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

2. Slide 2 Getting to and from Mars

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

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

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

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

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

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

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

10. Slide 10 Mapping of Major Mars Elements to Mission Phases

11. Slide 11 Major Drivers of Selected Mars Systems l 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. l In-Space Habitation Systems are driven by crew size and the duration of the Earth-Mars transfer trajectory l 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 l 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. Slide 12 Earth Departure Delta-V Standard (no abort option) Conjunction Trajectories

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

14. Slide 14 Mars Entry Velocity

15. Slide 15 Abort Type No Abort Prop Abort or Hybrid FRT Abort Outbound: TSH Outbound: ERV Outbound: TSH or ERV Outbound Abort?Ascent Abort?Outbound Config# Ares I / COTS# Ares V (Base)# EDS (Base)CEV (COTS) Yes MAV & ERV03 (2) 1 Yes TSH & ERV0331 Yes TSH, Earth Aerocapture 1332 (1) YesNo TSH, Earth Aerocapture 0331 Yes TSH & Droppable CEV 0432 NoYesTSH1332 (1) No TSH0331 Abort Dependencies – MOR Architectures

16. Slide 16 Ares Launch Vehicle Capability

17. Slide 17 Selected Mars Launch and Earth Departure Configurations Ares V Core & SRBs Ares V EDS Blue = Payload; Gray = EDS Fairing; Red = Nuclear Thermal Stage 1 Ares V, H 2 /O 2 2 Ares V, H 2 /O 2 1 Ares V, NTR2 Ares V, NTR

18. Slide 18 Launch and Earth Departure Performance

19. Slide 19 ESAS Launch Vehicle Mars Capability 1 CaLV1 CLV, 1 CaLV2 CaLV, 1 EDS2 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

20. Slide 20 Georgia Tech CE&R Aeroentry Analysis (Ventry = 4.63 km/s, L/D = 0.5) 15 m diameter aeroshell Aeroshell Sizing Impact on TMI Mass l CE&R aerocapture and aeroentry analysis indicated that aeroshell sizing (diameter) would have a major impact on maximum mass of Mars systems l ESAS CaLV has a fairing diameter of 8.4 meters, although larger fairings for Mars systems would likely be possible l For equal ballistic coefficient, the following entry mass limits likely apply for entry systems of the specified diameter: 8.4^2 = 7110^2 = 10012^2 = 14415^2 = 255 Diameter [m]8.4101215 Entry Mass Range [mt]25 to 3135 to 4451 to 6480 to 100 10 m 15 m 12 m 8.4 m

21. Slide 21 Aeroshell Sizing Impact on TMI Mass l CE&R aerocapture and aeroentry analysis indicated that aeroshell sizing (diameter) would have a major impact on maximum mass of Mars systems l ESAS CaLV has a fairing diameter of 8.4 meters, although larger fairings for Mars systems would likely be possible l For equal ballistic coefficient, the following entry mass limits likely apply for entry systems of the specified diameter: 8.4^2 = 7110^2 = 10012^2 = 14415^2 = 255 10 m 15 m 12 m 8.4 m Diameter [m]8.4101215 Entry Mass Range [mt]25 to 3135 to 4451 to 6480 to 100 1 LV, 1 EDS 1.5 LV, 1 EDS 2 LV, 1 EDS 2 LV, 2 EDS

22. Slide 22 Highly Elliptic Mars Orbits l Use of highly elliptic Mars orbits can decrease trans-Earth injection (TEI) delta-V relative to TEI from low Mars orbit l 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) l For the above chart, the pericenter is 3,700 km, equivalent to an altitude of ~300 km

23. Slide 23 On Mars

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

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

26. Slide 26

27. Slide 27

28. Slide 28 Mars Water

29. Slide 29 Motivation for Solar Power on Mars l Mars missions studies frequently select nuclear fission reactors for providing surface power based upon technical considerations n e.g., mass, complexity, volume, etc. l However, requiring nuclear power places a policy constraint on the critical path for human Mars missions n Need sustained funding to develop and test reactor n Need political approval for every Mars mission including a reactor l Such a constraint compounds already existing policy challenges n Increased dependence on the stance of political parties/politicians n Sensitivity to changes in public view-point with respect to nuclear power in general l Other options such as dynamic isotope power systems also suffer from approval and funding constraints l 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. Slide 30 Is Solar Power Technically Feasible? l Preliminary analysis during our CE&R study indicated that solar power would be feasible as the primary power source for human Mars missions n Reasonably straightforward for missions without ISRU, more challenging for missions with extensive ISRU l We will conduct a further study of the technical feasibility and consequences of relying upon solar power for human Mars missions n 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) l 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)

31. Slide 31

32. Slide 32 Mars Water Mars Elevation

33. Slide 33 Moon and Mars

34. Slide 34 What to do on Moon to prepare for Mars l Three main areas in which the Moon can offer aid in preparing for Mars, which could be considered objectives for lunar campaigns: 1. Testing systems, technologies, and procedures for Mars exploration in an environment distinct from Earth. 2. Increasing understanding of partial gravity (possibly coupled with radiation) impacts on crew health and performance. 3. Providing an intermediate milestone for human space exploration efforts.

35. Slide 35 How to measure those things l Metrics for each of the three areas of objectives for Mars preparation using the Moon are as follows: 1. Degree to which lunar systems are similar to Mars systems and fraction of Mars systems validated during lunar activities 2. 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 3. 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. Slide 36 Mars Exploration Elements l Following list of elements required for Mars exploration n Earth Launch and Entry Crew Cabin(s) n Heavy Lift Launch Vehicle and Earth Departure Systems n Descent Stage n Heatshields n Long-term Surface Habitat n Mars Ascent Vehicle (Cabin and Propulsion) n Earth Return Vehicle (Habitat and Propulsion) n EVA and Mobility Systems n Surface Power Systems l Following list of technologies beneficial for Mars missions n In-Situ Propellant Production/In-Situ Consumables Production n Mars ISRU Compatible Propulsion (e.g., CH 4 /O 2, C 2 H 4 /O 2 ) Items denoted in blue indicate high potential for Moon-Mars commonality

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

38. Slide 38 Commonality Strategy – Transportation Development Roadmap Design Philosophy: Maximize hardware commonality to minimize gap between lunar and Mars missions and overall development and production costs CEV + IPU (27 m 3 ): Integrated aeroshell Mars Mission Hardware LEO / ISS Mission Hardware Common in-space propulsion stage (LCH 4 / LOX): Core propulsion stage XL strap-on tanks XXL strap-on tanks (ERV) Heavy Lift Launch Vehicle: (“2 stages”, 100 mt to LEO) Short Lunar Mission Hardware Habitat core and inflatable pressurized tent for planetary surfaces: Long Lunar Mission Hardware Note: Block upgrades across phases are not depicted LEO propulsion stage: CEV launch vehicle: CEV power pack: LAT for CEV capsule: SDLV upper stage (125 mt to LEO), potentially EDS- derived: Mars landing gear & exoskeleton: Engine 1 (LCH 4 / LOX) Restartable, non-throttleable: Common Earth departure stage (LH 2 / LOX) Engine 2 (LCH 4 / LOX) Throttleable: Lunar landing gear & exoskeleton:

39. Slide 39 Base Moon-Mars Exploration System Commonality Concept l High-level commonality concept developed during Base Period using selected Moon and Mars architectures l Commonality focused on design reuse of complete elements, with modularity in “Yellow Stage” and habitat design l Develop high-level scheme to identify elements where commonality may be beneficial n Can be based upon elements with similar capabilities (or requirements) n Need to be careful which requirements are compared we.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 l Develop commonality concept in further detail n 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 Lunar Transportation ArchitectureMars Transportation Architecture

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

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

42. Slide 42 Moon-Mars Common System Vehicle Stacks Post-Earth departure commonality mass overhead relative to customized systems: Lunar Direct Return (Arch 1) Mars Orbit Rendezvous: Combined Trans. and Surf. Habs (Arch. 969) Short MissionLong Mission Lunar Crew Transfer System Lunar Long- Duration Surface Habitat Outbound Transfer & Surface Habitat Earth Return Habitat & Propulsion Mars Ascent Vehicle & Return CEV 81 100 112 106 59 39 36 34 9 21 9 mt 27 AS: 33 9 DS: 33 Hab: 49 TEIS: 57 Hab: 25 HS: 34 Number launches (HLLV+CEVLS): l Elements combine together to form vehicle stacks for variety of missions l Numbers at left represent wet mass in metric tonnes of elements in LEO n Earth Departure Stages have the same dry mass (11 mt) and maximum wet mass (112 mt) n CEVLV capacity 30 mt n Lunar HLLV capacity 100 mt n Mars HLLV upgraded to 125 mt l Low commonality overhead due to appropriate use of modularity to support variants 2+12+03+13+0 1%2%4%3%2% IMLEO commonality overhead relative to customized systems: 13%20%4% 3% 63% savings in unique element dry mass for common vs. custom system design For modest mass increase, Mars-back commonality offers significant savings in development and production

43. Slide 43 Extending ESAS Elements to Mars Missions l Based upon the significant capability of ESAS launch vehicles to deliver payloads to Mars, options to extend the remaining elements were assessed l In the baseline architecture presented: n Mars-dedicated aeroentry and propulsion systems are developed for aerocapture/descent, landing, ascent, and trans-Earth injection n 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 n 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 n Single launch logistics flights pre-position consumables, surface exploration equipment, and power infrastructure for initial mission, resupply consumables and spares for subsequent missions l Total number of CaLV launches required (no CLV launches are needed) is presented for a low launch demand and a high launch demand scenario n “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 n “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 n While not presented, intermediate launch demand options also exist with other crew size and technology combinations

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

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

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

47. Slide 47 What to do on Moon to prepare for Mars l Areas in which Moon can aid in preparing for Mars: 1. Testing systems, technologies, and procedures for Mars exploration in an environment distinct from Earth a)Gain operational and developmental experience; test operations procedures b)Test technologies for Mars systems (e.g., method of water recycling c)Test Mars subsystems (e.g., ECLSS, fuel cell) d)Test Mars systems (e.g., habitat, rover) 2. Increasing understanding of partial gravity (possibly coupled with radiation) impacts on crew health and performance 3. Providing an intermediate milestone for human space exploration efforts l Metrics for each area: 1. 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) 2. 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 3. 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)

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

49. Slide 49 Potential Areas of Moon-Mars Commonality l Habitat sub-systems l Habitat structure n Note: Mars mission will require relatively large habitat volume, should limit surface assembly due to safety considerations n Mars EDL system would place additional constraints on habitat dimensions, CG l Surface exploration systems (i.e., rovers, EVA aids, etc.) l EVA suits n Possible differences in thermal, note also different in-situ resources may impact design l Power systems (generation and storage) n For solar, storage is easier on Mars, generation harder l Operations commonality (time lag and associated impacts) n Should lunar missions operate in this manner from the start? l Resupply and logistics? l Propulsion? n Ascent, TEI: LCH 4 /LOX, C 2 H 4 /LOX, RP-1/LOX?, etc. n H 2 /LOX for descent? l 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. Slide 50 Mars Considerations for Lunar Systems l 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: n How is Mars different from the Moon? n How do the differences tend to impact Mars exploration systems? n What can be done with lunar systems to have higher relevance towards Mars systems?

51. Slide 51 Differences between Mars and the Moon (1) l Mars has an atmosphere n Mars systems require an aeroshell for EDL wAeroshell must be launchable at Earth (constrains diameter) wIntegrated system must be stable through atmospheric flight wTends to lead to compact, low cg descent stage and payloads for Mars wFor Moon: Consider designing payloads to meet Mars aeroshell constraints n CO 2 is easily available on Mars wMars systems would tend to use atmospheric-based ISRU, rather than more challenging, complex, and risk regolith-based ISRU wFor Moon: Consider use of Mars ISRU related processes in ECLS and power systems (e.g., water electrolysis, Sabatier) wFor Moon: Do not justify lunar regolith or water-ice based ISRU on the basis of relevance to Mars; do not predicate Mars ISRU on success with lunar ISRU n Mars surface radiation environment more benign than Moon wMars systems may require less shielding while on surface wFor Moon: Unclear if this carries implications; may impact common mobility system design in that Mars systems would not need as much shielding

52. Slide 52 Differences between Mars and the Moon (2) l Campaign- / mission-level differences n Mars missions will be longer and possibly feature larger crew wRequired surface habitat volume will be greater for Mars than Moon wFor Moon: Consider oversized (relative to lunar requirements) habitat or means for increasing habitat volume between Moon and Mars n Full Mars habitation capabilities are required from the start wMars systems would tend to limit assembly required to provide crew habitation (assembly might be used to augment capabilities) wFor Moon: Consider means of providing core crew habitation functionality without assembly n Mars missions do not have any opportunity for re-supply wLeads towards pre-positioning of required consumables and sufficient up- front planning to anticipate full needs of mission wFor Moon: Consider logistics strategy that does not rely on rapid re-supply, but uses pre-placement or pre-planned re-supply

53. Slide 53 Differences between Mars and the Moon (3) l Mars has approximately twice the gravity of the Moon n Mars surface operations will be under a higher gravitational load than lunar surface operations wTends towards desire for lighter space suits for Mars wTends towards decreased reliance on astronauts lifting, carrying items wMobility energy requirements will be higher for a given mass wMobility systems may become stuck more easily / more stuck on Mars wFor Moon: Consider means to decrease space suit mass further than may be needed for lunar ops; limit degree to which astronauts lifting is required wFor Moon: Aim to minimize mobility “payload” mass; consider means to recover higher weight mobility systems l Mars is further away from Earth than the Moon n Mars missions will feature a significant round-trip communications lag wMars missions operations will conducted in very different manner (relative to operations to date) in order to accommodate time lag; crew will have significantly greater autonomy, with Earth supporting and advising on a longer time-scale wMars systems will require greater automation in order to monitor system health and limit workload on surface crew wFor Moon: Consider means of simulating Earth-Mars communications lag in lunar operations; establish lunar surface operations methodology including time lag

54. Slide 54 Differences between Mars and the Moon (4) l Mars is further from Sun and has a ~24 hour day n Solar incidence is lower although eclipse time is shorter wSolar power generation more challenging on Mars wEnergy storage (e.g., regen fuel cells) is easier on Mars wFor Moon: Consider means to modularize lunar solar power generation system so that additional units can be included in Mars case; consider means to decrease energy storage capacity for Mars relative to Moon wNote: Nuclear power is not necessarily required for Mars l Mars thermal environment is different from the Moon n Regolith temperature more moderate on Mars than Moon, also have atmospheric interactions on Mars wDifferent thermal systems might be required between Moon and Mars (although more analysis needed) wFor Moon: Identify how relevant lunar systems are towards Mars, determine options for transitioning from lunar thermal environment to Mars l Mars may have or may have had indigenous life n Should restrict contamination of sites of scientific interest with organisms transported by humans; limit exposure of crew to martian organisms wTechniques for sterile sampling of biologically interesting sites may be required wTechniques for preventing biological transfer into crewed areas may be required wFor Moon: Consider testing of such techniques as part of lunar operations

55. Slide 55 l Concept to implement delay in Earth-Moon communications so that lunar surface missions are conducted in a manner and with an organization that is relevant to Mars n Delay would be in Earth to Moon comm. leg (not Moon to Earth), so that mission support could actively assess situation and switch to real-time comm. in an emergency n Primary role of mission support organization would be to assist crew and monitor longer time-scale processes – real time decision-making would reside on the lunar surface l Option to include an Earth based “artificial artificial intelligence” organization may be worthwhile to reduce need to develop onboard automation prior to lunar missions n Would be staffed by personnel simulating the role of “automation” n Should be tightly tied to organization developing automation software, to inform the development of software and test various techniques n Human involvement in this area would be decreased / eliminated over time l Could also have 2 “virtual astronauts” on Earth with real-time comm. links n Could assist in difference in crew size between Moon and Mars n Could monitor EVAs, teleoperate rovers, etc. n Could test various roles for crew members in a Mars mission Mars-Driven Lunar Missions Operations Concept Mission Support 4 to 20 min. lag

56. Slide 56 l Concept to implement delay in Earth-Moon communications so that lunar surface missions are conducted in a manner and with an organization that is relevant to Mars n Delay would be in Earth to Moon comm. leg (not Moon to Earth), so that mission support could actively assess situation and switch to real-time comm. in an emergency n Primary role of mission support organization would be to assist crew and monitor longer time-scale processes – real time decision-making would reside on the lunar surface l Option to include an Earth based “artificial artificial intelligence” organization may be worthwhile to reduce need to develop onboard automation prior to lunar missions n Would be staffed by personnel simulating the role of “automation” n Should be tightly tied to organization developing automation software, to inform the development of software and test various techniques n Human involvement in this area would be decreased / eliminated over time l Could also have 2 “virtual astronauts” on Earth with real-time comm. links n Could assist in difference in crew size between Moon and Mars n Could monitor EVAs, teleoperate rovers, etc. n Could test various roles for crew members in a Mars mission Mars-Driven Lunar Missions Operations Concept Mission Support Artificial Automation Virtual Astronauts 8 to 40 min. lag emergency comm.