" Primacord 2 EBWs
Nonelectric cap, typ. 21" Plywood template 5" × 21.5" dia. C4 Aluminum light shield
Instrumentation Crystal Pins 6" apart 2 columns PVF2 gauges @ center, radius, and mid-radius 4 sets
37" × 22" dia. propellant segment
7" × 22" dia. propellant segments
2’ and 4’ velocity gauges, 180° apart 3" × 22" dia. paraffin
FIGURE 6.36 Twenty-two inch critical diameter test apparatus.
6.3 Launch Abort Environments Affecting SNPSs
in the energetic material and deflagrations are explosive events characterized by subsonic shock in the energetic materials, perhaps a new definition for an explosive process may need to be developed (Merril, Nichols & Lee, 1999).
Additional computer modeling It appears that detonations in Class 1.3 composite propellants are unconfirmed (some physicists have stated that such detonations are impossible) rapid deflagration is entirely possible and has been observed experimentally. Class 1.3 presents less of an explosive hazard than its more volatile counterpart graded Class 1.1 (e.g., Trident missile propellant). There is data from flyer plate experiments at Lawrence Livermore National Laboratory (LLNL) that show that “violent” explosions were induced in 1.3 propellant from high speed impacts. The test data indicates these explosions were not from shock induced detonation (i.e., there was no SDT); instead the explosion was a rapid conflagration induced in test samples (Williams, 1987). The Solid Rocket Motor Ground Impact (SRMGI) model is an analytical/engineering code developed by RDA Nance, Crawford et al. (1990) to provide estimates of the effects of solid motor impacts. Previous work for the Air Force and NASA in analyzing the Titan 34D-9 failure, particularly ground impact explosions and blowout fragment ranges, led to the development of this code. It was subsequently expanded to include the aft segment of the Shuttle boosters, and Delta II Graphite Epoxy Motors (GEMs). Major subroutines determine the theoretical Sandia penetration depth (Sandia Labs developed penetration scale), gas generating rate void pressure, fragment accelerations and motion, venting, and final fragment velocities. Key assumptions involve percent increase in burn area and, for the stage impact, the cavity formation time, initial cavity size, and blowout fraction. Maximum fuel burn rates, soil burn area characteristics, and blowout fraction can be parametrically varied. The summary of critical velocities versus soil number and burn rate, as predicted by the modified SRMGI code, are presented in the semi-log plot in Figure 6.37. This figure shows the smaller GEMs require the highest speed (> 400 ft/sec) and hardest surfaces (S < 1.2) to produce runaway explosive. The Titan upper segments could go critical on somewhat softer surfaces at 300e400 ft/sec but require 500e600 ft/sec for soil numbers up to 2.4. Above S ¼ 2.4 the SRMGI code indicates that impacts of the Titan segments should not produce explosive deflagration even if burn rates are as high as 500 in/sec. Both the GEM and the Titan will not go critical if burn rates are less than 8e15 in/sec. The STS aft segments appear to present the greatest hazard. As indicated in the figure, the STS’s large size produced runaway condition for soil impacts from S z 1.4 up to S z 15 to 20 and for maximum fuel burn rates which are only 5e6 in/sec. For hard surfaces (S < 1.4) when the aft stage does not penetrate more than 6 ft (half the stage diameter, the minimum depth assumed in the code for this segment, impacting end on, to create a cavity) then it is assumed conditions would not be created for a runaway deflagration. The code, with all of its engineering approximations and assumptions, only indicates that conditions for runaway deflagration have been created.
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Runaway explosion regions
800
700 Impact velocity (ft/s)
GEM 18 Burn rate = 14 30
TITAN 8 1215 40 500
500
600 STS AFT SEGMENT (min. Br 5 in/s)
500
400
300
200 20
Muck
10
Bay mud
8
Top soil
6
Moist clay, fine sand
Alluvial soil
Packed soil
Hard clay
Dense wet sand
4 2 SANDIA Soil No.
1
Shale
Broken rock
0.8
Soft concrete, asphalt
0.6
Hard concrete, granite
324
FIGURE 6.37 SRMGI code runs for space booster SRMs.
Blast Overpressure Hazard A characteristic feature of explosions is blast. Gas explosions are characterized by rapid combustion in which high-temperature combustion products expand and affect their surroundings. In this fashion, the heat of combustion of a fueleair mixture (chemical energy) is partially converted into expansion (mechanical energy). Mechanical energy is transmitted by the explosion process into the surrounding atmosphere in the form of a blast wave (Lee & Green, 1988). If the combustion process within an explosion is relatively slow, then expansion is slow, and the blast consists of a low-amplitude pressure wave that is characterized by a gradual increase in gas dynamic state variables (e.g., pressure, density, and particle velocity) (Figure 6.38). If, on the other hand, combustion is rapid, the blast is characterized by a sudden increase in the gas dynamic state variables: a shock. Due to extensive military investigations on the destructive potential of high explosives, relating the explosive power of an accidental explosion to an equivalent TNT charge is usually the comparison of choice.
6.3 Launch Abort Environments Affecting SNPSs
Pressure
SHOCK LOW AMPLITUDE PRESSURE WAVE
t
FIGURE 6.38 Blast wave shapes.
TNT Equivalency The use of TNT equivalency methods for blast prediction purposes is quite simple. The available combustion energy is converted into an equivalent charge weight of TNT with the following formula: WTNT ¼ ae
Wf Hf ¼ am Wf HTNT
(68)
where Wf is the weight of the fuel involved (kg), WTNT is the equivalent weight of TNT or yield (kg), Hf is the heat of combustion of the fuel in question (J/kg), HTNT is the TNT blast energy (J/kg), ae is the TNT equivalency based on energy, and am is the TNT equivalency based on mass. If the equivalent weight of TNT is known, the blast characteristics, in terms of peak side-on overpressure of the blast wave, can be derived for varying distances from the explosion. Concerns about this peak overpressure caused by the explosion of either liquid or solid propellant are focused in the near-field where the nuclear source material resides. Far-field effects of explosive blast are more concerned with effects of pressure increases on facilities, equipment, window glass breakage and the safety of personnel in the immediate area of the launch pad or at farther distances where pressure waves can focus.
Blast scaling The upper half of Figure 6.39 represents how a spherical explosive charge of diameter d, produces a blast wave of side-on peak overpressure P, and positive-phase duration tþ, at a distance R from the charge center. Experimental observations show that an explosive charge of diameter Kd, produces a blast wave of identical side-on peak overpressure p, and positive-phase duration Ktþ, at a distance KR from the charge center (lower half of Figure 6.40). Consequently,
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R
d
I
P t+
KR
Kd KI
P Kt+
FIGURE 6.39 Blast wave scaling.
charge size can be used as a scaling parameter for blast. Charge size, however, is not a customary unit for expressing the power of an explosive charge; charge weight is more appropriate. Therefore, the cube root of the charge weight, which is proportional to the charge size, is used as a scaling parameter. If the distance to the charge and the duration of the wave are scaled with the cube root of the charge weight, the distribution of the blast parameters in a field can be graphically represented, independent of charge weight. This technique, which is common practice for high-explosive blast data, is called the Hopkinson scaling law (Hopkinson, 1915). The compilation of experimental data for TNT surface bursts are presented in Figures 6.40 and 6.41. Early tests of liquid oxygen and RP-1 fuel (kerosene) were conducted under project PYRO in the 1960s to substantiate this curve. Project PYRO, conducted in the late 1960s, consisted of a comprehensive program to determine the blast and thermal characteristics of the three liquid propellant combinations in most common use at the time in military missiles and space vehicles: liquid oxygen with RP-1, liquid oxygen with liquid hydrogen, and nitrogen tetroxide with Aerozine 50 (50% UDMH/50% N2H4) (Willoughby, Wilton & Mansfield, 1968). During the course of the program some 270 tests were conducted with these propellant combinations on weight scales ranging from 200 to 100,000 lb. This basic explosive test program was supplemented by analytical and statistical studies, laboratory-scale experimental studies, simulation tests with inert propellant combinations, and a series of high-explosive tests for calibration and evaluation purposes. Several test configurations were used. The Confinement by Ground Surface (CBGS) condition best simulated the
6.3 Launch Abort Environments Affecting SNPSs
25,000 LB DATA LARGE POOL DIAMETER DATA ALL OTHER DATA
400
200
100 80
PEAK OVERPRESSURE (psi)
60
40
20
10 8 6
4
2 Note: The curve is not fit to the data, but rather the TNT equivalent blast curve when the far field equivalent yield is extrapolated to the near field.
1
1
2
4
6 8 10 20 40 60 80 100 SCALED DISTANCE (ft/lb½) Source: Willoughby et al. (1968)
FIGURE 6.40 Project PYRO test: Confined by Ground Surface (CBGS) case.
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25,000 LB DATA 400
ALL OTHER DATA
200
100 80 60 PEAK OVERPRESSURE (psi)
328
40
20
10 8 6
4
2 Note: The curve is not fit to the data, but rather the TNT equivalent blast curve when the far field equivalent yield is extrapolated to the near field.
1
1
2
4
6
8
10
SCALED DISTANCE (ft/lb½)
FIGURE 6.41 Project PYRO test: Confinement by Missile (CBM) case.
20
40
60
80 100
Source: Willoughby et al. (1968)
6.3 Launch Abort Environments Affecting SNPSs
sequential ground impact of separate fuel and oxidizer propellant tanks. This condition could occur by over pressurization of tanks or by the fallback or topple-over of the entire launch vehicle. Mixing interaction between the two propellants involved consideration of two configurations for the CBGS. Examination of the various length-to-diameter (L/D, or L-over-D) ratios and propellant flow velocity combinations revealed that the two propellant masses could have led to two cases of study: when both propellants were vertical, and when both propellants were horizontal. Figure 6.42 represents the Confinement by the Missile (CBM) configuration in which case an internal failure is assumed to occur, and one propellant falls down into the other. Cause of this type of failure could be the rupture of a common bulkhead between fuel and oxidizer tanks due to over pressurization or fragments from
FIGURE 6.42 CBM case setup.
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a primary explosion (e.g., engine blowup) penetrating the bulkhead. To simulate this failure, a test setup as shown in Figure 6.42 was used with a small C-4 charge used to propel the breaker ram through the tempered glass diaphragm separating the oxidizer from the fuel. Duration of the CBM case was limited to the time that the propellants remained confined by the walls of the vehicle. This time was determined by the strength of the tankage, the rate of vaporization of the cryogenic material, the initial pressure in the tanks, and the initial ullage space. The mixing mode in this case is determined by the interaction of the upper propellant and lower propellant at the time of opening. The position and velocity distributions of the lower propellant at the moment of first contact are fairly well specified, since it was assumed to be in its original configuration and to have zero velocity. Those of the upper propellant, however, may have a large range of values, depending on how full the tanks are and how large an opening is created between them. If a relatively small opening is produced, the L/D ratio of the top propellant will be effectively very much greater than that of the bottom one, and its total weight (entering into the mixing at an early enough stage to be a matter of concern) will be much smaller than that for the bottom one. In addition, the top propellant will have an initial velocity given by the fluid head in the top tank and the pressure differential between the two tanks. As the opening between the tanks becomes larger, the L/D ratio and effective weight of the upper propellant more nearly approaches the values they originally had in the missile. In addition, at the time the opening between the tanks is the full cross section of the original tank, the velocity would reach the value given by the acceleration of gravity through the distance of the ullage space in the lower tank, provided there was no pressure difference between the two tanks. Where the opening between the tanks is significantly less than the full cross section of the vehicle, the different initial L/D values for each propellant can be treated by assuming that the L/D ratio refers to the overall vehicle geometry and by defining a Do/Dt ratio, where Do is the opening diameter and Dt the vehicle diameter. The CBGS scenario produces a hemispherical blast while the CBM case generates a spherical shape. When using these graphs, a scaled distance must be determined, equal to K ¼
R W
=
330
1 3
(69)
where R is the distance of interest in feet and W is the combined liquid propellant weight in pounds. Adjustment factors are applied to W to reflect the fraction of terminal yield and the geometry of the explosion wave (i.e., spherical or hemispherical). Empirical data on TNT equivalency of liquid propellant combinations were used to develop Table 6.10 from Air Force Manual 91-201 (AFSC, 2011). The net explosive weight of liquid propellant is the total quantity of fuel (e.g., liquid hydrogen, kerosene (RP-1)) and oxidizer (e.g., liquid oxygen, nitrogen tetroxide) that would mix and explode. Assume a prelaunch explosion of a Delta II first stage occurs at a distance of 69.3 ft from the nuclear power source installed in the spacecraft. Predict the overpressure
6.3 Launch Abort Environments Affecting SNPSs
Table 6.10 Liquid propellant TNT equivalence Propellant combination
TNT equivalence
LO2/LH2 B5H9 þ an oxidizer LO2/LH2 plus LO2/RP-1
60%1 60% Sum of (60% for LO2/LH2) þ (20% for LO2/RP-1) 20% up to 500,000 lb plus 10% over 500,000 lb 20% up to 500,000 lb plus 10% over 500,000 lb 20% up to 500,000 lb plus 10% over 500,000 lb 10% 10% 10% 10% 10% plus equivalence of the solid propellant 100% 100%
LO2/RP-1 LO2/NH3 B5H9 þ a fuel IRFNA and aniline IRFNA and UDMH IRFNA and UDMH plus JP-4 N2O4/UDMH plus N2H4 N2O4/UDMH plus N2H4 plus solid propellants Tetranitromethane (alone or in combination) Nitromethane (alone or in combination)
Recent testing at white sands may lower this value substantially.
environment at this distance assuming the propellants leak to the flame trench of the launch pad forming a pool that is subsequently ignited in a hemispherical explosion. The Delta II first stage propellant load of LO2 and RP-1 equals 212,574 lb from Table 6.5. Using a TNT equivalence rating of 20% for the LO2 and RP-1 combination from Table 6.10, a scaled distance can be calculated using equation (6) as follows: =
W
1 3
¼
69:3 feet ð212; 574 lb 20%Þ
=
R
1 3
ft
¼ 2:0
=
K ¼
1 3
lb
(70)
Looking at Figure 6.40 for the hemispherical case (CBGS) the overpressure at a scaled distance of 2.0 ft is approximately 300 psi. If the destruct system had inadvertently activated, opening the LO2 and RP-1 tanks to afford mixing and ignition at the tanks’ interface, a spherical explosion could be assumed and a factor of 0.5 would need to be applied to the total propellant weight, W in calculating the scaled distance. 69:3 =
ð212; 574 0:5 20%Þ
1 3
¼ 2:5
ft =
K ¼
1 3
lb
(71)
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10 00 STATIC OVERPRESSURE (psi)
332
WSTF PYRO Deep Hole 90%
100
50%
10%
10
1 0.10
10.0
1.0 SCALED DISTANCE
100.0
(ft/lb1/3)
FIGURE 6.43 Static overpressure for MMH/NTO propellant combination.
which relates to an overpressure of approximately 240 psi from Figure 6.40. Experimental data for more recent testing done with nitrogen tetroxide (N2O4) and monomethyl hydrazine (MMH) are shown along with PYRO deep hole results in Figure 6.43.
Impulse pressure The impulse, I, is the area under the positive portion of the overpressureetime response curve (Figure 6.39). This is the more damaging portion of the blast wave, therefore the impulse predicted will be used along with peak overpressure to characterize potential adverse effects of explosive commodities.
Solid propellant yield models Over the years, great effort has been expended on debris risk model input parameters to improve, understand, clarify, or justify the values used for assessing launch hazards. One of the most controversial issues remaining is the estimation of solid propellant impact yield. There are many models and methods available to estimate the explosive yield produced upon impact of the solid propellant at variable velocities and configurations, many of which have been used for Titan SRMUs, Atlas SRBs, and Delta II GEMs. Solid Rocket Motors have consistently driven explosive debris risks to the general public and personnel for launch area and downrange overflight accident scenarios. Solid propellant also tends to drive risks to payloads requiring SNPSs (e.g., Cassini, MERs, PNH, MSL). The Space Shuttle SRB in particular provides the highest launch area explosive debris risk value of all the other launch vehicles for one reason: the aft segment does not include a linear shaped charge (LSC) to split the case as is done for the forward segments. This leaves a contained mass of approximately 260,000 lb of 1.3 HTPB
6.3 Launch Abort Environments Affecting SNPSs
FIGURE 6.44 STAR-48 destruct test at White Sands Missile Center.
propellant descending to impact should destruct action be necessary during ascent. The short lived Ares replacement vehicle of NASA’s Constellation Program provided for extension of the standard LSC to incorporate the aft segment. However, how do I know my yield model is accurate? Considerable effort has been expended culling available historical data to validate our models, but the truth remains: there is very little controlled testing data. The only true test I know of was the 1964 Titan IIIC segment (~82K lbm 1.3 HTPB) strapped to a sled and slapped end-on into a concrete wall at 667 ft/sec. Models tend to over-predict the yield calculated from that resultant test data. Other yield “data” tends to come from investigations of real incidents when crater dimensions are used to back-calculate appropriate yield. Some small scale data exists from critical diameter testing but it is questionable if that data adequately relates to fallback yield. Nonetheless, all available data has been previously scrutinized, weighed, and factored into the methods of yield prediction. Why not get more data? Since 1964, when the sled test was accomplished, constraints to large scale high explosive testing have increased considerably (environmental limitations, high cost). A STAR-48 full scale destruct test, accomplished at WSMC for the MER nuclear risk assessment was an excellent demonstration of the difficulties and the rewards of this type of challenging data acquisition program (Figure 6.44).
Solid propellant computer modeling Special computer models have been developed to predict the blast environment provided when a solid propellant chunk impacts the ground. The PISCES-2DELK continuum mechanics model, developed by FOILS Engineering, is a hydrodynamics code which was used to envision the impact of a General Purpose Heat Source (GPHS) on sand or concrete followed by a direct impact or near-field impact of a large propellant chunk (Mukunka, 1996). It encompasses the cratering process, the response of the propellant to it, the subsequent fracturing and burning of the
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FIGURE 6.45 GPHS impact modeling via PISCES-2 DELK at 0 seconds.
propellant, and finally derivation of the flow conditions (pressure, density, etc.) radially away from the impact point. Figures 6.45e6.48 depict predictions of the response and deformation of the GPHS to these events.
Orbital Re-Entry Failures that can occur once the spacecraft has achieved orbital velocity can end the mission abruptly or over a longer delay ending in re-entry through Earth’s atmosphere, heating, and eventual breakup and impact.
FIGURE 6.46 GPHS impacting modeling via PISCES-2 DELK at 3.0 ms.
6.4 Containment Design
FIGURE 6.47 GPHS impacting modeling via PISCES-2 DELK at 10.0 ms.
FIGURE 6.48 GPHS impact modeling via PISCES-2 DELK.
6.4 Containment Design The safety issues for radioisotope heat sources and fission reactor based heat generation are fundamentally different and must be considered separately. For radioisotope
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systems, the fuel material is hazardous at launch, and containment of the fuel material, to the greatest degree possible, must be achieved for all accident environments. Radioisotope systems also produce heat continuously and this energy must be managed though all mission phases. For fission reactor systems, the nuclear fuel is relatively nonhazardous at launch and the hazard increases with reactor operating time as the fission product inventory builds up. For reactor systems, the safety approach focuses on avoiding inadvertent criticality ensuring that any fission product inventory is adequately decayed before accidental re-entry. Reactor systems produce heat on demand, therefore simplifying thermal management though the various missions phases. However, reactor systems are inherently more complex devices than radioisotope systems, leading to a wide variety of failure modes that must be considered. The safety approach for radioisotope systems has been well developed through the long history of the use in space applications. Radioisotope systems are designed to contain the fuel though the variety of insults than can occur in launch vehicle accidents and inadvertent re-entry.
Small Heater Containment Design A small application of radioisotope material in spacecraft systems involves placing Light Weight Radioisotope Heater Units (LWRHUs) in locations throughout the space vehicle that can benefit from a constant source of heat to maintain thermal control. The Mars Pathfinder Sojourner utilized three of these to keep its Warm
FIGURE 6.49 LWRHU layers.
6.4 Containment Design
Electronics Box (WEB) above critical temperatures during the cold Martian night. The Cassini spacecraft has 117 LWRHUs located in various places for thermal management (Furlong & Wahlquist, 1999). LWRHUs consist of 2.68 grams of PuO2 in a fuel pellet that produce 1 watt of heat output. Each fuel pellet is encapsulated in a platinum rhodium cladding and encased in a multilayer graphite containment to protect the pellet in the event of an accident (Figure 6.49).
General Purpose Heat Source Module The workhorse of the Radioisotope Thermoelectric Generator (RTG) is the General Purpose Heat Source (GPHS) (Figure 6.50).
RTG/MMRTG structure Containment is at the GPHS level, but is not as necessary at the RTG level. This is actually a benefit since in most accident cases the RTG can be broken apart leaving the individual GPHS modules to fall back to land with impact. Since they are designed for the re-entry heating, as individual “boxes” they are lighter and their areal density can decrease the impact speed (Figure 6.51).
Space Reactor System Containment Design The safety approach for reactor based systems is less developed due to the extremely limited application of this technology to date and the fact that no programs in recent history have gone to flight. Nonetheless, a number of programs in the past 20 years
Cap Aeroshell Cap
Fueled Floating Clad Membrane Fuel Pellet
Graphite Impact Shell
Lock Member
Carbon Bonded Carbon Fiber Sleeve Aeroshell Lock Screw Individual GPHS Module
FIGURE 6.50 General purpose heat source stack.
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FIGURE 6.51 Radioisotope Thermoelectric Generator (RTG).
have proposed the use of space reactor systems (Angelo, 1985) and the safety of such systems had been studied extensively (Ponomarev-Stepnoi, Talyzin & Usov, 2000). At launch, the uranium fuel of a reactor system is relatively non-hazardous as compared to the plutonium fuel for radioisotope systems. As the reactor operates, the fission product inventory builds up increasing the hazard. The safety of reactor systems, therefore, depends primarily on preventing operation in the biosphere and ensuring that there is sufficient decay time after reactor operation before it can return to the biosphere. The first element of ensuring safe operation of space reactor systems is the prevention of inadvertent criticality. Inadvertent criticality must be prevented for both normal credible accident conditions. The primary accident scenarios of concern for preventing criticality are compaction and water immersion. Compaction is a concern in impact accidents where the nuclear fuel could potentially be arranged to a critical configuration. Water immersion is a concern as flooding of the reactor core and reflection provided by water immersion can also lead to a critical configuration. In some cases, the reactor can be designed to be inherently insensitive to compaction and water immersion. If such design is not feasible, then active safety measure such as removable neutron poisons can be employed to address inadvertent criticality concerns. Additional operational measures are employed to prevent a radiologically hot reactor from entering the biosphere. Prelaunch, the reactor can only be operated for low power testing such that negligible radioactivity is produced. After launch, the reactor can only be operated when the system reaches a “sufficiently high orbit.” A sufficiently high orbit is defined as that which the orbital lifetime is adequate to allow for decay of the fission products before re-entry. For reactor disposal in Earth orbit, such disposal must be in a sufficiently high orbit and re-entry of a radiologically hot reactor is precluded.
6.5 Risk Assessment for Nuclear Missions The use of radioactive materials has enabled humankind to pursue scientific research and gain insights into our universe that would have been impossible or impractical by other means. However, these same materials that have proven
6.5 Risk Assessment for Nuclear Missions
so beneficial in outer space can pose a significant hazard to human health and the environment if an accident causes them to be dispersed. For this reason, a rigorous risk assessment process is performed for all launches involving radioactive material. The purpose of a risk assessment is twofold: first, during the design of the mission and the spacecraft, risk assessment enables mission decision makers to identify, assess, and manage possible safety concerns (and, in this context, especially radiological safety concerns) related to the mission. Then, once the mission and spacecraft design have been completed, the results from the risk assessment are presented to the administrative and political decision makers who participate in the launch approval process as evidence to document the safety of the mission. One way to view the risk assessment is as the means of adjudicating the competition between the extreme environments that might be encountered during a launch accident and the robust containment structures that are designed to prevent the radioactive materials from reaching the environment. The risk assessment examines the scenarios that have the potential to cause a release of radioactive material and estimates the likelihood that a release will occur, the characteristics of the release that might be expected, and the effects that the release might have on the public and the environment. Thus, a risk assessment is a broad and far-reaching analysis that is performed by a team with many different skills and tools. So what is risk? Risk can be viewed as the potential for realization of unwanted or adverse consequences. Risk estimates are based on the likelihood and consequences of an adverse event, considering uncertainty. Thus, a risk analyst must provide a rigorous and complete understanding of which adverse consequences can occur, the mechanisms by which such consequences can be produced (scenarios), the likelihood of occurrence for each scenario, the specific magnitude of consequences that would be expected should the consequence occur, and the degree of certainty involved in each of these estimates. For activities related to the use of radioactive materials in outer space, the principal adverse consequences of an inadvertent release of radioactive materials are the potential for environmental contamination and human health effects (especially latent cancer fatalities). It should be noted that spaceflight presents many risks other than those involved with the release of radioactive material. Examples include the use of toxic and flammable materials, asphyxiation in confined spaces, and other common industrial hazards. These risks must also be identified and managed to ensure the overall safety of persons and the environment. However, the focus of this section is specific to the special risks associated with the use of radioactive materials. The basic process for assessing risks for the launch of radioactive materials requires the analyst to: • •
Analyze the launch vehicle and mission to determine the frequency of events that might threaten the radioactive material. Determine the distribution of the potentially threatening environments in terms of their conditional likelihood, magnitude, time of occurrence, and spatial characteristics.
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• • • •
Determine the expected response of the structures intended to contain and isolate the radioactive material to each of the potential threatening environments. Determine the amount and characteristics of the radioactive release, if any. Determine how any released material will be transported through the environment cause contamination or to cause human exposure. Estimate the health effects of these potential radioactive exposures.
To be successful this process must involve the collaboration of experts from a wide variety of technical disciplines working together to understand complex nonlinear phenomena and provide straightforward insights for decision makers.
Characterizing the Insult Environment Spacecraft designers understand the importance of preventing the release of radioactive materials to the environment, so they design structures into the spacecraft to contain and isolate the radioactive material. The job of the risk analyst is to understand the conditions under which the radioactive material might escape these structures. The chief mechanism for escape is the introduction of energy into these structures that is sufficient either to breach the structure or to change the material form of the radioactive material inside to enable it to escape through some other pre-existing pathway. An example of the former is a crack in cladding material, while the latter may occur if heat causes the material to vaporize and escape through a vent. The introduction of energy to radioactive material or its protective structures is often called an “insult.” Energy insults can take the form of atmospheric overpressure, explosive shock, mechanical force, impact by fragments, impact with the ground, or thermal energy from burning propellants or the heat of re-entry. Historically, other sources of energy such as electric power or chemical reaction have usually not been present in sufficient magnitude to challenge the containment of the radioactive material except to the degree that they contribute to the creation of the other energy sources. The potential sources and magnitudes of these insult environments have been described earlier in this chapter (see section on “Launch Abort Environments Affecting SNPSs”). The risk analyst looks for mechanisms that can apply sufficient energy to damage the radioactive material or its protective structures and cause the material to be released to the environment. Understanding the threshold for “sufficient” energy for any individual insult type is extraordinarily challenging. For example, extensive testing and computational simulation programs may be required to characterize the behavior of the protective structures, with the objective of providing risk analysts with the most accurate possible understanding of how these materials would perform in a real accident should it occur. Limitations on testing and computational resources mean that such studies often involve the application of only a single insult type in any experiment or simulation. However, a real accident involving a launch vehicle or spacecraft produces an extremely chaotic environment in which multiple insults of various types might be created in rapid succession. The superposition of results from
6.5 Risk Assessment for Nuclear Missions
separate effects tests into a comprehensive picture of material behavior in this chaotic environment remains a significant challenge for risk analysts. In addition to understanding the type and magnitude of the insult environments that may cause the release of radioactive material from its protective structures, the risk analyst must also understand how those insult environments might come to exist. The analyst must use the list of insult environments which have the potential to result in radioactive releases when examining all aspects of the spacecraft, the launch vehicle, and the mission. It may be possible for accidents to create such environments not only during the launch itself, but also during the preparation of the spacecraft and launch vehicle for the launch event, and during any possible re-entry of the spacecraft into the Earth’s atmosphere. After many years of studying launch vehicle and spacecraft accidents that have occurred, and studying postulated accident scenarios from risk assessment studies for all phases of previous missions, a rich historical database has been developed to relate which insult types should be studied in which mission phases. This list serves as a generic starting point for consideration of insult environments because it incorporates information from a broad cross section of past experience. The potential for a launch vehicle to explode, with its resultant overpressure, possible explosive shock, fragment field, ground impact, and thermal insults must be considered for any mission phase when the radioactive material is installed in a fueled launch vehicle, but the characteristics of these insults (e.g., magnitude and spatial distribution) may vary dramatically depending on when the accident occurs and which events or malfunctions trigger the accident. The risk analyst must not stop, however, with the list of generic insult environments and their causes. Each space mission is a first-ever event using one-of-a-kind hardware that has been designed for a special purpose but never before completely operated as an integrated unit. Therefore, it is possible that some of the generic insult environments may not apply in the “usual” mission phases, and it is likely that unique new accident conditions may exist with the potential to cause insults to the radioactive material or its protective structures. For example, the Pluto New Horizons mission, launched in 2006, marked the first time a U.S. mission used a large solid rocket upper stage in close proximity to an RTG. This mission feature meant that previous assumptions regarding the mission time over which accidents might produce fragments of solid rocket propellant and solid rocket motor cases with the opportunity to impact the RTG had to be changed. In addition, the proximity of the solid rocket motor to the RTG meant that the impact of large high velocity propellant fragments on the RTG was far more likely for this mission than for any previous mission. Thus, the risk assessment team paid special attention to this unusual source of insults in order to ensure that it was appropriately captured in the mission risk profile. Another example of a mission-specific insult environment occurs when radioactive material is contained deep inside of a spacecraft. When radioactive material such as an RTG is highly exposed, the overpressure produced by the deflagration of liquid fuels may tear the RTG away from the spacecraft and allow it to fly
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free. In such situations, the RTG may land well clear of any burning launch vehicle debris, thus reducing the potential for mechanical insults such as the impact of launch vehicle debris or thermal insults from burning solid rocket propellant. When the radioactive material is contained deep inside of the spacecraft, there is significantly more potential for these insults, and also the potential for insults from the structure of the spacecraft itself. These examples demonstrate how the consideration of mission-specific insult environments is critical to be accurate depiction of mission risk.
Accident Likelihood In order to assess mission risks, the risk analyst must understand the insult environment “hazard curve” e that is, the relative likelihood with which insults of each type and magnitude will occur. To accomplish this, a mission accident assessment brings together information about the likelihood of various observed and postulated accident-initiating events with mission characteristics and physical response data so that the risk analyst can understand the circumstances under which each insult type and magnitude will occur. The process begins by identifying scenarios that could be important to mission risk. In general, analysts look for potential situations that might impart kinetic, thermal, or chemical energy to the radioactive material. For example, prior to launch, accidents during vehicle assembly might result in kinetic energy insults from accidental launch vehicle collapse or dropping materials from great heights, or thermal insults form accidental fires or failure to adequately cool RTGs. During launch, vehicle malfunctions may cause the launch vehicle to explode or crash, or they may cause the launch vehicle to fly off-course requiring actuation of the vehicle destruct systems. The vehicle destruct systems may also malfunction directly to cause the launch vehicle either to be destroyed at an unintended time or to fail to be destroyed when destruction is needed. These chaotic environments can result in many different sources of kinetic, thermal and chemical insults to the radioactive materials. Different types of insult environments occur once the spacecraft reaches suborbital or orbital altitudes, as the insult of re-entry heating becomes possible but reduced inventories of rocket fuels reduce the likelihood and severity of explosive insults. The situations under which these insults might occur are identified using systematic probabilistic risk assessment tools such as failure modes and effects analysis, hazards and operability analysis, and fault tree analysis. Tools like event sequence diagrams and event tree analysis are used to systematically explore how each type of accident might progress to specific insult-producing outcomes, or conversely, be mitigated before insults are produced. For example, if a scenario is identified that could produce a thermal insult to an RTG due to failure of cooling systems, the event sequence diagram or event tree would incorporate options for re-establishing cooling via alternative means or removing the RTG from the spacecraft before adverse temperatures could be reached.
6.5 Risk Assessment for Nuclear Missions
The likelihood of each accident initiating event is highly time-dependent not only between mission phases but even within a mission phase. For example, the likelihood of catastrophic failure for solid- and liquid-fueled rocket engines is often much higher during the first several seconds after ignition than at any other time during the mission. These failures, in turn, occur at relatively low altitudes near the launch pad where the falling materials may impact hard surfaces such as concrete and steel structures. Thus, the analyst must carefully account for the time profile of each failure type in order to produce an accurate assessment of mission risks. The frequency of specific accident initiating events is estimated by considering a wide variety of statistical and non-statistical evidence. Detailed examination of past space flight failures provides evidence regarding the overall frequency of such accidents. However, the historical database covers a wide variety of dissimilar launch vehicles and spacecraft, and some of these historical failures may not be relevant to the mission being assessed. Evidence also comes from the probabilistic risk and reliability studies that are performed for the launch vehicle, the spacecraft, and the ground support systems, as well as component and subsystem test data. In many cases, however, these studies and measured data may not be fully applicable to the mission being assessed. For example, a tank failure pressure may have been measured assuming a full load of fuel while the profile for the mission being assessed calls for this tank to be only 70% full. Thus, it is an inexact science to estimate the likelihood of each accident initiating event for each mission’s specific one-of-a-kind characteristics from the sea of partially-applicable data. Analysts accomplish this task using statistical techniques when possible (especially Bayesian analyses), but often must make use of expert judgment techniques because of the disparate nature of the underlying data. A thorough risk mission assessment examines a large number of accidents derived from a large number of possible initiating events. However, the insult environments for the radioactive material are often similar for accidents from many different initiating events, so analysts may group the accident scenarios so that accidents that produce similar event sequences and insult combinations can be analyzed together for their radioactive release potential. The result of this analysis can be thought of as a hazard curve for each insult type. It includes groups of accidents that are expected to produce similar combinations of insults, and a probability distribution for each group of accidents as a function of mission time. It also includes statistical distributions to represent the intensity and location of each insult type (e.g., fragment size, velocity and angular distribution). For U.S. launches, this information is documented in a Launch Vehicle Databook (Rumerman, 1999) that is unique to each mission.
Radioactive Release Characterization A group of accidents often presents the radioactive material and its protective structure with an environment that includes multiple sequential or even concurrent insults with physics that are coupled and nonlinear. The potential for radioactive release
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must be estimated for these complex and chaotic environments on the basis of separate effects data from experimental and computational sources. Where possible, limited sets of combined environment and multiple insult evaluations are performed to help analysts understand how best to use the data from the more numerous and complete separate effects sources. Otherwise, the separate effects data are applied to the material sequentially to predict the effects on the radioactive material and its protective structures. The objective of the analysis is to evaluate whether a particular sequence of insults will cause radioactive material to be released to the environment, and if so, to predict the characteristics of the release that are important to its potential transport through the environment and ability to cause human health effects. The key characteristics that are estimated are the mass of radioactive material released, its particle size distribution, its chemical composition, and the thermal environment of the release. The mass of radioactive material released depends on the performance of both the protective structure and the material itself. Breaches in the protective structures enable material to escape to the environment, and the released mass depends strongly on the size of the breach. But the released mass also depends on the physical form of the radioactive material. For example, material that has been crushed into a powder by a kinetic energy insult can escape through a breach more easily than material that is in the form of a pellet, extruded metal, or large chunks. Also, liquid or vaporized material can escape through pathways that are inaccessible to solids. The risk analysis makes use of models to predict how much material escapes under specific insult conditions, and these models are tied to both experimental and computational results. The scatter in the experimental data for such models can be great, and it is incumbent upon the risk analyst to understand and account for this source of uncertainty. Two other important characteristics of the released radioactive material are its particle size distribution and its chemical form. The particle size distribution is important because small particles (e.g., less than 10 microns in diameter) are easily transported by wind and are readily inhaled by humans, while large particles (e.g., greater than 100 microns in diameter) fall out of the atmosphere quickly and are not easily inhaled. The chemical form is also important for predicting the biological uptake of certain radioactive materials. The size distribution and chemical form of the material can be significantly altered by the insult environment. For example, while the initial form of the material may be a ceramic oxide pellet with very benign distribution and uptake characteristics, a kinetic insult may crush the ceramic to produce very fine particles, and a severe fire environment may vaporize the material and change its chemical form significantly. A final important characteristic of the radioactive release is its thermal characteristics. Materials that are released in a fire environment may be lofted high into the atmosphere and, thus, transported far greater distances in the environment than comparable releases in an ambient temperature environment. Each of these release characteristics can have a significant effect on the behavior of the material as it is transported through the environment and as it affects exposed
6.5 Risk Assessment for Nuclear Missions
humans. Since the insult environment is complex and nonlinear, and since there is significant statistical uncertainty in the release models, the radioactive release problem is not a candidate for closed form solution or even simple approximation. Instead, release predictions can be computed using a tool like a Monte Carlo discrete event simulation tool that explicitly considers the known uncertainties to estimate releases. The tool might select the mission elapsed time at which the accident occurs and position the spacecraft at the appropriate point. It might follow the spacecraft and radioactive material as it flies off-course and/or falls to the ground. It might generate a statistical distribution of fragments for any explosive environment and determine which fragments impact the radioactive material e either in the air of on the ground. It might also apply the appropriate overpressure insult and thermal environment to the material and compute the characteristics of any radioactive releases that may occur. This process can be repeated for a large number of Monte Carlo trials until the distribution of possible releases is sufficiently characterized. This same process is repeated for each group of accidents. This computational method produces far more possible releases than can be assessed in the radioactive material transport or health effects estimation analyses. A frequent characteristic of the set of computed releases is that an overwhelming majority of the releases are very small with characteristics that would be expected to cause minimal land contamination or human health effects. Large releases very rare, but since they may be orders of magnitude larger than typical releases, they can have a profound effect on statistics such as the mean release and mean consequences. Furthermore, these very rare trials often involve a combination of insults that are at or beyond the bounds of the experimental dataset, and are, thus, relatively uncertain. The sensitivity of the mean to computational artifacts and uncertain data is a fact that is difficult to effectively communicate to decision makers.
Radioactive Material Transport The released radioactive material is transported through the environment and may be deposited far from the original accident site. The objective of this part of the mission risk assessment is to understand how each release of radioactive material is transported through the environment, to estimate the exposure of humans to the radioactive material, and to predict the area and magnitude of land contamination. These results depend not only on the characteristics of the release described previously, but also on the weather conditions into which the release occurred and the population distribution around the accident site. The weather conditions into which the release occurs are critical. A release on a day when winds would blow the radioactive plume out to sea will produce minimal land contamination and human health effects. However, the same release on a day with winds blowing toward a population centre may produce very different results. The principal tools used in the radioactive material transport analysis are plume rise and atmospheric transport models that explicitly account for population. The analysis is
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based on the assumption that the distribution of possible weather conditions at the time of the potential accident can be appropriately estimated from the weather conditions from recent years for dates near the potential accident date (as related to the projected launch date). Typical analyses use weather data from the previous several years for the dates in question, resulting in 150e250 days of weather measured conditions as potential conditions into which the radioactive material might be released. Additional data are required to support the plume rise and transport models, such as model parameters related to the degree to which a plume spreads or the speed with which particles of a particular size are deposited from the atmosphere on the ground. There is significant uncertainty regarding which values of particular model parameters produce the best agreement with experimentally observed behavior, so the results of this analysis must reflect these uncertainties for decision makers. A complete analysis of radioactive material transport would evaluate all radioactive releases into all weather conditions using a variety of model parameters. Such an analysis is not possible within current computational capabilities, so a statistical subset radioactive releases and weather conditions must be analyzed. Methods for generating this statistical subset have included pseudo-Monte Carlo sampling and importance sampling. Since the data contain a few rare but large releases and a few weather trials that yield unusually large consequences, the potential for random combination of these rare events in a Monte Carlo analysis can make the mean and other statistics very sensitive to the presence or absence of specific random data pairings. Overcoming these sampling artifacts will require a significant increase in computational power to be applied to these analyses.
Health Effects Estimation A key metric in the assessment of risks for space missions involving radioactive material is the potential for released material to cause human health effects such as latent cancer or early fatality. To compute this metric, risk analysts use methods sanctioned by the International Commission on Radiological Protection (ICRP) (ICRP, 1999, 2005). These methods consider human exposure to radiation on an organ-by-organ basis for exposures related to radioactive materials that get into the body through pathways such as ingestion or inhalation, as well as exposure from radioactive materials outside of the body through pathways such as the “shine” from radioactive materials in the air or on the ground. The methods also consider mechanisms by which materials can be changed from one exposure pathway to another, such as when the material is dissolved in water or when material is resuspended in air after having been previously deposited on the ground. Since specific radionuclides are preferentially deposited in particular organs, the analysis is very material-specific. As stated previously, the vast majority of radioactive releases computed in typical mission risk assessments are very small. This leads to radiological doses to humans that are very small, and are increasingly small as one moves to greater distances from the accident site. However, these greater distances also hold the possibility for a much larger population to be exposed to these very small
6.5 Risk Assessment for Nuclear Missions
radiological doses e often an exceedingly small percentage of normal background radiation. The ICRP method holds that all nonzero radiological doses hold the potential to cause latent cancer, so these very small computable (but likely not measurable) radiological doses are applied to very large numbers of people to produce health effects, and these health effects from extremely small doses often constitute a large fraction of the overall mission risk. An alternative to the ICRP model holds that a radiation dose threshold exists below which the potential for latent cancer is negligible (a de minimis dose). Thus, persons who receive extremely low doses of radiation are not counted in the mission health effects. However, there is no scientific consensus as to whether a de minimis threshold exists and, if it exists, what its value should be. Furthermore, political decision makers are not of the same mind on this subject. For this reason, risk analysts may present results based on numerous hypothetical de minimis threshold values so that one can understand how this technical uncertainty affects the overall mission risk results.
Risk Integration The analysis method described in this paper leads to the definition of several possible accident scenarios, each with many potential radioactive releases, each of which could occur under a variety of weather conditions leading to different land contamination and human exposure potentials, each of which might lead to different of health effect estimates (depending on the value of the de minimis dose selected). These results are then integrated and summarized to provide a predictive risk profile. To accomplish this, the population of accidents, releases, and consequences is sorted and integrated into a complementary cumulative distribution function (CCDF). An example CCDF is shown in Figure 6.52. A point on a CCDF is defined as an ordered pair consisting of a probability (the ordinate) and a consequence value (the abscissa), and is interpreted as the probability that the consequences will exceed this value (thus, it is sometimes referred to as an “exceedence frequency” curve). A CCDF can be generated from a data set as the complement of the integral of the probability density function to a specific point. There are several benefits to displaying risk as a CCDF. First, it visually reinforces the fact that large consequences are very unlikely to occur. Also, when comparing among alternatives, comparing the CCDFs quickly shows differences between the risk profiles that cannot be seen by comparing single values such as the mean risk (e.g., one alternative has a lower likelihood of large consequences but a higher likelihood of accidents involving a small number of consequences). Finally, a CCDF directly answers typical risk-related questions, such as, “What is the probability that at least one person will die as a result of this mission?” and “How bad could a one-in-a-million accident be for this mission?” The answers to these questions are simply points on the CCDF. The integrated risk results may be composed of several CCDFs (e.g., one each for health effects and land area contamination, or one to represent the risks
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Mission Risk for the Pluto New Horizons Mission 1.E-02
Probability of Exceedence
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1.E-03
1.E-04
1.E-05
1.E-06 1.E-05
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FIGURE 6.52 Risk is represented as a complementary cumulative distribution function.
associated with each mission phase, or one for each analyzed de minimis threshold). In addition, mean risk values (computed as loss expectation values, or the integration of probability times consequence) and other single-value estimates (e.g., the probability of at least one heath effect) can be computed and presented to decision makers if requested, although the accompanying discussion must place these in their proper perspective.
6.6 International Protocols and U.S. Environmental Review Several international treaties, agreements, conventions, and principles adopted to date have some relevance to the use of space nuclear systems. Similarly, the environmental process formally established in the U.S. is applicable to the launch and use of space nuclear systems. These sections address these international protocols and the U.S. environmental process as they relate to the use of space nuclear systems.
International Protocols, Principles and Guidelines Applicable to the Launch and Use of Space Nuclear Systems The international community, primarily through the United Nations (UN), has taken an increasingly active role in establishing agreements on safety and environmental issues relating to space activities as well as on nuclear power considerations (Sholtis, Marshall, Bennett, Brown, Usov & Dawson, 2008). During the 1960s through the
6.6 International Protocols and U.S. Environmental Review
1980s, five international treaties were established to address space safety and environmental issues; namely: 1. Treaty on Principles Governing the Activities of States in the Exploration and Use of Outer Space, Including the Moon and Other Celestial Bodies (known as the Outer Space Treaty), adopted October 10, 1967; 2. Agreement on the Rescue and Return of Astronauts and the Return of Objects Launched into Outer Space (known as the Rescue Agreement), adopted December 3, 1968; 3. Convention on International Liability for Damage Caused by Space Objects (known as the Liability Convention), adopted October 9, 1973; 4. Convention on Registration of Objects Launched into Outer Space (known as the Registration Convention), adopted September 15, 1976; and 5. Agreement Governing the Activities of States on the Moon and Other Celestial Bodies (known as the Moon Treaty), which was not signed by the U.S., adopted July 11, 1984. The first four treaties govern activities in space, whether nuclear or non-nuclear, while the Moon Treaty does mention emplacement of radioactive materials on the Moon. In the late 1980s, in the wake of the Chernobyl nuclear power reactor accident in 1986, the UN established two conventions related to nuclear power; namely: 1. Convention on Early Notification of a Nuclear Accident, adopted October 27, 1987; and 2. Convention on Assistance in the Case of a Nuclear Accident or Radiological Emergency, adopted February 26, 1987. Space nuclear systems have received international attention ever since the unplanned re-entry of the former Soviet Union’s Cosmos 954, Radar Ocean Reconnaissance Satellite (RORSAT), with an onboard radioactively hot space reactor, into the Earth’s atmosphere and subsequent impact of the reactor debris onto the Great Slave Lake area in the Northwest Territories of Canada on January 24, 1978. As a direct result of this accident, in 1978 the UN Committee on the Peaceful Uses of Outer Space (COPUOS) established a Working Group on Nuclear Power Sources (NPS) in Space principally under its Scientific and Technical Subcommittee, but also under its Legal Subcommittee (Bennett, Sholtis & Rashkow, 1987). This Working Group on NPS in Space, which has convened periodically since the late 1970s, worked to formulate a Preamble and eleven Principles related to the safe use of NPS in space. That Preamble and eleven Principles were adopted by the UN General Assembly in June 1992 (UN Legal Subcommittee, 1992). A summary of the Preamble and Principles follows:
Preamble The Preamble recognizes that for some missions nuclear power systems are essential and affirms that the Principles only apply to nuclear power systems “. devoted to
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generation of electric power onboard space objects for non-propulsive purposes, which have characteristics generally comparable to those of systems used and missions performed at the time of the adoption of the Principles .” In other words, the Principles do not apply to nuclear propulsion systems or to new space nuclear systems. The Preamble also recognizes “. that this set of Principles will require future revision in view of emerging nuclear-power applications and of evolving international recommendations on radiological protection”.
Principle 1 Applicability of international law e basically states that the use of nuclear power systems will be carried out in accordance with international law.
Principle 2 Use of terms e defines a number of terms, in particular “. the terms ‘foreseeable’ and ‘all possible’ describe a class of events or circumstances whose overall probability of occurrence is such that it is considered to encompass only credible possibilities for purposes of safety analysis.” In addition, the definition of the term “general concept of defense-in-depth” allows flexibility in achieving this goal by allowing consideration of “. the use of design features and mission operations in place of or in addition to active systems, to prevent or mitigate the consequences of system malfunctions. Redundant safety systems are not necessarily required for each individual component to achieve this purpose. Given the special requirements of space use and of varied missions, no particular set of systems or features can be specified as essential to achieve this objective”.
Principle 3 Guidelines and criteria for safe use e this principle sets forth general goals for radiation protection and nuclear safety followed by specific design safety criteria for nuclear reactors and for radioisotope generators.
Principle 4 Safety assessment e requires a “thorough and comprehensive” safety assessment which is to be made publicly available prior to each launch.
Principle 5 Notification of re-entry e requires a timely notification of the re-entry of radioactive materials into the Earth’s atmosphere and provides a format for such notification.
Principle 6 Consultations e requires States providing information under Principle 5 to respond promptly to requests for further information or consultations sought by other States.
Principle 7 Assistance to States e requires States with tracking capabilities to provide information to the Secretary-General of the United Nations and to any concerned
6.6 International Protocols and U.S. Environmental Review
State, and requires the launching State to promptly offer assistance. After re-entry, other States and international organizations with relevant technical capabilities should also provide assistance to the extent possible when requested by an affected State.
Principle 8 Responsibility e States shall bear international responsibility for their use of space nuclear power systems.
Principle 9 Liability and compensation e holds the launching State and the State procuring such a launch internationally liable for any damage, including restoration “. to the condition which would have existed if the damage had not occurred”. Compensation includes “. reimbursement of the duly substantiated expenses for search, recovery and clean-up operations, including expenses for assistance received from third parties.”
Principle 10 Settlement of disputes e disputes “. shall be resolved through negotiations or other established procedures for the peaceful settlement of disputes, in accordance with the Charter of the United Nations.”
Principle 11 Review and revision e requires that “These Principles shall be reopened for revision by the Committee on the Peaceful Uses of Outer Space no later than two years after their adoption.” Principle 3 is perhaps the most important and most controversial of these Principles. During the adoption of the Principles, the U.S. delegation formally expressed reservations about the technical validity of these Principles to the United Nations. For example, in Section 1.3 of Principle 3, dose limits were established for accidents. Although the dose limits did not apply to “low probability accidents with potentially serious consequences” the U.S. delegation pointed out that Principle 3 should address risk (probability of an adverse consequence) rather than numerical dose limits (Lange, 1991a). In particular, the U.S. delegation made the point that “. this modification, by taking into account the probabilistic concept of risk, which is a central feature of a thorough safety assessment, relates the recommendation directly to the well-proven [space nuclear power system] practices of the United States” (Lange, 1991b). “In November 1990 the International Commission on Radiological Protection published new recommendations in the form of ICRP-60, which supersede the approach taken in Principle 3 when it was developed earlier that year” (Smith, 1992). The International Atomic Energy Agency (IAEA) independently supported the U.S. position by stating that “The sole use of the individual-related dose limits, rather than the complete ICRP system of radiation protection (including sourcerelated constraint), is, in the Agency’s view, inappropriate and does not conform
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with the aims of the ICRP recommendations. Secondly, as the ICRP has recently issued new recommendations on dose limitation . It might, therefore, be problematic to issue guidelines and criteria of safe use of [nuclear power systems] in outer space that would be outdated from their inception” (IAEA, 1991). The United States also proposed clarifying language in 1991 with the statement that “We believe that this clarification removes any doubts as to the intent behind the application of the term ‘defense-in-depth’. As was clear at the time that the Legal Subcommittee reached consensus on this principle, the Subcommittee did not intend to apply the terrestrial standards as such to space systems.” In 1991 the U.S. delegation proposed changes to the wording of Section 3.2 (relating to re-entry of radioisotope sources) “. to take into account the fact that the probability of accidental re-entry from a hyperbolic or highly elliptical orbit can be reduced to a very low value by mission design and operations” and to recognize “. the fact that the practical design objective for RTG containment systems is localization rather than zero release under all circumstances, and that there are practical limits from a costversus-risk standpoint on ‘complete’ clearing of radioactivity by a recovery operation” (Lange, 1991b). At a meeting of the Special Political Committee on October 28, 1992, the U.S. representative stated “The United States did not block the consensus recommendation of the Committee to forward the principles to the General Assembly, nor will the United States oppose their adoption here. On some points, however, it remains our view that the principles related to safe use of nuclear power sources in outer space do not yet contain the clarity and technical validity appropriate to guide safe use of nuclear power sources in outer space. The United States has an approach on these points which it considers to be technically clearer and more valid and has a history of demonstrated safe and successful application of nuclear power sources. We will continue to apply that approach” (Hodgkins, 1992). Although the UN Principles are non-binding, the United States and other nations are working together to establish generally acceptable international safety and environmental guidance relating to space nuclear power. International cooperation in space enterprises and the voice of the international community is expected to increase during the twenty-first century (Sholtis, Marshall, Bennett, Brown, Usov and Dawson, 2008).
The U.S. Environmental Process for the Launch and Use of Space Nuclear Systems The National Environmental Policy Act, or NEPA, was enacted within the U.S. in 1970 (NEPA, 1970 et seq). It requires all U.S. federal agencies to formally consider and document the consideration of planned programs, projects, actions, and activities that could have a significant impact on the quality of the human environment. Planning and pursuing a space mission which would be launched and operated with nuclear materials or systems onboard could potentially trigger the NEPA process (Sholtis et al., 2008).
6.6 International Protocols and U.S. Environmental Review
NEPA implementing guidance and procedures provide detailed guidelines for programs and projects to ensure NEPA requirements are met. The NEPA implementing guidance and procedures require that planned federal programs, projects, actions, and activities with potentially significant impacts complete an environmental analysis sufficient to support and document a formal decision on the matter. Moreover, NEPA requires that this environmental analysis be subjected to public comment, and that any decision in the matter must be made and documented after public comment and responses to those comments have been addressed. Lastly, NEPA requires that a formal, documented decision in the matter must be completed before any substantial expenditure of funds and any irrevocable environmental impacts occur. This required NEPA environmental analysis can be in one of two forms: (1) An Environmental Assessment (EA); or (2) an Environmental Impact Statement (EIS). An EIS involves much more in-depth analyses than an EA, but both must be prepared in Draft form, for public comment, and subsequently in Final form, where all public comments are explicitly addressed. In summary, the EA is a concise public document that provides sufficient information and analyses to determine if a more involved EIS needs to be prepared. Figure 6.53 provides an overview of the NEPA compliance flow process for a planned action involving nuclear materials and shows the trigger points for development of an EIS as opposed to an EA. If, after a Draft EA is completed, subjected to public comment, and updated as a Final EA, it is determined that there are no significant environmental impacts, then a Finding Of No Significant Impact (FONSI) is prepared and published in the Federal Register; NEPA compliance is complete at this point. If, on the other hand, a determination is made that significant environmental impacts do potentially exist, then an EIS is prepared e first as a Draft EIS, and subsequently after public comment, as a Final EIS. An EIS can be done without an EA preceding it if it is reasonably certain that an EIS will be required. See Figure 6.53. An EIS is a detailed and rigorous document, which provides a full discussion of potentially significant environmental impacts, based on the best information available at that time. It presents a discussion of the purpose of the proposed action and details the need for the action. The EIS provides a general understanding of how the action will be implemented, and what potential environmental impacts, both positive and negative, could result. It is also designed to inform decision makers and the public of a reasonable range of alternatives that are compatible with the purpose and need of the action. These alternatives must be compared with the proposed action in terms of their potential environmental impacts. In addition to alternatives, the EIS must also consider the no-action alternative. The EIS undergoes two public comment periods. The first comment period follows a Notice of Intent (NOI) to prepare an EIS which is published in the Federal Register, with a request for comments on the scope of the project. The NOI is also sent to any person or group who has indicated an interest in the project. The second public comment opportunity is a review of the Draft EIS (DEIS). Copies of the DEIS
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START
Sufficient Nuclear Material or Controversy?
NO
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Environmental Impact Statement (EIS)
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END
FIGURE 6.53 NEPA compliance process.
are sent to anyone who responded to the NOI and to any others who have indicated an interest in the project. Additionally, a Notice of Availability (NOA) is published in the Federal Register. Also, a DEIS is provided to any interested or involved federal and state agencies for review. Anyone can provide comments on the DEIS. Comments received are considered by the sponsoring agency, and agency responses are published in the Final EIS (FEIS). At the conclusion of the EIS process a Record of Decision (ROD) is prepared and filed, documenting the agency’s decision on whether to go forward with the action, an alternative, or no action. An EIS for space nuclear missions must provide a detailed analysis of the impact of a normal launch on the air, water, plants, etc. around the launch area. It must also discuss a reasonable range of postulated accidents and their potential impacts. Alternatives considered can include alternate launch vehicles, alternate power sources, and alternate flight trajectories. The launch radiological impact analysis for credible accidents at the launch site, downrange, and inadvertent re-entry from failure to achieve Earth orbit is based on best available information with respect to launch
6.7 Nuclear Mission Launch Approval
vehicle accidents, accident probabilities, the resulting potential accident environments, the nuclear system responses to those environments, and an assessment of the potential health impacts. Postulated re-entry accidents following Earth orbit insertion are also considered. The analysis is a combined effort of the launch vehicle owner, the launching organization (NASA or DOD), the Department of Energy, and other involved government agencies and supporting contractors. Although the analysis performed for an EIS is similar to those performed for the U.S. launch approval process for space nuclear missions, it is separate and distinct from it due to the different timing, purpose, and specifics of each process. For missions like MARS PATHFINDER that use Radioisotope Heater Units (RHUs) with no Radioisotope Power Systems (RPSs), either an EA or an EIS is required. The decision on which will satisfy NEPA compliance requirements is based on a number of factors, most importantly whether it is likely that the process will end with a FONSI. If the EA is completed and it is determined that there is a potential for significant environmental impacts, an EIS will be prepared. As already mentioned, the NEPA process differs significantly in several ways, and is wholly separate from the U.S. launch approval process for space nuclear missions. Because the NEPA process is completed early in the program cycle, the agency may have to use preliminary data and analyses instead of the later, more specific data or more refined analyses used for the launch approval process. In addition, the NEPA process differs in that it involves the public and a large number of federal, state, and local agencies. The U.S. launch approval process for space nuclear systems is not subject to NEPA. Instead, it involves an internal government process, with independent review and evaluation, ultimately leading to a launch decision by the White House; see section “Nuclear Mission Launch Approval” below.
6.7 Nuclear Mission Launch Approval The effects of a release of radioactive materials into the environment from an accident involving a space mission are not limited to quantifiable consequences such as land contamination and potential cancer fatalities. Such a release could also be a serious political matter for the country sponsoring the mission. In order to ensure that both the technical and political risks are appropriately weighed when such missions are authorized, the launch approval process for missions involving the use of radioactive materials should involve both independent scientific reviews of mission risks and participation by political leaders in the launch approval process. The United States formally established such a process in the form of a Presidential Directive in 1977 (PD/NSC-25, 1977). This directive establishes a launch approval process that requires approval from key political leaders following an independent assessment of mission risks. In order to facilitate this process, the directive establishes an Interagency Nuclear Safety Review Panel (INSRP) for each mission involving the use of radioactive
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materials. The INSRP serves two functions: first, to provide expert independent review of all nuclear-related safety and risk assessments performed for the mission, and second, to provide an independent estimate of the nuclear-related risks for use by decision makers in the launch approval process. The INSRP for each mission is an ad hoc panel formed at the request of the agency sponsoring the mission. The panel and its scientific and technical support personnel use the methods described previously in this chapter (see Section “Risk Assessment for Nuclear Missions”) to review information provided by the mission project team such as the mission characteristics, risk assessment methods, physical and structural assessment methods, and the supporting data on which the nuclear risk assessment will be based. A formal but robust technical dialogue between the INSRP and the mission project team ensues to ensure that the INSRP has ample opportunity to probe the methods, data and results of the mission project team, to perform independent analyses, and to ensure that safety questions and issues are resolved. The safety review and risk assessment process results in two independent evaluations of mission risk: a Final Safety Analysis Report from the mission project team, and a Safety Evaluation Report (SER) from the INSRP. Both of these documents are presented to decision makers for use in a multistep mission launch approval process that culminates in the Office of Science and Technology Policy (OSTP) in the Executive Office of the President of the United States. Launch approval authority rests with the President, although he generally authorizes his Science Advisor as the decision making authority for launch approval. The launch approval process generally takes at least 6 months. Both the risk assessments and the launch approval are generally valid only for a limited period of time (e.g., 1 month) in part because the assessments are based on specific weather conditions that are unique to this time interval. The analyses, reviews, and launch approval are all based on the assumption that the launch vehicle and spacecraft will meet specific quality standards. However, the launch vehicle and spacecraft will not be completed until mere hours before launch. Thus, the launch approval decision comes with conditions attached e namely that the launch vehicle will be prepared in such a way that mission personnel can certify that the necessary quality standards have been met.
Launch Vehicle Certification for Flight Besides the payload safety review requirements for a spacecraft that utilizes a SNPS, considerations have to be made for the launch vehicle chosen to provide “the ride” to the destination. The most valuable validation of the selected launch vehicle’s strengths and weaknesses comes from the program’s launch vehicle databook. This probabilistic risk assessment (PRA) based book provides detailed information about each of the launch vehicle subsystem failure modes and reliability values. The effects of these failure modes are then examined further in the Safety Analysis Report (SAR) by developing release scenarios and ultimately calculating radioisotope exposure risk. The many existing safety requirements and regulations that exist for launching
6.8 Nuclear Mission Launch Integration
a “normal” payload apply and mission-unique “special” requirements are usually listed in these. For the most part, however, these regulations lead a Range User, preparing to bring a radioisotope system to the launch area for assembly, test, and finally launch to one place e the aforementioned PD/NSC-25 requirement to empanel an INSRP and begin a thorough review of the program’s nuclear payload mission. As an example, NASA’s Procedural Requirements (NPR) document for General Safety Program Requirements (NPR 8715.3, latest version) provides details on how NASA complies with the necessary requirements of PD/NSC-25. NASA Policy Directive 8610.7 (latest version) provides requirements for selection of a launch vehicle to transport varying classes of payloads into orbit/space. Increased risks (based on hazard, cost, national interest) required increased scrutiny of launch vehicle design and greater flight experience to mitigate those risks.
6.8 Nuclear Mission Launch Integration The hazardous nature of the radioisotope-laden payload introduces unique challenges to the normal spacecraft processing and integration process. The rigorous safety review process that exists for “normal” payloads is followed for nuclear missions as well; however, one simple principle tends to apply e introduce the radioisotope system at the launch vehicle processing facility as late in the prelaunch schedule as possible.
Design Influence This basic assumption does not come without a price however, since design, fabrication, test, and integration of often specialized attach mechanisms to accommodate a late arrival requires ingenuity, schedule availability, and valued spacecraft mass allocation. The investment is rewarded though, especially in minimizing overall “exposure” time of launch processing workers required to conduct tasks in proximity to the nuclear system’s appointed spacecraft real estate. Radioisotope systems can be mounted on booms to provide distance from sensitive instrumentation. This also allows more efficient radiative thermal control by having heat bearing surfaces exposed to the space environment. Simple radioisotope heater units tend to be “buried” in equipment such as the Warm Electronics Box (WEB) on the MER Rovers Spirit and Opportunity. Thermal cooling loops are utilized to transfer the decay heat as necessary to maintain internal component temperatures within critical functional ranges.
Potential detrimental effects of destruct The normal design of Range Safety Systems (RSSs) are driven by the need to (a) halt thrusting rocket engines (solid or liquid) in order to control a failing vehicle’s terminal impact point; and (b) disperse propellants as necessary to minimize explosive yield events upon impact with the ground. Since the main idea is to destroy the pressurization capability of a thrusting vehicle, there is very little concern for the consequences to other, mission-necessary subsystems (e.g., telemetry, solar power, command
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computers) since once a need arises that warrants activation of the destruct systems, their utility becomes rapidly less important in the ensuing explosive consequence. It is a very different story for payloads that maintain a radioisotope component. One can argue that any final data that can be capture via telemetry downlink is invaluable to resolve the “What happened” investigation that will follow a launch anomaly. Conversely, the vulnerability of an onboard SNPS must be considered so that an action meant to mitigate risks to the public, personnel, and critical resources from falling debris, toxic effluent exposure, and distant focusing overpressure hazards does not consequently expose the radioisotope containments systems to extreme environments. Range safety destruct system ordnance components provide penetration energy to defeat tankage and break motor casings. However, what occurs at the “non-business end” of a conical shaped charge (CSC) can be hazardous to spacecraft mounted RTGs. Very small metallic casing pieces of the CSC can translate from the back plane of the device at near-rifle velocity and the proximity of spacecraft systems leave them vulnerable to impacts. Heavier mounting bosses and brackets can also be expelled at slower velocities but the higher “beta” value makes them also very lethal. The additional concern for evaluation of side effects from flight termination activation is very similar in the world of human spaceflight. Much has been learned about the chaotic nature of vehicle breakup and explosion effects to space vehicle crew which translates well to the nuclear payload sensitivity issue. Difficulties in understanding the near-field effects of liquid and solid propellant explosions can greatly increase uncertainties in modeling these abort scenarios to calculate overall risks.
Launch Site Processing Issues After a spacecraft has completed assembly and test, normally at the owners offsite facility (e.g., JPL, APL, Boeing, Lockheed), it is shipped in whole or in part to the launch area. For nuclear payloads the utilization of radioisotope materials adds constraints to the already hectic processing schedule. Special permitting, licenses, and security needs can accompany these payloads. Additional personnel hazards are included since many tasks require workers to approach the radioisotope hazard zones in order to carry out their unassociated subsystem work. Normally used protective equipment may be newly vulnerable to the inherently high temperatures that exist on the surfaces of the radioisotope systems. One example is the standard Self-Contained Atmospheric Protection Ensemble (SCAPE) worn by propellant handlers. These suits are not designed for use around the high temperature system so special procedures are used and care has to be taken when required to be near the radioisotope sources. Radiation zones are set up to indicate exposure limitations and special personnel access and monitoring procedures are used. Dosimetry badges can provide tracking and warning of overexposure. One of the greatest advantages of transferring the radioisotope components to the spacecraft late in the flow is that many of the most hazardous processes have been completed. Launch vehicle electrical connections have been completed and tested, ordnance devices have been installed and verified, propellant loading has been tested
6.8 Nuclear Mission Launch Integration
so tanks have been through a cycle validating integrity. Catastrophic accidents, such as the Brazil explosion, which could be extremely detrimental to the radioisotope system, are much less likely to occur.
Launch Day Concerns As the final countdown approaches, missions utilizing radioisotope materials will have enhanced requirements to monitor early failures. Tracking the position of the launch vehicle along its trajectory is accomplished via human observation, metric optics (i.e., cinetheodolites), onboard guidance and navigation (inertial measurement unit) data transmitted to the ground, and radar information that is fed into a tracking solution. Except for the visual methods (human observation and cinetheodolites) the tracking methods tend to provide a point estimate of the vehicle’s position in three dimensional space. This information is compared to flight constraints (destruct lines) set to mitigate risk to the public and personnel. Mission Flight Control Officers (MFCOs) monitoring the launch vehicle behavior initiate the RSS if prescribed criteria are met that include shutdown of liquid rocket engines and, as necessary, activation of onboard destruct ordnance. The area within a calculated risk contour in which debris would possibly fall during a launch anomaly, termed the Flight Caution Area (FCA) is cleared of all public and personnel. Only a few Launch Essential Personnel are allowed to remain inside this area and only if the predicted risk is within acceptable limits and the purpose is “essential” to support the launch. Unique to nuclear payload missions, observing the launch vehicle longitudinal orientation becomes very important. During a typical launch, an accident that might deliver debris impacts to the ground around the launch pad certainly could happen, but the limits of that area are predictable and cleared of people to mitigate any risk. For a nuclear mission, one of the most devastating accident scenarios that can lead to a maximum release of radioisotope material involves a full stack intact impact (FSII), especially in a nose down configuration. The payload would impact the ground first, delivering extreme forces to the containment systems, followed by the remaining structure of the rocket and subsequent explosions and fires. While the containment devices are built robustly, there can be a low probability of failure and the additional exposure to high temperature solid propellant fire can lead to vaporization of the material. To mitigate this risk it is important that an accident condition is quickly recognized and MFCO action is taken to initiate the vehicle destruct system preventing the FSII (Botts, 2003). The destructed case also can provide risk of radioisotope release and subsequent vaporization but to a much less extent. The Range Safety Advisory System (RSAS) is used to recognize the potential of anomalous flight behavior that may drive the launch vehicle to a FSII condition. During the early phase of launch ascent (T-0 to approximately Tþ50 seconds) MFCOs monitor a display (Figure 6.54) that provides a visual depiction of the longitudinal attitude as the vehicle lifts, initiates a pitchover maneuver downrange, and
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Atlas
RSAS UNCLASSIFIED
Data Loss
Analog
Guidance
Total Rotation
Mission Time
0
Vehicle Status
0.0
45
Vehicle Attitude Centaur Cleared Tower
Atlas Angle From Nominal
90
0.0
Alarm Status
Acceleration
Tone Alarm Active
135
Tone Alarm override
180
Atlas Rotation
0.0
Tone Alarm Test
0
45
Main Engine L
90
+6
+6
135
+4
180
+2
Chamber Pressures
+4
1
4 +6 +4 +2
Main Engine R
–2 –4 –6 –2
L
R
–4
2
3
–6
+2
5 –6 –4 –2
+2 +4 +6 –2 –4 –6
FIGURE 6.54 Range Safety Advisory System display.
progresses from over land flight to ocean overflight. A deviation of the real-time longitudinal vector that reaches a delta value of 45 degrees from the planned nominal direction indicates that the failure in progress has reached a point that is unrecoverable. The RSAS triggers an audible alarm and visual indication that destruct action is necessary. MFCOs visually verify the alarm condition via camera observation and human observer confirmation and follow quickly with RSS activation. Additional cues can be provided to the MFCO by observing the launch vehicle rocket nozzle gimbal angles (also shown). A failure of the guidance system or mechanical actuators might be indicated if the nozzle thrustline indicator (white dot) moves to an extreme angle and remains there. This tells the MFCO to expect a pitch/yaw response from the vehicle that can lead to a FSII condition. Core and
6.9 Symbols and Acronyms
attached booster rocket chamber pressures indicate any loss of thrust conditions that can cue the MFCO to an expected pitch/yaw response. Based on thrust to weight (T/ W) capability of the vehicle, a loss of thrust to critical rocket motors may indicate a fallback is imminent. Once the vehicle and payload have traveled downrange sufficiently to confirm only ocean impacts would occur in a flight failure, the RSAS can be shut down and more traditional long range tracking becomes primary.
Radiological Contingency Concerns Anomalies can occur during the period from final integration of the SNPS through initiation of transfer to an Earth escape trajectory and, if the mission includes, through possible return to Earth (to gain gravitational boost) until the nuclear payload is bound for the extraterrestrial destination. Planning for these accidents and the radiological consequences is a large part of extra effort that takes place. Impacts near the launch pad and launch range would be initially handled by the Launch Emergency Operations Center (LEOC) and the Radiation Contingency Center (RADCC). The LEOC would act to cordon the area preventing entry so no one can approach any of the harmful toxic, explosive or, in this case, radioactive materials. Off-site detection, warning, emergency response, and recovery would be managed via the RADCC with representatives from range, program, local, state, and federal emergency agencies that have established procedures to mitigate risks to the general public. RADCC personnel can also initiate U.S. state department responses to alert other nations of possible or pending re-entry failures. Efforts following the initial response to mitigate risk to humans would focus on locating, assessing, containing, and transporting any SNPS components in the cordon area. Longer term cleanup operations may be necessary for very extreme accident cases.
6.9 Symbols and Acronyms a Dh ε εAl εg εT 3 T q ma ms r ra s sc
entrainment coefficient change in enthalpy effective emissivity of fireball surface emissivity of aluminum structure effective gas emissivity turbulent energy dissipation rate volumetrically-averaged dissipation rate vortex azimuthal position within fireball dynamic viscosity of ambient air microsecond fireball density density of ambient air Stefan–Boltzmann constant ¼ 5.67 10–12 critical geometry
• Watts • • • cm2/sec3 cm2/sec3 radians g/cm,sec g/cm,sec g/cm3 g/cm3 W/cm2,K4 cm
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! y yq yr c J V
a A A AFSPS Al2O3 AP Ap APU ASRG atm b C cal CALE CBGS CBM CCDF Cd Ci cm COPUOS cp CPIA CTPB D dc de DEIS DELK Dt E Eb EIS EM
gas velocity vector angular component of vortex velocity vector radial component of vortex velocity vector dynamic shape factor of particle stream function del operator for cross product
cm/sec cm/sec cm/sec • • •
constant based on Reynold’s number surface area of spherical fireball cross-sectional area of propellant Affordable Fission Surface Power System aluminum oxide projected area of fireball effective projected area of particle Auxiliary Power Unit Advanced Stirling Radioisotope Generator atmosphere constant based on Reynold’s number number concentration of particles calorie C-language based Arbitrary Lagrangian–Eulerian Confined By Ground Surface Confined By Missile Complementary Cumulative Distribution Function drag coefficient Curie centimeter Committee On the Peaceful Uses of Outer Space [UN] constant-pressure specific heat Chemical Propulsion Information Agency carboxy-terminated polybutadiene maximum fireball diameter critical diameter diameter of volume-equivalent sphere for particles Draft EIS [PISCES-2] dimensional Eulerian–Lagrangian Coupled maximum fireball diameter enthalpy inflow from air entrainment blackbody emissive power (heat flux) Environmental Impact Statement electromagnetic
• cm2 cm2 • • cm2 Cm–3 • • atm • Cm–3 cal • • • • • Ci cm • J/mol,K • • feet cm cm • • feet Watts kW/m2 • •
6.9 Symbols and Acronyms
f FCA FEIS FONSI FPSE FPY FSAR FSII ft FTS FY g Gd2O3 GEM gnd GPa GPHS ha H2 H2O 2 ha hAl hf hF hp hr HTNT HTPB IAEA ICRP Isp IUS J K kbar Kd Ktþ kW kWe kWt L LANL
emissivity correction factor Flight Caution Area Final Environmental Impact Statement Finding of No Significant Impact Free Piston Stirling Engine full power year Final Safety Analysis Report Full-Stack Intact Impact feet Flight Termination System fiscal year acceleration due to gravity ¼ 981 gadolinium oxide Graphite Epoxy Motor ground gigapascal (109 pascal) General Purpose Heat Source enthalpy of ambient air hydrogen hydrogen peroxide enthalpy of entrained air heat transfer coefficient between Al structure and fireball molar enthalpy of fireball enthalpy of formation enthalpy of product species specified enthalpy of the reactants TNT blast energy hydroxy-terminated polybutadiene International Atomic Energy Agency International Commission on Radiological Protection specific impulse Inertial Upper Stage joule (kg$m2/sec2) Kelvin temperature kilobar explosive charge diameter blast wave positive-phase duration kiloWatts kiloWatts electric kiloWatts thermal fireball energy loss Los Alamos National Laboratory, New Mexico
• • • • • yr • • ft • yr cm/sec2 • • • GPa • J/mol • • J/mol W/cm2,K J/mol Watts J/mol J/mol J/kg • • • sec • J K kbar msec kW kW kW Watts •
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Lb lbf lbm lb LEOC LH2 LLNL LO2 LSC LWRHU m MEPAG MER mf MFCO MHW MMH mol ms MSFC N2H 4 N2O4 na n_ a n_ p n_ r NAC NASA NEPA NERVA nf NOA NOI 237 Np np Np Nr Ns NSC NTO NTP
effective thickness of fireball (mean beam length) pounds force pounds mass pounds Launch Emergency Operations Center liquid hydrogen Lawrence Livermore National Laboratory, California liquid oxygen Linear-Shaped Charge Light Weight Radioisotope Heater Unit meter Mars Exploration Program Analysis Group Mars Exploration Rover (Spirit and Opportunity) mass of fuel Mission Flight Control Officer multi-hundred Watt monomethyl hydrazine mole millisecond Marshall Space Flight Center, Alabama anhydrous hydrazine nitrogen tetroxide quantity of entrained air molar rate of air entrainment molar product production rate molar combustion rate of propellant reactants NASA Advisory Council National Aeronautics and Space Administration, USA National Environmental Protection Act, USA Nuclear Engine for Rocket Vehicle Applications quantity of combustion products and entrained air of fireball Notice of Availability Notice of Intent neptunium quantity of combustion products number of product species number of reactant species number of particle sizes National Security Council nitrogen tetroxide Nuclear Thermal Propulsion
cm lbf lbm lb • • • • • • m • • kg • • • mol ms • • • moles moles/sec moles/sec moles/sec • • • • moles • • • moles • • • • • •
6.9 Symbols and Acronyms
OMS OSTP p p PBAN Pb–Te PD PERMS PISCES-2DELK PIRAT PNH psia Pu 238 Pu 238 PuO2 PVF2 pwr q qAl r r R R RADCC rc Re RHU ROCETS RORSAT
RPS RSAS RSS ry s SCoRe-S SDT sec SER
Orbital Maneuvering System [Executive] Office of Science and Technology Policy pressure perimeter of propellant polybutadiene acrylonitrile lead telluride Presidential Directive Propellant Energetic Response to Mechanical Stimuli Physics Int’l multi-Spatial Code for Engineering and Science Propellant Impact Risk Assessment Team Pluto New Horizons pounds per square inch absolute plutonium plutonium-238 isotope plutonium-238 oxide polyvinylflouride power thermal radiation flux heat added to fireball from aluminum combustion fireball radius burning rate reactant enthalpy inflow universal gas constant ¼ 82 Radiological Control Center, NASA Kennedy Space Center truncated fireball radius Reynold’s number Radioisotope Heater Unit ROCket Engine Transient Simulation Radar Ocean Reconnaissance Satellite (U.S. name for Russian UPracm>fn9k SPutojl Altjco9k; Managed Satellite Activities) Radioisotope Power System Range Safety Advisory System Range Safety System vortex radius second Sectored Compact Reactor-Small Shock to Detonation Transition Second Safety Evaluation Report
• • atm cm • • • •
• • psia • • • • kW/m2 Watts cm in/sec Watts cm3,atm/mol,K • cm • • • •
• • • cm s • • S •
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SiGe SNPS SOPHY SPT SRB SRG SRM SRMGI SRMU SSME Stg STS t T T* Ta TAl tb tc Tc TE Th tT TVC 2DELK u 235
U UDMH UN UNCOPUOS UO2 U.S. V VLS W Wa Wf WG Wk Wp Wr WSMC
silicon germanium Space Nuclear Power System Solid Propellant Hazards Program Seebeck–Peltier–Thompson Solid Rocket Booster Stirling Radioisotope Generator Solid Rocket Motor Solid Rocket Motor Ground Impact [model] Solid Rocket Motor Upgrade (Titan IV) Space Shuttle Main Engine Stage Space Transportation System (Space Shuttle) time temperature T / 1000 ambient temp surface temperature of aluminum structure fireball liftoff time stem liftoff time cold temperature thermoelectric hot temperature duration of visible radiation Thrust Vector Control [PISCES]-2 dimensional Eulerian–Lagrangian Coupled fireball rise velocity uranium-235 isotope unsymmetrical dimethylhydrazine United Nations United Nations COPUOS uranium oxide United States fireball volume Veı´culo Lanc¸ador de Sate´lites (Brazilian Launch Vehicle) weight of fuel and oxidizer combined molecular weight of air ¼ 28.97 mean molecular weight of fireball mixture Working Group species k molecular weight mean molecular weight of products mean molecular weight of reactants White Sands Missile Center, New Mexico
• • • • • • • • • • • • seconds K K K K sec sec • • • sec • • cm/s • • • • • • cm3 • lb g/mol g/mol • g/mol g/mol g/mol •
6.9 Symbols and Acronyms
wt WTNT XDT y ya yp Yr z
weight TNT equivalent weight Unknown Detonation Transition molar quant. of comb. products produced per mole of reactant molar air fraction (na/nf) mole fraction of product species year height of fireball center
• • • • • • cm
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6.9 Symbols and Acronyms
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