EX13-98-036 

Reference Mission Version 3.0 

Addendum to the Human Exploration of 

Mars: The Reference Mission of the 

NASA Mars Exploration Study Team 

Exploration Office 

Advanced Development Office 

June, 1998 

National Aeronautics and 

Space Administration

Lyndon B. Johnson Space Center 

Houston, TX 77058
EX13-98-036

 
EX13-98-036 

 

Reference Mission Version 3.0 

Addendum to the Human Exploration of Mars: 

The Reference Mission of the NASA Mars Exploration Study Team

June, 1998 

Edited By: 

______________________________ 

Bret G. Drake 

Human Mission Study Lead, Exploration Office 

Approved By: 

______________________________ 

Douglas R. Cooke 

Manager, Exploration Office
EX13-98-036

 
EX13-98-036 

i

Contents 

A1.0 Introduction........................................................................................................................................1 

A2.0 Strategic Modifications.......................................................................................................................1 

A2.1 Reference Mission 1.0 Launch Strategy..........................................................................................2 

A2.1.1 System Repackaging .................................................................................................................5 

A2.1.2 System Mass Reductions...........................................................................................................6 

A2.1.3 Modified Launch Strategy.........................................................................................................7 

A2.2 Elimination of Initial Habitat Flight ................................................................................................7 

A2.2.1 Volume Augmentation ..............................................................................................................8 

A2.2.2 Redundancy Considerations......................................................................................................8 

A2.3 Revised Mission Strategy for Version 3.0.......................................................................................9 

A3.0 System Design Improvements ..........................................................................................................10 

A3.1 Incorporation of TransHab Improvements ....................................................................................10 

A3.2 System Improvements ...................................................................................................................12 

A3.2.1 In-Situ Resource Utilization .......................................................................................................12 

A3.2.2 Power Systems ........................................................................................................................12 

A3.2.3 Science Systems ......................................................................................................................14 

A3.2.4 EVA Systems ..........................................................................................................................15 

A3.3 Transportation System Improvements...........................................................................................16 

A3.3.1 Earth-to-Orbit Transportation .................................................................................................16 

A3.3.2 Trans-Mars Injection...............................................................................................................19 

A3.3.3 Aeroassist ................................................................................................................................21 

A3.3.4 Descent and Landing...............................................................................................................23 

A3.3.5 Ascent......................................................................................................................................26 

A3.3.6 Trans-Earth Injection...............................................................................................................26 

A3.3.7 Launch Packaging ...................................................................................................................26 

A4.0 Summary of Reference Mission Version 3.0....................................................................................28 

A5.0 Revolutionary Next Steps .................................................................................................................33 

A5.1 Solar Electric Propulsion...............................................................................................................33 

A5.1.1 Electric Propulsion Mission Concept......................................................................................34 

A5.1.2 Electric Vehicle Concepts .......................................................................................................35 

A5.2 Aerocapturing the TransHab .........................................................................................................36 

A5.3 Bimodal Nuclear Thermal Rocket (NTR) Propulsion...................................................................39 

5.3.1 Bimodal NTR Mission Concept................................................................................................39 

5.3.2 All Propulsive” Bimodal NTR Option Using TransHab............................................................41 

5.3.3 “LOX-Augmented” NTR Option...............................................................................................42 

5.3.4 Three-Magnum Scenario............................................................................................................43 

5.3.4.1 Mission Content ......................................................................................................................44 

5.3.4.2 Strategies and Technology Challenges ...................................................................................44 

5.3.4.3 Combination Lander Scenario.................................................................................................45 

5.3.4.4 Split Mission Scenario ............................................................................................................45 

5.3.5 All Solar Scenario ......................................................................................................................48 

A6.0 Continuing Work...............................................................................................................................50 
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 ii

List of Figures 

Figure A2-1 Cargo and Piloted Vehicles for Reference Mission Version 1.0 ......................... 3 

Figure A2-2 Reference Mission Version 1.0 Mission Sequence ............................................. 4 

Figure A2-3 Mars Surface Lander and Habitat Aeroshells for Version 1.0............................. 5 

Figure A2-4 Habitat Repackaging Strategy ............................................................................. 6 

Figure A2-5 Historical Space Habitat Pressurized Volume..................................................... 7 

Figure A2-6 Mars Surface Inflatable Habitat Concept............................................................. 8 

Figure A3-1 Payload Capability to 407 km.............................................................................. 17 

Figure A3-2 Magnum Launch Vehicle..................................................................................... 18 

Figure A3-3 Schematic of solid core NTR turbopump and power cycle ................................. 20 

Figure A3-4 NTR stage and aerobraked Mars payload for Version 3.0 .................................. 22 

Figure A3-5 Aeroassist Study Results for Version 3.0 ............................................................ 23 

Figure A3-6 Entry and Landing Parachute Study Results for Version 3.0 .............................. 25 

Figure A3-7 Launch and Packaging Configurations for Version 3.0....................................... 27 

Figure A4-1 Reference Mission Sequence for Version 3.0...................................................... 29 

Figure A5-1 Solar Electric Propulsion Mission Concept......................................................... 34 

Figure A5-2 SEP Crew Taxi Concepts..................................................................................... 36 

Figure A5-3 Conceptual Solar Electric Propulsion Vehicle..................................................... 37 

Figure A5-4 Launch manifest for the Solar Electric Propulsion Vehicle Concept .................. 37 

Figure A5-5 Potential Transhab Aerobrake Configurations .................................................... 38 

Figure A5-6 Schematic “Bimodal” NTR System..................................................................... 39 

Figure A5-7 “Bimodal” NTR Transfer Vehicle Option ........................................................... 40 

Figure A5-8 All Propulsive Bimodal NTR Carrying TransHab .............................................. 43 

Figure A5-9 Three-Magnum Combination Lander Scenario ................................................... 46 

Figure A5-10 Three-Magnum Combination Lander Launch Packages ..................................... 46 

Figure A5-11 Combination Lander Concept on Mars Surface .................................................. 47 

Figure A5-12 Three-Magnum Split Mission Scenario............................................................... 47 

Figure A5-13 Three Magnum Split Mission Launch Packages ................................................. 48 

Figure A5-14 Solar Irradiation at Mars...................................................................................... 49
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iii

List of Tables 

Table A3-1 Mass Reduction Benefits from the Transhab Study ............................................ 11 

Table A3-2 ISRU System Breakdown for Version 3.0........................................................... 13 

Table A3-3 Power System Improvements for Version 3.0 ..................................................... 13 

Table A3-4 Science Manifest for Version 3.0 ........................................................................ 15 

Table A3-5 Launch Vehicle Requirements............................................................................. 16 

Table A3-6 Payload Capability to 407 km.............................................................................. 19 

Table A4-1 Payload Mass Evolution from Version 1.0 to Version 3.0 .................................. 28 

Table A4-2 Earth Return Vehicle Mass Scrub for Version 3.0 .............................................. 30 

Table A4-3 Cargo Lander Mass Scrub for Version 3.0 .......................................................... 31 

Table A4-4 Piloted Lander Mass Scrub for Version 3.0......................................................... 32 

Table A5-1 Comparison of “Bimodal” NTR to Reference Mission Version 3.0 ................... 41 
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 iv

List of Acronyms 

ARC Ames Research Center 

EELV Evolved Expendable Launch 

Vehicle 

EEV Earth Entry Vehicle 

EMU EVA Mobility Unit 

EPM Electric Propulsion Module 

ERV Earth Return Vehicle 

EVA Extra Vehicular Activity 

Hab Habitat 

HEO High Earth Orbit 

HLLV Heavy Lift Launch Vehicle 

IMLEO Initial Mass in Low Earth Orbit 

ISRU In-Situ Resource Utilization 

JPL Jet Propulsion Laboratory 

JSC Johnson Space Center 

kg kilograms 

KSC Kennedy Space Center 

kWe kilowatts electric 

LaRC Langley Research Center 

lbf pound force 

L/D Lift-to-Drag Ratio 

LEO Low Earth Orbit 

LeRC Lewis Research Center 

LFBB Liquid Fly Back Booster 

LSS Life Support System 

m/s meters per second 

mt metric ton 

MAV Mars Ascent Vehicle 

MSFC Marshall Space Flight Center 

MTV Mars Transfer Vehicle 

NTP Nuclear Thermal Propulsion 

NTR Nuclear Thermal Rocket 

P/C Physical / Chemical 

PCRV Pressurized Control Research 

Vehicle 

PMAD Power Management and 

Distribution 

psi pounds per square inch 

PVA Photovoltaic Array 

RCS Reaction Control System 

SDV Shuttle Derived Vehicle 

SEP Solar Electric Propulsion 

SP Special Publication 

STS Space Transportation System 

TEI Trans-Earth Injection 

TMI Trans-Mars Injection 

TransHab Transit Habitat 

TPS Thermal Protection System 
EX13-98-036 

v

Foreword 

This Addendum to the Mars Reference Mission was developed as a companion document to NASA 

Special Publication 6107, “Human Exploration of Mars: The Reference Mission of the NASA 

Exploration Study Team.” The Addendum summarizes changes and updates to the Mars Reference 

Mission that were developed by the Exploration Office since the final draft of SP 6107 was printed in 

early 1998. 

The Reference Mission is a tool used by the Exploration Team and the exploration community to 

compare and evaluate approaches to mission and system concepts that could be used for human 

exploration missions. It is intended to identify and clarify system “drivers”, or significant sources of 

cost, performance, risk, and schedule variation. It does not represent a final or recommended approach 

to human Mars missions. Several alternative scenarios, including human exploration missions to the 

Moon, Asteroids, or other targets beyond Earth orbit as well as employing different technical 

approaches to solving mission and technology challenges, are currently under study by the Exploration 

Team. Comparing alternative approaches provides the basis for continual improvement to technology 

investment plans and general understanding of future human exploration missions. 

The Addendum represents a “snapshot” of work in progress in support of planning for future human 

exploration missions through May 1998. Annual publications of revisions to the Reference Mission are 

planned beginning in late 1998. 

Please direct all correspondence and inquiries about this document to: 

Exploration Office 

Attention: Reference Mission Data Manager 

Mail Code EX13 

NASA Johnson Space Center 

2101 Nasa Road 1 

Houston, Texas 77058-3696 
EX13-98-036

 vi
EX13-98-036 

1

A1.0 Introduction 

NASA Special Publication 6107 details the 

work of the Exploration Study Team through 

the spring of 1994∗ . As described in that 

report, the primary role of the Reference 

Mission is two-fold. First, it is used to form a 

template by which subsequent exploration 

strategies may be evaluated for consideration as 

alternate or complementary approaches to the 

human exploration of Mars. Second, the 

Reference Mission is intended to stimulate 

additional thought and development in the 

exploration community and beyond. 

In serving these two purposes, several 

components of the original Reference Mission 

(referred to as Version 1.0) have been modified 

to that which is presented in its current form 

Mars Reference Mission Version 3.0. The 

changes are manifested at the strategic, mission, 

and system levels of development, and augment 

or improve upon prior work done by NASA’s 

Exploration Study Team. To facilitate and 

document the ongoing work of the Exploration 

Team, this addendum will outline the current 

strategy (as of this addendum’s publication 

date) as well as provide a description of the 

current systems. Section two of this Addendum 

provides a brief overview of the changes to the 

reference approach which are strategic in nature, 

that is changes which cross many systems and 

elements. Section three provides a description 

of improvements to many of the individual 

systems and elements. Lastly, section four 

discusses several revolutionary mission 

approaches and technical options, currently 

under consideration by the exploration 

community, which can provide significant 

improvements in the mission architecture and 

mass estimates. 

A2.0 Strategic Modifications 

The original Reference Mission, compiled in 

the 1993-94 time frame, has been reviewed and 

improved in many facets of its design. 

Modifications to that strategy have been made 

to create a mission offering less risk, lower cost, 

and better technical approach than previous 

mission designs. This section will discuss the 

strategic modifications which have been made 

to the original Reference Mission, namely 

alteration of the launch strategy to greatly 

reduce the required size of the launch vehicle 

and revision of a mission strategy leading to the 

elimination of the initial habitat flight. 

 

∗

 See: www-sn.jsc.nasa.gov/marsref/ 
EX13-98-036

 2

A2.1 Reference Mission 1.0 Launch Strategy 

Perhaps the biggest assumption of the 

original Reference Mission centered on the 

launch system; specifically, a large, yet-to- bedeveloped launch vehicle was required to place 

the mission elements into low Earth orbit 

(LEO). The launch manifest for the mission 

elements is shown in Figure A2-1. As can be 

seen, a 200-metric- ton launch vehicle would be 

required to achieve a human mission in four 

launches. This scenario consists of three 

launches for the first trans-Mars injection (TMI) 

opportunity, followed by three launches at each 

subsequent opportunity. The first human 

mission consists of three cargo launches in the 

first injection opportunity followed by one 

piloted launch in the following opportunity, 

each manifested with the specific equipment as 

shown in the figure. 

To graphically illustrate how each of the 

four launches are conducted to support the first 

human mission, Figure A2-2 is provided. 

During the first mission opportunity in 2011, the 

three cargo vehicles are launched on a nearly 

Hohmann transfers from Earth to Mars. 

Reference Mission Version 1.0 was designed 

such that the Earth Return Vehicle (ERV-1), 

containing the return habitat, enters a parking 

orbit about Mars by utilizing an aerocapture 

maneuver upon arrival at Mars. The other two 

cargo elements, Cargo-1 and Hab-1, perform an 

aerocapture followed by aeroentry and landing, 

delivering the dry ascent vehicle and crew 

surface habitat to the Martian surface. These 

components are followed 26 months later (at the 

next injection opportunity in 2014) by a second 

surface habitat, Hab-2, piloted by a crew of six. 

The crew performs an aerocapture followed by 

aeroentry and landing to the surface in close 

proximity to the previously deployed surface 

assets (Cargo-1 and Hab-1). After completion 

of the 500-day surface mission, the crew 

ascends to Mars orbit and rendezvous in Mars 

orbit with the pre-deployed return vehicle 

(ERV-1). 

It was recognized that development of the 

large 200-metric ton launch vehicle posed a 

significant technology and development 

challenge to the mission strategy. Design of the 

large launcher raises several cost issues 

(development, new launch facilities, etc.), and 

the physical size of the launch vehicle is itself a 

potential limitation to implementing Version 1.0 

of the Reference Mission. The requirement of a 

heavy lift booster was driven primarily by the 

initial mass to Low Earth Orbit (IMLEO); 

therefore, an effort was initiated in the fall of 

1996 to reduce the required mass and volume of 

each launch. These efforts were undertaken 
EX13-98-036 

3

CARGO 1

Ascent Vehicle (Dry) & Lander

ISRU Plant & H2 Feedstock

Surface Power System

Pressurized &Rover

3 Tele-operated Science Rovers

Unpressurized Rover

DIPS Cart, Science

HAB 1

Hab Module and Lander

Surface Nuclear Power

Unpressurized Rover

500 Day Consumables

ERV 1

Return Hab LOX/CH4 TEI

180 + 500 Days Consumables

Piloted 1

Surface Hab and Lander

Unpressurized Rover

180 + 500 Days Consumables

15.0 m

20.6 m

4.7 m

16.3 m

20.6 m

4.7 m

19.0 m

20.6 m

4.7 m

16.3 m

20.6 m

4.7 m

205.1 t 216.6 t

246.6 t 208.2 t

Figure A2-1 Cargo and Piloted Vehicles for Reference Mission Version 1.0. 
EX13-98-036

 4 Opportunity 1 (2011): 3 flights Earth Opportunity 2 (2014): 1 flight

Return

Mars

Cargo-1

Return Habitat,

chemical TEI Stage

(to Mars orbit)

ERV-1

Crew of 6 Aerocaptures and Lands in Outbound Hab

Surface rendezvous with pre-deployed assets

Crew Ascends to 

Return Hab 

in capsule

Crew Direct Enters

 in capsule, Apollo-style.

Outbound Hab delivered to LEO on Large Launch Vehicle.

Crew of 6 delivered to LEO in Shuttle or other. 

Both rendezvous with NTR.

Cargo delivered to LEO on Large Launch

Vehicle, rendezvous with NTR.

Ascent Vehicle,

Prop Production,

Surface Exploration Gear,

(to surface)

Crew Hab/Lab

(to surface) Hab-1

Hab-2

MAV/EEV

Crew of returns to Earth

in Return Hab (ERV-1)

Figure A2-2 Reference Mission Version 1.0 Mission Sequence
EX13-98-036 

5

while balancing the need to minimize the 

number of launches to reduce ground launch 

costs and limit added operational complexity 

due to Low Earth Orbit (LEO) rendezvous and 

docking. In order to reduce the size of the 

launch vehicle, a critical examination of the 

payloads, in terms of their physical size and 

mass, was conducted. The goal of this 

modification was to remanifest the payload 

elements onto two smaller (80 metric ton class) 

launch vehicles rather than the single large 

vehicle. 

A2.1.1 System Repackaging 

Reducing the physical size of the launch 

elements is important from many aspects of the 

launch vehicle design, including reducing the 

mass of the systems and reducing the 

aerodynamic loads on the payload shroud. The 

geometry of the large (10 m diameter) aeroshell 

for the large launch vehicle, used for both the 

Mars lander and the surface habitat modules, is 

given in Figure A2-3. Of particular note is the 

unused volume between the lander / habitat and 

the aeroshell.

Figure A2-3 Mars Surface Lander and Habitat Aeroshells for Version 1.0. 
EX13-98-036

 6

A proposed solution to this excess volume is 

shown in Figure A2-4. In this design, the 

habitat structure is integrated with the Mars 

entry aeroshell and launch shroud. In addition 

to reducing the structural mass of the element, 

the integrated design serves several functions 

beyond those which were proposed in Version 

1.0 of the Reference Mission. Specifically, the 

integrated habitat / pressure hull with a thermal 

protection system (TPS): 

• serves as both an Earth ascent shroud and 

Mars entry aeroshell 

• eliminates the need for on-orbit assembly / 

verification of the aeroshell 

• allows for stowage in an 80-metric-ton-class 

launch vehicle. 

10 m 7.5 m

Figure A2-4 Habitat Repackaging Strategy. 

During the outbound and return interplanetary journeys, Reference Mission Version 

1.0 allows for 90 m3

 of pressurized volume per 

crew member. As can be seen in Figure A2-5, 

this value is consistent with data from previous 

space missions. It is desirable to maintain this 

living quality for the crew despite any 

subsequent changes which may occur to the 

original Reference Mission. 

A2.1.2 System Mass Reductions 

The second step in changing the launch 

strategy focused on reducing the system masses 

in order to reduce the mass delivery 

requirements for the launch system. The 

payload masses were critically examined, and 

any duplications were eliminated. In addition, 

studies were undertaken to scrub the system 

masses to achieve the required weight savings. 

The goal of this work was to reduce each 

payload delivery flight to accommodate the 

approximate volume and weight limitations of 

two 80-metric-ton launchers. These mass 

reductions are discussed in further detail in 

Section 3.0 of this Addendum. 
EX13-98-036 

7

Mission Duration (days)

0.1

1

10

100

1000

1 10 100 1000

Mercury

Voskhod

Apollo

LEM

Vostok Gemini

STS

Apollo

CM Soyuz

Skylab ISS

Salyut 7

Mir

Total

Pressurized

Volume

(m3)/crew

Long-Duration Spaceflight:

Ref. mission Total Vol. 90 m3/crew

V 1.0 (50)

20’ Test Chamber

Phase III (62)

Figure A2-5 Historical Space Habitat Pressurized Volume. 

A2.1.3 Modified Launch Strategy 

Reduction of the payload delivery flights’ 

mass and volume enables the opportunity to 

utilize a smaller launch vehicle. This 

repackaging allows the mission to change from 

a launch vehicle requiring a 200 metric-ton 

launch to two individual launches of magnitudes 

within the envelope of launch systems which 

can be evolved from current capabilities. This 

design, delivering the interplanetary propulsion 

system and cargo into Earth orbit separately, 

would require one rendezvous and docking 

operation prior to each outbound journey to 

Mars. While doubling the number of launches, 

this strategy eliminates the high costs of 

developing the large 200 metric-ton launch 

vehicle of Version 1.0. 

A2.2 Elimination of Initial Habitat Flight 

While reviewing the original mission 

strategy, the initial habitat lander (Hab-1) was 

identified as a launch component which could 

potentially be eliminated. During the Spring of 

1997 a team of engineers at the Johnson Space 

Center (JSC) investigated a concept of utilizing 

inflatable structures (known as the TransHab) 

instead of traditional hard aluminum structures 

for habitation systems (see Section 3 for more 

details of this concept). Results of this study 

demonstrated significant subsystem mass 

savings for the TransHab concept. Given the 

significant volume per unit mass increase 

provided by the inflatable TransHab concept, 

the attention of the Exploration Team returned 

to the launch packaging outlined in the original 
EX13-98-036

 8

Reference focusing on techniques of 

augmenting the surface living volume. 

A2.2.1 Volume Augmentation 

As noted earlier, a sufficient level of 

pressurized living volume is critical for crew 

health maintenance. A TransHab-derived 

inflatable structure would provide such 

augmentation, arriving on the Mars surface in 

the Cargo-1 flight two years before the crew. 

Pre-plumbed and ready for integration into the 

life support of the Piloted Crew Lander, the 

inflatable structure would simply need to be 

installed by the crew upon arrival, as depicted in 

Figure A2-6. 

The mass of the inflatable module 

(estimated at 3.1 metric-tons without crew 

accommodations or life support) could be 

substituted for the mass of the pressurized rover 

Figure A2-6 Mars Surface Inflatable 

Habitat Concept. 

(5 metric-tons) originally manifested on the 

Cargo-1 flight. The pressurized rover, deferred 

to the second Cargo delivery flight, would 

arrive a few months after the crew and would 

still be available for the majority of the mission. 

In essence, the redundancy of the pressurized 

rover (for the first Mars crew) has been traded 

for the elimination of an entire Mars-bound 

habitat flight. 

A2.2.2 Redundancy Considerations 

The concern that system redundancy would 

be reduced with the elimination of Hab-1 was 

mitigated by the redundancy already built into 

the Reference Mission. For example, several 

levels of redundancy are present in the mission 

architecture to address failure of the 

regenerative lift support system of a habitation 

module. Four levels of this redundancy are 

outlined below. 

• First level backup - In-Situ Resource 

Utilization processes generate enough water 

and oxygen for the entire surface mission to 

run “open loop.” 

• Second level backup - The Ascent Vehicle / 

ISRU plant on Cargo-2 of the subsequent 

mission, arriving to the surface a few 

months after the crew, could be used to 

supply life support rather than for propellant 

production. 
EX13-98-036 

9

• Third level backup - The surface could be 

abandoned for the orbiting Earth Return 

Vehicle, which has a sufficient food cache 

to last until the next trans-Earth injection 

window. 

• Fourth level backup - The Earth Return 

Vehicle (ERV-2) of the subsequent mission, 

arriving a few months after the crew, would 

provide an additional refuge for the crew if 

necessary. 

A2.3 Revised Mission Strategy for Version 3.0 

The strategic modifications to the Reference 

Mission described in this section have 

significantly reduced many of the barriers faced 

during the formulation of the original approach. 

The combination of repackaging the mission 

elements into smaller launch vehicles along 

with elimination of the initial habitat lander has 

allowed significant reduction in launch vehicle 

size, from 200-metric tons down to 80-metric 

tons, while only introducing two additional 

flights to the overall launch manifest. 
EX13-98-036

 10

A3.0 System Design Improvements 

Point of Contact: Bret Drake/JSC 

In order to accomplish the strategic changes 

discussed in the previous section, improvements 

to the system designs were required, specifically 

in terms of system mass reductions. 

Modifications to the systems were accomplished 

by the Exploration Study Team (Johnson Space 

Center, Marshall Space Flight Center, Lewis 

Research Center, Ames Research Center, 

Kennedy Space Center, and Langley Research 

Center), the JSC TransHab study team, and by 

the Human/Robotic Exploration Team (led 

jointly by JPL and JSC). Specific 

improvements include: 

• Incorporation of TransHab system designs 

• Mass scrub of many of the systems 

• Improvements of the transportation system 

designs 

A3.1 Incorporation of TransHab 

Improvements 

Point of Contact: Donna Fender/JSC 

In an effort to reduce the cost of human 

habitation in space, a group from the 

Engineering Directorate at JSC has been 

studying an economic and innovative habitation 

concept based on inflatable structure 

technology. In the spring of 1997, the 

improvements associated with the TransHab 

effort were identified as potential habitat 

options. Many of the subsystem improvements 

could be incorporated into the Reference 

Mission for both the interplanetary and surface 

phases of the mission. 

The Exploration Team has been working to 

quantify improvements identified in the 

TransHab study, specifically environmental life 

support system and structural improvements. It 

is important to note that the Reference Mission 

architecture and crew size has remained 

unchanged with the incorporation of the 

TransHab option. Some of the masses used by 

the TransHab team, however, have been scaled 

to match the duration of the Mars Reference 

Mission.

Advantages of incorporating the TransHab 

study into the current Mars exploration strategy 

are manifested primarily in mass reductions. 

These benefits are provided in Table A3-1. The 

results are presented for both the Piloted Crew 

Lander Surface Habitat and the Earth Return 

Vehicle, and are given in terms of percent 

changes from the Version 1.0 Reference 

Mission. Many of the subsystem mass 

estimates taken from the TransHab studies were 

of higher fidelity than those previously used by 

the Mars 
EX13-98-036 

11

Piloted Crew Lander

"Scrubbed Ref. Mission"* Version 3.0 Delta

Habitat Element 33657 kg 19768 kg -13889 kg -41% Mass Reduction due to included systems

P/C LSS 3000 4661 1661 55% 2778 dry plus 1883 fluids

Crew Accommodations 16157 11504 -4653 -29% Normalized to 680 days

EVA Equipment 1000 969 -31 -3% Normalized to Six EMUs

Comm/info management 1500 320 -1180 -79%

Power Dist. 500 275 -225 -45%

Thermal 2000 500 -1500 -75% "Mostly" included in LSS Mass

Structure 5500 1039 -4461 -81%

Crew 500 500 0 0%

Spares 3500 0 -3500 -100% Spares accounted for in elements

3kWe PVA/RFC Keep-Alive 1700 1700 0 0%

Unpressurized Rovers (3) 440 500 60 14% Payload assumed

EVA Consumables 2300 2300 0 0%

Crew+EVA Suits** 1300 0 -1300 -100% Double bookkept - should be deleted

Total Payload Mass 39397 kg 24268 kg -15129 kg -38%

Terminal Propulsion System 4200 4200 0 0%

Total Landed Mass 43597 kg 28468 kg -15129 kg -35%

Terminal Propellant 10800 7052.2 -3748 -35%

Aerobrake (15%) 8160 5328 -2832 -35%

Mars Entry Mass 62557 kg 40848 kg -21709 kg -35%

*As presented 1/13/97

**Double bookkept - should be deleted

Earth Return Vehicle

"Scrubbed Ref. Mission"* Version 3.0 Delta

Habitat Element 31395 kg 21615 kg -9781 kg -31% Mass Reduction due to included systems

P/C LSS 2000 4661 2661 133% 2778 dry plus 1883 fluids

Crew Accommodations 13021 10861 -2160 -17% 180d Consumables + 500 d add'l food

EVA Equipment 500 485 -15.5 -3% Normalized to Three EMUs

Comm/info management 1500 320 -1180 -79%

Power Dist. 500 275 -225 -45%

30 kWe PVA Power 2974 2974 0 0%

Thermal 2000 500 -1500 -75% "Mostly" included in LSS Mass

Structure 5500 1039 -4461 -81%

Science Equipment 900 500 -400 -44% Payload assumed

Spares 2500 0 -2500 -100% Spares accounted for in elements

Jettison Excess Consumables -6600 -6600 0 0%

Dock Earth Entry Vehicle & P/L 6900 6900 0 0%

ERV Mass at TEI 31695 kg 21915 kg -9781 kg -31%

TEI Dry Stage Mass 3500 3500 0 0%

Earth Return RCS Prop 1100 1100 0 0%

Earth Return Prop 31800 23231 -8569 -27%

Aerobrake (15%) 10169 7417 -2752 -27%

Mars Orbit Insertion Mass 77964 kg 56862 kg -21102 kg -27%

*As presented 1/13/97

Table A3-1 Mass Reduction Benefits from the TransHab Study. 
EX13-98-036

 12

Exploration Team, which accounts for some of 

the increases in mass values. 

A3.2 System Improvements 

A3.2.1 In-Situ Resource Utilization 

Points of Contact: Jerry Sanders and Todd 

Peters/JSC 

The fidelity of the In-Situ Resource 

Utilization system designs were improved 

during the Fall 1997 design cycle. An improved 

system design tool was developed which 

incorporates options and sizing routines for 

different products (fuels, oxidizers, water for 

life support, etc.), production processes, 

cryogenic fluid cooling, and tank sizing. With 

the increased fidelity of the model, the ISRU 

system mass estimates were adjusted downward 

for the plant itself (from 4802kfrom Reference 

Mission Version 1.0 to 3941kg) and upward for 

the hydrogen feedstock (from 4500kg to 

5420kg). These estimates reflect a plant that 

will produce both the ascent propellant and a 

surface life support system consumables water 

cache (23 metric tons). The power requirement 

for the In-Situ Resource Utilization system is 

driven by both the quantity of products required 

and the time required to produce the products. 

Sufficient time for product production is 

provided such that all required consumables are 

produced and stored in the surface systems prior 

to crew departure from Earth. Given these 

groundrules, the current estimate for the power 

required is on the order of 45 kWe. Further 

details describing the mass and power 

breakdown for the ISRU system are provided in 

Table A3-2 . 

A3.2.2 Power Systems 

Point of Contact: Bob Cataldo/LeRC 

During recent analysis efforts, the surface 

power system design was revisited in order to 

obtain mass and cost savings from the original 

system design. The Reference Mission Version 

1.0 surface power system design was based on 

the reactor technologies developed within the 

SP-100 program, however with 3-80 kWe 

closed Brayton cycle (CBC) engines operating 

at 1100 K. Numerous system trades about this 

original design were conducted considering 

power needs, radiation shielding, reactor types, 

operating temperatures, power conversion 

technologies, recuperation efficiencies, power 

distribution voltage, inlet temperature, and 

number of spare power engines. Updates to the 

original analysis, including operation at turbine 

inlet temperatures of 1300 K, enabled a 

reduction in overall system mass from 14.0 to 

10.7 metric tons. Although this assumes a 

temperature increase of approximately 150 K 
EX13-98-036 

13

beyond current Brayton technology, required 

reactor and fuels technologies remain consistent 

with those developed within the SP-100 and 

other DOE/NASA programs. In addition, a first 

order assessment of the mass impacts of 

utilizing the same reactor technology as the 

propulsion system was performed. If feasible 

and practical, only one development program 

would then be required for both the propulsion 

and power systems. A power system based 

upon a gas-cooled nuclear thermal propulsion 

engine was estimated to have a mass of 

Subsystem Mass Subsystem Power

Propellants Life Support Propellants Life Support

Compressor 496 kg 193 kg 5645 W 2893 W

Sabatier Reactor 60 kg 50 kg 0 W* 0 W*

Hydrogen Membrane Separator 29 kg 23 kg 288 W 225 W

Methane Water Separator 394 kg 315 kg 1690 W

Pyrolysis Unit 711 kg 1172 kg 3397 W 3911 W

Electrolysis Unit 277 kg 18734 W

Oxygen Liquefier 43 kg 2215 W

Methane Liquefier 41 kg 2093 W

Subtotal 2051 kg 1753 kg 32371 8719 W

Total System 3,805 kg 41,091 W

* Reaction is exothermic requiring startup power only (~10 kWe for 1 hour)

Table A3-2 ISRU System Breakdown for Version 3.0 

DRM Version 1.0 DRM Version 3.0

Reactor type SP-100 SP-100 

Gas cooled (common 

with propulsion system 

technology)

Heat transport method Liquid metal to gas Liquid metal to gas Direct

Power conversion 3- 1140 K CBC 3- 1300 K CBC 3- 1300 K CBC

Shield 4-pi, 5 REM/yr @ 360° 4-pi, 5 REM/yr @ 90°, 

50 REM/yr @270°

4-pi, 5 REM/yr @ 90°, 

50 REM/yr @270°

Distance from Base 2.0 km 2.5 km 2.8 km

Distribution line voltage 2000 V 5000V 5000V

System mass (mt) 12.2 9.3 10.5

Deployment cart (15%) 1.8 1.4 1.6

Total 14.0 10.7 12.1

Table A3-3 Power System Improvements for Version 3.0. 
EX13-98-036

 14

12.1 metric tons. Some of the more salient 

features of the three designs are shown in Table 

A3-3. Currently Reference Mission Version 3.0 

carries the heavier mass of the gas cooled 

reactor system. 

In addition to the system designs discussed 

above, other system level trades are being 

conducted. For instance, additional mass 

savings could result by using indigenous 

shielding materials such as soil and/or 

condensed CO2. The use of indigenous 

shielding would minimize the system mass 

differences shown in the table, since the shield 

mass is the major component of system mass 

variation. 

These concepts are being evaluated for their 

impact on the power system design itself as well 

as other systems that might be required to 

support this concept, such as, mobile equipment 

or refrigeration systems. In addition, smaller 

reactor concepts, such as a 50 kWe power 

system, have been assessed resulting in a total 

system mass as low as 5.6 metric tons for an SP100 based system. These smaller reactor 

concepts could be used for the initial mission 

phases, with multiple units providing higher 

power levels for more robust exploration 

activities, such as food production. 

A3.2.3 Science Systems 

Point of Contact: John Gruener/JSC 

A review of the science components for the 

Reference Mission was conducted during the 

Fall of 1997. The emphasis of this activity was 

to critically review the science manifest, seeking 

mass savings. The focus of the review was not 

to change the science strategy, but merely to 

seek methods of reducing the science manifest 

mass estimates. It is desirable to maintain a 

balance between mass reduction and science 

content. Due to the time limitations of the 

study, it was not possible to conduct detailed 

system designs for the various scientific 

instruments, instead, emphasis was placed on 

understanding the current science content as it 

pertains to previous systems designs and 

removing any undefined system content (50 kg), 

unnecessary undefined margins (250 kg), and 

undefined discretionary science (300 kg). A 

detailed science manifest of the first human 

mission for Reference Mission Version 3.0 is 

shown in Table A3-4. 
EX13-98-036 

15

Surface Science Equipment* DRM 1.0 DRM 3.0

 Field Geology Package 335 kg 300 kg 35 kg not accounted for

 Geoscienc Laboratory Eq. 125 kg 110 kg 15 kg not accounted for

 Exobiology Laboratory 50 kg 50 kg No change

 Traverse Geophysical Inst. 400 kg 275 kg 125 kg discretionary margin removed

 Geophysical/Meterology Inst. 200 kg 75 kg 125 kg discretionary margin removed

 10-Meter Drill 260 kg 260 kg No change

 Meterology Balloons 200 kg 200 kg Needs better definition

 Biomedical/Bioscience Lab 500 kg 500 kg Needs better definition

 Discretionary Science 300 kg 0 kg Removed

Total 2370 kg 1770 kg

Cruise Science Equipment*

 Particles & Fields Science 100 kg 100 kg No change

 Astronomy Instruments 200 kg 200 kg Estimate only

 Small Solar Telescope 100 kg 100 kg No change

 Biomedical Instruments 200 kg 200 kg Needs better definition

Total 600 kg 600 kg

* NASA Reference Publication 1345

Table A3-4 Science Manifest for Version 3.0. 

A3.2.4 EVA Systems 

Point of Contact: Robert Yowell/JSC 

The EVA consumables estimates for 

Reference Mission Version 3.0 were improved 

through the incorporation of a parametric sizing 

algorithm developed during the TransHab study. 

In addition, to gain further reductions an 

assumption was made that consumable mass 

would only be allocated for two emergency 

EVAs during transit, allowing for two, eighthour EVAs, performed by two crew. This 

resulted in a mass of 48 kg for the transit 

phases. The transit vehicles also include195 kg 

each of EVA support equipment (airlock, 

airlock systems, EMU spares). 

System synergism was also incorporated to 

gain further mass reductions for the surface 

phase of the mission. EVA consumable 

requirements were included in the sizing of the 

In-Situ Resource Utilization system such that 

additional oxygen was produced by the ISRU 

system to provide the necessary consumables 

for routine surface EVA exploration. Utilizing 

the locally produced oxygen could save 

approximately five metric tons for the surface 

phase of the mission alone (10.9 kg per twoperson eight-hour EVA). The current EVA 

consumable estimates are sufficient for one, 

eight-hour EVA per week, performed by two 
EX13-98-036

 16

crew. Additional consumables for a more 

robust exploration scenario, including food 

sticks, batteries, drink bags, visors, etc., but not 

oxygen which the ISRU provides, have not been 

included in the EVA estimates for Reference 

Mission 3.0. Estimates for the additional 

ancillary consumables, for more robust EVAs, 

will be incorporated in the next version of the 

Reference Mission. These changes resulted in a 

total of 446 kg for the surface phase of the 

mission. Therefore, the total mass of the EVA 

consumables is currently estimated at 932 kg 

versus 3000 kg in Version 1.0. Further 

examination of the assumptions used to reduce 

these masses is underway. 

The EVA dry mass was slightly reduced 

from 1000 kg to 940 kg based on inputs from 

the EVA Project Office at Johnson Space 

Center. This reflects a mass of 156 kg per suit. 

A3.3 Transportation System Improvements 

Point of Contact: Steve Richards/MSFC 

Re-examination of the performance and 

design characteristics of the transportation 

elements for the Reference Mission were led by 

engineers at the Marshall Space Flight Center 

with support from the Lewis Research Center, 

Ames Research Center, Langley Research 

Center, and Kennedy Space Center. Major 

modifications to the transportation elements, 

resulting in Reference Mission Version 3.0, are 

discussed in this section. 

A3.3.1 Earth-to-Orbit Transportation 

Points of Contact: Bill Eoff and David 

Smith/MSFC 

Human Mars mission launch costs are 

driven by initial mass in low-Earth-orbit 

(IMLEO); launch costs per pound of payload; 

launch vehicle development costs; and on orbit 

assembly costs. Earth-to-Orbit (ETO) metrics 

identified in DRM 3.0 required launch vehicle 

payload capability of 80 metric tons to minimize 

on orbit assembly costs and meet payload size 

requirements. Cost metrics of less than $1000 

launch cost per pound of payload and total 

mission costs of $6B for any launch vehicle 

development costs and all launch recurring costs 

have been designated as reasonable starting 

requirements to drive system designs, see Table 

A3-5. 

Reference Mission 3.0 Payload Requirements

 P/L Diameter: 7.5 m / 24.8 ft

 P/L Length: 27.7 m / 91.4 ft

 P/L Weight: 80 mt / 176 Klb

 Assembly Orbit (28.5 deg)407 km / 220 nmi

 Launch Rate: 6 per year

Table A3-5 Launch Vehicle Requirements 

During the design cycle for Reference 

Mission Version 1.0 numerous configurations 

were considered and a Shuttle derived vehicle 
EX13-98-036 

17

(SDV) with an inline core vehicle was selected. 

The SDV launch concept barely meets the $6B 

cost metric for total mission ETO costs because 

of the high core vehicle costs for Shuttle 

common hardware. In addition, recent analysis 

indicated that the SDV configuration exceeded 

the $1000/lb metric by a factor of two. 

Launch vehicle assessments for Reference 

Mission 3.0 focused on evaluating a core 

vehicle that is not Shuttle derived to decrease 

launch costs. Advances in launch vehicle 

technologies from the Reusable Launch Vehicle 

(RLV) and Evolved Expendable Launch 

Vehicle (EELV) programs could make it cost 

effective to develop a core vehicle that would 

potentially reduce the $6B ETO cost metric to 

$2.5B or less per current estimates. This new 

vehicle concept has been designated as 

“Magnum” to differentiate from the numerous 

other past launch vehicle studies. The current 

Magnum configuration is an inline core vehicle 

with two attached Shuttle boosters. The 

payload is aft mounted on the expendable core 

vehicle; a similar configuration as Titan IV but 

with over five times the payload capability for 

one third the launch costs, as shown in Figure 

A3-1. 

80 22 21 18 14

0

20

40

60

80

Magnum

EELV

VentureStar

Shuttle

Titan

Metric Tons

Figure A3-1 Payload Capability to 407 km. 

The Magnum vehicle configuration includes 

a core component which is 8.4 meters (27.5 ft) 

in diameter, the same as the Shuttle External 

Tank, to allow common use of Shuttle boosters 

and launch facilities, see Figure A3-2. By using 

Shuttle launch facilities and the proposed Liquid 

Fly Back Boosters (LFBB), recurring costs is 

estimated to be less than $1000 per pound of 

payload. A composite shroud is used to protect 

the payload during ascent and a small kick stage 

is used for circularizing the orbit. The current 

design of the Magnum launch vehicle provides a 

delivery capability of 85 metric tons (188 KLB) 

to 407 km (220 nmi) orbits at 28.5 degrees 

inclination or 80 metric tons (176 KLB) to 51.6 

degree inclination orbits. See Table A3-6 for 

additional Magnum performance data.
EX13-98-036

 18

Magnum Launch Vehicle

Liquid Flyback 

Booster (2)

Payload Fairing 

92 ft cyl x 25 ft I.D.

Circularization 

Stage

LO2 Tank

LH2 Tank

RS 68 

Engines (2)

Payload / Stage 

Adapter

Fwd Booster Attach

Aft Booster 

Attach

Thrust 

Structure

Figure A3-2 Magnum Launch Vehicle. 

Technology development and 

demonstrations for the Magnum launch vehicle 

concept are driven by the large vehicle size and 

low life cycle cost requirements. Current 

evaluations are focused on maximizing the costeffective application of technologies for 

engines, valves, composite tanks/structures, and 

other hardware or facilities under development 

or projected to be available on other programs 

such as RLV or EELV. The proposed Magnum 

technology development program would 

physically extend these technologies to fit 

Magnum. Tasks would need to be conducted to 

demonstrate 8.8 meter (27.5 ft) diameter 

composite fuel tank manufacturing techniques 

derived from techniques developed on 

substantially smaller tanks for RLV. Equivalent 
EX13-98-036 

19

Performance

or Mass (mt)

Inclination 

28.5°

Inclination 

51.6°

Shroud Drop @ 400 Kft 85.4 79.9

Integrated Shroud/Aerobrake

 Payload Only 75.7 70.2

 Shroud / Aerobrake Wt. 13.6 13.6

 Total Injected 89.3 83.7

Table A3-6 Payload Capability to 407 km. 

tasks would be conducted to demonstrate large 

composite shrouds using the Advanced Grid 

Stiffened (AGS) composite shroud 

manufacturing techniques first developed for 

EELV by the USAF Phillips Lab. Composite 

structures, propellant ducts and valve 

technologies would also need to be 

demonstrated. 

Though the Magnum configuration using 

LFBBs was selected to drive technology 

developments, the Magnum configuration is still 

open for assessment of alternate boosters, 

engines, etc. which would meet requirements. 

A3.3.2 Trans-Mars Injection 

Point of Contact: Stan Borowski (LeRC) 

A high performance trans-Mars injection 

(TMI) system is required to propel the cargo 

and piloted spacecraft payloads from their LEO 

assembly orbits to the desired trans-Mars 

trajectories and to stay within the mass (~80 

metric tons) and payload dimension (~7.6 m 

diameter x ~28 m length) limits of the Magnum 

launch vehicle. For Reference Mission 3.0 the 

solid core nuclear thermal rocket (NTR) was 

used for the Trans-Mars Injection stage. Other 

alternatives, such as a Solar Electric Propulsion 

concept, are currently under investigation as 

discussed in Section 5. 

Conceptually, the NTR engine is relatively 

simple (Fig. A3-3). High pressure hydrogen 

propellant flows from the turbo pumps cooling 

the nozzle, reactor pressure vessel, neutron 

reflector, control drums, core support structure 

and internal radiation shield, and in the process 

picks up heat to drive the turbines. The 

hydrogen exhaust is routed through coolant 

channels in the reactor core’s fuel elements 

where it absorbs the energy released by 

fissioning uranium atoms. The propellant is 

superheated (to 2,700-3,100 K), and then 

expanded out a supersonic nozzle for thrust. 

Controlling the NTR engine during its 

operational phases (startup, full thrust, and 

shutdown) is accomplished by matching the 

turbo pump-supplied hydrogen flow to the 

reactor power level. Control drums, located in 

the surrounding reflector region, regulate the 

number of fission-released neutrons that are 

reflected back into the core. An internal neutron 

and gamma radiation shield, containing interior 

coolant passages, is also placed between the 

reactor core and sensitive engine components to 
EX13-98-036

 20

prevent excessive radiation heating and material 

damage. 

Nozzle Reflector Control

Drum

Turbines

Pumps

LH2

Tank

Reactor

Radiation

Shield

Figure A3-3 Schematic of solid core NTR 

turbopump and power cycle. 

The TMI stage used in Reference Mission 

3.0 employs three 15 thousand pounds force 

(klbf) NTR engines, each weighing 2224 kg, for 

an engine “thrust-to-weight” ratio of ~3.1. The 

TMI stage utilizes a “tricarbide” fuel material 

composed of a solid solution of uranium, 

zirconium and niobium ceramic carbides. This 

fuel has been developed and extensively tested 

in Russia. During reactor tests, hydrogen 

exhaust temperatures of ~3100 K have been 

reported for run times of over an hour. For exit 

temperature in the range of 2900-3075 K, 

specific impulse values of ~940-960 seconds are 

estimated for the tricarbide NTR engine 

assuming a chamber pressure of 2000 psia, a 

nozzle area ratio of 300 to 1, and a 110% bell 

length nozzle. 

A “common” TMI stage design has been 

defined for both the Mars cargo and piloted 

missions. The single tank stage is sized for the 

energetically demanding 2009 fast transit 

piloted mission opportunity and is therefore 

capable of injecting heavier surface and orbital 

payload elements on minimum energy Mars 

cargo missions. The NTR TMI stage and its 

aerobraked Mars payloads are illustrated in Fig. 

A3-4. The TMI stage LH2 tank is cylindrical 

with √2/2 ellipsoidal domes. It has an inner 

diameter of 7.4 meters, an ~20 meter length, and 

a maximum LH2 propellant capacity of ~54 tons 

assuming a 3% ullage factor. The main TMI 

stage component is the LH2 tank which is 

covered by a 2 inch multilayer insulation (MLI) 

thermal protection system that minimizes 

propellant boiloff in low Earth orbit to ~0.043 

kg/m2/day. Avionics, fuel cell power, storable 

reaction control system and docking systems are 

located in the stage forward cylindrical adapter 

section. Rearward is the stage aft skirt, thrust 

structure, propellant feed system and NTR 

engines. The total TMI stage “dry mass” is 

estimated at ~23.4 metric tons and assumes the 

use of composite materials for the propellant 

tank and all primary structures. For the piloted 

mission, an external disk shield is added to each 

engine to provide crew radiation protection 

which increases the stage dry mass by ~3.2 

metric tons. 
EX13-98-036 

21

The cargo and piloted Mars spacecraft 

depart LEO using a “2-perigee burn” Earth 

departure scenario to reduce gravity losses 

however single burn departures are also easily 

accommodated. The total engine burn time for 

the TMI maneuver is ~35 minutes--about half 

that demonstrated in the Russian reactor tests. 

The common TMI stage can inject ~74 and 61 

metric tons of payload to Mars on each cargo 

and piloted mission, respectively. The range of 

initial mass in Low-Earth Orbit varies from 

~135 to 148 metric tons and the overall vehicle 

length is ~50 meters. Following the TMI 

maneuver and an appropriate cooldown period, 

the aerobraked Mars payload and spent TMI 

stage separate. The storable bipropellant RCS 

system onboard the TMI stage is then used to 

perform the final midcourse correction and 

disposal maneuvers which place the TMI stage 

onto a trajectory that will not reencounter Earth 

over the course of a million years. 

A3.3.3 Aeroassist 

Points of Contact: Jim Arnold and Paul 

Wercinski/ARC 

The purpose of the Summer/Fall 1997 

aeroassist study was to develop and end-to-end 

conceptual design for human aeroassist vehicles 

consistent with Reference Mission Version 3.0 

payloads and configurations. The emphasis of 

the study was to develop a reliable mass 

estimate for the aerobrake as well as to provide 

a better understanding of the technologies 

required for the eventual development of an 

aeroassist capability. 

The Aeroassist Summer/Fall study used the 

Design Reference Mission Version 3.0 Piloted 

Vehicle mission and trajectory for sizing the 

entry vehicle for aerocapture and descent from 

orbit. This trajectory had Mars entry speeds of 

7.6 km/s, consistent with a 180-day transit in 

one particular opportunity. A triconic aerobrake 

shape was chosen as a baseline to accommodate 

packaging requirements of the payload 

elements. It was determined that the triconic 

shape had sufficient lift-to-drag (L/D) capability 

to meet aerocapture and descent to surface 

requirements. An L/D = 0.6 was selected for a 

trim angle of attack of 47 degrees. The 

aerocapture at Mars was performed without 

exceeding the 5g maximum deceleration limit 

which is necessary to maintain crew health and 

performance during the aerobraking maneuver. 
EX13-98-036

 22

Figure A3-4 NTR stage and aerobraked Mars payload for Version 3.0 

Several Navier-Stokes 3-dimensional 

flowfield solutions were calculated for this 

shape using appropriate CO2 chemistry for 

reacting flows to perform a preliminary thermal 

protection system (TPS) sizing and trade study 

for an overshoot trajectory. Turbulent heating 

estimates were also performed and were 

identified as a large contributor to uncertainties 

in predicting heating distribution over the 

triconic vehicle. Aerodynamic trim was 

calculated as well and a center-of-gravity 

location near 49-53% length from the nose was 

needed for trim. Radiation from the shock layer 

was also estimated and found to be highly 

dependent on the reacting gas chemistry models 

used. Peak heating rates near the nose region 

were found to between 150-250 W/cm2. 

Turbulent flows can result in even higher 

heating rates downstream. For higher entry 

velocities, at 8.4 km/s, peak heating rates above 

350 W/cm2 were modeled, but more analysis is 

needed due to the higher contributions of 

radiative heating associated with higher entry 

speeds. Dust erosion effects were also studied 

and are expected to not be as large of an effect 

on TPS mass estimates in comparison to 

turbulent flow or radiative heating issues. 

Heatshield structure was only estimated by 

analogy with structure estimates for a Magnum 

shroud. Heatshield mass estimates (TPS and 
EX13-98-036 

23

Altitude vs Time, Aerocapture Trajectory(Ve=8.5 km/s)

POC: Mark Loomis, Paul Wercinski 650-604-6578, NASA-ARC

Simple Biconic

Base Diameter=7.5 m

Length=12.75 m

Temp. Deg. F

Mach Number

Surface Temperature and Flow Field

Mach Number 

8 Species, Real Gas Calculation (GASP)

95% C02, 5% N2, Laninar Flow

Non Ablating, Fully Catalytic Surface

140

120

100

80

60

40

20

0

Altitude, km

0 100 200 300 400 500 600

Time, s

 Overshoot Trajectory

Computed Point (V=4.7 km/s)

Figure A3-5 Aeroassist Study Results for Version 3.0 

structure) yielded mass fractions ranging from 

16 – 18% of the total entry vehicle mass. These 

estimates were used for an entry vehicle 

carrying 51 metric-tons of cargo. During the 

aeroassist study, emphasis was not only placed 

on developing a conceptual approach for human 

aeroassist, but effort was also devoted to 

determining key technologies required for 

aeroassist. The following technology needs the 

were identified from this study: 

• Robust 3D Conceptual Fluid Dynamic code 

capable of radiating, turbulent, and dusty 

flows 

• Reliable reacting rate/transport and radiation 

models 

• Transition and turbulent models 

• Validation methods 

• Guidance Navigation &Control. options on 

approach, L/D > 0.3 guidance capability, 

terminal descent and landing 

• “Human” rated TPS 

• 2D TPS sizing tools 

• Arc-jets for CO2 flows 

• Flight validation of TPS materials 

• High-fidelity integrated design tools 

supported by local experts across agency. 

A3.3.4 Descent and Landing 

Points of Contact: Carol Dexter and Larry 

Kos/MSFC, and Michelle Munk/JSC 
EX13-98-036

 24

Major changes to the descent system for 

Reference Mission 3.0 include: 1) improved 

estimates of the descent phase using parachutes, 

and 2) elimination of the lander mobility 

requirement. 

The descent and landing scheme in Version 

1.0 included the use of parachutes with a final 

landing delta-V of 1000 m/s. The entry to 

landing phase of the mission was re-examined 

in Version 3.0 and now includes a higher 

fidelity method which incorporates mass 

reductions. Preliminary results were obtained 

from combining a 3-degree-of-freedom entry 

simulation and a basic sizing algorithm. In the 

simulation, the Cargo-1 vehicle, the most 

massive lander, was deorbited and flown 

through the atmosphere. Viking-type 

parachutes were then deployed at about 8 km 

altitude when the vehicle was traveling roughly 

700 m/s. The sizing algorithm was then used 

parametrically determine the number and size of 

parachutes and engines required for three 

different target altitudes. The masses of the 

parachutes, engines, fuel, and aerobrake were 

calculated in the sizer and the total vehicle mass 

was used as the performance metric. The data 

generated in this analysis are shown in Figure 

A3-6. A comparison of the new vehicle using 

the parachute scheme versus the vehicle using 

the all-propulsive scheme showed a potential 

savings of ten metric tons. Further analysis of 

this descent and landing approach includes: 

• Verifying the results with an integrated 

simulation 

• Assessment of supersonic deployment of a 

cluster of large (on the order of 50-mdiameter) parachutes 

• Determination of vehicle dynamics 

• Consideration of aborts, engine-out 

situations, and hazard avoidance 

requirements 

Reference Mission Version 1.0 included the 

capability of the descent system to perform 

limited surface mobility. This capability was 

provided so that the two surface habitats could 

be brought together and essentially “docked” to 

integrate the livable volume for the crew. With 

the deletion of the initial habitat, the descent 

system surface mobility mechanisms are not 

required, thus significantly reducing the 

complexity of the descent system design. 
EX13-98-036 

25

60000

62000

64000

66000

68000

70000

72000

74000

76000

78000

0 50 100 150 200 250 300 350 400

Terminal Velocity (m/s)

Total System Mass (kg)

5 km Altitude

3 km Altitude

1 km Altitude

12,4

7,4

4,4

3,5 2,5 2,6 2,6

1,7

1,8

1,10

1,11

1,18

1,21

9,4

5,4

3,5 3,5 2,6 2,7

1,22

1,27

1,33

1,40

1,47

1,56

1,65

7,5

4,6

3,8

2,11

2,14

1,18

1,9 1,10

1,12

1,14

1,13

1,12

1,16

Figure A3-6 Entry and Landing Parachute Study Results for Version 3.0 

Given the improvements in the entry and 

landing scenario and deletion of the surface 

mobility requirements, the descent system was 

refined. The descent system employs four 

RL10-class engines modified to burn LOX/CH4. 

These are used to perform the post-aerocapture 

circularization burn and the final 632 meters per 

second of descent velocity change after 

parachute deployment. The descent engines are 

also used for orbital correction maneuvers 

during the transit from Earth to Mars, the orbit 

adjust and trim maneuvers after aerocapture, 

and the de-orbit burn prior to the atmospheric 

entry and landing. Architecture definitions for 

the engine include: 

• Specific impulse of 379 seconds 

• Mixture ratio of 3.5 

• Chamber pressure of approximately 600 psi 

• A nozzle area ratio of approximately 400 

• Thrust level of approximately 15,000 lbf. 

• Additional requirements are that the engines 

be capable of throttling and gimbaling 

although specific ranges for these 

parameters have not been determined. 

The descent system for Reference Mission 

Version 3.0 is capable of placing approximately 
EX13-98-036

 26

40 metric tons of cargo on the surface. The dry 

mass of this system is approximately 4.9 metric 

tons requiring 11 metric tons of propellant. 

A3.3.5 Ascent 

Points of Contact: Carol Dexter and Larry 

Kos/MSFC 

The major modification of the ascent stage 

for Reference Mission Version 3.0 is the 

incorporation of a common descent/ascent 

propulsion system approach. The ascent stage 

propulsion system shares common engines and 

propellant feed systems with the descent stage. 

This eliminates the need for a separate ascent 

propulsion system reducing the overall mass 

and subsequent cost. These common engines 

are the same RL10-class engines modified to 

burn LOX/ CH4 as the descent stage. These 

engines perform with an average specific 

impulse of 379 seconds throughout the ascent 

maneuver. The ascent propulsion system will 

require approximately 39 metric tons of 

propellant to accomplish the approximately 

5,625 meters per second of velocity change 

required for a single-stage ascent to orbit and 

rendezvous with the previously deployed ERV. 

The structure and tanks needed for this 

propellant and the other attached hardware 

elements have a mass of 4.1 metric tons, 

including the mass of the engines but not the 

crew capsule. 

A3.3.6 Trans-Earth Injection 

Point of Contact: Larry Kos/MSFC 

Improvements were also made to the TransEarth Injection (TEI) stage for Reference 

Mission Version 3.0. The TEI stage uses two 

RL10-class engines modified to burn LOX/ 

CH4, similar to the descent stage. These 

engines perform with an average specific 

impulse of 379 seconds throughout the TEI 

maneuver. The TEI stage requires 

approximately 29 metric tons of propellant and 

has a dry mass of 5.9 metric tons. 

A3.3.7 Launch Packaging 

Point of Contact: Larry Kos/MSFC 

During 1997 the Exploration Transportation 

Team led my the Marshall Space Flight Center 

performed a packaging and launch configuration 

analysis of the Reference Mission Version 3.0 

payload elements. The focus of the packaging 

analysis was to determine the overall launch 

sequence and payload dimensions to ensure that 

the mission elements would fit within the 

overall payload dimensions and launch strategy 

of the Magnum Launch Vehicle. An overview 
EX13-98-036 

27

of the launch packaging analysis is provided in 

Figure A3-7. As can be seen in the figure, the 

overall launch sequence of the mission elements 

begins approximately 97 days prior to the 

opening of the Trans-Mars Injection window. 

This timeline is driven primarily by the launch 

processing of the payload elements and launch 

vehicle. For this analysis, 30 days were allotted 

for element processing between launches. A 

more thorough analysis of the ground 

processing is currently underway to determine a 

better estimate for the processing timeline. 

8.6 m

–67 days / TMI:

 mab = 10.2 mt

 mretHab = 29.1 mt

TEI Stage (30klbf

 total):

 (boil-off: 0.3%/mo ave.)

 mdry = 5.9 mt

 mp = 28.9 mt

 24 RCS thrusters

 mpyld = 74.1 mt

–37 days / TMI:

Ltank = 20 m (typ)

TMI Stage:

 (boil-off: 1.8%/mo LEO)

 mdry = 23.4 mt

 mp = 50.0 mt

 mstage = 73.4 mt

3 15 klbf

 NTP engines

12 RCS thrusters

28 m

(max)

28 m

(max)

–97 days / TMI:

 mab = 9.9 mt

 mecrv = 4.8 mt

Ascent Stage (60klbf

 total):

 mdry = 4.1 mt

 mp = 38.4 mt

Surface Payload:

 mcargo = 31.3 mt

 (incl. mLH2 = 5.4 mt)

Descent Stage (60klbf

 total):

 mdry = 4.9 mt

 mp = 11.0 mt

 24 RCS thrusters

 mpyld = 66.0 mt

–7 days / TMI:

TMI Stage:

 mdry = 23.4 mt

 mp = 45.3 mt

 mstage = 68.6 mt

3 15 klbf

 NTP engines

12 RCS thrusters

–67 days / TMI:

 mab = 13.6 mt

 mcrew = 0.5 mt

Surface Payload:

 mtransHab = 28.9 mt

 mmisc = 1.5 mt

Descent Stage (60klbf

 total):

 mdry = 4.9 mt

 mp = 11.4 mt

 24 RCS thrusters

 mpyld = 60.8 mt

–37 days / TMI:

TMI Stage:

 mdry = 26.6 mt

 mp = 50.0 mt

 mstage = 76.6 mt

3 15 klbf

 NTP engines

12 RCS thrusters

2011 TMI Stack1: 147.5 mt 2011 TMI Stack 2: 134.7 mt 2014 TMI Stack : 137.5 mt

Figure A3-7 Launch and Packaging Configurations for Version 3.0 
EX13-98-036

 28

A4.0 Summary of Reference Mission 

Version 3.0 

From the work of the original Reference 

Mission (Version 1.0), the strategy for the 

human exploration of Mars has evolved from its 

original form to one of reduced system mass, 

use of a smaller, more reasonable launch 

vehicle, and use of more current technology. 

The steps which have been taken by the 

Exploration Team are motivated by the need to 

reduce the mass of the payload delivery flights, 

as well as the overall mission cost, without 

introducing additional mission risk. By 

eliminating the need for a large heavy-lift 

launch vehicle and deleting the redundant 

habitat delivery flight in Version 3.0 , two 

launches from the Earth were eliminated. The 

net result is a current Version 3.0 Reference 

Mission which requires an injected mass of 

approximately one-half that of the 1993/94 

Reference Mission (Table A4-1). 

The modifications which have been made to 

the Reference Mission have resulted in 

significant reductions in total initial mass in 

low-Earth-orbit without significantly altering 

the overall mission architecture. A complete 

overview of the current Reference Mission 

Version 3.0 architecture is provided in Figure 

A4-1. 

A comparison of the mass breakdown for 

the various flights are provided in Table A4-2 

through Table A4-4. The masses of Reference 

Mission 1.0 and 3.0 are provided for 

comparison. 

Reference Mission Version 1.0 3.0

 First Opportunity:

 Cargo Lander (Cargo-1) 90,190 kg 66,043 kg

 Habitat (Hab-1) 90,598 kg N/A 

 Earth Return Vehicle-1 (ERV-1) 131,374 kg 74,072 kg

 Second Opportunity:

 Crew+Habitat 2 (Piloted-1) 89,980 kg 60,806 kg

TOTAL 402 mt 201 mt

Table A4-1 Payload Mass Evolution from Version 1.0 through Version 3.0 
EX13-98-036 

29

2011 - 2 Cargo 

Missions 

Launched

2014 - Crew 

transit habitat 

launched

Crew reaches Mars

in 130-180 days on

fast transit trajectory

Trans-Mars

injection and

Cruise

Earth Return

Vehicle

aerocaptures into

Mars orbit

Cargo lander with

propellant production

plant, power systems,

inflatable hab, ascent

vehicle lands on Mars

Crew

Arrival

Surface science concentrates on the search 

for life. Deep drilling, geology and 

microbiology investigations are supported

 by both EVA and by surface laboratories.

Crew departure.

Ascent

vehicle uses

locally produced

methane and

LOX.

Ascent Vehicle rendezvous

with Earth Return Vehicle 

in Mars Orbit. 130-180

day return trip to Earth

ends with direct entry and

precision parafoil landing.

Figure A4-1 Reference Mission Sequence for Version 3.0 
EX13-98-036

 30

Earth Return Vehicle Reference Version 1.0 Final Version 3.0

 Habitat Element 51974 kg 26581 kg

 Life Support System 6000 kg 4661 kg TransHab Study 7/97

 Crew Accomm. + Consumables 22500 kg 12058 kg TransHab Sizer from K. Kennedy

 Health Care 2500 kg 0 kg Included in Crew Acc. & Cons.

 EVA equipment 4000 kg 243 kg TransHab Sizer from K. Kennedy

 Comm/info management 1500 kg 320 kg TransHab Study 7/97

 30 kw PVA power system 3474 kg 3249 kg B. Cataldo 12/97 & TransHab Study 7/97

 Thermal Control system 2000 kg 550 kg TransHab Aerobrake Study 11/97

 Structure 10000 kg 5500 kg Scaled from Boeing 7.6 x 16.2 Mars Hab

 Science equipment 900 kg 600 kg 12\97 Scrub

 Spares 3500 kg 1924 kg Rodriggs/Munk 12/97

SUBTOTAL 56374 kg 29105 kg

 Excess consumables kg -7392 kg Jettison 560 days of food before TEI

 Crew & Samples from Surface 500 kg kg Crew and Payload from surface

ERV MASS AT TEI 63274 kg 27042 kg

 TEI stage drymass 5200 kg 4806 kg MSFC update 1/98

 Propellant mass 52000 kg 28866 kg MSFC update 1/98

 Earth return RCS propellant 0 kg 1115 kg MSFC update 1/98

 Aerobrake 17300 kg 10180 kg ARC update 1\98

TOTAL TEI MASS 120474 kg 61829 kg

TOTAL MOI MASS 131374 kg 74072 kg

 NTR Propulsion System 28900 kg 23400 kg

 Shadow Shield 0 kg 0 kg

 TMI Propellant 86000 kg 50000 kg

TOTAL INITIAL MASS 246274 kg 147472 kg

Table A4-2 Earth Return Vehicle Mass Scrub for Version 3.0
EX13-98-036 

31

Cargo Lander 1 Reference Version 1.0 Final Version 3.0

 Earth Entry/Mars Ascent Capsule 5500 kg 4829 kg JSC 12/97 Updates (X-38, ACRV, EVA)

 Ascent stage dry mass 2550 kg 4069 kg MSFC update 1/98

 ISRU plant 4802 kg 3941 kg T. Peters update 1/98; CH4, O2, H2O

 Hydrogen feedstock 4500 kg 5420 kg T. Peters 1/98

 PVA keep-alive power system 300 kg 825 kg B. Cataldo 11/97

 160 kw nuclear power plant 12498 kg 11425 kg B. Cataldo 12/97

 1.0 km power cables, PMAD 1900 kg 837 kg B. Cataldo 12/97

 Communication system 820 kg 320 kg TransHab 7/97

 Pressurized Rover 15500 kg 0 kg Delayed

 Inflatable Laboratory Module kg 3100 kg Derived from TransHab Study 7/97

 15 kwe DIPS cart 1100 kg 1500 kg DRM v 2.0

 Unpressurized rover 440 kg 550 kg PSS ESDB May 1991, p.486

 3 teleoperable science rovers 1320 kg 1500 kg DRM v 2.0

 Water storage tank 1220 kg 150 kg T. Peters 1/98

 Science equipment 3800 kg 1770 kg J. Gruener Update 12/97

TOTAL CARGO MASS 56250 kg 40236 kg

 Vehicle Structure 0 kg 3186 kg MSFC Update 1/98

 Terminal propulsion system 4670 kg 1018 kg MSFC Update 1/98

TOTAL LANDED MASS 60920 kg 44440 kg

 Propellant 11970 kg 10985 kg MSFC Update 1/98

 Forward Aeroshell 17300 kg 9918 kg ARC Update 1/98

 Parachutes and mechanisms kg 700 kg 4 parachues (to go with 4 engines)

TOTAL ENTRY MASS 90190 kg 66043 kg

 NTR Propulsion System 28900 kg 23400 kg

 Shadow Shield 0 kg 0 kg

 TMI Propellant 86000 kg 45300 kg

TOTAL INITIAL MASS 205090 kg 134743 kg

Table A4-3 Cargo Lander Mass Scrub for Version 3.0 
EX13-98-036

 32

Crew Lander Reference Version 1.0 Final Version 3.0

 Habitat element 2 53400 kg 28505 kg

 Life Support System 6000 kg 4661 kg TransHab Study 7/97

 Health Care 2500 kg 0 kg Included in Crew Accommodations

 Crew Accommodations 22500 kg 12058 kg TransHab Sizer from K. Kennedy

 EVA equipment 4000 kg 243 kg TransHab Sizer from K. Kennedy

 Comm/info management 1500 kg 320 kg TransHab Study 7/97

 Power 500 kg 3249 kg B. Cataldo 12/97 & TransHab Study 7/97

 Thermal 2000 kg 550 kg TransHab Aerobrake Study 11/97

 Structure 10000 kg 5500 kg Scaled from Boeing 7.6 x 16.2 Mars Hab

 Science 900 kg kg 12/97 scrub

 Spares 3500 kg 1924 kg Rodriggs/Munk 12/97

 Crew 500 kg 500 kg 6 - 183 lb people

 3 kw PVA keep-alive power 1700 kg 0 kg Included above

 Unpressurized rover 3 440 kg 550 kg DRM v 2.0

 EVA consumables kg 446 kg TransHab Sizer from K. Kennedy

 EVA suits kg 940 kg R. Yowell est. 12/97 - 156 kg/suit

TOTAL PAYLOAD MASS 56040 kg 30941 kg

 Vehicle structure kg 3186 kg MSFC Update 1/98

 Terminal propulsion system 4670 kg 1018 kg MSFC Update 1/98

TOTAL LANDED MASS 60710 kg 35145 kg

 Propellant 11970 kg 11381 kg MSFC Update 1/98

 Forward Aeroshell 17300 kg 13580 kg ARC Update 1/98

 Parachutes and mechanisms kg 700 kg 4 parachues (to go with 4 engines)

TOTAL ENTRY MASS 89980 kg 60806 kg

 NTR Propulsion System 28900 kg 23400 kg

 Shadow Shield 3300 kg 3200 kg

 TMI Propellant 86000 kg 50000 kg

TOTAL INITIAL MASS 208180 kg 137406 kg

Table A4-4 Piloted Lander Mass Scrub for Version 3.0 
EX13-98-036 

33

A5.0 Revolutionary Next Steps 

The Mars Reference Mission described in 

NASA Special Publication 6107, as modified by 

the updates described in the this addendum, 

provides a general framework for the human 

exploration of Mars. Since the original framing 

of the Reference Mission, other approaches 

have been brought forward as potential mission 

and technology options. These approaches, 

currently being analyzed by the Exploration 

Team, seem to be promising alternatives for 

accomplishing the primary objectives set forth 

in the original mission plan. The major mission 

alternatives currently under investigation 

include: 

• A Solar Electric Propulsion (SEP) option for 

performing the Earth departure phase of the 

mission 

• An approach for capturing the inflated 

TransHab into Mars orbit 

• Derivatives of the Nuclear Thermal Rocket 

concept which produces both propulsive 

thrust and continuous power 

• Techniques for minimizing launch mass 

perhaps to meet a three-Magnum launch 

scenario, and 

• All solar power scenarios. 

A5.1 Solar Electric Propulsion 

Points of Contact: Kurt Hack and Leon 

Geffert/LeRC, and Jeff George/JSC 

Many different approaches have been 

developed utilizing both solar electric and 

nuclear electric propulsion as a method of 

transporting both cargo and crew to and from 

Mars. These approaches focused on how an 

electric vehicle could be utilized to perform all 

of the major trajectory phases of the mission, 

including trans-Mars injection, Mars orbit 

capture, and trans-Earth injection. Although 

highly efficient from a propellant utilization 

standpoint, the relatively high power levels 

required to achieve fast-piloted trips generated 

two major challenges: 1) The vehicles were very 

large requiring significant on-orbit assembly 

and/or deployment, and 2) The technology 

requirements were significant (lightweight, 

multi-megawatt-class nuclear or solar 

powerplants; efficient and durable thrusters 

scaled to power levels on the order of 500 

kWe). These two significant challenges 

eliminated the electric propulsion vehicle as the 

primary propulsion concept for human Mars 

missions. 

During the Spring of 1997 an alternate 

concept of utilizing electric propulsion was 

proposed by the Lewis Research Center. After 

examining the payload delivery requirements of 
EX13-98-036

 34

the Reference Mission, it was determined that a 

compromise approach would be to utilize 

electric propulsion to perform the bulk of the 

trans-Mars injection, rather than all mission 

phases. This would minimize the disadvantages 

of previous approaches while still providing 

significant mission benefits. 

A5.1.1 Electric Propulsion Mission Concept 

The solar electric approach currently under 

investigation by the Exploration Team utilizes 

the high efficiency of electric propulsion where 

it provides the most benefit - boosting cargo out 

of the Earth’s gravity well. The overall mission 

strategy for the electric propulsion option is 

fundamentally the same as that of the Reference 

Mission: two cargo elements are launched in 

the first mission opportunity, followed by a 

piloted vehicle in the subsequent opportunity. 

The only major difference occurs in the 

replacement of the nuclear thermal TMI stage 

with a solar electric “tug” and small chemical 

kick stage. An overview of the mission concept 

is shown in Figure A5-1.

Electric Propulsion (EP) space

tug performs low-thrust transfer

for Mars-bound cargo to High

Earth Orbit (many months

transfer)

Crew delivered in “small”

chemically-propelled

transfer vehicle - X-38

derived (few days

rendezvous time)

HEO

Remainder of trans-Mars

injection performed by

chemically-propelled system

Space tug returns for

refueling and next

assignment (faster or more

efficient return since no

payload present)

LEO

Figure A5-1 Solar Electric Propulsion Mission Concept 
EX13-98-036 

35

Injection of cargo and piloted mission 

elements to Mars begins with the electric 

propulsion spiral phase. Due to the inherent 

high specific impulse at low thrust 

characteristics of electric propulsion, mission 

elements cannot be directly injected toward 

Mars via a traditional short impulsive burn. 

Orbital energy is instead continuously added 

over a period of approximately nine months, 

with the vehicle and payload following a spiral 

trajectory from an initial circular low Earth orbit 

(LEO) to a final elliptical high Earth orbit 

(HEO). A small chemical stage is then used to 

provide the final injection of the mission cargo 

toward Mars. The now-unloaded solar electric 

vehicle then returns to LEO to await a repeat 

sortie of the piloted vehicle element in the 

succeeding mission opportunity. 

Delivery of the crew to Mars requires a 

slight modification to the front-end of the 

mission. As with the cargo missions, the 

electric propulsion vehicle is used to boost the 

piloted vehicle, sans crew, into a high Earth 

orbit. The crew is not transported in the vehicle 

during this phase for two primary reasons. 

First, during the spiral boost phase of the 

mission, the vehicle traverses the harsh Van 

Allen radiation belts many times - far too 

excessive for piloted missions. Second, the 

spiral phase takes several months to perform, 

significantly increasing the exposure of the crew 

to the debilitating effects of zero-gravity. 

Rather than employing countermeasures, these 

effects are minimized by delivering the crew in 

a high speed taxi to the piloted vehicle after it 

has been boosted to the final high Earth 

departure orbit. After a short rendezvous and 

checkout period, the piloted vehicle, like the 

previous cargo vehicles, is injected to Mars with 

a small chemical stage. 

A5.1.2 Electric Vehicle Concepts 

Vehicle concepts for the electric propulsion 

option are currently under investigation by the 

Exploration Team. During the selection and 

analysis process, emphasis is being placed on 

developing a concept which can be deployed 

easily, do not require significant advancements 

in technology, and is a low cost approach. 

Conceptual vehicle designs for the crew taxi and 

solar electric vehicle are shown in Figures A5-2 

and A5-3. The concepts shown are still under 

investigation and will continue to evolve as 

advancements in the analysis are made. A 

summary of the mission mass estimates for the 

solar electric vehicle concept are provided in 

Figure A5-4. 
EX13-98-036

 36

Mass (kg) X-38 Derivative Capsule

Crew module mass 11,340 6,500

Stage propellant mass 11,721 6,890

Stage inert mass 2,418 1,588

Gross mass 25,479 14,979

Figure A5-2 SEP Crew Taxi Concepts. 

A5.2 Aerocapturing the TransHab 

Point of Contact: Bill Schneider/JSC 

The other major mission option currently 

under investigation is the approach of 

aerocapturing the inflated TransHab into Mars 

orbit. During the Fall of 1997 a “Skunk Works” 

study team composed of experts from the 

Johnson Space Center, Langley Research 

Center, Ames Research Center, and Marshall 

Space Flight Center, conducted a study of the 

TransHab aerocapture concept. The goals of the 

study were to design a lightweight aeroshell 

system capable of capturing the inflated 

TransHab into Mars orbit and to determine the 

best system for crew return to Earth. Aeroentry 

and landing were not considered during this 

study and were left for follow-on analysis. 

Two aeroshell concepts were analyzed during 

the study: The Ellipsled, which uses the 

structure from the Magnum launch vehicle 

shroud (requiring no on-orbit assembly); and the 

Spherical Dome, which is Shuttle-launched and 

assembled (see Figure A5-5). In order to 

estimate the total system mass, the analysis 

included investigations of the entry flight 

dynamics, thermal protection system, structural 

design, and assembly operations. 

Analysis conducted showed that both the 

Ellipsled and the Spherical Dome could 

accomplish an aerocapture at Mars with positive 

margins. However, a number of factors has led 

to a selection of the slender shape concept. 

Results from the study indicated that the mass 

fraction (ratio of the mass of the aerobrake to 

the mass of the aerobrake and payload) ranged 

from 14.6% to 15.5%. These results closely 

matched those from the previous aeroshell 

analyses (see section A3.4), indicating 

reliability in the mass estimates. Aerobrake 

mass fractions of this magnitude provide 

significant mission advantages by reducing the 

total mass required for the mission. 

Analyses of the TransHab Aerobrake 

concepts are still in progress. Several factors 
EX13-98-036 

37

Figure A5-3 Conceptual Solar Electric Propulsion Vehicle 

2011

HLV

2014

HLV

2014

HLV

2014

STS

/LFBB

SPM 16.0MT

EPM 60.0MT

Tot. 75.9MT

MTV 80.3MT

Tot. 80.3MT

TMI 29.2MT

EPM 47.1MT

Tot. 76.3MT

Taxi 26.8MT

Tot. 26.8MT

2011

HLV

MTV 52.6MT

TMI 20.5MT

Tot. 73.1MT

IMLEO:

382.9 MT (HLV)

26.8 MT (STS)

IMLEO:

382.9 MT (HLV)

26.8 MT (STS)

2011

HLV

MTV 55.8MT

TMI 21.4MT

Tot. 77.3MT

Total Launches:

5 HLV

1 STS

Total Launches:

5 HLV

1 STS

MTV - (Injected) Mars Transfer Vehicle

TMI - Trans Mars Injection Stage/Propellant

SPM - Solar Power/Bus Module

EPM - Electric Propulsion Module (wet)

Taxi - LEO to HEO Crew Taxi

2011: One SEP stage delivers

two Cargo MTV/TMI’s to

39,709 x 800 km @ 51.6 deg.

2014: Same SEP Stage is refueled

and delivers piloted MTV/TMI to

70,761 x 800 km @ 51.6 deg.

Figure A5-4 Launch manifest for the Solar Electric Propulsion Vehicle Concept.
EX13-98-036

 38

remain to be investigated to complete the study, 

including: 

• Modifications to the TransHab. The initial 

effort focused on aerocapturing the habitat 

originally designed by the JSC team in the 

Spring of ‘97. Modifications to the 

TransHab, including structural 

modifications to operate on the surface of 

Mars, and the addition of a crew flight deck, 

were not addressed. 

• Entry and Landing. The initial study 

focused only on the aerocapture phase of the 

mission, and did not address the issues 

associated with the entry and landing 

phases, such as static and low-speed 

dynamic stability, parachute deployment, 

terminal engine requirements, or landing 

accuracy. If it is not feasible to land the 

inflatable TransHab in its current 

configuration, modifications and additional 

vehicle elements may have to be introduced 

into the architecture. 

• Assessment of the impacts of decreasing the 

Earth-Mars transit times from 200-days to 

180-days to be consistent with previous 

analysis performed by the Ames Research 

Center. 

TransHab-to-Aeroshell Attachment. Detailed 

structural design of the attachments between the 

TransHab and the aeroshell will serve to further 

refine the mass estimates of the system, and 

must address packaging, deployment, and 

accessibility issues. 

Figure A5-5 Potential TransHab Aerobrake Configurations. 
EX13-98-036 

39

A5.3 Bimodal Nuclear Thermal Rocket 

(NTR) Propulsion 

Point of Contact: Stan Borowski/LeRC 

Although most of the current work is 

focused on the Solar Electric concept, the NTR 

approach is being maintained for comparison. 

The solid core NTR propulsion system 

represents a “rich source of energy” in that it 

contains substantially more uranium-235 fuel in 

its reactor core than it consumes during its 

primary propulsion maneuvers. By configuring 

the NTR engine as a “bimodal” system, 

abundant electrical power can also be generated 

for a variety of spacecraft needs. During power 

generation, the reactor core operates in 

essentially an “idle mode” with a thermal power 

output of ~100 kilowatts. The reactor thermal 

energy is subsequently removed and routed to a 

turbo-alternator-compressor Brayton power 

conversion unit using a helium-xenon working 

fluid, as shown in Figure A5-6. A space 

radiator system rejects waste heat and also 

reduces decay heat propellant loss following 

propulsive burns. 

Thermal Propulsion

He/Xe

Brayton Power Conversion

is Primary Option. Other

Options Can Also Be

Adapted

Compressor

Generator

Cryogenic H2 

Propellant Tank

H2 Turbopump

Radiator Heat Turbine

Exchanger

Power Conditioning

Payloads

On-Board 

Systems

Attitude Control

Ion Thrusters

Figure A5-6 Schematic “Bimodal” NTR 

System 

5.3.1 Bimodal NTR Mission Concept 

An option to Reference Mission Version 3.0 

that utilizes bimodal NTR transfer vehicles in 

place of the expendable NTR stages is being 

evaluated. A common “core” stage, used on 

cargo and piloted vehicles alike, is outfitted 

with three 15 klbf bimodal NTR engines 

capable of providing up to 50 kilowatts of 

electrical power (kWe) using any two engines 

The bimodal core stage is not jettisoned after 

the TMI maneuver but remains with the cargo 

and piloted payload elements providing 

midcourse correction (MCC) propulsion and all 

necessary power during transit. Near Mars, the 

bimodal stage separates from the aerobraked 

payloads and performs its final disposal 

maneuvers. A key difference between 

Reference Mission 3.0 and the bimodal option is 
EX13-98-036

 40

the absence of the aerobraked LOX/methane 

(CH4) TEI stage which is replaced by an “all 

propulsive” bimodal NTR-powered Earth 

Return Vehicle (ERV) illustrated in Figure A5-

7. 

The bimodal stage LH2 tank is slightly 

shorter than the expendable TMI stage tank at 

19 meters and has a maximum LH2 propellant 

capacity of ~51 tons with a 3% ullage factor. A 

turbo-Brayton refrigeration system is located in 

the forward cylindrical adaptor section to 

eliminate LH2 boiloff during the lengthy (~4.3 

year) ERV mission. A 12 kWe Brayton 

refrigeration system is included to remove the 

~100 watts of heat flux penetrating the 2 inch 

MLI system in low-Earth-orbit where the 

highest heat flux occurs. Enclosed within the 

conical aft radiator section of the bimodal core 

stage is a closed Brayton cycle (CBC) power 

conversion system employing three 25 kWe 

Brayton rotating units (one for each bimodal 

reactor) which operate at ~2/3 of rated capacity, 

thus providing an “engine out” capability. The 

turbine inlet temperature of the He-Xe working 

gas is ~1300 K and the total system specific 

mass is estimated to be ~30 kg/kWe. 

A mass comparison of the bimodal NTR 

transfer 

Figure A5-7 “Bimodal” NTR Transfer Vehicle Option
EX13-98-036 

41

Magnum 

Launch Flight Element 2011 Cargo Lander 2011 ERV * 2014 Crew Lander Totals

Mission Type DRM Bimodal DRM Bimodal DRM Bimodal DRM Bimodal

Payload 66.0 65.0 74.1 25.5 60.8 56.4 200.9 146.9

 - Surface/"In-Space" - 40.2 - 40.2 - 29.1 - 25.5 - 30.9 - 28.4 - 100.2 - 94.1

#1 - Transportation - 25.8 - 24.8 - 45.0 - 29.9 - 28.0 - 100.7 - 52.8

"In - Line" 

Propellant/Tankage - - - 20.1 - 5.3 - 25.4

(LH2 &/or LOX)

NTR TMI stage

#2 ("Modified" DRM 68.6 73.6 73.4 79.0 76.6 79.0 218.7 231.6

uses "bimodal" NTRs)

Total : 134.6 138.6 147.5 124.6 137.5 140.7 419.6 403.9

# Magnums 22222266

* 2011 ERV mission using "bimodal" NTRs for MOC and TEI is lighter than DRM by ~23 t and eliminates DDT&E and recurring

costs for LOX/CH4 TEI stage, also recurring cost for 30 kWe PVA and aerobrake.

** Common "Bimodal" NTR TMI stage provides 50 kWe power capability to the ERV, Crew and Cargo lander missions. Also

supplies MCC burns for these missions. For cargo lander, the "Bimodal" stage refrigeration/heat rejection systems can be used

to cryocool 4.5 t of "seed" LH2 and dump "waste heat" from 15 kWe DIPS power cart.

Table A5-1 Comparison of “Bimodal” NTR to Reference Mission Version 3.0 

vehicles and the Reference Mission Version 3.0 

vehicles is shown in Table A5-1. 

The mass values assume a “2-perigee burn” 

Earth departure scenario. Overall, the bimodal 

approach has a lower “three-mission” initial 

mass than Reference Mission 3.0. In addition, 

the bimodal approach can reduce the operational 

complexity of the mission (eliminates solar 

array deployment/retraction) as well as 

eliminating the need for an aerobrake and 

injection stage for the Earth Return Vehicle. 

5.3.2 All Propulsive” Bimodal NTR Option 

Using TransHab 

Another option to the Reference Mission 3.0 

under consideration is the use of a bimodal NTR 

stage to propulsively capture all payload 

elements into Mars orbit. This “all propulsive” 

NTR option provides the most efficient use of 

the bimodal engines which can supply abundant 

power to the spacecraft and payloads in Mars 

orbit for long periods. Propulsive capture into 

the reference “250 km by 1 sol” elliptical Mars 

parking orbit also makes possible the use of a 
EX13-98-036

 42

standardized, reduced mass “aerodescent” shell 

because of the lower payload entry velocity 

(~4.5 km/s) encountered. From this orbit, the 

triconic aerobrake mass varies by only ~400 kg 

for a 20 ton increase in payload mass (see

Section 3.3.3). 

The attractiveness of the “all propulsive” 

bimodal NTR option is further increased by the 

utilization of the lightweight, inflatable 

“TransHab” module discussed in Section 3.1. 

The substitution of TransHab for the heavier, 

hard-shell habitat module introduces the 

potential for propulsive recovery of the Earth 

Return Vehicle in Earth orbit and its reuse on 

subsequent missions. TransHab use also allows 

the crew to travel to and from Mars on the same 

bimodal transfer. In Mars orbit, the crew 

transfer vehicle rendezvous with the “unpiloted” 

habitat lander which is now delivered as a cargo 

element by the bimodal stage. The absence of 

crew from the bimodal habitat lander eliminates 

the need for outbound consumables and engine 

crew radiation shields and allows it to carry offloaded surface habitation and science equipment 

previously carried on the cargo lander. 

A three-dimensional image of the bimodal 

transfer vehicle used on the piloted mission is 

shown in Figure A5-8. The TransHab is ~9.7 

meters long and inflates to a diameter of ~9.5 

meters. Its total mass is ~24.3 metric tons 

which includes the crew and their consumables. 

The total length and initial mass of the piloted 

transfer vehicle is ~54 meters and ~141 metric 

tons, respectively. A smaller, “in-line” 

propellant tank is used on the bimodal transfer 

vehicles that deliver the ~46 metric ton habitat 

and ~54 ton cargo landers into Mars orbit. The 

habitat and cargo transfer vehicles are ~56 

meters long and have a LEO mass of ~129 

metric tons and 144 metric tons, respectively. 

5.3.3 “LOX-Augmented” NTR Option 

An enhanced NTR option, known as the 

“LOX-augmented” NTR (LANTR), is presently 

under study by NASA which combines 

conventional LH2-cooled NTR and supersonic 

combustion ramjet (scramjet) technologies. The 

LANTR concept utilizes the large divergent 

section of the NTR nozzle as an “afterburner” 

into which LOX is injected and supersonically 

combusted with reactor preheated hydrogen 

emerging from LANTR’s choked sonic throat--

essentially “scramjet propulsion in reverse.” By 

varying the oxygen-to-hydrogen mixture ratio 

(MR), the LANTR engine could potentially 

operate over a wide range of thrust and specific 

impulse values while the reactor core power 

level remains relatively constant. For those 

missions where volume (not mass) constraints 

limit bimodal stage performance, LANTR 

propulsion can help to increase “bulk” 
EX13-98-036 

43

propellant density and total thrust output, while 

decreasing the engine burn times. LOX 

augmentation would be particularly beneficial 

during the TMI burn to reduce gravity losses. 

Following this maneuver, the spent “in-line” 

LH2 tank and a small LOX tank attached to 

it could be jettisoned as a single unit. On all 

subsequent burns, the LANTR engines would 

operate on only LH2 (MR = 0). Cold flow 

experimental injector tests and reactive 

computational fluid dynamics analyses are 

currently underway at NASA Lewis Research 

Center in preparation for future hot flow tests 

aimed at demonstrating concept feasibility. 

Figure A5-8 All Propulsive Bimodal NTR Carrying TransHab 

5.3.4 Three-Magnum Scenario 

Point of Contact: Andrew Petro/JSC 

During the Spring of 1998 a special design 

study was conducted to define the elements, 

mission content, and technology required to 

accomplish a human Mars mission which could 

be accommodated for launch within the mass 

and volume capacity of three heavy-lift launch 

vehicles. The reference launch vehicle used in 

the study was the Magnum launch vehicle and 
EX13-98-036

 44

so this mission concept is referred to as the 

“Three-Magnum Mars Mission”. The design 

team was directed to employ a solar electric 

propulsion (SEP) stage for delivering the Mars 

mission elements to a high apogee Earth 

departure orbit and to not employ nuclear 

propulsion for any maneuvers. 

This study was unusual in the approach of 

designing to a fixed constraint for Earth launch 

mass. The most significant result of the study 

was the identification of the technology 

challenges which must be met to achieve the 

launch mass goal. 

The capacity of the Magnum launch vehicle 

defined for this study was 89.5 metric tons for 

launch packages which employ the launch 

shroud as an aeroshell, and 85.5 metric tons for 

payloads which do not include the shroud as 

payload. The payload capability quoted is for 

launch from the Kennedy Space Center to a 

circular orbit of 400 kilometers at an inclination 

of 28.5 degrees. The dimensions of the 

Magnum shroud were defined as an outer 

diameter of 8.4 meters and a length of 28 

meters. 

5.3.4.1 Mission Content 

The mission defined in this study included a 

crew of four people, a scientific payload of 1770 

kg and two unpressurized rovers with a mass of 

650 kg each. The missions were conjunctionclass with outbound and inbound transit 

durations of 180 to 200 days and Mars surface 

stay times of 520 to 580 days. The elements 

were designed to accomplish missions in six out 

of the eight opportunities in the synodic cycle. 

The other two opportunities would require an 

additional propulsive stage of approximately 16 

metric tons. 

Several different mission scenarios were 

considered and two were documented for the 

study: a Combination Lander Scenario in which 

all elements are sent to Mars in a single 

opportunity, and a Split Mission Scenario in 

which some elements are deployed at Mars in 

the first opportunity and the crew travels to 

Mars in the next opportunity. The Split Mission 

Scenario is similar to the Design Reference 

Mission 3.0 whereby propellant for Mars ascent 

is produced at Mars. 

5.3.4.2 Strategies and Technology Challenges 

Several strategies were used to constrain the 

total mission mass with respect to the Design 

Reference Mission and to achieve the launch 

mass target. 

• Crew reduced from 6 to 4 persons 

• Initial departure orbit apogee raised from 

39,000 km to 120,000 km 

• Hydrogen fuel is used for all maneuvers. 
EX13-98-036 

45

In addition, several technology development 

challenges were identified as necessary to 

achieve the launch mass target. 

• Structures, tanks, and aeroshells with a 

reduction in mass of up to 50% over current 

technology 

• High performance power generation system 

for space and surface operations (100 

kg/kWe) 

• Long-term hydrogen storage with near zero 

boil-off for up to four years 

• Lightweight chemical propulsion engines 

with a specific impulse of 480 sec. 

• Deployable solar electric propulsion system 

with a megawatt-capacity solar array 

5.3.4.3 Combination Lander Scenario 

This scenario is illustrated in Figure A5-9 

and the launch packages with element masses 

are shown in Figure A5-10. Figure A5-11 is a 

three-dimensional drawing of the Combination 

lander concept as it would be deployed on the 

surface. This lander includes the crew module 

for descent and ascent along with the surface 

habitat. 

5.3.4.4 Split Mission Scenario 

The Split Mission Scenario is similar to the 

Design Reference Mission scenario but it 

includes all of the strategies and technology 

challenges mentioned above. The major 

differences in this scenario are 1) the predeployment of the return vehicle in Mars orbit, 

2) pre-deployment of the ascent vehicle on the 

surface of Mars, 3) the production of propellant 

on Mars, and 4) the use of methane rather than 

hydrogen for Mars ascent. The scenario is 

illustrated in Figure A5-12 and the launch 

packages and element masses are shown in 

Figure A5-13. 

5.3.4.5 Summary 

By incorporating the aggressive technology 

goals, two mission scenarios were defined 

which could meet the three-Magnum launch 

mass and volume constraint. It should be noted 

that each scenario also requires a Space Shuttle 

launch at the beginning of the mission to deliver 

the crew and their high-Earth orbit taxi and also 

a Shuttle mission at the end to recover the crew 

in low Earth orbit. This three-launch strategy is 

reliant on the key technologies described 

previously. An effort is currently underway to 

better understand the difficulty of the 

technology challenges as they compare to 

current state-of-the-art, the risks associated with 

these technologies, development costs, and the 

architectural impacts of potential technology 

fall-backs if it is believed that the technology 

development cannot be completed as needed. 
EX13-98-036

 46

Outbound

Return

1 Earth Launch

(MI) SEP

(M2) Combo Lander

(M3) Transit Hab/TEI

5 TMI

Transit Hab/TEI with Crew

Combo Lander

6 Aerocapture at Mars

Combo Lander

6 Aerocapture at Mars

Transit Hab/TEI with Crew

7 Mars Orbit Rendezvous

Transit Hab/TEI with Crew

& Combo Lander

8 Descent & Landing

Combo Lander with Crew

3 Earth Launch

Taxi with Crew

9 Surface Mission

4 HEO Rendezvous

Crew in Taxi with 

Combo Lander

Transit Hab/TEI

2 SEP Spiral

SEP

Combo Lander

Transit Hab/TEI

10 Mars Ascent

Ascent Vehicle with Crew

separates from Combo Lander

11 Mars Orbit Rendezvous

Ascent Vehicle with Crew

& Transit Hab/TEI 12 TEI

Transit Hab/TEI with Crew

13 Aerocapture at Earth

Transit Hab with Crew

14 Earth Orbit Rendezvous

Transit Hab with Crew & Shuttle

Earth

Mars

Earth

Figure A5-9 Three-Magnum Combination Lander Scenario 

Transit Habitat 27,403

Aeroshell 8,750

TEI 40,017

TMI 12,572

Total 88,742

SEP 85,500

Total 85,500

Ascent Crew Module 2,704

Ascent Stage 10,681

Surface Habitat 20,293

Surface Payload 3,070

Surface/Transit Power 3,025

Lander Stage 13,166

Aeroshell 6,901

TMI 29,418

Total 89,258

28 m

8.4 m

ALL MASSES IN KG

SEP Stage

Figure A5-10 Three-Magnum Combination Lander Launch Packages 
EX13-98-036 

47

Inflated Surface Habitat

with Body-Mounted Radiator

Empty Descent Tanks

Surface Power System

Power System Radiator

2 Unpressurized Rovers

Ascent Crew Module

Figure A5-11 Combination Lander Concept on Mars Surface 

Outbound - 1st Opportunity

Return

1 Earth Launch

(MI) SEP

(MII) Ascent Vehicle

 & Return Hab/TEI 3 TMI

Ascent Vehicle & Return Hab/TEI 4 Aerocapture at Mars

Ascent Vehicle

9 HEO Rendezvous

Crew in Taxi with 

Transit/Surface Hab

2 SEP Spiral

SEP

Ascent Vehicle 

& Return Hab/TEI

14 Mars Ascent

Ascent Vehicle with Crew

15 Mars Orbit Rendezvous

Ascent Vehicle with Crew

& Return Hab/TEI

16 TEI

Return Hab/TEI with Crew 17 Aerocapture at Earth

Return Hab with Crew

18 Earth Orbit Rendezvous

Return Hab with Crew & Shuttle

Earth

Mars

4 Aerocapture at Mars

Return Hab/TEI

Remains in orbit

5 Descent & Landing

Ascent Vehicle

Ascent propellant produced

Outbound - 2nd Opportunity

Earth

6 Earth Launch

(MIII) Transit/Surface Hab 10 TMI

Transit/Surface Hab with Crew

11 Aerocapture at Mars

Transit/Surface Hab with Crew

8 Earth Launch

Taxi with Crew

7 SEP Spiral

SEP

Transit/Surface Hab 12 Descent & Landing

Transit/Surface Hab with Crew

13 Surface Mission

Earth

Figure A5-12 Three-Magnum Split Mission Scenario 
EX13-98-036

 48

Surface/Transit Power 3,025

Hab Lander 7,958

Rover 650

Aeroshell 5,124

TMI 12,874

SEP 2nd Load 23,282

Total 75,150

SEP 1st Flight 53,220

TMI 19,711

Total 72,931

Ascent Vehicle 9,105

Ascent Lander 5,044

ISRU Plant 4,868

Ascent Aeroshell 3,203

Return Habitat 25,458

Surface Payload 2,420

Hab Aeroshell 6,184

TEI Stage 15,927

Total 72,209

JSC Sizing with LeRC Model 

ALL MASSES IN KG

TMI

SEP Stage

Return Hab

TEI

Ascent 

Vehicle

Surface Hab

TMI

SEP Propulsion Unit 

for Second Flight

• First Opportunity • Second Opportunity

Figure A5-13 Three-Magnum Split Mission Launch Packages 

5.3.5 All Solar Scenario 

Another alternative strategy under 

consideration is an approach where total 

reliance would be place on propulsion and 

power concepts based solely on chemical and 

solar technologies. Of particular importance is 

the power generation strategy which has relied 

on the same technology base (SP-100) from the 

original reference through Reference Mission 

3.0. This power strategy has been one of 

providing a robust power generation and storage 

capability to enable significant mass reductions. 

This technique of trading mass for power has 

been manifested in the Reference Mission in the 

form of advanced technologies such as in-situ 

resource utilization, bioregenerative closedloop life support systems, and long-range 

pressurized rovers. These high power demands 

necessitated the use of advanced power 

concepts such as surface nuclear reactors and 

dynamic isotope power sources. 

A major challenge of an all-solar human 

Mars mission is the lack of solar irradiation at 

Mars. As can be seen in Figure 5-14 the solar 

flux at the surface can be as low as 6.5% of that 
EX13-98-036 

49

in low-Earth orbit. The reduction of solar flux 

is due to the distance of Mars from the sun, the 

presence of the atmosphere, and potential dust 

storms. 

0

200

400

600

800

1000

1200

1400

1600

Earth

Orbit

Mars

Orbit

Clear Cloudy

(Local

Storm)

Cloudy

(Global

Storm)

Solar Irradiance (W/sqm)

.

Mars Surface

100%

43%

22%

13% 6.5%

Figure 5-14 Solar Irradiation At Mars. 

Analysis of an all-solar approach will 

include: 

• Developing a mission approach where the 

surface element power needs can be reduced 

to the lowest level possible 

• Understanding the sensitivities of advanced 

solar cell technologies 

• Analysis of solar power generation system 

setup and maintenance, such as cleaning due 

to dust accumulation 

• Analysis of the impacts of elimination of 

advanced technologies (in-situ resource 

utilization, long-range rovers, food 

production, etc.) on the overall mission 

approach, including risk. 

Analysis of this all-solar mission approach 

is currently under way. Results of this study 

will be included in the next update of the 

Reference Mission. 
EX13-98-036

 50

A6.0 Continuing Work 

A “reference mission” is a continual work 

in progress, provided to the space community to 

critique and build upon. Future human 

exploration analysis activities will focus on 

exploration targets such as the Moon, Asteroids, 

or other destinations beyond Earth orbit. 

Further addenda will be published which 

document these changes to the mission, changes 

which will undoubtedly be made before it 

becomes a reality. Through this process, the 

Exploration Team can nurture the design of 

human exploration missions which offer a safe 

and economical strategy for taking the next step 

in humanity’s exploration and development of 

space.