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IS20

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Marshall Space Flight Center, Alabama

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i

NASA/TM—2010–216450

Ares V Utilization in Support of a Human 

Mission to Mars

J.B. Holladay, J.P. Jaap, R.M. Pinson, S.D. Creech, and R.M. Ryan 

Marshall Space Flight Center, Marshall Space Flight Center, Alabama

T.S. Monk and K.E. Baggett

Zero Point Frontiers Corp., Huntsville, Alabama

M.D. Runager 

Jacobs ESTS Group/Triumph Aerospace, Huntsville, Alabama

I.J. Dux and K.J. Hack

Glenn Research Center, Cleveland, Ohio

J.M. Hopkins and C.E. Brown

Kennedy Space Center, Kennedy Space Center, Florida

T.A. Manning

Ames Research Center, Moffett Field, California

November 2010

National Aeronautics and

Space Administration

Marshall Space Flight Center • MSFC, Alabama 35812
ii

Available from:

NASA Center for AeroSpace Information

7115 Standard Drive

Hanover, MD 21076 –1320

443–757–5802

This report is also available in electronic form at

<https://www2.sti.nasa.gov>

Acknowledgments

The work documented in this Technical Memorandum (TM) was performed during the summer of 2010 

as a joint effort between Ares V team members at Marshall Space Flight Center, Glenn Research Center, Kennedy 

Space Center, Ames Research Center, Zero Point Frontiers Corporation, and Jacobs ESTS Group/Triumph Aerospace. 

The authors would like to particularly acknowledge Constellation Program support from the Johnson Space Center, 

persons who worked on the Mars DRA 5.0 team, and those involved with generating the further detailed “7-Launch 

NTR Space Transportation System for NASA’s Mars DRA 5.0” AIAA paper, as these studies provided a baseline 

from which this TM was created.
iii

TABLE OF CONTENTS

1. INTRODUCTION ............................................................................................................. 1

2. MARS ARCHITECTURE OVERVIEW: DESIGN REFERENCE 

ARCHITECTURE 5.0 ....................................................................................................... 2

3. ARES V OVERVIEW: LAUNCH VEHICLE OPTIONS ................................................... 5

3.1 Ares V Concept History ............................................................................................... 5

3.2 Concepts Assessed for Mars Design Reference Architecture 5.0 ................................... 6

3.3 Concept Assessed During Phase A-Cycle 3′ ................................................................. 6

3.4 Key Mass Assumptions for Phase A-Cycle 3′ Mars Design 

Reference Mission Assessment ..................................................................................... 8

3.5 Performance of Phase A-Cycle 3D for Low-Earth Orbit Applications ......................... 9

3.6 Payload Performance to Final Circular Orbit ............................................................... 12

4. MARS MISSION ANALYSIS: MARS TRANSFER VEHICLE OPTIONS ..................... 19

4.1 Nuclear Thermal Propulsion Option ............................................................................ 19

4.2 Chemical Propulsion Options ....................................................................................... 24

5. MARS MISSION ANALYSIS: KEY CHALLENGES....................................................... 37

5.1 Required Launch Vehicle Build, Storage, and Transportation Rate .............................. 37

5.2 Required Launch Vehicle Launch Rate ........................................................................ 37

5.3 Payload Mass Required ................................................................................................ 42

5.4 Mission Reliability ....................................................................................................... 44

5.5 On-Orbit Assembly Operations/Functionality Required .............................................. 50

5.6 Earth Departure Stage as Mars Transfer Vehicle Stage ................................................ 51

5.7 Technology Developments Required ............................................................................ 54

6. LUNAR AND MARS MISSION SYNERGISTIC POTENTIAL ..................................... 63

6.1 Infrastructure Sharing .................................................................................................. 63

6.2 Earth Departure Stage as Trans-Mars Insertion Stage ................................................. 65

6.3 Opportunity to Engage Commercial and International Partners .................................. 67

6.4 Enabling Alternate Missions ........................................................................................ 69

6.5 Public Perception ......................................................................................................... 69
iv

TABLE OF CONTENTS (Continued)

7. FORWARD WORK ........................................................................................................... 70

7.1 Ares V Performance—Trajectory ................................................................................. 70

7.2 Earth Departure Stage—Subsystems ............................................................................ 70

7.3 Shroud ......................................................................................................................... 71

7.4 Mars Campaign Ground Operations ............................................................................ 71

7.5 Safety and Mission Assurance ...................................................................................... 72

7.6 Mars and Near-Earth Orbit Synergistic Potential ........................................................ 72

REFERENCES ....................................................................................................................... 75
v

LIST OF FIGURES

 1. Mars DRA 5.0 top-level trade tree .......................................................................... 2

 2. DRA 5.0 long-stay Mars mission overview (seven-launch NTP strategy) ................ 3

 3. Timeline of major Ares V changes/reviews since ESAS ........................................... 5

 4. Potential shroud options assessed for DRA 5.0 ....................................................... 7

 5. PA-C3D Ares V concept (baseball card) .................................................................. 8

 6. Conceptual LEO mission profile .............................................................................. 13

 7. Representative LEO payload using Shuttle OMS assumptions. 

A 1.2-t knockdown assumption is included in the data ............................................ 17

 8. LEO performance growth options ........................................................................... 18

 9. TMI stage propulsion options assessed during PA-C3′ ............................................ 19

10. TMI propulsion options discussed ........................................................................... 20

11. NTP launch sequence and payload mass requirements ............................................ 20

12. Ares V payload to LEO delivery sequence (minimum participation, NTP cargo) .... 22

13. Ares V payload to LEO delivery sequence (minimum participation, NTP crew) ...... 22

14. Ares V payload to LEO delivery sequence (maximum participation, NTP cargo) ... 24

15. Ares V payload to LEO launch No. 1 (maximum participation, NTP crew) ............ 25

16. Ares V payload to LEO launch No. 2 (maximum participation, NTP crew) ............ 25

17. Ares V payload to LEO launch No. 3 (maximum participation, NTP crew) ............ 26

18. DRA 5.0 chemical propulsion module option ......................................................... 27

19. Propellant delivery and transfer option (EDS-as-MTV) .......................................... 29

20. Dual-use shroud/encapsulated payload launch (chemical cargo) ............................. 30
vi

LIST OF FIGURES (Continued)

21. Propellant launch for cargo MTV (first chemical cargo) ........................................... 31

22. Shared propellant launch for first MTV (second chemical cargo) ............................. 31

23. Shared propellant launch for second MTV (two and one-half chemical cargo) ........ 32

24. MOI/TEI module launch, first launch (first chemical crewed) .................................. 33

25. TransHab/truss/consumables/“second TMI stage,” second launch 

(second chemical crewed) .......................................................................................... 33

26. Propellant launch for ‘second TMI stage’ (third chemical crewed) ............................ 34

27. ‘First TMI Stage’ launch (fourth chemical crewed) ................................................... 34

28. Commercially delivered propellant option (EDS tanker derivative) .......................... 36

29. General decision tree for Ares V CS extensibility to Mars DRM .............................. 38

30. MAF CS processing timeline .................................................................................... 38

31. Proposed Ares V flight rate assumptions (NTP or chemical) .................................... 39

32. Proposed Ares V Mars Ops Con (chemical or nuclear) ............................................. 40

33. Preliminary proposed GOP Mars architecture (upgrades 

to the lunar architecture) .......................................................................................... 41

34. Mass and volume requirements for multiple missions ............................................... 43

35. Trade tree for EDS as MTV stage ............................................................................. 53

36. Cargo MTV propellant inventory at TMI ................................................................. 57

37. Crewed MTV propellant inventory at TMI ............................................................... 58

38. Mass loss of various boiloff rates (200-t initial, 180-day assembly) .......................... 58

39. Mass loss of various boiloff rates for NTP crewed MTV .......................................... 59

40. Mass loss of various boiloff rates for chemical cargo MTVs .................................... 59
vii

LIST OF FIGURES (Continued)

41. Benefit of adding sixth launch to cargo MTV assembly sequence ............................ 60

42. Benefit of adding fifth launch to crewed MTV assembly sequence ........................... 61

43. Benefit of adding fifth launch to crewed MTV assembly sequence ........................... 62

44. One-and-a-half launch lunar architecture from Ares V Ops Con ............................ 63

45. Chemical option (i.e., EDS-as-MTV, or unofficially scavenger option) .................... 64

46. EDS concept for Mars architecture ......................................................................... 65

47. EDS concept for lunar architecture ......................................................................... 65
viii
ix

LIST OF TABLES

 1. Performance summary for the elliptical orbits (all orbits to 29° inclination) ............ 10

 2. Performance summary for the circular orbits (all orbits to 29° inclination) ............. 11

 3. Delta-V budget for circularization burn ................................................................... 12

 4. Delta-V budget for disposal burn ............................................................................ 12

 5. Circularization system for 220-nmi circular orbit from 

the 120-×220-nmi orbit ........................................................................................... 14

 6. Circularization system for 405-nmi circular orbit from 

the 120-×405-nmi orbit ........................................................................................... 15

 7. Disposal propellant and system mass sizing ............................................................. 16

 8. Effects of number of ML, HBs, and days in workweek for OSF option 

as it relates to flight rate and launch spacing, respectively ........................................ 42

 9. Risk elements .......................................................................................................... 46

10. Launch manifests with assigned activities (each of which incurs a risk 

indicated in table 1) .................................................................................................. 47

11. Loss-of-campaign risk for three architectures, with commercial 

tanker assumed to carry 20 t of propellant per launch, with 15 days 

between commercial launches .................................................................................. 48

12. Loss-of-campaign risk sensitivity against commercial tanker payload 

and gap between commercial tanker launches ......................................................... 49

13. Sample section of EDS functionality/capability matrix ........................................... 52

14. Technology investments assumed for chemical options in PA-C3′ ........................... 55

15. Current sample of international relationships ......................................................... 68
x

LIST OF ACRONYMS AND SYMBOLS

ACO Advanced Concepts Office

ACS attitude control system

AF Air Force

AIAA American Institute of Aeronautics and Astronautics

AR&D automated rendezvous and docking

ARF Assembly Refurbishment Facility

ATLAST Advanced Technology Large-Aperture Space Telescope

CaLV Cargo Launch Vehicle

CARD Constellation Architecture Requirements Document

CEV Crew Exploration Vehicle

CFM Cryogenic Fluid Management

CG center of gravity

CLV Crew Launch Vehicle

CMG Control Moment Gyroscope

CS core stage

CVS Concept Validation Study

CxP Constellation Program

CY calendar year

DAV descent/ascent vehicle

DDT&E design, development, test, and evaluation

DES discrete event simulation

DRA Design Reference Architecture

DRM Design Reference Mission

ECO engine cutoff

EDL entry, descent, and landing

EDS Earth Departure Stage

EPF Eastern Processing Facility

EOI Earth orbit insertion

ESAS Exploration Systems Architecture Study

ESMD Exploration Systems Mission Directorate

ETO Earth-to-orbit
xi

LIST OF ACRONYMS AND SYMBOLS (Continued)

FOM figure of merit

GG gravity gradient

GLONASS Global Navigation Satellite System

GN&C Guidance, Navigation, and Control

GOP Ground Operations Project

GPS Global Positioning System

GRAM Global Reference Atmospheric Model

He helium

HB high bay

HLLV heavy lift launch vehicle

IDAC-3 Integrated Design Analysis Cycle-3

IMLEO initial mass in low-Earth orbit

IMU inertial measurement units

ISRU in situ resource utilization

ISS International Space Station

KSC Kennedy Space Center

LAT Lunar Architecture Team

LCCR Lunar Capabilities Concept Review

LEO low-Earth orbit

LPE Launch Pad Element

LH2 liquid hydrogen

LOM loss-of-mission

LOX liquid oxygen

MAF Michoud Assembly Facility

MAV Mars ascent vehicle

MDAV Mars descent/ascent vehicle

MECO main engine cutoff

ML mobile launcher

MLE Mobile Launch Element

MMOD micrometeoroid and orbital debris

MOI Mars orbit insertion

MPS main propulsion system

MSFC Marshall Space Flight Center
xii

LIST OF ACRONYMS AND SYMBOLS (Continued)

MTV Mars transfer vehicle

NEO near-Earth object

NEP nuclear-electric propulsion

NERVA Nuclear Engine for Rocket Vehicle Application

NPF Nuclear Processing Facility

NSO nuclear safe orbit

NSSP Next Step in Strategic Partnership

NTP nuclear thermal propulsion

NTR nuclear thermal rocket

O&M operations and maintenance

OML outer mold line

OMS orbital maneuvering system

OMV orbiting maneuvering vehicle

Ops Con Operational Concepts

OSF Offline Stacking Facility

PA-C3 Phase A-Cycle 3

PA-C3′ Phase A-Cycle 3 Prime

PA-C3′′ Phase A-Cycle 3 Double Prime

PA-C3D Phase A-Cycle 3 point-of-departure concept

PBAN polybutadiene acrylonite

PMAD power management and distribution

PMBT propellant mean bulk temperature

POD point-of-departure

POST program to optimize simulated trajectories

PS payload shroud

QD quantity distance

R&D rendezvous and docking

RCS reaction control system

RPOD rendezvous, proximity operations, and docking

RPSF Rotation Processing and Surge Facility

S&MA Safety and Mission Assurance

SBU Sensitive But Unclassified

SHAB surface habitat
xiii

LIST OF ACRONYMS AND SYMBOLS (Continued)

SM service module

SPE Spacecraft Processing Element

SRB solid rocket booster

SRPE Solid Rocket Processing Element

SRR Systems Requirements Review

SRRE Spacecraft Recovery and Retrieval Element

SSC Stennis Space Center

SSME Space Shuttle Main Engine

SSPF Space Station Processing Facility

ST storage time

STS Space Transportation System

TCS thermal control system

TEI trans-Earth injection

TIM Technical Interface Meeting

TLI trans-lunar injection

TM Technical Memorandum

TMI trans-Mars injection

TransHab transit habitat

TRV3D trajectory vehicle, Cycle 3, D-vehicle configuration

VAB Vehicle Assembly Building

VI Vehicle Integration

VIE Vertical Integration Element

VIF Vertical Integration Facility

ZBO zero boiloff
xiv
1

TECHNICAL MEMORANDUM

ARES V UTILIZATION IN SUPPORT OF A HUMAN MISSION TO MARS

1. INTRODUCTION

Identified by numerous studies over the past 20+ yr as a necessary development that serves 

as the foundation upon which a human space flight exploration strategy must be built, a heavy lift 

vehicle like the Ares V has the potential to open the solar system to human exploration. Initial 

efforts by NASA’s Exploration Systems Mission Directorate (ESMD) and the Constellation Program (CxP) to utilize the performance capability offered by the Ares V launch vehicle for a Mars 

mission strategy was undertaken in 2007 for what was to become Mars Design Reference Architecture 5.0 (DRA 5.0). This early assessment showed that the performance capability of the Ares V 

was necessary in order to reduce the number of launches required for the Mars architecture to 

a reasonable level (7 flights for nuclear and 11 or 12 flights for chemical). 

With over 4 yr of increased knowledge of the Ares V concept, an evolved vehicle configuration, and established Ares V element teams in place for almost a year-and-a-half, the Ares V 

team assessed the capability offered by Ares V against the Mars Design Reference Mission (DRM) 

established in the Exploration Systems Architecture Study (ESAS)1 and later the Constellation 

Architecture Requirements Document (CARD). This provided valuable insight into the available 

performance capability to low-Earth-orbit (LEO) and the available performance growth options 

available; the required functionality to launch, maintain, and assemble a Mars transfer vehicle 

(MTV) in LEO; and the capability of the Earth Departure Stage (EDS) to possibly perform the 

trans-Mars injection (TMI) maneuver if required. In addition, it allowed the Ares V team to assess 

the ability of the vehicle to be built, stored, transported, and launched at the required rate for the 

Mars architecture. The initial assessment in these focus areas will be discussed in this Technical 

Memorandum (TM), and future work will be identified that will provide confidence that the Ares V 

launch vehicle can help meet the challenges presented by an exploration strategy for human missions to Mars.
2

2. MARS ARCHITECTURE OVERVIEW: DESIGN REFERENCE ARCHITECTURE 5.0

While the trade space is quite large for potential scenarios for Mars exploration missions, 

a valuable resource that was leveraged by the Ares V team was Mars DRA 5.0. This architecture 

for sustained human exploration of Mars was worked throughout 2007 and 2008 and publically 

released in 2009.2 As outlined in that study, a preferred approach for a human mission to Mars 

includes utilizing a conjunction class mission design (typically referred to as a ‘long-stay’ mission), 

predeploying cargo assets in both Mars orbit and on the surface before launching the crew, utilizing aerocapture for Mars orbit insertion (MOI) to some extent (in this case for the cargo assets), 

and utilizing in situ resource utilization (ISRU) to some extent in order to decrease the amount of 

required landed mass on the surface. These high-level trades were discussed in detail in DRA 5.0 

with the remaining top-level trade explored being the in-space propulsion technique used for the 

TMI maneuver and the MOI and trans-Earth injection (TEI) maneuvers. This top-level trade tree 

can be seen in figure 1.

Human Exploration

of Mars

Conjunction Class

Long Surface Stay

Predeploy Predeploy

Opposition Class

Short Surface Stay

Special Case

1-yr Round Trip

Aerocapture Propulsive Aerocapture Propulsive Aerocapture Propulsive Aerocapture Propulsive

ISRU No

ISRU ISRU NoISRU ISRU NoISRU ISRU NoISRU ISRU NoISRU ISRU NoISRU ISRU NoISRU ISRU NoISRU

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

NTR

Electric

Chemical

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35 36 37 38 39 40 41 42 43 44 45 46 47 48

NTR - Nuclear Thermal Rocket Electric=Solar or Nuclear Electric Propulsion

All-Up All-Up

Mission

Type

Cargo

Deployment

Mars Capture

Method

(Cargo)

Mars Ascent

Propellant

Interplanetary

Propulsion

(no hybrids

in Phase 1)

Figure 1. Mars DRA 5.0 top-level trade tree.
3

As shown in figure 2, the nuclear thermal propulsion (NTP) option was the preferred 

approach identified in DRA 5.0. The primary reason for this was the relatively large reduction in 

launches required using the NTP and chemical propulsion options explored in that study. While the 

NTP option required only 7 launches to meet the delta-velocity (dV) requirement for TMI with the 

payloads assumed, the chemical propulsion options required up to 12 launches. However, while the 

technology was recognized in DRA 5.0 for the benefit it would provide, these chemical options did 

not assess the potential of utilizing cryogenic propellant transfer in LEO. This option is the basis 

for the chemical options presented in this TM.

10

11

12

13

1

14

8

9

4

5

3

2

7

6

≈540 Days on Mars

≈26 mo ≈30 mo Orion Direct

Earth Return

ISRU/Propellant

Production for MAV

AC/EDL of MDAV/Cargo Lander

Habitat Lander AC

Into Mars Orbit

Cargo 

≈350 Days

to Mars

Cargo MTVs

Crewed MTV

4 Ares-V Cargo Launches 3 Ares-V Cargo Launches

Ares-1 Crew Launch

Crew: Jettisons TMI

Drop Tank ≈180 Day

Transfer Out to Mars

Crew: ≈180 Days

Back to Earth

Crew: Uses Orion/SM to

Transfer to Hab Lander:

Then EDL on Mars

MAV Ascent

to Orbit

Crew: Jettisons

DM and Contingency

Consumables Prior

to TEI

 Ares-V Mars Requirements:

• ≈140 t to LEO (407-km circ)

• Shroud Size: 12-m D×42.5-m L

• PL Envelope: 10-m D×35.5-m L

Figure 2. DRA 5.0 long-stay Mars mission overview (seven-launch NTP strategy).

The mission architecture presented in DRA 5.0 consists of the following 14 main 

characteristics: 

(1) Ares V launches components of two cargo vehicles to LEO. The first launch for each 

of the cargo vehicles delivers a “dual-use” shroud with an encapsulated payload (an ascent vehicle/

cargo lander for one MTV and a lander/surface habitat for the other MTV), and the remaining 
4

launches deliver the propulsive element needed to provide the dV necessary to inject that shroud 

and encapsulated payload on an outbound trajectory to Mars.

(2) These two cargo MTVs, after being assembled in LEO, await the most energy efficient 

alignment of Earth and Mars for the given synodic opportunity, and then the propulsive element 

performs the TMI maneuver. This maneuver places the cargo MTVs on a low-energy transfer that 

takes on the order of 9–12 mo to transfer between the Earth and Mars.

(3)–(5) The dual-use shroud is used for an aerocapture maneuver to insert the shroud and 

payload into a Mars orbit. The ascent vehicle/cargo lander performs an entry, descent, and landing 

(EDL) maneuver and begins to make propellant for the Mars-to-Mars-orbit ascent. The lander and 

crew habitat within the dual-use shroud wait in Mars orbit for crew arrival and eventual EDL with 

the crew inside.

(6) About 2 yr later (on the order of 26 mo), a crewed MTV is being assembled in LEO. 

The Ares V is once again responsible for launching very large payloads to LEO in support of this 

MTV, which consists of a transit habitat (TransHab) module, contingency consumables, solar 

arrays, a docking mechanism for rendezvous with an Orion crew capsule (launched on a crew 

launch vehicle (CLV)), MOI and TEI propulsive components, and TMI propulsive elements.

(7)–(8) Once this vehicle is assembled, a CLV delivers the crew to the crewed MTV for transit to Mars. While the cargo MTVs used a low-energy transfer to Mars, the crewed MTV performs 

a more demanding TMI maneuver to place the crew on a shorter duration transfer to Mars—on 

the order of 6 mo as opposed to 9–12 mo. This reduces the amount of time that the crew is subject 

to the space environment (microgravity, radiation, etc.).

(9) Propulsive capture (rather than aerocapture) is used to place the crew in Mars orbit. 

They then rendezvous and transfer to the lander and crew habitat, which then performs the EDL 

maneuver to the surface of Mars. The crewed MTV stays in Mars orbit for the eventual trip back 

to Earth.

(10)–(13) The crew stays on the surface of Mars for a long duration, on the order of 

540 days. Eventually, they board the ascent vehicle to transfer to Mars orbit, rendezvous, and 

transfer over to the waiting crewed MTV. The TEI maneuver is performed, which is a higher than 

minimum-energy transfer that places them on an approximately 6-mo-long trajectory back 

to Earth.

(14) The crew utilizes the Orion capsule for reentry back to the surface of Earth. The total 

crew trip duration is about 900–1,000 days. 
5

3. ARES V OVERVIEW: LAUNCH VEHICLE OPTIONS

3.1 Ares V Concept History

Originally defined in ESAS as the cargo launch vehicle (CaLV), the Ares V concept was 

envisioned as a 27.5-ft-diameter/Space Shuttle Main Engine (SSME) based, heavy-lift launch 

vehicle with two solid rocket boosters (SRBs) (similar to the Space Transportation System (STS)). 

However, subsequent studies showed a single-engine upper stage, referred to as the EDS, increased 

payload injection to LEO and represented an optimized solution as a translunar injection (TLI) 

stage. As knowledge of functionality required, costs, schedule, and performance requirements 

matured, the vehicle underwent design changes leading up to the Lunar Capabilities Concept 

Review (LCCR) (“ESAS-LCCR Ares V; Ares V Pre-Phase-A Refinements,” Powerpoint Presentation, 2009) in the summer of 2008. This served as the Mission Concept Review for Ares V and 

technically served as the transition from conceptual design studies to an Ares V Phase A concept. 

Currently, the Ares V is moving toward a Systems Requirements Review (SRR) in July 2011 that 

will formally begin the transition to a Systems Design Review (SDR) and into a Phase B project. 

Figure 3 shows the progress and revisions made on the Ares V vehicle since ESAS.

• 45.0-t Lander

• 20.0-t CEV

• No Loiter in LEO

• 8.4-m OML

• 5 SSMEs/2 J2S

• Launch Vehicle 27.3

• Increase Ares I/

 Ares V Common-

 ality

• Ares I: 5 Seg 

 RSRB/J-2X

 Instead of air-

 start SSME

• Ares V: 1 J-2X

• ≈$4.25-B Life Cycle

 Cost Savings

• 5-Engine Core

• Increased Common-

 ality With Ares I 

 Booster

• 30–95 Days LEO

 Loiter Assessed

• Lunar Architecture

 Team (LAT) 1/2 

 Studies

• Mission Delta-Vs

 Increased

• Increase Margins

 From TLI Only to

 Earth Through TLI

• Loiter Penalties for

 30-Day Orbit

 Quantified

• LAT 1/2 Studies

• Lunar/Mars

 Systems Benefit

• Tank Assembly

 Tooling Common-

 ality

• Transition From Conceptual 

 Design to Engineering Assess-

 ment (57.01.14 ACO Vehicle to 

 Engineering Concept Validation 

 Study (CVS) Vehicle)

• PA-C1 Vehicle: TRV1-A

• Engineering Shortfall Assessment

 Defined a “Lower Bound” for

 Ares V Performance

• PA-C2 Focused on Understanding

 Sensitivities and Further Charac-

 terizing Performance

• PA-C2 Assessed Vehicle: TRV2

• 6 Core Engines

• 5.5-Segment PBAN

• 45.0-t Lander

• 20.2-t CEV

• ≈6-t Perf. Margin

• 4-Day LEO Loiter

• Ares I Common MGAs

Original ESAS

Capability

CY-06 Budget

Trade

Detailed Trade

of SSME

Verses RS-68

IDAC-3

Trade Space

EDS Diameter

Change From

8.4 m to 10 m

LCCR

Recommended

Option (51.00.48) Ares V Road to SRR

LCCR Capability

2005 2006 2007 2008 2009

ESAS LCCR Ares V MCR PA-C1 F2F PA-C2 F2F

Figure 3. Timeline of major Ares V changes/reviews since ESAS.
6

3.2 Concepts Assessed For Mars Design Reference Architecture 5.0

Performance analysis for the Ares V launch vehicle was assessed for DRA 5.0 during the 

summer and fall of 2007. During that timeframe (about 1.5 yr after ESAS), Ares V was involved in 

what was known as Integrated Design Analysis Cycle-3 (IDAC-3). The Ares V had recently transitioned from a 27.5-ft diameter/SSME based core stage (CS) to a 33-ft diameter/RS-68 engine based 

CS. At that time, many potential booster options, various CS lengths and diameters, etc. were being 

explored. For DRA 5.0, a reference vehicle was chosen that incorporated five RS-68B engines on 

a 33-ft diameter CS, an Ares I-like five-segment polybutadiene acrylonite (PBAN) SRB, a single 

J-2X on a 27.5-ft EDS, and a 27.5-ft diameter shroud (dubbed the 45.0.2 concept). While keeping 

that base configuration, several other vehicle options were assessed for “payload to LEO” sensitivity, as well as several orbit options for payload delivery.

A very important finding during the DRA 5.0 assessment was the characterization of the 

impact of the larger shrouds required for delivering ‘Mars-sized’ payloads to LEO. Several shroud 

options were assessed that varied both diameter and overall length. In addition, the dual-use 

shroud was assessed, which provided Earth-to-orbit (ETO) ascent protection, in-space thermal 

protection, MOI thermal protection, and thermal protection to the payload during the Mars entry, 

descent, and landing phase. Furthermore, a fuel-stage delivery ‘nose cone only’ type of shroud was 

assessed for LEO propellant delivery options. Figure 4 depicts a few of these shroud options, but 

they are discussed more fully in the DRA 5.0 addendum and follow-on technical papers listed.

3.3 Concept Assessed During Phase A-Cycle 3′

As analysis on the Ares V concept continued through to the more recent Ares V analysis 

cycle (Phase A-Cycle 3′ (PA-C3′)), the Mars mission architecture remained locked to that presented 

in Mars DRA 5.0, including two cargo MTVs that are predeployed (utilizing aerocapture for MOI) 

before a crewed MTV departs on the following synodic opportunity (utilizing propulsive capture 

for MOI). As such, MTV options were chosen within that architecture to include nuclear thermal 

propulsion (NTP) and chemical propulsion options intended to reduce the number of flights. The 

Ares V launch vehicle was employed to deliver components for these MTVs, such as payload and 

propellant. Each propulsion option required a different number of flights (e.g., NTP consisted 

of a seven-launch campaign as presented in DRA 5.0 and a competitive chemical option that was 

found during PA-C3′ consisted of a nine-launch campaign). 

During PA-C3, a range of concepts was assessed that represented the best estimate for 

Ares V concepts that could potentially balance all of the performance, cost, and schedule figures 

of merit (FOMs) for the lunar DRMs. Among these, a concept emerged as a preferred point-ofdeparture (POD) concept—PA-C3D. This concept utilized five-segment ‘fast-burn’ PBAN SRBs, 

a 33-ft diameter CS with five RS-68B-E/O regeneratively cooled engines, a 33-ft diameter EDS with 

a single J-2X, and a 33-ft diameter composite tangent-ogive shroud. This concept was the chosen 

option to assess against the Mars DRM. The major attributes of this concept can be seen in 

figure 5.
7

Option C

142.56 m

Option B

128.45 m

Option A

110.31 m

120 m

100 m

80 m

60 m

40 m

20 m

0 m

VAB Door Height 118.9 (390)

Figure 4. Potential shroud options assessed for DRA 5.0.3

In identifying the value of assessing the Mars DRM, four fundamental questions were 

posed. The answers to the following four questions aided in characterizing the synergy available 

between the lunar and Mars architectures, mainly the utilization of a common Ares V launch 

vehicle (with required modifications identified at a qualitative level):

(1) What is Ares V PA-C3D initial mass in LEO (IMLEO) capability as a function of 

altitude? Assess a range of orbit options (elliptical phasing, elliptical disposal, and direct inject to 

circular orbits) to determine the overall vehicle performance (with system estimates for circularization and disposal at the various options).

(2) What are the IMLEO growth potential/options for Ares V PA-C3D? Identify reasonable 

growth paths for PA-C3D if more performance is required.

(3) What is the impact of in-space pre-TMI burn functions on EDS design? Identify 

assumed functions required to insert and maintain the payload until the TMI maneuver.

(4) What are the extensibility options for Ares V EDS to the Mars TMI vehicle?

Identify assumed functions required to perform the TMI maneuver.
8

Payload Shroud (PS)

Predicted Dry Mass

15,142 lbm (6,868 kg)

Manufacturing Variation

2.00% - 303 lbm (137 kg)

Earth Departure Stage (EDS)

Propellants

LOX/LH2

Loaded Propellants

555,394 lbm (251,922 kg)

Available for Impulse

545,468 lbm (242,665 kg)

Predicted EDS Dry Mass

67,769 lbm (30,739 kg)

EDS Manufacturing Variation

2.00% - 1,355 lbm (615 kg)

Predicted EDSE Dry Mass

5,544 lbm (2,515 kg)

EDSE Manufacturing Variation

1.02 % - 57 lbm (26 kg)

Solid Rocket Booster (SRB)

Propellant

PBAN

Characteristics

Steel Case, 8-673 Trace

Loaded Propellant (each)

1,387,203 lbm (629,225 kg)

Predicted Inert Mass (each)

239,222 lbm (180,509 kg)

Manufacturing Variation

0.94% - 2,249 lbm (1,020 kg)

No. of Boosters/Type

2/5-Segment SRMs + Spaces

(100s9_673TROG_Fto40K)

Integrated Vehicle

Ares V Gross Weight

7,722,450 lbm (3,502,845 kg)

TRV-3-D min Gross 

Payload Delivered

142,968 lbm (64,849 kg)

LEO Delivery

241 km (130 Nmi) circ at 29.0%

Altair Mass (Launch -LEO)

99,209 lbm (45,000 kg)

Altair Mass (Post-TLI)

99,107 lbm (44,954 kg)

Orion Mass

44,500 lbm (20,185 kg)

FPR

9,799 lbm (4,445 kg)

Fuel Bias

683 lbm (310 kg)

Boiloff

7,271 lbm (3,298 kg)

Approximate Burnout 

Mass With Engine

75,514 lbm (38,760 kg)

No. Engines/Type

1/J-2X

Primary Mode Thrust (Ascent)

294,000 lbf (1,307,777 N)

Secondary Mode Thrust (TLI)

239,166 lbf (1,063,863 N)

Primary Mode I

sp (Ascent)

448 s (4,393 4 N-s/kg)

Secondary Mode I

sp (TLI)

448 s (4,393 4 N-s/kg)

Core Stage (CS)

Propellant

LOX/LH2

Loaded Propellants

3,399,091 lbm (1,541,797 kg)

Available for Impulse

3,346,736 lbm (1,530,831 kg)

Predicted CS Dry Mass

254,353 lbm (115,373 kg)

CS Manufacturing Variation

1.00% - 2,544 lbm (1,154 kg)

Predicted CSE Dry Mass (each)

15,161 lbm (6,872 kg)

CSE Manufacturing Variation

1.03% - 156 lbm (71 kg)

FPR

None: All Carried by EDS

Fuel Base

4,942 lbm (3,837 kg)

Approximate Burnout 

Mass With Engine

397,604 lbm (130,350 kg)

No. Engines/Type

5/RS-68-B-E/0

Engine Thrust Modes (each)

108% - 3,603,087 N (810,005 lbf)

57% - 1,929,474 N (433,763 lbf)

Variation - 18,233 N (4,099 lbf)

Engine I

sp Modes (each)

108% - 4,148 N-s/kg (423.0 s)

57% - 14,169 N-s/kg (425.1 s)

Variation - 25 N-s/kg (2.6 s)

PA-C3D

72.1

(22.0)

77.0

(23.5) 235.0

(31.4)

194.5

(59.3)

395.4

(120.51)

395.4

(120.51)

160.3

(148.9)

83.3

(25.4)

11.3

(3.4)

384.1

(117.1)

Figure 5. PA-C3D Ares V concept (baseball card).

3.4 Key Mass Assumptions For Phase A-Cycle 3′ Mars Design Reference Mission Assessment

Several key high-level assumptions were made to facilitate the analysis performed relative to 

the Mars DRM. First, it was determined that available performance would be decremented by the 

mass of the system required to circularize the payload to the final assembly altitude. This allowed 

the EDS to be decoupled from that particular system until a final delivery orbit and assembly orbit 

were determined. Furthermore, the mass allocation for the system required to jettison and dispose 

of the empty EDS was taken from the available performance for similar reasons. These systems will 

be further described in the following section.

At the EDS element level, the entire burnout mass of the stage was upsized by 18% (from 

25.5 t to 30 t) to account for known system level impacts that could not be analyzed during the 
9

time constraints of PA-C3′. This 4.5 t was an initial estimate to account for structural growth to 

accommodate the larger payloads and payload shrouds (PSs), upsized reaction control system/attitude control system (RCS/ACS) components for increased loiter durations and larger payloads, and 

other hardware impacts.

Even after taking these performance penalties (and others as discussed in the next section) 

into consideration, the PA-C3D configuration delivered more than required in support of the Mars 

DRM as outlined in the CARD. This CARD requirement of 125 t of payload was used as the basis 

for constructing the mission scenarios presented herein. Of the 125 t of available payload, when onorbit propellant transfer was utilized, an additional 5-t decrement was implemented to account for 

rendezvous/docking hardware, transfer lines, etc., both on the delivering tank side and the receiving 

EDS side. Obviously, more detailed analysis is required to support the mass allocations given to 

allow for these EDS capabilities.

3.5 Performance of Phase A-Cycle 3D for Low-Earth Orbit Applications

Throughout the PA-C3 and PA-C3′ analysis, there have been several trajectory types that 

have been tailored based on the types of analysis being performed along with the available time to 

complete the analysis. Different studies use the trajectory type that is most applicable to the study’s 

goals. The trajectory type chosen for the IMLEO study was the performance reference trajectory. 

While this is similar to the minimum performance trajectory, the main difference is that a knockdown is calculated for the minimum performance reference trajectory and not for the performance 

reference. Therefore, the performance reference is not applicable for comparing to requirement 

metrics and the CARD. The data produced in the IMLEO analysis is more suited for comparing 

between the different types of orbits and launch conditions than comparing directly to a target 

mass value.

The payload to LEO is maximized for a performance reference trajectory. The liquid 

engines, RS-68s and the J-2X, are run at the minimum guaranteed specific impulse (Isp) and nominal thrust levels. The SRB trace is degraded following Shuttle methods and is based on the nominal 

burn rate at a 61-degree propellant mean bulk temperature (PMBT). Other characteristics of this 

trajectory type include February GRAM mean monthly winds and atmosphere, predicted mean 

mass, and the aerodynamic database in the nominal as delivered form. For all PA-C3 and PA-C3′

trajectories, the maximum dynamic pressure was limited to 800 lbf psf via throttling of the RS-68s. 

For the IMLEO study, the performance reference trajectories consisted of a single ascent 

phase, generated in the program “Program to Optimize Simulated Trajectories” (POST). POST is 

a generalized point-mass, discrete-parameter, targeting and optimization program. It provides the 

capability to target and optimize point-mass trajectories for a powered or unpowered vehicle near 

an arbitrary rotating, oblate planet. POST has been used successfully to solve a wide variety of 

atmospheric ascent and reentry problems. Its generality is evidenced by its multiple phase simulation capability, which features generalized planet and vehicle models. This flexible simulation capability is augmented by an efficient discrete parameter optimization capability that includes equality 

and inequality constraints.
10

Mass to LEO was maximized for five different orbits, two elliptical orbits of 120×220 nautical miles (nmi) and 120×405 nmi, and three circular orbits with altitudes of 130 nmi, 220 nmi, and 

405 nmi. The performance summaries for the two elliptical orbits are presented in table 1. The 

insertion altitude was allowed to vary and the optimal insertion for both elliptical orbits was near 

the perigee, though not exactly at perigee. Propellant was off-loaded from the fully loaded EDS 

tanks in order to optimize the performance to LEO. The optimal loading was around 70% for the 

two elliptical cases. The 120-×220-nmi orbit delivered 175 t, while the 120-×405-nmi orbit delivered 171 t. These performance values include everything taken to this orbit (the EDS burnout mass, 

circularization hardware and propellant, RCS/ACS propellant load, and all necessary disposal 

masses). It is the mass that arrives at these orbits and should be divided among the different parts 

as warranted.

Table 1. Performance summary for the elliptical orbits (all orbits to 29° inclination).

Insertion Orbit 120×220 nmi 120×405

Vehicle designation TRV3D TRV3D

Trajectory description Performance Reference—

PA-C3D Vehicle Configuration

Performance Reference—

PA-C3D Vehicle Configuration

Trajectory date 4/9/2010 4/12/2010

Total mass at RS-68B ignition (lbm) 7,746,845.3 7,751,465.5

 SRB loaded propellant 473,947 473,947

 SRB inert mass 2,774,406 2,774,406

 Core stage impulse propellant 3,351,666 3,351,666

 Core stage jettison mass 375,718.4 375,718.4

 Shroud jettison mass 14,268 14,268

 EDS ascent impulse propellant 358,609.3 371,172.9

 EDS ascent FPR propellant 13,434.7 13,306.6

Gross LEO mass is everything delivered to LEO: Includes EDS burnout mass, circularization hardware and propellant, RCS/ACS 

 propellant load, and disposal mass. It is not the final payload of the vehicle and not the value to compare to requirements.

Gross LEO mass (lbm) 384,795.9 376,980.6

Gross LEO mass (kg) 174,540.5 170,995.5

Total core usable propellant (impulse + start-up) 3,351,666 3,351,666

Total EDS usable propellant (impulse + FPR) 372,044 384,479.5

EDS usable propellant capacity (lbm) 544,784.8 544,784.8

Actual EDS propellant loading (%) 68.3 70.6

Delta EDS propellant (lbm) –172,740.8 –160,305.4

J-2X burn time (s) 550.3 569.3

Insertion altitude (nmi) 121.5 122.1

Maximum acceleration (gs) 3.36 3.35

Maximum dynamic pressure (psf) 800 800

The performance summaries for the circular orbits are presented in table 2. The 130-nmi 

and 220-nmi orbits also used approximately 70% of the loaded EDS propellant, while the 405-nmi 

orbit used 83%. The mass delivered to the orbits was 176 t to the 130-nmi orbit, 166 t to the 
11

220-nmi orbit, and 145 t to the 405-nmi orbit. The performance dropped off significantly with the 

higher circular orbit altitudes. Since the 120-×405-nmi orbit achieved 26 t more than a direct insertion into the 405-nmi orbit, it would be advantageous to examine the costs of circularizing from 

the 120-×405-nmi orbit to the desired final circular orbit. 

Table 2. Performance summary for the circular orbits (all orbits to 29° inclination).

Insertion Orbit 130-nmi Circular 220-nmi Circular 405-nmi Circular

Vehicle designation TRV3D TRV3D TRV3D

Trajectory description Performance Reference—

PA-C3D Vehicle Configuration

Performance Reference—

PA-C3D Vehicle Configuration

Performance Reference—

PA-C3D Vehicle Configuration

Trajectory date 4/22/2010 4/22/2010 4/22/2010

Total mass at RS-68B ignition (lbm) 7,745,468.5 7,753,871.7 7,759,173.6

 SRB loaded propellant 473,947 473,947 473,947

 SRB inert mass 2,774,406 2,774,406 2,774,406

 Core stage impulse propellant 3,351,666 3,351,666 3,351,666

 Core stage jettison mass 375,718.4 375,718.4 375,718.4

 Shroud jettison mass 14,268 14,268 14,268

 EDS ascent impulse propellant 354,786.4 384,108.7 438,016.5

 EDS ascent FPR propellant 13,473 13,113.9 12,148.2

Gross LEO mass is everything delivered to LEO: Includes EDS burnout mass, circularization hardware and propellant, RCS/ACS propellant load, and disposal mass. 

 It is not the final payload of the vehicle and not the value to compare to requirements.

Gross LEO mass (lbm) 387,203.8 366,643.7 319,003.5

Gross LEO mass (kg) 175,632.7 166,306.8 144,697.5

Total core usable propellant (impulse + start-up) 3,351,666 3,351,666 3,351,666

Total EDS usable propellant (impulse + FPR) 368,259.4 397,222.6 450,164.7

EDS usable propellant capacity (lbm) 544,784.8 544,784.8 544,784.8

Actual EDS propellant loading (%) 67.6 72.9 82.6

Delta EDS prop (lbm) –176,525.5 –147,562.2 –94,620.1

J-2X burn time (s) 544.5 589.1 671.4

Insertion altitude (nm) 130 220 405

Maximum acceleration (gs) 3.4 3.3 3.3

Maximum dynamic pressure (psf) 800 800 800

The final goal is to insert the payload into an ending circular orbit with the elliptical orbits 

considered as transfer orbits. This is accomplished by performing a single circularization burn at 

the apogee of the elliptical orbits. For preliminary analysis purposes, a dV budget was calculated 

for this burn for both of the elliptical orbits and is presented in table 3. In order to approximate the 

impact of gravity losses, the pure impulsive dV was increased by 5%. The circularization systems on 

the vehicle will need the capability to cover at least the amount of the listed dV.

Some of the equipment and mass that arrive in the circular orbit may not need to remain in 

the final orbit (i.e., the circularization system and EDS); hence, these items will need to be disposed 
12

Table 3. Delta-V budget for circularization burn.

Desired Circular Altitude 

(nmi)

Elliptical Perigee Altitude 

(nmi)

Elliptical Apogee Altitude 

(nmi)

dV Budget

(ft/s)

220 120 220 183

405 120 405 500

of properly. For this study, the disposal method consisted of a single deorbit burn that adjusts the 

circular orbit into an elliptical orbit with a 30-nmi altitude perigee. This perigee altitude is low 

enough to have the equipment reenter the atmosphere without any additional burns. The dV budgets recommended for these deorbit burns are presented in table 4. As with the circularization dV

budget, a 5% increase was included to approximate the impact of gravity losses.

Table 4. Delta-V budget for disposal burn.

Starting Circular Altitude

(nmi)

Deorbit Elliptical Perigee

(nmi)

dV Budget

(ft/s)

130 30 190

220 30 354

405 30 668

3.6 Payload Performance to Final Circular Orbit

As already discussed, the ascent performance results shown in tables 1 and 2 represent the 

total weight injected into each insertion orbit analyzed. The EDS dry weight along with the propellant and system weights necessary to circularize the payload into its final orbit must come out 

of this total weight. Additionally, the system must be sized to deorbit the EDS, provided it is not 

being used to support TMI. Figure 6 illustrates the conceptual mission profile assumed for the calculation of payload to the final circular orbit altitude.

The large number of possible on-orbit engine configurations led to a parametric sweep 

of the Isp and mass fraction of this system. The sweep included an Isp range of 280 through 450 s 

along with propellant mass fractions between 0.5 through 0.8. To evaluate the circularization 

requirements for the 220-nmi and 405-nmi apogee elliptical orbits, the dVs from table 3 were 

assumed. Using the Isp and required dV, the ideal rocket equation is used to find the propellant 

necessary to perform circularization. That propellant amount is then used with the assumed mass 

fraction to calculate the system’s inert weight. Two sets of numbers were generated for each insertion orbit, one representing the case where the entire EDS is circularized with the other sizing the 

system when the EDS is not circularized with the rest of the vehicle. For all cases, including direct 

circular insertion, the EDS is assumed to have an inert mass of 66,138 lbm, (≈30 t) not including 
13

Elliptical Phasing Orbit

Final Circular Altitude

Coast Period

Highly Exaggerated—Not to Scale

MECO1

dVdeorbit

dVcirc

Figure 6. Conceptual LEO mission profile.

the on-orbit propulsion system. Results of the circularization system sizing to the 220-nmi altitude 

circular orbit from the 120-×220-nmi orbit are presented in table 5. The sizing results from the 

120-×405-nmi orbit to the 405-nmi circular orbit are presented in table 6. Note: The payloads 

listed in tables 5 and 6 do not include the disposal system mass.

System requirements for the disposal of an empty EDS were then calculated separately 

using the dV information in table 7 using the same methodology, Isp, and mass fraction ranges 

from the circularization parametric sweep. The propellant and system mass results for disposal 

of an empty EDS from the three circular obits can be found in table 7.
14

Table 5. Circularization system for 220-nmi circular orbit from the 120-×220-nmi orbit.

Circularization performed with Total stack (Payload + Empty EDS) (lbm)

pmf

I

sp (s)

280 300 350 400 450

Circular system burnout mass 0.5 7,738 7,227 6,203 5,433 4,833

Circular system propellant mass 7,738 7,227 6,203 5,433 4,833

Payload (subtotal) 303,170 304,192 306,240 307,780 308,980

Circular system burnout mass 0.6 5,159 4,818 4,135 3,622 3,222

Circular system propellant mass 7,738 7,227 6,203 5,433 4,833

Payload 305,750 306,601 308,308 309,591 310,591

Circular system burnout mass 0.7 3,316 3,097 2,658 2,328 2,071

Circular system propellant mass 7,738 7,227 6,203 5,433 4,833

Payload 307,592 308,322 309,785 310,885 311,742

Circular system burnout mass 0.8 1,934 1,807 1,551 1,358 1,208

Circular system propellant mass 7,738 7,227 6,203 5,433 4,833

Payload 308,974 309,613 310,893 311,855 312,605

Circularization performed with Payload ONLY (lbm)

pmf

I

sp (s)

280 300 350 400 450

Circular system burnout mass 0.5 6,408 5,984 5,136 4,499 4,002

Circular system propellant mass 6,408 5,984 5,136 4,499 4,002

Payload 305,831 306,677 308,373 309,648 310,642

Circular system burnout mass 0.6 4,272 3,990 3,424 2,999 2,668

Circular system propellant mass 6,408 5,984 5,136 4,499 4,002

Payload 307,967 308,672 310,085 311,148 311,976

Circular system burnout mass 0.7 2,746 2,565 2,201 1,928 1,715

Circular system propellant mass 6,408 5,984 5,136 4,499 4,002

Payload 309,492 310,097 311,308 312,219 312,929

Circular system burnout mass 0.8 1,602 1,496 1,284 1,125 1,001

Circular system propellant mass 6,408 5,984 5,136 4,499 4,002

Payload 310,637 311,165 312,225 313,022 313,643
15

Table 6. Circularization system for 405-nmi circular orbit from the 120-×405-nmi orbit.

Circularization Performed With Total Stack (Payload + Empty EDS) (lbm)

pmf

I

sp (s)

280 300 350 400 450

Circular system burnout mass 0.5 20,353 19,031 16,372 14,365 12,797

Circular system propellant mass 20,353 19,031 16,372 14,365 12,797

Payload 270,125 272,769 278,086 282,100 285,238

Circular system burnout mass 0.6 13,569 12,687 10,915 9,577 8,531

Circular system propellant mass 20,353 19,031 16,372 14,365 12,797

Payload 276,909 279,113 283,544 286,889 289,503

Circular system burnout mass 0.7 8,723 8,156 7,017 6,157 5,484

Circular system propellant mass 20,353 19,031 16,372 14,365 12,797

Payload 281,755 283,644 287,442 290,309 292,550

Circular system burnout mass 0.8 5,088 4,758 4,093 3,591 3,199

Circular system propellant mass 20,353 19,031 16,372 14,365 12,797

Payload 285,390 287,042 290,366 292,874 294,835

Circularization Performed With Payload Only (lbm)

pmf

I

sp (s)

280 300 350 400 450

Circular system burnout mass 0.5 16,782 15,692 13,499 11,845 10,551

Circular system propellant mass 16,782 15,692 13,499 11,845 10,551

Payload 277,268 279,448 283,832 287,142 289,729

Circular system burnout mass 0.6 11,188 10,461 9,000 7,896 7,034

Circular system propellant mass 16,782 15,692 13,499 11,845 10,551

Payload 282,862 284,678 288,332 291,090 293,246

Circular system burnout mass 0.7 7,192 6,725 5,785 5,076 4,522

Circular system propellant mass 16,782 15,692 13,499 11,845 10,551

Payload 286,857 288,414 291,546 293,910 295,758

Circular system burnout mass 0.8 4,195 3,923 3,375 2,961 2,638

Circular system propellant mass 16,782 15,692 13,499 11,845 10,551

Payload 289,854 291,216 293,957 296,025 297,642
16

Table 7. Disposal propellant and system mass sizing.

Disposal Performed With Empty EDS From 130-nmi Circular (lbm)

pmf

I

sp (s)

280 300 350 400 450

Circular system burnout mass 0.5 1,381 1,289 1,107 969 862

Circular system propellant mass 1,381 1,289 1,107 969 862

Circular system burnout mass 0.6 920 860 738 646 575

Circular system propellant mass 1,381 1,289 1,107 969 862

Circular system burnout mass 0.7 592 553 474 415 370

Circular system propellant mass 1,381 1,289 1,107 969 862

Circular system burnout mass 0.8 345 322 277 242 216

Circular system propellant mass 1,381 1,289 1,107 969 862

Disposal Performed With Empty EDS From 220-nmi Circular (lbm)

pmf

I

sp (s)

280 300 350 400 450

Circular system burnout mass 0.5 2,549 2,382 2,047 1,795 1,598

Circular system propellant mass 2,549 2,382 2,047 1,795 1,598

Circular system burnout mass 0.6 1,699 1,588 1,365 1,197 1,065

Circular system propellant mass 2,549 2,382 2,047 1,795 1,598

Circular system burnout mass 0.7 1,092 1,021 877 769 685

Circular system propellant mass 2,549 2,382 2,047 1,795 1,598

Circular system burnout mass 0.8 637 596 512 449 399

Circular system propellant mass 2,549 2,382 2,047 1,795 1,598

Disposal Performed With Empty EDS From 405-nmi Circular (lbm)

pmf

I

sp (s)

280 300 350 400 450

Circular system burnout mass 0.5 4,728 4,423 3,810 3,346 2,983

Circular system propellant mass 4,728 4,423 3,810 3,346 2,983

Circular system burnout mass 0.6 3,152 2,949 2,540 2,231 1,988

Circular system propellant mass 4,728 4,423 3,810 3,346 2,983

Circular system burnout mass 0.7 2,026 1,896 1,633 1,434 1,278

Circular system propellant mass 4,728 4,423 3,810 3,346 2,983

Circular system burnout mass 0.8 1,182 1,106 952 836 746

Circular system propellant mass 4,728 4,423 3,810 3,346 2,983
17

To provide a baseline estimate of the usable payload delivered to each circular orbit, the 

impacts of both the circularization and disposal systems were combined assuming a propellant 

mass fraction of 0.8 and an Isp of 300 s.4 These values are similar to the Shuttle orbital maneuvering system and provide a reasonable first guess for this type of system. The payload curves in figure 7 include a couple of important assumptions. First is the assumption of an EDS inert mass 

of 66,138 lbm (30 t), which may not include all necessary weight changes necessary to support the 

Mars DRM. Second, due to the use of the performance reference trajectory instead of a minimum 

performance trajectory, an estimated LEO performance knockdown factor of 2,645 lbm (1.2 t) 

is included to approximate the minimum performance trajectory. The net result is that this early 

estimate provides confidence that the PA-C3D configuration can deliver >125 t to LEO in support 

of the Mars DRM.

As part of the PA-C3′ Mars DRM assessment activity, reasonable vehicle performance 

growth options were captured in trade tree format for future assessment if payload mass grows 

beyond the performance capability of the PA-C3D configuration’s capability. These growth options 

were not explored in detail during this study, but in most cases the thousands of conceptual designs 

that have been performed on the Ares V concept since ESAS have characterized the general performance increase that they could potentially provide. Figure 8 is a representative trade tree of 

the growth options that would be considered for increased LEO performance capability.

150

140

130

120

110

100

90

80

0 50 100 150 200 250 300 350 400 450

Final Payload (t)

Final Circular Orbit (nmi)

Direct Circular Injection

Elliptical Transfer Orbit With EDS Mass Removed

Ellipitcal Transfer Orbit With EDS, Circ And Disposal Masses Removed

CARD Requirement

Figure 7. Representative LEO payload using Shuttle OMS assumptions. 

A 1.2-t knockdown assumption is included in the data.
18

Trajectory Optimization for LEO

(Orbits/Throttling/Engine Cutoff/etc.)

Shroud Options

Increased Propellant Options

General

Larger Diameter

Densified Propellants

Increased Height

(Requires VAB Modification)

AlLi Alloy (2195 or 2050)

Composite Materials in NonCryogenic/Unpressurized

Structure

RS-68B-E/0

Other High Thrust/High

I

sp LOX/LH2 Engine Engine Options

Structure Options

Core Stage Options

Booster Options

EDS Options

Performance Growth Options

Number/Type of Engines

Number of Engines

Nonrecoverable

Composite Case

Optimized Thrust Trace

Nozzle Optimization

Increased Propellant Mass

HTPB Propellant

Other Booster Types

Stage Optimized for

LEO Payload Delivery

Figure 8. LEO performance growth options.
19

4. MARS MISSION ANALYSIS: MARS TRANSFER VEHICLE OPTIONS

As part of the initial ground rules and assumptions set at the beginning of the PA-C3′

cycle for the Mars DRM activity, it was determined that all post-TMI payloads would be set to 

their assessed values during the DRA 5.0 study and follow-on activities. This locked the general 

parameters for the main activity during the 7-wk assessment period of PA-C3′, which dealt with 

understanding the requirements of the Ares V vehicle to launch and possibly maintain the main 

propulsion system for the TMI maneuver (and perform the TMI maneuver if necessary). Two main 

propulsion system types were assessed for applicability to the TMI maneuver for the architecture 

presented in DRA 5.0: NTP and chemical propulsion (liquid oxygen/liquid hydrogen (LOX/LH2)). 

Figure 9 shows a general breakdown of the options considered throughout PA-C3′.

NTP Option

(7 Launches)

Chemical Options

(9–11 Launches)

TMI Modules

(11 Launches)

EDS With Residuals

(11 Launches)

In-Space Cryogenic

Propellant Transfer

EDS Refuel

(10 Launches)

Scavenger

(9 Launches)

Commercial Tanker

(4 Launches + TBD

Commercial Launches)

Figure 9. TMI stage propulsion options assessed during PA-C3′.

As the launch count required for the various options was assessed, the “carry-forward” 

options were further condensed to depict the most attractive chemical propulsion options from 

a launch count and lunar-EDS evolution perspective. As shown in figure 10, these include both 

“minimum Ares V participation” and “maximum Ares V participation” NTP options and a chemical propulsion option that utilized in-space cryogenic propellant transfer and its derivative that 

relied on commercial vendors to supply the needed propellant. These four options will be the focus 

of this TM.

4.1 Nuclear Thermal Propulsion Option

The NTP option is considered the reference approach for the DRA 5.0 architecture, and 

it will be referenced as a specific implementation of that technology based on DRA 5.0 in this 

TM (engine type, engine performance parameters, propellant/stage mass assumptions, subsystem 

assumptions, etc.).5 The general launch sequencing is depicted in figure 11.
20

NTP Option

(7 Ares V Launches) Chemical Option

Min. Ares V

Participation

Max. Ares V

Participation

In-Space Cryogenic

Propellant Transfer

Scavenger

(9 Ares V Launches)

Commercial Tanker

(4 Ares V Launches + TBD

Commercial Launches)

Figure 10. TMI propulsion options discussed.

Cargo SHAB MTV (IMLEO 236 t)

Cargo DAV MTV (IMLEO 236 t)

Crewed MTV (IMLEO 326 t)

Payload: Crew Habitat + Orion

Mass: ≈60 t

• Total Ares V Launches: 7 (4 for two Cargo MTVs and 3 for Crewed MTV)

• Launch Rate: ≈4–6 per yr per mission

• Conjunction-class long duration mission: ≈490–570 days at Mars

• Earth orbit: 407-km circular; Mars orbit: 250 km x 33,793 km (ref. DRA5.0)

Delta-V Requirements (included for reference only)

• Mission dV Budget (largest total dV across 18-yr synodic opp. (2028–2046)):

 • Cargo: Earth Departure dV ≈3.9 km/s; C3 ≈15 km2/s2

 • Crew: Earth Departure dV ≈4.1 km/s; C3 ≈19 km2/s2

Payload: NTP Module, Propellant Drop Tank and Truss

Mass: 123 t

Payload: NTP Stage and Propellant

Mass: 139 t

Payload: NTP Stage and Propellant

Mass: 133.2 t

Payload: NTP Stage and Propellant

Mass: 133.2 t

Launch 1

Day 1

Launch 3

Day 60

Launch 2

Day 30

Launch 4

Day 90

≈26 mo Later

Launch 5

Day 1

Launch 6

Day 30

Launch 7

Day 60

Crew Launch

Vehicle Required

Mission Profile (Reference DRA5.0 and AIAA 7-Launch Architecture)

Payload: Dual-Use Shroud and Encapsulated Payload (DAV)

Mass: 102.8 t

Payload: Dual-Use Shroud and Encapsulated Payload (SHAB)

Mass: 102.8 t

Figure 11. NTP launch sequence and payload mass requirements.
21

Delivering twice the Isp of its chemical propulsion counterparts, nuclear propulsion systems significantly reduce the mass injected to LEO requirements of the Ares V launch system. This 

increase in operating efficiency of the system and the resulting reduction in mass required allows 

for the same payload to be injected to Mars with a reduced number of Ares V launches because a 

larger percentage of the in-space mass is payload rather than the propellant required to send that 

payload to its destination. The general requirements of the Ares V launch system were divided into 

two main categories for this MTV option: (1) Minimum Ares V participation and (2) maximum 

Ares V participation.

While either the minimum participation or maximum participation would have a significantly different implication for the requirements on the Ares V launch vehicle, either option would 

fit within the mission architecture depicted in figure 11.

4.1.1 Minimum Ares V Participation

The first scenario for delivering the required payloads and propulsion elements for the NTP 

option to LEO involves utilizing the Ares V EDS in much the same way as the Ares I upper stage 

is utilized for the lunar DRM. The EDS would only be required to deliver the large payloads to 

an elliptical orbit (exact orbit to-be-determined) and require the payload and/or auxiliary systems 

added to the payload to provide all of the on-orbit functionality required to deliver the payload 

to the ultimate loiter/assembly orbit (circularization maneuvers, rendezvous/proximity operations 

maneuvers, etc.), maintain the payload for the entire loiter/assembly duration (power, ACS, etc.), 

perform any functionality required prior to the TMI maneuver (micrometeoroid and orbital debris 

(MMOD) avoidance, reboost, control moment gyroscope (CMG) desaturation if applicable, etc.), 

and ultimately perform the TMI maneuver and the ensuing NTP stage disposal maneuver. In this 

scenario, the Ares V EDS is assumed to function much like a traditional payload-to-LEO delivery 

service. In addition, it would be an option for the EDS to perform some, if not all, of the circularization maneuver and its own jettison and disposal maneuvers, but that assumption was not 

included in the minimum participation scenario.

A representative LEO delivery sequence is shown in figure 12. In this notional assembly 

sequence, the Ares V EDS (or second stage) is shown delivering the NTP element of the cargo 

MTV to an elliptical orbit on the first launch followed by the dual-use shroud with encapsulated 

payload on the second launch. In this depiction, the orbit is assumed to be a stable phasing elliptical orbit so the payload will have several orbits for potential contingency operations should they 

be needed. The EDS is jettisoned at the earliest possible convenience, and it performs a disposal 

maneuver for reentry, eventually breaking up and splashing down over an open body of water. 

The expected EDS lifetime for both cases is ≤5 days.

Where the cargo MTVs consist of two Ares V launches each, the crewed MTV requires 

three Ares V launches followed by an Ares I (CLV) to deliver the crew to the assembled crewed 

MTV. A notional assembly sequence for this mission is depicted in figure 13. The NTP element 

is delivered first, then the truss structure/LH2 “drop tank” and about 94 t of LH2, and then the 

TransHab with contingency consumables, associated canister, and docking mechanism.5 Again, 

the EDSs that are responsible for delivering the crewed MTV components into a stable phasing 

elliptical orbit have an expected lifetime of ≤5 days. 
22

Circularization

Maneuver

Circularization Maneuvers

Circular at 410 (220) Docking

Wait for

Departure

Window

Notional

Minimum Ares V participation

Launches for NTP cargo MTV

Altitudes approximate: km (nmi)

MTV

Departure

(TMI) Rendezvous

Maneuvers

Loiter for

Disposal

Opportunity

ECO (EOI)

225×375

(120×200)

ECO (EOI)

225×410

(120×220) Loiter for

Disposal

Opportunity

Disposal Orbit:

410×56

(220×30)

Disposal Orbit:

375×56

(200×30)

Core

Stage

First

Stage

1 2 Core

Stage

First

Stage

Shroud

Contingency operations

for safing nuclear

materials not shown

EDS lifetime ≈5 days

Disposal

Maneuver Payload

Separation

Each MTV component is launched

with GN&C and OMS capability

MTV performs reboost and

orbit debris avoidance as needed

Nuclear MTV stages have

redundant/assured safe

disposal capability

Figure 12. Ares V payload to LEO delivery sequence (minimum participation, NTP cargo).

Circularization

Maneuver

Circularization Maneuvers

Circular at 410 (220) Docking Docking

Wait for

Departure

Window

Notional

Minimum Ares V participation

Launches for NTP cargo MTV

Altitudes approximate: km (nmi)

MTV

Departure

(TMI) Rendezvous

Maneuvers

Rendezvous

Maneuvers

Loiter For

Disposal

Opportunity

ECO (EOI)

225×375

(120×200)

ECO (EOI)

225×375

(120×200)

ECO (EOI)

225×410

(120×220) Loiter For

Disposal

Opportunity Loiter For

Disposal

Opportunity

Disposal Orbit:

410×56

(220×30)

Disposal Orbit:

410×56

(220×30)

Disposal Orbit:

375×56

(200×30)

Core

Stage First

Stage

1 3 2

Core

Stage

Core

Stage

First

Stage

Upper

Stage

First

Stage

Shroud Shroud Shroud

Contingency operations for safing

EDS lifetime ≈5 days nuclear materials not shown

Disposal

Maneuver

Payload

Separation Correct Phase

Angle Before Final

Circularization

Correct Phase

Angle Before Final

Circularization

Orion Phasing

And Rendezvous

Maneuvers

Not Shown

Each MTV component is launched

with GN&C and OMS capability

MTV performs reboost and

orbit debris avoidance as needed

Nuclear MTV stages have redundant/

assured safe disposal capability

Figure 13. Ares V payload to LEO delivery sequence (minimum participation, NTP crew).
23

There would be additional opportunities to further optimize the delivery orbit for this 

scenario; however, the stable phasing elliptical chosen for this case was selected because of 

the following reasons:

• The MTV is a high-value National asset. Inserting into an orbit with a 45-minute lifetime 

(negative perigee for ensured EDS reentry on first pass) allows little time to debug and repair 

a failure (separation problems, failure of the MTV’s orbital maneuvering system (OMS) system, 

failure of the MTV’s avionics, etc.). Inserting into a stable orbit allows time to repair many of 

the possible failures.

• Inserting into a 130-×220-nmi orbit versus a 30-×220-nmi orbit saves significant dV on the 

behalf of the auxiliary propulsion systems of each of the MTV components.

• EDS is a fully capable stage. No new capability is required above those required for the lunar 

DRMs and deorbit from LEO is a lunar requirement for contingencies. 

It is also fairly evident from looking at these depictions that it may be in the best interest 

of the total system architecture for the EDS or some other auxiliary propulsion stage to perform 

some, if not all, of the circularization maneuver; rendezvous, proximity operations, and docking 

(RPOD) maneuvers; and perhaps some on-orbit functionality in order to decrease the requirements 

levied on the payload elements (and decrease mass at TMI). This thought process led to the development of the maximum Ares V participation approach.

4.1.2 Maximum Ares V Participation

The other scenario explored for delivering NTP components to LEO was termed the maximum Ares V participation option. Under this scenario, the Ares V EDS and added auxiliary 

propulsion systems would be responsible for delivering the payloads to their final loiter/assembly 

orbit. In addition, the EDS was assessed for its ability to add some on-orbit functionality that only 

applied during the LEO assembly operation phase of the mission architecture and was not needed 

after the TMI maneuver. These types of functions (and some contingency capabilities) seemed best 

placed on the EDS so they could be provided by that stage, but the mass associated with providing 

those functions could be jettisoned along with the EDS before the TMI maneuver was performed—

increasing overall mission efficiency. These functions include final orbit insertion/circularization, 

RPOD functionality, ACS/RCS during rendezvous if applicable, MMOD avoidance maneuvers if 

applicable, CMG desaturation, and its own jettison/disposal functionality. 

A general assembly sequence for the cargo case of this option is depicted in figure 14. In this 

case, the first Ares V is responsible for delivering the NTP TMI propulsion stage for the cargo mission to LEO. Two of these are delivered for the two cargo MTVs (as shown in fig. 11). The second 

Ares V (for this cargo MTV) is responsible for delivering the dual-use shroud with encapsulated 

payload to the final assembly orbit. Once assembled, the TMI maneuver is performed to inject 

the payload on a transfer to Mars. 

It can be seen in this notional case, that the EDS is assumed to inject the NTP stage into a 

stable phasing elliptical orbit and then auxiliary propulsion systems added to the EDS perform the 

circularization maneuver at the desired altitude (currently assumed to be 20 nmi less than the final 
24

Circularization

Maneuver

Circularization

Maneuvers

Circular at 410 (220) Undocking

Undocking

Docking

Notional

Minimum Ares V participation

Launches for NTP cargo MTV

Altitudes approximate: km (nmi)

MTV

Departure

(TMI)

Rendezvous

Maneuvers

ECO (EOI)

225×375

(120×200) ECO (EOI)

225×410

(120×220) Before docking, EDS is controlling

attitude of NTP stage via RCS

EDS OMS for debris avoidance

Up to 30 Revs

to Correct Phase Angle

FS, CS, and Shroud

Disposal not Shown

Disposal Orbit:

410×56

(220×30)

Disposal Orbit:

410×56

(220×30)

Core

Stage First

Stage

#1

#1

#3

#2

#2

Contingency operations for safing

nuclear materials not shown

EDS lifetime ≈5 days

EDS lifetime up to 150 days

Default attitude is GG.

Disposal

Desposal Maneuver

Maneuver

Arrays

Deployed

After docking, MTV is controlling

attitude via CMG EDS OMS for debris

avoidance EDS OMS for reboost

All postinsertion maneuvers

using EDS OMS

Figure 14. Ares V payload to LEO delivery sequence (maximum participation, NTP cargo).

assembly altitude). Furthermore, the second Ares V launch delivers the dual-use shroud and encapsulated payload to a stable phasing elliptical orbit. This payload and EDS are maintained in the 

orbit until the appropriate phase error reduction is achieved between it and the NTP stage delivered 

on the first launch. When that reduction in phase angle is realized, the auxiliary propulsion system 

performs the circularization maneuver at the desired loiter/assembly altitude (currently assumed to 

be a 220-nmi circular orbit), at which time the EDS is discarded. 

The auxiliary propulsion systems on EDS1 (delivered with the NTP stage) are responsible for performing the gradual bumps in altitude required to rendezvous with the payload at the 

final assembly altitude. Finally, after docking is achieved between the NTP stage and the dual-use 

shroud/payload, the EDS is responsible for maintaining the loiter altitude (pre-TMI), any MMOD 

avoidance maneuvers that may be needed, and perhaps CMG desaturation if used to maintain attitude on orbit. Overall, these types of LEO functions are best placed on a system that has the ability 

to be jettisoned and disposed of before the TMI maneuver is performed (the EDS in this case). In 

that manner, the mass required to perform those functions is discarded to increase overall mission 

efficiency.

The crewed MTV assembly sequence is shown in figures 15–17. The same types of functions 

are also provided for that particular assembly sequence. 

4.2 Chemical Propulsion Options

While benefits in mass requirements in LEO because of increased operational efficiency 

have been shown for the NTP option, some chemical propulsion options also offer advantages. 
25

Circularization

Maneuvers

Circular at 410 (220)

Loiter for

Next Arrival

Notional

Maximum Ares V participation

1st Launch for NTP crew MTV

Altitudes approximate: km (nmi)

ECO (EOI)

225×410

(120×220)

Before docking with habitat, EDS 

is controlling attitude of NTP stage 

via RCS EDS OMS for debris 

avoidance EDS OMS for reboots

Core

Stage

Shroud

First

Stage

#1

Contingency 

operations for safing

nuclear materials 

not shown

This EDS stage remains until 

departure for Mars; lifetime 180 days

Default attitude is GG.

All postinsertion maneuvers

using EDS OMS

Figure 15. Ares V payload to LEO launch No. 1 (maximum participation, NTP crew).

Circularization

Maneuver Circular at 410 (220) Docking

Default attitude is GG.

Undocking

Disposal

Maneuver

Notional

Maximum Ares V participation

2nd Launch for NTP crew MTV

Altitudes approximate: km (nmi)

Loiter for

Next Arrival

Rendezvous

Maneuvers

ECO (EOI)

225×375

(120×200)

Disposal Orbit:

410×56

(220×30)

Core

Stage

First

Stage

#2

#2

#2

#1

Shroud

Contingency operations

for safing nuclear

materials not shown

EDS lifetime up to 30 days

Proximity

Operations

Proximity

Operations

Up to 30 Revs

to Correct

Phase Angle MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost

All postinsertion maneuvers

using EDS OMS

Figure 16. Ares V payload to LEO launch No. 2 (maximum participation, NTP crew).

The two that were studied in detail during PA-C3′ and carried forward for further analysis both 

involve cryogenic propellant transfer in LEO but provide maximum utilization of a lunar-EDS 

derived, in-space transportation system architecture. However, another option was also assessed 
26

Circularization

Maneuver

Circular at 410 (220)

Docking

Default attitude is GG.

Orion

arrival and

docking

Disposal

Maneuver Disposal

Maneuver

Notional

Maximum Ares V participation

3rd Launch for NTP crew MTV

Altitudes approximate: km (nmi)

MTV

departure

(TMI)

Rendezvous

Maneuvers

ECO (EOI)

225×410

(120×220)

Disposal orbit:

410×56

(220×30)

Disposal orbit:

410×56

(220×30)

Core

Stage First

Stage

#3

#3

≈30 Days

#1

#1

#1

#3

#3

Shroud

Contingency operations

for safing nuclear

materials not shown

EDS lifetime up to 10 days EDS lifetime up to 180 days

Proximity

Operations

Proximity

Operations

Up to 30 revs

to Correct

Phase Angle

MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost

All postinsertion maneuvers

using EDS OMS

Figure 17. Ares V payload to LEO launch No. 3 (maximum participation, NTP crew).

for comparison purposes, which was presented in much more detail in the DRA 5.0 final report. 

This was the chemical propulsion module option discussed below.

4.2.1 Chemical Propulsion Modules Assessed for Mars Design Reference Architecture 5.0

In developing the cargo and crewed MTVs, separate propulsive elements were identified 

for each major mission maneuver: TMI, MOI, and TEI. All TMI modules are meant to retain 

commonality among the cargo and crewed MTVs, and each module utilizes five RL10-B2 engines 

and is jettisoned after performing its burn. The MOI modules use two RL10-B2 engines for the 

all-propulsive crew and cargo missions. TEI modules similarly use two RL10-B2 engines and also 

perform plane changes while in Mars orbit. 

The two cargo MTVs require two TMI modules each (20 total RL-10s), as specified in DRA 

5.0. The crewed MTV employs multiple propulsive stages, namely three TMI modules (15 RL-10s), 

one TEI module (2 RL-10s), and one MOI module (2 RL-10s). With three TMI modules required 

for each crewed MTV, a two-burn TMI maneuver is necessary; therefore, the two outer stages 

perform the first portion of the TMI burn (and then are jettisoned), followed by the TMI module 

located in the center providing the remaining dV necessary to place the crewed MTV elements into 

a trans-Mars coast. 

This approach and the associated number of propulsion modules that were assumed to 

be delivered to LEO requires an 11-launch architecture with an additional crew launch vehicle 

to deliver the crew into LEO once the crewed MTV is assembled and the TMI window is open. 
27

While this option is detailed more fully in the Mars DRA 5.0 final report and addendum, analysis 

by the Ares V team of this mission architecture led to alternative in-space transportation system 

approaches (while still using traditional chemical propulsion system approaches) that were found 

to reduce the number of launches, total mass to orbit requirements, and overall cost and complexity of developing dedicated in-space transportation systems. The two most competitive options 

were deemed the propellant tanker option (also called the scavenger option) and the derivative 

commercial tanker option.

For comparative purposes, the launch and assembly sequence for an approximate DRA 5.0 

chemical propulsion module case is depicted in figure 18.

Cargo SHAB MTV (IMLEO 325 t)

Cargo DAV MTV (IMLEO 325 t)

Crewed MTV (IMLEO 540 t)

Payload: TEI and MOI; 125 t

Payload: TMI Module; 121 t

Payload: TMI Module; 121 t

Payload: TMI Module; 121 t

Payload: TransHab; 53 t

• Total Ares V Launches: 11 (6 for two Cargo MTVs and 5 for Crewed MTV)

• Launch Rate: 6–8 per yr per mission

• 2 TMI modules required for each Cargo vehicle; 3 TMI modules required 

 for each Crew vehicle (2 burn TMI, 2 outer, then center)

• Conjunction long duration stay: 180 day transits with ≈490–570 days at Mars

• Earth orbit: 407-km circular; Mars orbit: 250 km x 33,793 km

• Assuming Trade tree branch 3: AC (Cargo MOI) and PC (Crew MOI); 

 Predeploy; partial-ISRU; Nuclear surface power

• Cargo vehicles support the crew mission of the next mission opportunity 

 (during the same launch year)

Payload: TMI Module

Mass: 108 t

Payload: TMI Module

Mass: 108 t

Payload: TMI Module

Mass: 108 t

Payload: Surface Habitat

Mass: 108 t

Launch 1

Day 1

Launch 2

Day 60

Launch 5

Day 120

Launch 3

Day 30

Launch 4

Day 90

Launch 6

Day 150

≈26 mo Later

Launch 8

Day 30

Launch 7

Day 1

Launch 9

Day 60

Launch 11

Day 120

Launch 10

Day 90

Crew Launch

Vehicle Required

Mission Profile (Reference DRA5.0 and AIAA 7-Launch Architecture)

Payload: Surface Lander

Mass: 108 t

Payload: TMI Module

Mass: 108 t

Figure 18. DRA 5.0 chemical propulsion module option.

4.2.2 Earth Departure Stage as Mars Transfer Vehicle Trans-Mars Insertion Stage: Propellant 

Tanker With In-Space Propellant Transfer 

While the chemical propulsion module option found that a dedicated in-space RL-10-

derived stage may better serve agency needs for a chemical propulsion option for sustained human 
28

exploration of Mars, a less costly, less complex, and potentially more reliable approach may involve 

also utilizing the same assets for the Mars exploration campaign that are to be built and operated 

for a decade or more for the lunar campaign. The main advantage for a chemical propulsion system with an EDS heritage would be a clear evolution path for the lunar EDS when applied to the 

Mars DRM and the shared mission commonality which that would entail (in the case of concurrent lunar and Mars missions for instance or for increased reliability confidence). This extensibility 

would result in a much greater return-on-investment in the lunar exploration elements that would 

otherwise be underutilized for the Mars campaign. In addition, the cost of the ETO and in-space 

transportation elements would be reduced to the needed technological investments above-andbeyond those planned for development during the lunar missions. This EDS-derived propellant 

tanker option was given the informal moniker scavenger option due to the fact that all available 

performance up to the CARD requirement of 125 t was used for either payload, propellant, docking and propellant transfer hardware, or a combination thereof. The propellant tanker option 

consists of 9 Ares V launches with an option to grow to 10 or 11 to add greater margin to the 

architecture, if required. This is reasonably competitive with the seven-launch NTP architecture 

as proposed in DRA 5.0 in terms of total Ares V launches required to carry out the proposed mission. It is currently assumed that both will require an additional CLV to deliver the crew once the 

crewed MTV is assembled in LEO. 

As shown in figure 19, the first five launches deliver components for the two cargo MTVs. 

In this scenario, the first and second Ares V launches deliver the dual-use shrouds with encapsulated payloads (102.8 t each) for the cargo MTVs, along with the EDS that will perform the TMI 

maneuver for each, which is assumed to have a small amount of propellant remaining in the tanks 

(currently assumed to be ≈22 t of LOX/LH2 propellants) once inserted into the LEO assembly 

orbit. The third and fourth launches deliver ≈120 t of propellant (in other EDS modules) that is 

transferred to the 22 t of residual propellant from the first EDS TMI stage (for each cargo MTV). 

The third and fourth launch EDSs are then jettisoned and disposed of. These launches result in 

about 142 t of total propellant loaded in the EDS TMI stage. A total of 192 t of propellant is 

needed to meet the dV requirement for the cargo mission (with current assumptions for stage mass, 

payload mass, dV, etc.), so a fifth and final launch acts as a split-tanker stage EDS that divides 100 t 

between the two cargo MTVs, (50 t allocated to each). Obviously, this shared launch has more performance capacity so it would also be able to deliver on the order of 10 t of additional propellant 

to serve as margin or decrease sensitivity to boiloff assumptions made for this assessment. Furthermore, instead of sharing a launch between the two cargo MTVs, a dedicated second propellant 

delivery launch could be planned to provide much more propellant to each MTV to either decrease 

the need for zero boiloff (ZBO) and/or provide margin for the propellant transfer functionality. 

Once this EDS meets with and tanks 50 t to each cargo MTV, it is jettisoned and disposed of. The 

net result of these five Ares V launches is two cargo MTVs assembled in LEO, each comprised of 

a 102.8-t, dual-use, shroud and encapsulated payload and a single EDS loaded with 192 t of LOX/

LH2 propellant that will perform the TMI maneuver.

Also shown in figure 19, the crewed MTV requires four Ares V launches (with current 

assumptions). The first launch delivers the TEI and MOI stages (assumed at 125 t from previous 

analysis) and has fully exhausted all propellant in the EDS en route to LEO. The second launch 

delivers a TransHab, small truss structure and contingency consumables to LEO, in total ≈50 t. 
29

Cargo SHAB MTV (IMLEO 325 t)

Cargo DAV MTV (IMLEO 325 t)

Assuming ≈77% Filled EDS(s)

DRA5.0 TMI Prop. Module

Assuming ≈48% and ≈77% 

Filled EDS(s)

Crewed MTV (IMLEO 555 t With 4 Launches, 566 t total)

Payload: TEI and MOI Module

Mass: 125 t

Payload: Transhab Module + TMI Stage 2 + Prop

Mass: 155 t (≈49.4 t + 30 t + ≈75 t (Prop))

Payload: Propellant (Tanks TMI Stage and Deorbits)

Mass: 120 t TMI Propellant (to TMI Stage 2)

Payload: EDS (TMI Stage 1 + Adapter)

Mass: 155 t (120 t (Prop)) + 35 t (Stage))

• 9 Total Ares V Launches: 5 for two Cargo MTVs and 4 for Crewed MTV

• Launches 3 and 4 carry propellant to TMI stages on Cargo MTVs. Launch 5 

 carries 100-t Tanker Stage to half tank TMI stages with 50 t each (up to 60 t)

• Cargo vehicles support the crew mission of the next mission opportunity 

 (during the same launch year)

• Earth orbit: 220-nmi (407-km) circular; Mars orbit: 250 km x 33,793 km

Payload: Propellant (Tanks TMI Stage and Deorbits)

Mass: 120 t TMI Propellant

Payload: Tanker Stage (Delivers Fuel to Both Cargo MTVs)

Mass: 100 t (50 t to Each Cargo MTV (TMI Stage), Then Deorbits)

Payload: Surface Habitat + TMI Stage (≈22-t Propellant

Mass: 155 t (102.8 t + 52 t)

Launch 1

Day 1Launch 3

Day 60

Launch 5

Day 120

Launch 2

Day 30

Launch 4

Day 90 Launch 6

Day 1

≈26 mo Later

Launch 8

Day 60

Launch 7

Day 30

Launch 9

Day 90

Crew Launch

Vehicle Required

Mission Profile (Reference DRA5.0 and AIAA 7-Launch Architecture)

Payload: DAV + TMI Stage (≈22-t Propellant)

Mass: 155 t (102.8 t + 52 t)

Payload: Propellant (Tanks TMI Stage and Deorbits)

Mass: 120 t TMI Propellant

Figure 19. Propellant delivery and transfer option (EDS-as-MTV).

This leaves an ≈75 t total performance capability that would otherwise be unutilized. Therefore, this 

second launch is assumed to also deliver an EDS that has a residual propellant load of 75 t, which 

serves as the second stage of the two-burn TMI maneuver for the crewed MTV. The third launch 

is a dedicated propellant launch, which transfers 120 t of propellant to the previously 75 t filled 

EDS. This results in about 195 t of LOX/LH2 propellants in that stage. Once the 120 t of propellant is transferred, the empty EDS is jettisoned and disposed of. The fourth launch brings up the 

first stage of the two-burn TMI maneuver and appropriate docking hardware to connect to the 

other EDS. It is currently assumed that the docking hardware will be on the order of 5 t, which 

leaves ≈120 t of capability for delivering propellant on this launch. That gives a total of 120 t in 

the first TMI stage EDS and 195 t in the second TMI stage EDS resulting in 315 t of total LOX/

LH2 propellants (still slightly short of the dV requirement but close for this fidelity analysis). 

Another option for the crewed MTV is a dedicated propellant tanker for the first TMI stage. 

It is currently assumed that it will be delivered with 120 t of propellant, which is less than 50% of 

the EDS tank capacity. Another launch would provide up to 120 t of additional propellant to this 

EDS in order to increase the dV capability of the assembled stack, provide margin for the transfer 

of propellant in-orbit, and/or decrease the need for a ZBO thermal control system. 
30

Figures 20–23 depict additional details on this launch and assembly sequence for both cargo 

MTV cases. Each cargo MTV would be assembled over a period of months (current assumption 

is 30 days between Ares V launches). The first component that would be delivered is the dual-use 

shroud with encapsulated payload, as shown in figure 20. This payload is injected into the final 

assembly/loiter orbit, and the associated EDS for this launch will serve as the platform for receiving 

the propellants that will be delivered on later Ares V launches. 

Circularization

Maneuver Circular at 410 (220)

Default attitude is GG.

Arrays

Deployed

MTV

Activated

Notional

Maximum Ares V participation

1st Launch for NTP cargo MTV

Altitudes approximate: km (nmi)

ECO (EOI)

225×410

(120×220)

Core

Stage

First

Stage

EDS lifetime 150 days

MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost (if needed)

All postinsertion maneuvers

using EDS OMS

Payload Mass 103 t

Propellant residual: 22 t

Figure 20. Dual-use shroud/encapsulated payload launch (chemical cargo).

Figure 21 shows the first dedicated propellant tanker launch. The assumption depicted for 

delivering the maximum propellant to orbit is a tank set located above the EDS that also functions 

as the outer-mold line (OML) of the vehicle (similar to the Space Transportation System (STS) 

external tank). This is done in order to decouple the performance capability of the vehicle from the 

actual geometric constraints of the EDS tanks. There are other methods for delivering propellant 

to LEO including loading the EDS tanks to 100% capacity and burning the optimum amount of 

propellant to deliver the maximum amount of EDS residual propellant to orbit (no payload forward of EDS, only a nose cone) or using an encapsulated tank set (traditional shroud as the OML) 

that may be the optimum solution for constructing and delivering propellant transfer hardware 

to orbit. Most of these details require much more in-depth analysis. Figure 22 shows the shared 

tanker launch.

The key feature of this shared propellant launch (shown in figures 22 and 23) is its ability to rendezvous and dock with both cargo MTVs on-orbit. Therefore, it is assumed that it will 

be launched into a phasing elliptical orbit, rendezvous/dock/transfer propellant to the first cargo 

MTV, then jettison and transfer to a lower orbit in order to phase appropriately to the second 
31

Circularization

Maneuver

Circular at 410 (220)

Docking

Default attitude is GG.

Loiter for

Departure

Disposal

Maneuver

Notional

Maximum Ares V participation

2nd Launch for cargo MTV

Altitudes approximate: km (nmi)

Propellant transferred: 120 t

Total propellant: 142 t

Propellant trasfer during coast uses 

RCS for settling ullage and He pressure

to force liquid transfer.

After Docking,

Stage-to-Stage

Propellant Transfer

Transfer

Remaining

Propellants

Activate

Cryogenic

Cooling

Rendezvous

Maneuvers

ECO (EOI)

225×375

(120×200)

Disposal Orbit:

410×56

(220×30)

Core

Stage First

Stage

Fuel

Transfer

During Burn

Tanks jettisoned at 2nd apogee

by retrorockets, resulting in

375×110 (200×60) reentry orbit

Tanks Jettisoned

≈30 Days

EDS lifetime 10 days

Proximity

Operations

Up to 30 Revs

to Correct

Phase Angle

MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost (when needed)

All postinsertion maneuvers

using EDS OMS

Figure 21. Propellant launch for cargo MTV (first chemical cargo).

Circularization

Maneuver

Circular at 410 (220)

Docking

Default attitude is GG.

Loiter for

Departure

MTV Departure

(TMI)

After transferring 1/2 of propellant to 1st cargo MTV vehicle,

the EDS tanker maneuvers to a phasing orbit at 300x375 (160x200)

Notional

Ares V is MTV propulsion stage

Launch of shared tanker

Transfer to 1st cargo MTV

Altitudes approximate: km (nmi) Propellant residual: 100 t

Propellant transferred: 50 t

Total propellant: 192 t

Propellant headroom: 25 t

Propellant transfer during coast uses 

RCS for settling ullage and He pressure

to force liquid transfer.

After Docking,

Stage-to-Stage

Propellant Transfer

Transfer

Remaining

Propellants

Activate

Cryogenic

Cooling

Rendezvous

Maneuvers

ECO (EOI)

225×375

(120×200)

Core

Stage

First

Stage

Fuel

Transfer

During Burn

Tanks jettisoned at 2nd apogee

by retrorockets, resulting in

375×110 (200×60) reentry orbit

Tanks Jettisoned

≈30 Days

Proximity

Operations

Up to 30 Revs

to Correct

Phase Angle

MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost (when needed)

All postinsertion maneuvers

using EDS OMS

Figure 22. Shared propellant launch for first MTV (second chemical cargo).
32

Circularization

Maneuver

Circular at 410 (220)

Docking

Default attitude is GG.

MTV Departure

(TMI)

Disposal

Maneuver

Notional

Ares V is MTV propulsion stage

Rendezvous of shared tanker

Transfer to 2nd cargo MTV

Altitudes approximate: km (nmi)

Loiter for

Departure

Rendezvous

Maneuvers

Disposal Orbit:

410×56

(220×30)

Propellant transferred: 50 t

Total propellant: 192 t

EDS lifetime ≈20 days

Proximity

Operations

After Docking,

Stage-to-Stage

Propellant

Transfer

After transferring 1/2 of propellant to 1st cargo MTV vehicle,

the EDS tanker phases to 2nd MTV cargo vehicle, then rendezvous

Phasing orbit: 300×375 (160×200).

≈30 Days

MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost (when needed)

All postinsertion maneuvers

using EDS OMS

Figure 23. Shared propellant launch for second MTV (two and one-half chemical cargo).

cargo MTV. In this manner, the auxiliary propulsion system will have to be sized appropriately to 

perform these additional in-space maneuvers, including any changes in the other orbital elements 

of the vehicle.

Figures 24–27 show the assembly sequence currently assumed for the crewed MTV. The first 

launch (depicted in fig. 24) is the MOI/TEI chemical propulsion stages, which are delivered to the 

final assembly/loiter orbit. Subsequence launches are delivered to this orbit.

The second launch of the crewed MTV assembly sequence is depicted in figure 25. This 

launch delivers the crew TransHab (a small truss structure that connects to a contingency consumables canister and provides the aft interface for docking with the MOI/TEI modules delivered on 

the first launch) and the consumables located in the canister that provides provisions for the mission duration in the case of the crew not being able to descend to the surface of Mars. The summation of all of these components is less than 50 t; therefore, the EDS is loaded with an additional 

≈75 t of LOX/LH2 propellant. Once the payload is successfully docked with the MOI/TEI modules, 

the EDS is undocked and redocks at the aft end of the vehicle to serve as the second stage for the 

TMI maneuver.

The third launch (fig. 26) depicts the delivery of a dedicated tanker launch for the second 

TMI stage. Figure 27 shows the delivery of the first TMI stage and the CLV delivery of the crew 

to the completed crewed MTV stack. 
33

Circularization

Maneuver Circular at 410 (220)

≈30 Days Unattended

Default attitude is GG.

Arrays

Deployed

MTV

Activated Disposal

ManeuverDisposal Orbit: 410×56 (220×30)

Notional

Maximum Ares V participation

1st Launch for EDS crew MTV

ECO (EOI) Altitudes approximate: km (nmi)

225×410

(120×220)

Core

Stage

Shroud

First

Stage EDS lifetime 2–4 days

MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost (if needed)

All postinsertion maneuvers

using EDS OMS

Payload Mass 125 t

Propellant headroom: 20 t

(not utilized)

Figure 24. MOI/TEI module launch, first launch (first chemical crewed).

Circularization

Maneuver

Circular at 410 (220)

Docking

Default attitude is GG.

Loiter for

Next Arrival

EDS Docks Transhab at Front,

Then Docks at Rear

Notional

Ares V is MTV propulsion

2nd Launch for EDS crew MTV

Altitudes approximate: km (nmi)

Propellant residual: 75 t

Activate

Cryogenic

Cooling

Rendezvous

Maneuvers

ECO (EOI)

225×375

(120×200)

Core

Stage

Shroud

First

Stage

≈30 Days Unattended

Proximity

Operations

Up to 30 Revs

to Correct

Phase Angle

MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost (when needed)

All postinsertion maneuvers

using EDS OMS

Figure 25. TransHab/truss/consumables/‘second TMI stage,’ second launch 

(second chemical crewed).
34

Circularization

Maneuver

Circular at 410 (220)

Docking

Default attitude is GG.

Loiter for

Next Arrival

Disposal

Maneuver

Notional

Ares V MTV propulsion stage

3rd Launch for crew MTV

Altitudes approximate: km (nmi)

Propellant transferred: 120 t

Total propellant: 195 t

Propellant transfer during coast uses 

RCS for settling ullage and He pressure

to force liquid transfer.

After Docking,

Stage-to-Stage

Propellant Transfer

Transfer

Remaining

Propellants

Activate

Cryogenic

Cooling

Rendezvous

Maneuvers

ECO (EOI)

225×375

(120×200)

Disposal Orbit:

410×56

(220×30)

Core

Stage

First

Stage

Fuel

Transfer

During Burn

Tanks jettisoned at 2nd apogee

by retrorockets, resulting in

375×110 (200×60) reentry orbit

Tanks Jettisoned

≈30 Days

EDS lifetime 10 days

Proximity

Operations

Up to 30 Revs

to Correct

Phase Angle

MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost (when needed)

All postinsertion maneuvers

using EDS OMS

Figure 26. Propellant launch for ‘second TMI stage’ (third chemical crewed).

Circularization

Maneuver

Circular at 410 (220)

MTV Departure

(TMI)

Docking

Default attitude is GG.

Loiter for

Orion Arrival

Notional

Ares V is MTV propulsion stage

4th Launch for crew MTV

Altitudes approximate: km (nmi)

Propellant transferred: 120 t

Total propellant: 315 t

Propellant transfer during coast uses 

RCS for settling ullage and He pressure

to force liquid transfer.

Transfer

Remaining

Propellants

Activate

Cryogenic

Cooling

Rendezvous

Maneuvers

ECO (EOI)

225×375

(120×200)

Core

Stage

First

Stage

Fuel

Transfer

During Burn

Tanks jettisoned at 2nd apogee

by retrorockets, resulting in

375×110 (200×60) reentry orbit

Tanks Jettisoned

≈30 Days

Orion Phasing,

and Redenzvous

Manuevers

not Shown

Proximity

Operations

Up to 30 Revs

to Correct

Phase Angle

MTV controlling attitude via CMG

EDS OMS for debris avoidance

EDS OMS for reboost (when needed)

All postinsertion maneuvers

using EDS OMS

Figure 27. ‘First TMI Stage’ launch (fourth chemical crewed).
35

4.2.3 Earth Departure Stage as the Trans-Mars Insertion Stage: Commercial Tanker

This option is comparable to the previously mentioned propellant tanker option, but it 

reduces the number of required Ares V launches by utilizing commercial tankers. As the Ares 

V propellant tanker option evolved, it was noticed that most of the required launches are dedicated propellant delivery launches for the TMI stages (five of the nine launches). As a potentially 

enhancing capability, it became evident that if a commercial market for LOX/LH2 propellants 

indeed exists 20–25 yr in the future, the dedicated Ares V launches to deliver this propellant could 

be replaced by this commercial market. Therefore, the absolute minimum number of required 

launches for Ares V was found, which was four. This consists of the very large masses and the platform for the commercial tanker services to deliver the propellants to (i.e., empty or nearly empty 

EDSs). In addition, there will be numerous, but not yet determined, commercial launches necessary 

to provide the additional propellant to the TMI stages for the MTVs. 

Two of these four Ares V launches will deliver the dual-use shroud and encapsulated payload along with an empty (or nearly empty) EDS. These two launches will serve as the two cargo 

MTVs. The later two launches would deliver the TransHab module (associated elements, small 

truss, contingency consumables, etc.), the MOI/TEI modules, and two empty (or nearly empty) 

EDS elements that will serve as the crewed MTV TMI stages. Upon R&D with the MTV, these 

EDS elements would await more propellant from the commercial tankers. 

While this unfetters NASA from the responsibility of delivering all required mass to LEO, 

there are also numerous other considerations such as the interactions between commercial and 

government vehicles, ability to accept multiple propellant loads with a single EDS, reliability of the 

significant increase in required launches per mission, and a host of others. If only NASA-to-NASA 

interfaces are desired, commercial industry must position and maintain a depot in space where 

fuel may be extracted, such that a NASA launch vehicle interfaces with a NASA-assembled MTV. 

Further, one might consider the cost associated with the number of commercial launches to deliver 

up to 200 t of propellant to each cargo MTV and up to 250 t to the crewed MTV, as opposed to the 

marginal cost of two or three additional Ares V flights within a year. Nevertheless, it would allow 

for the establishment of a relatively stable market for up to 650 t of LOX/LH2 cargo to LEO on a 

2-yr repeating cycle and engage the commercial (and potentially international) industry in the overall Mars mission strategy—a potentially attractive development that may help ensure the longevity 

of the Mars campaign. Furthermore, the capability would still exist to deliver the propellants with 

the Ares V tanker if the commercial market does not present itself or if there are unforeseen insurmountable challenges associated with this propellant delivery method. Figure 28 further describes 

this option.

The general assembly sequences are the same as the EDS-as-MTV option, but it can be seen 

that several commercial tankers will replace the dedicated propellant launches. These assembly 

sequences have not been created.
36

Cargo SHAB MTV (IMLEO 325 t)

Cargo DAV MTV (IMLEO 325 t)

Assuming ≈77% Filled EDS(s)

Assuming ≈48% and 

≈77% Filled EDS(s)

Crewed MTV (IMLEO 566 t

Including Orion)

Payload: Provided by Commercial Tanker

Mass: up to 145+ t (50+ t + 195+ t)

• 4 Total Ares V Launches: 2 for the two Cargo MTVs and 2 for Crewed MTV)

• Commercial industry tasked with delivering payload; Not a responsibility 

 carried by this mission launch architecture

• Planning can include developing the backup capability to deliver needed 

 propellant to carry out the mission

• Cargo vehicles support the crew mission of the next mission opportunity 

 (durring the same launch year)

Payload: TransHab + EDS (TMI Stage)

Mass: 155 t (≈50 t + (70 t + 35 t))

Payload: TEI and MOI + Empty TMI Stage

Mass: 155 t (125 t + 30 t)

Payload: Surface Hab. + TMI (22-t Prop.)

Mass: 155 t (102.8 t + 52 t)

Launch 1

Day 1

Launch 2

Day 30

Launch 3

Day 1

≈26 mo Later

Launch 4

Day 30

Crew Launch Vehicle Required

After Propellant Delivery

Mission Profile (Reference DRA5.0 and AIAA 7-Launch Architecture)

Payload: DAV + TMI Stage (22-t Prop.)

Mass: 155 t (102.8 t + 52 t)

Payload: Provided by Commercial Tanker

Mass: up to ≈192+ t (Each Cargo MTV)

Figure 28. Commercially delivered propellant option (EDS tanker derivative).
37

5. MARS MISSION ANALYSIS: KEY CHALLENGES

5.1 Required Launch Vehicle Build, Storage, and Transportation Rate

The number of vehicles that will be manufactured in order to meet mission requirements 

must be addressed as part of ground operations in order to successfully drive towards Mars architecture. The options assessed for Mars, and emphasized herein, are the NTP and chemical propulsion systems, with between four and six consecutive launches, at a rate of one every 30 days. 

In order to address extensibility of the current Ares V concept to the Mars DRM, the 

CS element team investigated modifications needed to support some of the particularly stressing 

aspects of that particular mission including various manufacturing, transportation, and storage 

approaches required to support the higher flight rates. As the largest of the Ares V elements, the 

CS is particularly interested in these operational changes and the necessary testing infrastructure 

required to develop the Ares V vehicle (at the Marshall Space Flight Center (MSFC), Stennis 

Space Center (SSC), and/or others) and ground infrastructure needed to successfully assemble 

and launch the vehicle at this increased rate (at the Kennedy Space Center (KSC)).

For this assessment, the PA-C3D candidate POD vehicle was used as a baseline, as seen 

in figure 29. This offered a CS from which others could be derived, such as a vehicle with more 

(or less) dV output from the CS. Driving factors were identified in order to determine this, including the amount of propellant or dry mass required and performance of the engines (e.g., six 

RS-68B-E/0s may be employed to provide a higher dV from a PA-C3D-derived CS). It was determined that, if allowed to grow, the CS would exceed the Vehicle Assembly Building (VAB) height 

limit, and the Michoud Assembly Facility (MAF) diameter limit. These considerations may 

require decisions on additional investments in infrastructure, vehicle concept/configuration 

changes, or a new vehicle design.

The CS manufacturing team generated a plan for MAF to support the Mars DRA 5.0 missions. This plan maintains the current manufacturing and facility approaches planned for the Ares 

projects. These plans, which are documented in a 2009 MAF manufacturing study, include the 

capability of producing a CS every 3 mo as a one-shift operation. Although this rate cannot keep 

pace with consecutive 30-day launch intervals, the MAF timeline, as seen in figure 30,6 shows the 

need for operations to start soon enough to build ahead and store the necessary quantity of stages. 

This approach will require storage facilities to be constructed either at MAF or KSC. If stored 

at MAF, stages will ship to KSC every 30 days, probably requiring additional barges. If stored at 

KSC, each stage will ship upon completion from MAF every 3 mo. 

5.2 Required Launch Vehicle Launch Rate

The preliminary Mars Ground Operations Project’s (GOP) Operational Concepts (Ops 

Con) for the Mars DRM builds off launch rate assumptions made during the Mars DRA 5.0 
38

Yes

Yes

Yes

Yes

No

No

No

No Major Impact(s) to Lunar

Approach to KSC Ops? Major Impact(s) to

Lunar Approach

to Mfg at MAF? Major Impact(s) 

to Lunar Approach

to Testing?

Clean Sheet

PA-C3 D CS Derived

PA-C3D Vehicle

More Delta V

From CS

Less Delta V

From CS

Does PA-C3D Vehicle Satisfy

Performance Requirements?

Higher Dry Mass (Low Tech Risk)

Dry Mass Reduction

Propellant Offload Less Propellant 

More Propellant 

Lower Performing Engines

Higher Performing Engines

Fewer Engines

More Engines

New Engine

Slush LH2

Drop Tanks

Shorter Tanks

VAB Modifications

Composite Unpressurized Structure

Aluminum Lithium/Al 2050

Larger Diameter With Major Manufacturing Facility Mods

e.g., RS-68B

RS-68 Upgrades

Fewer SRB Segments

SRB Spacer Removed

Cross Beam Above LOX Tank

Figure 29. General decision tree for Ares V CS extensibility to Mars DRM.

Storage time (ST)

ST

ST

ST

ST

ST

150 Days

*Each solid bar represents processing of 1 core stage to completion. 6 shipments

Figure 30. MAF CS processing timeline.

assessment. The launch manifest objectives in DRA 5.0 include either 9 consecutive launches 

for the NTP option or 12 consecutive launches for the chemical option supporting the back-toback cargo and crew Mars campaigns, as shown in figure 31. For this study, launches were assumed 

to be every 30 days. 

In order to determine a preliminary Mars GOP architecture that would meet the launch 

spacing and flight rate recommendations, several assumptions were needed. The vehicle would 
39

Cargo Campaign 1

Cargo Campaign 1

≈26 mo

≈26 mo

Chemical Launches

NTR Launches

Cargo Campaign 2

Cargo Campaign 2

Crew Campaign 1

Crew Campaign 1

1 2 3 4 5 6 7

1 2 3 4 5

1 2 3 4 5

6 7

8 9 10 11 12

6 7 8 9 10

13

1 2 3 4 5

Figure 31. Proposed Ares V flight rate assumptions (NTR or chemical).

fit within the size constraints of the VAB (high bay (HB) heights, widths, etc.). The CS and EDS 

arrive in the VAB in time for launch Vehicle Integration (VI) (ship-to-integrate). The ship-tointegrate concept is chosen to minimize flight hardware processing at the launch site and to minimize infrastructure to support storage of flight elements. No concurrent lunar missions or Ares 

V launches for other programs would be available for the GOP while supporting a Mars launch 

campaign.

This study assumes the lunar program has been completed and the follow-on program 

is the Mars campaign. The Mars program begins with the following lunar program assets already 

in place: 

• Two Ares V mobile launchers (MLs) and two Ares V VAB HBs.

• Two transporters.

• Two launch pads (LC39 Pads A and B).

• The Space Station Processing Facility (SSPF). 

The preliminary GOP Mars architecture was developed utilizing existing lunar campaign 

assets, and then it identified any additional launch and processing infrastructure needed to meet 

the Mars DRA 5.0 requirements. This exercise also met the GOP objectives to lower design, development, test, and evaluation (DDT&E) and operations and maintenance (O&M) costs going into 

a new program.
40

The proposed Mars GOP architecture is based on a preliminary deterministic analysis 

using the first cut of the Ares V ground operations processing timeline. With the limited number of 

processing requirements defined, the Ares V timeline used Ares I timeline analogs where applicable, 

along with historical data (where available), and subject matter experts’ engineering expertise. This 

deterministic assessment analyzed four options: 

(1) The baseline lunar Ops Con.

(2) SRB Offline Stacking Facility (OSF) Ops Con.

(3) Vertical Integration Facility (VIF) Ops Con.

(4) The baseline lunar Ops Con without ML launch mounts.

The GOP favored the SRB OSF to be the proposed Ares V Mars architecture after taking 

into account development cost, O&M cost, life cycle cost, and mission manifest satisfaction. This 

option would allow the GOP to meet the proposed manifest with the fewest assets and facilities 

with the lowest upfront development cost. The subsequent preliminary Ares V Mars Ops Con, 

based on the proposed architecture, is depicted in figure 32.

Offline Processing/Servicing/Stacking Stacking/Integrated Test Launch Recovery 

SRB Retrieval

Recovery

LC-39 VAB

• Booster Processing 

Hangar AF/ARF/RPSF

• Booster Offline Stacking 

Offline Stacking Facility

***Assumes “Option A” Fairing*** 

NPF

EDS

• Haz Servicing

• Fairing Integration

• Encapsulation

• Buildup and checkout

• Non-Haz Servicing

• Element Integration

Chemical Payload

Payload

Fairing

Nuclear

Payload

Core Stage

RSRM

and SRB

SSPF EPF

• Buildup and Checkout

• Haz/Non-Haz Servicing

• Fairing Integration

• Encapsulation

• Booster Transfer to ML

• Vehicle Stacking and

 Integration 

• Integrated Test

• Hyper Load

• Cryo Load

• Launch Count 

OR 

Figure 32. Proposed Ares V Mars Ops Con (chemical or nuclear).
41

To meet the launch spacing assumption of 30-day centers for Mars missions, several 

upgrades to the GOP lunar architecture would be required. If a Nuclear Processing Facility 

(NPF) is not part of the lunar architecture, one would need to be brought on line to support 

nuclear payloads processing for either chemical or NTP missions and to support nuclear stage 

handling for NTP missions. If an NPF is needed for the lunar program, it should be sized or modifiable to support a future Mars program. In support of the Mars manifest, the GOP will need one 

additional (for a total of three) Ares V MLs and one two-bay OSF. The proposed additions to the 

lunar architecture are depicted in figure 33.

Offline

Stacking Facility 

NPF

SPE

(Spacecraft Processing Element)

MLE

(Mobile Launch Element)

LPE

(Launch Pad Element)

SRRE

(Space Recovery and Retrieval Element)

Orion Recovery Ship

SSPF

ARF

RPSF

Hanger AF

SRB Recovery Ships

MPPF

Payload Processing

Facility

or

VIE

(Vertical Integration Element)

VAB

VAB High Bays

SRPE

(Solid Rocket Processing Element)

NEW 

NEW 

NEW 

Ares V

ML

Ares V

ML

Ares V

ML

Ares l

ML

Ares V Pad 

Ares l Pad 

2 1

4 3

Figure 33. Preliminary proposed GOP Mars architecture (upgrades to the lunar architecture).

Bringing on line an OSF allows ground operations to stack the SRBs in a facility designed 

to minimize quantity distance (QD) restrictions and to perform SRB stacking operations outside 

the critical path. DDT&E and O&M costs for this new facility would be lower than that of a much 

larger VIF. This deterministic assessment also resulted in the fewest Ares V MLs and VAB HBs 

needed to support the launch manifest and flight rate. Since every option required the use of an 

NPF, this facility was not a discriminator in the Ops Con and architecture decisions.

Preliminary deterministic analysis that supports a flight rate of 11 launches per year with 

a minimum of 26-day launch spacing between launches is a 5-day, 3-shift workweek. This work 

schedule scenario allows for surge capabilities, if needed, by working a 6-day, three-shift workweek 

or the 7-day, three-shift workweek. A 6-day, 3-shift workweek is required to meet the chemical 

manifest, which allows for a maximum flight rate of 13 per year with a minimum 22-day spacing 
42

between the launches. The preliminary deterministic launch rate and launch spacing analysis results 

for the OSF option are shown in table 8.

Table 8. Effects of number of ML, HBs, and days in workweek for OSF option 

as it relates to flight rate and launch spacing, respectively.

Maximum Flight Rate

Scenario

×1 Pad

×2 ML (+2 HB) ×3 ML (+2 HB) ×3 ML (+3 HB) ×4 ML (+2 HB) ×4 ML (+3 HB)

Offline facility

× 1 OSF bay

5-day week No MOD 7.2 7.2 7.2 7.2 7.2

W/Mod

6-day week No MOD 8.6 8.6 8.6 8.6 8.6

W/Mod

7-day week No MOD 10.0 10.0 10.0 10.0 10.0

W/Mod

× 2 OSF bays

5-day week No MOD 8.9 11.3 11.3 11.3 11.3

W/Mod

6-day week No MOD 10.6 13.5 13.5 13.5 13.5

W/Mod

7-day week No MOD 12.3 15.6 15.6 15.6 15.6

W/Mod

Minimum Launch Spacing

Scenario

×1 Pad

×2 ML (+2 HB) ×3 ML (+2 HB) ×3 ML (+3 HB) ×4 ML (+2 HB) ×4 ML (+3 HB)

2 3 3 4 4

2 2 3 2 3

1 1 1 1 1

Offline facility

× 1 OSF bay

Standard

(5-day work week)

40 40 40 40 40

Standard

(6-day work week)

34 34 34 34 34

Standard

(7-day work week)

29 29 29 29 29

× 2 OSF bays

Standard

(5-day work week)

33 26 26 26 26

Standard

(6-day work week)

27 22 22 22 22

Standard

(7-day work week)

24 19 19 19 19

5.3 Payload Mass Required

The Mars DRM is very challenging in many aspects—durations on orbit (LEO, transMars coast, and in the Mars system), a bioastronautics perspective, the technology developments 

required to support the mission, and many others. Perhaps one that is most often discussed is the 
43

large launch mass requirement to complete a human mission to Mars. Some mission architectures 

have chosen to use many smaller vehicles to achieve this mass requirement, often at the expense 

of mission reliability, mission complexity, payload packaging considerations, infrastructure investments for the large launch rate, and overall mission cost. However, the Ares V system is capable 

of delivering very large masses to LEO, allowing for more contingency planning, increasing the 

mission reliability, and allowing for cost sharing with other on-going efforts by the Agency (space 

science mission, lunar mission, near-Earth object (NEO) missions, and others).

Figure 34 allows for a very top-level comparison of these driving space exploration mission 

payloads in comparison with other potential missions.7 Where most LEO applications are volume 

driven (ISS habitation modules or a next-generation space telescope), the exploration missions 

have very high mass and volume requirements. Having a heavy-lift launch system that is capable 

of delivering these very large, very massive payloads to LEO in support of the mission is a necessity for sustained exploration of the solar system.

Payload Mass 

0 m3

10 t 40 t 70 t 100 t 130 t 160 t 

2,500 m3

500 m3

1,000 m3

1,500 m3

2,000 m3

Payload 

Volume 

 8 m 

 10 m 

12 m 

 

NTP Stage for 

Mars MTV 

Chemical Stage 

for Mars MTV 

EDS for Mars MTV 

Mars

Transhab 

Mars Cargo in 

Dual-Use Shroud 

 5 m 

NEOs 

Mars Sample Return 

ISS LEO Augmentation 

(Double or More) 

Ultraviolet

Observatory 

Lunar

Landers 

ATLAST

(Telescope) 

SAFIR

Telescope 

Relay/Comm

Constellation 

Fuel Depot

Demo 

ISS LEO

Augmentation

(Single) 

Shroud

Diameter

Figure 34. Mass and volume requirements for multiple missions.
44

5.4 Mission Reliability

Achieving an acceptable mission reliability estimation will be a key challenge when undertaking initial human missions to Mars. While the DRA5.0 architecture is structured to reduce the 

likelihood of crew loss (cargo deployed before crew commits to mission, contingency consumables 

to allow the crew the ability to not commit to surface operations, etc.), the mission still consists 

of many systems across multiple MTVs working in concert to successfully perform a complex and 

challenging mission. The Ares V Mars campaign risk assessment is a top-level, preliminary estimate of the risk of campaign loss during the ground to LEO and TMI segments of the Mars campaign. The focus is on launch vehicle risk, LEO operations risk, and TMI engine burn risk from 

a loss of vehicle or delivered payload component perspective. Campaign loss is assumed to occur 

with the first loss of vehicle or payload component. All technology components are assumed to 

be fully mature such that their reliability does not improve any further with each use.

The campaign risk is assessed for the following three campaign architectures: 

(1) NTP (maximum Ares V participation).

(2) Chemical scavenger. 

(3) Commercial tanker.

The sensitivity of the commercial tanker campaign risk to tanker propellant payload capacity and interval between flights is also examined. Loss-of-crew risk is not evaluated in this assessment due to uncertainty in the NTP failure modes and due to its likely being a minor consideration 

compared to the loss-of-campaign risk.

The common characteristic among these campaigns is that a total of three MTVs are 

assembled in orbit and sent to Mars. The differences lie in the number of launches needed to supply the components and propellant for the MTVs, the propulsion systems (nuclear or chemical) 

driving the MTVs, and the duration the MTV components are forced to loiter in LEO as they wait 

for other components or propellant. The first two MTVs will carry cargo payloads to the Mars 

surface. After successful deployment of these payloads on the surface of Mars and after the return 

of the launch window (in about 26 mo), the third MTV, which carries the crew TranHab and crew, 

will be assembled in LEO from components and propellant and sent to Mars.

Under the assumptions made about the missions, the loss-of-campaign risk is found to be 

the least for the NTP campaign (17%, 1 in 5.9), followed by the chemical scavenger (22%, 1 in 4.6), 

and the commercial tanker (78%, 3 in 4). The risk increases with number of launches, which range 

from 8 launches for the NTP campaign to 41 for the commercial tanker campaign (assuming an 

18.5-t tanker payload). The LEO loiter risk sustained by MTV components, which is found to form 

a significant fraction of overall loss-of-campaign risk, may offer the best opportunity for reducing 

overall risk if one considers campaign architecture modifications that minimize MTV component 

loiter time or reduce the amount of loiter risk per day of loiter for each component. 
45

5.4.1 Ground Rules and Assumptions

The following seven assumptions were adopted to identify leading campaign risks:

(1) The first loss of vehicle, payload component, or propellant in the campaign results in 

a complete loss-of-campaign. No spares, relaunches, or repairs are assumed. Likewise, the consequence of a launch or delivery failure, other than the complete campaign loss, is not assessed.

(2) No launch, in-orbit, or operational delays are assumed.

(3) Common risks are assumed across launches and components. All Ares V launches 

are treated as carrying identical loss-of-mission (LOM) risk. All components are assumed to carry 

EDS-like loiter risk in LEO. All commercial tanker launchers are assumed to carry equal risk 

(though different than Ares V), with orbital operations assumed to carry the same risk as their 

Ares V counter parts.

(4) LOM risk probabilities for rendezvous, docking, and undocking failures incorporate 

retries. 

(5) Loiter is assumed to carry duration-based risk per component. In-space components are 

assumed to carry the same loiter risk, and the duration is based on the accumulated time due to the 

launch-gap durations. Postdocking/pre-TMI burn loiter is assumed to be negligible.

(6) Launch gaps are assumed to be 30 days for Ares V, 15 days* for commercial tanker, 

and 1 day for the CLV.

(7) 14-day loiter in LEO after MTV assembly for checkout.

(8) Commercial tanker propellant delivery assumptions: 18.5 t* per flight (based on Delta 

IV Heavy), 201 t each required for MTV1 and MTV2, and 250 t required for MTV3.

(9) A 26-mo interval is assumed between the Earth departure window utilized by the cargo 

and Mars habitat transfer vehicles (MTV1 and MTV2) and the window used by the crew transfer 

vehicle (MTV3). All launches from the ground are scheduled ahead of the respective TMI burns 

to maintain the 26-mo interval.

*Commercial tanker payload and launch intervals are varied in the commercial tanker sensitivity study.

5.4.2 Per-Flight Mission Operations and Risk Inputs

The mission operations for which risks are assumed are given in table 9.8 They are grouped 

as follows: (1) Launch and delivery, (2) LEO loiter, and (3) MTV TMI risks. Only LEO loiter risks 

are accumulated by loiter duration and number of loitering components (as component days). 

All other risks are assumed to be demand based. Launch vehicle and MTV TMI risks account for 
46

burn durations. Ares V ascent risks and LEO loiter risks are taken from the PA-C3′ loss-of-mission 

estimates. CLV ascent risks are based on Ares I. Commercial tanker ascent failure risk is assumed 

to be 1 in 100. The LEO loiter risk is based on thermal control system (TCS), RCS, avionics and 

communications, fuel cell, and structural MMOD failures. The values are scaled directly from the 

4-day loiter risks assumed in the Ares V PA-C3′ lunar sortie mission. Automated rendezvous and 

docking information is based on Russian heritage figures, and no adjustments are made for crew 

docking, which would allow for manual override dockings. Undockings are considered very low 

risk and are not explicitly included. Ascent and orbital fuel transfer and solar array deployment use 

placeholder failure probabilities. MTV NTP operational risk during TMI burn is primarily based 

on RL-10 risk, given the likely similarities of NTP and RL-10 engines. MTV chemical burn (used 

for scavenger and commercial tanker campaigns), which utilizes the J-2X engine, is assigned risk 

based on the restart burn failure rates assumed in the Ares V lunar sortie. The duration from first 

to last launch varies from approximately 850 days (NTP) to over 1,700 days (commercial tanker), 

assuming a 26-mo (780-day) Earth-Mars alignment window between the TMI burns of the first 

and third MTVs.

Table 9. Risk elements.

Mission Operation Risk Risk Unit 1 in

Launch and Delivery Ares V ascent 8.6×10–3 1 in 120

CLV 2.8×10–3 1 in 360

Comm. tanker 1.0×10–2 1 in 100

Ascent fuel transfer 1.0×10–3 1 in 1,000

Solar deploy 2.0×10–3 1 in 500

Rendezvous and docking 1.5×10–3 1 in 690

Orbital fuel transfer 2.0×10–3 1 in 500

LEO Loiter LEO Loiter 1.3×10–4 1 in 7,400

MTV

TMI

TMI MTV1/2 NTP 2.9×10–3 1 in 350

TMI MTV3 NTP 4.6×10–3 1 in 220

TMI MVT1/2 chemical 2.0×10–3 1 in 500

TMI MVT3 chemical 3.5×10–3 1 in 290

5.4.3 Loss-of-Campaign Risk Estimation

The overall campaign risk (based on the first loss-of-mission experienced in the delivery 

of any of the components or propellant) during LEO loiter or in the failure of an MTV during 

TMI burn is computed by accumulating the failure probabilities of all the operations in all of the 

launches. The relationship between launch and operations by demand or duration is given in table 

10 for the three campaigns. 

5.4.4 Campaign Comparisons

The comparison of loss-of-campaign risk is shown in table 11. The risk is further decomposed into risk carried by the launch vehicles and their payload (including loiter and MTV risks). 
47

Table 10. Launch manifests with assigned activities (each of which incurs a risk indicated in table 9).

Launch Qty

Ares V 

or CLV or 

CHLLV

Planned 

Day of 

Launch Payload

Ascent—Core—Engine

Ascent—Core—Nonengine

Ascent ED—Engine

Ascent EDS—Nonengine

CLV—Engine

CLV—Nonengine

CHLLV—Engine

CHLLV—Nonengine

Ascent Fuel Transfer

Solar Deploy

Rendezvous

Docking

Orbital Fuel Transfer

Undocking

Loiter

TMI MTV1/2 NTP Engine

TMI MTV1/2 NTP Nonengine

TMI MTV3 NTP Engine

TMI MTV3 NTP Nonengine

TMI MTV1/2 Chem. J-2x

TMI MTV1/2 Chem. Nonengine

TMI MTV3 Chem. J-2x

TMI MTV3 Chem. Nonengine

NTP

1 1 Ares V 0 NTP propulsion 1 1 1 1 1 1 1 1 134 3 1

2 1 Ares V 30 NTP propulsion 1 1 1 1 1 1 1 1 134 3 1

3 1 Ares V 60 Mars cargo 1 1 1 1 1 1 14

4 1 Ares V 90 Mars habitat 1 1 1 1 1 1 14

5 1 Ares V 779 NTP propulsion 1 1 1 1 1 1 136 3 1

6 1 Ares V 809 NTP fuel 1 1 1 1 1 1 1 45

7 1 Ares V 839 TransHab module 1 1 1 1 1 1 1 15

8 1 CLV 840 CEV 1 1 1 1 1 14

Chemical Scavenger

1 1 Ares V 0 Mars habitat + TMI stage 1 1 1 1 1 268 1 1

2 1 Ares V 30 Mars cargo + TMI stage 1 1 1 1 1 208 1 1

3 1 Ares V 60 Tanker EDS 1 1 1 1 1 1 1 1 1 2 0

4 1 Ares V 90 Tanker EDS 1 1 1 1 1 1 1 1 1 2 0

5 1 Ares V 120 Tanker EDS 1 1 1 1 1 1 1 2 2 3 0

6 1 Ares V 809 TEI and MOI stages 1 1 1 1 2 1 210 1 1

7 1 Ares V 839 EDS TMI stage 1 1 1 1 1 1 1 1 1 75

8 1 Ares V 869 TransHab module 1 1 1 1 1 1 1 1 45

9 1 Ares V 899 TMI stage 1 1 1 1 1 1 1 1 15

10 1 CLV 900 CEV 1 1 1 1 1 0 14

Chemical Commercial Tanker (18.5 t)

1 1 Ares V 0 Mars habitat + TMI stage 1 1 1 1 1 1,978 1 1

2 1 Ares V 30 Mars cargo + TMI stage 1 1 1 1 1 2,008 1 1

3 11 CHLLV 60 Propellant 11 11 11 11 11 11 0

4 11 CHLLV 105 Propellant 11 11 11 11 11 11 0

25 1 Ares V 1,094 TEI, MOI, TMI (empty) stages 1 1 1 1 2 1 2,025 1 1

26 1 Ares V 1,124 TransHab module 1 1 1 1 1 1 1 1 1,290

40 14 CHLLV 1,154 Propellant 14 14 14 14 14 14 0

41 1 CLV 1,740 CEV 1 1 1 1 2 2 0
48

Table 11. Loss-of-campaign risk for three architectures, with commercial tanker assumed to carry 

20 t of propellant per launch, with 15 days between commercial launches.

Campaign NTP

Chemical 

Scavenger

Commercial 

Tanker

Launches (Ares V + comm. + CLV) 7+0+1 9+0+1 4+35+1

Overall loss of campaign 17%, 1 in 5.9 22%, 1 in 4.6 78%, 1 in 1.3

Risk breakdown by launcher type

Ares V + payload 16%, 1 in 6.1 21%, 1 in 4.7 65%, 1 in 1.5

Commercial + payload N/A N/A 38%, 1 in 2.6

CLV + payload 1 in 120 1 in 120 1 in 140

Breakdown by launcher mission phase

Launch ascent only 1 in 16 1 in 13 1 in 3.0

Postascent delivery 1 in 43 1 in 23 1 in 7.7

Loiter (component days) 510 840 7,300

Loiter 6.6%, 1 in 15 10%, 1 in 10 63%, 1 in 1.6

TMI burn 1 in 32 1 in 120 1 in 120

It is also broken down by mission segment across all launches of each campaign. The differences 

between the NTP and chemical scavenger are small, with NTP gaining an edge across all breakdown categories except “CLV + payload” and TMI burn, both of which have minor overall impact. 

The numbers of launches, including CLV, are 8 for NTP and 10 for chemical scavenger. The chemical commercial tanker results indicate considerably worse risk for campaign loss. For this comparison, a tanker payload of 18.5 t, which corresponds to current Delta IV heavy lift capability, and 

a tanker launch spacing of 45 days were assumed. The 18.5-t assumption requires that there be 

36 tanker launches or 41 launches overall.

The mission phase risk breakdown in table 11 indicates that, in all cases, both launch 

(ascent) and LEO loiter carry a similar order of risk and represent the leading risks across all three 

campaign architectures. LEO loiter risk is carried exclusively by the payload of the Ares V launches 

in all cases, and commercial tanker launches are assumed to carry no loiter risk directly (except 

through the orbital phasing operation prior to docking and propellant transfer). Because there are 

only four Ares V launches in the commercial tanker case, its loiter risk is imposed entirely on those 

flights. Even without tanker risk factored in, the Ares V + payload campaign loss risk, at 67%, 1 in 

1.5, still far exceeds the risk incurred by the NTP and chemical scavenger architectures. This is due 

to the total number of loiter component days imposed on the installed components in LEO as the 

36 tankers fill up the MTV tanks with propellant. In all campaigns, loiter risk is significant; in the 

commercial tanker case, it is the dominant nontanker source of risk.

The commercial tanker campaign loiter risk is directly dependent on the number of tanker 

launches required (a function of the tanker payload) and the launch spacing between the tanker 

launches. In table 12, the results of a sensitivity study of commercial tanker loss-of-campaign risk 

are given for tanker launch gaps of 15, 30, and 45 days, and for 18.5-, 30-, and 120-t payloads. The 

15-day launch gap is a best-case optimistic launch rate, and the 45-day gap would correspond to 

a more currently viable launch rate. The 18.5-t payload is representative of the current commercial 
49

Table 12. Loss-of-campaign risk sensitivity against commercial tanker payload and gap 

between commercial tanker launches.

Campaign Tanker Gap 18.5 t 30 t 120 t (Ares V)

Tanker launches (MTV1, 2, 3) 36 (11, 11, 14) 23 (7, 7, 9) 6 (2, 2, 2)

Overall loss of campaign 15 days 59%, 1 in 1.7 47%, 1 in 2.2 20%, 1 in 4.9

Ares V + payload 33%, 1 in 3.0 25%, 1 in 3.9 13%, 1 in 7.6

Commercial + payload 38%, 1 in 2.6 27%, 1 in 3.7 7.8%, 1 in 13

Total loiter (comp. days) 870 1,800 640

Overall loss of campaign 30 days 70%, 1 in 1.4 55%, 1 in 1.8 24%, 1 in 4.1

Ares V + payload 51%, 1 in 1.9 39%, 1 in 2.6 17%, 1 in 5.8

Commercial + payload 38%, 1in 2.6 27%, 1 in 3.7 7.8%, 1 in 13

Total loiter (comp. days) 5,000 3,300 1,000

Overall loss of campaign 45 days 78%, 1 in 1.3 64%, 1 in 1.6 28%, 1 in 3.6

Ares V + payload 65%, 1 in 1.5 50%, 1 in 2.0 21%, 1 in 4.7

Commercial + payload 38%, 1 in 2.6 27%, 1 in 3.7 7.8%, 1 in 13

Total loiter (comp. days) 7,300 4,700 1,400

heavy lift capability, 30 t is a likely commercial payload capability in the near future, and a 120-t 

payload corresponds to the Ares V payload capacity. The range of loss-of-campaign risk varies 

from 78% (1 in 1.3) for a 18.5-t payload at 45-d intervals to 20% (1 in 4.9) for the 120-t payload at 

15-d intervals. The use of Ares V as a tanker at 15-d launch intervals brings the tanker campaign 

architecture in line with the NTP and scavenger options. However, the use of the Ares V diminishes 

the campaign architecture’s definition as a commercial tanker option.

One can argue that, of the three campaign architectures, the commercial tanker approach 

offers the greatest opportunity for backup flights should there be a propellant delivery failure. In 

one scenario, if one assumes that the commercial tanker launchers can always be relaunched and 

that the Ares V cannot, then the entire loss-of-campaign risk is indicated by the “Ares V + payload” risk alone. If one considers the 18.5- and 30-t cases of the commercial tanker but omits the 

tanker launch risk, then the commercial tanker campaign architecture can begin to approach the 

NTP and scavenger range of campaign risk if the tanker launch gap is in the 15-day range (this 

neglects the marginal increase in loiter risk by any relaunches).

5.4.5 Conclusions

Loss-of-campaign risk was estimated for three campaign architectures: NTP, chemical 

scavenger, and chemical commercial tanker. The risk estimate assumes that a loss-of-campaign is 

observed with the first loss of a launch vehicle, payload, or MTV, and that component reliability 

does not mature with each launch and use. Of these campaigns, the NTP architecture was found 

to bear the least risk (of loss of campaign (17% or about 1 in 6), followed by the chemical scavenger (22% or about 1 in 5), and the commercial tanker (78% for a 45-day launch gap and a 18.5-t 

tanker payload). The loss-of-campaign risk is found to be strongly dependent on the total number 

of launches and accumulated loiter time, which is partly a function of launch gap. Given the strong 
50

dependence found on loiter duration, it is recommended that the basis of the loiter risk per component day and the assignment of loiter risk be reexamined to verify their importance in overall 

campaign risk. If loiter risk persists in its impact, it becomes a viable target for redesign of the 

architecture in the effort to reduce overall loss-of-campaign risk. One such architecture that may 

bear examination in this context is the propellant depot variant of the commercial tanker, where 

the tankers fill a permanent propellant storage facility in LEO to which assembled MTVs would 

dock, acquire needed propellant, and embark toward Mars with a minimum of LEO loiter. While 

reducing loiter duration and the risk attributed to that key driver, the risk associated with the number of launches will be unaffected. Furthermore, more analysis will have to be performed to assign 

a reasonable risk value to the propellant depot itself and its associated accumulated loiter duration 

in LEO.

5.5 On-Orbit Assembly Operations/Functionality Required

The EDS Mars Extensibility Team conducted an assessment of the potential roles of the 

Ares V EDS element in support of a Mars mission campaign as laid out in Mars DRA 5.0 and 

the on-going Ares V assessment of the CxP Mars DRM during PA-C3′ and PA-C3′′, including 

on-orbit assembly operations for NTP, chemical, and EDS-as-MTV-stage campaigns. The objective was to identify the functionality required, to assign functions to EDS or MTV, and to develop 

EDS-extensibility options in the Mars DRM trade space.

Ground rules and assumptions of the assessment included the following:

• Basic Ares V EDS conceptual design for the PA-C3D configuration is used as a pointof-reference for assessing additional functionality requirements.

• EDS includes the J-2X engine as designed for lunar DRMs (LOX/LH2 engine).

• Within each campaign, a similar EDS configuration is utilized when possible.

• MTV payloads provide for their own needs.

• Appropriate guidance, navigation, and control (GN&C)/ACS/RCS components are included 

on EDS stage configurations as applicable.

• Assessment includes only nominal missions (no contingencies).

• Disposal burns are not addressed.

EDS Extensibility Team members included EDS element management, project lead systems 

engineers, VI systems integration and vehicle operations engineers, support systems engineers, and 

discipline engineers from propulsion, cryogenic fluid management (CFM), avionics, and operations.

Product deliverables resulting from the TIM include an EDS functionality/capability matrix, 

hardware/configuration changes matrix, and trade/technology trees. Functions identified and 

assessed for on-orbit assembly operations include the following:

• Attitude control.

• Debris avoidance (≥150-day loiter).

• Active rendezvous.

• Passive rendezvous.
51

• Docking for assembly.

• Deorbit and disposal.

• Power.

• Communications.

• Propellant settling.

• Reboost.

• Long-term CFM.

• MMOD protection.

• Insert payload to orbit.

• Undocking.

• Separation.

• Propellant exchange or transfer (chemical and EDS-as-MTV-stage only).

Assessment of the above functions included identifying the current EDS capability coverage 

to the functions, identifying any major issues with EDS providing a given function, and identifying 

major trades, analyses, or other drivers that need to be assessed to determine functional capability. 

The information was recorded in the EDS functionality/capability matrix (see table 13). Trade/technology trees were also constructed for each of the campaigns (see figure 35).

Extensibility Team members recorded discipline observations and system impacts. Key 

trades and required forward work/future analysis was identified in several areas. Key system level 

impacts noted include the following:

• Need for auxiliary propulsion system, including substantial mass and fuel requirement.

• Power requirements for ZBO capability.

• Increased radiation hazard for a ≥150-day loiter.

• Nuclear disposal requirement for all campaigns.

• In-space propellant transfer.

• Avionics required for automated docking.

• Significant increase in number and complexity of communication interfaces.

5.6 Earth Departure Stage as Mars Transfer Vehicle Stage

The EDS Mars Extensibility Team assessed the NTP, chemical and EDS-as-MTV stage 

campaigns for Mars exploration. Of the three campaigns, the EDS-as-MTV stage has numerous 

challenges that are specific to this campaign. These are summarized as follows:

• J-2X not available for circularization burn because second burn has to be TMI.

• Cooperative debris avoidance of multiple EDSs.

• EDS-to-EDS interfaces, including structural and communication.

• Disposal of stages providing second part of TMI.

• Propellant settling for multiple EDSs.

• Propellant transfer between stages.

• Propellant transfer between drop-tanks and main tanks.
52

Table 13. Sample section of EDS functionality/capability matrix.

Chemical MTV Assembly (c&h) EDS as MTV Stage ((c&h) NTP MTC Assembly

Nomenclature:• Mission: NTP, CMTV, EDSMTV• cargo (c), human)(h)• Element: MTV, EDS, DISS Current Capability Issues Trades & Tech Current Capability Issues Trades & Tech Current Capability Issues Trades & TechDocking for Assembly • No capability in 

 current EDS. • EDS does not  currently own pay- load interface

• MTV will need to have  structural docking  system• Part of the adapter  to EDS may stay with  MTV elements for  later assembly of  MTV.

• Proximity maneuvers.• Off-CG control.• Constraints imposed  by MTV docking  systems and proce- dures.• Additional H/W.

No capability current  in EDS.

• Extra systems (MPS,  structure, etc.) have to be  jettisoned prior to RPOD.• Configuration of EDS-to- EDS berthing/docking  hardware a potential  design, structural avionics,  and ops issue when also  combined with drop tanks  concept.• Slightly different implica- tions of this issue for  2-EDS inline cargo and  multi-EDS human mission.

• Configurational trades  to account for different  MTV configurations  and docking options.• Proximity maneuvers.• Constraints imposed  by MTV docking sys- tems and procedures.• Additional H/W.

No capability current  in EDS. EDSS does  not currently own  payload interface.

Assumption that NTP  (launch 1) power system  provides power for NTP  element docking (power  LH2, line connection, etc.).  If not, what are implications  to EDS?• MTV will need to have  structural docking system.• Part of the adapter to EDS  may stay with MTV ele- ments for later assembly  of MTV (especially human  launch 3). 

• Proximity maneuvers.• Constraints imposed  by MTV docking sys- tems and procedures.• Additional H/W.

Deorbit and disposal Limited capability

 in current EDS (no aux engines).

• Different back-away,  deorbit, disposal  scenarios.

• What additions to  RCS H/W system  needed?• Sizing of propellant to  Aux and RCS system.

Limited capability  in current EDS  (no Aux engines).

• Different back-away,  deorbit (for stages doing  first part of TMI), disposal  scenarios, deep space  disposal for stages provid- ing second part of TMI.

• What additions to RCS  H/W system needed?• Sizing of propellant to  aux and RCS system.

Limited capability  in current EDS (no Aux engines).

• Different back-away,  deorbit, disposal scenarios.• Need to be able to perform  disposal to higher NSO for  NTP and surface nuclear  power system on cargo  mission. Raising full stack  to NSO could be prohibitive  to this option without an  additional EDS, which has  lots of other implications.

• What additions to  RCS H/W system  needed?• Sizing of propellant to  Aux and RCS system.

Power Limited duration 

 power, which is  currently fuel cell/ batteries  (mission 1).

Current system cannot support ≥150 day loiter.

• New solar power array  system. Can it be  kitted?• Can we use MTV  power in certain  configurations?

Limited duration  power, which is  currently fuel cell/ batteries  (mission 1).

Significantly more power  required to accommodate  ZBO of main propellant  tank. How much power is  needed?

Power system options. Limited duration 

 power, which is  currently fuel cell/ batteries  (mission 1).

Current system cannot sup- port 180+ day loiter.

• New solar power  array system. Can it  be kitted?• Can we use MTV  power in certain  configurations?

Communications Limited capability 

 in current EDS.

• How does one EDS  communicate with  other EDSs?• What are comm.  Implications for  AR&D?• Knowledge of MTV  health and status  during ascent and  around RPOD.

• Who has control  during various  phases?• Comm interfaces.

Limited capability  in current LEO.

What communication needs  are there during highly  elliptical and Earth escape  phases?

Limited capability in  current EDS .

• Do the NTP elements  (before transhab) have  independent capability to  comm. Health and status  to MS, or does EDS ascend  that data?• What are comm. Implica- tions for AR&D?• Knowledge of MTV health  and status during ascent  and around RPOD.

• Who has control dur- ing various phases?• Comm interfaces.
53

EDS Mars

Extensibility Trades

and Assessments EDS as MTV Stage

Payload to Orbit

Payload Maneuvering and Orbit Maintenance

Attitude Control

EDS Capability Needed (Sp. MTV CMG Desat.)

Cooperative With MTV/OMV

EDS RCS/ACS

Combined

Separate

Debris Avoidance

EDS Capability Needed

Cooperative Avoidance Multiple EDSs

Active Rendezvous

Rendezvous Maneuvers

Impact to Avionics/HW

Implication to Aux and RCS/ACS

Passive Rendezvous Passive to What Level

Docking For Assembly

Proximity Maneuvers

Off-CG Control

Constraints Imposed by MTV Docking Systems

Additional HW

MTV Configurations and Docking Options

Prop Settling

Sensed Acceleration Rates Above 0.04 g

Below 0.04 g

Slosh Issues With RPOD

Prop Requirement

RCS/ACS or Aux

Reboost

Prop Requirement

Engines

Deorbit And Disposal

Needed Additions to RCS HW

Sizing of Prop to Aux and RCS System

Assured Disposal Of Nuclear Surface Power System

On-Orbit Operations

Power Power System Options

Long-Term CFM

Power Required for ZBO

Synergy With MTV ZBO Systems

RCS/ACS/Aux

Cryogenic Prop

Other Prop

Dedicated Tanks

Prop Scavenging/Main Tanks

Propellant Transfer

Which Set of Tanks Drives Which Maneuver

When is Transfer Necessary

Transfer Prop From Drop Tanks Direct to Engine

Refill Main Tanks

MMOD Protection

What Systems Need Protection Solar Panel Damage Tolerance

PNP Required (0.999?)

MTV Payload Adapter for Forward Skirt Protection

Stage Tending

Avionics

Health Monitoring

Interfaces

TMI

NTP MTV

EDS Mars Extensibility.mmap—4/29/2010—Mindjet

Chemical MTV

Figure 35. Trade tree for EDS as MTV stage.
54

• Disposal of drop-tanks.

• Multiple passes (at least two, perhaps four) through the Van Allen radiation belts.

• Restart conditioning of multiple J-2Xs.

• Staging for multiple J-2Xs during TMI.

• Longer single-burn duration and total burn-time duration for J-2X.

• Assured nuclear disposal (surface fission reactor power source).

5.7 Technology Developments Required

Many technologies are required to be developed to support the Mars mission concept; there 

is little debate in that regard. The common item that is mentioned early in any conversation on this 

matter concerns the in-space transportation system developments required. This can range from 

relatively low-technology investments using traditional chemical propulsion systems (either the 

DRA5.0-derived chemical propulsion modules or partially filled Ares V EDSs), to a more moderate technology development program that develops and incorporates in-space cryogenic propellant transfer, to even more challenging development programs from a variety of aspects (technical, 

political, development cost, etc.) that utilize NTP, nuclear-electric propulsion (NEP), solar electric 

propulsion (SEP), or plasma-based electric propulsion (e.g., the Variable Specific Impulse Magnetoplasma Rocket (VASIMR)). 

A relatively low technology investment portfolio is needed for the propulsion system itself 

if a path is chosen along the lines of the DRA5.0 RL-10 derived Chemical Propulsion Modules 

or if a J-2X derived stage is used. The use of these engine systems has been demonstrated for 

decades, but modest investment will be needed to ensure overall system integrity over a long–

duration assembly operation, control of propellant boiloff during assembly, autonomous rendezvous and docking (AR&D), etc. Furthermore, a J-2X derived solution will need to determine the 

optimum method for propellant delivery to LEO whether it be propellant residuals in the main 

tanks, usage of drop tanks, or other methods. 

A more moderate technology investment approach can be made in the chemical rocket 

arena. A technology development that could potentially decrease long-term costs, possibly establish a commercial market, and provide a performance benefit is cryogenic propellant delivery and 

transfer. Development programs have shown the feasibility of transferring propellant in orbit 

(e.g., Orbital Express, and noncryogenic propellant)9 and plans are in place to further demonstrate 

the feasibility of transferring cryogens (CRYogenic Orbital TEstbed (CRYOTE)10 and Fiscal Year 

2011 proposal for technology development programs). When coupled with the development 

of an in-space transportation platform such as the Ares V EDS, the possibilities are enormous—

up to 250 t of propellant powering a system of almost 300,000-lbf thrust and ≈450-s Isp. 

Calculations have shown the potential of delivering almost 200 t to the Moon or over 150 t 

to Mars (per EDS utilized). Furthermore, it takes an ≈80% full EDS to deliver the cargo payloads 

assessed in DRA5.0, and two EDSs (one ≈75% full and one at ≈50% full) to deliver the crewed elements of that Mars architecture, as mentioned previously. Overall, this technological development 

would open up the inner solar system to human exploration by using the propulsion systems that 

have been designed, developed, tested, and used over the past few decades. Furthermore, it has the 
55

potential to use Agency investments made for initial human activities on the Moon. This can result 

in a decrease in the marginal cost of the Mars campaign by using the predeveloped assets to the 

maximum extent possible and increase mission reliability by demonstrating the in-space propulsion 

system by a decade or more before committing to the Mars program.

From the von Braun era of rocket propulsion onward, NTP was thought to be the new era 

of in-space transportation systems. The Nuclear Engine for Rocket Vehicle Application (NERVA) 

development program demonstrated the feasibility of the technology, but other aspects of the 

system of a nontechnical nature have since stunted the development of this technology. Efforts in 

the early 1990s to underscore the need for this system type (Synthesis Group Report) and the sincecancelled Project Prometheus of the early 2000s still have not led to the full development of this 

system (let alone a domestic in-space demonstration of its capability). Present day excitement exists 

in the areas of NEP and/or plasma-based electric propulsion. However, even these technologies 

require large nuclear systems for electric power when scaled up to the class of propulsion required 

for a human mission and the mass requirements that entails. 

Beyond coupling in-space cryogenic propellant transfer with chemical propulsion or developing nuclear-based systems, several other technology developments were assumed for the assessment (and in DRA5.0). These are further depicted in table 14. This table can be viewed two distinct 

ways. First, if these technologies are developed for a ‘Mars first campaign’ a lot of them are backwards compatible with any other potential DRM that the agency may undertake in the foreseeable 

future (Moon, L2, NEO, etc.). Second, it can be viewed as an incremental technology development 

roadmap, walking through the inner solar system until a suite of technologies can be utilized to 

undertake the demanding Mars mission. 

Table 14. Technology investments assumed for chemical options in PA-C3′.

Technology Lunar Cargo Lunar Crew L2 NEO Mars Beyond Mars

Propulsion enhancement ● ● ● ● ● ●

Multiuse EDS Planned Planned ● ● ● ●

Light(er) weight structures ● ● ● ● ● ●

Automated R&D ◘ ● ◘ ● ● ●

Near zero boiloff rates ◘ ● ◘ ● ● ●

Build/launch rates ◘ ◘ ◘ ◘ ● ●

Larger shrouds ○ ○ ● ● ● ●

In-space cryogenic propellant transfer ○ ○ ○ ◘ ● ●

Multiuse shroud ○ ○ ○ ○ ● ◘

Notes: ●–Directly applicable

 ◘–Somewhat applicable

 ○–Not applicable
56

The rationale behind the items includes (at a very top level) the following:

• Propulsion enhancements include any improvements made in J-2X performance that can be 

applied across the board if that engine is utilized for the in-space transportation propulsive element. Furthermore, any upgrades in the CS or booster set would also have direct implications 

across the board if a common system is utilized, leveraging these assets in an appropriate manner. This allows for maximum return on investment of taxpayers’ dollars in the realm of transportation systems.

• Multiuse EDS implies an EDS (or second stage) that can provide a multitude of services to 

both the payload and as part of the mission architecture. It is currently planned to provide ETO 

ascent, maintenance of the Altair lander payload during a 4-d loiter period, passive control during an R&D maneuver with the Orion crew exploration vehicle (CEV), and finally a TLI maneuver to inject the Altair lander/Orion CEV to the Moon. This same rationale could potentially be 

applied across the board for inner-solar system exploration.

• Lighter weight structures would obviously have implications across the DRM set if a common 

system is used for exploration.

• Automated R&D (or autonomous R&D) would be useful across the DRM set if some type 

of assembly operation is required (including 1.5 launch or dual-launch scenarios).

• Will address ZBO in further detail.

• Build/launch-rate improvements could potentially allow the Agency to undertake multiple mission options concurrently or allow a more demanding mission (such as the Mars DRM) to be 

undertaken.

• Larger shrouds would obviously be required for the more demanding missions, such as Mars and 

beyond Mars, but it would also be useful for placing large space telescopes or other payloads in 

either the Sun-Earth L2 or Earth-Moon L2 (Lagrangian point). It may also be required to deliver 

assets to LEO in support of a human mission to a NEO.

• Finally, the dual-use shroud concept has shown the potential to provide ETO ascent protection, some thermal/space environment protection during the long loiter/trans-Mars coast period, 

thermal protection during initial MOI, and thermal protection for the entry, descent, and landing 

phase required to deliver the lander assets to the surface of Mars. This type of shroud may also 

provide a better systems approach to delivering other assets to destinations with an atmosphere.

ZBO is a fundamental assumption that was assumed in DRA 5.0 and in follow-on studies, and it may be a driving technology need for the type of architecture presented. If the LOX/

LH2 (or LH2 in the case of the NTP option) boils off (temperature increases cause a phase change 

from liquid to gas, gas increases pressure in the tank, increase in pressure activates pressure relief 

valves, and pressure release results in mass loss), the amount of propellant required to provide the 

appropriate dV may not be there when needed. This either requires more propellant than is actually 
57

required to be delivered to LEO so some boiloff may occur without jeopardizing mission success, 

or a thermal control system capable of keeping the cryogenic propellants cool may be required to 

prevent the phase change from occurring in the first place.

The amount of cryogenic propellant required for the cargo MTV is almost 90 t of LH2

for the NTP option or over 190 t of LOX/LH2 propellants for the chemical option. This is further 

demonstrated in figure 36, where over one-third of the mass at TMI is propellant for the NTP 

option, and about 60% of the mass is propellants in the chemical case. For the crewed MTV 

(fig. 37), the NTP option requires about 120 t for the TMI maneuver (about one-third of the total 

mass), while the chemical option requires about 315–340 t (about 60%). Even a small percentage 

of these propellants boiled off over a period of months would be detrimental to the mission 

architecture.

Total ≈326 t 

Total ≈540 t 

Total ≈566 t 

3 Ares V Launches 5 Ares V Launches 4 Ares V Launches 

Crewed MTV Comparison

TMI Stack TMI Stack TMI Stack IMLEO: Not

Utilized for TMI

IMLEO: Not

Utilized for TMI

IMLEO: Not

Utilized for TMI

Mass (t)

NTP DRA5.0 Chem. Option Chemical Scavenger Option

EDS Disposal Mass TMI Propellant

Stage Mass (everything else) Payload (dual-use shroud + payload)

600

500

400

300

200

100

0

Figure 36. Cargo MTV propellant inventory at TMI.

A very simple method of finding the potential impact of various boiloff rates includes 

assuming an allowed ‘average boiloff rate per day.’ As shown in figure 38, various boiloff rates per 

day over a long duration on orbit have various impacts to the system. Assuming 200 t of propellant is delivered to LEO on day 0 and it is allowed to boiloff at the various rates over 180 days, the 

impact can range from 70 t of propellant lost (rate of loss equal to 0.25% per day) to virtually no 

loss in propellants at all or ZBO. Therefore, a quantification of ZBO could be described as an average boiloff rate of 0.001% of propellant loss per day over the entire assembly operation.

However, the propellants for the mission options assessed are not always delivered on the 

first launch. This requires a slightly different approach to calculating the boiloff rate allowed for
58

Total ≈236 t 

Total ≈325 t Total ≈325 t 

2 Ares V Launches 3 Ares V Launches 2.5 Ares V Launches 

Cargo MTV Comparison

TMI Stack TMI Stack TMI Stack IMLEO: Not

Utilized for TMI

IMLEO: Not

Utilized for TMI

IMLEO: Not

Utilized for TMI

Mass (t)

NTP DRA5.0 Chem. Option Chemical Scavenger Option

EDS Disposal Mass TMI Propellant

Stage Mass (everything else) Payload (dual-use shroud + payload)

350

300

350

200

150

100

50

0

Figure 37. Crewed MTV propellant inventory at TMI.

225

200

175

150

125

0 20 40 60 80 100 120 140 160 180 200

Propellant Mass in LEO (t)

Loiter Period (days)

0.001% Prop Load

0.01% Prop Load

0.05% Prop Load

0.10% Prop Load

0.15% Prop Load

0.20% Prop Load

0.25% Prop Load

Figure 38. Mass loss of various boiloff rates (200-t initial, 180-day assembly).
59

200

175

150

125

100

75

50

25

0

0 20 40 60 80 100 120 140

Propellant Mass in LEO (t)

Loiter Period (days)

0.01% Prop Load

0.10% Prop Load

0.25% Prop Load

Required Prop. Load

Figure 39. Mass loss of various boiloff rates for NTP crewed MTV.

Propellant Needed for Cargo dV (192 t)

250

225

200

175

150

125

100

75

50

25

0

0 50 100 150 200 250 300 350 400

Propellant Mass in LEO (t)

Loiter Period (days)

0.01% Prop Load

0.10% Prop Load

0.25% Prop Load

Figure 40. Mass loss of various boiloff rates for chemical cargo MTVs.
60

Propellant Needed for Cargo dV (192 t)

275

250

225

200

175

150

125

100

75

50

25

0

0 50 100 150 200 250 300 350 400 450

Propellant Mass in LEO (t)

Loiter Period (days)

0.01% Prop Load

0.10% Prop Load

0.25% Prop Load

Figure 41. Benefit of adding sixth launch to cargo MTV assembly sequence.

the mission. The NTP crewed MTV assembly sequence is shown figure 39, where about 87 t of LH2

propellant is delivered on the first launch and then about 94 t of LH2 propellant is delivered on the 

second launch. It can be seen that a boiloff rate as low as 0.01% per day would be close to meeting 

mission requirements, but a rate greater than that would cause the MTV to be significantly short of 

the propellant required to meet the dV target.

For the chemical MTV option, the same relationship basically holds true. As shown in figure 40, a boiloff rate of 0.01% would allow for the cargo MTV to also meet its dV requirement.

For the chemical option though, a sixth dedicated launch could be added in order to 

have two dedicated propellant tankers delivering LOX/LH2 to the two cargo MTVs. This could 

be planned as an option to decrease the mission sensitivity to ZBO, but still plan to have ZBO 

developed to decrease the number of launches required for the mission. As shown in figure 41, 

the boiloff allowance could be increased to 0.10% boiloff per day while still meeting the required 

propellant to perform the mission. Furthermore, mission planning could be arranged in such a way 

that the last propellant launch could arrive but transfer would not take place until the boiloff in 

each tank reduced the propellant load to a sufficient level that would not exceed the tank capacity 

of the TMI stage (or 80% fill constraint assumed during PA-C3′).

The same relationship holds true for the chemical option crewed MTV assembly sequence. 

Increasing from four to five launches for that MTV assembly allows for up to 0.10% boiloff per day 

while still meeting mission requirements (fig. 42).
61

Propellant Needed for Crewed dV (342 t)

450

400

350

300

250

200

150

100

50

0

0 50 100 150 200 250 300 350 400

Propellant Mass in LEO (t)

Loiter Period (days)

0.01% Prop Load

0.10% Prop Load

0.25% Prop Load

Figure 42. Benefit of adding fifth launch to crewed MTV assembly sequence.

Another integral benefit of this cryogenic propellant transfer approach is the ability to 

decrease sensitivity to the launch spacing requirement. DRA 5.0 assumed that launches had to 

occur within 30 days of one another to carry out the mission. However, propellant transfer and the 

ability to deliver ‘top-off propellant’ later in the sequence also allows for the 30-day launch center 

approach to be increased to either 45 days or 60 days, depending on how much boiloff the program 

is willing to accept and the technological development progress of ZBO systems. Figure 43 depicts 

sensitivity to the launch spacing requirement.

As shown in figure 43, the ability to transfer propellants onorbit can significantly reduce 

the sensitivity of the mission architecture to key design variables (ZBO of cryogenic propellants 

and spacing between launches). Furthermore, it is not required to increase the launches if further 

developments in either area result in ZBO or infrastructure to support increased launch rates.
62

200

180

160

140

120

100

80

60

40

20

0

170 day total 245 day total 320 day total 200 day total 290 day total 380 day total

30 day centers 45 day centers 60 day centers 30 day centers

4 Ares V Launches for Crewed MTV 5 Ares V Launches for Crewed MTV

45 day centers 60 day centers

Mass (t)

Boiloff Allowed

0.01%

0.10%

0.25%

Allowable Boiloff (Black Column) =

Delivery Capability—Prop. for dV

Cargo MTVs

Crewed MTV

140

120

100

80

60

40

20

0

200 day total

0.04% 0.03% 0.02%

N/A N/A N/A

0.21% 0.15% 0.12%

0.22% 0.16% 0.13%

290 day total 380 day total 230 day total 335 day total 440 day total

30 day centers 45 day centers 60 day centers 30 day centers

5 Ares V Launches (2.5 per Cargo MTV) 6 Ares V Launches (3 per Cargo MTV)

45 day centers 60 day centers

Mass (t)

Boiloff Allowed

0.01%

0.10%

0.25%

Colored Bars are Total Propellant Boiloff in LEO 

Using 3 Rates (Expressed as % Boiloff Per Day)

Boiloff, Launch Ctr, 

and Add’l. Flight 

Impact Comparison

–Preliminary NFTU!!!

Assumes 2×Launch Center

overall contingency and 20 days 

final assembly and checkout 

(80, 110, and 140 days total dwell 

applied prior to TMI). Note: NASA/

SP—2009–566-ADD levies 120 day 

contingency, rationale is unclear.

Allowable Boiloff Rate

EDS Current

Requirement

0.25% LOX

0.70% LH2

TLI performance sensitivity

indicated LOX boiloff more

significant impact. More detailed

comparison/analysis required

for TMI impact evaluation.

Figure 43. Benefit of adding fifth launch to crewed MTV assembly sequence.
63

6. LUNAR AND MARS MISSION SYNERGISTIC POTENTIAL

6.1 Infrastructure Sharing

Identifying the Mars architecture depended, in large part, on the commonalities between 

that and the lunar architecture already being assessed. It has been highly desirable to maintain as 

many commonalities as possible between the two transportation architectures in order to maximize 

the synergy, thereby reducing complexity and cost. As the Mars DRA 5.0 specifies, this synergy is 

defined by the subsystem technologies, space transportation elements, and the use of common ETO 

launch vehicles. Figures 44 and 45 illustrate the DRMs for the Moon and Mars, respectively, from 

the Ares V Ops Con (the chemical option is depicted in figure 45).

In figure 44, one may see that one Ares V (in addition to one Ares I with a crew of four) 

is required for an extended lunar surface stay. Once in LEO, the EDS and Altair rendezvous and 

dock with Orion. The EDS is employed to perform the TLI and then it is jettisoned. After the 

lunar stay, the TEI is performed by Orion to return to Earth.

Low Lunar Orbit 

Ares l-V Launch Sequence 

Shown. Option to Reverse

Sequence to be Preserved.

≈8 to ≈180 Day Surface Stays

Separate Cargo Mission

Provides Supporting Resources

Ascent Stage

Expended

Service

Module

Earth Departure Expended

Stage Expended

Moon

Earth 

Orion 

Orion 

Performs

TEI

Entry

Landing

EDS Performs TEI

Altair Performs LOI

EDS, Altair 

DRMs Vehicles are not to scale 

Earth Rendezvous Orbit 

LLO 100 km (54 nm)

Figure 44. One-and-a-half launch lunar architecture from Ares V Ops Con.
64

In figure 45, a similar assembly takes place in LEO, though on a larger scale. Two cargo 

MTVs and one crewed MTV are assembled and loiter in LEO until enough propellant is delivered 

(via other Ares V launches) to perform a TMI maneuver during the next synodic opportunity. This 

would require 9–11 Ares V launches for the chemical option. Likewise, seven Ares V launches, plus 

an additional Ares I crewed launch, would be required for this proposed Mars NTP campaign. 

≈540 Days on Mars

≈180 Days

≈180 Days

≈350 Days

≈26 mo ≈30 mo

Crew Transfers to Habitat

Lander Using Orion, TransHab

Orbit Mars During Surface Stay.

2nd Stage Inserts MTV Into

Trans-Mars Orbit

1st Stage Inserts MTV Into

Intermediate Earth Orbit

Ascent Stage and ISRU Module

Land on Arrival Habitat Orbits 

Until Crew Arrival.

Cargo 1 & 2 MTV Missions Crew Ares V disposed After LEO delivery

Ares V Flights

Ares V becomes Earth Departure Stage

Drop Tanks

A A B B C A B C D

Figure 45. Chemical option (i.e., EDS-as-MTV, or unofficially scavenger option).

If systems are independently developed for Moon and Mars architectures, multiple obstacles may need to be addressed, such as delayed Mars operations due to a technology gap between 

the two missions. Similar to the impending Space Shuttle gap, the ability (cost and labor) to curtail lunar operations to enable Mars missions would likely occur after lunar missions. Even with 

a renewed investment in a Mars mission, there would be no heritage technology or experience with 

systems designed for such.

In order to develop a sustainable architecture and space program, technologies developed 

for lunar missions should carry considerable commonalities to those required for Mars missions. 

Several aspects of the architecture offer potential for such synergy, including a common EDS 

module acting as either a TLI or TMI stage. Based on driving requirements, common lunar-Mars 

systems may be decomposed into such elements with similar capabilities.
65

6.2 Earth Departure Stage as Trans-Mars Insertion Stage

There are many benefits associated with employing an EDS as the TMI module. Sections 4.2 

and 5.6 of this TM address the application of EDS for Mars extensibility options in more detail. 

Eliminating the need to develop a new propulsive stage for TMI would have considerable benefits. 

The Mars DRM requires near ZBO technology to allow the cryogenic fluid management system to 

function during the TMI maneuvers. This technology development would prevent otherwise necessary modifications to EDS propellant tanks or MPS. Figures 46 and 47 represent upgrades made 

to the EDS if it would serve as a multiarchitecture stage.

Figure 46. EDS concept for Mars architecture.

Figure 47. EDS concept for lunar architecture.

The lunar EDS and TMI stage subsystems were compared as part of the Mars DRA 5.0 

study. The subsystems identified were: Trajectory and Mission Analysis, Propulsion, Structures, 

Thermal, Power, and Avionics. A summary of the comparison results follows.
66

6.2.1 Trajectory and Mission Analysis

The Mars missions would need to allow for significantly larger midcourse correction maneuvers. In addition to this, there would also be a difference seen in the main propulsion system (MPS) 

dV requirements. The LEO loiter time would also be extended for the Mars mission due to the time 

and support needed for assembly of the MTVs.

6.2.2 Propulsion

Currently, there are many similarities between the in-space stage being considered for the 

Mars mission and the Ares V upper stage (i.e., EDS). All modules would use the same LOX/LH2

propellant combination, and the propulsion systems would include helium (He) and/or autogenous tank pressurization systems for the pump-fed engines. If one system is employed for both 

the TLI and TMI, decisions will have to be made regarding the engine choice. The current lunar 

EDS design uses a single J-2X engine; whereas, the Mars stages could use several RL10-B2 engines 

or a single J-2X. The RL-10 engine is characterized by its expander cycle that produces a nominal 

thrust of 24,750 lbf at an Isp of 462.2 s. The current J-2X engine specifications state a much higher 

thrust of 294,000 lbf with a 448-s Isp rating.

6.2.3 Structures

Material choice is an important factor when assessing structures, though most components 

would be sized by the launch and staging loads. Components of the system need to be designed 

for the Mars orbit environment, including possible material degradation and the effects of thermal 

cycling. The material choice must account for occurrences such as microcracking during launch 

and outgassing. MMOD protection would also be required for the Mars mission. 

6.2.4 Thermal

Many of the following thermal features are common to both the lunar EDS and the Mars 

TMI stages:

• The heat collection and transport systems (e.g., circulating fluid systems, cold plates, 

and heat exchangers).

• Heat rejection systems (e.g., space radiators).

• The passive control systems (heaters, insulations, passive devices, phase-change materials, 

special materials and coatings, etc.).

However, the lunar EDS will be designed with passive cryogenic propellant storage while the 

Mars TMI stage will require ZBO active cryogenic propellant storage. The development of a twostage cryocooler operating at 20 K for LH2 storage will be required for the Mars mission active 

cryogenic propellant storage. The active approach affords essentially indefinite storage duration at 

the price of increased complexity, system hardware mass, power consumption, and heat rejection.
67

6.2.5 Power

Both the lunar EDS and the Mars TMI stages are powered by solar arrays. The lunar EDS 

contains a power management and distribution (PMAD) system to power itself and the lander 

during LEO loiter and TLI. There is no cryogenic cooling requirement on the EDS; it has an allpassive thermal system. On the other hand, the TMI stages will contain solar arrays and a PMAD 

system to power themselves only during LEO loiter and TMI and have no payload power capability. However, since the TMI stages are actively cryogenic cooled, they will require significantly more 

power than the EDS (about four times more), which means a much larger PMAD system than in 

the EDS. Also, because of the longer assembly and loiter time in LEO, the TMI solar arrays will 

need to handle more radiation degradation and MMOD damage.

6.2.6 Avionics

The following control features are common between the lunar EDS and Mars TMI stages:

• Both designs contain solar power generation and have navigation and communication components for LEO (sun and star trackers, inertial measurement units (IMUs), S-band transceivers, 

Global Positioning System (GPS), etc.); however, there are several differences. In particular, 

the instrumentation and data handling components will need to be larger in capacity in the TMI 

for the Mars mission to accommodate the active thermal cooling system. Also, the TMI requires 

an R&D system and low-gain interstage communication system for docking and assembly 

operations.

• The lunar EDS avionics system provides the GN&C capabilities of the stack during LEO loiter. 

It has a one-fault-tolerant system since the CEV takes over navigation control after docking, 

providing a two-fault-tolerant system for the crewed TLI operation. The Mars TMI stage avionics systems also consists of a one-fault-tolerant GN&C system to navigate the stage to assembly 

orbit and dock with the stack, but the reboost module will provide GN&C for the stack during 

assembly and testing.

• The TEI stage will be the controlling element during TMI burns and thereafter and will provide 

the two-fault-tolerant GN&C that is required with a crew aboard.

These operational differences mean that the two avionic systems will have differences 

in system architecture, interconnections, and software.2

6.3 Opportunity to Engage Commercial and International Partners

The role of global partnerships has evolved as the space industry has matured. Fueled by a 

competitive surge during the Cold War, America and the Soviet Union were incentivized to become 

the first nation in space. Since then, many actions have been taken to foster healthier relationships 

with space industries of fellow nations. The current program of record includes a gap in which the 

United States will rely on Russia to provide a crew-launch capability to the ISS, as needed. Table 

15 illustrates current international relationships including multiple Space Station endeavors, signed 

treaties, and promised cooperation on projects such as satellite assembly and maintenance.
68 Table 15. Current sample of international relationships.

China Russia India United States Canada Brazil Japan Thailand France

China Jointly tabled a draft 

treaty on the prevention of weapons and 

use of force in outer 

space. Possible joint 

venture on the study 

of the Moon.

Have agreed to discuss 

expanded cooperation 

and to start a dialogue on 

human space flight and 

exploration.

Signed agreement to provide Africa with data from 

jointly held “Earth resources 

satellite” through ground 

station in South Africa, 

Egypt, and Spain.

Developed extensive space 

exchanges and cooperation 

including progress in space 

science, Earth science. life 

science, satellite application, and satellite research.

Russia Jointly tabled a draft 

treaty on the prevention of 

weapons and use of force 

in outer space. Possible 

joint venture on the study 

of the Moon.

Space agreements 

signed for Global 

Navigation Satellite 

System (GLONASS).

Apollo-Soyuz, ISS 

(main players).

Interagency agreement drafted for a 

joint space program.

India Space agreements 

signed for GLONASS.

Next Steps in Strategic 

Partnership (NSSP)**; joint 

statement pledging to build 

closer ties in space exploration, satellite navigation, 

and commercial space; 

Chandrayaan-1.

THEOS* Joint mission to understand 

tropical weather phenomena.

United States Have agreed to discuss 

expanded cooperation 

and to start a dialogue on 

human space flight and 

exploration.

Apollo-Soyuz, ISS 

(main players).

Next Steps in Strategic 

Partnership (NSSP)**; 

joint statement pledging 

to build closer ties in 

space exploration, 

satellite navigation, 

and commercial space; 

Chandrayaan-1.

ISS

Canada Space panel regularly 

meets.

Brazil Signed agreement to 

provide Africa with data 

from jointly held “Earth 

resources satellite” through 

ground station in South 

Africa, Egypt, and Spain.

Japan Interagency agreement drafted for a 

joint space program.

ISS Space panel 

regularly 

meets.

THEOS*

Thailand THEOS* THEOS*France Developed extensive space 

exchanges and cooperation 

including progress in space 

science, Earth science, life 

science, satellite application, and satellite research.

Joint mission to understand tropical weather 

phenomena.

 *THEOS: First remote sensing satellite that shares timely and indepth information with Sentinal Asia, a regional project on satellite information for disaster management.**NSSP: Proposed expanded engagement on civilian nuclear activities, civilian space programs, and high-technology trade based on a series of reciprocal steps.
69

Such exploration missions as the Moon and Mars would undoubtedly benefit from international collaboration and commercial partnerships. There is considerable opportunity to not only 

foster the growth of commercial industry but to sustain it with technology that crosses the border 

between one mission architecture and the next. The design of such missions would require highly 

standardized components and processes, where the designing, development, manufacturing, and 

testing are prioritized such that all partners agree, in turn minimizing system complexity and cost 

for each partner. Furthermore, it is an unprecedented opportunity to involve and inherently incorporate the capabilities of the commercial industry and international partners in an overall Mars 

mission strategy.

6.4 Enabling Alternate Missions

This synergy would allow for opportunities to directly validate Mars elements during lunar 

missions (while encouraging advanced technological development for use on the Moon) instead of 

postponing development for Mars missions. The technologies employed herein may also have realworld applications. There is experience to be gained in routine manufacturing and system operation 

that would decrease risk and improve reliability.

6.5 Public Perception

A seamless transition from lunar to Mars architecture would also prevent workforce disruption. Such a direct correlation between Moon and Mars exploration would have highly visible 

milestones and applicability to more than one goal. Capabilities for flexible implementation allow 

the public to foresee the technology potential in future missions and for multiple mission opportunities. Without singularly focused technology development, these capabilities may easily transfer 

to Earth-focused applications.

Commonality between lunar and Mars exploration systems would significantly accelerate the onset of Mars exploration with early development of needed technologies. This would 

also reduce, if not evade, any development gap between lunar and Mars missions, even allowing 

for simultaneous exploration on the Moon and Mars. Life-cycle costs would also be considerably 

decreased, and it may be possible to show the benefits of developing assets for lunar exploration 

that can then be applied at smaller marginal cost to NEO exploration, which can then be applied 

at a much reduced marginal cost for Mars exploration. In this manner, the Mars exploration 

strategy would not have to carry the full development cost otherwise required in a ‘Mars Forward’ 

strategy.
70

7. FORWARD WORK

The Ares V team identified several areas of forward work on the Mars DRM. In addition 

to updating the requirements data set for a Systems Requirements Review (SRR), further developing the Ares V Operational Concepts Document, creating functional flow block diagrams of 

required functionality/capabilities, and increasing the team’s exposure to the Mars DRM, several 

areas of specific focus were identified. These include the trajectory work, vehicle subsystems, the 

ground operations concept, the Safety and Mission Assurance (S&MA) team, and the synergistic 

potential between Mars and NEO DRMs.

7.1 Ares V Performance—Trajectory 

Current performance estimates looked at both direct insertion circular cases and 120-nmi 

perigee insertion orbits. The elliptical orbit perigee was chosen to be high enough to allow EDS 

and payload to be inserted into a stable orbit. However, additional work is necessary to establish the potential performance gain from lowering the insertion orbit perigee below 120 nmi. In 

the extreme, this would result in negative perigee altitudes, with the optimal perigee being traded 

between ascent performance improvement and the propellant necessary to circularize the stack. 

Depending on the payload being flown, an additional constraint may be necessary to ensure the 

vehicle is inserted into an altitude sufficient to keep heating rates below payload requirements. 

Additionally, lower altitude perigees allow for the EDS to be disposed of without a deorbit maneuver, provided the EDS will not be needed on orbit. Further work will be necessary to consider how 

reliability and payload concerns affect the trade of insertion orbit used, as these concerns may not 

allow for a nonstable insertion orbit to be used. Initially, the IMLEO study included a –10-nmialtitude perigee, but those cases were not run.

Although the EDS may not be needed for on-orbit operations, it may still be necessary to 

insert it into a higher perigee stable orbit due to payload and operational considerations. In these 

cases, it would be necessary to perform a disposal maneuver from the elliptical insertion orbit. For 

the current analysis, only disposal from the final circular orbit was considered. Further work will 

be necessary to develop the elliptical orbit disposal dV budget and incorporate the performance 

impacts into the LEO performance capability.

7.2 Earth Departure Stage—Subsystems

There are opportunities for continued work in the EDS subsystem realm. Possible subsystem mass 

impacts to EDS may be evaluated, since 30 t was assumed for an EDS burnout mass (post-TMI) 

for the purposes of dV, payload, and total mass assessments. There is also a 5-t placeholder (in 

addition to the 30 t) for required hardware to dock EDS and mechanisms required to transfer 

propellant. This implies an 18% increase over the lunar EDS burnout mass (post-TLI), which was 

approximately 25.5 t. This additional mass required for Mars missions is due to RCS/ACS and/

or auxiliary propulsion, ZBO power and cryogenic coolers, structural mass impacts for larger 
71

payloads and shrouds, and ensured disposal of the empty EDS (for a two-burn TMI). Currently, 

there is no heritage technology that supports such ZBO capabilities, though it is necessary to 

enable the mission. In addition, in-space cryogenic propellant transfer is also of interest, since the 

proposed chemical propulsion options would utilize the EDS and reduce the number of Ares V 

launches. Future work may entail adding additional margin into the mission architecture by appropriately characterizing these fundamental technologies.

The engine of the EDS may also improve with continued analysis. The current J-2X 

burn duration requirements could be modified to burn for ≈9–10 min for ETO ascent followed 

by a required burn for ≈10–11 min for TMI. The total burn duration would then be on the order 

of 20 min; however, further analysis is needed.

7.3 Shroud

Further work might include identifying the impact of required shrouds. Perhaps a larger 

shroud would be needed to encapsulate both the TransHab and MOI/TEI modules. There is also 

the opportunity to assess impacts of maintaining shroud to orbit in a “multiuse shroud” case. Previous studies show an approximate 80% retained shroud mass impact to payload (1 lbm of shroud 

mass retained to orbit results in 0.8 lbm of payload reduction).

7.4 Mars Campaign Ground Operations 

Forward work on Mars campaign ground operations are as follows:

• Continue following the vehicle design and modify the Ops Con and architecture as appropriate. 

Should the vehicle outgrow the VAB physical constraints, the vehicle architecture changes, or the 

GOP determines the Ops Con is not feasible for any other reason, alternate ground-system architectures will need to be determined and evaluated.

• The GOP will continue making updates to the Ares V timeline and rerunning the flight rate and 

launch spacing models as applicable.

• The option of a dual-use pad should be studied to determine its feasibility and impacts on the 

Mars manifest.

• Additional studies to understand facility modification periods and their effects on availability 

are needed. Commodity (LH2, helium, etc.) quantities and availability for a higher manifest need 

to be studied further to determine impacts to launch availability. The amount of commodities 

required could drive additional ground systems and infrastructure to support a Mars campaign.

• Should launch site storage of any flight elements be required due to launch schedule delays or 

supply-chain production that is faster than integration/use, additional studies would be needed 

to look at the impacts to the GOP architecture to support such changes.
72

• Nuclear processing at KSC and a better understanding of ground processing requirements 

as related to nuclear stages also needs further study.

• Continued studies of the Ares V ML launch mounts are needed to coincide with future development of OSF operations to address feasibility of the preliminary proposed architecture. This 

study should include a more detailed analysis of ML launch mount feasibility and its impacts 

on the proposed Ops Con.

• Concurrent support of other missions, such as simultaneous lunar missions or other payloads, 

may affect the analysis and results.

• A discrete event simulation (DES) will provide better results by including common cause variability, modeling interactivities between VAB HBs, and accounting for delay probabilities. A DES 

will likely indicate requirements for additional ground systems resources to accommodate more 

realistic manifest performance.

7.5 Safety and Mission Assurance

Within the context of transportation, LEO assembly, and MTV TMI, there are a number 

of areas that should be addressed to allow the Mars campaign risk assessment to better reflect 

likely risks. The first such area is component reliability maturation. The risk figures used in this 

assessment assume fully matured technology and do not account for the reduced reliability associated with technology introduced early in its production cycle. A comprehensive study would 

include the effects of liquid and NTP engine reliability maturation, with sensitivity to the number 

of precursor and test flights with those engines.

The second area is the impact of launch and operational delays on campaign risks (e.g., 

the accumulation of launch delays that force an MTV to miss its TMI window or cause excess 

loiter and total boiloff to exceed the propellant margin). While the notion of relaunch in the event 

of a tanker launch failure was discussed above; relaunch of Ares V is not addressed.

Finally, given the large loiter risk and the high value of MTV assets, repair in LEO may 

become a necessary contingency. When considering the full extent of loss-of-campaign risk (and 

other FOMs) and its impact on campaign architecture, the post-TMI campaign elements of transMars cruise, Mars vicinity and surface operations, Earth transit, and Earth entry, descent, and 

landing should be represented at the very least in the form of a proxy in the risk model. All of these 

areas constitute elements that are under consideration for inclusion in a campaign risk simulation 

model currently under development. Loss-of-crew risk will also be assessed.

7.6 Mars and Near-Earth Object Synergistic Potential

Similar to lunar and Mars synergy, there is potential for forward work when evaluating 

NEO extensibility options. Similar to the Mars architecture, a NEO mission would depend in large 

part on the commonalities between that and the Mars (and potentially lunar) architecture already 

in place or being assessed. Delta-velocity requirements would be assessed and compared to other 
73

potential DRMs, including details of the Earth-to-NEO transfer dV and the rendezvous and eventual Earth-return (TEI) dV. Propellant transfer evaluation for a NEO DRM may emphasize similar 

technologies employed for a Mars DRM such as the decision to use in-space cryogenic propellant 

transfer. Propellant fill efficiency and cryogenic propellant boiloff impact has been assessed for the 

Mars DRM up to 5% propellant mass loss and/or 0.25% average boiloff rate per day, respectively.
74
75

REFERENCES

 1. “NASA’s Exploration Systems Architecture Study (ESAS)—Final Report; Part 6: Launch 

Vehicles and Earth Departure Stages,” NASA/TM—2005–214062, Marshall Space Flight Center, AL, November 2005.

 2. Mars Architecture Steering Group: “Human Exploration of Mars Design Reference 

Architecture 5.0 Addendum,” B.G. Drake (ed.), NASA/SP—2009–566, July 2009.

 3. Feldman, S.; Borowski, S.; Engelund, W.; et al.: “Analysis of Shroud Options in Support 

of the Human Exploration of Mars,” Aerospace Conference, 2010 IEEE, 2010.

 4. Jenkins, D.: Space Shuttle: The History of Developing the National Space Transportation 

System—The Beginning Through STS-50, Walsworth Publishing Company, Marceline, 

Missouri, 1992.

 5. Borowski, S.K.; McCurdy, D.R.; and Packard, T.W.: “7-Launch NTP Space Transportation 

System for NASA’s Mars Design Reference Architecture (DRA) 5.0,” AIAA–2009–5308, 

August 2009.

 6. “Ares V Michoud Assembly Facility Manufacturing Master Plan Study,” The Austin Company, 

August 31, 2009.

 7. Gonzales, A.: “Alternate Ares V Payload Accommodation,” July 2010.

 8. Fragola, J.R.; Maggio, G.; Frank, M.V.; et al.: “Probabilistic Risk Assessment of the Space 

Shuttle, Phase 3: A Study of the Potential of Losing the Vehicle During Nominal Operation,” 

NASA–CR–197808, Marshall Space Flight Center, AL, 1995.

 9. Orbital Express, <http://en.wikipedia.org/wiki/Orbital_Express>, Accessed September 21, 

2010.

10. Gravlee, M.; Wollen, M.; Walls, L.; and Rhys, N.: “CRYOTE (Cryogenic Orbital Testbed) 

Concept,” AIAA–2009–6440, Pasadena, CA, September 2009.
76

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Ares V Utilization in Support of a Human Mission to Mars

J.B. Holladay, J.P. Jaap, R.M. Pinson, S.D. Creech, R.M. Ryan, 

T.S. Monk,* K.E. Baggett,* M.D. Runager,** I.J. Dux,*** 

K.J. Hack,*** J.M. Hopkins,† C.E. Brown,† and T.A. Manning‡

George C. Marshall Space Flight Center

Marshall Space Flight Center, AL 35812

National Aeronautics and Space Administration

Washington, DC 20546–0001

Unclassified-Unlimited

Subject Category 15

Availability: NASA CASI (443–757–5802)

Prepared by the Science & Exploration Vehicle Office, Science & Mission Systems Office

*Zero Point Frontiers Corp., **Jacobs ESTS Group/Triumph Aerospace, ***Glenn Research Center, 

†Kennedy Space Center, ‡Ames Research Center

M–1299

Technical Memorandum

NASA/TM—2010–216450

Mars, launch vehicle, Ares V, EDS, Mars mission, scavenger, chemical campaign

01–11–2010

UU 90

NASA

U U U

During the analysis cycles of Phase A-Cycle 3 (PA-C3) and the follow-on 8-wk minicycle of PA-C3′, the Ares V team assessed the Ares V PA-C3D configuration to the 

Mars Design Reference Mission as defined in the Constellation Architecture Requirements Document and further described in Mars Design Reference Architecture 

5.0 (DRA 5.0) that was publicly released in July 2009. The ability to support the reference approach for the crewed Mars mission was confirmed through this analysis 

(7-launch nuclear thermal propulsion (NTP) architecture) and the reference chemical approach as defined in DRA 5.0 (11- or 12-launch chemical propulsion module 

approach). Additional chemical propulsion options were defined that utilized additional technology investments (primarily in-space cryogenic propellant transfer) 

that allowed for the same mission to be accomplished with 9 launches rather than the 11 or 12, as documented in DRA 5.0 and associated follow-on activities. This 

nine-launch chemical propulsion approach showed a unique ability to decouple the architecture from major technological developments (such as zero-boiloff technology or the development of NTP stages) and allowed for a relaxing of the infrastructure investments required to support a very rapid launch rate (30-day launch 

spacing as documented in DRA 5.0). As an enhancing capability, it also shows promise in allowing for and incorporating the development of a commercial market 

for cryogenic propellant delivery on orbit, without placing such development on the critical path of beyond low-Earth orbit exploration. The ability of Ares V to 

support all of the aforementioned options and discussion of key forward work that is required to fully understand the complexities and challenges presented by the 

Mars mission is further documented herein.

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