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A high-energy sample return Earth re-entry demonstrator to address planetary protection issues
Acta Astronautica 61 (2007) 198 – 202 www.elsevier.com/locate/actaastro
A high-energy sample return Earth re-entry demonstrator to address planetary protection issues Andy Phippsa,∗ , Alex da Silva Curiela , Martin Sweetinga , Arthur Smithb , Steve Lingardc a Surrey Satellite Technology Ltd, Surrey Space Centre, Guildford, Surrey GU2 7YE, UK b Fluid Gravity Engineering Ltd, 83 Market St, St Andrews, Fife, KY16 9NX, UK c Vorticity Ltd, Hampden House, Monument Business Park, Chalgrove, Oxfordshire, OX44 7RW UK
Available online 26 March 2007
Abstract For many of the proposed planetary exploration missions, robotically collected sample materials must be returned to the Earth for more detailed analysis. For the more interesting locations in the solar system, full planetary protection guidelines must be taken into account to prevent cross contamination. So far, few missions have demonstrated re-entry from interplanetary space, which requires a considerably greater velocity change than the more common Earth orbit re-entry missions, and so far none have dealt with the planetary protection issues. Missions such as the proposed Mars Sample Return mission must depend on reliable systems to carry out a sample return. Such a complex safety critical system can of course not be used for the first time on a real mission, nor can it be tested under Earth laboratory conditions. This paper proposes a design concept for an European Space Agency testbed to help validate the necessary technology. The spacecraft will carry scientific instruments to help characterise the environment, so that subsequent missions can rely on this. © 2007 Elsevier Ltd. All rights reserved.
1. Introduction
1.1. Spacecraft concept
The proposed mission is launched into the Earth orbit, and it carries an upper stage to propel it to interplanetary speeds, before being re-entered into the Earth’s atmosphere. The proposed re-entry probe weighs just 45 kg, and carries a sealed container with up to 1 kg sample material. Design trades related to the probe design and the recovery operation is discussed. The paper concludes that the demonstration mission will allow many issues and risks in sample return missions to be rehearsed and retired.
Trades have identified that the Earth-entry Vehicle Demonstrator is best implemented as a two stage spacecraft, providing a cost efficient solution (Fig. 1). The Hyperbolic Re-entry Probe (HRP) simulates the sample return vehicle returning from Mars, and a Carrier and Delivery System (CDS) which accelerates the probe and delivers it in an Earth-bound trajectory simulating that for a vehicle returning from Mars. Due to the complexity and criticality of the vehicle operations, operations from the European Space Operations Centre are baselined. 1.2. Trajectory analysis
∗ Corresponding author. Tel.: +44 1483 803803.
E-mail address: [email protected] (A. Phipps). 0094-5765/$ - see front matter © 2007 Elsevier Ltd. All rights reserved. doi:10.1016/j.actaastro.2007.01.015
Due to the Earth–Mars relative positions, suitable launch windows for efficient low-energy Mars transfers
A. Phipps et al. / Acta Astronautica 61 (2007) 198 – 202
199
Fig. 1. Earth Vehicle Demonstrator.
only present themselves approximately every two years. Based on an assessment of typical launch dates of potential sample return missions, Earth entry velocities from 12 to 12.8 km/s need to be considered. The design has therefore taken into account the worst case figure of 12.8 km/s. In order to demonstrate such a re-entry event, a spacecraft can be launched, and accelerated to re-enter at such a velocity. The choice of an initial retrograde orbit offers the advantage that less propellant needs to be carried, as the Earth rotational velocity effectively increases the relative re-entry velocity. For a prograde orbit, an orbital velocity of 13.3 km/s would need to be achieved, as opposed to an orbital velocity of 12.3 km/s for a retrograde orbit. Starting orbits of GTO were considered, as this provides ample opportunities as a piggyback or secondary passenger, but such orbits are prograde and would lead to a larger and more expensive mission. A dedicated retrograde Highly Elliptical Orbit was therefore considered as the baseline instead. A zero degree inclination was selected to provide contingency and multiple opportunities in time to reach the same re-entry site on the equator. Once in a highly elliptical orbit, the apogee is raised further in a stepwise fashion using spacecraft engines. This increases the perigee velocity, and provides schedule margin for the re-entry event planning. It also ensures the spacecraft spends ample time at apogee for the final major re-entry burn, which minimises gravity losses for this burn. The re-entry event itself commences with a final firing of the spacecraft engines at apogee. This places the spacecraft composite into a hyperbolic trajectory that enters the Earth atmosphere, for entry over the equator. During the descent, the spacecraft trajectory is closely tracked, and final corrections are made at a point
Fig. 2. The 300,000 km apogee orbit. (1) apogee raising manoeuvres, (2) apogee re-entry burn.
that balances accuracy of the navigation solution, leaving sufficient time for the corrections to have efficient results. At this point the Hyperbolic Return Vehicle is separated from the Carrier vehicle, and they re-enter on slightly different trajectories, with the aim to burn up the Carrier vehicle in the atmosphere. Additional factors that have been considered in the trajectory trades include • exposure of the spacecraft to the radiation belts; • effects of gravitational perturbations by the moon; • sensitivity to thruster alignment, timing and magnitude errors; • ease and accuracy of obtaining navigation information leading up to the re-entry event. As a result, a 300,000 km apogee was selected for the mission, with the sequence illustrated in Figs. 2 and 3. This results in a two-day timeline from the re-entry burn until the entry into the Earth atmosphere.
200
A. Phipps et al. / Acta Astronautica 61 (2007) 198 – 202
1.3. Launcher selection A number of launchers have been considered where compatible with a low cost launch (Table 1). For the mission baseline, a dedicated PSLV was selected, providing the best compromise between availability, control over precise launch parameters and cost. 1.4. Landing site selection Several landing sites were considered, including land, ocean and air retrieval. Near equatorial landing sites provide the lowest energy requirements for the spacecraft, but appear to be restricted to a water landing when taking into account the complexities and costs of sample retrieval. Although a water landing is attractive in terms of the design of the vehicle for terrestrial impact, the risk of damage to the sample container due to a
Fig. 4. Final re-entry of the mission components.
hard landing on debris, ships or even sea mammals in the ocean cannot be accepted, hence design for a hard landing is an absolute necessity. Northern Canada was considered, but non-equatorial sites would require extra propellant to be carried. The Australian desert does not require significant propellant to be carried, and the Woomera test range was finally considered as a suitable retrieval site for the purposes of the study. This site also permits the Carrier and Delivery Vehicle to be landed in the ocean (Fig. 4). Based on a worst case launch into Geostationary Transfer Orbit, the total propellant requirements are set by the necessary velocity increments of up to 3.7 km/s (Table 2). 2. Spacecraft design
Fig. 3. Point (3) Region of final acceleration thrust.
The launch mass of the spacecraft composite is 950 kg, of which over 70% accounts for the propellant carried. The Hyperbolic Entry Probe itself weighs 46 kg, with a 5 kg sample canister payload. The EVD vehicle is spin stabilised during all propulsive firing events, and de-spun to lower spin rates just
Table 1 Launcher Option
Mass capability
Heritage
Cost (USD)
PSLV
∼ 1050 kg into GTO
∼ $25 million
Soyuz-Fregat Ariane 5 Sylda Ariane 5 Cyclade
1350 kg into GTO > 1000 kg to GTO Soft limit of 700 kg to GTO
8 launches since 1993. First GTO launch in 2003 Yes Developed for Smart-1 mission
$35–$50 million
¥35 million ∼ ¥11 million for 350 kg platform
A. Phipps et al. / Acta Astronautica 61 (2007) 198 – 202 Table 2 Delta-V requirements Initial orbit
Trajectory Apogee raising DV Apogee DV Acceleration DV CDS avoidance DV Total Delta-V
201
weeks, and that some COTS-based avionics of SSTL’s Giove-A (in MEO orbit) heritage could be readily used. PSLV GTO (I = 17.7◦ ) Minimum value (m/s)
Including 10% margin (m/s)
666.1 387.0 2290.9 TBD 3344.0
732.7 425.6 2520.0 TBD 3678.4
Fig. 5. Propulsion system: (1) helium pressurant tank, (2) 500 N main engine, (3) conospherical propellant tank (10 N attitude thrusters not shown).
only before the final re-entry event. Spin-vector determination is accomplished through a slit-sun, and Earth Crossing detector. Gyros will be carried to provide additional safety checking information. Due to the large variation of mission mass, active nutation damping will be performed using thrusters. The Carrier Vehicle is of conventional design, but required to be highly mass efficient in a way similar to interplanetary missions. The spacecraft structure is based on a central Carbon Fibre Composite thrust tube, with line-charge separation systems, leading to a dry mass of the spacecraft of 300 kg. Four body mounted solar arrays are carried, charging a lithium-ion battery. The propulsion system comprises a conventional liquid bi-propellant system employing a 500 N main engine. In addition, 14 of 10 N thrusters will be used to provide attitude control (Fig. 5). A US miniaturised transponder was baselined, allowing the mission to utilise 13 m dish antennas within the ESOC. Despite an expected radiation dose rate of 10 krad, it is expected that the mission duration is in the order of
3. Hyperbolic re-entry vehicle design Extensive trades were performed to establish the effects of capsule ballistic coefficient with shape, stability, mass, thermal protection system selection, landing dispersion and development of the imposed spacecraft constraints. For the high loads and fluxes of the Earth return a rigid capsule and TPS are necessary. With the allowed level of deceleration and accurate surface location capability, a ballistic entry is by far the simplest and most robust. Classical theory shows that the minimum heat load is delivered to a ballistic vehicle if it has either a very low ballistic coefficient or a very high ballistic coefficient. For the low ballistic coefficient case, the deceleration is undertaken in a low atmospheric density and thus heating levels are low. For the high ballistic coefficient vehicle little energy is lost and heat loads are also low, but impact speeds are high. In practice, the low ballistic coefficient ballistic vehicle is most suited to scientific exploration missions due to landing velocity constraints. Three mature concepts were initially examined: the 60◦ cone of Huygens and Beagle2, the 45◦ cone of Pioneer Venus and the 30◦ cone from the CAESAR study. The 60◦ conical shape was selected, with a flat back, and small hemispherical cover for the payload and avionics. The concept is compatible with both land and water landings. The baseline atmospheric entry angle is approximately 20◦ . At 120 km, the spacecraft enters the atmosphere at a relative velocity of 12.8 km/s. The peak heat flux occurs at 59 km, approximately 15 s after passing through 120 km altitude. At this stage it is essential that the vehicle is correctly aligned so that the heat flux impinges on the heat shield rather than the rear of the vehicle. The peak acceleration of 118 g occurs approximately 3.5 s later. By 50 s after passing 120 km, the vehicle has decelerated to Mach 1.0 at a height of approximately 32 km. Vehicle stability is likely to be at its worst at Mach1; however, vehicle stability is not an absolute requirement at this phase in the mission, although the ability to ensure a stable attitude before landing is. The time from atmosphere interface to landing is approximately 9 min for both EVD and MSR if no parachute is provided. With the addition of a parachute the descent time could be increased significantly; however, the dispersion would also increase.
202
A. Phipps et al. / Acta Astronautica 61 (2007) 198 – 202
4. Summary and conclusions
Fig. 6. Hypersonic entry vehicle.
In the case of a delivery system error leading to a vertical entry, the peak heat flux of 35 MW/m2 would occur 6 s after passing through 120 km. The peak deceleration of 420g would occur 1 s later. The overall descent time would be 1 min less than the nominal. Although the loads are more severe than the baseline scenario, the sample container is perfectly capable of withstanding the entry ‘g’ loads and the carbon-phenolic heat shield material has been used against higher heat fluxes than this (Fig. 6). It is essential for the vehicle to be recovered after landing, and to that extent a GPS/radio beacon or a GPS/sonar beacon is baselined for land/ocean landing.
A concept for a demonstrator for the Earth re-entry element of a Mars Sample Return mission has been discussed. Such an Earth re-entry mission has not been thoroughly demonstrated before at such velocities, and the demonstration mission will allow many issues and risks in sample return missions to be evaluated, rehearsed and retired. An entry velocity of 12.8 km/s must be demonstrated, whilst safely returning a 1 kg sample in a 5 kg sealed canister under planetary protection conditions. A PSLV launch places the EVD into an initial GTO. Through a series of phasing orbits the apogee is increased to 300,000 km, where an apogee burn brings the spacecraft back into a retrograde Earth intercept trajectory. As a result of this complex trajectory the use of ESOC has been baselined. This series of trajectory manoeuvres require a high performance restartable liquid propulsion system. A land recovery in Australia has been baselined for the purposes of this study. Acknowledgements The study team would like to acknowledge funding and support from the European Space Agency for this work.
A high-energy sample return Earth re-entry demonstrator to address planetary protection issues Andy Phippsa,∗ , Alex da Silva Curiela , Martin Sweetinga , Arthur Smithb , Steve Lingardc a Surrey Satellite Technology Ltd, Surrey Space Centre, Guildford, Surrey GU2 7YE, UK b Fluid Gravity Engineering Ltd, 83 Market St, St Andrews, Fife, KY16 9NX, UK c Vorticity Ltd, Hampden House, Monument Business Park, Chalgrove, Oxfordshire, OX44 7RW UK
Available online 26 March 2007
Abstract For many of the proposed planetary exploration missions, robotically collected sample materials must be returned to the Earth for more detailed analysis. For the more interesting locations in the solar system, full planetary protection guidelines must be taken into account to prevent cross contamination. So far, few missions have demonstrated re-entry from interplanetary space, which requires a considerably greater velocity change than the more common Earth orbit re-entry missions, and so far none have dealt with the planetary protection issues. Missions such as the proposed Mars Sample Return mission must depend on reliable systems to carry out a sample return. Such a complex safety critical system can of course not be used for the first time on a real mission, nor can it be tested under Earth laboratory conditions. This paper proposes a design concept for an European Space Agency testbed to help validate the necessary technology. The spacecraft will carry scientific instruments to help characterise the environment, so that subsequent missions can rely on this. © 2007 Elsevier Ltd. All rights reserved.
1. Introduction
1.1. Spacecraft concept
The proposed mission is launched into the Earth orbit, and it carries an upper stage to propel it to interplanetary speeds, before being re-entered into the Earth’s atmosphere. The proposed re-entry probe weighs just 45 kg, and carries a sealed container with up to 1 kg sample material. Design trades related to the probe design and the recovery operation is discussed. The paper concludes that the demonstration mission will allow many issues and risks in sample return missions to be rehearsed and retired.
Trades have identified that the Earth-entry Vehicle Demonstrator is best implemented as a two stage spacecraft, providing a cost efficient solution (Fig. 1). The Hyperbolic Re-entry Probe (HRP) simulates the sample return vehicle returning from Mars, and a Carrier and Delivery System (CDS) which accelerates the probe and delivers it in an Earth-bound trajectory simulating that for a vehicle returning from Mars. Due to the complexity and criticality of the vehicle operations, operations from the European Space Operations Centre are baselined. 1.2. Trajectory analysis
∗ Corresponding author. Tel.: +44 1483 803803.
E-mail address: [email protected] (A. Phipps). 0094-5765/$ - see front matter © 2007 Elsevier Ltd. All rights reserved. doi:10.1016/j.actaastro.2007.01.015
Due to the Earth–Mars relative positions, suitable launch windows for efficient low-energy Mars transfers
A. Phipps et al. / Acta Astronautica 61 (2007) 198 – 202
199
Fig. 1. Earth Vehicle Demonstrator.
only present themselves approximately every two years. Based on an assessment of typical launch dates of potential sample return missions, Earth entry velocities from 12 to 12.8 km/s need to be considered. The design has therefore taken into account the worst case figure of 12.8 km/s. In order to demonstrate such a re-entry event, a spacecraft can be launched, and accelerated to re-enter at such a velocity. The choice of an initial retrograde orbit offers the advantage that less propellant needs to be carried, as the Earth rotational velocity effectively increases the relative re-entry velocity. For a prograde orbit, an orbital velocity of 13.3 km/s would need to be achieved, as opposed to an orbital velocity of 12.3 km/s for a retrograde orbit. Starting orbits of GTO were considered, as this provides ample opportunities as a piggyback or secondary passenger, but such orbits are prograde and would lead to a larger and more expensive mission. A dedicated retrograde Highly Elliptical Orbit was therefore considered as the baseline instead. A zero degree inclination was selected to provide contingency and multiple opportunities in time to reach the same re-entry site on the equator. Once in a highly elliptical orbit, the apogee is raised further in a stepwise fashion using spacecraft engines. This increases the perigee velocity, and provides schedule margin for the re-entry event planning. It also ensures the spacecraft spends ample time at apogee for the final major re-entry burn, which minimises gravity losses for this burn. The re-entry event itself commences with a final firing of the spacecraft engines at apogee. This places the spacecraft composite into a hyperbolic trajectory that enters the Earth atmosphere, for entry over the equator. During the descent, the spacecraft trajectory is closely tracked, and final corrections are made at a point
Fig. 2. The 300,000 km apogee orbit. (1) apogee raising manoeuvres, (2) apogee re-entry burn.
that balances accuracy of the navigation solution, leaving sufficient time for the corrections to have efficient results. At this point the Hyperbolic Return Vehicle is separated from the Carrier vehicle, and they re-enter on slightly different trajectories, with the aim to burn up the Carrier vehicle in the atmosphere. Additional factors that have been considered in the trajectory trades include • exposure of the spacecraft to the radiation belts; • effects of gravitational perturbations by the moon; • sensitivity to thruster alignment, timing and magnitude errors; • ease and accuracy of obtaining navigation information leading up to the re-entry event. As a result, a 300,000 km apogee was selected for the mission, with the sequence illustrated in Figs. 2 and 3. This results in a two-day timeline from the re-entry burn until the entry into the Earth atmosphere.
200
A. Phipps et al. / Acta Astronautica 61 (2007) 198 – 202
1.3. Launcher selection A number of launchers have been considered where compatible with a low cost launch (Table 1). For the mission baseline, a dedicated PSLV was selected, providing the best compromise between availability, control over precise launch parameters and cost. 1.4. Landing site selection Several landing sites were considered, including land, ocean and air retrieval. Near equatorial landing sites provide the lowest energy requirements for the spacecraft, but appear to be restricted to a water landing when taking into account the complexities and costs of sample retrieval. Although a water landing is attractive in terms of the design of the vehicle for terrestrial impact, the risk of damage to the sample container due to a
Fig. 4. Final re-entry of the mission components.
hard landing on debris, ships or even sea mammals in the ocean cannot be accepted, hence design for a hard landing is an absolute necessity. Northern Canada was considered, but non-equatorial sites would require extra propellant to be carried. The Australian desert does not require significant propellant to be carried, and the Woomera test range was finally considered as a suitable retrieval site for the purposes of the study. This site also permits the Carrier and Delivery Vehicle to be landed in the ocean (Fig. 4). Based on a worst case launch into Geostationary Transfer Orbit, the total propellant requirements are set by the necessary velocity increments of up to 3.7 km/s (Table 2). 2. Spacecraft design
Fig. 3. Point (3) Region of final acceleration thrust.
The launch mass of the spacecraft composite is 950 kg, of which over 70% accounts for the propellant carried. The Hyperbolic Entry Probe itself weighs 46 kg, with a 5 kg sample canister payload. The EVD vehicle is spin stabilised during all propulsive firing events, and de-spun to lower spin rates just
Table 1 Launcher Option
Mass capability
Heritage
Cost (USD)
PSLV
∼ 1050 kg into GTO
∼ $25 million
Soyuz-Fregat Ariane 5 Sylda Ariane 5 Cyclade
1350 kg into GTO > 1000 kg to GTO Soft limit of 700 kg to GTO
8 launches since 1993. First GTO launch in 2003 Yes Developed for Smart-1 mission
$35–$50 million
¥35 million ∼ ¥11 million for 350 kg platform
A. Phipps et al. / Acta Astronautica 61 (2007) 198 – 202 Table 2 Delta-V requirements Initial orbit
Trajectory Apogee raising DV Apogee DV Acceleration DV CDS avoidance DV Total Delta-V
201
weeks, and that some COTS-based avionics of SSTL’s Giove-A (in MEO orbit) heritage could be readily used. PSLV GTO (I = 17.7◦ ) Minimum value (m/s)
Including 10% margin (m/s)
666.1 387.0 2290.9 TBD 3344.0
732.7 425.6 2520.0 TBD 3678.4
Fig. 5. Propulsion system: (1) helium pressurant tank, (2) 500 N main engine, (3) conospherical propellant tank (10 N attitude thrusters not shown).
only before the final re-entry event. Spin-vector determination is accomplished through a slit-sun, and Earth Crossing detector. Gyros will be carried to provide additional safety checking information. Due to the large variation of mission mass, active nutation damping will be performed using thrusters. The Carrier Vehicle is of conventional design, but required to be highly mass efficient in a way similar to interplanetary missions. The spacecraft structure is based on a central Carbon Fibre Composite thrust tube, with line-charge separation systems, leading to a dry mass of the spacecraft of 300 kg. Four body mounted solar arrays are carried, charging a lithium-ion battery. The propulsion system comprises a conventional liquid bi-propellant system employing a 500 N main engine. In addition, 14 of 10 N thrusters will be used to provide attitude control (Fig. 5). A US miniaturised transponder was baselined, allowing the mission to utilise 13 m dish antennas within the ESOC. Despite an expected radiation dose rate of 10 krad, it is expected that the mission duration is in the order of
3. Hyperbolic re-entry vehicle design Extensive trades were performed to establish the effects of capsule ballistic coefficient with shape, stability, mass, thermal protection system selection, landing dispersion and development of the imposed spacecraft constraints. For the high loads and fluxes of the Earth return a rigid capsule and TPS are necessary. With the allowed level of deceleration and accurate surface location capability, a ballistic entry is by far the simplest and most robust. Classical theory shows that the minimum heat load is delivered to a ballistic vehicle if it has either a very low ballistic coefficient or a very high ballistic coefficient. For the low ballistic coefficient case, the deceleration is undertaken in a low atmospheric density and thus heating levels are low. For the high ballistic coefficient vehicle little energy is lost and heat loads are also low, but impact speeds are high. In practice, the low ballistic coefficient ballistic vehicle is most suited to scientific exploration missions due to landing velocity constraints. Three mature concepts were initially examined: the 60◦ cone of Huygens and Beagle2, the 45◦ cone of Pioneer Venus and the 30◦ cone from the CAESAR study. The 60◦ conical shape was selected, with a flat back, and small hemispherical cover for the payload and avionics. The concept is compatible with both land and water landings. The baseline atmospheric entry angle is approximately 20◦ . At 120 km, the spacecraft enters the atmosphere at a relative velocity of 12.8 km/s. The peak heat flux occurs at 59 km, approximately 15 s after passing through 120 km altitude. At this stage it is essential that the vehicle is correctly aligned so that the heat flux impinges on the heat shield rather than the rear of the vehicle. The peak acceleration of 118 g occurs approximately 3.5 s later. By 50 s after passing 120 km, the vehicle has decelerated to Mach 1.0 at a height of approximately 32 km. Vehicle stability is likely to be at its worst at Mach1; however, vehicle stability is not an absolute requirement at this phase in the mission, although the ability to ensure a stable attitude before landing is. The time from atmosphere interface to landing is approximately 9 min for both EVD and MSR if no parachute is provided. With the addition of a parachute the descent time could be increased significantly; however, the dispersion would also increase.
202
A. Phipps et al. / Acta Astronautica 61 (2007) 198 – 202
4. Summary and conclusions
Fig. 6. Hypersonic entry vehicle.
In the case of a delivery system error leading to a vertical entry, the peak heat flux of 35 MW/m2 would occur 6 s after passing through 120 km. The peak deceleration of 420g would occur 1 s later. The overall descent time would be 1 min less than the nominal. Although the loads are more severe than the baseline scenario, the sample container is perfectly capable of withstanding the entry ‘g’ loads and the carbon-phenolic heat shield material has been used against higher heat fluxes than this (Fig. 6). It is essential for the vehicle to be recovered after landing, and to that extent a GPS/radio beacon or a GPS/sonar beacon is baselined for land/ocean landing.
A concept for a demonstrator for the Earth re-entry element of a Mars Sample Return mission has been discussed. Such an Earth re-entry mission has not been thoroughly demonstrated before at such velocities, and the demonstration mission will allow many issues and risks in sample return missions to be evaluated, rehearsed and retired. An entry velocity of 12.8 km/s must be demonstrated, whilst safely returning a 1 kg sample in a 5 kg sealed canister under planetary protection conditions. A PSLV launch places the EVD into an initial GTO. Through a series of phasing orbits the apogee is increased to 300,000 km, where an apogee burn brings the spacecraft back into a retrograde Earth intercept trajectory. As a result of this complex trajectory the use of ESOC has been baselined. This series of trajectory manoeuvres require a high performance restartable liquid propulsion system. A land recovery in Australia has been baselined for the purposes of this study. Acknowledgements The study team would like to acknowledge funding and support from the European Space Agency for this work.