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An ‘entry level’ mission to a near Earth object
Acta Astronautica 59 (2006) 845 – 857 www.elsevier.com/locate/actaastro
An ‘entry level’ mission to a near Earth object Andy Phipps∗ , Max Meerman, James Wilhelm, Dave Gibbon, James Northam, Alex da Silva Curiel, Jeff Ward, Martin Sweeting Surrey Satellite Technology Ltd., Surrey Space Centre, Guildford, Surrey, GU2 7XH, UK Available online 13 September 2005
Abstract Near Earth Objects (NEOs) are comets or asteroids that intersect or pass near to our planet posing a real and underestimated danger to mankind. While the probability of impact is low, the consequences of such an impact could be apocalyptic. Various programs are underway to discover these kilometer-sized objects from Earth. However, once targets of interest have been identified a fly-by or orbiting spacecraft is required to understand the objects’ mass, morphology and composition. Fly-past NEO missions represent the simplest interplanetary missions and need not be high cost. An ‘entry level’ mission has been conceptually designed able to deliver a 10 kg science ‘reference’ payload to NEO fly-by for a total mission cost (including launch and operations) of ¥20 million (FY2003). This paper outlines the platform architecture, cost and cost drivers, and describes the key technology trades to be performed and the developments required to extend current Low Earth Orbit (LEO) technology to a deep space mission. It concludes by identifying the top-level trade-offs to be made in order to enhance the science return of the mission. © 2005 Elsevier Ltd. All rights reserved.
1. Introduction Asteroids, like comets, are the remnant population of planetesimals from which the planets accumulated 4.6 billion years ago. Many lie in the main asteroid belt some between the orbits of Mars and Jupiter. However, collisions between asteroids eject fragments, and over millions of years many of these have been nudged into orbits that are within the Earth’s vicinity and as such are termed Near Earth Asteroids (NEAs). Those that have the potential to make threateningly close approaches to Earth (< 7.5 million km) and have an ∗ Corresponding author.
E-mail address: [email protected] (A. Phipps). 0094-5765/$ - see front matter © 2005 Elsevier Ltd. All rights reserved. doi:10.1016/j.actaastro.2005.07.021
absolute magnitude of 22.0 or less—nominally corresponding to objects with a diameter greater than 150 meters—are termed Potentially Hazardous NEOs. As of July 2003 there are 519 such objects known [1]. Following the Shoemaker Levy comet impact on Jupiter in 1994 (Fig. 1), interest in asteroids and comets has increased. While impacts of this magnitude are extremely rate, the fact that this impact was observable in our lifetime was alarming. In September 2000, the task force on Potentially Hazardous NEOs published its final report [2] recommending “that the Government explore, with likeminded countries, the case for mounting a number of coordinated space rendezvous missions based on relatively inexpensive microsatellites, each to visit
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2.1. Target selection Out of the 519 NEO’s labeled as ‘potentially hazardous’ 4179 Toutatis was selected as the reference target. Being a contact binary it is both scientifically interesting, and provides a timely close approach to Earth in 2008. 2.2. Launch The lowest cost launch opportunity—offering constraints that do not adversely drive the mission—is offered by a Proton secondary launch into GeoStationary Orbit (GSO). This approach to escape from Earth orbit is both novel and cost effective. Several low cost alternatives exist including the Ariane-5 Cyclade and Dnepr although some development may be required. Fig. 1. Comet Shoemaker-Levy 9 impact with Jupiter in 1994.
2.3. Trajectory simulation a different type of NEO to establish its detailed characteristics”. This paper addresses this recommendation outlining the platform architecture, cost and cost drivers of an ‘entry level’ mission to a NEO; describing the key technology trades to be performed; and the developments required to extend cost-effective LEO technology to a deep space mission. It concludes by identifying the top-level trade-offs to be made to enhance the science return of the mission.
2. Mission and system design The mission objectives are to: demonstrate the capability to intercept a NEO in deep space for low cost. An ‘entry level’ mission will be developed capable of supporting a 10 kg science ‘reference’ payload. Many of the successful mission design philosophies from earlier SSTL spacecraft have been followed [3]. In addition, to reduce risk the spacecraft will be designed for omni-directional communications and all-aspect power generation, significantly reducing mission risk and facilitating a less demanding operations regime. This conceptual platform design will be heavily based upon existing research on a low cost lunar orbiter and interplanetary mission [4,5].
Trajectory simulation was undertaken using Satellite Tool Kit (Astrogator) making the following assumptions: • Initial launch into GSO in March 2008. • Period of waiting is assumed in GSO for correct orbit geometry. • 20 Newton thruster firing over 3◦ true anomaly of each orbit. • 15 phasing orbits, leading to Earth escape after 10 weeks. • A plane change maneuver of 80 m s−1 ensures the spacecraft departs from the ecliptic plane to encounter the target. Simulation suggests a conservative approach to Earth escape (through a series of phasing orbits) is practical readily achieving a fly-past of the target. From this initial launch a practical Earth escape requires around 1300 m s−1 and target encounter requires around 1500 m s−1 . This is achievable with, commercially available, modest cost, chemical propulsion (Figs. 2 and 3). 2.4. Environment An assessment was made into the radiation and thermal environment the spacecraft will encounter and the
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Table 1 Radiation analysis assumptions
Fig. 2. Initial GSO launch and Earth escape trajectory.
Fig. 3. Fly-past on sunlight side (target Eros).
applicability of Commercial Off-The-Shelf Technology (COTS technology) for this mission. Total dose effects are caused by accumulated charge within the semiconductor material which manifests itself as changes in threshold voltage and consequently increasing leakage current. While in GSO the spacecraft is subject to geomagnetically trapped radiation from the Van Allen belts—largely consisting of trapped electrons in the outer belt. The spacecraft remains in this initial orbit for several months awaiting the correct departure geometry. While in interplanetary space it is subject to an increase in total dose from sparse but highly penetrative galactic cosmic rays and infrequent but intense solar storms.
Feature
Specification
Effective shielding thickness (mm aluminum) Launch data Orbit phases: GSO-escape (weeks) Interplanetary transfer (weeks) Post NEO fly-past/operation (weeks)
4 Mar-08 11 25 6
A prediction of the radiation environment the spacecraft will encounter was analysed using ESA’s SPacecraft ENVironment Information System (SPENVIS) code. Radiation models AP-8, AE-8 and JPL-91 (representing trapped protons, trapped electrons and solar proton fluencies) were utilized. A shielding thickness of 2 mm from sub-system module walls was assumed—leading to an effective thickness of 4 mm when considering the effects of adjacent sub-systems and propellant tanks (Table 1). Analysis suggests the expected cumulative mission radiation dose would be in the region of 6.0 krad (Si). However, the AP-8 & AE-8 models tend to overpredict the trapped electron and proton radiation environment and a de-rating factor of 2 times is generally applied [6]. While typical ‘soft’ COTS technology fails near 5–10 krad (Si) suitably selected technology can withstand as much as 100 krad (Si). A further reduction in total radiation dose can be made with the application of spot shielding—high density metal (usually tantalum or copper) applied to the component surface. Single event upset (SEU) occurs as a result of highenergy charged particles inducing charge deposition within a semiconductor device. This may result in a change in digital state or permanent damage due to an over-current condition, termed single event latch-up (SEL). The selected subsystems are SEU tolerant, both at subsystem level (by use of error detection and correction, multiple voting systems etc.) and at system level (the spacecraft will revert into a safe mode in the event of serious problem arising). The spacecraft will suffer a variable amount of SEUs depending on its orbital position with an increased risk whilst crossing the radiation belts. The likelihood of SEL occurring within the spacecraft (based upon previous occurrences in LEO) has
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Table 2 Design drivers and assumptions Design driver
Comments
V capability of 1500 (scaleable 2000 + m s−1 )
Solutions limited to chemical propulsion
Launch vehicle accommodation 1.4 m diameter, 1.4 m tall
Proton design for direct injection into GEO
Proton launch
400 kg maximum mass limit
Off-the-shelf propulsion components
No development of new tanks or thrusters
Design must be capable of stacking two spacecraft if Ariane or alternative launch baselined
Spacecraft must be small and sufficiently strong enough Antennas have to be checked for mechanical interference
Capable of spin stabilization
Moment of inertia aligned for correct spin axis
Structure simple to analyse, low cost to manufacture, easy to assemble, preferably with re-use of existing designs
Variety of platform structural concepts to be assessed
Largely dual redundant design
Avionics only
Platform should be capable of supporting a 1 m diameter communications antenna (if required)
Design must cater for compatibility with stacking, and Earth and asteroid target pointing
been estimated as once per spacecraft every three years. This SEL may be destructive, or, in approximately 80% of cases, benign. In many cases, powering down and then powering up the subsystem will eradicate the effects. The use of fast acting over-current sensing power switches is not envisaged at this stage. While not offering definitive protection these may mitigate the chances of burn-out occurring but increase system complexity (which in itself increases the probability of these effects occurring). The spacecraft thermal environment is characterized by long periods of slow rotation in sunlight (barbeque mode) and short periods of infrequent eclipse (lasting up to 72 min). Simulations suggest the interior of the spacecraft shows a high degree of thermal stability and although some optimization remains, it is possible to conclude that a passive thermal control system would meet the requirements. This would offer minimal cost and complexity with high reliability. The exterior of the spacecraft is also within the allowable range of temperatures for the units and subsystems located there. During eclipse or during periods of spacecraft anomaly (such as computer reset) the environment will be characterized by long periods with
only basic functions operating and hence very small amounts of power being dissipated. Electrical heaters will be required to sustain the temperature, and would be required to dissipate around 50 W.
2.5. Structure The structural design is subject to a number of early assumptions and mission drivers. This is to limit the number of possible solutions that have to be conceptually designed and analysed. These mission drivers and assumptions are identified (Table 2). A V requirement of 1300.1500 m s−1 drives the propellant mass fraction to 48–53% of the total mass of the spacecraft (assuming a Isp of 205 S). Assuming monopropellant hydrazine 160 L of propellant storage is required. No existing SSTL platform is capable of meeting this storage requirements however, to minimize development cost, the current Enhanced Microsatellite platform was conceptually modified to support this propellant (see Fig. 4). While compatible with the launch vehicle limitations, this solution was found to be
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Fig. 5. Four tanks in allowable outline. Fig. 4. Enhanced microsatellite modified to meet NEO mission requirements.
unsatisfactory because • The single string design violated the 120 kg mass limit on the structure, a largely dual redundant system produces a mass of 180 kg. • The propulsion system is not scaleable much beyond 1300 m s−1 . • The design will not readily lend itself to low cost spin stabilization. With the propulsion system is a major structural driver it has been decided to start with four off-the-shelf tanks (baselined for an existing project). In effect, the spacecraft will be designed around these (Fig. 5). Four tanks will fit inside the allowed cylinder of 1.4 m diameter, and there are several structural options possible. Four layouts were selected for an initial assessment: a thrust tube cylinder, an octagon, an open strut structure and an square load-bearing thrust tube configuration (based upon a geostationary spacecraft configuration). All have sufficient volume for external solar panels and antennas (Fig. 6). Advantages and disadvantages of these are • The cylindrical thrust tube (a) is simple to manufacture, but less simple to assemble and test avionics and the propulsion systems. This design builds upon in-house development and test heritage.
Fig. 6. Structural concepts: (A) thrust tube, (B) octagon, (C) open strut design, (D) square load-bearing thrust tube.
• The octagon (b) will be slightly heavier and more expensive than the cylinder, but it will be simpler to assemble the complete spacecraft. • The strut structure (c) is a new development, but simple to analyse and simple to assemble.
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A. Phipps et al. / Acta Astronautica 59 (2006) 845 – 857 Table 3 Mass budget Feature
Mass (kg)
Payload Propulsion Power Communications On-board data handling Environment AODCS and safety Structure (including harness and solar arrays) Margin
10.0 193.7 16.2 11.6 2.8 7.1 8.5 63.3 31.3
Total
344.4
honeycomb skin will be 1 mm thick, while overall thickness will be 20 mm. The mass budget is detailed above demonstrating compatibility with the 400 kg launch vehicle limit. A 10 kg science reference payload is included along with an additional 10% mass margin (Table 3). 2.6. Propulsion Fig. 7. Cylinder option internal view.
• The square load-bearing thrust tube structure (d) is currently being developed for an existing geostationary mission, but requires the addition of two extra tanks. It is not well suited to spin stabilization, and two platforms cannot be accommodated into the same launch slot. In order to progress the concept design cylindrical thrust tube (a) is selected because of the lower design and manufacturing costs, but also for the in-house experience with this technology. However, the strut design promises both low mass, low cost and facilitates scaling of the propulsion system and could be considered if the study is taken further (Fig. 7). The configuration developed has a diameter of 1.3 m, and a height of 0.6 m. The thrust tube is made of 2 mm thickness aluminum alloy sheet and is capable of supporting an additional spacecraft mounted on top enabling a dual launch. A cross-shaped honeycomb frame is located in the centre of the cylinder. The main function of this cross frame is to support the propulsion tanks and the avionics modules. The
Following the initial mission specific analysis the major propulsion system design drivers emerge as • Minimum specific impulse 200 S. The mission cannot be feasibly performed with anything lower. • Minimum thrust level 1 N, although greater than 10 N is preferred, this limits the exposure to higher radiation environments. • Low cost and low risk. Given these requirements the propulsion system options can quickly be narrowed down. Electric propulsion systems (typically hall effect thrusters and gridded-ion thrusters) produce thrust levels at least an order of magnitude lower than required (at reasonable power levels). Additionally they cannot be considered as low cost options. The specific impulse and thrust level requirements steer the solution towards a chemical propulsion system. Given the cost and risk driver, the optimum solution is to use a conventional blowdown monopropellant hydrazine system. A larger V could be supported, however with a more complex and expensive bipropellant solution.
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The baseline design uses four 60 L propellant tanks, each containing 41 kg of hydrazine. The tanks are 480 mm diameter spheres and are very similar to the oxidizer tanks flown on NASA’s NEAR-Shoemaker spacecraft. As most of the propulsive manoeuvres are performed during the spin phase of the mission a very simple internal sponge can be used over the tank outlet to provide propellant during the de-spun mission phase. All other equipments and thrusters would be flight qualified off-the-shelf items. The main 20 N engine would be supplemented by 6 or 8 smaller 1 N attitude and spin control thrusters. Each propellant tank has a maximum capacity of 45 kg of hydrazine in blowdown mode. Utilising this additional propellant load the V capability increases to 1600 m s−1 . With the addition of a propellant pressurization system the V can be further increased, whilst retaining the same tank and structure configuration. 220 kg of hydrazine could be used in regulated mode to give around 1850 m s−1 V. V can be further increased by using a conventional bipropellant MMH/NTO system. For a launch mass of 443 kg the V increases to around 2400 m s−1 . 2.7. Attitude determination and control system For cost minimization and simplicity the spacecraft will be spin stabilized during propulsive burns and the transfer phase, but during encounter will be de-spun. Assuming a spin rate of 10 rpm a propellant budget of 10 m s−1 is required during thrusting phases. This V being provided by the on-board hydrazine system. The different ADCS modes required to be supported are: initial de-tumble, spin up and thrust, and de-spin. The complement of sensors and actuators detailed in Table 4 is sufficient for the mission. 2.8. Communications, ranging and orbit determination The communication system is primarily driven by the 1.2 million km range during the Toutatis encounter, introducing a free space loss of around 220 dBs. To minimise cost, the communications system architecture employs existing SSTL LEO qualified S-band equipment. The downlink system consists of dual redundant 6 W transmitters. BPSK is used with convolutional
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encoding. The uplink system consists of two dual redundant receivers. These use CPFSK with no coding in common with LEO satellites. Two low gain antennas on each of the uplink and downlink will provide an omni-directional pattern for reliable communications at all spacecraft attitudes. A 0.8 m parabolic antenna will provide a high rate up/downlink facility. Use of this will be limited to periods when the satellite can be oriented to point the antenna towards Earth. One transmitter and one receiver can be switched to the high gain antenna. The standard SSTL 3.7 m S-band groundstation antenna is proposed with a G/T of 13 dB/K. An uplink transmit power of 200 W is required. This will require the installation of a high power amplifier at the current SSTL antenna. The low gain data rate will be 50 bps on both the uplink and downlink. A high gain rate of 9600 bps is assumed for the uplink and 19 200 bps for the downlink. These rates are based on what can be achieved using an antenna the size of the current SSTL groundstation antenna. Table 5 shows figures for each link. The link margin for the low gain uplink is very low. Use of a proposed 6 m dish on the groundstation and, or uprating the spacecraft transmitter to 10 W will solve this problem. Low data rates require low phase noise close to the local oscillator carrier on all transmitters and frequency converters. For the 50 bps links, a review and potentially an upgrade will be required for both the current groundstation and spacecraft equipment. SSTL’s existing LEO spacecraft rely heavily on GPS for positional fixes. Outside of LEO a pseudo random ranging system on the communication link will be employed. This is more compatible with current SSTL techniques and requires less infrastructure on the spacecraft than tone ranging. Pseudo random noise ranging was successfully demonstrated by Miller on AO-40 [7]. Total clock jitter in the groundstation equipment and satellite transponder should be as low as possible. A clock jitter of one hundredth of a bit is assumed [7]. A bit rate of 50 bps will give a coarse downrange range resolution (determined by bit ‘length’) of 3000 km, successive measurements will reduce these errors. With the assumed clock jitter fine ranging accuracy is 30 km. To avoid ambiguity a code length
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Table 4 ADCS sensors and actuators Type Sun sensor Star camera IMU (Gyros & Accelerometer) 1 N thrusters
Notes
SSTL AST20 Systron Donner MONARC
Closed loop propulsion control
Table 5 Link performance Link
Rate [bps]
Margin [dB]
Data per day
OBC reload
Low gain uplink Low gain downlink High gain uplink High gain downlink
50 50 9600 19 200
0.4 3.3 5.3 5.6
500 k 500 k 90 M 180 M
50 h – 15 min –
Table 6 Ranging performance Parameter
Low rate
Hi rate
Unit
Data rate Range at encounter Round trip time Minimum code length Selected code length Maximum range obtainable Code cycle time (initial correlation time) Coarse resolution Down-range accuracy Cross-range accuracy
50 1,200,000 8.0 400 1024 3,072,000 3000 1/100 of a bit 30 9830
9600 1,200,000 8.0 76800 76800 1,200,000 15.63 1/100 of a bit 0.16 51.11
bps km S bits
of at least 1024 bits is required at encounter distance (see Table 6). Utilising two ranging groundstations (Guildford UK and Colorado Springs USA) a 7300 km baseline is created. At an encounter range of 1.2 million km a crossrange accuracy of 9800 km is achieved. This is useful for coarse orbit determination using the omni antennas however, the use of the 9600 bps high gain link improves the down-range and cross-range to 0.16 and 51 km, respectively, again successive measurements will improve this accuracy. 2.9. Power The requirements for the power system design are to utilize current LEO technology where practical. Once
km km km km
outside of Earth orbit there are no eclipses, but in GSO up to 72 min can be expected. Because no influence over the launch timing can be expected for a secondary launch, the power system must be sized for this situation. Power demand has been assessed during the various phase of the mission. Initially, during commissioning, 40 W is required rising to 45 W during operations, largely due to greater demand from the downlink power amplifier and on-board data handling sub-system. During extended periods of eclipse all unnecessary sub-systems are powered down leaving receivers, network and computing sub-systems operating, as well as heaters to maintain the propellant tanks and plumbing above the minimal operating temperature. This consumes 51 W for the duration of the eclipse.
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The power system design is based upon LEO technology with minimal development—this provides the most cost-effective solution not necessarily the highest efficiency. Ten single junction (GaAs/Ge technology) solar panels provide all-aspect power generation allowing the spacecraft to slew to any solar aspect angle for extended periods of time, so reducing mission complexity and risk. An unregulated topology provides 26–35 V to the platform sub-systems, 5 V being provided by a dual redundant power conditioning module. A battery charge regulator for each solar panel feeds power to the bus and battery, dual redundant power conditioning module supplies a regulated 5 V output to the various subsystems. Two 28 V Nicad batteries provide a total power storage capacity of 264 W h. With an anticipated worse case depth of discharge of 50% the lifetime of the batteries will be sufficient at around 500 cycles. Sub-system operation is controlled by schedule files uploaded to the primary OBC. Should a negative power budget exist such that the battery voltage drops below a pre-determined point non-critical subsystems are automatically powered down to avoid battery depletion until the anomaly is resolved. 2.10. On-board data handling The mission objectives are met by two flight-proven 386 computers each with up to 64 Mbyte RAM. Each program memory is EDAC protected. The OBCs constantly watch-dog each other to ascertain if failure has occurred requiring a re-boot. Telemetry and command interface between subsystems is provided for by a Controller Area Network (CAN) bus (4 kbits s−1 ) and high speed data linking is via a 20 mbps LVDS link. As a result of the low rate uplink requiring around 50 h to re-load the OBC a fast re-load will be possible flight code will be loaded from an on-board PROM. Modifications to flight code (typical patches being 50–150 kbytes) can be uploaded in 2–7 h. 2.11. Imager The imager has been conceptually designed to provide proof of capability rather than being sci-
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ence driven. The requirements are: flyby velocity ∼10 km s−1 , typical Albedo 4%, wide dynamic range, flyby range 100–1000 km. A planar array sensor (rather than linear scan) has been selected so that the requirements of ADCS stability do not become a driver, although existing imager technology has been discounted due to the cost and mass. The requirements can be met by a 2-D CMOS array sensor (used in a snapshot mode) and developed from an existing system—offering a wide dynamic range, VGA resolution (640 × 480 pixels), and low complexity control electronics. This provides mass and power savings resulting in a low-cost demonstration of imaging capability for this mission. In the current configuration more than 200 images can be stored on the 386, on-board compression would substantially increase this. 2.12. Operations A major cost driver for typical interplanetary missions is mission operations. This mission will be based upon SSTL’s standard LEO operations philosophy which is established around autonomous operations with operator intervention only in the event of an anomaly, or to verify critical operations. Modifications to the existing operations infrastructure will be required to support orbit determination and manoeuvre planning. Ranging information will be used to reconstruct the orbit and differences from the desired orbit acting as orbit corrections. Time tagged schedule files will be periodically uploaded detailing key parameters including ADCS requirements, propulsion firing duration etc. It is assumed the Proton launch vehicle will place the spacecraft into an orbit visible from Guildford, UK.
3. Science enhancements This mission performs a simple NEO fly-past as a demonstration of capability and has been developed with a 10 kg science ‘reference’ payload. The Space Studies Board National Research Council published a study report outlining the scientific goals of a NEO research program to understand the orbital distribution, physical characteristics, composition and origin of such objects [8]. The physical
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characteristics (shape, size, albedo, spin characteristics and mass) as well as composition of NEOs can be readily studied during fly-by. This study assesses the capability of the developed platform to perform low cost science to meet these science requirements. This includes assessment of the mission drivers as well as the mission cost impact. It considers an off-the-shelf X-ray spectrometer and magnetometer, as well as the platform modifications that are required to perform a crude mass assay of a NEO. 3.1. X-ray spectrometer The use of the Compact Imaging X-ray Spectrometer (CIXS) on this mission has been explored, this instrument is ideal to provide high quality spectroscopic and imaging information of a NEO [9]. While the instrument can be readily accommodated, and is not too demanding on spacecraft resources, it will drive the mission design in the following manner: 1. During encounter the CIXS instrument must point at the target while simultaneously the solar monitor must be provided with an unobstructed view of the Sun (with an accuracy of ±1◦ ) to provide the background solar flux count. The current platform concept does not support both of these pointing requirements. 2. The current CIXS instrument uses a pushbroom array imposing a requirement on the spacecraft attitude modes. 3. The ideal range for this unmodified instrument is 25 km placing requirements on the ranging and imager targeting system. 3.2. Magnetometer It has been estimated that the asteroid Eros has a magnetic field strength (at 35 km altitude) of 500–18 000 nT [10]. The 3-axis magnetometer selected has a sensitivity of ±10 nT and a range of ±60 T and is suitable to perform rudimentary science. To minimize cost and complexity it is assumed that it will be spacecraft mounted, rather than boom mounted. It is further assumed the intrinsic spacecraft magnetic field can be largely calibrated out prior to encounter. Again, the close fly-by distance
of >100 km drives the ranging and image targeting system. Data is generated at 12 bit resolution for each of the three axis, and even with an increased sample rate of 200 Hz the quantity of data generated is not a mission driver. The instrument cost is < ¥30, 000, however the design, but steps necessary to ensure spacecraft magnetic cleanliness (outlined in [10]) and calibration costs are considerably higher (estimated at several ¥100 k). 3.3. Mass assay While mass can be determined by a variety of methods, the study examined a crude method utilizing onboard equipment as this would not significantly drive up mission cost. The mass of a NEO can be determined by the orbit deflection of the spacecraft during flyby—Hyperbolic trajectory flight path turning angle Eq. (1) [11]. Sin() =
1 1+
rca V 2 GM
,
(1)
where is the turning angle, rca the radius of closest approach, V the velocity relative to object, G the universal gravitational constant, and M the mass. This equation assumes the radius of closest approach to the NEO can be accurately measured— requiring either a laser rangefinder or radar ranging system, the study assumed the use of on-board RF equipment and tried to establish the accuracy of results, the development required and the likely cost impact. The existing RF system potentially could be modified to act as a primitive altimeter. Conceptually, distance to the target can be established by transmitting a pseudo random code at the target, and storing the returned data in the OBC memory, this would then be download at a later date for correlation. However, developments of this system would be required, these include: same frequency receiver, switchable bandwidth filters to cope with large variations in Doppler shift, and modifications to increase the system gain. It is estimated these developments may require ¥500–1000 k and consequently it may prove lower risk and cost to procure and modify an existing altimeter, perhaps designed for an aircraft or UAV.
A. Phipps et al. / Acta Astronautica 59 (2006) 845 – 857 Table 7 Mission cost estimate Non recurring engineering Flight model AIV LEOP Ground segment
¥1, 212, 000 ¥8, 460, 000 ¥377, 000 ¥8, 633, 000 ¥543, 000
Total mission cost
¥19, 225, 000
The spacecraft will be target pointing during the encounter, the 0.8 m high gain antenna with a beamwidth of 13◦ at S-band will not drive the attitude pointing requirements. Making the assumptions stated below, there is a miniscule orbit deflection of 1.98 × 10−8 degrees for a typical NEO. Assumption: Fly-past velocity 10 km s−1 NEO mass assumed to be 5 × 1013 kg Fly-by distance 100 km High-rate communications (0.16×51 km error ellipse) The duration of time it takes to observe the orbit deflection largely depends upon the orbit geometry. If the deflected spacecraft appears largely in the crosstrack errors then it takes nearly one year to observe any appreciable orbit deflection. If the orbit geometry can be selected such that the deflection appears in the down-range errors then the deflection can be observed in less than a day—a substantially longer time would be needed to characterize the orbit better leading to a better mass estimate.
4. Mission cost Detailed cost estimates for the mission appear in Table 7. Assumptions made • Commercial figures (customer: SSTL). • 2003 prices (¥1.42 = £1.00). • Costing estimates include use of two existing groundstations. The ‘entry level’ mission can be developed, launched and operated for a total mission cost of less than ¥20
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million. Assuming full use of one of the 64 Mbyte memory during the encounter the mission cost per returned data bit is ¥0.04. The cost drivers being launch, propulsion, solar panel procurement, structure development, and communication and ranging system development, For a mission that includes the science instruments (outlined above) the mission cost increases to ¥25 million (this figure does not include the cost associated with the scientific analysis of the results).
5. Technology developments The mission enabling developments for this ‘entry level’ mission are listed • Development of a thrust tube structural concept. • Communications system: operation at low data rates will require significant modification to the existing equipment. The implementation of a 200 W high power amplifier at the dish must be traded against a larger 6 m antenna. In order to improve the low rate downlink the existing spacecraft s-band high power amplifier must be modified from 4.3 W RF output to 6 W. • Implementation of a pseudo random ranging system. • Software developments include: NEO imager search algorithm, software to perform closed loop control over thruster firings as well as development of a lost-in-space algorithm. The following list details the key enabling technologies required to lower mission cost and enhance the mission science return bit per Euro. • Propulsion Technology—a conventional monopropellant system cost is around ¥4 million a bipropellant system ¥6 million. The use of solar electric propulsion or HTP-kerosene bipropellant system (both being developed at SSTL) offers the opportunity to increase V or cut mission cost. • Modifications to the existing LEO qualified star imager to operate in deep space would reduce procurement cost. • Advanced communications coding and data compression algorithm.
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• Greater on-board storage capability through the use of a SSDR or hard disc drive. • Greater efficiency power system including the development of the direct energy transfer topology and the use of lithium-ion or lithium polymer storage technology. • Development of LEO imager technology for science gathering.
6. Conclusion An ‘entry level’ interplanetary platform has been conceptually designed to undertake a wide variety of missions specifically a NEO fly-past mission being studied. The study has established that the total mission cost is less than ¥20 million (FY2003) including platform, operations, payload and launch. A 10 kg science ‘reference’ package has been included which is able to return 64 Mbyte of science in its fleeting fly-past of the NEO target. A large number of trade-offs are possible, including the accommodation of up to 34 kg of payload mass. The propulsion system is sized to provide 1500 m s−1 for the specific target studied. However, substantial margin exists in the design allowing up to 1600 m s−1 V (using off-the-shelf monopropellant technology) or 2320 m s−1 V (using bipropellant technology). Consequently, the platform is capable of supporting lunar exploration missions [4,5] or multiple fly-bys of opportune targets. For the ‘entry level’ mission the principle cost drivers have been minimized. A non-optimal secondary launch into GSO has been selected and trajectory and environmental analysis suggests this leads to the lowest cost solution. Current off-the-shelf LEO platforms are unsuitable for the mission as they are insufficiently mass and volumetrically optimized, forcing the development of a new platform. This is based upon a central thrust tube for cost and risk minimization. Many LEO avionics are applicable for the mission, but the communication system (including groundstation) must be modified to support data rates down to 50 bps and a new pseudo random ranging system developed. The required accuracy of encounter is a cost driver requiring higher data rates and consequently larger antennas or higher power
Fig. 8. ‘Entry level’ NEO platform.
amplifiers. A further cost driver is GSO eclipse survival, this drives the power storage requirements up to 150 W h. The radiation environment surprisingly is not a significant driver (< 3 krad (Si)) with the use of COTS technology components and systems (Fig. 8).
Acknowledgements The authors are indebted to the following for their assistance: Luis Gomes, Mark Allery, Chris Jackson, Kevin Morgan, Neville Bean, Craig Clark, John Buckley and Guy Richardson, and Fred Kennedy USAF.
References [1] http://neo.jpl.nasa.gov/neo/pha.html [2] Report of the Task Force on Potentially Hazardous Near Earth Objects. http://www.nearearthobjects.co.uk/report/ resourcestask_intro.cfm, September 2000. [3] J. Wertz, Larson, Wiley. Reducing Space Mission Cost. PoSAT-1 by Dr. Jeff Ward, Chapter 13.4. ISBN 1-88188305-1. [4] A. Phipps, et al., Low cost lunar orbiter system design, IEEE Aerospace Conference, Montana, USA, 7–14 March 2003.
A. Phipps et al. / Acta Astronautica 59 (2006) 845 – 857 [5] A. Phipps, et al., The Moon and Inner Planets—Getting there for Low Cost, 4th IAA International Conference on Low-Cost Planetary Missions, May 2–5, 2000. [6] J. Vette, AE/AP Trapped Particle Flux Maps 1966–1980, http://nssdc.gsfc.nasa.gov/space/model/magnetos/aeap.html, 25/04/02. [7] J. Miller, Phase III ranging System, http://www.amsat.org/ amsat/articles/g3ruh/123.html
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[8] Space Studies Board National Research Council, The Exploration of Near-Earth Objects. National Academy Press, ISBN 0-309-06083-4, 1998. [9] http://sspg1.bnsc.rl.ac.uk/Share/documents.html [10] D. Lohr, et al., The NEAR Magnetic field Instrument, John Hopkins APL Technical Digest, vol. 19(2), 1998. [11] Bate, Mueller, White. Fundamentals of Astrodynamics, 0-486-60061-0, 1971.
An ‘entry level’ mission to a near Earth object Andy Phipps∗ , Max Meerman, James Wilhelm, Dave Gibbon, James Northam, Alex da Silva Curiel, Jeff Ward, Martin Sweeting Surrey Satellite Technology Ltd., Surrey Space Centre, Guildford, Surrey, GU2 7XH, UK Available online 13 September 2005
Abstract Near Earth Objects (NEOs) are comets or asteroids that intersect or pass near to our planet posing a real and underestimated danger to mankind. While the probability of impact is low, the consequences of such an impact could be apocalyptic. Various programs are underway to discover these kilometer-sized objects from Earth. However, once targets of interest have been identified a fly-by or orbiting spacecraft is required to understand the objects’ mass, morphology and composition. Fly-past NEO missions represent the simplest interplanetary missions and need not be high cost. An ‘entry level’ mission has been conceptually designed able to deliver a 10 kg science ‘reference’ payload to NEO fly-by for a total mission cost (including launch and operations) of ¥20 million (FY2003). This paper outlines the platform architecture, cost and cost drivers, and describes the key technology trades to be performed and the developments required to extend current Low Earth Orbit (LEO) technology to a deep space mission. It concludes by identifying the top-level trade-offs to be made in order to enhance the science return of the mission. © 2005 Elsevier Ltd. All rights reserved.
1. Introduction Asteroids, like comets, are the remnant population of planetesimals from which the planets accumulated 4.6 billion years ago. Many lie in the main asteroid belt some between the orbits of Mars and Jupiter. However, collisions between asteroids eject fragments, and over millions of years many of these have been nudged into orbits that are within the Earth’s vicinity and as such are termed Near Earth Asteroids (NEAs). Those that have the potential to make threateningly close approaches to Earth (< 7.5 million km) and have an ∗ Corresponding author.
E-mail address: [email protected] (A. Phipps). 0094-5765/$ - see front matter © 2005 Elsevier Ltd. All rights reserved. doi:10.1016/j.actaastro.2005.07.021
absolute magnitude of 22.0 or less—nominally corresponding to objects with a diameter greater than 150 meters—are termed Potentially Hazardous NEOs. As of July 2003 there are 519 such objects known [1]. Following the Shoemaker Levy comet impact on Jupiter in 1994 (Fig. 1), interest in asteroids and comets has increased. While impacts of this magnitude are extremely rate, the fact that this impact was observable in our lifetime was alarming. In September 2000, the task force on Potentially Hazardous NEOs published its final report [2] recommending “that the Government explore, with likeminded countries, the case for mounting a number of coordinated space rendezvous missions based on relatively inexpensive microsatellites, each to visit
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2.1. Target selection Out of the 519 NEO’s labeled as ‘potentially hazardous’ 4179 Toutatis was selected as the reference target. Being a contact binary it is both scientifically interesting, and provides a timely close approach to Earth in 2008. 2.2. Launch The lowest cost launch opportunity—offering constraints that do not adversely drive the mission—is offered by a Proton secondary launch into GeoStationary Orbit (GSO). This approach to escape from Earth orbit is both novel and cost effective. Several low cost alternatives exist including the Ariane-5 Cyclade and Dnepr although some development may be required. Fig. 1. Comet Shoemaker-Levy 9 impact with Jupiter in 1994.
2.3. Trajectory simulation a different type of NEO to establish its detailed characteristics”. This paper addresses this recommendation outlining the platform architecture, cost and cost drivers of an ‘entry level’ mission to a NEO; describing the key technology trades to be performed; and the developments required to extend cost-effective LEO technology to a deep space mission. It concludes by identifying the top-level trade-offs to be made to enhance the science return of the mission.
2. Mission and system design The mission objectives are to: demonstrate the capability to intercept a NEO in deep space for low cost. An ‘entry level’ mission will be developed capable of supporting a 10 kg science ‘reference’ payload. Many of the successful mission design philosophies from earlier SSTL spacecraft have been followed [3]. In addition, to reduce risk the spacecraft will be designed for omni-directional communications and all-aspect power generation, significantly reducing mission risk and facilitating a less demanding operations regime. This conceptual platform design will be heavily based upon existing research on a low cost lunar orbiter and interplanetary mission [4,5].
Trajectory simulation was undertaken using Satellite Tool Kit (Astrogator) making the following assumptions: • Initial launch into GSO in March 2008. • Period of waiting is assumed in GSO for correct orbit geometry. • 20 Newton thruster firing over 3◦ true anomaly of each orbit. • 15 phasing orbits, leading to Earth escape after 10 weeks. • A plane change maneuver of 80 m s−1 ensures the spacecraft departs from the ecliptic plane to encounter the target. Simulation suggests a conservative approach to Earth escape (through a series of phasing orbits) is practical readily achieving a fly-past of the target. From this initial launch a practical Earth escape requires around 1300 m s−1 and target encounter requires around 1500 m s−1 . This is achievable with, commercially available, modest cost, chemical propulsion (Figs. 2 and 3). 2.4. Environment An assessment was made into the radiation and thermal environment the spacecraft will encounter and the
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Table 1 Radiation analysis assumptions
Fig. 2. Initial GSO launch and Earth escape trajectory.
Fig. 3. Fly-past on sunlight side (target Eros).
applicability of Commercial Off-The-Shelf Technology (COTS technology) for this mission. Total dose effects are caused by accumulated charge within the semiconductor material which manifests itself as changes in threshold voltage and consequently increasing leakage current. While in GSO the spacecraft is subject to geomagnetically trapped radiation from the Van Allen belts—largely consisting of trapped electrons in the outer belt. The spacecraft remains in this initial orbit for several months awaiting the correct departure geometry. While in interplanetary space it is subject to an increase in total dose from sparse but highly penetrative galactic cosmic rays and infrequent but intense solar storms.
Feature
Specification
Effective shielding thickness (mm aluminum) Launch data Orbit phases: GSO-escape (weeks) Interplanetary transfer (weeks) Post NEO fly-past/operation (weeks)
4 Mar-08 11 25 6
A prediction of the radiation environment the spacecraft will encounter was analysed using ESA’s SPacecraft ENVironment Information System (SPENVIS) code. Radiation models AP-8, AE-8 and JPL-91 (representing trapped protons, trapped electrons and solar proton fluencies) were utilized. A shielding thickness of 2 mm from sub-system module walls was assumed—leading to an effective thickness of 4 mm when considering the effects of adjacent sub-systems and propellant tanks (Table 1). Analysis suggests the expected cumulative mission radiation dose would be in the region of 6.0 krad (Si). However, the AP-8 & AE-8 models tend to overpredict the trapped electron and proton radiation environment and a de-rating factor of 2 times is generally applied [6]. While typical ‘soft’ COTS technology fails near 5–10 krad (Si) suitably selected technology can withstand as much as 100 krad (Si). A further reduction in total radiation dose can be made with the application of spot shielding—high density metal (usually tantalum or copper) applied to the component surface. Single event upset (SEU) occurs as a result of highenergy charged particles inducing charge deposition within a semiconductor device. This may result in a change in digital state or permanent damage due to an over-current condition, termed single event latch-up (SEL). The selected subsystems are SEU tolerant, both at subsystem level (by use of error detection and correction, multiple voting systems etc.) and at system level (the spacecraft will revert into a safe mode in the event of serious problem arising). The spacecraft will suffer a variable amount of SEUs depending on its orbital position with an increased risk whilst crossing the radiation belts. The likelihood of SEL occurring within the spacecraft (based upon previous occurrences in LEO) has
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Table 2 Design drivers and assumptions Design driver
Comments
V capability of 1500 (scaleable 2000 + m s−1 )
Solutions limited to chemical propulsion
Launch vehicle accommodation 1.4 m diameter, 1.4 m tall
Proton design for direct injection into GEO
Proton launch
400 kg maximum mass limit
Off-the-shelf propulsion components
No development of new tanks or thrusters
Design must be capable of stacking two spacecraft if Ariane or alternative launch baselined
Spacecraft must be small and sufficiently strong enough Antennas have to be checked for mechanical interference
Capable of spin stabilization
Moment of inertia aligned for correct spin axis
Structure simple to analyse, low cost to manufacture, easy to assemble, preferably with re-use of existing designs
Variety of platform structural concepts to be assessed
Largely dual redundant design
Avionics only
Platform should be capable of supporting a 1 m diameter communications antenna (if required)
Design must cater for compatibility with stacking, and Earth and asteroid target pointing
been estimated as once per spacecraft every three years. This SEL may be destructive, or, in approximately 80% of cases, benign. In many cases, powering down and then powering up the subsystem will eradicate the effects. The use of fast acting over-current sensing power switches is not envisaged at this stage. While not offering definitive protection these may mitigate the chances of burn-out occurring but increase system complexity (which in itself increases the probability of these effects occurring). The spacecraft thermal environment is characterized by long periods of slow rotation in sunlight (barbeque mode) and short periods of infrequent eclipse (lasting up to 72 min). Simulations suggest the interior of the spacecraft shows a high degree of thermal stability and although some optimization remains, it is possible to conclude that a passive thermal control system would meet the requirements. This would offer minimal cost and complexity with high reliability. The exterior of the spacecraft is also within the allowable range of temperatures for the units and subsystems located there. During eclipse or during periods of spacecraft anomaly (such as computer reset) the environment will be characterized by long periods with
only basic functions operating and hence very small amounts of power being dissipated. Electrical heaters will be required to sustain the temperature, and would be required to dissipate around 50 W.
2.5. Structure The structural design is subject to a number of early assumptions and mission drivers. This is to limit the number of possible solutions that have to be conceptually designed and analysed. These mission drivers and assumptions are identified (Table 2). A V requirement of 1300.1500 m s−1 drives the propellant mass fraction to 48–53% of the total mass of the spacecraft (assuming a Isp of 205 S). Assuming monopropellant hydrazine 160 L of propellant storage is required. No existing SSTL platform is capable of meeting this storage requirements however, to minimize development cost, the current Enhanced Microsatellite platform was conceptually modified to support this propellant (see Fig. 4). While compatible with the launch vehicle limitations, this solution was found to be
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Fig. 5. Four tanks in allowable outline. Fig. 4. Enhanced microsatellite modified to meet NEO mission requirements.
unsatisfactory because • The single string design violated the 120 kg mass limit on the structure, a largely dual redundant system produces a mass of 180 kg. • The propulsion system is not scaleable much beyond 1300 m s−1 . • The design will not readily lend itself to low cost spin stabilization. With the propulsion system is a major structural driver it has been decided to start with four off-the-shelf tanks (baselined for an existing project). In effect, the spacecraft will be designed around these (Fig. 5). Four tanks will fit inside the allowed cylinder of 1.4 m diameter, and there are several structural options possible. Four layouts were selected for an initial assessment: a thrust tube cylinder, an octagon, an open strut structure and an square load-bearing thrust tube configuration (based upon a geostationary spacecraft configuration). All have sufficient volume for external solar panels and antennas (Fig. 6). Advantages and disadvantages of these are • The cylindrical thrust tube (a) is simple to manufacture, but less simple to assemble and test avionics and the propulsion systems. This design builds upon in-house development and test heritage.
Fig. 6. Structural concepts: (A) thrust tube, (B) octagon, (C) open strut design, (D) square load-bearing thrust tube.
• The octagon (b) will be slightly heavier and more expensive than the cylinder, but it will be simpler to assemble the complete spacecraft. • The strut structure (c) is a new development, but simple to analyse and simple to assemble.
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A. Phipps et al. / Acta Astronautica 59 (2006) 845 – 857 Table 3 Mass budget Feature
Mass (kg)
Payload Propulsion Power Communications On-board data handling Environment AODCS and safety Structure (including harness and solar arrays) Margin
10.0 193.7 16.2 11.6 2.8 7.1 8.5 63.3 31.3
Total
344.4
honeycomb skin will be 1 mm thick, while overall thickness will be 20 mm. The mass budget is detailed above demonstrating compatibility with the 400 kg launch vehicle limit. A 10 kg science reference payload is included along with an additional 10% mass margin (Table 3). 2.6. Propulsion Fig. 7. Cylinder option internal view.
• The square load-bearing thrust tube structure (d) is currently being developed for an existing geostationary mission, but requires the addition of two extra tanks. It is not well suited to spin stabilization, and two platforms cannot be accommodated into the same launch slot. In order to progress the concept design cylindrical thrust tube (a) is selected because of the lower design and manufacturing costs, but also for the in-house experience with this technology. However, the strut design promises both low mass, low cost and facilitates scaling of the propulsion system and could be considered if the study is taken further (Fig. 7). The configuration developed has a diameter of 1.3 m, and a height of 0.6 m. The thrust tube is made of 2 mm thickness aluminum alloy sheet and is capable of supporting an additional spacecraft mounted on top enabling a dual launch. A cross-shaped honeycomb frame is located in the centre of the cylinder. The main function of this cross frame is to support the propulsion tanks and the avionics modules. The
Following the initial mission specific analysis the major propulsion system design drivers emerge as • Minimum specific impulse 200 S. The mission cannot be feasibly performed with anything lower. • Minimum thrust level 1 N, although greater than 10 N is preferred, this limits the exposure to higher radiation environments. • Low cost and low risk. Given these requirements the propulsion system options can quickly be narrowed down. Electric propulsion systems (typically hall effect thrusters and gridded-ion thrusters) produce thrust levels at least an order of magnitude lower than required (at reasonable power levels). Additionally they cannot be considered as low cost options. The specific impulse and thrust level requirements steer the solution towards a chemical propulsion system. Given the cost and risk driver, the optimum solution is to use a conventional blowdown monopropellant hydrazine system. A larger V could be supported, however with a more complex and expensive bipropellant solution.
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The baseline design uses four 60 L propellant tanks, each containing 41 kg of hydrazine. The tanks are 480 mm diameter spheres and are very similar to the oxidizer tanks flown on NASA’s NEAR-Shoemaker spacecraft. As most of the propulsive manoeuvres are performed during the spin phase of the mission a very simple internal sponge can be used over the tank outlet to provide propellant during the de-spun mission phase. All other equipments and thrusters would be flight qualified off-the-shelf items. The main 20 N engine would be supplemented by 6 or 8 smaller 1 N attitude and spin control thrusters. Each propellant tank has a maximum capacity of 45 kg of hydrazine in blowdown mode. Utilising this additional propellant load the V capability increases to 1600 m s−1 . With the addition of a propellant pressurization system the V can be further increased, whilst retaining the same tank and structure configuration. 220 kg of hydrazine could be used in regulated mode to give around 1850 m s−1 V. V can be further increased by using a conventional bipropellant MMH/NTO system. For a launch mass of 443 kg the V increases to around 2400 m s−1 . 2.7. Attitude determination and control system For cost minimization and simplicity the spacecraft will be spin stabilized during propulsive burns and the transfer phase, but during encounter will be de-spun. Assuming a spin rate of 10 rpm a propellant budget of 10 m s−1 is required during thrusting phases. This V being provided by the on-board hydrazine system. The different ADCS modes required to be supported are: initial de-tumble, spin up and thrust, and de-spin. The complement of sensors and actuators detailed in Table 4 is sufficient for the mission. 2.8. Communications, ranging and orbit determination The communication system is primarily driven by the 1.2 million km range during the Toutatis encounter, introducing a free space loss of around 220 dBs. To minimise cost, the communications system architecture employs existing SSTL LEO qualified S-band equipment. The downlink system consists of dual redundant 6 W transmitters. BPSK is used with convolutional
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encoding. The uplink system consists of two dual redundant receivers. These use CPFSK with no coding in common with LEO satellites. Two low gain antennas on each of the uplink and downlink will provide an omni-directional pattern for reliable communications at all spacecraft attitudes. A 0.8 m parabolic antenna will provide a high rate up/downlink facility. Use of this will be limited to periods when the satellite can be oriented to point the antenna towards Earth. One transmitter and one receiver can be switched to the high gain antenna. The standard SSTL 3.7 m S-band groundstation antenna is proposed with a G/T of 13 dB/K. An uplink transmit power of 200 W is required. This will require the installation of a high power amplifier at the current SSTL antenna. The low gain data rate will be 50 bps on both the uplink and downlink. A high gain rate of 9600 bps is assumed for the uplink and 19 200 bps for the downlink. These rates are based on what can be achieved using an antenna the size of the current SSTL groundstation antenna. Table 5 shows figures for each link. The link margin for the low gain uplink is very low. Use of a proposed 6 m dish on the groundstation and, or uprating the spacecraft transmitter to 10 W will solve this problem. Low data rates require low phase noise close to the local oscillator carrier on all transmitters and frequency converters. For the 50 bps links, a review and potentially an upgrade will be required for both the current groundstation and spacecraft equipment. SSTL’s existing LEO spacecraft rely heavily on GPS for positional fixes. Outside of LEO a pseudo random ranging system on the communication link will be employed. This is more compatible with current SSTL techniques and requires less infrastructure on the spacecraft than tone ranging. Pseudo random noise ranging was successfully demonstrated by Miller on AO-40 [7]. Total clock jitter in the groundstation equipment and satellite transponder should be as low as possible. A clock jitter of one hundredth of a bit is assumed [7]. A bit rate of 50 bps will give a coarse downrange range resolution (determined by bit ‘length’) of 3000 km, successive measurements will reduce these errors. With the assumed clock jitter fine ranging accuracy is 30 km. To avoid ambiguity a code length
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Table 4 ADCS sensors and actuators Type Sun sensor Star camera IMU (Gyros & Accelerometer) 1 N thrusters
Notes
SSTL AST20 Systron Donner MONARC
Closed loop propulsion control
Table 5 Link performance Link
Rate [bps]
Margin [dB]
Data per day
OBC reload
Low gain uplink Low gain downlink High gain uplink High gain downlink
50 50 9600 19 200
0.4 3.3 5.3 5.6
500 k 500 k 90 M 180 M
50 h – 15 min –
Table 6 Ranging performance Parameter
Low rate
Hi rate
Unit
Data rate Range at encounter Round trip time Minimum code length Selected code length Maximum range obtainable Code cycle time (initial correlation time) Coarse resolution Down-range accuracy Cross-range accuracy
50 1,200,000 8.0 400 1024 3,072,000 3000 1/100 of a bit 30 9830
9600 1,200,000 8.0 76800 76800 1,200,000 15.63 1/100 of a bit 0.16 51.11
bps km S bits
of at least 1024 bits is required at encounter distance (see Table 6). Utilising two ranging groundstations (Guildford UK and Colorado Springs USA) a 7300 km baseline is created. At an encounter range of 1.2 million km a crossrange accuracy of 9800 km is achieved. This is useful for coarse orbit determination using the omni antennas however, the use of the 9600 bps high gain link improves the down-range and cross-range to 0.16 and 51 km, respectively, again successive measurements will improve this accuracy. 2.9. Power The requirements for the power system design are to utilize current LEO technology where practical. Once
km km km km
outside of Earth orbit there are no eclipses, but in GSO up to 72 min can be expected. Because no influence over the launch timing can be expected for a secondary launch, the power system must be sized for this situation. Power demand has been assessed during the various phase of the mission. Initially, during commissioning, 40 W is required rising to 45 W during operations, largely due to greater demand from the downlink power amplifier and on-board data handling sub-system. During extended periods of eclipse all unnecessary sub-systems are powered down leaving receivers, network and computing sub-systems operating, as well as heaters to maintain the propellant tanks and plumbing above the minimal operating temperature. This consumes 51 W for the duration of the eclipse.
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The power system design is based upon LEO technology with minimal development—this provides the most cost-effective solution not necessarily the highest efficiency. Ten single junction (GaAs/Ge technology) solar panels provide all-aspect power generation allowing the spacecraft to slew to any solar aspect angle for extended periods of time, so reducing mission complexity and risk. An unregulated topology provides 26–35 V to the platform sub-systems, 5 V being provided by a dual redundant power conditioning module. A battery charge regulator for each solar panel feeds power to the bus and battery, dual redundant power conditioning module supplies a regulated 5 V output to the various subsystems. Two 28 V Nicad batteries provide a total power storage capacity of 264 W h. With an anticipated worse case depth of discharge of 50% the lifetime of the batteries will be sufficient at around 500 cycles. Sub-system operation is controlled by schedule files uploaded to the primary OBC. Should a negative power budget exist such that the battery voltage drops below a pre-determined point non-critical subsystems are automatically powered down to avoid battery depletion until the anomaly is resolved. 2.10. On-board data handling The mission objectives are met by two flight-proven 386 computers each with up to 64 Mbyte RAM. Each program memory is EDAC protected. The OBCs constantly watch-dog each other to ascertain if failure has occurred requiring a re-boot. Telemetry and command interface between subsystems is provided for by a Controller Area Network (CAN) bus (4 kbits s−1 ) and high speed data linking is via a 20 mbps LVDS link. As a result of the low rate uplink requiring around 50 h to re-load the OBC a fast re-load will be possible flight code will be loaded from an on-board PROM. Modifications to flight code (typical patches being 50–150 kbytes) can be uploaded in 2–7 h. 2.11. Imager The imager has been conceptually designed to provide proof of capability rather than being sci-
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ence driven. The requirements are: flyby velocity ∼10 km s−1 , typical Albedo 4%, wide dynamic range, flyby range 100–1000 km. A planar array sensor (rather than linear scan) has been selected so that the requirements of ADCS stability do not become a driver, although existing imager technology has been discounted due to the cost and mass. The requirements can be met by a 2-D CMOS array sensor (used in a snapshot mode) and developed from an existing system—offering a wide dynamic range, VGA resolution (640 × 480 pixels), and low complexity control electronics. This provides mass and power savings resulting in a low-cost demonstration of imaging capability for this mission. In the current configuration more than 200 images can be stored on the 386, on-board compression would substantially increase this. 2.12. Operations A major cost driver for typical interplanetary missions is mission operations. This mission will be based upon SSTL’s standard LEO operations philosophy which is established around autonomous operations with operator intervention only in the event of an anomaly, or to verify critical operations. Modifications to the existing operations infrastructure will be required to support orbit determination and manoeuvre planning. Ranging information will be used to reconstruct the orbit and differences from the desired orbit acting as orbit corrections. Time tagged schedule files will be periodically uploaded detailing key parameters including ADCS requirements, propulsion firing duration etc. It is assumed the Proton launch vehicle will place the spacecraft into an orbit visible from Guildford, UK.
3. Science enhancements This mission performs a simple NEO fly-past as a demonstration of capability and has been developed with a 10 kg science ‘reference’ payload. The Space Studies Board National Research Council published a study report outlining the scientific goals of a NEO research program to understand the orbital distribution, physical characteristics, composition and origin of such objects [8]. The physical
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characteristics (shape, size, albedo, spin characteristics and mass) as well as composition of NEOs can be readily studied during fly-by. This study assesses the capability of the developed platform to perform low cost science to meet these science requirements. This includes assessment of the mission drivers as well as the mission cost impact. It considers an off-the-shelf X-ray spectrometer and magnetometer, as well as the platform modifications that are required to perform a crude mass assay of a NEO. 3.1. X-ray spectrometer The use of the Compact Imaging X-ray Spectrometer (CIXS) on this mission has been explored, this instrument is ideal to provide high quality spectroscopic and imaging information of a NEO [9]. While the instrument can be readily accommodated, and is not too demanding on spacecraft resources, it will drive the mission design in the following manner: 1. During encounter the CIXS instrument must point at the target while simultaneously the solar monitor must be provided with an unobstructed view of the Sun (with an accuracy of ±1◦ ) to provide the background solar flux count. The current platform concept does not support both of these pointing requirements. 2. The current CIXS instrument uses a pushbroom array imposing a requirement on the spacecraft attitude modes. 3. The ideal range for this unmodified instrument is 25 km placing requirements on the ranging and imager targeting system. 3.2. Magnetometer It has been estimated that the asteroid Eros has a magnetic field strength (at 35 km altitude) of 500–18 000 nT [10]. The 3-axis magnetometer selected has a sensitivity of ±10 nT and a range of ±60 T and is suitable to perform rudimentary science. To minimize cost and complexity it is assumed that it will be spacecraft mounted, rather than boom mounted. It is further assumed the intrinsic spacecraft magnetic field can be largely calibrated out prior to encounter. Again, the close fly-by distance
of >100 km drives the ranging and image targeting system. Data is generated at 12 bit resolution for each of the three axis, and even with an increased sample rate of 200 Hz the quantity of data generated is not a mission driver. The instrument cost is < ¥30, 000, however the design, but steps necessary to ensure spacecraft magnetic cleanliness (outlined in [10]) and calibration costs are considerably higher (estimated at several ¥100 k). 3.3. Mass assay While mass can be determined by a variety of methods, the study examined a crude method utilizing onboard equipment as this would not significantly drive up mission cost. The mass of a NEO can be determined by the orbit deflection of the spacecraft during flyby—Hyperbolic trajectory flight path turning angle Eq. (1) [11]. Sin() =
1 1+
rca V 2 GM
,
(1)
where is the turning angle, rca the radius of closest approach, V the velocity relative to object, G the universal gravitational constant, and M the mass. This equation assumes the radius of closest approach to the NEO can be accurately measured— requiring either a laser rangefinder or radar ranging system, the study assumed the use of on-board RF equipment and tried to establish the accuracy of results, the development required and the likely cost impact. The existing RF system potentially could be modified to act as a primitive altimeter. Conceptually, distance to the target can be established by transmitting a pseudo random code at the target, and storing the returned data in the OBC memory, this would then be download at a later date for correlation. However, developments of this system would be required, these include: same frequency receiver, switchable bandwidth filters to cope with large variations in Doppler shift, and modifications to increase the system gain. It is estimated these developments may require ¥500–1000 k and consequently it may prove lower risk and cost to procure and modify an existing altimeter, perhaps designed for an aircraft or UAV.
A. Phipps et al. / Acta Astronautica 59 (2006) 845 – 857 Table 7 Mission cost estimate Non recurring engineering Flight model AIV LEOP Ground segment
¥1, 212, 000 ¥8, 460, 000 ¥377, 000 ¥8, 633, 000 ¥543, 000
Total mission cost
¥19, 225, 000
The spacecraft will be target pointing during the encounter, the 0.8 m high gain antenna with a beamwidth of 13◦ at S-band will not drive the attitude pointing requirements. Making the assumptions stated below, there is a miniscule orbit deflection of 1.98 × 10−8 degrees for a typical NEO. Assumption: Fly-past velocity 10 km s−1 NEO mass assumed to be 5 × 1013 kg Fly-by distance 100 km High-rate communications (0.16×51 km error ellipse) The duration of time it takes to observe the orbit deflection largely depends upon the orbit geometry. If the deflected spacecraft appears largely in the crosstrack errors then it takes nearly one year to observe any appreciable orbit deflection. If the orbit geometry can be selected such that the deflection appears in the down-range errors then the deflection can be observed in less than a day—a substantially longer time would be needed to characterize the orbit better leading to a better mass estimate.
4. Mission cost Detailed cost estimates for the mission appear in Table 7. Assumptions made • Commercial figures (customer: SSTL). • 2003 prices (¥1.42 = £1.00). • Costing estimates include use of two existing groundstations. The ‘entry level’ mission can be developed, launched and operated for a total mission cost of less than ¥20
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million. Assuming full use of one of the 64 Mbyte memory during the encounter the mission cost per returned data bit is ¥0.04. The cost drivers being launch, propulsion, solar panel procurement, structure development, and communication and ranging system development, For a mission that includes the science instruments (outlined above) the mission cost increases to ¥25 million (this figure does not include the cost associated with the scientific analysis of the results).
5. Technology developments The mission enabling developments for this ‘entry level’ mission are listed • Development of a thrust tube structural concept. • Communications system: operation at low data rates will require significant modification to the existing equipment. The implementation of a 200 W high power amplifier at the dish must be traded against a larger 6 m antenna. In order to improve the low rate downlink the existing spacecraft s-band high power amplifier must be modified from 4.3 W RF output to 6 W. • Implementation of a pseudo random ranging system. • Software developments include: NEO imager search algorithm, software to perform closed loop control over thruster firings as well as development of a lost-in-space algorithm. The following list details the key enabling technologies required to lower mission cost and enhance the mission science return bit per Euro. • Propulsion Technology—a conventional monopropellant system cost is around ¥4 million a bipropellant system ¥6 million. The use of solar electric propulsion or HTP-kerosene bipropellant system (both being developed at SSTL) offers the opportunity to increase V or cut mission cost. • Modifications to the existing LEO qualified star imager to operate in deep space would reduce procurement cost. • Advanced communications coding and data compression algorithm.
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• Greater on-board storage capability through the use of a SSDR or hard disc drive. • Greater efficiency power system including the development of the direct energy transfer topology and the use of lithium-ion or lithium polymer storage technology. • Development of LEO imager technology for science gathering.
6. Conclusion An ‘entry level’ interplanetary platform has been conceptually designed to undertake a wide variety of missions specifically a NEO fly-past mission being studied. The study has established that the total mission cost is less than ¥20 million (FY2003) including platform, operations, payload and launch. A 10 kg science ‘reference’ package has been included which is able to return 64 Mbyte of science in its fleeting fly-past of the NEO target. A large number of trade-offs are possible, including the accommodation of up to 34 kg of payload mass. The propulsion system is sized to provide 1500 m s−1 for the specific target studied. However, substantial margin exists in the design allowing up to 1600 m s−1 V (using off-the-shelf monopropellant technology) or 2320 m s−1 V (using bipropellant technology). Consequently, the platform is capable of supporting lunar exploration missions [4,5] or multiple fly-bys of opportune targets. For the ‘entry level’ mission the principle cost drivers have been minimized. A non-optimal secondary launch into GSO has been selected and trajectory and environmental analysis suggests this leads to the lowest cost solution. Current off-the-shelf LEO platforms are unsuitable for the mission as they are insufficiently mass and volumetrically optimized, forcing the development of a new platform. This is based upon a central thrust tube for cost and risk minimization. Many LEO avionics are applicable for the mission, but the communication system (including groundstation) must be modified to support data rates down to 50 bps and a new pseudo random ranging system developed. The required accuracy of encounter is a cost driver requiring higher data rates and consequently larger antennas or higher power
Fig. 8. ‘Entry level’ NEO platform.
amplifiers. A further cost driver is GSO eclipse survival, this drives the power storage requirements up to 150 W h. The radiation environment surprisingly is not a significant driver (< 3 krad (Si)) with the use of COTS technology components and systems (Fig. 8).
Acknowledgements The authors are indebted to the following for their assistance: Luis Gomes, Mark Allery, Chris Jackson, Kevin Morgan, Neville Bean, Craig Clark, John Buckley and Guy Richardson, and Fred Kennedy USAF.
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A. Phipps et al. / Acta Astronautica 59 (2006) 845 – 857 [5] A. Phipps, et al., The Moon and Inner Planets—Getting there for Low Cost, 4th IAA International Conference on Low-Cost Planetary Missions, May 2–5, 2000. [6] J. Vette, AE/AP Trapped Particle Flux Maps 1966–1980, http://nssdc.gsfc.nasa.gov/space/model/magnetos/aeap.html, 25/04/02. [7] J. Miller, Phase III ranging System, http://www.amsat.org/ amsat/articles/g3ruh/123.html
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