Human Mars Missions

1. Human Mars Missions Paul Wooster paul@developspace.net July 18, 2007

2. Getting to and from Mars

3. Earth Surface Earth Orbit Earth-Mars Transfer Mars-Earth Transfer Mars Orbit Mars Surface Earth-Mars-Earth Transportation Pathways Likely primary mission pathway Potential abort pathway Unlikely pathway

4. The crew uses the TSH for Earth-Mars transfer and Mars surface stay Return from Mars surface to Earth in ERV Only feasible if ISPP is utilized because of large ERV habitat mass Example: Zubrin’s “Mars Direct” 2 vehicle designs required: TSH, ERV Mars surface rendezvous, all assets on the surface and accessible to crew Split mission with pre-deployment is required because of the use of ISPP: ERV needs to be fully fueled and ready for Earth return before crew leaves Earth Direct Return Marssurface Marsorbit TSH ERV Broken lines: uncrewed operations Required crew transferbetween vehicles Solid lines: crewed operations Different colors indicatedifferent vehicles

5. Earth-Mars transfer in Transfer Habitat 1 Surface stay in surface habitat (SH) Return from Mars surface to Earth in transfer habitat (TH-2) deployed one opportunity before This strategy is only feasible with ISPP because of large return habitat mass Mission mode requires either a single landing site or a chain of landing sites 2 vehicle designs required: SH and TH Mars surface rendezvous, all assets on the surface and accessible to crew Split mission required due to need for pre-deployed TH-2 Makes most sense for a base: surface infrastructure (habitat, mobility, power system) could be emplaced once and re-supplied Direct Return + Surface Habitat Marssurface Marsorbit SH TH-1 TH-2

6. Earth-Mars transfer and Mars surface stay in TSH Ascent to Mars orbit in MAV Return from Mars orbit to Earth in ERV This strategy could work both with and without ISPP because the MAV crew compartment could be made very light (>5 mt) Example: Mars semi-direct, Mars DRM 1.0, 3.0 The Mars staging orbits would likely be dependent on the utilization of ISPP: ISPP: highly elliptic Mars staging orbit No ISPP: circular low Mars staging orbit 3 vehicle designs required: TSH, MAV, ERV Mars surface and Mars orbit rendezvous Split mission with pre-deployment is required in case of ISPP for MAV ascent propellants Transfer & Surface Hab and MOR Marssurface Marsorbit TSH MAV ERV

7. Earth-Mars and Mars-Earth transfers in Interplanetary Transfer Hab (ITH) Mars descent and ascent in Mars Descent & Ascent Vehicle (MDAV) Option for propulsive abort to orbit exists during Mars powered descent Surface stay in surface hab (SH) This strategy works both with ISPP and without The Mars staging orbits would likely be different for ISPP and no ISPP: ISPP: highly elliptic Mars staging orbit No ISPP: circular low Mars staging orbit 3 vehicle designs required: ITH, MDAV, SH Mars surface and Mars orbit rendezvous Makes most sense for a base: surface infrastructure (habitat, mobility, power system) could be emplaced once and re-supplied Mars Descent / Ascent Vehicle and MOR Marssurface Marsorbit MDAV SH ITH

8. Earth-Mars and Mars-Earth transfers in Interplanetary Transfer Hab (ITH) Mars descent in MDAV-2, Mars ascent in pre-deployed MDAV-1 Surface stay in pre-deployed surface hab (SH) This strategy would only be chosen if ISPP is utilized for the MDAV Mission mode requires either a single landing site or a chain of landing sites The Mars staging orbits would likely be different for ISPP and no ISPP: ISPP: highly elliptic Mars staging orbit No ISPP: circular low Mars staging orbit 3 vehicle designs required: ITH, MDAV, SH Mars surface and Mars orbit rendezvous Split mission required due to pre-deployed MDAV Makes most sense for a base: surface infrastructure (habitat, mobility, power system) could be emplaced once and re-supplied Pre-Deployed Mars Descent / Ascent Vehicle and MOR Marssurface Marsorbit MDAV-2 MDAV-1 SH ITH

9. Earth-Mars and Mars-Earth transfers in Interplanetary Transfer Hab (ITH) Mars landing & surface stay in Landing & Surface Hab (LSH) Ascent to Mars orbit in MAV This strategy works both with ISPP and without The Mars staging orbits would likely be different for ISPP and no ISPP: ISPP: highly elliptic Mars staging orbit No ISPP: circular low Mars staging orbit 3 vehicle designs required: ITH, LSH, ERV Mars surface and Mars orbit rendezvous Landing & Surface Habitat and MOR Marssurface Marsorbit LSH MAV ITH

10. Mapping of Major Mars Elements to Mission Phases

11. Major Drivers of Selected Mars Systems • Earth Launch and Departure Systems are driven by the total payload which must be launched towards Mars and the Delta-V associated with the required Earth-Mars transfer trajectory. • In-Space Habitation Systems are driven by crew size and the duration of the Earth-Mars transfer trajectory • Mars Aero-Systems for Mars orbit capture and Mars entry and descent are driven by the payloads they must support and the Mars atmospheric entry velocities of the Earth-Mars transfer trajectories they must withstand • Mars Landing Systems are dependent upon the payload they must deliver to the surface and the state (velocity and altitude) at which they must begin operation

12. Earth Departure Delta-V Standard (no abort option) Conjunction Trajectories

13. Earth Departure Delta-V Conjunction Trajectories with Earth-Mars-Earth Abort Option

14. Mars Entry Velocity

15. Abort Dependencies – MOR Architectures No Abort Outbound: TSH Prop Abort or Hybrid Outbound: ERV Abort Type FRT Abort Outbound: TSH or ERV

16. Ares Launch Vehicle Capability

17. Selected Mars Launch and Earth Departure Configurations 1 Ares V, H2/O2 2 Ares V, H2/O2 1 Ares V, NTR 2 Ares V, NTR Ares V EDS Ares V Core & SRBs Blue = Payload; Gray = EDS Fairing; Red = Nuclear Thermal Stage

18. Launch and Earth Departure Performance

19. 1 CaLV 1 CLV, 1 CaLV 2 CaLV, 1 EDS 2 CaLV, 2 EDS ~40 mt TMI ~50 mt TMI ~60 mt TMI ~90 mt TMI ~30 mt MO ~37 mt MO ~45 mt MO ~67 mt MO ~20 mt MS ~25 mt MS ~30 mt MS ~45 mt MS TMI – Trans-Mars Injection; MO – Mars Orbit; MS – Mars Surface ESAS Launch Vehicle Mars Capability

20. 15 m diameter aeroshell Aeroshell Sizing Impact on TMI Mass • CE&R aerocapture and aeroentry analysis indicated that aeroshell sizing (diameter) would have a major impact on maximum mass of Mars systems • ESAS CaLV has a fairing diameter of 8.4 meters, although larger fairings for Mars systems would likely be possible • For equal ballistic coefficient, the following entry mass limits likely apply for entry systems of the specified diameter: 8.4 m 10 m 15 m 12 m Georgia Tech CE&R Aeroentry Analysis (Ventry = 4.63 km/s, L/D = 0.5) 8.4^2 = 71 10^2 = 100 12^2 = 144 15^2 = 255

21. Aeroshell Sizing Impact on TMI Mass • CE&R aerocapture and aeroentry analysis indicated that aeroshell sizing (diameter) would have a major impact on maximum mass of Mars systems • ESAS CaLV has a fairing diameter of 8.4 meters, although larger fairings for Mars systems would likely be possible • For equal ballistic coefficient, the following entry mass limits likely apply for entry systems of the specified diameter: 8.4 m 10 m 15 m 12 m 2 LV, 2 EDS 2 LV, 1 EDS 1.5 LV, 1 EDS 1 LV, 1 EDS 8.4^2 = 71 10^2 = 100 12^2 = 144 15^2 = 255

22. Highly Elliptic Mars Orbits • Use of highly elliptic Mars orbits can decrease trans-Earth injection (TEI) delta-V relative to TEI from low Mars orbit • In this analysis, a ~10 hour period highly elliptic orbit is assumed in elliptic orbit cases, which decreases TEI delta-V by 1,000 m/s (this delta-V is added to Mars ascent requirements) • For the above chart, the pericenter is 3,700 km, equivalent to an altitude of ~300 km

23. On Mars

24. Moon-Mars Thermal Comparison(Regolith Surface Temperature)

25. Moon-Mars-Boston Thermal Comparison(Regolith/Concrete Surface Temperature)

28. Mars Water

29. Motivation for Solar Power on Mars • Mars missions studies frequently select nuclear fission reactors for providing surface power based upon technical considerations • e.g., mass, complexity, volume, etc. • However, requiring nuclear power places a policy constraint on the critical path for human Mars missions • Need sustained funding to develop and test reactor • Need political approval for every Mars mission including a reactor • Such a constraint compounds already existing policy challenges • Increased dependence on the stance of political parties/politicians • Sensitivity to changes in public view-point with respect to nuclear power in general • Other options such as dynamic isotope power systems also suffer from approval and funding constraints • If solar power can be used as primary power source on Mars it could increase the political feasibility and sustainability of human Mars missions Note: These policy constraints would also apply to nuclear thermal and nuclear electric propulsion

30. Is Solar Power Technically Feasible? • Preliminary analysis during our CE&R study indicated that solar power would be feasible as the primary power source for human Mars missions • Reasonably straightforward for missions without ISRU, more challenging for missions with extensive ISRU • We will conduct a further study of the technical feasibility and consequences of relying upon solar power for human Mars missions • Factors to be considered include required day and night-time power/energy levels, landing location (latitude), seasons and distance to Sun, atmospheric opacity and light scattering (including statistical nature of dust storms), array type (tracking, fixed-incline, fixed-horizontal) • Relying on solar power would likely place constraints on where missions could be conducted, available power levels, and may drive towards a single base (if a large investment is required to emplace the solar power system)

32. Mars Water Mars Elevation

33. Moon and Mars

34. What to do on Moon to prepare for Mars • Three main areas in which the Moon can offer aid in preparing for Mars, which could be considered objectives for lunar campaigns: • Testing systems, technologies, and procedures for Mars exploration in an environment distinct from Earth. • Increasing understanding of partial gravity (possibly coupled with radiation) impacts on crew health and performance. • Providing an intermediate milestone for human space exploration efforts.

35. How to measure those things • Metrics for each of the three areas of objectives for Mars preparation using the Moon are as follows: • Degree to which lunar systems are similar to Mars systems and fraction of Mars systems validated during lunar activities • Number of crew exposed to particular durations of lunar partial gravity (e.g., # at 1 month, # at 3 months, # at 6 months, etc.), with longer durations preferred • Date of initial occurrence of high visibility events (e.g., lunar vicinity flight, human lunar landing, long-duration mission, total surface time of Mars mission)

36. Mars Exploration Elements • Following list of elements required for Mars exploration • Earth Launch and Entry Crew Cabin(s) • Heavy Lift Launch Vehicle and Earth Departure Systems • Descent Stage • Heatshields • Long-term Surface Habitat • Mars Ascent Vehicle (Cabin and Propulsion) • Earth Return Vehicle (Habitat and Propulsion) • EVA and Mobility Systems • Surface Power Systems • Following list of technologies beneficial for Mars missions • In-Situ Propellant Production/In-Situ Consumables Production • Mars ISRU Compatible Propulsion (e.g., CH4/O2, C2H4/O2) Items denoted in blue indicate high potential for Moon-Mars commonality

37. Moon-Mars Exploration System Commonality Distinct Moon, Mars Exploration Systems,Lunar Operations Maintained • If distinct systems are developed for Moon and Mars, we may: • Significantly delay Mars operations • Need to curtail lunar operations to enable Mars (development, operations), resulting in a Moon-Mars mission gap • Never get to Mars at all, because the renewed major investment is not sustainable • By developing a common Moon-Mars exploration system, we can overcome these obstacles and also: • Directly validate key Mars elements during lunar missions • Gain experience in routine production and system operation, decreasing cost and risk • Avoid workforce disruption during transition from Moon to Mars, and possibly continue lunar operations during Mars missions • Provide direct tie between Moon and Mars exploration in the eyes of the public and Congress Distinct Moon, Mars Exploration Systems,Lunar Missions Curtailed Common Moon-Mars Exploration System,Option to Maintain Lunar Missions

38. Commonality Strategy – Transportation Development Roadmap LEO / ISS Mission Hardware Short Lunar Mission Hardware Common in-space propulsion stage (LCH4 / LOX): Core propulsion stage XL strap-on tanks XXL strap-on tanks (ERV) LAT for CEV capsule: CEV + IPU (27 m3 ): Lunar landing gear & exoskeleton: LEO propulsion stage: Engine 1 (LCH4 / LOX) Restartable, non-throttleable: CEV power pack: CEV launch vehicle: Engine 2 (LCH4 / LOX) Throttleable: Design Philosophy: Maximize hardware commonality to minimize gap between lunar and Mars missions and overall development and production costs Common Earthdeparture stage(LH2 / LOX) Heavy Lift Launch Vehicle: (“2 stages”, 100 mt to LEO) Mars Mission Hardware SDLV upper stage (125 mt to LEO), potentially EDS-derived: Mars landing gear & exoskeleton: Long Lunar Mission Hardware Habitat core and inflatablepressurized tent forplanetary surfaces: Integrated aeroshell Note: Block upgrades across phases are not depicted

39. Base Moon-Mars Exploration System Commonality Concept Lunar Transportation Architecture Mars Transportation Architecture • High-level commonality concept developed during Base Period using selected Moon and Mars architectures • Commonality focused on design reuse of complete elements, with modularity in “Yellow Stage” and habitat design • Develop high-level scheme to identify elements where commonality may be beneficial • Can be based upon elements with similar capabilities (or requirements) • Need to be careful which requirements are compared • e.g., for a propulsion stage, the combination of delta-v, payload, and thrust characterize the capability (to first order); taken in isolation they do not • Develop commonality concept in further detail • Trades must be performed between modularity/platforming or “stretchable” options relative to a single design for many use cases Note: While commonality shown for a particular pair of architectures, approach is not unique to those chosen

40. Extensible Destination Vicinity Propulsion System Integrated view Exploded view Core LCH4 / LOX propulsion stage -4 non-throttleable LCH4/LOX engines -4 common-bulkhead tanks (CBH) -primary and secondary structure Derived LCH4 / LOX propulsion stages Lunar / Mars descent Mars ascent Mars TEI -4 non-throttleable LCH4/LOX engines -8 common-bulkhead tanks -primary and secondary structure -additional primary structure -4 non-throttleable LCH4/LOX engines -4 common-bulkhead tanks -4 spherical extension tanks (2 CH4, 2 O2) -primary and secondary structure -dedicated exoskeleton for Mars TEI -4 throttleable LCH4/LOX engines -8 common-bulkhead tanks -primary and secondary structure -dedicated Moon and Mars landing gears and exoskeletons

41. Common Destination Vicinity Propulsion System Mars Ascent Vehicle Mars Moon Crew Transport • Modular solution for Destination Vicinity Propulsion System • Common propulsion stage core employed in all use-cases (sized by Lunar Ascent & TEI) • Duplicate set of tanks (relative to core) provides additional propellant for Lunar/Mars Descent and Mars Ascent • Extra-large set of strap-on tanks used for TEI from Mars on Earth Return Vehicle • Descent stage structural ring and landing gear specific to destination due to distinct loading conditions • Common ascent engines, common descent engines for Moon [2 engines] and Mars [4 engines] Transfer and Surface Habitat Earth Return Vehicle Surface Habitat

42. Moon-Mars Common System Vehicle Stacks Lunar Direct Return (Arch 1) Mars Orbit Rendezvous: Combined Trans. and Surf. Habs (Arch. 969) Short Mission Long Mission Outbound Transfer & Surface Habitat Mars Ascent Vehicle & Return CEV Earth Return Habitat & Propulsion Lunar Long-Duration Surface Habitat Lunar CrewTransfer System • Elements combine together to form vehicle stacks for variety of missions • Numbers at left represent wet mass in metric tonnes of elements in LEO • Earth Departure Stages have the same dry mass (11 mt) and maximum wet mass (112 mt) • CEVLV capacity 30 mt • Lunar HLLV capacity 100 mt • Mars HLLV upgraded to 125 mt • Low commonality overhead due to appropriate use of modularity to support variants 9 9 9 mt Hab: 49 27 21 Hab: 25 AS: 33 DS: 33 TEIS: 57 DS: 33 36 34 HS: 34 HS: 34 HS: 34 112 106 59 106 39 81 112 100 106 106 Post-Earth departure commonality mass overhead relative to customized systems: 1% 2% 4% 3% 2% IMLEO commonality overhead relative to customized systems: 63% savings in unique element dry mass for common vs. custom system design 13% 20% 4% 4% 3% Number launches (HLLV+CEVLS): 2+1 2+0 3+1 3+0 3+0 For modest mass increase, Mars-back commonality offers significant savings in development and production

43. Extending ESAS Elements to Mars Missions • Based upon the significant capability of ESAS launch vehicles to deliver payloads to Mars, options to extend the remaining elements were assessed • In the baseline architecture presented: • Mars-dedicated aeroentry and propulsion systems are developed for aerocapture/descent, landing, ascent, and trans-Earth injection • The CEV is extended to provide habitation during Earth launch and entry, and during Earth-Mars and Mars-Earth transit in combination with the LSAM-derived Mars Landing and Ascent Vehicle crew compartment, as part of a dual-launch, dual-heatshield crew transportation system • A large surface habitat capable of supporting up to 6 crew members is positioned to the Martian surface prior to the arrival of the crew in a single launch • Single launch logistics flights pre-position consumables, surface exploration equipment, and power infrastructure for initial mission, resupply consumables and spares for subsequent missions • Total number of CaLV launches required (no CLV launches are needed) is presented for a low launch demand and a high launch demand scenario • “Low demand” could be achieved with 4 crew (in one crew transportation system), methane-oxygen propulsion for maneuvers near Mars, and ISRU for both consumables production and propellant production • “High demand” could be achieved either with 4 crew (in two crew transportation systems), hypergolic propulsion, and no ISRU, or with 6 crew (in two crew transportation systems), methane-oxygen propulsion, and no ISRU • While not presented, intermediate launch demand options also exist with other crew size and technology combinations

44. Conceptual Mars Exploration Architecture Based on Lunar Elements Mars Surface Mars Orbit Interplanetary Transfer Earth Vicinity Mars Crew Transportation System Concept Trans-Mars Configuration Launch Configuration Crew Transport Surface Habitat Logistics Flights

45. Conceptual Mars Exploration Architecture Based on Lunar Elements Mars Surface Mars Orbit Interplanetary Transfer Earth Vicinity Low demand: 4 crew, methane-oxygen prop, ISRU High demand: 4 crew, hypergolic prop, no ISRU OR 6 crew, methane-oxygen prop, no ISRU Crew Transport Surface Habitat Logistics Flights

46. Mars Considerations for the Moona.k.a. Mars-Back to the Moon Additional Considerations for building and extensible moon/Mars exploration system

47. What to do on Moon to prepare for Mars • Metrics for each area: • Degree to which lunar systems, subsystems, etc. are similar to Mars systems and fraction of Mars systems validated during lunar activities (with validated earlier being better) • Number of crew exposed to particular durations of lunar partial gravity (e.g., # at 1 month, # at 3 months, # at 6 months, etc.), with longer durations preferred • Date of initial occurrence of high visibility events (e.g., lunar vicinity flight, human lunar landing, long-duration mission, total surface time of Mars mission) • Areas in which Moon can aid in preparing for Mars: • Testing systems, technologies, and procedures for Mars exploration in an environment distinct from Earth • Gain operational and developmental experience; test operations procedures • Test technologies for Mars systems (e.g., method of water recycling • Test Mars subsystems (e.g., ECLSS, fuel cell) • Test Mars systems (e.g., habitat, rover) • Increasing understanding of partial gravity (possibly coupled with radiation) impacts on crew health and performance • Providing an intermediate milestone for human space exploration efforts

48. Measuring Readiness for Mars • It may be useful to define a “Mars Exploration Readiness Level” (MERL) to determine Mars preparedness • Would serve a similar function to the TRL scale for measuring technologies • Having such a scale would enable: • Reviewing status • Monitoring progress • Determining criteria for tests • Lunar testing would be one means of increasing MERL • Laboratories, terrestrial Mars-analogs, LEO, deep-space, and Mars could also serve as test environments • MERL could be applied to areas 1 and 2 from prior listing (i.e., technical readiness and human physiological readiness) • MERL could be a hierarchical measure, and at top level would depend upon overall Mars architecture(s) under consideration Level Required by NASAto commence Mars missions MERL Time

49. Potential Areas of Moon-Mars Commonality • Habitat sub-systems • Habitat structure • Note: Mars mission will require relatively large habitat volume, should limit surface assembly due to safety considerations • Mars EDL system would place additional constraints on habitat dimensions, CG • Surface exploration systems (i.e., rovers, EVA aids, etc.) • EVA suits • Possible differences in thermal, note also different in-situ resources may impact design • Power systems (generation and storage) • For solar, storage is easier on Mars, generation harder • Operations commonality (time lag and associated impacts) • Should lunar missions operate in this manner from the start? • Resupply and logistics? • Propulsion? • Ascent, TEI: LCH4/LOX, C2H4/LOX, RP-1/LOX?, etc. • H2/LOX for descent? • Note: Lunar ISRU is intentionally not on this list While these areas offer potential for commonality, the list does not indicate what should be done in designing lunar systems to make them common with Mars

50. Mars Considerations for Lunar Systems • In order to determine Mars-related considerations for the development of lunar systems (so as to aid in lunar system design decision making), we ask: • How is Mars different from the Moon? • How do the differences tend to impact Mars exploration systems? • What can be done with lunar systems to have higher relevance towards Mars systems?