- Email: [email protected]
Project 242: Fission fragments direct heating for space propulsion—Programme synthesis and applications to space exploration
Acta Astronautica 82 (2013) 153–158
Contents lists available at SciVerse ScienceDirect
Acta Astronautica journal homepage: www.elsevier.com/locate/actaastro
Project 242: Fission fragments direct heating for space propulsion—Programme synthesis and applications to space exploration M Augelli a,n, G F Bignami b,c, G Genta d a
Centre National d’Etudes Spatiales, 18 Avenue Edouard Belin, 31401 Toulouse, France Institute for Advanced Study of Pavia (IUSS), Lungo Ticino Sforza 56, 27100 Pavia, Italy Committee on Space Organisation (COSPAR - ICSU), 2 place Maurice Quentin, 75039 Paris Cedex 01, France d Politecnico di Torino, C. Duca degli Abruzzi 24, 10129 Torino, Italy b c
a r t i c l e i n f o
abstract
Article history: Received 20 December 2011 Received in revised form 11 March 2012 Accepted 3 April 2012 Available online 8 May 2012
The status and the main results achieved by Project 242 are presented. Project 242 is a programme (funded by ASI—1999/2002 from an idea of Carlo Rubbia) that studied a new concept of space propulsion motor by using direct conversion of the kinetic energy of fission fragments into increasing of enthalpy of a propellant gas. Project 242 studied the application of this propulsion system to a manned mission to Mars. Preliminary results were very satisfactory and it has been observed that a propulsion system with these characteristics could make the mission feasible. Results for other unmanned missions to the outer solar system are also presented. & 2012 Elsevier Ltd. All rights reserved.
Keywords: Nuclear propulsion Manned mission Human exploration Mars
1. Introduction In 1998 Carlo Rubbia proposed to the Italian Space Agency (ASI) and to the scientific community a new concept of propulsion based on the direct heating of a propellant gas by fission fragments generated by a fissile material. After an accurate internal evaluation, ASI decided to create a technical working group (WG) dedicated to a deep assessment study. The result of the working group was a high interest in the proposed concept and in possible space applications. The WG did not find any show stoppers and recommended to start a specific programme to pursue the development of the fission fragments propulsion system. This programme, funded by ASI, was started under the name Project 242 (P242) [1]. The objectives of the programme were to perform the detailed design study of the propulsion system and to pave the way to the realisation of a
n
Corresponding author. E-mail address: [email protected] (M. Augelli).
0094-5765/$ - see front matter & 2012 Elsevier Ltd. All rights reserved. http://dx.doi.org/10.1016/j.actaastro.2012.04.007
propulsion module prototype to be tested in a dedicated ground facility. The programme was funded from 1999 to 2002 when it was stopped due to changes in the ASI strategy and politics. In case an opportunity to restart the programme becomes available, the first part of the activities needs to be finalised. After this, the second part of the programme shall be started: the design and the construction of a new large test facility fully dedicated to this test. The third and last step of Project 242 should be the realisation of a flight propulsion rocket module. The scope of this paper is to recall the programme status at the time it has been put on-hold, to show the main achievements and to prepare the work to be done in the case the programme would be restarted. 2. The nuclear motor 2.1. The original idea A space propulsion system can be characterised by two main parameters: (1) the specific impulse and (2) the thrust. Present propulsive concepts can only assure
154
M Augelli et al. / Acta Astronautica 82 (2013) 153–158
systems with really high thrusts but low specific impulses (e.g. chemical propulsion, Isp ¼300–500 s) or systems with really high specific impulses but low thrust (e.g. electric propulsion, Isp up to 10,000 s). Some existing nuclear propulsion concepts like the Nuclear Engine for Rocket Vehicle Application (NERVA) are also limited to specific impulses not much higher than chemical propulsion, having some thermodynamics and gasdynamics similarities and limitations. Propulsion systems that are able to provide high specific impulses and at the same time relatively significant thrust could make possible large missions to deep space within an acceptable duration as well as human exploration to the Earth’s nearby planets. The original idea of Carlo Rubbia, that is now protected by some patent claims, was based on the principle that direct energy transfer from fission fragments (FF) to the propulsive gas is by far more efficient than using basic heat exchanges. A simple gasdynamic expansion through a nozzle delivers remarkable performances if compared to existing propulsion motors. This is essentially because a few kilograms of nuclear fuel are able to generate an amount of energy that is many times the one produced by the largest existing chemical rockets.
3.2. Neutron flux enhancement (from a diffuser to the n-hohlraum)
3. Main physics principles and design choices
Several studies [2] and some experiments at CERN (TARC) have shown that highly diffusive but neutron transparent media can be used to effectively store neutrons in a relatively small volume, confined by diffusion (n-hohlraum). Secondary neutrons from fission must be first thermalised, since at these energies the fission probability is the largest one. Therefore, in order to speed-up the thermalisation process, an appropriate diffusing medium of low atomic mass is chosen (high lethargy). The need to have at the same time the fissile fuel concentrated in very thin layers (in order to extract the most of FF energy in the gas) and surrounded by a relatively large volume of a high diffusive medium (in order to reach criticality in very thin layer) let us to individuate a reference configuration for making the engine working: cavity carved inside a volume filled with diffusing medium, highly transparent to neutrons (see a sketch in Fig. 1). The inner walls of the cavity are covered by a very thin layer of highly fissionable deposited material which acts as a neutron source. The neutron ‘‘ping–pong’’ between the cavity walls contributes to enhance the interaction probability, allowing the neutrons to cross many times the thin fuel layer, in such a way that a ultra-thin fissile layer acquires definitively a finite fission probability.
3.1. Direct heating by fission fragments
3.3. Choice of the fissionable fuel element
The concept is essentially to use directly the heavy ionising fragments produced in the nuclei fission to powerfully heat a selected propellant gas. About 88% of the nuclear energy due to fission is carried by two fission fragments, the rest being essentially released in the form of gamma rays and kinetic energy of the emitted neutrons ( E2.5 neutrons for 235U, E3.26 neutrons for 242mAm). The penetration range of FF is extremely short ( E10 mm in solid fuel) and usually energy is locally converted into heat inside the fuel. This means that in order to have a direct extraction of the FF from the fuel to heat directly the gas, an extraordinary thin layer of fuel (of the order of thousands atomic layers) is required and consequently a special neutron dynamics has to be assessed in order to ensure that even an ultra-thin fissile layer can reach criticality. Computation results [1] allow to state that the energy fraction that can be extracted as FF is about 34% for a fuel foil 1 mg/cm2 thick (1100 atomic layers) and falls to 24% for a 3 mg/cm2 layer thick. The fissionable layer will be arranged as a coating on the inner walls of the chamber’s trough which the propellant gas flows. The heating of the gas is performed by the ionisation energy loss of the FF which is emitted by the fissile layer. The specific ionisation losses in the gas are more than a factor 20 larger than in the fuel layer, since the speed of the FF is generally smaller than that of the inner electrons of the high Z layer. Thus, a modest thickness of gas (less than 0.5 mg/cm2) is sufficient to extract most of the energy from the FF. As a consequence even with very low gas pressure (few bars) the FF can be completely absorbed in a range of few centimetres.
Many potentially interesting Actinide elements have been surveyed, trying to identify the best choice according to the following criteria: (1) large fission cross section; (2) high fission probability; (3) acceptably long lifetime and (4) reasonable production procedure. On this basis, at least three elements appeared to be of interest for the application, namely in order of merit 242mAm, 251Cf and 239Pu. Pure 241Am grows in reactor spent fuel as a decay element of 241Pu (t1/2 ¼ 14.35 y), while 242mAm can be produced in suitable quantities by irradiating 241Am in an epi-thermal reactor, after isotopic separation. The irradiation has to be performed in a special neutral matrix, in order to suppress locally the neutron spectrum in the energy region where the 242mAm is fissionable, leaving the higher energy neutrons for which 241Am has a good capture cross section. In all aspects 242mAm seems to be the best candidate [1] as the fuel element and it has been selected. This is the reason of the programme name: Project 242. In parallel the multiplication factor for the n-Holhraum has been evaluated [1]. Computations show, as expected, that in practice the actual thickness of the layer should be significantly larger than the minimum critical thickness in order to make allowance for the fuel decay during burning. Also for this aspect 242mAm is largely better than 239Pu. 3.4. Propellant gas From a performance point of view, hydrogen is by far the best propellant (propulsive gas) since for a gasdynamic acceleration process (through an expansion nozzle) the specific impulse is inversely proportional to the
M Augelli et al. / Acta Astronautica 82 (2013) 153–158
Carbon Diffuser
155
Thin Fuel Layer (1 m of Am242m)
FF
FF
Empty volume
FF
FF
FF
FF
Fig. 1. n-Hohlraum concept.
square root of the molecular mass of the accelerated gas. So, the baseline propellant is hydrogen with helium as a possible backup if in the detailed conception phase it will be found that hydrogen creates a too severe environment for the solid parts. Also for nuclear processes, hydrogen as propellant gas is a very good choice. The selected motor specific impulse defines the stagnation temperature in the motor chamber. In turn the state equation for hydrogen will give the required specific enthalpy necessary to the gas to reach this temperature. In a perfect gas, enthalpy grows linearly with temperature. In a molecular gas there are however two main transitions, that are only slightly pressure dependent and that can contribute, if appropriately used, to the enthalpy balance. They are (1) molecular dissociation, occurring at about 4000 K, leading to the formation of atomic hydrogen and (2) spontaneous dissociation, starting from 10,000 K giving an incipient plasma. The energy transferred during the heating process to the gas allowing those transitions can be recovered in the nozzle during the expansion process (where the gas is cooling) only if the corresponding recombination is fast enough. As a first result [3], we can state that in most cases plasma recombination is really fast, but molecular recombination is not. So, for performance estimation, an expanded gas with plasma recombination and without molecular recombination was taken into account. Isp reduction is of the order of a factor 1.2 wrt equilibrium.
counteracting the heating process. This emission is composed of line radiation plus a continuum radiation. The first component comes from transitions between quantum levels (line spectrum). The second one is generated by transitions from levels to continuum and free–free transitions (plasma). One can verify that line radiation is globally absorbed in the gas and it appears as an additional term to the heat conductibility. On the contrary, for the continuum radiation, the gas is found to be a relatively transparent medium, so this component is transferred back to the walls. In addition radiative power is a very fast growing function of the temperature above 10,000 K, mainly due to the growing population of free electrons. More in detail, some computations have been performed in a P242 representative configuration heating a flowing gas with a constant specific power, and it has been seen that above 9500 K (corresponding to onset of plasma associated radiation) the required power to increase the temperature is diverging, if radiative losses are considered [3]. This prescribes a real limitation to the maximum attainable gas temperature and subsequently to the deliverable specific impulse because, in order to reduce the effect of these power losses, the heating specific power must be increased. Thus, as explained before, a thinner layer of gas propellant is heated very fast. This reduces the specific thickness of the gas under heating and makes it transparent to line radiation. As a consequence, radiative losses increase again.
3.5. Losses in the hot hydrogen
3.6. Motor configuration
During heating, by the fact that temperature is raising, the hot hydrogen starts to emit strong radiation eventually
Based on preliminary studies and on a first optimisation for the reference mission described later, a proposed
156
M Augelli et al. / Acta Astronautica 82 (2013) 153–158
Fig. 2. Layout of the proposed motor configuration.
motor configuration has been defined. It consists of a certain number of tubes packed together. Each tube is coated with a thin fissionable layer and opened at one end to let the gas to escape. The tubes are located inside a neutron reflecting structure (made, for example, of C or BeO), which ensures neutron containment and criticality. Some control rods are necessary to control the neutron multiplication. They are inserted in the reflector. A cooling system to keep the americium foil and the neutron diffuser to the operating temperature could flow around the tubes and transfer the produced heat to a radiating panel. A schematic layout is provided in Fig. 2 also with the baseline design values (mainly Isp of 2700 s and thrust of 3200 N). The total thermal power is 230 MW. For this configuration the propulsion power is around 43 MW and the thermal power to be dissipated is around 190 MW, giving a power radiator mass of about 3 t. Some fluid-dynamics analyses have been performed on a single FF-heated tube. By the fact that the heating process is almost uniform along the radius, it shall be more convenient to have propellant gas flowing from the wall to the tube centreline in a radial direction. A technological application of carbon–carbon composite materials could be a solution to provide a porous wall. Carbon–carbon is also a good material for neutronics. The full fluid-dynamics calculations, including FF heating and radiation, show [3] that even taking into account the back-heating towards the walls due to the gas conductivity, the maximum temperature inside the tube is ever reached on the axis (along the radius before reaching the axis) with a quasi uniform net power distribution.
consideration a reference mission for the human exploration of Mars. 4.1. Reference mission As already emphasised, the novelty of this kind of propulsion system (high Isp and good thrust level) allows to imagine a manned mission to Mars along new lines because higher velocity changes are possible. The main hypotheses of this mission are:
Use of existing (or already in development) launch
systems. It is considered not realistic to imagine the development of a new gigantic launcher (like Saturn V for the Apollo programme) to reach the target orbit in a single launch. Modular approach with automatic or assisted in-orbit integration (e.g. in the proximity of ISS using robotic arms and Extra Vehicle Activities) to build the interplanetary ships, with the possibility to use a multiple spacecrafts approach: separate cargo and crew. Overall mission time duration for crew of the order of one year. This is considered a limit to ensure a good human survival probability. In particular for the risk of radiation exposure. All the manoeuvres are performed using the FF heated motor. No need of risky aerocapture and/or aerobraking manoeuvres. All the necessary propellant to return to Earth is sent from Earth. In-situ production or extraction, on the Mars surface, to prepare the propellant for return is avoided.
4. Application to space missions During the first optimisation loop, the preliminary design characteristics of the motor were defined taking into
The computed mission consists of three ships assembled in orbit: 2 cargo (TUG 1&2) and 1 manned (Clipper). 20 launches of existing heavy launch vehicles (STS and its
M Augelli et al. / Acta Astronautica 82 (2013) 153–158
replacement, Ariane5, Delta4, etc.) will be necessary to ferry into LEO the needed modules to assemble the ‘‘fleet’’. Cargo spaceships are sent before on a slow transit trajectory. They contain everything necessary to allow the crew Mars descent, in orbit ascent and safe return to Earth. When the cargo load has reached its final position in Low Mars Orbit (LMO), the ‘‘GO!’’ to leave LEO is given to the ‘‘clipper’’ spaceship that will travel on a fast transit (high energy) trajectory. The crew will reside on Mars for a period of the order of a month, and then ascend to LMO and fast-track back to Earth. Total LEO-to-LEO time is about one year. From Mars, the reusable part of the cargo spaceships return to Earth. Once everything is back in LEO, possibly docked to the ISS, after refurbishment and maintenance, it is ready to be re-used for the next astronomical window. The relevant trajectories are presented in Fig. 3. How this mission is attainable with the FF heated motor is well explained in Fig. 4. Orbital mechanics allow only two kinds of opportunities between Earth and Mars, but more generally between two orbits with common focus: conjunction and opposition. As shown in Fig. 4a it is evidenced that if we want to have crew residing for a short time on Mars, we have to achieve high velocity changes. With the existing propulsion technologies (chemical propulsion) the opposition
157
opportunities are not accessible, so the mission is not possible (in Fig. 4b chemical propulsion is logarithmically out of the plot scale). With a standard nuclear thermal reactor like Nerva, only the low energy–long stay class of transfer orbits could be reached, but with hard limitations: aerocapture and aerobraking on Mars and on Earth, in-situ production of propellant on Mars for the crew to return to Earth, and seating of crew on a bomb-grade motor with high neutron fluxes (E3.2 1019 n/s). This motor concept is very effectively shielded. The nuclear radiation produced during the functioning is sufficiently small to ensure the safety of the crew even for long missions. Fig. 4b shows also that reasonable high thrust levels could dramatically reduce propellant consumption (reduction of gravity losses). The thrust is the whole motor thrust (3200 N for the baseline configuration). Table 1 shows a synthetic mass budget for the mission with the option of TUG motors to return to Earth. 4.2. Other possible missions to the outer solar system Without performing a detailed mission analysis nor a new optimisation of the motor parameters, it could be interesting to see how the present propulsion system
Fig. 3. Trajectories for the TUG and Clipper spaceships.
Fig. 4. (a) Opposition and conjunction opportunities (Mars resident time vs. velocity change). (b) Propellant consumption as a function of Isp and Thrust.
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M Augelli et al. / Acta Astronautica 82 (2013) 153–158
Table 1 Mass budget of the mission including return to Earth of TUG engines. Trip
Ship
MLEO (t)
Complete mission including Outbound transfer 1 Tug1 142 2 Tug2 245 3 Clipper 147 Return 1 2 3
transfer Tug1 Tug2 Clipper
31 33 70
MLMO (t)
LH2 (t)
Duration (days)
Missions to the outer solar system Saturn 2309 (6.3 yr) Saturn’s satellites 2348 (6.4 yr) Jupiter 1066 (2.9 yr) Jupiter’s Galilean Satellites 1099 (3 yr) Uranus 5853 (16 yr)
Orbit
TUG engines return 102 176 92
40 69 55
19 31 15
Slow transit (Hohmann) Slow transit (Hohmann) Fast transit
42 45 162
11 12 92
5 6 25
Slow transit (Hohmann) Slow transit (Hohmann) Fast transit
Table 2 Results of missions to the outer solar system for automatic exploration/ science. Departure is from LEO to a low altitude orbit around the target celestial body. Mass in LEO can be up to 20 t. Target
Engine on (d)
DV (km/s)
MLEO/MTGT
15.24 8.20 18.75 7.09 15.61
3 1.8 4 1.7 3.1
could make easier and more realistic automatic missions to the outer solar system. In Table 2 some potentially interesting missions to the outer solar system are presented. The comparison with in progress or past missions having similar targets can easily show the advantage of the proposed propulsive system. 5. Conclusions Missions to explore the solar system demand high performances to the on-board propulsion systems. Current
available motors are very limited in terms of power, thus are not able to deliver a high specific impulse with a relatively acceptable thrust level. This is the main requirement to transform a dream into a real mission. In particular if manned. Project 242 has started evaluating the potentiality of a fission fragment heated motor and its application to an ambitious human exploration of Mars, starting its design. The programme is on-hold since 2002. If the opportunity to restart the programme would become available, once consolidated the obtained results and finalised the started activities, two main cornerstones are awaiting along the path: (1) the design and the construction of a new large ground test facility and (2) the realisation of a flight propulsion rocket module. References [1] M. Augelli, et al., Report of the Working Group on a Preliminary Assessment of a New Fission Fragment Heated Propulsion Concept and its Applicability to manned missions to the planet Mars (Project 242)—ASI—Rome, 15 March, 1999. [2] C. Rubbia, Neutrons in a Highly Diffusive Medium: A New Propulsion Tool for Deep Space Exploration? CERN Colloquium (27.08.1998). [3] C. Rubbia, Fission Fragments Heating for space Propulsion—CERN Technical Note CERN SL-Note—2000-036. EET.
Contents lists available at SciVerse ScienceDirect
Acta Astronautica journal homepage: www.elsevier.com/locate/actaastro
Project 242: Fission fragments direct heating for space propulsion—Programme synthesis and applications to space exploration M Augelli a,n, G F Bignami b,c, G Genta d a
Centre National d’Etudes Spatiales, 18 Avenue Edouard Belin, 31401 Toulouse, France Institute for Advanced Study of Pavia (IUSS), Lungo Ticino Sforza 56, 27100 Pavia, Italy Committee on Space Organisation (COSPAR - ICSU), 2 place Maurice Quentin, 75039 Paris Cedex 01, France d Politecnico di Torino, C. Duca degli Abruzzi 24, 10129 Torino, Italy b c
a r t i c l e i n f o
abstract
Article history: Received 20 December 2011 Received in revised form 11 March 2012 Accepted 3 April 2012 Available online 8 May 2012
The status and the main results achieved by Project 242 are presented. Project 242 is a programme (funded by ASI—1999/2002 from an idea of Carlo Rubbia) that studied a new concept of space propulsion motor by using direct conversion of the kinetic energy of fission fragments into increasing of enthalpy of a propellant gas. Project 242 studied the application of this propulsion system to a manned mission to Mars. Preliminary results were very satisfactory and it has been observed that a propulsion system with these characteristics could make the mission feasible. Results for other unmanned missions to the outer solar system are also presented. & 2012 Elsevier Ltd. All rights reserved.
Keywords: Nuclear propulsion Manned mission Human exploration Mars
1. Introduction In 1998 Carlo Rubbia proposed to the Italian Space Agency (ASI) and to the scientific community a new concept of propulsion based on the direct heating of a propellant gas by fission fragments generated by a fissile material. After an accurate internal evaluation, ASI decided to create a technical working group (WG) dedicated to a deep assessment study. The result of the working group was a high interest in the proposed concept and in possible space applications. The WG did not find any show stoppers and recommended to start a specific programme to pursue the development of the fission fragments propulsion system. This programme, funded by ASI, was started under the name Project 242 (P242) [1]. The objectives of the programme were to perform the detailed design study of the propulsion system and to pave the way to the realisation of a
n
Corresponding author. E-mail address: [email protected] (M. Augelli).
0094-5765/$ - see front matter & 2012 Elsevier Ltd. All rights reserved. http://dx.doi.org/10.1016/j.actaastro.2012.04.007
propulsion module prototype to be tested in a dedicated ground facility. The programme was funded from 1999 to 2002 when it was stopped due to changes in the ASI strategy and politics. In case an opportunity to restart the programme becomes available, the first part of the activities needs to be finalised. After this, the second part of the programme shall be started: the design and the construction of a new large test facility fully dedicated to this test. The third and last step of Project 242 should be the realisation of a flight propulsion rocket module. The scope of this paper is to recall the programme status at the time it has been put on-hold, to show the main achievements and to prepare the work to be done in the case the programme would be restarted. 2. The nuclear motor 2.1. The original idea A space propulsion system can be characterised by two main parameters: (1) the specific impulse and (2) the thrust. Present propulsive concepts can only assure
154
M Augelli et al. / Acta Astronautica 82 (2013) 153–158
systems with really high thrusts but low specific impulses (e.g. chemical propulsion, Isp ¼300–500 s) or systems with really high specific impulses but low thrust (e.g. electric propulsion, Isp up to 10,000 s). Some existing nuclear propulsion concepts like the Nuclear Engine for Rocket Vehicle Application (NERVA) are also limited to specific impulses not much higher than chemical propulsion, having some thermodynamics and gasdynamics similarities and limitations. Propulsion systems that are able to provide high specific impulses and at the same time relatively significant thrust could make possible large missions to deep space within an acceptable duration as well as human exploration to the Earth’s nearby planets. The original idea of Carlo Rubbia, that is now protected by some patent claims, was based on the principle that direct energy transfer from fission fragments (FF) to the propulsive gas is by far more efficient than using basic heat exchanges. A simple gasdynamic expansion through a nozzle delivers remarkable performances if compared to existing propulsion motors. This is essentially because a few kilograms of nuclear fuel are able to generate an amount of energy that is many times the one produced by the largest existing chemical rockets.
3.2. Neutron flux enhancement (from a diffuser to the n-hohlraum)
3. Main physics principles and design choices
Several studies [2] and some experiments at CERN (TARC) have shown that highly diffusive but neutron transparent media can be used to effectively store neutrons in a relatively small volume, confined by diffusion (n-hohlraum). Secondary neutrons from fission must be first thermalised, since at these energies the fission probability is the largest one. Therefore, in order to speed-up the thermalisation process, an appropriate diffusing medium of low atomic mass is chosen (high lethargy). The need to have at the same time the fissile fuel concentrated in very thin layers (in order to extract the most of FF energy in the gas) and surrounded by a relatively large volume of a high diffusive medium (in order to reach criticality in very thin layer) let us to individuate a reference configuration for making the engine working: cavity carved inside a volume filled with diffusing medium, highly transparent to neutrons (see a sketch in Fig. 1). The inner walls of the cavity are covered by a very thin layer of highly fissionable deposited material which acts as a neutron source. The neutron ‘‘ping–pong’’ between the cavity walls contributes to enhance the interaction probability, allowing the neutrons to cross many times the thin fuel layer, in such a way that a ultra-thin fissile layer acquires definitively a finite fission probability.
3.1. Direct heating by fission fragments
3.3. Choice of the fissionable fuel element
The concept is essentially to use directly the heavy ionising fragments produced in the nuclei fission to powerfully heat a selected propellant gas. About 88% of the nuclear energy due to fission is carried by two fission fragments, the rest being essentially released in the form of gamma rays and kinetic energy of the emitted neutrons ( E2.5 neutrons for 235U, E3.26 neutrons for 242mAm). The penetration range of FF is extremely short ( E10 mm in solid fuel) and usually energy is locally converted into heat inside the fuel. This means that in order to have a direct extraction of the FF from the fuel to heat directly the gas, an extraordinary thin layer of fuel (of the order of thousands atomic layers) is required and consequently a special neutron dynamics has to be assessed in order to ensure that even an ultra-thin fissile layer can reach criticality. Computation results [1] allow to state that the energy fraction that can be extracted as FF is about 34% for a fuel foil 1 mg/cm2 thick (1100 atomic layers) and falls to 24% for a 3 mg/cm2 layer thick. The fissionable layer will be arranged as a coating on the inner walls of the chamber’s trough which the propellant gas flows. The heating of the gas is performed by the ionisation energy loss of the FF which is emitted by the fissile layer. The specific ionisation losses in the gas are more than a factor 20 larger than in the fuel layer, since the speed of the FF is generally smaller than that of the inner electrons of the high Z layer. Thus, a modest thickness of gas (less than 0.5 mg/cm2) is sufficient to extract most of the energy from the FF. As a consequence even with very low gas pressure (few bars) the FF can be completely absorbed in a range of few centimetres.
Many potentially interesting Actinide elements have been surveyed, trying to identify the best choice according to the following criteria: (1) large fission cross section; (2) high fission probability; (3) acceptably long lifetime and (4) reasonable production procedure. On this basis, at least three elements appeared to be of interest for the application, namely in order of merit 242mAm, 251Cf and 239Pu. Pure 241Am grows in reactor spent fuel as a decay element of 241Pu (t1/2 ¼ 14.35 y), while 242mAm can be produced in suitable quantities by irradiating 241Am in an epi-thermal reactor, after isotopic separation. The irradiation has to be performed in a special neutral matrix, in order to suppress locally the neutron spectrum in the energy region where the 242mAm is fissionable, leaving the higher energy neutrons for which 241Am has a good capture cross section. In all aspects 242mAm seems to be the best candidate [1] as the fuel element and it has been selected. This is the reason of the programme name: Project 242. In parallel the multiplication factor for the n-Holhraum has been evaluated [1]. Computations show, as expected, that in practice the actual thickness of the layer should be significantly larger than the minimum critical thickness in order to make allowance for the fuel decay during burning. Also for this aspect 242mAm is largely better than 239Pu. 3.4. Propellant gas From a performance point of view, hydrogen is by far the best propellant (propulsive gas) since for a gasdynamic acceleration process (through an expansion nozzle) the specific impulse is inversely proportional to the
M Augelli et al. / Acta Astronautica 82 (2013) 153–158
Carbon Diffuser
155
Thin Fuel Layer (1 m of Am242m)
FF
FF
Empty volume
FF
FF
FF
FF
Fig. 1. n-Hohlraum concept.
square root of the molecular mass of the accelerated gas. So, the baseline propellant is hydrogen with helium as a possible backup if in the detailed conception phase it will be found that hydrogen creates a too severe environment for the solid parts. Also for nuclear processes, hydrogen as propellant gas is a very good choice. The selected motor specific impulse defines the stagnation temperature in the motor chamber. In turn the state equation for hydrogen will give the required specific enthalpy necessary to the gas to reach this temperature. In a perfect gas, enthalpy grows linearly with temperature. In a molecular gas there are however two main transitions, that are only slightly pressure dependent and that can contribute, if appropriately used, to the enthalpy balance. They are (1) molecular dissociation, occurring at about 4000 K, leading to the formation of atomic hydrogen and (2) spontaneous dissociation, starting from 10,000 K giving an incipient plasma. The energy transferred during the heating process to the gas allowing those transitions can be recovered in the nozzle during the expansion process (where the gas is cooling) only if the corresponding recombination is fast enough. As a first result [3], we can state that in most cases plasma recombination is really fast, but molecular recombination is not. So, for performance estimation, an expanded gas with plasma recombination and without molecular recombination was taken into account. Isp reduction is of the order of a factor 1.2 wrt equilibrium.
counteracting the heating process. This emission is composed of line radiation plus a continuum radiation. The first component comes from transitions between quantum levels (line spectrum). The second one is generated by transitions from levels to continuum and free–free transitions (plasma). One can verify that line radiation is globally absorbed in the gas and it appears as an additional term to the heat conductibility. On the contrary, for the continuum radiation, the gas is found to be a relatively transparent medium, so this component is transferred back to the walls. In addition radiative power is a very fast growing function of the temperature above 10,000 K, mainly due to the growing population of free electrons. More in detail, some computations have been performed in a P242 representative configuration heating a flowing gas with a constant specific power, and it has been seen that above 9500 K (corresponding to onset of plasma associated radiation) the required power to increase the temperature is diverging, if radiative losses are considered [3]. This prescribes a real limitation to the maximum attainable gas temperature and subsequently to the deliverable specific impulse because, in order to reduce the effect of these power losses, the heating specific power must be increased. Thus, as explained before, a thinner layer of gas propellant is heated very fast. This reduces the specific thickness of the gas under heating and makes it transparent to line radiation. As a consequence, radiative losses increase again.
3.5. Losses in the hot hydrogen
3.6. Motor configuration
During heating, by the fact that temperature is raising, the hot hydrogen starts to emit strong radiation eventually
Based on preliminary studies and on a first optimisation for the reference mission described later, a proposed
156
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Fig. 2. Layout of the proposed motor configuration.
motor configuration has been defined. It consists of a certain number of tubes packed together. Each tube is coated with a thin fissionable layer and opened at one end to let the gas to escape. The tubes are located inside a neutron reflecting structure (made, for example, of C or BeO), which ensures neutron containment and criticality. Some control rods are necessary to control the neutron multiplication. They are inserted in the reflector. A cooling system to keep the americium foil and the neutron diffuser to the operating temperature could flow around the tubes and transfer the produced heat to a radiating panel. A schematic layout is provided in Fig. 2 also with the baseline design values (mainly Isp of 2700 s and thrust of 3200 N). The total thermal power is 230 MW. For this configuration the propulsion power is around 43 MW and the thermal power to be dissipated is around 190 MW, giving a power radiator mass of about 3 t. Some fluid-dynamics analyses have been performed on a single FF-heated tube. By the fact that the heating process is almost uniform along the radius, it shall be more convenient to have propellant gas flowing from the wall to the tube centreline in a radial direction. A technological application of carbon–carbon composite materials could be a solution to provide a porous wall. Carbon–carbon is also a good material for neutronics. The full fluid-dynamics calculations, including FF heating and radiation, show [3] that even taking into account the back-heating towards the walls due to the gas conductivity, the maximum temperature inside the tube is ever reached on the axis (along the radius before reaching the axis) with a quasi uniform net power distribution.
consideration a reference mission for the human exploration of Mars. 4.1. Reference mission As already emphasised, the novelty of this kind of propulsion system (high Isp and good thrust level) allows to imagine a manned mission to Mars along new lines because higher velocity changes are possible. The main hypotheses of this mission are:
Use of existing (or already in development) launch
systems. It is considered not realistic to imagine the development of a new gigantic launcher (like Saturn V for the Apollo programme) to reach the target orbit in a single launch. Modular approach with automatic or assisted in-orbit integration (e.g. in the proximity of ISS using robotic arms and Extra Vehicle Activities) to build the interplanetary ships, with the possibility to use a multiple spacecrafts approach: separate cargo and crew. Overall mission time duration for crew of the order of one year. This is considered a limit to ensure a good human survival probability. In particular for the risk of radiation exposure. All the manoeuvres are performed using the FF heated motor. No need of risky aerocapture and/or aerobraking manoeuvres. All the necessary propellant to return to Earth is sent from Earth. In-situ production or extraction, on the Mars surface, to prepare the propellant for return is avoided.
4. Application to space missions During the first optimisation loop, the preliminary design characteristics of the motor were defined taking into
The computed mission consists of three ships assembled in orbit: 2 cargo (TUG 1&2) and 1 manned (Clipper). 20 launches of existing heavy launch vehicles (STS and its
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replacement, Ariane5, Delta4, etc.) will be necessary to ferry into LEO the needed modules to assemble the ‘‘fleet’’. Cargo spaceships are sent before on a slow transit trajectory. They contain everything necessary to allow the crew Mars descent, in orbit ascent and safe return to Earth. When the cargo load has reached its final position in Low Mars Orbit (LMO), the ‘‘GO!’’ to leave LEO is given to the ‘‘clipper’’ spaceship that will travel on a fast transit (high energy) trajectory. The crew will reside on Mars for a period of the order of a month, and then ascend to LMO and fast-track back to Earth. Total LEO-to-LEO time is about one year. From Mars, the reusable part of the cargo spaceships return to Earth. Once everything is back in LEO, possibly docked to the ISS, after refurbishment and maintenance, it is ready to be re-used for the next astronomical window. The relevant trajectories are presented in Fig. 3. How this mission is attainable with the FF heated motor is well explained in Fig. 4. Orbital mechanics allow only two kinds of opportunities between Earth and Mars, but more generally between two orbits with common focus: conjunction and opposition. As shown in Fig. 4a it is evidenced that if we want to have crew residing for a short time on Mars, we have to achieve high velocity changes. With the existing propulsion technologies (chemical propulsion) the opposition
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opportunities are not accessible, so the mission is not possible (in Fig. 4b chemical propulsion is logarithmically out of the plot scale). With a standard nuclear thermal reactor like Nerva, only the low energy–long stay class of transfer orbits could be reached, but with hard limitations: aerocapture and aerobraking on Mars and on Earth, in-situ production of propellant on Mars for the crew to return to Earth, and seating of crew on a bomb-grade motor with high neutron fluxes (E3.2 1019 n/s). This motor concept is very effectively shielded. The nuclear radiation produced during the functioning is sufficiently small to ensure the safety of the crew even for long missions. Fig. 4b shows also that reasonable high thrust levels could dramatically reduce propellant consumption (reduction of gravity losses). The thrust is the whole motor thrust (3200 N for the baseline configuration). Table 1 shows a synthetic mass budget for the mission with the option of TUG motors to return to Earth. 4.2. Other possible missions to the outer solar system Without performing a detailed mission analysis nor a new optimisation of the motor parameters, it could be interesting to see how the present propulsion system
Fig. 3. Trajectories for the TUG and Clipper spaceships.
Fig. 4. (a) Opposition and conjunction opportunities (Mars resident time vs. velocity change). (b) Propellant consumption as a function of Isp and Thrust.
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Table 1 Mass budget of the mission including return to Earth of TUG engines. Trip
Ship
MLEO (t)
Complete mission including Outbound transfer 1 Tug1 142 2 Tug2 245 3 Clipper 147 Return 1 2 3
transfer Tug1 Tug2 Clipper
31 33 70
MLMO (t)
LH2 (t)
Duration (days)
Missions to the outer solar system Saturn 2309 (6.3 yr) Saturn’s satellites 2348 (6.4 yr) Jupiter 1066 (2.9 yr) Jupiter’s Galilean Satellites 1099 (3 yr) Uranus 5853 (16 yr)
Orbit
TUG engines return 102 176 92
40 69 55
19 31 15
Slow transit (Hohmann) Slow transit (Hohmann) Fast transit
42 45 162
11 12 92
5 6 25
Slow transit (Hohmann) Slow transit (Hohmann) Fast transit
Table 2 Results of missions to the outer solar system for automatic exploration/ science. Departure is from LEO to a low altitude orbit around the target celestial body. Mass in LEO can be up to 20 t. Target
Engine on (d)
DV (km/s)
MLEO/MTGT
15.24 8.20 18.75 7.09 15.61
3 1.8 4 1.7 3.1
could make easier and more realistic automatic missions to the outer solar system. In Table 2 some potentially interesting missions to the outer solar system are presented. The comparison with in progress or past missions having similar targets can easily show the advantage of the proposed propulsive system. 5. Conclusions Missions to explore the solar system demand high performances to the on-board propulsion systems. Current
available motors are very limited in terms of power, thus are not able to deliver a high specific impulse with a relatively acceptable thrust level. This is the main requirement to transform a dream into a real mission. In particular if manned. Project 242 has started evaluating the potentiality of a fission fragment heated motor and its application to an ambitious human exploration of Mars, starting its design. The programme is on-hold since 2002. If the opportunity to restart the programme would become available, once consolidated the obtained results and finalised the started activities, two main cornerstones are awaiting along the path: (1) the design and the construction of a new large ground test facility and (2) the realisation of a flight propulsion rocket module. References [1] M. Augelli, et al., Report of the Working Group on a Preliminary Assessment of a New Fission Fragment Heated Propulsion Concept and its Applicability to manned missions to the planet Mars (Project 242)—ASI—Rome, 15 March, 1999. [2] C. Rubbia, Neutrons in a Highly Diffusive Medium: A New Propulsion Tool for Deep Space Exploration? CERN Colloquium (27.08.1998). [3] C. Rubbia, Fission Fragments Heating for space Propulsion—CERN Technical Note CERN SL-Note—2000-036. EET.