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IS20

George C. Marshall Space Flight Center

Marshall Space Flight Center, Alabama

35812
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i

NASA/TM—2011–216458

Executive Summary of Ares V: 

Lunar Capabilities Concept Review 

Through Phase A-Cycle 3

J.B. Holladay

Marshall Space Flight Center, Marshall Space Flight Center, Alabama

K.E. Baggett and S.M. Feldman

Zero Point Frontiers Corp., Huntsville, Alabama

February 2011

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>
iii

TABLE OF CONTENTS

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

2. ARES V VEHICLE EVOLUTION .................................................................................... 4

3. TRADE SPACE SUMMARY AND CONCEPT EVOLUTION ....................................... 12

3.1 Lunar Capabilities Concept Review Trade Space .......................................................... 12

3.2 Phase A-Cycle 1 Trade Space ........................................................................................ 14

3.3 Phase A-Cycle 2 Trade Space ........................................................................................ 14

3.4 Phase A-Cycle 3 Trade Space ........................................................................................ 15

3.5 Phase A-Cycle 3 Prime Trade Space ............................................................................. 16

3.6 Phase A-Cycle 3 Double Prime Trade Space ................................................................. 20

3.7 Overall Trades Assessment ........................................................................................... 22

4. POINT-OF-DEPARTURE DESCRIPTION ..................................................................... 23

4.1 Family-of-Vehicle Concept ........................................................................................... 23

4.2 Phase A-Cycle 3D—Description of the Candidate Point-of-Departure Vehicle ............ 25

4.3 Ares V Point-of-Departure Performance Capability ..................................................... 33

5. ARES V MARS EXTENSIBILITY ................................................................................... 34

5.1 Mars Transfer Vehicle Options: Nuclear Thermal Propulsion, Chemical, 

and Earth Departure Stage as Mars Transfer Vehicle Stage ......................................... 34

5.2 In-Space Earth Departure Stage ................................................................................... 37

5.3 Docking Adaptor and Propellant Transfer ................................................................... 38

5.4 Lunar and Mars Synergy .............................................................................................. 43

REFERENCES........................................................................................................................ 45
iv

LIST OF FIGURES

 1. Components of the Cx architecture .......................................................................... 1

 2. Heritage systems utilized on Ares V .......................................................................... 3

 3. Development history of Ares V from ESAS to LCCR .............................................. 4

 4. Recent Ares V vehicle concepts leading to the 51-series concept ............................... 7

 5. Minimum performance assessed ............................................................................... 10

 6. Ares V LCCR trade space (51-series vehicles) ........................................................... 13

 7. Shroud separation analysis: (a) Quad sector, (b) trisector, and (c) bisector ................ 15

 8. Ares V CS base engine layout trade ........................................................................... 16

 9. PA-C3D CS with horizontal engines ......................................................................... 17

10. Mass and volume requirements for multiple missions ............................................... 21

11. Maximum dynamic pressure versus vehicle payload .................................................. 22

12. Family of vehicles POD configuration for PA-C3′′ .................................................... 24

13. PA-C3D candidate POD vehicle (PA-C3′) ................................................................. 25

14. Expanded view of PA-C3D booster concept ............................................................. 26

15. Ares V POD booster: Five-segment, PBAN propellant, steel case ............................. 27

16. CS and booster sizing ............................................................................................... 28

17. Ares V CS element view ............................................................................................ 29

18. CS aft end expanded view ......................................................................................... 29

19. Ares V EDS element view ......................................................................................... 31

20. Ares V shroud element view ...................................................................................... 32
v

LIST OF FIGURES (Continued)

21. Ares V avionics and software .................................................................................... 33

22. NTP launch sequence and payload mass requirements ............................................. 35

23. Chemical propulsion module option ......................................................................... 36

24. Propellant delivery and transfer option (EDS-as-MTV-stage) ................................... 37

25. EDS concept for (a) lunar and (b) Mars architecture ................................................ 38

26. N-point docking mechanism with (a) and (c) latching claw, and (b) probes .............. 39

27. EDS-as-MTV model for loads and stress analysis ..................................................... 41

28. Chemical propulsion model for loads and stress analysis .......................................... 42

29. Dedicated in-space chemical propulsion stage and truss detail .................................. 42

30. Docking adaptor mechanism and ring structure ....................................................... 43
vi

LIST OF TABLES

1. Ares V VI major trades summary ............................................................................... 18

2. Family of vehicles concept comparison ...................................................................... 23

3. Performance analysis for candidate POD configurations ............................................ 24

4. Phasing trajectory sample functions ........................................................................... 40
vii

LIST OF ACRONYMS AND SYMBOLS

3DOF three degrees of freedom

ACO Advanced Concepts Office

ACU actuator control unit

Al-Li aluminum-lithium

APO Ares Project Office

ATP authority-to-proceed

BCPDU booster control and power distribution unit

BDM booster deceleration motors

BSM booster separation motors

CAD computer aided design

CARD Constellation Architecture Requirements Document

CEV Crew Exploration Vehicle

CFD computational fluid dynamics

CFM cryogenic fluid management

CLV Crew Launch Vehicle

CS core stage

CSB core stage bulkhead

CSCU core stage control unit

CSE core stage engine

CVS Concept Validation Study

Cx Constellation

CxAT Constellation Architecture Team

CxP Constellation Program

DACU data acquisition and control unit

DARU data acquisition and record unit

DDT&E design, development, test, and evaluation

DFIS development flight instrumentation system

DoD Department of Defense
viii

LIST OF ACRONYMS AND SYMBOLS (Continued)

DRA 5.0 Design Reference Architecture 5.0

DRM Design Reference Mission

EDS Earth Departure Stage

EELV evolved expendable launch vehicle

EPS electrical power system

ESAS Exploration Systems Architecture Study

ET external tank

ETO Earth-to-orbit

FEM finite element model

FOM figures of merit

FPR flight performance reserve

FSS flight safety system

GLOM gross liftoff mass

GN&C guidance, navigation, and control

HLLV heavy-lift launch vehicle

HTPB hydroxyl-terminated polybutadiene

IDA integrated design and analysis

IDAC integrated design and analysis cycle

IDAWG integrated design and analysis working group

IMLEO initial mass in low-Earth orbit

IOP ignition overpressure

IS imaging system

IS1 interstage 1

IS2 interstage 2

IS3 interstage 3

ISS International Space Station

KSC Kennedy Space Center

LCCR Lunar Capabilities Concept Review

LEO low-Earth orbit

LH2 liquid hydrogen
ix

LIST OF ACRONYMS AND SYMBOLS (Continued)

LOI lunar orbit insertion

LOX liquid oxygen

LRB liquid rocket booster

LV launch vehicle

MAF Michoud Assembly Facility

MCR Mission Concepts Review

MER mass estimating relationship

MGA mass growth allowances

MMOD micrometeoroids and orbital debris

MOI Mars orbit insertion

MPS main propulsion system

MTV Mars transfer vehicle

NEO near-Earth orbit

NTP nuclear thermal propulsion

OD outer diameter

OFI operational flight instrumentation

OML outer mold line

OMS orbital maneuvering system

PA-C1 Phase A-Cycle 1

PA-C2 Phase A-Cycle 2

PA-C3 Phase A-Cycle 3

PA-C3′ Phase A-Cycle 3 prime

PA-C3′′ Phase A-Cycle 3 double prime

PA-C3A Phase A-Cycle 3A (launch vehicle)

PA-C3B Phase A-Cycle 3B (launch vehicle)

PA-C3C Phase A-Cycle 3C (launch vehicle)

PA-C3D Phase A-Cycle 3D (launch vehicle)

PA-C3E Phase A-Cycle 3E (launch vehicle)

PA-C3F Phase A-Cycle 3F (launch vehicle)

PA-C3G Phase A-Cycle 3G (launch vehicle)
x

LIST OF ACRONYMS AND SYMBOLS (Continued)

PA-C3H Phase A-Cycle 3H (launch vehicle)

PA-C4 Phase A-Cycle 4

PBAN polybutadiene acrylonitrile

POD point-of-departure

RFCS radio frequency communication system

RS recovery system

RSRB reusable solid rocket booster

S&A safe and arm

SBKF shell buckling knockdown factor

SFA Shortfall Assessment

SRB solid rocket booster

SRR Systems Requirements Review

SSME Space Shuttle Main Engine

STS Space Transportation System

TCS thermal control system

TEI trans-Earth injection

TLI translunar injection

TMI trans-Mars injection

TPS thermal protection system

TransHab transit habitat

TVC thrust vector control

TVEC thrust vector control electronics

U.S. United States

USAF United States Air Force

VI vehicle integration

ZBO zero boiloff
xi

NOMENCLATURE

dV delta velocity

Isp specific impulse

Q dynamic pressure
xii
1

TECHNICAL MEMORANDUM

EXECUTIVE SUMMARY OF ARES V: LUNAR CAPABILITIES CONCEPT REVIEW 

THROUGH PHASE A-CYCLE 3

1. INTRODUCTION

This Technical Memorandum (TM) is intended to provide an overview of the current 

Ares V heavy-lift launch vehicle (HLLV) including its mission, role, and interfaces within the 

Constellation (Cx) architecture. It is meant to act as both a stand-alone record and a supplemental 

summary, such that follow-on appendices might be coupled with the Executive Summary to form 

a narrative. Each appendix will function independently of the others (Detailed Element Description, Payload Integration, etc.)

The Ares V HLLV provides the heavy lift capability for the Cx Program’s (CxP’s) exploration architecture. To use proven technologies, components, and infrastructure from the Saturn, 

Space Shuttle, and contemporary launch vehicle (LV) programs was a goal established during the 

Exploration Systems Architecture Study (ESAS) for Ares V and the Ares I Crew LV (CLV). Also 

where feasible, the Ares V project is directed to seek commonality between the Ares LVs to minimize development and operational costs and improve safety and reliability. The vehicle 

components of the Cx architecture are shown in figure 1. 

Ares V Cargo 

Launch Vehicle

Ares I CLV

Altair Lunar 

Lander

Orion Crew Exploration

Vehicle (CEV)

Earth Departure Stage (EDS)

Figure 1. Components of the Cx architecture.1
2

In order to carry out the National Space Policy directive to replace the Space Shuttle and 

complete the International Space Station (ISS), the Constellation development plan focuses on 

developing the Ares I and making it operational by 2015. Ares V is scheduled for a 2011 authorityto-proceed (ATP) decision that would enable it to make its first flight in 2018 and support lunar 

exploration in 2020.

The primary mission of the Ares V is to launch the Altair lunar lander into low-Earth orbit 

(LEO) and then send the lander and the Orion CEV into a translunar injection (TLI) trajectory 

to the Moon. (The Orion CEV is launched separately on Ares I.) In addition to crewed missions, 

Ares V will also launch automated cargo landers into LEO and then to specific lunar destinations. 

While retaining the goals of heritage hardware and commonality, the Ares V configuration continues to be refined through a series of internal trades to be discussed later in this TM. The previous 

Ares V point-of-departure (POD) configuration was an iteration of a study recommended by the 

Ares projects and approved by the CxP during the Lunar Capabilities Concept Review (LCCR)/

Ares V Mission Concept Review (MCR) in June 2008. 

In the current mission profile, the Ares V is launched from Kennedy Space Center (KSC) 

in Florida. Following booster and core stage (CS) separation, the Ares V EDS engine ignites at 

altitude, followed by separation of the payload shroud. Shroud separation occurs last in the staging 

sequence prior to reaching LEO to avoid recontact with the LV stack. The EDS delivers the EDS/

Altair stack into a stable LEO loiter orbit. Concurrently, the Orion CEV, launched by the Ares I, 

performs a rendezvous-and–dock maneuver with the EDS/Altair stack. After successful docking, 

ground controllers complete a system checkout of the EDS before it reignites its engine to perform 

the TLI burn and send the mated EDS/Altair/Orion stack to the Moon. The EDS is discarded after 

completion of the TLI burn, which marks the end of the Ares portion of the lunar mission. The 

current concept of operations calls for an Ares V launch as early as 90 min after Ares I, with three 

subsequent launch opportunities over the following 3 days, one launch opportunity per day. Ares V 

is currently designed for a 4-day loiter with TLI on the fourth day.

The design of Ares V shapes, and is shaped by, the requirements and designs of the other Cx 

components. The Ares V first stage booster is designed to share hardware, technologies, and manufacturing and operational facilities found in the Shuttle boosters. The Ares V EDS will share the 

J-2X engine and various subsystems now being developed for the Ares I upper stage. The Ares V 

CS design also employs five or six commercial RS-68 engines now used on the Delta IV. In the case 

of all those common components (shown in fig. 2), the Ares V application will require modifications for the Ares V mission that requires interface with the relevant hardware and management 

organizations. Ares V must also interface with the Orion and Altair projects regarding basic weight 

and volume requirements and numerous other design parameters such as the payload adapter and 

utilities supplied to Altair; structural, thermal, and acoustic loads; on-orbit power and thermal 

requirements; etc.
3

Upper Stage Derived 

Vehicle Systems

J-2X Upper Stage Engine

First Stage 

(Five-Segment SRB)

Elements 

From SRB

U.S. Air Force (USAF)

RS-68B Engine

From Delta IV RS-68

Ares I

25.5 t (56.2 klbm) to LEO Boeing Delta IV Ares V

71.1 t (156.7 klbm) to TLI (with Ares I)

63 t (138.5 klbm) to Direct TLI

187.7 t (413.8 klbm) to LEO

Note: Vehicles Not to Scale

Figure 2. Heritage systems utilized on Ares V.

The Constellation Architecture Requirements Document (CARD) provides the mass 

requirements for both the lunar sortie (crewed) and lunar cargo design reference missions (DRMs).2

For the lunar sortie mission, the CARD specifies an Orion control mass of 20.2 t 

(44,500 lbm) and a lunar lander control mass of 45 t (99,208 lbm). The sortie mission assumes a 

LEO destination orbit of 242 km (130 nmi) at a 29° inclination. The CARD loiter duration is not 

specified but has continued to evolve with program trades from 95 days to 14 days. For the LCCR 

trades, it was further reduced to 4 days. The TLI maneuver begins at a minimum 185-km (100-nmi) 

altitude with a delta velocity (dV) requirement of 3,175 m/s (10,417 ft/s) plus gravity loss.

For the cargo mission, the CARD specifies a cargo lander control mass of 53.6 t 

(118,168 lbm) and a total TLI payload mass of 54.6 t (120,372 lbm). The cargo mission assumes 

a phasing orbit Earth-to-orbit (ETO) destination. A loiter requirement is unnecessary because 

Orion is not part of the cargo mission operations concept; however, a few revolutions in LEO are 

anticipated to allow for system checkout prior to the TLI burn. It is worth noting that the SaturnV 

TLI payload capability was 48.6 t (107,445 lbm) for the Apollo 17 mission.

The CARD also imposes additional requirements on the Ares V such as the use of the fivesegment solid rocket booster (SRB) and five RS-68B engines in the CS and the Mars mission mass 

requirements. 
4

2. ARES V VEHICLE EVOLUTION

The first designs for a heavy lift capability that would eventually be named Ares V were 

studied during ESAS, which began in 2005.3 From ESAS to the concept approved during LCCR 

as the Ares V MCR POD concept, NASA studied more than 1,700 configurations of the Ares V. 

Following the LCCR POD, several trade studies were performed that resulted in an updated POD 

known as the Phase A-Cycle 3D (PA-C3D) concept. This section will summarize the evolution 

of Ares V from the ESAS trades up to the PA-C3D concept, including the 51.00.39 concept that 

served as the entry point to the LCCR trade study. An overview of the Ares V development history is shown in figure 3, including the LCCR trade space options and recommended POD concept 

approved by CxP, both of which will be detailed in later sections. A description of the major trades 

leading to the pre-51.00.39 concept follows.

2005 2006 2007 2008

• 1 J-25+ EDS

EDS Engine 

Change

Diameter 

Change

• 8.4-m OML

• SSME Core

• 2 J-25+ EDS

• 45-t Lander

• 20-t CEV

• No Loiter in LEO

Original ESAS

Capability

Change to RS-68B 

Core Engines

• Core dia: 8.4 m to 10 m

• Fvac: 490 k to 750 k

• Isp: 452.1 to 414.7 s

• Booster: HTPB to PBAN 

propellant (Ares I 53-06 

Thrust Trace)

IDAC 3 Trade Space

• SRM Propellant

• Stage Materials

• Extra Strap-ons

• Engine Type

• 10+ m dia. Core

• Number of Stages

• Shroud Material

• Shroud Size

LCCR Options

• 5.5 Segments

• Inert Spacers

• HTPB SRB Prop

• Composite Case 

SRB

• 6-Engine Core

• 10-m Diameter 

EDS and Shroud

Recommended 

Option

• 6-Engine Core

• 5.5-Segment PBAN

Figure 3. Development history of Ares V from ESAS to LCCR.

NASA studied commercial, government, and concept LV architecture systems prior to 2005, 

culminating in the release of the ESAS final report. In a trade tree pruning exercise, the ESAS team 

evaluated the following eight options:

(1) Nonassisted versus assisted takeoff.
5

(2) Vertical versus horizontal takeoff.

(3) In-flight propellant tanking versus no tanking.

(4) Rocket versus rocket and air breathing versus air breathing propulsion.

(5) Expendable versus partially reusable versus fully reusable systems.

(6) Single-stage versus two-stage versus three-stage concepts.

(7) ‘Clean-sheet’ versus derivative systems.

(8) Evolved expendable LV- (EELV-) derived vehicles versus both side-mount and in-line 

Space Shuttle-derived vehicles versus clean-sheet LV architectures. 

Figures of merit (FOMs) used in the studies were: Cost, reliability, safety, programmatic 

risk, mission performance, and schedule. These FOMs were applied to drive out the best option in 

the analysis. Additional considerations included legal requirements from the NASA Authorization 

Act of 2005, workforce skills, and industrial capabilities. After a thorough analysis of the entire 

exploration architecture requirements, EELV solutions were decided to be less safe, less reliable, 

and more costly than the Space Shuttle-derived solutions. The ESAS concluded that NASA should 

pursue the Space Shuttle-derived architecture for exploration due to several advantages relating 

to safety, reliability, and cost. The Space Shuttle-derived approach also allowed NASA to leverage 

significant existing ground infrastructure investments and personnel with significant human space 

flight experience. Overall, the Space Shuttle-derived approach was found to be the most affordable, 

safest, and reliable, both by leveraging proven human-rated vehicles and infrastructure elements 

and by using common elements across the architecture.

The ESAS-recommended Ares V vehicle, designated Concept 27.3, included two fivesegment steel-case SRBs with hydroxyl-terminated polybutadiene (HTPB) propellant, which has 

a higher specific impulse (Isp), density, and better mechanical properties than the polybutadiene acrylonitrile (PBAN) fueled Space Shuttle SRB. This Ares V concept had an 8.4-m (27.5-ft) 

diameter Space Shuttle external tank (ET) derived CS powered by five RS-25 Space Shuttle Main 

Engines (SSME) redesigned to be low cost and expendable. The 8.4-m (27.5-ft) diameter EDS was 

powered by two liquid oxygen (LOX)/liquid hydrogen (LH2) J-2S+ engines. Based on the 1970s 

era J-2S development program, the J-2S+ was intended to be a simplified version of the J-2 engine 

used for the Saturn upper stages. Both the CS and EDS had aluminum-lithium (Al-Li) structures 

and propellant tanks. The Ares V variant had a gross liftoff mass (GLOM) of nearly 2,900 t 

(6.4 Mlbm). It was based on a 45-t (99,000-lbm) lunar lander, a 20-t (44,000-lbm) CEV, and no 

loiter capability in LEO.

In the subsequent NASA studies to refine the ESAS recommendations, the architecture was 

simplified to reduce the number of new development programs. Further analysis of EDS performance showed that using a single J-2S+ engine provided more performance than two J-2S+ engines 

because the additional thrust provided by two engines during the ascent burn did not make up for 
6

the second engine’s mass during the less-thrust-to-weight-sensitive TLI burn. When Ares I propulsion changed from a four-segment booster to a five-segment booster for the first stage and from 

the RS-25 to a more powerful evolution of the J-2 (dubbed J-2X) for the upper stage, it opened the 

trade space on Ares V. A single J-2X replaced the J-2S+ engine on the Ares V EDS. The RS-68B, 

a variant of the commercial engine flying on the Boeing Delta IV vehicle, was leveraged for the 

Ares V CS. The RS-68 was designed as a simple, expendable engine with a high production rate. 

Using the RS-68 offered the opportunity to partner with the Department of Defense (DoD) to 

lower unit costs and gain flight maturity on Delta IV engine upgrades prior to Ares V flights. Program savings were estimated to be approximately $4.25 billion over the RS-25 SSME-based ESAS 

concept due to the high cost of producing a nonrecovered, nonrefurbished SSME.

Because of the RS-68B’s lower efficiency, the CS was enlarged from 8.4 m (27.5 ft) to 10 m 

(33 ft) in diameter to hold the additional propellants needed and to accommodate the larger nozzle 

and exhaust clearances needed for the larger engine cluster. The lower initial and recurring costs 

of the RS-68B and the cost, technical, schedule, and reliability risks involved with redesigning 

the RS-25 for altitude start, outweighed the cost of developing Saturn-class tooling and facilities 

needed to manufacture and process the larger CS. The booster design also reverted from HTPB to 

PBAN solid propellant for its better technical maturity and commonality with Ares I. The resulting 

Ares V configuration, designated Concept 33.8.64, had a GLOM of 3,300 t (7.3 Mlbm) and was 

nearly 110 m (362 ft) tall. It exceeded the payload performance of the RS-25 solution by approximately 4 t (8,800 lbm) to TLI and enhanced the commonality between the Ares vehicles, improving 

both development and operational efficiencies.

Concept 33.8.64 evolved in a series of configuration trades involving shroud diameter, direct 

lunar missions, CEV and upper stage on the Ares V, added gravity losses on TLI burns, and Flight 

Performance Reserve (FPR) allocation change. The resulting POD was the 45.0.2 concept vehicle. 

Nearly 111 m (365 ft) tall, the new concept served as a benchmark to determine the effects of 

engine upgrades, SRB variations, alternate materials, added stages, added boosters, added engines, 

and increased stage diameter. The effort established the impact of several changes that would be 

important to later trades including composite tanks and structures, additional CS engines (CSEs), 

additional SRBs or liquid rocket boosters (LRBs), and the addition of an S-II-class second stage as 

a third liquid stage.

However, the study also concluded that composite propellant tanks carried a high technical risk. HTPB boosters and a third stage carried undesirably high design, development, test, 

and evaluation (DDT&E) costs (the third stage was a cost and reliability issue). Additional SRBs 

incurred undesirable high launch pad modification costs, and vehicles more than 122 m (400 ft) tall 

led to prohibitive KSC facility costs.

The 46 and 47 series were both studies of three-stage vehicles with four and five J-2X engine 

second stages and shortened and lengthened CSs, respectively. Variants within those series traded 

the use of the commercial RL-10B2 engine on the third stage, six RS-68B core engines, nested 

tanks, and other changes. The three-stage designs offered higher performance and reduced loads 

through the TLI phase. They also allowed the lunar orbit insertion (LOI) and TLI maneuvers 

to be performed by the third stage, which would reduce the size of Altair. However, the addition 
7

of a second stage with four to five J-2X engines (instead of one) and a unique third stage added 

significant costs. The cost benefits to the Altair project resulting from Ares V, assuming the LOI 

functionality, were shown to be minimal. Propulsion systems, particularly the number of engines, 

are primary contributors to LV reliability, and the increased number of engines for these threestage options resulted in an overall lower vehicle reliability. 

The 45 series then served as the starting point for trades that became the 51 series of 

Ares V concepts, which, in turn, served as the basis of the LCCR trade space formally assessed 

in June 2008. This LCCR also acted as the Ares V MCR, ultimately identifying a POD vehicle and 

satisfying MCR criteria. Figure 4 shows the common features and notable variants of the 45-, 46-, 

47-, and 51-series concepts.

Notable Variants

00 All Al-2219 Construction

01 None

02 Al-Li + Composite Baseline

03 Composite EDS

04 Composite Core

05 2 Delta IV Booster Strap-ons

06 Regenerative RS-68B

07 6 RS-68B

08 36-ft Core; 6 RS-68 Core

09 Comp PBAN 166-06

10 HTPB-8A

11 150-in Diameter SRMs

12 Mix Ratio Shift for TLl Burn

13 4 SRMs

14 39-ft Core; 6 RS-68B Core

100 HTPB-8A; 36-ft Diameter, 

 6 RS-68B Core

45.00.XX

Common Features:

10-m Core

8.4-m Combined EDS

2 5-seg. PBAN, Steel

5 RS-68B Engines

1 J-2X EDS

Al-Li Construction

Notable Variants

00 None

01 5 RL-10B2 Core

02 6 RS-68B Core

03 LOX Aft. Core

04 5 RL-10B2 Core; Suborbital 

Burn

05 Nested, LOX Aft Core Tanks

06 5 RL-10B2 Core; Suborbital 

Burn; LOX Aft Core

100 Nested, LOX Aft. Core Tanks, 

6 RS-68B; Suborbital Burn

46.00.XX

Common Features:

10-m ‘Short’ Core

10-m Second Stage

8.4-m Dedicated EDS

2 5-seg. PBAN, Steel

5 RS-68B Engines

5 J-2X Second Stage

1 J-2X EDS

Al-Li Construction

Notable Variants

00 None

01 5 RL-10B2, Prop. Offload EDS

02 5 RL-10B2; Suborbital Burn

03 6 RS-68B; 5 RL-1082; 

Suborbital Burn

47.00.XX

Common Features:

10-m ‘Long’ Core

10-m Second Stage

8.4-m Dedicated EDS

2 5-seg. PBAN, Steel

5 RS-68B Engines

5 J-2X Second Stage

1 J-2X EDS

Al-Li Construction

Notable Variants

26 HTPB-30907; Part Comp EDS

27 HTPB-30907; % Seg. Spacer; 

 Part Comp EDS

28 HTPB-30907; % Seg. Spacer; 

 6 RS-68B

29 HTPB-30907; % Seg. Spacer; 

 6 RS-68B; Part Comp EDS

30 HTPB-30907; 100% Comp EDS

31 HTPB-30907; % Seg. Spacer; 

100% Comp EDS

32 HTPB-30907; % Seg. Spacer; 

 6 RS-68B; 100% Comp EDS

33 HTPB-30907; % Seg. Spacer; 

 6 RS-68B, Comp Dry Core; 

 100% Comp EDS

34 PBAN-26207; Part Comp EDS

48 HTPB-30907; Part Comp EDS

50 5.5 Seg. PBAN-33307; Comp 

Non-Pres Core & EDS

58 PBAN-26207; % Seg. Spacer; 

 6 RS-68B, Comp Non-Pres Core 

 & EDS

59 HTPB-30907; % Seg. Spacer; 

 6 RS-68B, Comp Non-Pres Core 

 & EDS

60 5.5 Seg. PBAN-33307; 6 RS-688; 

Comp Non-Pres Core & EDS 

51.00.XX

Common Features:

10-m Core

10-m Combined EDS

2 5-seg. PBAN, Steel

5 RS-68B Engines

1 J-2X EDS

Al-Li Construction

Composite Shroud

Mix Ratio Shift for TLI

Figure 4. Recent Ares V vehicle concepts leading to the 51-series concept.
8

Common features to all 51-series configurations are a 10-m (33-ft) diameter outer mold line 

(OML), composite materials for the payload shroud and all dry structures, and metallic (Al-Li) 

propellant tanks for the EDS and CS. The 51-series vehicles reflect the following changes in ground 

rules and assumptions to the 45.0.2 concept:

• 4-day to 14-day loiter period.

• 222-km (120-nmi) to 242-km (130-nmi) injection orbit.

• 8.4-m (27.5-ft) to 10-m (33-ft) EDS diameter.

• 8.4-m (27.5-ft) to 10-m (33-ft) payload shroud.

The 51-series trades were driven by the Performance Enhancement Study findings regarding increased CS propellant load, SRB propellant and length, and the addition of a sixth CSE. 

The 51.00.39 concept was selected as the entry POD for the LCCR and is characterized by its 10-m 

(33-ft) standard CS with five RS-68B engines and two five-segment steel-case PBAN-propellant 

reusable SRBs. Its TLI payload capability, in conjunction with Ares I, was 63.6 t (140,214 lbm). 

As a direct result of the 51.00.39 vehicle falling short of the CARD performance requirement for 

the Ares V vehicle, other concepts were presented at the LCCR that met or exceeded the derived 

65.2-t (143,741-lbm) TLI requirement (Altair 45 t (99,028 lbm) + Orion 20.2 t (44,553 lbm)). These 

options are discussed further in section 3.1, LCCR Trade Space. Notably, the 51.00.48 concept was 

the preferred option as the new Ares V POD. It was characterized by longer SRBs (five and a half 

segments as opposed to five segments on the 51.00.39), which in turn allowed the CS to increase 

in length. Furthermore, the additional propellant load in this longer CS benefited greatly from the 

addition of a sixth RS-68B. This configuration provided about 71.1 t (156,749 lbm) to TLI, exceeding the 65.2 t (143,741 lbm) requirement by almost 6 t (13,227 lbm). While not meeting the desired 

LCCR goal of 75 t (165,347 lbm) to TLI, this provided about 6 t (13,227 lbm) of margin to the 

program, and it allowed for closure of the overall mission architecture.

Following the LCCR, the Ares Project Office (APO) desired a bottoms-up engineering 

assessment of the LCCR vehicle. What came to be known as the Concept Validation Study (CVS) 

followed, allowing the engineering expertise from across the Agency to focus on specific design 

traits of the Ares V vehicle and provide confidence in the LCCR POD vehicle. Several areas of 

focus were identified including the functionality required during the loiter period, restartability of 

the EDS engine, a new aerodynamic model of the POD utilizing the Agency wind tunnel assets, 

the development of a computational fluid dynamics (CFD) model, and several others. Overall, this 

more detailed analysis provided the initial step from an advanced conceptual design using basic 

structural analysis, three degrees of freedom (3DOF) physics-based trajectory tools, and subsystem 

design using mass estimating relationships (MERs) to an engineering-based model using Agency 

assets in several engineering disciplines. This more detailed analysis assessed the performance of 

the LCCR/MCR vehicle at 68.3 t (150,576 lbm), approximately 4% different from the concept 

assessed for the LCCR. This reduced the project margin from almost 6 t (13,227 lbm) to slightly 

over 3 t (6,614 lbm).
9

Following the CVS, the APO established formal operating cycles for the Ares V vehicle with 

a plan for proceeding to the next project milestone—the Ares V Systems Requirements Review 

(SRR). The vehicle was divided into seven Level 4 elements (reusable SRB (RSRB), CS, CSE, EDS,

EDS engine, shroud, and avionics/software), which were integrated by the vehicle integration (VI) 

team. The establishment of formal analysis cycle products, interim checkpoints, and working technical interchange meetings characterized these cycles. The first of these was Phase A-Cycle 1 

(PA-C1) that ran from January 2009 through the end of April 2009. This cycle built upon the 

knowledge gained during the CVS but put it in the framework of an overall operating rhythm. 

While the analysis team further developed the models needed to assess the performance of the 

vehicle, products were established that focused on requirements of the system, the concept of 

operations of the system, and functional analysis. Additional analysis of the vehicle resulted in the 

TLI performance of the vehicle to drop from 68.3 t (150,576 lbm) down to 66.8 t (147,269 lbm), 

a reduction of 1.5 t (3,307 lbm), and resulting in the loss of half of the remaining project margin.

This reduction in TLI payload caused the APO to focus Phase A-Cycle 2 (PA-C2) on establishing lower and upper bounds on the performance of the Ares vehicle while continuing the development of the products focused on requirements, operability, and functional analysis. This cycle 

ran from May 2009 through the end of September 2009. First, the shortfall assessment (SFA) was 

an effort to establish a lower bound on performance by using additional margins and mass growth 

allowances (MGA) where appropriate, assessing the vehicle at higher loads, increasing reserves and 

required TLI dV, decreasing engine performance, and varying other design criteria such as propellant density, propellant boiloff rates during the loiter period, etc. The SFA came to the conclusion 

that a reasonable lower bound for Ares V performance would be in the range of 58 t (127,868 lbm) 

to TLI, over 13 t (28,660 lbm) less than the LCCR POD and 7 t (15,432 lbm) less than the CARD 

requirement. 

The upper bound was established by focusing the PA-C2 study configuration on a vehicle 

with evolved performance characteristics, such as a regeneratively cooled nozzle RS-68 (known as 

the RS-68B-E/0) with an increased Isp and a five-segment RSRB with shared commonality with 

the Ares I booster (same case lengths, propellant, avionics, TVC, nozzle length, etc.) but a more 

optimized propellant grain design for use on the Ares V. To maintain the propellant loading of the 

CS without increasing the diameter, an inert (empty steel case) spacer was added to the top of the 

booster. This provided the height necessary to attach the booster between the LOX and LH2 tanks 

of the CS. In addition, design assumptions that increased performance (or decreased loads on the 

vehicle) were the constraining of maximum dynamic pressure (Q) to 800 psf, increasing the knockdown factor used for global buckling calculations, utilizing the pressure inside the tank to alleviate 

flight loads for the purposes of tank design/analysis, and others. Overall, these changes resulted 

in an upper bound of 73.7 t (162,481 lbm), 8.5 t (18,739 lbm) (more than the CARD requirement. The establishment of the bounds, ranging from 58 t (127,868 lbm) to almost 74 t (163,142 

lbm), provided confidence that as the design analysis maturity increased (and design uncertainty 

decreased) the Ares V had options available to meet the 65.2 t (143,741 lbm) CARD requirement. 
10

While PA-C2 was underway, there was a great deal of uncertainty surrounding the architecture with respect to the Presidential Commission to assess the progress of the CxP and provide alternatives for space exploration, known as the United States (U.S.) Human Space Flight 

Plans Committee or the Augustine Committee. Thus, the APO management team approached 

Phase A-Cycle 3 (PA-C3) in a flexible manner. PA-C3 study configurations were established that 

would optimize the Ares V vehicle with respect to multiple key FOMs—specifically the shortest 

development time (and likely lowest cost) and maximum performance—with a combination of 

either. In this manner, specific recommendations from the Augustine Committee could be implemented based on the key drivers that those recommendations were derived from. Figure 5 shows 

the specific study configurations for PA-C3, in addition to the LCCR Study, PA-C1, and PA-C2 

vehicles that were previously assessed. A process was established to conceptually size a vehicle 

based on trajectory modeling, loads characterization, mass properties best practices, and element 

trades/analysis input. This provided a key understanding of vehicle options and performance trade 

space. PA-C3 ultimately assessed eight vehicles, and the characteristics of each may be seen in 

figure 5. 

LCCR (51.00.48)

6x RS-68B

Ares I Booster

Comp. Biconic Shroud

LEO Injected Mass: 188 t

TLI Mass: 71 t

LCCR Study (≈12/08)

Performance by ACO

PA-C1 (2/09–4/09)

Performance assessed using

TRV-1A ‘reference’ trajectory

PA-C1

6x RS-68B

Ares I Booster

Composite EDS

Comp. Biconic Shroud

LEO Injected Mass: 181 t

TLI Mass: 67 t

PA-C2 (6/09–9/09)

Performance assessed using

‘reference’ trajectory

PA-C2

6x RS-68B-E/0

Short Burn Booster

Composite EDS

Comp. Tangent Ogive Shroud

LEO Injected Mass: 198 t

* TLI Mass: 74 t

PA-C2 TLI Performance adjusted to 

account for updates to GR&A based on 

cycle analyses. Updated TLI Mass: 70 t

PA-C3 B

Five RS-68B-E/0

Ares I Booster

Metallic EDS

PA-C3 A

Five RS-68B

Ares I Booster

Metallic EDS

Ares I 

Booster

Five-RS-68B Five-RS-68B-E/0 Six-RS-68B-E/0

Metallic 

EDS

Composite 

EDS

Composite HTPB 

Booster

Short Burn 

Booster

PA-C3 C

Five RS-68B

Short Burn Booster

Metallic EDS

PA-C3 H

Five RS-68B

Composite Booster

Metallic EDS

PA-C3 F

Five RS-68B-E/0

Composite Booster

Composite EDS

PA-C3 G

Six RS-68B-E/0

Composite Booster

Composite EDS

PA-C3 E

Six RS-68B-E/0

Short Burn Booster

Composite EDS

 * PA-C3 Configurations have composite, 

tangent ogive shroud.

** Not analyzed. Performance estimated 

by interpolation.

General Note: Different sets of GR&A 

used for LCCR. PA-C1, PA-C2, and 

PA-C3 analyses, leading to different 

performance for similar configurations

PA-C3 D

Five RS-68B-E/0

Short Burn Booster

Composite EDS

** PA-C3 n

Six RS-68B-E/0

Ares I Booster

Composite EDS

PA-C3 Configuration ‘menu’

Quickest Development

Closest to LCCR 

GR&A

PA-C3 Family (11/09–9/10)

Minimum Performance Assessed

Figure 5. Minimum performance assessed.
11

The overall vehicle maintained common attributes from the previous PA-C2 vehicle: The 

33-ft OML, two SRBs, one J-2X on the EDS, etc. The number of CSEs was reduced from six 

to five for the Phase A-Cycle 3D (PA-C3D) vehicle based on the increased performance assessed 

during PA-C2. However, the types of CSEs and RSRB were varied, which are currently thought to 

be the critical path items for the Ares V vehicle development. Phase A-Cycle 3A (PA-C3A) was perceived to be the shortest development time, utilizing the Ares I booster directly (and the resulting 

development schedule and cost savings from doing so) and the NASA/Air Force upgraded RS-68B 

CSE. If more performance were recommended, then the regeneratively cooled, enhanced-Isp RS68B-E/0 CSE would be used, along with an RSRB grain design tailored for use on the Ares V. This 

vehicle configuration was dubbed PA-C3D. A combination of these RSRB and CSE options made 

up Phase A-Cycle C3B (PA-C3B) and PA-C3C.

Following the analyses performed for the B and C vehicles, Phase A-Cycle 3E (PA-C3E) 

through Phase A-Cycle 3H (PA-C3H) vehicles were identified as extensions of the POD family 

concept. The descriptions of these vehicles are detailed in section 4.1 along with their resulting 

performances.

During PA-C3, which ran from approximately October 2009 through early March 2010, 

analysis was completed on the PA-C3A and PA-C3D configurations with the assumption that performance for the PA-C3B and PA-C3C configurations would fall in between those values. With the 

less efficient RS-68B, reduction of one CSE, and the less optimized Ares I booster, the performance 

of the PA-C3A configuration was about 55.2 t (121,695 lbm) to TLI. With the more efficient RS68B-E/0 engine and a more optimized thrust trace on the SRB, the PA-C3D configuration delivered 

about 64.8 t (142,860 lbm) to TLI, less than 0.5 t (1,102 lbm)from the CARD requirement. This is 

near the center of the performance distribution found during PA-C2, mainly due to the reduction 

of one CSE, the use of minimum guaranteed performance values on the engines, increased flight 

performance reserves, and other design assumptions. With the performance of this vehicle near the 

CARD performance requirement, it was decided that this vehicle would serve as the new Ares V 

POD design. This vehicle is further detailed in section 4.2.
12

3. TRADE SPACE SUMMARY AND CONCEPT EVOLUTION

This section focuses on the trade space and the primary concept vehicles assessed during 

the LCCR/Ares V MCR through PA-C3. The first trade space discussed is based on the 51.00.39 

concept used as the entry point for the LCCR and is briefly described in section 3.1. PA-C1 analysis allowed for the full transition from an advanced conceptual design to an engineering-based 

model with established analysis products, to include assessment of the POD concept approved at 

the LCCR/Ares V MCR in June 2008. Architecture evolution from the LCCR (51.00.48) led to the 

more detailed configuration discussed in section 3.2. PA-C2 focused on verifying the design process, 

validating performance requirements, and assessment of key Ares V sensitivities until September 

2009. Trade studies performed during this time are described in section 3.3. Following PA-C2, 

refined performance targets were matched to programmatic cost and schedule requirements resulting in a PA-C3 configuration. Multiple iterations of this configuration were performed producing 

the current POD LV employed, called the PA-C3D vehicle. The following descriptions highlight 

the cycles and introduce the trades associated with each. More detailed descriptions of the major 

trades performed for the Mars extensibility study (Phase A-Cycle 3 Prime (PA-C3′) through Phase 

A-Cycle 3 Double Prime (PA-C3′′)) are also given since this served as the umbrella analysis for 

other trades and studies. 

3.1 Lunar Capabilities Concept Review Trade Space 

At the beginning of the Cx lunar study, the Ares V LCCR trade space focused on six vehicles based on the 51-series configuration. The 51-series configuration established a set of features 

common to all vehicles within the trade space, specifically:

• 10-m (33-ft) diameter OML for the central stack.

• Composite dry structures for the CS, EDS, and shroud.

• Metallic propellant tanks for the CS and EDS.

• A single J-2X EDS engine.

• At least five CS RS-68B engines.

• 9.7-m (31.8-ft) shroud barrel length.

The trade space was created by combining variations of both the CS and booster into different configurations of the Ares V vehicle. As one dimension of the trade space, two variations of 

the CS were considered. The first variation consists of a standard size CS with five RS-68B engines, 

and the second variation represents an extended CS with six RS-68B engines. As the second dimension of the trade space, three booster variations were leveraged to fully define the Ares V trade 
13

space. These booster variations were the five-segment, PBAN steel booster; the 5.5-segment, PBAN 

steel booster; and the five-segment, HTPB composite booster. As was discussed in the introduction, numerous vehicle permutations were analyzed leading up the 51.00.39 concept. The other five 

concepts that made up the LCCR trade space were the 51.00.40, 51.00.41, 51.00.46, 51.00.47, and 

51.00.48. Figure 6 summarizes the common features to the 51-series trade space and the distinguishing elements of each trade space concept.

5-Segment 

PBAN

Steel Case

Reusable

5-Segment 

HTPB

Composite Case

Expendable

5.5-Segment 

PBAN

Steel Case

Reusable

Booster

Standard Core

+ 5 RS-68B Engines

Opt. Core Length +

6 RS-68B Engines Common Design Features

Composite dry structures for 

CS, EDS, and shroud

Metallic cryotanks for CS and 

EDS

RS-68B performance:

I

sp = 414.2 s

Thrust = 797 klbf @ vac

J-2X performance:

I

sp = 448.0 s

Thrust = 294 klbf @ vac

Shroud dimensions:

Barrel diameter = 10 m

Usable diameter = 8.8 m

Barrel length = 9.7 m

1.5 launch TLI capability

Cargo TLI capability

Initial LCCR Study Reference Alternative New POD

Recommend for New POD

Core

51.00.39

51.00.40

51.00.41

51.00.46

51.00.47

51.00.48

63.6 t

69.7 t

67.4 t

61.5 t

74.2 t

66.3 t

68.6 t

60.2 t

71.1 t

63.0 t

Spacers: 1

Spacers: 1

Spacers: 0

5.0 t

4.5 t

3.7 t

–2.3 t–3.1 t

6.1 t5.6 t

Figure 6. Ares V LCCR trade space (51-series vehicles).

The 51.00.39 concept vehicle is the entry POD to the Cx Architecture Team (CxAT) lunar 

study. This vehicle features a standard CS with five RS-68B engines and a five-segment, PBAN steel 

booster. It offers 63.6 t (140,200 lbm) of payload to TLI for crewed missions. The 51.00.39 does not 

meet the minimum performance requirement.

The 51.00.40 concept vehicle features a standard CS with five RS-68B engines and a fivesegment, HTPB composite booster. The 51.00.40 concept can deliver 69.7 t (153, 700 lbm) of 

payload to TLI for crewed missions. While this concept surpasses the minimum performance 

requirement, two other concepts within the trade space offer greater performance. The 51.00.40 

was therefore eliminated from the trade space.

The 51.00.41 concept vehicle is a variant of the 51.00.39 concept. It has the same CS configuration and features an additional half-segment, PBAN steel booster. Given that the 51.00.41 

provides 67.4 t (148,600 lbm) of TLI payload performance, it was removed from the final 

consideration.
14

The 51.00.46 concept vehicle features the optional, extended CS with six RS-68B engines. 

It employs the standard five-segment, PBAN steel booster. The TLI payload capability is 68.6 t 

(151,200 lbm). 

The 51.00.47 concept vehicle was chosen as an alternate concept during the LCCR. Its 

design features the extended CS with six RS-68B engines and the five-segment, HTPB composite 

booster. This particular combination of CS and booster option offers the greatest TLI payload performance for crewed missions of 74.2 t (163,600 lbm). While this concept delivers the most payload 

to TLI, its design is dependent upon a successful infusion of composite technology development 

and incurs significant cost increases.

The 51.00.48 concept vehicle was chosen as the new POD concept at the LCCR. Its primary 

attributes were the extended CS and the 5.5-segment, PBAN steel booster. The 51.00.48 did not 

require significant funding for technology development nor did it incur the largest production and 

DDT&E costs. The 51.00.48 vehicle met the CARD performance requirements plus some margin, 

but it did not fully meet the desired TLI payload goal of 75 t (165,300 lbm). Following the LCCR, 

further HLV design iterations were performed in order to reach a desired POD for the PA-C1 

concept.

3.2 Phase A-Cycle 1 Trade Space

In order to achieve a higher fidelity design and mission analysis, the transition from the 

LCCR/MCR vehicle to the PA-C1 vehicle was marked by the development of key driving requirements, establishment of integration and element teams with associated analysis products through 

an integrated design and analysis (IDA) working group (IDAWG), processes for conducting trade 

studies and other analysis, requirements/functional analysis/operational concepts development and 

verification/validation efforts, etc. 

The PA-C1 vehicle was a variant of the 51.00.48 concept with the same LOX/LH2 CS 

configuration, two 5.5-segment, PBAN steel-case boosters; the LOX/LH2 EDS; and the payload 

shroud. At a height of 116.74 m (383 ft), the PA-C1 vehicle configuration necessarily added height 

to the CS to match a 0.3-m (1-ft) increase in length to the SRBs (to satisfy the ground rule of keeping the exit of the nozzles coplanar). In addition, performance dictated an EDS length increase 

of 0.24 m (0.8 ft) using updated ground rules and assumptions, while the payload shroud was 

increased by 0.03 m (0.1 ft). These extensions resulted in an overall vehicle height increase of 

0.58 m (1.9 ft) over the LCCR/Ares V MCR vehicle. 

Trades during the PA-C1 timeframe were performed mainly at the element level. These 

include the SRB length (five-segment Ares I/five-segment optimized for Ares V versus 5.5 segments), 

EDS and CS material trades, shroud geometry trades, etc.

3.3 Phase A-Cycle 2 Trade Space

PA-C2 was meant to provide the technical basis for both engineering and optimizing the 

Ares V concept. This cycle demonstrated an increasing focus on detailed analysis with sensitivity 
15

to analytical assumptions and hardware decisions. Trades at vehicle and element levels resulted 

in modifications from the PA-C1 configuration. The PA-C2 configuration featured two fivesegment, PBAN steel boosters optimized for use on the Ares V (shorter burn time without 

a ‘throttle bucket’) and a new vehicle height of 116.37 m (381.8 ft). Analysis showed that opportunities for mass savings existed with the CS and EDS LH2 tank design, mainly from the use of 

pressure within the tank for relief of flight loads, an increase in the shell buckling knockdown factor (SBKF), an evolved Al-Li alloy (Al-Li 2050 versus Al-Li 2195), and/or an isogrid design versus 

an orthogrid design to reduce the material thickness required. Other trades focused on specific 

design criteria such as CS ullage value increases from 2% to 3% nominal, and the averaged propellant boiloff value for LOX reduced from 0.35% per day down to 0.25% per day based on updated 

analysis. As the PA-C1 study investigated both the biconic and tangent ogive shroud geometries, 

the PA-C2 utilized the tangent ogive configuration and explored shroud internal acoustic treatment 

effects, shroud barrel length and payload packaging, and shroud separation options, which are 

depicted in figure 7.

(a) (b) (c)

Figure 7. Shroud separation analysis: (a) Quad sector, (b) trisector, and (c) bisector.

3.4 Phase A-Cycle 3 Trade Space

As the PA-C2 configuration acted to identify vehicle performance sensitivities, PA-C3 

sought to optimize the system by integrating DDT&E/performance sensitivities. Vehicle configurations chosen for PA-C3 included options for reducing cost and schedule and achieving more 

performance. The modifications made to the PA-C2 configuration include changes to develop an 

Ares V vehicle more quickly but with reduced payload to TLI. Featured changes include increasing 

the nominal ullage to 4% and other mass impacts such as deletion of shroud acoustic blankets and 

additional CS thermal protection. A separation study was performed to characterize booster-stage 

clearance based on a variable number of booster deceleration motors and failure criteria. As part 

of an integrated propellant tank study, several trades initially thought to be critical were ultimately 

removed and held as opportunities. Such trades included the shell-buckling knockdown factor, 

the effects of using an isogrid pattern as opposed to orthogrid, and multiple material options with 

thickness limitations. 
16

In addition, the number of CSEs was reduced from six to five and a detailed trade of the 

CSE/SRB geometric arrangement was performed, as shown in figure 8. This study assessed the 

realized decrease in radiation and plume impingement concerns from PA-C2. A CS engine element 

trade was performed to characterize RS-68 option trades (e.g., RS-68B, RS-68 regenerativelycooled nozzle with 21.6:1 and 30:1 expansion ratios, and RS-68B-E/0) to determine performance. 

This trade also assessed controllability with various engine gimbal and engine-out scenarios. The 

results showed vehicle effects assuming different engines and layouts. 

Fixed Engine

Gimbaling Engine

Area Reserved for ±10°

Actuator Tilt About Both 

Gimbal Axes

Thrust Structure Ring 

Frame (396-in OD)

CSE Attach Point (1/4)

LOX Turbine Inlet

LH2 Turbine Inlet

CSB

AA

BB

CC

DD

EE FF GG

Figure 8. Ares V CS base engine layout trade.

As mentioned, two configurations known as PA-C3A and PA-C3D were analyzed during the PA-C3 trade studies. Both configurations maintain a 10-m (33-ft) diameter tangent ogive 

shroud, EDS, and CS. PA-C3A utilizes five RS-68B CSEs with a lower Isp (due to the decreased 

nozzle expansion ratio and ablative nozzle relative to the RS-68B-E/0), and two five-segment, 

steel-case Ares I boosters. PA-C3D employs five RS-68B-E/0 CSEs with a higher Isp, and two 

five-segment, steel-case Ares V short burn boosters. These combinations would allow for further 

evolution of the Ares V vehicle from a lower performing, quicker-to-construct configuration 

(PA-C3A) to a higher-performing configuration (PA-C3D).

3.5 Phase A-Cycle 3 Prime Trade Space

An intermediate cycle followed PA-C3, as the analysis was not yet suited for Phase ACycle 4 (PA-C4). The Mars extensibility study was born during this cycle and answered four 

primary questions: 

(1) What is Ares V PA-C3D IMLEO capability as a function of altitude and inclination?

(2) What are the IMLEO growth potential/options for Ares V PA-C3D?
17

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

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

Several other trades were performed as part of this prime cycle, including the initial mass 

in LEO (IMLEO) trajectory study. This trade helped to characterize the Ares V PA-C3D vehicle’s 

capability to deliver mass to various insertion orbits, including elliptical and circular orbits. 

Additionally, dV budgets were determined for the circularization and deorbit burns. 

Barge transportation and storage were also assessed for the CS during this cycle. Transportation by water was deemed a feasible solution due to volume and mass of the PA-C3D CS. This 

trade asked whether the assets still exist to transport such a structure, since water transportation 

has been utilized for Saturn V, EELVs, and Shuttle. The only operational covered barge is the Pegasus, once used for transporting ETs from Michoud Assembly Facility (MAF) to KSC. The Ares V 

PA-C3D CS is too large for transport on the Pegasus barge, both in length and diameter. Similarly, 

the Delta Mariner was considered but it cannot fit a 10-m (33-ft) diameter stage on a transporter 

into its cargo bay. Figure 9 shows the CS with horizontal engines positioned on a Pegasus-like 

barge. This barge has a 1.5-m (4.5-ft) wider opening to the cargo bay than the Pegasus and allows 

the engines to have a reasonable clearance at the top of the opening. This orientation allows the 

barge height to be no larger than the Pegasus and is therefore good for transportation purposes. 

37.5- × 38.5-ft Stage

41.5- × 40.5-ft Opening

54.5-ft Wide Barge

Figure 9. PA-C3D CS with horizontal engines.

Another trade assessed during PA-C3′ was throttle optimization. The primary objective 

of this effort was to produce a better estimate of optimum maximum Q, knowing that the vehicle 

gains performance with increasing allowable maximum Q assuming fixed mass. Another analysis 
18

was completed to determine the payload that the current Ares V vehicle can deliver to different 

C3 values. The results of these trades are summarized in table 1 (Ares V Configuration Analysis 

Take-Aways, Private Communication, October 2010). 

Table 1. Ares V VI major trades summary.

Trade Problem Results

IDA Aerothermal

Base heating Determine significance of base heating given various nozzle configurations.PA-C3 RS-68 nozzle heating reduced when compared with 

six in a circle configuration because distance from SRB plumes 

is greater; radiation slightly more severe for PA-C3D versus 

PA-C3A; PA-C3 base heating is generally high compared with 

historical launch vehicles due to RS-68B engines, but determined to not be significant issue.

Plume impingement Identify the implications of low-altitude 

SRB separation.

Recontact of the SRB with the continuing stage as a result of 

plume impingement and subsequent rotation is unlikely given the 

following conditions, assuming high dynamic pressure and offset 

Core: Below 150 kft if RS-68 nozzle exit is 5 ft forward of SRB 

nozzle; Below 100 kft altitude if RS-68 nozzle exit is 10 ft forward 

of SRB nozzle; Below 80 kft altitude if RS-68 nozzle exit is 15 ft 

forward of SRB nozzle (based upon PA-C3D CSE arrangement).

Acoustics

Ignition over pressure (IOP) Comparison of PA-CA and PA-C3D 

vehicle configurations to predict 

transient overpressures. 

IOP environment for both PA-C3A and PA-C3D vehicle configurations is high; PA-C3A SRB IOP environment is higher while PAC3D CSE IOP environment is higher. Strongest IOP response at 

Shroud. Sound suppression system is recommended. Subscale 

testing for suppression system and unsteady pressure CFD 

model is needed. 

Liftoff analysis Comparison of PA-CA and PA-C3D 

vehicle configurations.

No significant differences in acoustic levels between vehicles; 

PA-C3D environment about 1 dB lower than PA-C2 environments. Scale model testing recommended to quantify water 

suppression benefits.

Acoustics CFD for IOP Average pressure values at several 

locations are used to compare steady 

and unsteady procedures.

Steady State: Larger variations between the turbulence models 

were found in the jet, MLP, and flame trench regions. Unsteady 

State: Initial comparison shows that unsteady analysis may be 

necessary in noise generating regions. 

Layout Studies

Engine base layout Determine most viable options by 

working engine layouts in parallel.

Each examined option could be workable, but some were not 

viable because CSE fuel and oxidizer intake ports were blocked 

by structural ring frames. A five-engine configuration with a 

centered bell was chosen for PA-C3, with a 10° control pattern, 

assuming RS-68B-E/0 nozzle geometry. 

Booster-to-CS study Identify impact on the geometry/

design of the CS thrust structure, with 

engine count and gimbal limits.

Integrated trade study was performed to determine optimal 

engine layout and associated thrust structure (based on engine 

proximity to booster nozzle). Separation study characterized 

booster-stage clearance based on variable number of booster 

deceleration motors and failure critiers.

CS transportation and storage Compare current transportation and 

storage to proposed approach for 

PA-C3D vehicle. Road, rail, and air 

transportation cannot support the size/

mass of PA-C3D CS.

Transportation by water is seemingly the only realistic way to 

transport the PA-C3D CS hundreds of miles at a time. Could 

alleviate congestion by transporting several at once for missions 

with high launch rates (thus transportation rates).
19

Table 1. Ares V VI major trades summary (Continued).

Trade Problem Results

IDA Performance Studies

Study configuration performance 

characterization

Establish process to conceptually size 

vehicle based on trajectory modeling, loads characterization, mass 

properties best practices, and element 

trades/analysis input.

Provided key understanding of vehicle options and performance 

trade space and determined vehicle geometry based on optimal 

propellant loading.

Vehicle-to-ground system Identify potentially significant impact 

on launch vehicle design, launch 

structure(s), and/or ground ops 

assumptions.

Found early impacts of the vehicle to existing ground systems 

that helped to reduce the potential for significant design changes 

later with CFD models that are directly extensible to other HLV 

configurations. Analysis was performed to determine the need 

and location of mobile launcher stays and dampers. Rollout 

and wind-shear loads modeling was also performed to identify 

potential solutions for reducing these effects.

Performance versus C3 curves Characterize Ares V PA-C3D vehicle’s 

capability to deliver mass to different 

C3 energies.

Preliminary conclusions were that there were no show stoppers 

with the nominal separation case. More aero data is required for 

failure cases.

Delta V split Determine dV budgets for circularization and deorbit burns.A single burn at apogee is required to circularize the orbit from 

the elliptical delivery orbit. For disposal from the circular orbit, a 

single burn can be performed to lower perigee to 30 nm, allowing 

the spacecraft to enter Earth’s atmosphere near perigee. 

Other Studies

LOC/LOM analysis Assess loss of crew probability from 

crewed Ares V failures.

Improved failure models for crewed risk assessment for LOC. 

Reduced conservatism in failure propagation. Baseline PA-C3A/D 

LOC probability estimate decreased since PA-C2. A & D nearly 

the same. MMOD strike ranked 4th cause of LOC (5% of total).

Booster separation study Perform sensitivity study of separation 

dynamics. 

Throttle optimization Improve fidelity of throttle profile 

analysis. Produce better estimate of 

optimum max Q point for PA-C3D 

vehicle configuration.

Second max Q event occurs at higher altitude for lower max Q 

cases. SRB propellant load decreases noticeably with decreasing max Q. Small increase in core propellant load with decreasing max Q. Highest nominal max Q approximately 1,000 psf with 

no throttle used. Minimum analyzed max Q approximately equal 

to 650 psf (similar to PA-C3A vehicle nominal max Q).

Mars Tank Mass Study

Pressure relief of flight loads Determine potential mass savings 

available from consideration of internal 

tank pressure in structural analyses.

CS showed greater sensitivity to the in-flight pressure relief of 

flight loads rather than an increase in SBKF. 

Shell buckling knockdown factor Evaluate mass savings for application 

of 0.65, 0.75, and 0.85 SBKFs.

SBKF was assessed during PA-C2 but deemed not critical for 

continued evaluation into PA-C3. 

Alternative materials Evaluate potential of thick plate Al-Li 

alloy to provide mass savings on 

buckling-critical cryotank structure.

Al-Li 2195 is adequate for worst case analysis assumptions 

on the EDS LH2 barrel. Al-Li 2195 shows a mass savings over 

Al-2219. Al-Li 2050 shows a mass savings in the CS LH2 tank 

due primarily to the thickness limitations associated with Al-Li 

2195. 
20

Table 1. Ares V VI major trades summary (Continued).

Trade Problem Results

Mars Transfer Vehicle Options

NTP Assess trade space with various 

propulsion elements: NTP.

Using the Ares V as a heavy launch vehicle, the NTP option 

delivers the necessary components for cargo and crewed MTVs 

in nine launches (assuming an additional Ares I launch for crew).

Chemical Assess trade space with various propulsion elements: Dedicated in-space 

stage.

Using the Ares V as a heavy launch vehicle, the chemical option 

that utilizes dedicated in-space stages (as opposed to EDS modules) delivers the necessary components for cargo and crewed 

MTVs in 11 launches (assuming an additional Ares I launch for 

crew).

EDS-as-MTV-stage Assess trade space with various propulsion elements using EDS as TMI.Using the Ares V as a heavy launch vehicle, the chemical option 

that utilizes EDS modules as TMI stages delivers the necessary 

components for cargo and crewed MTVs in seven launches 

(assuming an additional Ares I launch for crew).

Other Transfer Vehicle Options

Propellant boiloff Assess boiloff technology since all inspace transportation assessed by the 

Ares V team rely on this fundamental 

technology development to enable the 

missions.

NTP MTV Option: 1.3-t – 27-t ZBO potential impact for cargo 

MTV; 1.9-t – 41.5-t ZBO potential impact for crewed MTV. Chemical MTV Option: 2.1-t – 46-t ZBO potential impact for cargo MTV; 

2.7-t – 60-t ZBO potential impact for crewed MTV. The chemical option may also be impacted due to its required in-space 

cryogenic propellant transfer, accounting for 1% – 5% mass loss. 

Both NTP and chemical options would thereby accrue vehicle 

performance but would require at least one additional launch per 

MTV. 

Launch spacing sensitivity Determine if more launches allow for 

less sensitivity to ZBO for the chemical propulsion option.

Cargo MTV: For five launches (2.5 each), EDS can have about 

10 t of boiloff and still meet dV requirement. This increases to 

about 70 t of allowed boiloff for six launches. Crewed MTV: 

For five launches, 90+ t of boiloff would be allowed to meet the 

dV requirement. Therefore, an additional launch to assemble 

the cargo MTVs could reduce the sensitivity to the boiloff rate 

required for cargo MTV EDS, especially for the initial Mars 

missions, and similarly for the crewed MTV. 

Propellant transfer feasibility Assess EDS positioning during fuel 

transfer.

Nose-to-tail was recommended at the conclusion of the study 

assuming separation from the MTV stack and centrifugal acceleration.

Mars shroud study — loads Illustrate effects of heavier LEO 

payloads on vehicle loads.

Large shroud and heavy payload significantly change vehicle 

loads. May need additional throttling in trajectory to reduce 

moment.

3.6 Phase A-Cycle 3 Double Prime Trade Space

To complement the Mars study, a few spinoff trades were assessed such as identifying the 

aerodynamic effects for the Mars shroud, propellant boiloff, and launch spacing sensitivity. The 

fundamental technology developments that enable a mission such as Mars are zero boiloff (ZBO) 

cryogenic propellant thermal control systems (TCSs). The trade described the sensitivity of the 

MTVs to these technologies. For either the NTP or chemical option, large quantities of cryogenic 

propellants would be in LEO for extended periods of time. This would require a dedicated TCS 

to keep the propellants below the boiling point to reduce boiloff to a reasonable level. This introduces the possibility of having dedicated launches to make up for significant losses. The sensitivity 

to ZBO was found by assessing the impact of various boiloff rates (0.01%, 0.1%, and 0.25%) for 

propellant delivery flights assuming launch spacing intervals of 30, 45, and 60 days. 
21

In late 2010, several more studies were performed and they added fidelity to the PA-C3 

package. One such study acted as a supplemental part of the Mars extensibility study conducted 

during PA-C3′. The docking adaptor trade assessed the functionality and design of an automated 

rendezvous and docking system. Using the chemical scavenger nine-launch option (fig. 24) as 

a reference baseline for the analysis, the objective was to add detail to the docking adaptor and 

propellant transfer required for the crewed MTV concept. 

A payload survey was also conducted to identify potential HLLV payloads for the purposes 

of building a robust advocacy for HLLVs. This approach included payload mass estimates, physical 

characteristics, accommodation requirements, and launch dates. Shrouds were identified to support the study efforts and vehicle configurations were evaluated to ensure that the launch capability 

would meet payload needs. The heavy-lift payload survey is shown in figure 10. 

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 10. Mass and volume requirements for multiple missions.
22

Another trade that was performed during PA-C3′′ was an assessment of maximum Q versus 

the vehicle payload. Injected TLI and TMI payload mass were compared to maximum dynamic 

pressure assuming a 100-psf dispersion between nominal and design maximum Q, as seen in 

figure 11. 

67.0

66.9

66.8

66.7

66.6

66.5

66.4

66.3

66.2

66.1

66.0

65.9

65.8

65.7

65.6

65.5

65.4

65.3

65.2

65.1

65.0

10

8

6

4

2

0

–2

–4

–6

–8

–10

650 700 750 800 850 900 950 1,000

Maximum Normal Dynamic Pressure (psf)

Injected TLI/TMI Payload Mass (t)

Delta Vehicle Predicted Mass (t)

Injected TLI/TMI Payload Mass

Delta Vehicle Predicted Mean Mass

Figure 11. Maximum dynamic pressure versus vehicle payload.

3.7 Overall Trades Assessment

Table 1 highlights major trades that were assessed as part of the cycles previously mentioned 

(PA-C1 through PA-C3′′) (Ares V Configuration Analysis Take-Aways, Private Communication, 

October 2010). The results given reflect the latest status of the trades. Table 1 is intended for 

summarizing purposes only. For more details, please refer to the original packages. 
23

4. POINT-OF-DEPARTURE DESCRIPTION

4.1 Family-of-Vehicle Concept

The intent of a POD vehicle is to identify a baseline configuration that can be carried forth 

for analysis purposes. The lunar transportation POD was meant to provide crew and cargo delivery 

to and from the Moon, provide capacity and capabilities consistent with candidate surface architectures, provide sufficient performance margins, remain within program constraints, and result in 

acceptable levels of risk. 

During the initial stages of PA-C3, the candidate POD vehicle was identified as the PA-C3D 

Ares V vehicle. Since then, shorter cycles have been added, namely PA-C3′ and PA-C3′′. These 

additional cycles have allowed the VI team to further its efforts in defining a reference POD before 

delving into PA-C4. A family–of-vehicles concept was introduced in PA-C3′′and incorporates 

vehicles PA-C3A through PA-C3H. The differences between these Ares V vehicles may be seen in 

table 2. Table 3 presents the performance analysis for candidate POD configurations. Figures 12 

and 13 show the evolved family-of-vehicle concept, and the PA-C3D vehicle respectively. 

Table 2. Family of vehicles concept comparison.

Family of Vehicles

A B C D E F G H

Propulsion

Five engines

Six engines

RS-68B

RS-68B-E/0

Booster

Ares I booster

Short burn booster

Composite/HTPB

EDS

Metallic EDS

Composite EDS
24

Table 3. Performance analysis for candidate POD configurations.

Mass (t)

TLI LEO

LCCR

PA-C1

PA-C2

PA-C3A

PA-C3D

PA-C3B

PA-C3C

PA-C3E

PA-C3F

PA-C3G

PA-C3H

71

67

70

55.22

64.80

59.46

58.96

68.86

70.23

73.72

64.19

188

181

198

161.3

180.3

170.4

169.5

188.8

191.8

199.06

180.8

PA-C3 H

Five RS-68B

Composite Case/

HTPB Booster 

Metallic EDS

PA-C3 A

Five RS-68B

Ares I Booster 

Metallic EDS

PA-C3 B

Five RS-68B-E/0

Ares I Booster 

Metallic EDS

PA-C3 C

Five RS-68B

Short-Burn Booster 

Metallic EDS

PA-C3 D

Five RS-68B-E/0

Short-Burn Booster 

68.9 t Composite EDS

188.8 t

64.8 t

180.3 t

59 t

169.5 t

59.5 t

170.4 t

55.2 t

161.3 t

64.2 t

180.8 t

33-ft. Diameter OML

2 SRBs

1 J-2X

73.7 t

199.1 t

70.2 t

191.8 t

PA-C3 E (PA-C2-like)

Six RS-68B-E/0

Short-Burn Booster 

Composite EDS

PA-C3 G

Six RS-68B-E/0

Composite Case/ 

HTPB Booster

Composite EDS

PA-C3 F

Five RS-68B-E/0

Composite Case/ 

HTPB Booster

Composite EDS

(Quickest Development)

(Baseline)

Performance assessed for each 

vehicle (total mass after EDS jettison).

LEO injected mass (estimates that 

include EDS mass but vehicles not 

optimized for LEO trajectory).

Figure 12. Family of vehicles POD configuration for PA-C3′′.
25

Ogive Shroud

EDS

J2-X

CS and Five RS-68B

Interstage

Loller Skirt

Two Five-Segment Boosters

3.4 m

(11.3 ft)

25.4 m

(83.3 ft)

35.3 m

(115.8 ft)

48.9 m

(160.3 ft)

61.2 m

(200.8 ft)

120.5 m

(395.4 ft)

22 m

(72.1 ft)

23.5 m

(77 ft)

117.1 m

(384.1 ft)

71.6 m

(235 ft)

59.3 m

(194.5 ft)

Figure 13. PA-C3D candidate POD vehicle (PA-C3′).

4.2 Phase A-Cycle 3D—Description of the Candidate Point-of-Departure Vehicle

This section describes the physical features and performance capability of the integrated 

vehicle and the individual elements.

4.2.1 Ares V Candidate Point-of-Departure Physical Descriptions

The Ares V LV concept PA-C3D is a two-and-a-half stage LV that delivers the Altair lunar 

lander and the EDS to LEO for rendezvous with the Orion crew vehicle system. The major components comprising the PA-C3D concept are the five-segment, PBAN-propellant SRBs, the LOX/LH2

CS, the LOX/LH2 EDS, and the tangent ogive payload shroud. Figure 13 highlights the elements 

and major features.

For the purposes of PA-C3′′, the dimensions for the PA-C3B, C, E, F, and G vehicles remain 

the same as the respective vehicles from which they are derived (i.e., PA-C3A and PA-C3D). 

PA-C3A was modified during this cycle to share a common booster with the PA-C3D vehicle and, 

as such, the vehicles all have the same overall length. However, it should be noted that the CS 

length for PA-C3A is still shorter than PA-C3D because the RS-68B-E/0 engine nozzle length 

differs from the RS-68B engine nozzle by 0.34 m (1.1 ft). Also, configurations that include a composite booster (i.e., PA-C3F and PA-C3G) may slightly alter vehicle dimensions and the OML. 

The thrust trace for the boosters is currently being revisited and will ultimately dictate the expansion ratio and duration of the booster burn.
26

4.2.2 Ares V Point-of-Departure Core Stage Booster 

4.2.2.1 Ares V POD Steel Booster Description. The POD booster configuration for the 

Ares V vehicle is designated as the five-segment steel case configuration (fig. 14). The OML of the 

five-segment steel case configuration is similar to the Ares I SRB, which is based on the foursegment Shuttle SRB. The propellant is the heritage space transportation system (STS)/Ares I 

PBAN propellant, with a propellant grain design that is optimized for use on the Ares V. Furthermore, an inert spacer (steel case only, no propellant) is used at the forward end of the booster 

to permit a longer CS with additional CS propellant. 

Nose Cap (Pilot and drogue parachutes)

Frustum (Three Main Parachutes)

Forward Skirt

Spacer

Igniter/Safe and Arm (S&A)

Forward Segment

Forward-Center Segment

Center-Center Segment

Aft-Center Segment

Aft Segment

Aft Skirt

ET Attach Ring

Systems Tunnel

Avionics

Forward 

Separation 

Motors

Nozzle

Aft Separation 

Motors

Avionics

Figure 14. Expanded view of PA-C3D booster concept.

The Ares V booster configuration uses heritage steel case cylinders and domes. The aft skirt, 

forward skirt, frustum, and nose cone are Shuttle SRB heritage configurations. The CS-to-booster 

attach cylinder of the aft segment and the attach ring/struts are Shuttle SRB heritage. The thrust 

vector control (TVC) system is Shuttle SRB heritage. Boosters would be separated similar to the 

Shuttle with STS/Ares I forward and aft booster separation motors (BSMs).

The Ares V booster is planned for recovery to refurbish components similar to the Shuttle 

SRB. The parachutes and all equipment required for reentry, splashdown, and recovery are included 

in the configuration. 
27

Nose Cone

New 150-ft Diameter 

Parachutes

Same Aft Skirt and 

TVC as Shuttle

Modern Electronics

Same Propellant as Shuttle (PBAN)— 

Optimized for Ares Application

Same Cases and Joints as Shuttle

BDM

Wide Throat Nozzle

Figure 15. Ares V POD booster: Five-segment, PBAN propellant, steel case.

4.2.2.2 Steel Booster Evolution. The initial five-segment steel case booster configuration for 

the Ares V vehicle used the identical motor design (five-segment RSRMV) as defined for the Ares I 

booster. The aft skirt, forward skirt, frustum, nose cone, and other hardware were used as defined 

on Shuttle SRB heritage configurations.

The subsequent booster configuration started with the Ares I booster design. The grain 

and nozzle designs were reoptimized to assist in satisfying the payload performance requirements 

for Ares V. The Ares V vehicle needs substantially more impulse during the atmospheric phase of 

flight than the CS and the Ares I booster can provide. More impulse equates to higher thrust traces 

versus time and larger propellant weight capacities.

Initially, the Ares V vehicle CS configuration was lengthened to provide more impulse by 

adding an empty (no propellant) one-half segment to the booster design. This length change to 

the booster permitted the CS to be lengthened to add more propellant. The second iteration added 

propellant to the empty one-half segment and, with the addition of the longer CS tanks, provided 

a substantial additional impulse during the first phase of ascent. Thus, the LCCR configuration 

was composed of five normal-sized booster segments and one-half segment. The last iteration 

leading to the current five-booster segment configuration was driven by the results of the CSE and 

booster combination trade studies performed. The five-segment booster is utilized on the PA-C3D 

POD to allow for commonality with the Ares I, but the propellant grain design is optimized for use 

on the Ares V. An optimized-length inert spacer replaced the empty one-half segment inert space to 

allow for a 71.6-m (235-ft) CS length, as seen in figure 16. The Ares V booster nozzle configuration 

has an expansion ratio of 7.22, which is currently the same as the Ares I for lunar missions.
28

Additional 

Spacer Length to 

Keep Nozzle 

Exits Coplanar

Standard 

Forward Skirt for 

Five-Segment 

Booster

PA-C3 “A” CS & CSB PA-C3 “D” CS & CSB

Spacer

Five Fueled 

Segments

RS-68B RS-68B-E/0

CS/CSB 

Interface

Figure 16. CS and booster sizing.

4.2.3 Ares V Point-of-Departure Core Stage 

The Ares V CS is a liquid propulsion element that provides thrust during the first 5 min 

of powered ascent. The CS utilizes a LOX/LH2 oxidizer/fuel combination that leverages much of 

the Space Shuttle ET design. This element is 71.6 m (235 ft) long, has a constant OML diameter 

of 10 m (33 ft), and consists of both Al alloy dry structures and Al-Li 2195 propellant tanks. The 

subsystems that make up the CS are the main propulsion system (MPS), aft skirt/thrust structure, 

LH2 tank, intertank, LOX tank and forward skirt, avionics and power, purge, thermal protection 

systems (TPSs), range safety, and development flight instrumentation for early flight vehicles 

(figs. 17 and 18). The CS has a dry mass of 149.7 t (330,108 lbm) and can carry a propellant mass 

of 1,520.2 t (3,351.7 klbm) of useable impulse propellant when fully loaded for ascent. At CS 

burnout, this element has a mass of 170.4 t (375.7 klbm).

To support the CS’s five RS-68 engines, its MPS system consists of systems for fill/drain, 

LOX/LH2 propellant measurement and feed lines, tank pressurization, engine prechill, propellant 

vales, helium purge, pogo suppression, and hydraulics. An option is being considered to create 

a new RS-68B-E/0 that outputs 3,603 kN (810,000 lbf) of thrust operating at 108% in vacuum; it 

also provides an Isp of 423 s. This engine also dumps its gas generator exhaust into a new regeneratively cooled nozzle.

The CS aft skirt is an Al cylindrical structure that joins the MPS/thrust structure 

to the LH2 tank.
29

• CS With Five RS-68Bs

 • An Option is Being Worked With Five RS-68B-E/0s

• Propellant Per Max VAB Height

Figure 17. Ares V CS element view.

Aft Skirt and Fairings

CSE and MPS

CSE Heat 

Shields LH2 Tank 

Aft Dome 

and Barrel

Thrust Structure

Figure 18. CS aft end expanded view.

The LH2 tank is 44.5 m (146.1 ft) long and is constructed of Al-Li barrels and domes 

(end-caps on domes may be Al 2219). Pressure relief valves are located on both forward and 

aft domes to assist in propellant settling. A sump at the bottom of the aft dome drains the fuel 

through a single feed line that goes to the MPS LH2 manifold.

The intertank physically separates the CS LH2 and LOX tanks and provides a physical 

interface to the solid boosters. Built with Al alloy 2090, the intertank carries the weight of the 

structures above and provides the structural interface for the SRB forward attach points. There are 

two attach fittings on opposite sides of the intertank that interface with the SRB forward attach 

mechanisms; an internal crossbeam joins these two attach fittings and provides structural support 

for the high loads induced by the SRBs.
30

The LOX tank is 17.1 m (56.1 ft) long and is constructed of Al-Li barrels and domes. Pressure 

relief valves are located on both the forward and aft domes to assist in propellant pressurization. 

A sump at the bottom of the aft dome drains the oxidizer through two feed lines that go to the two 

aft LOX manifolds.

The CS forward skirt is the forward-most component on the CS and is attached to the aft 

end of the interstage. This component houses the CS avionics used to monitor and control the 

CSEs and both SRBs and to communicate with the avionics system on the EDS instrument unit. 

The forward skirt also supports CS separation. Retrorockets may be evenly distributed around the 

circumference of the forward skirt to support a clean separation from the EDS; alternate locations 

for these separation motors are on the aft compartment or intertank.

4.2.4 Ares V Point-of-Departure Earth Departure Stage

The PA-C3D EDS is an upper stage propulsion element that provides the sole source of 

thrust once the solid boosters and CS have separated from the EDS-Altair stack. The EDS has 

a 10-m (33-ft) outer diameter, measures approximately 23.5 m (77 ft) long, and is comprised of 

composite (graphite epoxy IM7) dry structures and Al-Li 2195 propellant tanks. The primary 

functions of the EDS are to insert the EDS-Altair stack into LEO, provide resources as needed 

to Altair through the launch phase, perform loiter operations for up to 4 days in LEO, dock with 

Orion, and perform the TLI burn for the EDS-Altair-Orion stack. The EDS is pressurized to 

condition the propellant and restart the J-2X engine for the TLI burn. The primary EDS structures 

are the interstage, loiter skirt, EDS MPS, aft skirt/thrust structure, LOX tank, intertank, LH2 tank, 

and forward skirt. See figure 19 for a view of the major EDS components and design. The EDS has 

a dry mass of 36.1 t (79.5 klbm) including the loiter equipment and carries 240.4 t (530.1 klbm) of 

useable impulse propellant mass, of which 143.7 t (316.8 klbm) is burned during ascent. At EDS 

burnout, this element has a mass of 25.4 t (56.1 klbm).

The interstage is a composite cylindrical structure that interfaces with the CS. Together with 

the loiter skirt, the interstage houses the J-2X engine and its extended nozzle. While segmented to 

the EDS, the interstage remains attached to the CS throughout launch. A separation ring at the 

forward end is activated during CS separation, allowing the CS and interstage to separate together 

from the EDS-Altair stack.

The loiter skirt is a cylindrical structure that encompasses the J-2X engine. It is located 

between the interstage and aft skirt/thrust structure. This composite structure contains fuel cells/

batteries and other subsystem resources to maintain the EDS and Altair during its loiter operations 

in LEO. The PA-C3D POD concept assumes that the loiter skirt is jettisoned just prior to the TLI 

maneuver.

The EDS MPS provides the needed propulsion for the LEO insertion and TLI insertion 

burns. The MPS consists of the J-2X engine and the associated feed system, which interfaces with 

the EDS LOX and LH2 tanks.
31

Payload Adapter (owned by Altair)

LH2 Tank

Intertank

Aft Skirt

EDS J-2X Engine

Loiter Skirt w/Thermal 

Radiators

LOX Tank

Forward Skirt

Interstage

Usable Propellant: 251.9 t (555.2 klbm)

Dry Mass: 24.2 t (53.5 klbm)

Burnout Mass: 26.6 t (58.7 klbm)

Number of Engines: 1

Engine Type: J-2X

• Al-Li propellant tanks.

• Composite dry structure.

• 10-m (33-ft) outer diameter.

• Derived from Ares I upper stage.

• 4-day on-orbit loiter capability prior to TLI.

• Maintains Orion/Altair/EDS stack attitude in LEO prior to TLI burn.

• EDS provides 1.5 kW of power to Altair from launch to TLI.

Figure 19. Ares V EDS element view.

The aft skirt and thrust structure provide the structural interface between the J-2X engine 

and the EDS. The aft skirt is a short, tapered, conical section constructed from composite material 

and accommodates the small diameter of the EDS LOX tank. The thrust structure provides the 

attach points for rigidly mounting the J-2X engine. 

The LOX tank is constructed from aft and forward Al-Li domes, which together measure 

6 m (18 ft) long and 8 m (26 ft) in diameter. The LOX tank can hold up to 206.7 t (455,800 lbm) 

of oxidizer. The tank interior contains slosh baffles and a vortex baffle at the outlet. Unlike the CS 

propellant tanks, the EDS LOX tank must manage its cryogenic propellant to eliminate undesirable propellant loss during the 4-day loiter period. To this end, cryogenic fluid management (CFM) 

technologies will be incorporated into the tank to preserve propellant.

The EDS intertank separates the LOX and LH2 propellant tanks. It is manufactured from 

composite materials and tapers from the 10-m EDS outer diameter (at the LH2 tank) down to the 

diameter of the LOX tank.

The LH2 tank aft skirt is a short-barrel section that joins the intertank to the LH2 tank. 

It measures approximately 0.61 m (2 ft) long and is made from composite material.

The LH2 tank is comprised of one barrel section and two domes, which together measure 

approximately 10 m (33 ft) in length. The LH2 tank can hold 42.1 t (92,800 lbm) of LH2 fuel. The 

tank is an Al-Li metallic structure. The tank interior also leverages CFM technologies to prevent 

propellant boiloff during the 4-day loiter period. Micrometeoroids and orbital debris (MMOD) 
32

shielding is integrated into the external side of the barrel section to protect against the debris 

environment during loiter.

In the POD, the EDS forward skirt houses the avionics system that provides primary data 

and command and control for the Ares V integrated vehicle throughout all phases of flight. The 

forward skirt provides the primary interface to the payload shroud and adapter.

4.2.5 Ares V Point-of-Departure Payload Shroud 

The PA-C3D shroud is a composite structure consisting of a cylindrical barrel section and 

a nose cone. The shroud measures 22 m (72 ft) in length and 10 m (33 ft) in diameter across the 

barrel section. The POD vehicle exhibits a tangent ogive nosecone with a rounded tip. This baseline 

configuration features a quad-petal design for effective shroud separation, as seen in figure 20. The 

material construction consists of a composite sandwich with an Al honeycomb core and a painted 

cork TPS bonded to the outside. The shroud weighs a total of 6.7 t (14,839 lbm). This includes the 

structure, TPS, and separation mechanisms, of which approximately 6.5 t (14,268 lbm) is jettisoned 

during the ETO ascent.

• Composite sandwich construction (carbon-epoxy face 

sheets. Al honeycomb core).

• Painted cork TPS bonded to outer face sheet with RTV.

• Payload access ports for maintenance, payload 

consumables and environmental control (while on 

ground).

Thrust Rail Vertical Separation System 

Payload Umbilical Separation

Point of Departure (Ogive)

Frangible Joint 

Horizontal Separation

Quad Sector Design

POD Geometry: Ogive

Design: Quad Sector

Barrel Diameter: 10 m (33 ft)

Barrel Length: 9.7 m (32 ft)

Total Length: 22 m (72 ft)

Figure 20. Ares V shroud element view.

Preliminary sizing of the TPS, structures, and subsystems was performed during previous 

assessments. This PA-C3 trade study considers all identified driving factors including mass, payload 

volume, performance, cost, manufacturing, transportation, etc., with a focus on acoustic impacts 

and assessment of shroud shape refinement.
33

4.2.6 Ares V Point-of-Departure Avionics and Software

The Ares V avionics and software maintain control over the entire vehicle stack throughout 

the entire launch profile. The avionics control all aspects of flight including fault recovery, abort 

conditions, and any data interfaces to Altair or Orion. Avionics provides three-axis attitude control for CS flight phases. Roll control stability is maintained by the SRBs and CSEs while attached. 

Figure 21 identifies the Ares V avionics and software.

1. Flight and Command and Telemetry Computers

2. Data Acquisition and Control Unit (DACU)

3. Operations Recorder

4. Radio Freq Communication System (RFCS)

5. Flight Safety System (FSS)

6. Guidance, Navigation, and Control Units (GN&C)

7. Thrust Vector Control Electronics (TVCE)

8. Operational Flight Instrumentation (OFI)

9. Imaging System (IS)

10. Electrical Power System (EPS)

11. DFIS

12. EDS Separation Electronics

1. Data Acquisition and Record Unit (DARU)

2. Booster Control and Power Distribution Unit (BCPDU)

3. Actuator Control Unit (ACU)

4. Ignition and Separation Controller (ISC)

5. FSS

6. GN&C

7. OFI

8. IS

9. Recovery System (RS)

10. EPS

11. DFIS

1. CS Control Units (CSCU)

2. DACU

3. FSS

4. GN&C

5. OFI

6. TVCE

7. IS

8. EPS

9. DFIS

10. CS Separation Electronics

1. RS-68 Analog Engine Controllers

1. J-2X Engine Controller

1. Development Flight Instrumentation System (DFIS) 

Including Imaging System

2. Operational Flight Instrumentation (OFI)

3. Shroud Separation Electronics EDS

Altair Lunar Lander

Payload Shroud

J-2X Engine

Two SRBs

CS

Six RS-68B Engines

Figure 21. Ares V avionics and software.

4.3 Ares V Point-of-Departure Performance Capability

The performance capability was optimized for the PA-C3D vehicle by sizing the EDS while 

maintaining the extended CS size. The CS remained fixed because the structural attach point of the 

five-segment booster (with the inert spacer) constrained the position of the intertank thrust beam, 

thereby constraining the size of the propellant tanks. The mission profile is to launch Altair and 

a partially loaded EDS into a 240-km (130-nmi) circular orbit for rendezvous with Orion, which has 

been launched on Ares I. After docking, the integrated stack (i.e., Orion, Altair, and EDS) assumes 

to leave from a 185-km (100-nmi) circular orbit for TLI. The J-2X engine operates for this burn at 

the lower propellant mixture ratio of 4.5, which lowers the thrust from 1.31 MN (294 klbf) used on 

the ETO portion of the mission to 1.06 MN (238 klbf) for TLI, while the Isp is currently assumed to 

stay approximately constant at 448 s. The TLI payload capability for the lunar sortie DRM using 

the PA-C3D concept for a 1.5-launch scenario (rendezvous in LEO with Orion launched on Ares I) 

is 64.8 t (142,800 lbm).
34

5. ARES V MARS EXTENSIBILITY

5.1 Mars Transfer Vehicle Options: Nuclear Thermal Propulsion, Chemical, 

and Earth Departure Stage as Mars Transfer Vehicle Stage

The reference HLLV (Ares V) for NASA’s CxP served as the POD for the Mars Design 

Reference Architecture 5.0 (DRA 5.0).4 The assessment of the Ares V vehicle as it pertains to Mars 

missions remained locked to the architecture presented in the Mars DRA 5.0, consisting of two 

cargo MTVs that are predeployed (utilizing aerocapture for Mars orbit insertion (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 that were intended to reduce 

the number of flights to a reasonable and sustainable level. These options included nuclear thermal 

propulsion (NTP) and chemical propulsion. The Ares V LV 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 the Mars DRA 5.0 

and a competitive chemical option consisted of a nine-launch campaign). 

The chemical options assessed can be broadly categorized as those that use the EDS for 

TMI and those that do not use the EDS (propulsion modules are delivered to LEO). Figures 22 

through 24 illustrate the launch sequencing for the NTP option, chemical option (as presented 

in the Mars DRA 5.0), and EDS-as-MTV-stage option, respectively. 

The Ares V contributions to a Mars campaign are highlighted in this work with the goal of 

understanding the Ares V capability within the trade space. Using the EDS as the in-space transportation element was more fully characterized and was of special interest during the assessment. 

This work consisted of characterizing LEO performance as a function of insertion and 

final assembly altitudes, number of launches required for the campaign, and EDS functionality 

evaluated against the lunar configuration (e.g., as additional loiter duration, on-orbit capabilities 

required, and payload operations).
35

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

CLV 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

Cargo SHAB MTV (IMLEO 236 t)

Cargo DAV MTV (IMLEO 236 t)

Crewed MTV (IMLEO 326 t)

Figure 22. NTP launch sequence and payload mass requirements.
36

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 

requiredfor 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

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

Payload: Surface Lander

Mass: 108 t

Payload: TMI Module

Mass: 108 t

CLV Required

Cargo SHAB MTV (IMLEO 325 t)

Cargo DAV MTV (IMLEO 325 t)

Crewed MTV (IMLEO 540 t)

Figure 23. Chemical propulsion module option.
37

DRA5.0 TMI Prop. Module

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

CLV 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

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

Cargo SHAB MTV (IMLEO 325 t)

Cargo DAV MTV (IMLEO 325 t)

Assuming ≈48% and ≈77% Filled EDSs

Assuming ≈77% Filled EDSs

Figure 24. Propellant delivery and transfer option (EDS-as-MTV-stage).

5.2 In-Space Earth Departure Stage

The EDS Mars Extensibility team assessed the NTP, chemical, and EDS-as-MTV-stage 

campaigns for Mars exploration. There are also many benefits associated with employing an EDS 

as the TMI module. 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. Figure 25 illustrates 

upgrades made to the lunar EDS if it were to serve as a multiarchitecture stage. 
38

(a) (b)

Figure 25. EDS concept for (a) lunar and (b) Mars architecture.

The EDS-as-MTV-stage has several challenges that are specific to this campaign. These are 

summarized as follows:

• J-2X not available for circularization burn because the second burn must be TMI.

• Cooperative debris avoidance of multiple EDS modules.

• EDS-to-EDS interfaces, including structural and communication.

• Disposal of stages providing the second part of TMI.

• Propulsive settling for multiple EDS modules.

• Propellant transfer between stages.

• Propellant transfer between drop tanks and main tanks.

• 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).

5.3 Docking Adaptor and Propellant Transfer

As a supplemental part of the Mars extensibility study conducted during PA-C3, the functionality and design of an automated rendezvous and docking system was identified and assessed. 

Using the chemical scavenger nine-launch option (fig. 24) as a reference baseline for the analysis, 
39

the objective was to add detail to the docking adaptor and propellant transfer required for the 

crewed MTV concept. Despite several docking mechanism options, the representative docking 

mechanism baselined for this effort was referred to as the N-point docking mechanism and is 

illustrated in figure 26.

(a) (b) (c)

Figure 26. N-point docking mechanism with (a) and (c) latching claw, and (b) probes.

This mechanism allows for docking, undocking, and redocking capabilities, providing dV

upon undocking if required. Telemetry supports precision capture and there may be several points 

for docking for redundancy or fail-safe contingency planning.

A timeline was required in order to assess the functional analysis of the docking sequence. 

The docking adaptor sequence of events was carried during the effort as follows:

(1) Phasing trajectory.

(2) Far-range rendezvous.

(3) Close-range rendezvous.

(3) Mating.

(4) Propellant transfer.

(5) Demating and separation.

(6) Disposal. 

Once this sequence was defined, a list of functions was generated then vetted among team 

members. The functions were established with the assumptions that the MTV integrated stack 

undocked in orbit from the stages involved with transferring propellant and that the propellant 

was transferred using centrifugal acceleration. A sample matrix of operations and functions is 

given in table 4 and shows the performers, functions, needs, and suppliers necessary for the phasing 

trajectory event during in-space docking. 
40

Table 4. Phasing trajectory sample functions.

Event, Action Performer Functions Needs Supplier

Phasing Trajectory

Reduction of orbital phase angle 

between rendezvous vehicle and 

target spacecraft

Tanker rendezvous vehicle In-space movement Propulsion thrust, vectorable, magnitude, varyingOnboard reaction control 

system or Orbital Manuevering System (OMS)

Tanker rendezvous vehicle In-space navigation Gyroscope, GPS, radar, 

sensors, star trackers

Avionics (flight computer, 

software)

Tanker rendezvous vehicle In-space communications Telemetry, data, commandsAvionics, ground data 

system, space network

Tanker rendezvous vehicle In-transit cryo propellant 

management

Delivery proper prop load 

to target spacecraft

Perform navigation updates Tanker rendezvous vehicle Communications Telemetry, data, 

commands

Avionics, ground data 

system, space network

Tanker rendezvous vehicle Computations State vectors etc. Avionics (flight computer, 

software)

EDS perform alignments EDS target spacecraft

EDS begin free drift EDS target spacecraft

MTV undocks, assumes its own 

attitude control, goes into passive 

mode.

MTV payload (Go to payload OPS)

EDS assumes its own attitude control, performs backaway manuever.EDS target spacecraft

EDS perform station keeping 

manuevers

EDS target spacecraft Propulsion Propulsion thrust, vectorable, magnitude, varyingOnboard reaction control 

system

EDS go to docking attitude EDS target spacecraft Propulsion Propulsion thrust, vectorable, magnitude, varyingOnboard reaction control 

system

EDS target spacecraft Navigation Gyroscope, GPS, radar Avionics (flight computer, 

software)

EDS target spacecraft Communication Telemetry, data, commandsAvionics, ground data 

system, space network

The owners of these functions were then identified. Despite focus on Ares V vehicle components, this functional analysis has laid the groundwork for future work on HLLVs, regardless 

of specific stages such as the EDS used here. 

Propellant transfer was also investigated for the chemical scavenger nine-launch option. 

A functional analysis was performed that included a feasibility study recommending a nose-to-tail 

configuration for the EDS in-space positioning during fuel transfer. This analysis identified the 

trade space for a fuel transfer scenario; however, higher fidelity analysis is necessary to assess the 

best stage-positioning choice, considering length of feed lines, MPS plumbing location and accessibility, etc. Six refueling positions were considered: Nose-to-tail, nose-to-nose, side-to-side (in 

parallel), tail-to-tail, side mount (nose-to-side perpendicular), and side mount (tail-to-side perpendicular). A propellant transfer trade tree was generated that incorporated such stage orientations, 

propellant settling types, and propellant transfer types.
41

A preliminary mass estimate of the docking adaptor was given with support from the 

design analysis team. The requested mass of the docking adaptor included the docking mechanism 

(N-point docking mechanism), truss structures, and any additional avionics. The computer aided 

design (CAD) models of the integrated vehicle (figs. 27–29) and the N-point docking mechanism 

(fig. 30) allowed for a structural loads assessment that preceded a stress analysis. The stress analysis 

performed for the truss structures assumed 7075-T6 Al, 4- × 0.25-in round tubing and heights of 

approximately 5.79 m (19 ft).

The thinnest skins were assumed for these analyses and only axial loads were assessed. Primary and secondary structure masses were assumed, but no mass was built-in to account for fault 

tolerances or uncertainty factors for structures or mechanisms. Forward work would include additional margin, material trades (composites), moment assessment, and finite element models (FEM) 

of the integrated MTV stack.

TransHab

Trans-Earth Injection (TEI)

TEI Truss Structure (IS3)

TMI1

TMI1 Truss Structure (IS1)

MOI

MOI Structure (IS2)

Figure 27. EDS-as-MTV model for loads and stress analysis.
42

Figure 28. Chemical propulsion model for loads and stress analysis.

Figure 29. Dedicated in-space chemical propulsion stage and truss detail.
43

Figure 30. Docking adaptor mechanism and ring structure.

5.4 Lunar and Mars Synergy

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 cost and complexity. As the Mars DRA 5.0 specifies, this synergy is 

defined by the subsystem technologies (i.e., trajectory and mission analyses, propulsion, structures, 

thermal, power, and avionics), space transportation, elements, and the use of common ETO LVs.

If systems are independently developed for Moon and Mars architectures, multiple obstacles may need to be addressed (e.g., 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 postlunar missions. Even with 

a renewed investment in a Mars mission, there would be no heritage technology or experience 

with systems designed for independent Moon and Mars architectures.

In order to develop a sustainable architecture and resulting space program, technologies 

developed for lunar missions should carry considerable commonalities to those needed for Mars 

missions. Several aspects of the architecture offer potential for such synergy. This includes 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. 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 real-world applications. There is experience to be gained in routine manufacturing and system operation that would 

decrease risk and improve reliability. Technologies that would be developed for lunar and Mars 

DRMs may also prove compatible with other alternative missions, such as near-Earth orbit (NEO) 

destinations or missions that extend even beyond Mars. 
44
45

REFERENCES

1. “Ares V Operational Concepts Document,” CxP 72336, Marshall Space Flight Center, AL, 

2010.

2. “Constellation Architecture Requirements Document,” CxP 70000, Marshall Space Flight 

Center, December 2006.

3. “NASA’s Exploration Systems Architecture Study—Final Report,” NASA/TM—2005–214062, 

November 2005.

4. “Human Exploration of Mars Design Reference Architecture 5.0 Addendum,” B.G. Drake (ed.), 

NASA/SP—2009–566, Johnson Space Center, Houston, TX, July 2009.
46

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Executive Summary of Ares V: Lunar Capabilities Concept Review 

Through Phase A-Cycle 3 

J.B. Holladay, K.E. Baggett,* and S.M. Feldman*

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 Ares Projects Office

*Zero Point Frontiers Corporation, Huntsville, Alabama

M–1307

Technical Memorandum

NASA/TM—2011–216458

Executive Summary, PA-C3, Ares V, trade space, point-of-departure, Mars extensibility

01–02–2011

UU 58

NASA

U U U

This Technical Memorandum (TM) was generated as an overall Ares V summary from the Lunar Capabilities Concept Review (LCCR) 

through Phase A-Cycle 3 (PA-C3) with the intent that it may be coupled with separately published appendices for a more detailed, integrated narrative. The Ares V has evolved from the initial point of departure (POD) 51.00.48 LCCR configuration to the current candidate POD, PA-C3D, and the family of vehicles concept that contains vehicles PA-C3A through H. The logical progression from concept 

to POD vehicles is summarized in this TM and captures the trade space and performance of each. The family-of-vehicles concept was 

assessed during PA-C3 and offered flexibility in the path forward with the ability to add options deemed appropriate. A description 

of each trade space is given in addition to a summary of each Ares V element. The Ares V contributions to a Mars campaign are also 

highlighted with the goal of introducing Ares V capabilities within the trade space. The assessment of the Ares V vehicle as it pertains 

to Mars missions remained locked to the architecture presented in Mars Design Reference Authorization 5.0 using the PA-C3D vehicle 

configuration to assess Mars transfer vehicle options, in-space EDS capabilities, docking adaptor and propellant transfer assessments, 

and lunar and Mars synergistic potential. 

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