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i

NASA/TM—2012–217467

Cryogenic Propellant Storage and Transfer

Technology Demonstration: Prephase A 

Government Point-of-Departure Concept Study

J.A. Mulqueen, B.M. Addona, D.A. Gwaltney, K.A. Holt, R.C. Hopkins, J.A. Matus,

P.S. McRight, C.G. Popp, S.G. Sutherlin, and H.D. Thomas 

Marshall Space Flight Center, Huntsville, Alabama

M.F. Baysinger and C.D. Maples

Qualis Corporation, Huntsville, Alabama

P.D. Capizzo

Raytheon, Huntsville, Alabama

L.L. Fabisinski 

International Space Systems, Huntsville, Alabama

L.S. Hornsby

Jacobs Engineering, Huntsville, Alabama

T.K. Percy

Science Applications International Corporation, Huntsville, Alabama

S.D. Thomas

Dynetics, Huntsville, Alabama

September 2012

National Aeronautics and

Space Administration

Marshall Space Flight Center • Huntsville, 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/login/wt/>

TRADEMARKS

Trade names and trademarks are used in this report for identification only. This usage does not constitute an official 

endorsement, either expressed or implied, by the National Aeronautics and Space Administration.
iii

TABLE OF CONTENTS

1. STUDY OVERVIEW .......................................................................................................... 1

1.1 Study Purpose ............................................................................................................... 1

1.2 Study Timeline .............................................................................................................. 1

1.3 Study Topics ................................................................................................................. 2

1.4 Needs, Goals, and Objectives ........................................................................................ 3

1.5 Study Results ................................................................................................................ 4

2. STUDY TOPIC REVIEW .................................................................................................. 8

2.1 Mission Justification ..................................................................................................... 8

2.2 Mission Concept ........................................................................................................... 12

2.3 Demonstration Flight System ....................................................................................... 26

2.4 Study Accomplishments and Forward Work ................................................................. 85

REFERENCES ....................................................................................................................... 87
iv
v

LIST OF FIGURES

 1. Study timeline .......................................................................................................... 2

 2. Study teams .............................................................................................................. 3

 3. Study task flow ......................................................................................................... 5

 4. CPST Government POD mission concept ................................................................ 7

 5. Configuration used in DAS to determine the area/mass ratio ................................... 13

 6. CPST Government POD demonstration timeline .................................................... 19

 7. CPST Government POD propellant loss profile ....................................................... 20

 8. Thirty-day transfer demonstration series timeline .................................................... 21

 9. Prelaunch to spacecraft separation communications architecture ............................. 24

10. Spacecraft-controlled deorbit communications architecture ..................................... 25

11. Ground station S-band coverage .............................................................................. 26

12. CPST primary sections ............................................................................................. 27

13. Size and dimensions ................................................................................................. 28

14. Taurus II launch configuration ................................................................................. 28

15. On-orbit configuration ............................................................................................. 29

16. Helium pressurized subsystem schematic ................................................................. 33

17. Direct autogenous schematic .................................................................................... 35

18 Stored autogenous schematic ................................................................................... 36

19. Vent subsystem schematic—storage tank side shown only......................................... 37

20. Vent probe—external view ........................................................................................ 38
vi

21. Vent probe (payload on launch vehicle) cross section ............................................... 38

22. Vent probe in nonpropulsive vent use—shown in disengaged position ..................... 39

23. LH2 storage subsystem schematic ............................................................................. 40

24. LH2 storage tank assembly diagram ......................................................................... 41

25. LH2 transfer subsystem schematic ............................................................................ 43

26. Alternative LH2 transfer system schematic—cryogenic transfer

subsystem (three-way valve configuration; simplified rep.) ........................................ 44

27. CPST fluid system schematic .................................................................................... 46

28. Assumed propellant inventory over the mission timeline .......................................... 49

29. Assumed storage tank ullage percentage over the mission timeline ........................... 50

30. Helium pressure over the mission timeline ................................................................ 51

31. CPST payload 5-ft storage tank end dome—cover geometry .................................... 53

32. CPST payload tank top dome cover conceptual layout ............................................ 54

33. CPST payload storage tank bottom dome cover layout ............................................ 54

34. CPST payload configuration .................................................................................... 56

35. CPST finite element model ....................................................................................... 57

36. CPST orbital plot ..................................................................................................... 60

37. CFM/thermal system description ............................................................................. 61

38. CPST temperature results ......................................................................................... 64

39. A typical experiment data acquisition settled demo timeline 

(306 min) .................................................................................................................. 70

40. Data acquired and stored after downlink with X-band at 7 Mbps, 

four 8-min links per day ........................................................................................... 72

LIST OF FIGURES (Continued)
vii

41. Data acquired and stored after downlink with X-band at 10 Mbps, 

two 8-min links per day ............................................................................................ 72

42. Data acquired and stored after downlink with S-band at 3.5 Mbps, 

ten 8-min links per day ............................................................................................. 73

43. Pathfinder POD payload avionics and power system ................................................ 74

44. Six-month power profile ........................................................................................... 77

45. Propellant transfer power profile .............................................................................. 77

46. Power system schematic ............................................................................................ 78

47. CPST POD configuration ......................................................................................... 80

LIST OF FIGURES (Continued)
viii

LIST OF TABLES

 1. Study topics .............................................................................................................. 3

 2. Primary flight objectives ........................................................................................... 4

 3. CPST mission concept descope options (prioritization based 

on minimizing technical impact on flight test objectives) .......................................... 6

 4. Mission relevance to OCT roadmaps ........................................................................ 8

 5. Mission relevance to top challenges in OCT roadmaps ............................................ 9

 6. Mission relevance to human architecture destinations .............................................. 9

 7. Mission analysis ground rules and assumptions ....................................................... 12

 8. Input parameters for the LTIME satellite lifetime program ...................................... 14

 9. Effect of launch and orbit inclination on orbit lifetime ............................................ 14

10. Resulting area/mass ratios used in the orbit lifetime analysis .................................... 15

11. Orbit lifetime and stationkeeping estimates .............................................................. 16

12. Orbit insertion errors for two example launch vehicles ............................................. 16

13. The effect of launch vehicle insertion errors on orbit lifetime ................................... 17

14. Preliminary mission propellant estimate ................................................................... 18

15. CPST Government POD propellant loss and heating rates ...................................... 20

16. Settled transfer demonstration timeline .................................................................... 22

17. Unsettled transfer demonstration timeline ............................................................... 22

18. S-band coverage summary ........................................................................................ 26

19. CPST payload key design parameters ....................................................................... 31

20. CPST fluid system ground rules and assumptions .................................................... 31
ix

LIST OF TABLES (Continued)

21. Pressure-fed system pressure budget ......................................................................... 47

22. Pump-fed system pressure budget ............................................................................. 47

23. TVS system pressure budget ..................................................................................... 48

24. Configuration options .............................................................................................. 52

25. CPST payload storage tank end dome and internal components .............................. 53

26. Falcon 9 load conditions .......................................................................................... 57

27. Structural analysis mass results ................................................................................ 58

28. Ground rules and assumptions ................................................................................. 59

29. CPST avionics power dissipation .............................................................................. 62

30. CFM/thermal summary ........................................................................................... 63

31. Propellant loss and heating rates .............................................................................. 65

32. CFM/thermal input power ....................................................................................... 65

33. Estimated data storage rates over a 7-day experiment ............................................... 69

34. Data volumes acquired with duty durations: CCSDS overhead, 15%,

and prephase A data margin, 30% ............................................................................ 71

35. Avionics mass summary ........................................................................................... 75

36. Power requirements .................................................................................................. 76

37. Power system mass summary .................................................................................... 79

38. CPST POD mass summary—CPS Pathfinder ........................................................... 81

39. CFM subsystem mass ............................................................................................... 83

40. Spacecraft bus mass summary .................................................................................. 84
x

LIST OF ACRONYMS, SYMBOLS, AND DESIGNATORS

ACS altitude control system

AFC automated fluid coupling

APIC advanced programmable interrupt controller

ARC Ames Research Center

BAA Broad Area Announcement

BAC broad-area cooling

CAD computer automated design

CCAFS Cape Canaveral Air Force Station

C&DH communications and data handling

CDR Critical Design Review

CFM cryogenic fluid management

CFMS cryogenic fluid management system

CFS cryogenic fluid system

COPV composite overwrap pressurant vessel

CPS cryogenic propulsion stage

CPST cryogenic propellant storage and transfer

CV check valve

CY calendar year

DAS Debris Assessment Software

DDT&E design, development, test, and engineering

ECDU experiment computer/data unit

ESEOD equivalent sharp edge orifice diameter

FEM finite element model

GN2 gaseous nitrogen
xi

GN&C guidance, navigation, and control

GR&A ground rules and assumptions

GRC Glenn Research Center

GSFC Goddard Space Flight Center

HEFT Human Exploration Framework Team

HEX heat exchanger

JSC Johnson Space Center

J-T Joule-Thomson

KSC Kennedy Space Center

LAD liquid acquisition device

LEO low-Earth orbit

LH2 liquid hydrogen

LO2 liquid oxygen

LTIME Life TIME (an orbit lifetime tool)

MCR mission concept review

MDP maximum design pressure

MEL master equipment list

MEOP maximum expected operating pressure

MGA mass growth allowance

MLI multilayer insulation

MOC Mission Operations Center

MSFC Marshall Space Flight Center

NEA near Earth asteroid

NGO needs, goals, and objectives

OCT Office of the Chief Technologist

OD outer diameter

LIST OF ACRONYMS, SYMBOLS, AND DESIGNATORS (Continued)
xii

PCSA passive cryogenic storage analyzer 

PDM power distribution module

PDR preliminary design review

POD point of departure

QD quick disconnect

QPSK quadrature phase shift keying

RBO reduced boiloff

RCS reaction control system

RFMG radio frequency mass gauging

RV relief valve

RW reaction wheel

SOFI spray-on foam insulation

STK Satellite Tool Kit

TDRS tracking and data relay satellite

TDRSS tracking and data relay satellite system

TOC Technology Operations Center

TRL Technology Readiness Level

TVS thermodynamic vent system

WBS Work Breakdown Structure

WFF Wallops Flight Facility

V vent valve

VME Versa Module Eurocard (computer bus standard)

ZBO zero boiloff

LIST OF ACRONYMS, SYMBOLS, AND DESIGNATORS (Continued)
xiii

NOMENCLATURE

A area (m2)

Isp specific impulse (s)

T temperature

V velocity

V∞ hyperbolic excess speed

ΔV delta velocity
xiv
1

TECHNICAL MEMORANDUM

CRYOGENIC PROPELLANT STORAGE AND TRANSFER TECHNOLOGY 

DEMONSTRATION: PREPHASE A GOVERNMENT 

POINT-OF-DEPARTURE CONCEPT STUDY

1. STUDY OVERVIEW

1.1 Study Purpose

The primary purpose of this study was to define a point-of-departure (POD) prephase A

mission concept for the cryogenic propellant storage and transfer (CPST) technology demonstration mission to be conducted by the NASA Office of the Chief Technologist (OCT). The mission 

concept includes identification of the cryogenic propellant management technologies to be demonstrated, definition of a representative mission timeline, and definition of a viable flight system 

design concept. The resulting mission concept will serve as a POD for evaluating alternative mission concepts and synthesizing the results of industry-defined mission concepts developed under 

the OCT contracted studies. This mission concept provides a balanced approach between meeting 

mission flight test objectives and project funding constraints. The study objectives are summarized 

as follows:

• Develop a viable mission concept that provides a balanced approach to meeting mission 

objectives within project constraints.

• Satisfy CPST mission needs, goals, and objectives.

• Provide a framework for evaluating alternative CPST mission concepts.

• Provide a framework for technology development and ground test planning.

• Provide a foundation for a synthesized reference concept based on the OCT contracted studies

timeline.

1.2 Study Timeline

The 6-mo Government POD study was conducted from March to October 2011, building 

on the results of several other prephase studies conducted in 2010 as part of the CPST project formulation. The mission concept developed in this study was based on the same study guidelines as 

the OCT contracted studies to ensure the results of this study would be of the same content and 
2

Figure 1. Study timeline. 

format as the industry studies. Figure 1 shows the relationship of the Government POD study and 

the industry Broad Area Announcement (BAA) contracted studies. The results of the Government 

POD study will be synthesized with the results of the contracted studies to define a reference mission concept that will serve as the basis for defining the actual flight system in accordance to the 

CPST project acquisition process.

1.3 Study Topics

The study addressed eight study topics which are listed in table 1. Responsibilities for the 

study topics were spread across four teams from five NASA Centers, totaling over 60 study participants. The responsibilities of each team are shown in figure 2. The first three study topics addressed 

the basis for the technology demonstration mission. The CPST project management and technology teams provided key inputs to describe the need and justification for the proposed on-orbit 

technology demonstration. The technology team identified the demonstration technologies and 

developed the technology maturation plan. The mission concept, including mission orbit analysis 

and mission timeline, was defined for the study design team. The fourth study topic included the 

technical design analysis of the demonstration flight system, including the definition of technology 

components of the cryogenic fluid system (CFS) and the supporting spacecraft subsystem bus. The 

final four topics addressed programmatic considerations for the POD mission concept. The preliminary mission cost estimate included estimates of the design development and testing of the flight 

system and mission launch, and operations costs. The schedule assessment included development, 

testing, and integration of the technology components and flight system. Study topic seven identified potential partnerships and partnership strategies with other government agencies and industry. 

Study topic eight included identification of top-level project risks and an initial assessment of risk 

mitigation strategies.
3

Table 1. Study topics.

Study Topic/Product Area Responsible Team

(1) Mission justification Project analysis

(2) Technology maturation Technology analysis

(3) Mission concept Design analysis

(4) Demonstration flight system Design analysis

(5) Mission cost estimate Project analysis

(6) Project schedule Project analysis

(7) Government and industry partnerships Project analysis

(8) Project risk identification Safety and mission assurance

Figure 2. Study teams.

1.4 Needs, Goals, and Objectives

The POD mission concept defined in this study is based on the CPST project needs and 

goals listed in sections 1.4.1 and 1.4.2. 

1.4.1 Project Needs

There is a need for a cost-effective flight test in order to raise the Technology Readiness 

Levels (TRLs) of cryogenic fluid management (CFM) technologies by demonstrating on-orbit 

capability and processes used for storing and transferring cryogenic fluids (e.g., liquid oxygen 

(LO2) and liquid hydrogen (LH2) in a microgravity environment. 
4

1.4.2 Project Goals

Specifically, these technologies will help NASA satisfy its strategic goals and objectives 

in space exploration and scientific discovery by increasing the performance of current and future 

spacecraft systems. Technology maturity will be raised to a minimum TRL of 6 and a goal of 

TRL 7. The CPST project has defined two primary goals:

(1) Demonstrate long-duration, in-space storage of cryogenic propellants.

(2) Demonstrate in-space transfer of cryogenic propellants.

	 To achieve the CPST project goals, sets of flight test objectives were defined. The prioritization

system of these objectives is based on the perceived values of future technology system stakeholders.

Anticipated stakeholders include the most likely applications for the CPST technology such as in-space

cryogenic propulsion stages (CPSs) and in-space cryogenic propellant depots.

Table 2 is a listing of prioritized flight test objectives.

Table 2. Primary flight objectives. 

Primary Flight Test Objectives

P1 LH2–Cryogenic fluid storage Minimize storage boiloff LH2 in microgravity

P2 LH2–Cryogenic fluid acquisition Demonstrate acquisition and bubble-free flow of LH2 in microgravity

P3 LH2–Cryogenic fluid quantity gauging Demonstrate mass gauging of LH2 in microgravity

P4 LH2–Cryogenic fluid transfer Demonstrate transfer of LH2 in microgravity (settled and unsettled conditons)

P5 LO2–Cryogenic fluid storage Demonstrate zero boiloff storage of LO2 in microgravity

P6 LO2–Cryogenic fluid acquisition Demonstrate acquisition and bubble-free flow of LO2 in microgravity

P7 LO2–Cryogenic fluid quantity gauging Demonstrate mass gauging of LO2 in microgravity

P8 LO2–Cryogenic fluid transfer Demonstrate transfer of LO2 in microgravity (settled and unsettled conditions)

Secondary Flight Test Objectives

S1 Instrumentation–leak detection Demonstrate leak detection capability for LO2 and LH2 in microgravity

S2 Instrumentation–flow measurement Demonstrate flow measurement of LO2 and LH2 in microgavity 

S3 Tank pressurization methods Demonstrate tank pressurization and pressure control approached for LH2 and LO2

in microgravity

1.5 Study Results

	 The study approach followed an iterative process in which the study topics were divided into

subtasks to evaluate mission options and perform design trade studies. Many design iterations were

required due to changes in mission ground rules and constraints during the course of the study. Since

many of the programmatic constraints were not defined at the beginning of the study, the initial approach

was to define an upper and lower cost bound for the mission concept. The study task flow is shown in

figure 3.
5

Mission

Justification

Technology and Test

Mission Concept

Demonstration

Flight System

Needs, Goals 

and Objectives

Technology

Maturation

Mission

Timeline

Launch Vehicle

Options

Ground and Mission

Operations

CFS

Sizing

Flight System

Mission Concept

Mission Concept

Achieves all NGO’s Achieves Subset of NGO’s

• Significant and Partnership

Required

• Total Cost >$250 M

• NASA Funds ~$250 M

• Partner Funds Remainder

• Moderate and Partnership

Required

• Total Cost >$250 M

• NASA Funds ~$250 M

• Partner Funds Remainder

Lower Cost Bound

Thermal/Structural

Modeling

Spacecraft Bus

Options

Project Schedule

Cost Estimate

Partnership

Risk Identification

Figure 3. Study task flow. 

The design iterations also included subsystem design trades and consideration of mission 

descope options to reduce the development and operational costs of the technology demonstrations. The options were prioritized to minimize the impact on the established flight test objectives. 

The evaluation of descope options considered four factors:

(1) Potential reduction in project cost.

(2) Potential reduction in project schedule.

(3) Potential reduction in technology infusion to future programs.

(4) Potential reduction in probability of project success.

The descope options are summarized in table 3. The descope options that were incorporated 

into the final POD mission concept are shown in blue text in the bottom rows in the table. The 

most significant descope that was implemented was the decision to scale back the demonstration to 

LH2 only. This eliminated the four primary flight test objectives associated with LO2. It was determined that by reducing the demonstration to a single fluid, the overall value of the demonstration 

would be maintained since the LH2 demonstrations represent the more challenging technologies.
6

Table 3. CPST mission concept descope options (prioritization based on minimizing

technical on flight test objectives).

Descope Factor

Affected 

Flight Test 

Objectives

Project 

Cost 

Reduction

Project 

Life Cycle 

Schedule

Technology 

Infusion

Programmatic

Risk

Comments

NGO 

Flight Test 

Objectives 

Affected by 

Descope

Potential 

Reduction 

of DDT&E, 

Integration 

and

Operations 

Cost

Potential 

Reduction 

of DDT&E, 

Integration 

and

Operations

Potential 

Reduction of 

Technology 

Infusion 

into Future 

Missions

Potential 

Reduction 

of Project 

Success 

Probability

>$20 M >6 mo High Low

$4-20 M 2-5 mo Medium Medium

Ranking Critieria <$4M <1 mo Low High

Reduce pre-CDR ground test None Increases technology 

development risk

Reduce post-CDR ground test None Increases flight system risk

Eliminate active cooling P 1, 5

Demonstrate LO2 only P 1–4

Reduce mission duration P 1, 5 Reduce from 6 mo to 3 mo

Eliminate propellant transfers P 4, 8; S 2, 3

Reduce pressure-fed settled transfers P 4, 8; S 2, 3

Reduce pressure-fed unsettled transfers P 4, 8; S 2, 3

Reduce pump-fed settled transfers P 4, 8; S 2, 3

Reduce pump-fed unsettled transfers P 4, 8; S 2, 3

Eliminate low-g automated leak detection S 1 Leak detection, not leak rate

Reduce mass gauging experiment P 1-8 Linked to all flight test 

objectives

Reduce video requirements P2, 4, 6, 8 Use nonvideo flow-bubble 

sensor

Reduce number of transfer cycles P 4, 8; S 2, 3 Current POD has two 

transfer demos

Demonstrate LH2 only* P 5–8 LH2 is the preferred single 

fluid demo

Eliminate automated fluid coupling* P 4, 8 Simplifies system

Eliminate CFS redundancy* None Increases flight system risk

Eliminate autogenous tank pressurization* S 3 Reduced number of 

components

*Implemented on CPS-Pathfinder POD concept.

Figure 4 shows the final POD mission concept. The flight system consists of an LH2 technology 

demonstration system and a spacecraft subsystem bus. The technology demonstration system consists 

of the CFM payload and propellant. The spacecraft bus consists of all the subsystem components necessary to control and operate the spacecraft for a 6-mo, on-orbit demonstration mission. The design of the 

spacecraft bus was provided by the Goddard Space Flight Center (GSSF) team. The CPST spacecraft 

was sized so that it could be launched on a Taurus II launch vehicle departing from the NASA Wallops 

Flight Facility (WFF). The preliminary mission cost estimate, which includes development, integration, 

launch, and mission operations for the demonstration mission, was $422 million.
7

GSFC Provided

Spacecraft Bus

LH2 Technology

Demo System

 Element

CFM payload mass

CFM propellant (LH2)

Spacecraft bus

Launch mass

Mass (kg)

1,207 

260 

831 

2,298 

Spacecraft Size

Length:

Diameter:

4.4 m

2 m 

Mission Duration: 6 Months

Figure 4. CPST Government POD mission concept.
8

2. STUDY TOPIC REVIEW

2.1 Mission Justification

2.1.1 Mission Justification and Rationale

During the course of this study, team members assessed CFM technology gaps in light of 

the full range of mission architecture scenarios. In the early phases of the study, these assessments 

were based on expert knowledge of previously envisioned mission architectures available in the 

open literature. As the study progressed, the assessments expanded to include all available versions 

of the OCT technology roadmaps as well as presentations generated by the Human Exploration 

Framework Team (HEFT) and its successor, the Human Architecture Team. As shown in table 4, 

cryogenic storage technologies are enabling for at least seven and as many as ten of the OCT roadmaps while expulsion and transfer technologies are enabling for at least four and as many as nine 

roadmaps. Also, table 5 highlights the mapping of these cryogenic technologies to the ‘top 10’ 

lists in the OCT roadmaps with direct or significant relevance to nine challenges in three different 

roadmaps. Assessment of the range of HEFT architectures indicated that long-duration cryogenic 

storage (i.e., for storage durations beyond current state-of-the-art, about 9 hr) is a key enabler for 

missions beyond low-Earth orbit (LEO), as illustrated in table 6.

Table 4. Mission relevance to OCT roadmaps.

Probably

required

Active

cooling

Passive

cooling

Tank gauging

Liquid

acquisition

Cryo Strorage

Technologies

Cryogenic

Expulsion 

and

Transfer

Technologies

Feeding

conditioning

Auto

couplings

Probably

required

Probably

required

Probably

required

May be

required

Probably

required

May be

required

May be

required Required

Required Required Required Required Required Required Required

Probably

required N/A May berequired Required Required Required Required N/A Required Required

N/A N/A N/A

N/A

N/A

Probably

required

May be

required

May be

required Required Required Required Required

Probably

required

Probably

required

N/A May be

required

May be

required Required Required N/A Required

Probably

required

Probably

required

May be

required N/A Required N/A Required N/A N/A

N/A Required Required Required

Launch

Propulsion

Systems

In-Space

Propulsion

Technologies

Space

Power

and

Energy

Systems

Human

Health,

Life 

Support,

and

Habitation

Systems

Human

Exploration

Destination

Systems

Science

Instrumentation,

Observation,

and Sensor

Systems

Modeling,

Simulation,

IT and

Processing

Materials,

Structures,

Mechanical

Systems,

and

Manufacturing

Ground and

Launch

Systems

Processing

Thermal

Management

Systems

TA01 TA02 TA03 TA06 TA07 TA08 TA11 TA12 TA13 TA14

OCT Roadmaps That Benefit From CRYOSTAT Technologies*

CRYOSTAT Technologies

* Based on paragraph-by-paragraph assessment of OCT roadmaps. This table is a summary of detailed evaluation.
9

Table 5. Mission relevance to top challenges in OCT roadmaps.

Long-term, in-space

cryogenic propellant

storage and transfer*

Advance in-space

cryogenic engines

and supporting

components**

Nuclear thermal

components and

systems**

Advanced 

space-storable 

propellants**

TA02 (In-Space Propulsion) TA14 (Thermal

Management Systems) TA01 (Launch Propulsion)

Low-conductivity

structures and supports*

Innovative thermal

components and

loop architectures**

Nontoxic RCS**

Advanced main

propulsion system

components**

20K cryocoolers and

propellant tank

integration*

#2 #2 #4

#4 #3

#7

#8

#4

#5

CPST demo directly addresses challenge.

CPST demo addresses significant part of challenge.

*

**

Table 6. Mission relevance to human architecture destinations.

Technology Applicability

for Future Destinations*

LO2/LH2 reduced boiloff

flight demo

LO2/LH2 reduced boiloff

and other CPS tech

development

LO2/LH2 zero boiloff tech

development

Technology Need

May be

required

May be

required

May be

required

May be

required

Probably

required

Probably

required

Probably

required

May be

required

N/A Required

N/A N/A

N/A N/A N/A N/A

Required Required Required Required Required Required

Required Required Required Required Required Required Required

N/A

LEO

Advanced

LEO Cis-Lunar

Minimum

NEA

Full Near

Earth

Asteroid

Mars

Orbit

Mars

Moons

Mars

Surface

Lunar

Surface

Sortie

Lunar

Surface

Outpost

*Adapted from chart 26 of HEFT Final Briefing dated 1/18/2011, posted at: <http://www.nasa.gov/pdf/511089main_HEFT_Final_Brief_508_20110111.pdf>

LEO Beyond LEO Moon NEA Mars

Destination

Analysis of these multiple viewpoints led to the conclusion that long-duration cryogenic 

propellant storage and transfer is not only a suite of technologies that is broadly beneficial for 

a range of spacecraft systems, but is most notably an enabler for key architecture elements such 

as long-duration CPSs, in-space cryogenic propellant storage facilities (e.g., tankers and depots), 

and nuclear thermal propulsion systems. 
10

	 The fact that CFM technology advancement has stalled at TRLs 5 and 6 due to lack of a flight

demonstration is well documented.In fact, the technology gaps identified in this present study bear

remarkable similarity to the gaps identified in the COLDSAT studies conducted by three independent

contractor teams in the early 1990’s and summarized in 1997 by Glenn Research Center (GRC). The

intervening years have witnessed cryocooler technologies development, vastly improved analytical

tools, significantly expanded 1-g test databases, and improved gauging technologies; however, key

technologies that are dependent on the acceleration environment (i.e., 1-g versus microgravity) are still

awaiting advancement to TRLs 6 and 7. Whether NASA selects an architecture based on a heavy-lift

vehicle or an on-orbit cryogenic storage capability, the need for a flight demonstration of cryogenic propellant storage is required in the near term to bring full-scale development risks to acceptable levels. The

study team has identified the CPS preliminary design review (PDR) as the most likely driving constraint

for the CPST flight date. A flight program such as the CPS needs TRLs 6 and 7 in the PDR timeframe to

avoid unacceptable levels of development risks. The team believes the CPST flight should be complete

and the flight data correlated in time for CPST conclusions to be considered with the CPS preliminary

design package development. Assuming the CPS PDR occurs in the fourth quarter of CY 2017, and

allowing time for PDR package development, this line of reasoning leads to a CPST flight that ends no

later than the second quarter of CY 2017, indicating a launch date in the early fourth quarter of CY 2016.

2.1.2 Value and Benefits for Future Applications

The ability to store and expel (or transfer) subcritical cryogenic liquids in space over long 

periods of time has significant benefits for NASA missions. Since cryogenic oxygen and hydrogen 

have much higher performance characteristics than conventional storable propellants, a capability 

to store and use LO2 and LH2 in the propulsion systems of long-duration spacecraft opens a path 

to shorter transit times and/or increased payload masses for long-duration missions to the Moon, 

Mars, and other destinations. Considering that many propellants compatible with in situ propellant production must be stored as cryogenic liquids to meet vehicle mass and volume requirements, 

these new cryogenic capabilities would also help bridge the gap to ‘live-off-the-land’ architectures, 

such as the envisioned production of methane, hydrogen, or oxygen by carbon dioxide reduction 

or water electrolysis. These same storage, acquisition, and gauging technologies are also useful for 

accelerated missions to Mars and outer planets, since nuclear thermal propulsion systems require 

the use of these CFM technologies to assure that their propellants (such as xenon) are stored 

within acceptable volume and dry-mass constraints.

The ability for long-term, low-boiloff cryogenic storage also enables extended duration 

upper stages that are a confirmed interest to the commercial launch providers and their customers, 

including the Department of Defense. These longer duration, upper-stage missions enable a wider 

range of orbit selections for Earth-orbiting satellites. 

2.1.3 Mission Goal Prioritization

During this study, a team of subject matter experts ranked the mission objectives in light of 

perceived stakeholder needs. This team represented a cross section of NASA’s CFM technologists 

as well as spacecraft fluid system and propulsion system development experts. This community of 

experts represented experience in CFM ground testing, the design and operation of both storable 
11

and cryogenic propulsion systems, and the design of propellant transfer systems. Viewing the most 

likely and most immediate stakeholders to be the CPS or an in-space propellant storage function 

(i.e., a tanker or depot), the group ranked the CPST mission objectives to ensure that the most 

daunting development risks and the technology gaps without feasible workarounds were ranked 

highest. The group was careful to assure that the ranking reflected ‘mission pull’ rather than ‘technology push.’ The results of this ranking are shown in table 2.

	 Hydrogen technologies ranked higher than oxygen technologies. Closing the hydrogen technology gaps significantly reduces the risk of failing to conquer the oxygen gaps, due to hydrogen’s significantly lower boiling point and decreased surface tension properties. Storage ranked highest, since failure

to achieve significantly reduced boiloff (RBO) rates would be a first-order driver in sizing (and paying

for) future architecture elements.

	 Acquisition and expulsion of bubble-free cryogenic liquids ranked below storage due to the

catastrophic hazard associated with feeding entrained bubbles to a rocket engine and due to the complex interaction of surface tension and thermodynamics in the design and operation of the acquisition

device(s) in a long-duration, microgravity environment. It is also worth mentioning that acquisition is an

enabler for systems that purport to use cryogenic liquids as reaction control system (RCS) propellants

(not merely as main engine propellants), since the ability to settle main engine propellants would require

a robust acquisition capability within the RCS.

	 Gauging in microgravity ranked third. The gauging function is critical for the efficient operation of a fully capable storage system or propulsion system. The risk is initially offset by the somewhat

inefficient workaround of carrying additional RCS propellant to settle the cryogenic liquids, enabling the

use of somewhat lower risk, settled gauging methods. There are scenarios (such as the need for multiple

gauging operations near the end of a depot/tanker transfer operation) where settled gauging would be an

unwieldy, if not completely impractical, solution. Hence, this third-place ranking still represents a very

significant need.

	 Finally, the overall transfer technology suite ranked fourth—not because this technology is unimportant, but because the others represented more significant development risks in the collective opinion

of the expert team. It is noteworthy that the team concluded that the demonstration of an automated fluid

coupling in flight did not rise to the level of a flight objective. Close examination revealed that this coupling can be effectively matured through ground tests (including cryogenic flow testing and 6-degree-offreedom mate/demate simulations) and does not require the microgravity environment to achieve TRL 6,

since the microgravity environment does not significantly affect the design or operation of the coupling.

2.1.4 Satisfaction of Mission Objectives

	 The POD architecture settled on a hydrogen-only technology demonstration based on the higher

ranking hydrogen technology objectives as well as the challenging budget. As explained above, the

hydrogen-only approach does not close the oxygen technology gaps, but it does significantly reduce the

risks associated with the development of cryogenic oxygen capabilities. Within the fiscal realities of the

technology demonstration mission, the hydrogen technologies demonstration is considered the very best

return on investment for the eventual stakeholders in these technologies.
12

2.2 Mission Concept

2.2.1 Mission Analysis

2.2.1.1 Ground Rules and Assumptions. The guiding ground rules for the mission analysis work are listed in table 7. Since the target users of this technology are space missions that will 

require storage of cryogenic propellants on both LEO, translunar, and interplanetary trajectories, 

the thermal team decided that the best relevant thermal environment for the demonstration would 

be a LEO with an inclination that included both sun and shadow. As a result, the mission analysis team ground ruled a low-Earth circular orbit with a target inclination that would allow time in 

both sun and shadow, and decided to let the lifetime requirement determine the initial orbit altitude. Experience suggested that the target altitude for a circular orbit would most likely be in the 

range of 300 to 600 km. 

Table 7. Mission analysis ground rules and assumptions.

Target orbit LEO circular, ≈ 500-km altitude (determined by 

lifetime requirement), with inclination relevant to 

future space missions that will require cryogenic 

propellant storage

Nominal mission 

duration

6 mo

Target orbital lifetime 1 yr

Controlled reentry Yes; the team assumes that a controlled reentry 

is required due to the anticipated mass of the 

spacecraft

Launch year 2016

For this study, the target orbital lifetime is defined as the estimated time that the spacecraft 

will orbit before atmospheric reentry, assuming no stationkeeping. Based on data from previous 

studies and on the advice of the Orbital Debris Program Office at Kennedy Space Center (KSC), 

analysts set the target orbit lifetime to be twice the target spacecraft lifetime or 1 yr beyond the 

end of the mission, whichever was less. This resulted in a target orbit lifetime of 1 yr. Based on the 

anticipated mass of the spacecraft, and in conjunction with data from previous studies, the team 

decided to assume that a controlled reentry would be required to keep the risk of human casualty 

below the required probability of 1:10,000. Since the design was not complete until the end of the 

study, the team recommends that the assumption of controlled reentry be reexamined in follow-on 

work. A preliminary reentry analysis completed by the Mission Design Lab at GSFC supports the 

controlled reentry assumption, but not all spacecraft components were included.

The assumed launch year of 2016 results in conservative lifetime estimates since the solar 

flux values are expected to be decreasing from 2016 (which is the earliest that CPST is expected to 

launch) through 2020. Therefore, if the lifetime requirements are met for the year 2016, then they 

should be met or exceeded for all subsequent years up to 2020.
13

2.2.1.2 Mission Profile. Following launch, the spacecraft will operate for 2 wk in checkout 

mode. Assuming satisfactory completion of the checkout operations, the next month on orbit will 

involve the passive CFM storage demonstration. Afterwards, the spacecraft will perform the active 

portion of the CFM demonstrations. Following an optional extended mission phase of up to 6 mo, 

the spacecraft will perform a controlled reentry into the Earth’s atmosphere.

2.2.1.3 Orbit Lifetime Analysis and Target Orbit Altitude. Orbit lifetime analysis ensures 

that the spacecraft will remain in the required orbit and not be overcome by aerodynamic drag or 

other perturbations over the required duration of the mission. Given a 6-mo mission, a 1-yr orbit 

lifetime requirement conservatively results in an initial orbit altitude that will easily meet the mission requirements. The mission analysis team used NASA’s Debris Assessment Software (DAS) 

version 2.0, developed at Johnson Space Center (JSC), and Life TIME (LTIME), an orbit lifetime 

tool developed at Marshall Space Flight Center (MSFC), to estimate the orbit lifetime and decay 

values for CPST. Stationkeeping requirements came from a simple custom tool developed within 

the MSFC Advanced Concepts Office. 

The aerodynamic area tool in DAS provided a convenient way to estimate the area/mass 

ratio for CPST. The user inputs the basic geometry of the spacecraft, specifies an appropriate 

aerodynamic attitude, and lets DAS determine the aerodynamic area. The configuration used in 

DAS to determine the aerodynamic area is shown in figure 5. Assuming an aerodynamic attitude 

of ‘random tumbling’ resulted in an aerodynamic area of 14.16 m2. This attitude represents the 

best match among the options of random tumbling, gravity gradient, and fixed attitude, since the 

actual spacecraft attitude will be solar inertial with the aft end of the spacecraft facing the Sun. 

This solar inertial attitude results in the spacecraft slowly tumbling relative to the oncoming 

atmosphere during flight.

15

14

13

12

11

10

9

8

7

6

5

4

3

2

1

0

0 1 2 3 4 5 6 7 8 9 10 11 12 13

Units

Solar Array:

12-m2

 Total Area

Thermal Radiators:

4.4-m2

 total area

Cross-Sectional Area Plot

Figure 5. Configuration used in DAS to determine the area/mass ratio.
14

The resulting DAS output for the aerodynamic area is an input into LTIME, along with 

the spacecraft mass, orbit, aerodynamic parameters, and other values as listed in table 8. Analysts 

compared a few test cases of LTIME with two other computer applications that can also estimate 

lifetime: DAS and the Satellite Tool Kit (STK). Since the results compared well, the team used 

LTIME due to its speed and more user-friendly output. A comparison of the tool outputs indicated 

that a nominal solar flux along with a drag coefficient of 1.8 would result in LTIME output closely 

matching the other tools.

Table 8. Input parameters for the LTIME satellite lifetime program.

Epoch 1/1/2016

Orbit altitude 300–600 km, circular

Inclination 28.5 deg (see text for explanation)

Longitude of ascending node, 

argument of perigee, and mean 

anomaly

0

Area 14.16 m2

Mass 1,500–2,750 kg

Drag coefficient 1.8 (see text for explanation)

Predicted solar flux Nominal (see text for explanation)

Initially, the spacecraft was to be launched from Cape Canaveral Air Force Station 

(CCAFS) to a 28.5-deg inclination. However, midway through the study, launching from WFF 

became a possibility, so the team investigated the effect of orbit inclination on orbit lifetime. Based 

on the data listed in table 9, the team concluded that orbit inclination has a minimal effect on lifetime (up to 40 deg), with launch year having a much larger influence. Thus, all orbit lifetime runs 

in LTIME were with an inclination of 28.5 deg.

Table 9. Effect of launch and orbit inclination on orbit lifetime.

Launch Year

Altitude: 500 km

Area/Mass: 0.0050

Inclination (deg)

0 28.5 40

2017 7 7.1 7.1

2017.5 6.6 6.6 6.7

2018 6.1 6.2 6.2

2018.5 5.6 5.7 5.7

2019 5.1 5.2 5.3

2019.5 4.7 4.7 4.8

2020 4.2 4.2 4.3

2021 3.3 3.3 3.4

2022 2.9 3 3.1

2023 3.4 3.5 3.7
15

The range of masses used in the analysis was an attempt to cover a wide range of area/mass 

ratios in anticipation of the mass and/or configuration changing. Even if the configuration changes 

considerably, as long as the resulting area/mass ratio is covered in previous LTIME runs, no additional runs are required. The resulting area/mass ratios are listed in table 10. The value in bold 

represents the estimated values for the final configuration of this study.

Table 10. Resulting area/mass ratios used

in the orbit lifetime analysis.

Mass (kg) Area/Mass (kg/m2)

1,500 0.0094

1,750 0.0081

2,000 0.0071

2,200 0.0064

2,500 0.0057

2,750 0.0051

Many runs of LTIME resulted in a large table of data listing the expected orbit lifetimes for 

various area/mass ratios. An abbreviated table is provided in table 11. One can see that for the current configuration’s area/mass ratio of 0.0064, an initial circular orbit altitude of 400 km provides 

plenty of margin on the orbit lifetime since area/mass ratios of 0.0060 and 0.0080 result in lifetimes 

of 1.54 and 1.11 yr, respectively. Note that this analysis so far does not include the effect of launch 

vehicle insertion errors, which are discussed below. Also, since the predicted solar flux values are 

decreasing throughout 2016–2020, these tabulated values will increase for launches beyond 2016.

Table 11 also lists estimated stationkeeping delta-velocity (∆V) requirements. These values 

are rough estimates given by taking the orbit decay rates from LTIME and putting these into a custom tool that determines the required ∆V based on a user-specified maximum amount of altitude 

decay before boosting the orbit back to its original value. (Only corrections to altitude are considered here.) In this study, analysts set the maximum altitude decay value to 0.2 km before reboost 

was necessary. However, while the values were calculated and included in the results, the team’s 

opinion is that ‘stationkeeping is not required for CPST’ since specific altitude and inclination 

requirements are not part of the requirements, the only requirement being lifetime.

The team’s final step in determining the recommended target initial orbit altitude was to 

consider launch vehicle insertion errors. Since at least one of the candidate launch vehicles contains 

a solid stage, the insertion errors may be large enough to decrease the orbit lifetime substantially. 

Insertion errors from two candidate vehicles, the Minotaur IV and the Taurus II, are listed in 

table 12. While the current CPST design is too heavy for the Minotaur IV, it is included here 

for comparison purposes.
16

Table 11. Orbit lifetime and stationkeeping estimates.

Area/Mass Ratio

Initial Circular 

Orbit Altitude 

(km)

∆V Required for 

Stationkeeping

(m/s)

Altitude After 

0.5 yr With No 

Stationkeeping 

(km)

Orbit Lifetime 

With No

Stationkeeping

(yr)

0.0050 300 71 Reentry 0.16

350 24 168 0.55

400 8.6 382 1.90

450 3.4 444 6.37

500 1.6 497 9.88

0.0060 300 84 Reentry 0.13

350 28 Reentry 0.44

400 10 377 1.54

450 4 442 5.65

500 1.8 497 8.96

0.0070 300 99 Reentry 0.11

350 33 Reentry 0.33

400 12 377 1.11

450 4.7 442 3.97

500 2 497 7.90

0.0080 300 113 Reentry 0.10

350 38 Reentry 0.33

400 14 366 1.11

450 5.4 439 3.97

500 2.2 496 7.90

Table 12. Orbit insertion errors for two example launch vehicles.

Insertion Errors (km)

Minotaur IV Taurus II

Insertion apse 18.5 15

Noninsertion apse 92.6 80

These insertion errors can greatly affect the orbit lifetime. To estimate the magnitude of the 

effect, analysts ran several additional LTIME runs with the estimated insertion errors included. 

The results are listed in table 13. For the purposes of this exercise, the team used an area/mass 

ratio of 0.0070; time constraints did not allow running a variety of area/mass ratios. Launching 

the spacecraft on a Taurus II to a target circular orbit of 400 km could result in a worst-case orbit 

of 320× 385 km, and a decrease in orbit lifetime from 470 to 135 days, which no longer meets the 

lifetime requirement. One solution is to target a 450-km circular orbit, so that even the worst-case 

resulting orbit will have sufficient lifetime (466 days in this case) to satisfy the mission requirements.
17

Table 13. The effect of launch vehicle insertion errors on orbit lifetime.

Target Orbit

(and lifetime)

Actual Orbit

Minotaur IV Taurus II

Apogee 400 381.5 km 385 km

Perigee 400 307.4 km 320 km

Lifetime* 470 110 days 135 days

Apogee 450 431.5 km 435 km

Perigee 450 357.4 km 370 km

Lifetime* 1,760 376 days 466 days

Apogee 500 481.5 km 485 km

Perigee 500 407.4 km 420 km

Lifetime* 3,040 1,350 days 1,750 days

* Lifetime assumes 2016 launch, area/mass of 0.007.

Therefore, taking into consideration launch vehicle insertion errors, launch date, solar flux, 

and the possible area/mass ratios of the CPST configuration, ‘the mission analysis team recommends a target circular orbit altitude of 450 km.’ Additional options that could be investigated 

include, but are not limited to, using high-accuracy orbit insertion options for the Taurus II, 

having the spacecraft perform propulsive maneuvers to correct for insertion errors, and having 

the spacecraft perform stationkeeping.

2.2.1.4 End-of-Life Disposal and Risk of Human Casualty. The team completed a preliminary assessment of the Risk of Human Casualty as specified in NASA-STD-8719.14, Process 

for Limiting Orbital Debris, NASA’s DAS, version 2.0.1, was used in the analysis.1 While not all 

spacecraft components were included in the assessment, those that were resulted in a risk of human 

casualty of 1:2,100, which exceeds the NASA standard of 1:10,000. Therefore, the analysis shows 

that a controlled reentry will probably be required.

2.2.1.5 Preliminary Propellant Requirements. The propellant was estimated to accomplish 

three major tasks: propellant settling during the fluid transfer and mass gauging experiments, 

collision avoidance, and the end-of-life disposal of the spacecraft. For the propellant settling, the 

assumed initial mass was 2,210 kg. The fluid transfer experiments required an average of 3×10–5 g 

for a total of 312 min, while the mass gaugings required 1×10–3 g for 48 min. This resulted in 

a propellant settling ∆V of 34 m/s. A collision avoidance maneuver was included with a ∆V of 

3 m/s and an additional 0.2 m/s for attitude control.

Before the controlled reentry, most of the cryogenic propellant must be dumped overboard. 

Three percent of the initial load is assumed to remain in the tanks and lines. The deorbit (from 

the 450-km circular orbit) is separated into three maneuvers. The first two burns, both at 450 km, 

reduce the perigee altitude to 150 km. The final maneuver is performed such that the flight path 

angle is –1.2° at an altitude of 60 km. A preliminary finite-burn analysis resulted in a total deorbit 

∆V of 144.3 m/s with an additional 14.3 m/s for attitude control.
18

The mission events and corresponding propellant requirements are listed in table 14. Margins of 30% were added to the settling and avoidance ∆Vs. More fidelity went into the deorbit 

analysis, justifying a smaller 5% margin. The required maneuver propellant was 192 kg. A residual 

amount of 5% was added to account for the propellant trapped in the tanks and lines, resulting 

in a total loaded propellant requirement of 202 kg. For these propellant computations, a constant 

specific impulse (Isp) of 220 s was assumed, which is typical for monopropellant propulsion systems. A more detailed analysis is required to track the variation of the Isp and thrust during the 

various events, especially the controlled reentry.

Table 14. Preliminary mission propellant estimate.

Event

Mass

(kg)

MPS ∆V

(m/s)

ACS ∆V

(m/s)

Margin

(%)

Total ∆V

(m/s)

Propellant

(k/g)

Propellant settling 2,210 – 34 30 44.2 44.8

Collision avoidance 2,175 3 0.2 30 4.1 4.2

Deorbit 1,923 144.3 14.3 5 166.5 142.9

Maneuver propellant: 192

Residuals: 10

Total propellant requirement: 202

2.2.2 Technology Demonstration Timeline

The CPST mission timeline was developed to satisfy the identified needs, goals, and objectives as well as the requirement for a mission duration of 6 mo.

2.2.2.1 Storage Demonstration Timeline. The top-level mission timeline is shown in figure 6 

and begins with on-orbit arrival. There will be an initial checkout period for the vehicle and the 

CFS systems lasting an estimated 10 days. At or about the same time, the passive storage demonstration will begin with an initial settled mass gauging event and activation of the CFS data collection system. It is vital that thermal performance data collection begin as soon as possible in order 

to characterize the post-ascent thermal transient due to venting and offgassing of the storage tank 

insulation system. The passive storage demonstration is planned to last 1 mo. This should allow 

adequate time for data collection through the thermal transient period and subsequent steady-state 

period.
19

Mission Demonstration

Month

1 2 3 4 5 6

Spacecraft and CFM Demo Systems Checkout

LH2 Storage Tank Passive CFM Demo

LH2 Storage Tank Active CFM Demo

LH2 Transfer Demos

Storage Tank Passive CFM Demos Include:

– Determination of passive thermal control performance

– Settled mass gauging

– Unsettled mass gauging

– Low-conduction structural concepts

Storage Tank Active CFM Demos Include:

– Determination of active thermal control performance 

– Settled mass gauging

– Unsettled mass gauging

– Low-conduction structural concepts

Propellant Transfer Demos Include:

– Pump-fed propellant transfer

– Pressure-fed propellant transfer

– Settled propellant transfer

– Unsettled propellant transfer

– Transfer tank and transfer system conditioning

– Transfer rate measurement and vapor detection

– Settled and unsettled liquid acquisition

– Tank expulsion demos

Tanks Sized to Provide (at least):

6-Mo Storage Demo for LH2

Two Transfer Demo Series for LH2 

Figure 6. CPST Government POD demonstration timeline.

Upon completion of the passive storage demonstration, the active storage demonstration 

will begin with another settled mass gauging event and startup of the storage tank’s active cooling 

system. The remainder of the mission will be conducted with active cooling of the storage tank. 

The active storage demonstration is expected to last 5 mo, with adequate opportunities for collection of steady-state thermal performance data under active cooling.

2.2.2.2 Transfer Demonstration Timeline. During the active storage demonstration, two 

30-day transfer demonstration series will be completed. Each transfer demonstration series will 

include four, two-way transfers of LH2 between the storage and transfer tanks (i.e., transfer from 

storage tank to transfer tank followed by return transfer from transfer tank to storage tank). Two 

transfers will be performed under settled conditions—one pressure fed and the other pump fed. 

Similarly, two transfers will be performed under unsettled conditions—one pressure fed and the 

other pump fed. In addition, the transfer tank and transfer system will be chilled from the spacecraft internal ambient temperature to an initial transfer hardware target temperature prior to 

each two-way transfer. After each two-way transfer, the transfer tank and transfer system will be 

warmed to the spacecraft internal ambient temperature or other warm target temperature in preparation for the next chilldown and two-way transfer.
20

The mission propellant inventory profile is shown in figure 7. The corresponding propellant 

loss and heat load breakdown by location are given in table 15. These show that the propellant loss 

at the end of mission is less than the initial propellant load of 260 kg. Predicted tank heat loads are 

based on analytical results from the CPST Government POD vehicle flight thermal model (June 

2011), constructed using Thermal Desktop® with RadCAD® and SINDA/FLUINT. A margin of 

50% is applied to all heat loads.

Transfer

Demo

Series

Passive

CFM

Storage

Demo

Transfer

Demo

Series

Active CFM Storage Demo

300

250

200

150

100

50

0 0 1 2 3

Mission Time (mo)

Propellant Mass (kg)

4 5 6

LH2 Inventory

Figure 7. CPST Government POD propellant loss profile.

Table 15. CPST Government POD propellant loss and heating rates.

Propellant

Loss

(kg)

Steady-State Heating Rates

Tot Q

Tank

(W)

Str Q

Tank

(W)

Pen Q

Tank

(W)

Rad Q

Tank

(W)

Tot Q

BAC

(W)

Str Q

BAC

(W)

PenQ

BAC

(W)

Rad Q

BAC

(W)

LH2

Storage tank

 Passive storage

 Active storage (reduced boiloff)

24.2

107.2

4.13

3.66

2.50

2.50

0.61

0.61

1.02

0.55

– 

2.10 

–

–

–

–

–

2.10

Transfer tank (includes chilldown

 and residuals)

50.4 2.02 1.31 0.53 0.18 – – – –

Transfer system (includes 

 chilldown and residuals

57.5 10.01 – – 10.01 – – – –

Storage tank residual 7.8 – – – – – – – –

Total propellant loss 247 – – – – – – – –

Interception of conductive heat loads at the tank supports and penetrations with active 

cooling was not included in this analysis, due to the lack of a vetted heat exchanger concept. 

It is hoped that one or more viable heat exchanger concepts will be proven in the CPST WBS 4.0 
21

Technology Maturation activities. The increased conservation of propellant resulting from the 

interception of conductive heat loads could enable a third transfer demonstration series in the 

CPST mission timeline. This would be a valuable thermal design feature and should be addressed 

in future efforts.

Figure 8 depicts the content of each transfer demonstration series in greater detail. The four, 

two-way transfers with associated chilldown and settled mass gauging activities are shown notionally. Unsettled mass gauging is shown as an ongoing activity during all settled and unsettled periods, since the associated cost in terms of power and data handling is relatively low. The two-way 

transfers will be scheduled approximately a week apart to allow for downlink and review of CFS 

data and possible preparation of new instruction for uplink before the next two-way transfer.

Propellant transfer demos also include:

– Transfer rate measurement and bubble detection

– Settled and unsettled liquid acquisiton

– Propellant expulsion demos

– Unsettled steady-state waiting periods preceding

 unsettled and settled transfers

Settled mass gauging

Unsettled mass gauging

System chilldown

Settled pump-fed transfer

Settled pressure-fed transfer

Unsettled pump-fed transfer

Unsettled pressure-fed transfer

Downlink and review data

Transfer

Demonstration 01 02 03 04 05 06 07 08 09 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 25 26 27 28 29 30

Day

Figure 8. Thirty-day transfer demonstration series timeline.

Tables 16 and 17 provide even more detailed breakdowns of events for individual settled 

and unsettled two-way transfers, respectively. Approximately 5 hr are required to complete each 

two-way LH2 transfer. The sequence of events, the duration of individual events (LH2 transfers, 

pretransfer chilldowns, unsettled waiting periods, and mass gauging events) and target acceleration 

levels for the CFM Technology team provided settled operations.

Due to the mass impact associated with providing settled conditions, the allowable duration 

of settled conditions will be limited. For purposes of this study it has been assumed that 30 min 

will be required for each two-way settled transfer and 2 min will be required for each settled mass 

gauging event. Axial thrust levels of 1 mg and 30 μg have been selected for settled operations. No 

axial thrust will be provided for unsettled operations. Total durations for each axial thrust level are 

shown at the bottom of tables 16 and 17.
22

Table 16. Settled transfer demonstration timeline.

Total Time at 0.001 g’s: 24 min for CPST Government POD (total associated with all settled transfers). 

Total Time at 0.00003 g’s: 216 min for CPST Government POD (total associated with all settled transfers). 

Elapsed Time

(min)

Event Duration

(min) Event

Thrust Acceleration

(g’s)

2 2 Settled mass gauging 0.00003

4 2 Settled mass gauging 0.001

19 15 Unsettled steady state waiting period 0

 31 2 × 6 Chilldown prep 0.00003

139 18 × 6 Chilldown 0

141 2 Settled mass gauging 0

156 15 Settled propellant transfer from storage tank 0.00003

160 2 Settled mass gauging 0.001

158 2 Settled mass gauging 0.00003

175 15 Unsettled steady state waiting period 0

177 2 Chilldown prep 0.00003

285 18 × 6 Chilldown 0

287 2 Settled mass gauging 0.00003

302 15 Settled propellant transfer to storage tank 0.00003

304 2 Settled mass gauging 0.00003

306 2 Settled mass gauging 0.001

Table 17. Unsettled transfer demonstration timeline.

Total Time at 0.001 g’s: 24 min for CPST Government POD (total associated with all unsettled transfers). 

Total Time at 0.00003 g’s: 96 min for CPST Government POD (total associated with all unsettled transfers). 

Elapsed Time

(min)

Event Duration

(min) Event

Thrust Acceleration

(g’s)

2 2 Settled mass gauging 0.00003

4 2 Settled mass gauging 0.001

16 2 × 6 Chilldown prep 0.00003

124 18 × 6 Chilldown 0

126 2 Settled mass gauging 0.00003

141 15 Unsettled steady-state waiting period 0

156 15 Unsettled propellant transfer from storage tank 0

158 2 Settled mass gauging 0.00003

160 2 Settled mass gauging 0.001

162 2 Chilldown prep 0.00003

270 18 × 6 Chilldown 0

272 2 Settled mass gauging 0.00003

287 15 Unsettled steady-state waiting period 0

302 15 Unsettled propellant transfer to storage tank 0

304 2 Settled mass gauging 0.00003

306 2 Settled mass gauging 0.001
23

2.2.2.3 Minimum Mission Timeline. During the course of the study, it was determined that 

by starting the first transfer demonstration series 2 wk earlier, contingency storage and transfer 

demonstrations could be accomplished within the first 2 mo of the mission timeline, if necessary.

2.2.3 Mission Communications Architecture

The communications architecture covers the hardware, software, spacecraft and CFM command and control terminals, and networks required to operate CPST. The major components and 

their functions are as follows:

• Ground stations (White Sands (WS1), Wallops, and Santiago (AGO) with Dongara (AUWA) 

and Hawaii (USHI) as backup):

– Receive spacecraft and CFM data from CPST.

– Send commands to CPST (minimal command processing).

– Send the spacecraft data to the Missions Operations Center (MOC).

– Retransmit data upon request.

• Data distribution networks—combination of NASA integrated services network and the local 

area network:

– Provide connectivity between ground station(s) and the MOC and Technology Operations 

Center (TOC).

– Meet any needs for dedicated voice communication between MOC and TOC.

• MOC—overall command and control:

– Convert CFM objectives into spacecraft and CFM commands in order to perform mission 

operations (command fabrication).

– Perform command stream validation.

– Analyze and monitor real-time telemetry from the spacecraft for passes that are scheduled.

– Analyze real-time and stored health and status to generate trending data.

– Ensure healthy operation of the spacecraft bus.

– Manage the spacecraft onboard data recorder.

– Detect, report, and troubleshoot anomalous conditions.

– Transmit scripts, commands, data tables, and software updates to CPST via the ground 

station(s).

– Receive spacecraft data from CPST via the ground station(s).

– Station scheduling.

– Maintain data dictionaries.

– Archive scripts, commands, data tables, and software updates that were sent to CPST data 

storage and archive.

• Flight dynamics function:

– Perform targeting for controlled reentry.

– Provide acquisition data to the tracking networks.

– Provide time history of orbit and attitude based on downlinked spacecraft telemetry.
24

• TOC—demonstration objectives and analysis:

– Provide demonstration objectives and onboard data recorder recommendations to MOC.

– Receive CFM data from MOC.

– Analyze CFM health and status data post pass to ensure healthy operation of the CFM 

and recommend modification to operational processes.

– Detect, report, and support trouble shooting of anomalous conditions in the CFM.

– Support CFM sensor calibrations, algorithms, trending, and built-in tests.

2.2.4 Prelaunch to Spacecraft Separation

Figure 9 shows a simplified view of the communications architecture during this phase. 

CPST will undergo cryogenic loading and prelaunch checkout at the launch site prior to launch. 

This will require MOC, TOC, and/or an engineering support room at the launch site. CPST will 

not have telemetry interleaved with the vehicle data, but due to the uniqueness of cryogenic loading at the pad, there may be umbilical disconnects that are controlled through the vehicle launch 

sequencer. Also, the cryogenic loading information may need to be routed to the launch vehicle 

control room for the launch vehicle team to review.

= Telemetry

= CPST Data

= Timing/Tracking

= Umbilical Data

TDRS (TBD)

Launch Vehicle

Control Room

WSTF

Launch Vehicle

Ground Station

Note: Assuming no CPST video

or telemetry through vehicle.

Assuming only command during

ascent is destruct. NISN

MOC, TOC, Engineering

Support Room – TBD

Figure 9. Prelaunch to spacecraft separation communications architecture.

During ascent, the use of a tracking and data relay satellite (TDRS) is to be determined. 

The launch vehicle will send launch vehicle telemetry to its ground station. At this point it is 

assumed that CPST will record all data during ascent for later transmission.
25

2.2.5 Spacecraft Separation to Controlled Deorbit

Figure 10 shows a simplified view of the communications architecture in this phase.

GPS TDRS

SSA

5 kbps telemetry

2 kpbs cmd

* TDRSS is only

 used for critical

 operations.

 Normal operations

 use ground stations

Wallops and Santiago

Backup

Hawaii and Dongara

GRC CPST

MOC/Distribution

Center

White

Sands

Complex

S-band telemetry: 

4 Mbps

QPSK

cmd: 2 Kpbs

Command

Science & Hskpg

Command

Housekeeping

Figure 10. Spacecraft-controlled deorbit communications architecture.

The first 2 wk after spacecraft separation, during on-orbit checkout, a tracking and data 

relay satellite system (TDRSS) would be used in addition to the ground stations, but after that 

period, normal operation would ensue and ground stations would be used primarily. 

The MOC distributes all CFM data to the TOC for their review and analysis.

2.2.6 S-Band

S-band is required for uplinking commands to CPST. Uplink of commands can be done 

through TDRSS or one of the ground stations. The S-band frequency offers 4 Mbps of telemetry 

downlink to the ground stations and 5 kbps telemetry downlink through TDRSS. After the fluid 

transfers, all the CFM data and video can be downlinked within 24 hr; therefore, a higher frequency band (such as X-band) is not required.

The analysis in figure 10 shows the ground coverage for White Sands (WS1), Wallops, 

and Santiago (AGO) with Dongara (AUWA) and Hawaii (USHI) as backup. This analysis used 

a 500-km altitude, 40-deg inclination, and 5-deg minimum station elevation. The S-band coverage 

summary is shown in table 18.
26

!

Figure 11. Ground station S-band coverage.

Table 18. S-band coverage summary.

Ground Station 

Passes

per Day*

Average Pass 

Duration** 

(min)

Maximum 

Pass Duration 

(min)

Total Minutes

per Day

White Sands 6 8.95 9.74. 53.68

Wallops 5.21 9.26 9.78 48.26

Santiago 5.71 9 9.76 51.42

Dongara 6.71 8.43 9.74 56.63

Hawaii 5 8.24 9.71 41.22

Station coverage for the S-band assumption of >5°

elevation of the GRC CPST satellite in a 500-km circular orbit inclined 40°

 * Numbers based on a random 14-day coverage; only passes >5 min are considered.

** Under favorable conditions, passes may be longer but cannot always be assured.

The peak data volume per day is 18 Gbits, which required about nine passes to downlink 

to the ground. The peak occurs after a fluid transfer experiment. At other times, the data volume 

is about 6 Gbits per day, which requires three passes to downlink.

2.3 Demonstration Flight System

2.3.1 Configuration Overview

The layout of the CPST configuration can be divided into two main sections: a CFS 

and the spacecraft bus (fig. 12). The CFS is the experiment part of the spacecraft. The spacecraft 

bus provides the platform for the CFS to conduct its required mission and test objectives.
27

CFS—Experiment

Spacecraft Bus

Figure 12. CPST primary sections.

2.3.1.1 Design Drivers. The main design constraint was to meet the LH2 volume requirements of the storage and transfer tanks. In addition to the volume requirements, the other design 

driver was the maximum allowable payload size of the expected launch vehicle (fig. 13). Early in 

the conceptual phase a Minotaur IV launch vehicle was chosen. The storage tank was reduced in 

diameter to allow for the thermal isolating support structure and anticipated thermal insulation 

thicknesses. A final diameter of 1.7 m was selected. After examining the transfer tank sizes and 

dimensions of the pressurant tanks, it was decided to place the storage tank at the top of the spacecraft. The transfer tank is located directly below the storage tank in a diameter transition zone. The 

spacecraft bus mounting diameter and the outside diameter of the CFS creates the zone. This area 

was also selected to mount the eight pressurant tanks. This was done to keep the center gravity as 

low as possible and to group the various CFS components together.

2.3.1.2 Final Launch Configuration. When the final preliminary design was completed, the 

Minotaur IV launch vehicle did not have sufficient lift capability for spacecraft. The Taurus II was 

then chosen for the baseline design. This shroud has a larger diameter of 3.4 m (fig. 14). The design 

could possibly be resized to this new shroud size, which would make the tank dimensions and overall height different. It was decided to stay with the 2-m-diameter size since it would still fit and with 

the time constraints involved. 
28

4.4 m

0.8 m

2.8 m

Ø2 m

LH2Storage Tank

260 kg LH2

1.7-m Diameter; 1.1-m Barrel

0.6-m Dome Height

LH2 Transfer Tank

13-kg Capacity

0.72-m-Diameter Spherical

Figure 13. Size and dimensions.

Ø3.4 m

7.6 m

Figure 14. Taurus II launch configuration.
29

2.3.1.3 On-Orbit Configuration. Once the spacecraft reaches its desired altitude and checkouts are completed, the solar arrays and radiators are deployed. The solar arrays consist of eight 

separate two-panel-folding arrays, which are stowed against each side of the spacecraft bus. For 

this level of study, the two double-sided radiators are assumed to be located in the CFM experiment section in close proximity to the thermal subsystem to minimize plumbing (structural shell 

removed in all figures) (fig. 15).

Solar Arrays (8)

Radiators (2)

Figure 15. On-orbit configuration.

2.3.2 Payload Fluid System Overview

2.3.2.1 Description. The CPST payload fluid system includes all of the necessary functions 

and features to allow demonstration of LH2 cryogenic fluid storage at RBO levels under microgravity conditions, as well as propellant transfer and propellant quantity gauging functions under 

microgravity and settled conditions. To affect reduced propellant boiloff demonstrations, passive 

and active thermal control design features are incorporated, including a thermodynamic vent system (TVS), active propellant mixing, multilayer insulation (MLI), low heat leakage structural support struts, and reduced fluid penetrations for the main storage propellant tank. The CPST payload 

fluid system also employs a cryocooler broad-area cooling (BAC) architecture to actively refrigerate 
30

the external surface area of the storage propellant tank. This intercepts heat entering into this tank 

through the MLI over the surface area of the tank, reducing hydrogen propellant boiloff. 

Helium pressurization is provided to support pressurized propellant transfer and pump-fed 

transfer functions between the storage and transfer propellant tanks. For the pressure-fed transfer function, adequate driving pressure is provided from the source propellant tank to affect the 

required propellant flow rate to the destination propellant tank. Pressurization control is provided 

through closed-loop control of fast response pressure control valves (‘bang-bang’ valves). It is 

assumed that the helium high-pressure storage and pressure control valves will be thermally conditioned on orbit (heated to approximately –20 to –50 oF) to stay within the temperature capability 

of existing heritage components and associated soft goods. For the pump-fed transfer function, 

active pumping using a motor-driven, impeller-type pump generates propellant flow rate. Cryogenic 

transfer is based on unidirectional flow in transfer lines due to the constraint to use separate connections for liquid entry into and exit from the propellant tanks. Bi-directional flow configurations 

are possible, but with additional valve arrangement complexity. As a result, the unidirectional flow 

configuration is the simplest approach and has been incorporated into the payload fluid system 

design. 

The propellant transfer function includes line and transfer tank chill-in conditioning operations for both the settled and unsettled acceleration environments using pulses of propellant flow. 

Venting once the injected propellant has vaporized and absorbed energy from the line/tank follows 

each pulse. Cooling of lines and the transfer tank is achieved through automated iterative pulse/

vent cycles until temperatures are reached which support no-vent fill operations. The vent system 

is designed to use a common vent line for both the TVS venting and venting of the two propellant 

tanks. The vent exhaust is nonpropulsive to minimize vehicle disturbances during these operations. 

Valving for transfer and venting subsystems utilizes common design to the greatest extent possible 

to reduce cost and technical risk. Also, many of these valves employ a mechanical latching capability to reduce valve electrically-induced heat loads into the propellant systems. Common or derivative pump designs are also assumed for the TVS mixing and propellant transfer functions. These 

pump designs need to be tolerant of short-duration, two-phase flow and low margin on cavitation 

(net positive suction pressure). Derivatives of off-the-shelf laboratory pedigree pump designs are 

assumed. Key CPST payload fluid design parameters are shown in table 19.
31

Table 19. CPST payload key design parameters.

CFS design pressure levels: Comment

Helium pressurization 4,500–250 psi

Storage and transfer tank subsystems 

and transfer subsystem

15–40 psia

Vent subsystem: 0–40 psia

 Cryocooler/BAC circuit Up to 250 psia At operating temperature of ~90 K

Propellant flow rates:

 TVS mixing 244 lbm/hr LH2

 Propellant transfer 78 lbm/hr LH2 With capability to increase up to TVS 

mixing flow rates

Propellant storage tank sizing basis:

 Propellant condition Saturated LH2 at 15 psia

 Loading ullage volume 10% of useable propellant volume

 Combined tank internal component

 Volume and loading uncertainty

5% of useable propellant volume

 Storage tank volume 148.6 ft3

 Transfer tank volume 7.1 ft3 5% of storage tank volume

2.3.2.2 Payload Fluid System Ground Rules and Assumptions. The CPST payload fluid 

system ground rules and assumptions are provided in table 20.

Table 20. CPST fluid system ground rules and assumptions

CFM GR&A

Ground replenishment of cryogens Ground replenishment until T–0

Venting of cryogens Active venting through umbilical until T–0

Ground purge Dry GN2 purge provided for LH2 and LO2 tank MLI until T–0

Tank venting during ascent Active venting to relieve pressure

Tank venting on orbit Active and passive venting through a nonpropulsive vent

Valve power control Fixed current supplied to open valves, holding (reduced 

level) current used when valve needs to be powered open for 

extended duration

Ground actuation of valves Capability to actuate all fill and drain valves and vent valves 

from ground

Use of latching valves Latching valves used wherever media is a cryogen

Vehicle structural/thermal design Similar to CPS nonstructural tank concept

Cryogenic fluid management design Similar to COLDSAT plus active system

Target storage conditions Normal boiling point

CFM performance demos Active CFM and passive CFM

Transfer modes for transfer demos Pressure-fed and pump-fed

Thrust accelerations for transfer demos Zero thrust, 0.00003 g axial, 0.001 g axial

Other demos Mass gauging, liquid acquisition

Post-ascent thermal transient Passive CFM initiated ASAP to assess thermal transient

Margin on all heat loads 50%
32

Key CPST payload ground rules and assumptions include the use of T–0 umbilicals for 

cryogenic servicing activities and storage of the LH2 at normal boiling point (saturated) conditions.

2.3.2.3 Payload Fluid System Hardware Selection Methodology. To minimize costs, strong 

preference was given to selecting off-the-shelf components or components with flight heritage 

or qualification. Additionally, minimizing the total number of different components was also 

preferred to reduce potential design, development, test, and engineering (DDT&E) expenses. 

An individual component design therefore would be used in as many locations within the system 

as practical.

2.3.2.4 Payload Fluid System Redundancy Rationale. The CPST payload is designated as 

a class D mission per NPR 8705.4, Risk Classification for NASA Payloads.2 As such, the payload 

is allowed to design for single-point failure (zero failure tolerance) with mitigation of ‘use of high 

reliability parts, additional testing, or by other means.’ Additionally, ‘single string and selectively 

redundant design approaches may be used,’ but all required ‘applicable NASA safety directives and 

standards’ must be met. Given these design requirements, the CPST payload employs a zero failure 

tolerant architecture as its baseline, in order to reduce cost, although selective areas were addressed 

for failure tolerance redundancy. Specifically, two failure modes were evaluated for fault tolerance: 

internal leakage /failed open position and failed closed position.

2.3.2.4.1 Internal Leakage or Failed Open Position Failure Modes. Components were 

placed in a series configuration to mitigate the risk of excessive leakage or a failed open valve 

that would result in a loss of mission or pose a safety concern. Examples of components in this 

arrangement are helium pressurization control valves, ullage transfer valves, vent valves, the 

helium fill and drain valve and helium quick-disconnect configuration, the LH2 fill and drain valve 

and LH2 quick-disconnect configuration, and the BAC fill and drain valve and sealed plug 

configuration.

2.3.2.4.2 Failed Closed Position Failure Modes. Components were placed in a parallel configuration if a failed closed condition would result in a safety concern, such as overpressurizing the 

fluid system. Examples of components in this arrangement include the burst disk/relief valve and 

vent valve configuration. Due to budget and mass constraints, parallel component arrangements 

could not be made to prevent against loss of the mission if a component were to inadvertently fail 

closed. Conducting protoflight ground testing prior to launch plans mitigation of this risk. This 

test methodology would allow any latent defects in the component design or control system to be 

discovered and rectified prior to on-orbit use.

2.3.2.4.3 Redundancy Not Required. Several of the components in the fluid system do not 

require series or parallel redundancy. One example of this is the transfer pump, whereas the loss 

of the pump would not prevent the transfer of fluids from one tank to the other. Transfer can still 

be accomplished via the use of the helium pressurization system. The other example of no redundancy required is with the backpressure relief valves, which is a check valve used in a relief function. These valves are used to prevent overpressurization of liquid trapped in line segments. Since 

these valves do not have a credible failure mode of failing closed, there is no need to consider this 

mode for redundancy. 
33

2.3.2.5 CPST Payload Fluid System Functions. As identified above, key CPST payload fluid 

system functions include passive thermal conditioning during storage (including TVS and propellant mixing functions), active thermal conditioning during storage (employing a refrigerated cycle 

using a cryocooler and BAC circuit for heat intercept), propellant transfer (between the storage and 

transfer propellant tanks, including line and transfer tank conditioning), propellant acquisition 

(to support propellant transfer and mixing operations), helium pressurization and propellant tank 

venting operations (as required to support propellant transfer and storage objectives), and propellant quantity gauging (under microgravity and settled acceleration environments).

2.3.2.5.1 Overview. To achieve the mission objection of pressurized transfer, a helium 

pressurization subsystem was incorporated into the CPST payload design. A diagram showing 

the baseline design is shown in figure 16.

QD-1100

0.25 in OD×0.035 in Wall 0.25 in OD×0.035 in Wall

0.25 in OD×0.035 in Wall 0.25 in OD×0.035 in Wall

F-1103

(25-μ Absolute)

V-1104

V-1106

OR-1112

OR-1114

V-1108

V-1110

V-1101 T-1102

To Cryogenic

Storage System

To Cryogenic

Receiver System

Figure 16. Helium pressurized subsystem schematic.

The helium pressurization subsystem is designed to provide pressurization capabilities to 

both the storage and transfer tanks via a set of helium tanks manifolded together. Each cryogenic 

tank is pressurized via a separate line segment containing series redundant pressurization control 

valves with a downstream flow control orifice. All helium being supplied to the storage and transfer 

tanks is flowed through a 25-μ absolute filter prior to entering the pressurization control valves. 

This is to minimize the potential for internal contamination that could result in excessive internal 

leakage of the control valves.
34

2.3.2.5.2 Ground Loading Operation. The helium pressurization subsystem is loaded 

with helium prior to integration into the payload shroud. The ground interface quick-disconnect 

QD-1100 is engaged with the ground half disconnect and 28 VDC is applied to the helium fill and 

drain valve V-1101 to allow helium to be loaded into the system until the pressure is ≈4,050 psia. 

Once the helium is loaded, power is removed from V-1101 and the ground half quick disconnect 

is demated from QD-1100.

2.3.2.5.3 Design History. There are two possible component types—a mechanical regulator 

or a normally closed, two-way valve—to control the flow of helium to the cryogenic tanks. Ultimately, a two-way normally closed valve (also called bang-bang valves) was chosen for the CPST 

pressurization control valves. This decision was based on several factors, such as allowing use of 

common valves in a fluid system, lower cost component, lower design risk, and more flexibility 

on cryogenic tank pressure.

The current design uses the same valve for both the helium pressurization valve and helium 

fill and drain valve. Due to the limited funding available to develop and operate the spacecraft, 

commonality of components was desired to reduce DDT&E costs associated with each component 

design.

Regulators compared to two-way valves tend to be more expensive and have more failure 

modes. Additionally, mechanical regulators have more limited operational limits such as flow rate 

and regulated outlet pressure. With two-way valves, the commanded open time can be adjusted to 

get the desired tank pressure. Based on analysis conducted, the time to overpressurize one of the 

tanks was ≈1 s, long enough for the command logic to open and close the pressurization control 

valve.

2.3.2.5.4 Alternate Pressurization Methods. Alternative pressurization methods were 

evaluated as part of the prephase A study. The method evaluated was using autogenously generated 

pressure to pressurize the storage and transfer tanks. There were two autogenous pressurization 

techniques evaluated: direct autogenous and stored autogenous:

(1) Direct autogenous—The direct autogenous approach would pump LH2 through control valves into a heat exchanger to gasify the liquid and send it directly back into the LH2 tank. 

A schematic of this concept is shown in figure 17.
35

HEX-1902 Heat Source HEX-1906

V-1901

V-1900 V-1904

CV-1903

OR-1908 OR-1909

CV-1907

V-1905

Figure 17. Direct autogenous schematic.

The disadvantages with this type of pressurization system are as follows: 

• A high power usage is required to vaporize the liquid, heating, and therefore boiloff of the stored 

cryogen due to heat energy input used for vaporization.

• Collapse of the vapor back into liquid form, resulting in lower than designed pressurization 

of the cryogenic tank. Based on these concerns, the direct autogenous pressurization approach 

was not used for the flight vehicle design.

(2) Stored autogenous—The stored autogenous approach was an enhanced version of the 

direct autogenous method discussed above. In the stored autogenous method, LH2 was pumped to 

a predetermined quantity into an intermediate storage tank. This intermediate tank, which is isolated from the cryogenic tank by control valves, is then heated to vaporize the liquid stored within 

the tank. Due to the long times between transfers, the heat needed to boil off the liquid is generated 

by the environment which eliminates the need for additional power capabilities on the spacecraft. 

After the liquid has vaporized, the control valves would pulse open, just like the helium pressurization valves, to regulate cryogenic tank pressure. An advantage of this approach is that the LH2

used to pressurize is essentially recycled once it collapses back to liquid temperatures. A schematic 

of the stored autogenous approach is shown in figure 18. The stored autogenous method was not 

used due to funding and payload mass limitations for the vehicle.
36

V-1906 V-1907 V-1908

V-1902

V-1901

CV-1903

FM-1900

V-1905

RV-1909

T-1904

From Cryogenic

Storage System

To Cryogenic

Storage System

From Transfer Tank

Vent System

To Cryogenic

Storage System

Figure 18. Stored autogenous schematic.

2.3.2.5.5 Venting—Overview. To ensure there is no tank rupture due to overpressurization, 

a vent subsystem was incorporated into the CPST payload design. A diagram showing the baseline 

design is shown in figure 19.
37

0.75 in OD×0.028 Wall

2 in OD×0.035 in Wall

FM-1208

VP-1209

RV-1200

CV-1206

V-1202

V-1203

2 in OD×0.035-in Wall

From Receiver Tank

Item

From Cryogenic

Storage System

(TVS Vent)

From Cryogenic

Storage System

(Storage Tank Vent)

From Cryogenic

Storage System

(TVS Vent)

Launch Vehicle

Vent Interface

Figure 19. Vent subsystem schematic—storage tank side shown only.

The vent subsystem is designed to manifold vents from both TVS and cryogenic tanks into 

a single vent out of the vehicle. Controlled venting on the cryogenic tanks can be accomplished 

by commanding the vent valves V-1202 and V-1203 to the open position (for the storage tank). 

A check valve, CV-1206, prevents the cavity between the two valves from overpressurizing if there 

is trapped cryogenic liquid between the two valves when the vent valves are in the closed position.

Passive venting of the cryogenic tanks is accomplished using a burst disc /relief valve combination per cryogenic tank. The burst disc minimizes the internal leakage that can occur through 

a closed relief valve prior to the disc being ruptured. The relief valve and burst disc are set to open 

lower than the rated pressure of the cryogenic tanks.

Since the TVS valve vents also are connected with the tank vents, consideration was made 

to ensure that any backpressure generated in the vent lines was included in the TVS valve design 

parameters. A pressure of 5 psia was determined to be in the vent lines during a cryogenic tank 

venting procedure.

2.3.2.5.6 Vent Probe Concept. One of the challenges of venting in space is not generating 

undesired torques on the spacecraft. The concept selected to minimize these torques was what has 

been called the vent probe. An external view, looking at the side of the vent probe, is shown 

in figure 20.
38

Figure 20. Vent probe—external view.

An additional requirement of the vent probe is to duct the vent fluid through a vent line 

when the CPST vehicle is attached to the launch vehicle and then be a nonpropulsive vent after 

the CPST separates from the launch vehicle. To accomplish both of these vent requirements, the 

vent probe is designed to seal in a sleeve fixed to the launch vehicle to route the vent fluid through 

the vent line on the launch vehicle. To ensure a nonpropulsive vent path, the flow within the vent 

probe will be turned 180 deg to eliminate most of the fluid momentum before it is expelled from 

the CPST vehicle. Additionally, there is a large volume internal to the probe to maintain a low fluid 

exit velocity. Figure 21 shows the vent probe in the configuration when integrated to the launch 

vehicle. Figure 22 shows the vent probe configuration when used as a nonpropulsive vent.

Deflector (CPST Side)

Probe (CPST Side)

Sleeve (Launch Vehicle Side)

Flow Path

Flow Flow

Figure 21. Vent probe (payload on launch vehicle) cross section.
39

Flow

Flow

Flow

Figure 22. Vent probe in nonpropulsive vent use—shown in disengaged position.

2.3.2.5.7 Payload Fluid System Liquid Hydrogen Storage—Overview. The LH2 storage 

subsystem contains both the LH2 storage tank and ground fill and drain connection. The LH2 will 

be maintained in the storage tank during flight to verify minimal boiloff, accomplished by both 

passive and active storage techniques.

The burst disc /relief valve (RV-1200), vent valves (V-1202 and V-1203), pressurization check 

valve (CV-1413), and ullage transfer valve (V-1414) are close coupled with the top of the storage 

tank to minimize propellant migration into the lines.

The TVS valve (V-1404), LH2 fill and drain valve (V-1402), and circulation isolation valve 

(V-1411) are close coupled with the bottom of the storage tank to also minimize propellant migration into the lines. To reduce the heat leak into the propellants, the number of storage tank penetrations was minimized. Separate propellant transfer lines were required due to the need to pull liquid 

from the liquid acquisition device (LAD) and flow liquid back into the storage tank outside the 

LAD. A schematic of the LH2 storage subsystem is shown in figure 23.
40

FM-1415

FM-1406

P-1409

CV-1421

CV-1423

T-1400

CV-1413

V-1414

V-1411

V-1404

CV-1403

V-1402

V-1416

FM-1418

QD-1401

LGT-1432

Light Source VID-1433

Video Camera

F-1424

(25-μ Absolute)

0.50 in OD× 0.035 in Wall

LH2

Storage Tank

0.50 in OD× 0.035 in Wall 0.50 in OD× 0.035 in Wall

0.50 in OD× 0.035 in Wall

0.50 in OD× 0.035 in Wall

To Transfer System

To/From Ullage

Transfer

From Helium

Pressurization

System To Vent System

Figure 23. LH2 storage subsystem schematic.

2.3.2.5.8 Ground Loading Operation. The LH2 storage subsystem is loaded with LH2

using a T–0 umbilical on the launch pad. The ground interface quick disconnect (QD-1401) is 

engaged with the ground half disconnect and 28 VDC is applied to the LH2 fill and drain valve 

(V-1402) to allow LH2 to be loaded into the system until the pressure is ≈20 psia. Once the LH2

is loaded, power is removed from V-1402 and the ground half quick disconnect is demated from 

QD-1401. LH2 is not loaded into the LH2 receiver tank while on the ground.
41

2.3.2.5.9 Liquid Hydrogen Storage Tank Assembly. Contained within the LH2 storage 

system is the storage tank, which is designed to be a modular system containing all tank valving, 

LAD, helium gas diffuser, internal mass gauging sensors, and thermal controls. A diagram of the 

components included in the LH2 tank assembly can be seen in figure 24.

Cryocooler and BAC

Circuit

Figure 24. LH2 storage tank assembly diagram.

2.3.2.5.10 Payload Passive Thermal System. To affect reduced propellant boiloff demonstrations, passive and active thermal control design features are incorporated, including a TVS, 

active propellant mixing, MLI, low heat leakage structural support struts, and reduced fluid penetrations for the main storage propellant tank. The TVS utilizes a Joule-Thomson (J-T) expansion 

device for vented gas cooling which, in conjunction with a heat exchanger, provides cooling of the 

stored LH2 and pressure control. Propellant is acquired for the TVS function under settled and 

unsettled acceleration conditions via surface tension LADs installed inside the tanks, which also 

provide propellant for planned propellant transfer operations. Penetrations into the propellant 

tanks are isolated with close-coupled components to preclude propellant migrations into the lines. 

These include the vent system burst disk/relief valve and vent valves, the pressurization check valve, 

the ullage transfer valve, the TVS control valve, the LH2 fill and drain valve, the tank isolation 

valve, the transfer check valve, and the circulation isolation valve. Composite/metallic hybrid struts 

are used to thermally isolate while structurally supporting both the storage and transfer propellant 

tanks. To provide acceptable conditioning of the storage tank during the ground and ascent phases, 

the tank is encased in a layer of spray-on foam insulation (SOFI), and the MLI is actively purged 

with GN2 (to prevent moisture condensation within the MLI). Only the storage propellant tank 

is loaded with LH2 for the ground and ascent phases of the mission.
42

2.3.2.5.11 Payload Active Thermal System. The CPST payload fluid system also employs 

a cryocooler-BAC network architecture to actively refrigerate the external surface area of the storage propellant tank. This intercepts heat entering into this tank through the MLI over the surface 

area of the tank, reducing hydrogen propellant boiloff. Due to the low technology readiness of 

a 20 oF cryocooler system, the CPST payload employs a state-of-the-art 90 K cryocooler architecture. Because of this, the payload realizes an RBO design implementation for the LH2, as opposed 

to a zero boiloff (ZBO) design. Cryocooler options include pulse tube, Stirling, and reverse turboBrayton cooler designs. The pulse tube and Stirling (regenerative) coolers require a gas circulator to 

supply the BAC network. The reverse turbo-Brayton (recuperative) cooler can utilize the inherent 

circulation loop; therefore, a separate circulator is not required. Vibration of any of these cryocooler systems is minimal. The cryocooler refrigerant, either helium or neon at 90 K, is circulated 

through an array of small-diameter cooling tubes embedded within the MLI on the storage tanks. 

These tubes, mounted to a thin aluminum foil shield, which encapsulates the total surface area of 

the tank (as mounted inside the MLI), comprise the BAC. The BAC is mounted within the MLI 

layers such that it effectively intercepts heat conducted through the MLI (at a position predicted 

to be approximately at the 90 K temperature of the refrigerant). Options are under consideration 

to additional heat interception by active cooling of the structural support struts and the fluid and 

instrumentation penetrations via straps to the BAC.

2.3.2.5.12 Payload Fluid System Liquid Hydrogen Transfer. To move LH2 from the storage 

tank to the transfer tank, there are connecting lines, valves, and pumps to accomplish this transfer. 

Due to the need to use separate exit and entry connections to the tanks and minimize design complexity, unidirectional flow paths were used for the transfer subsystem. A bi-directional configuration would have been possible, but with a more complicated valve arrangement. 

It can be seen in the LH2 transfer subsystem schematic, shown in figure 25, that there 

are parallel flow paths per flow direction. These parallel paths are needed to demonstrate both 

pressurized and pump-fed transfer modes.
43

!

0.50 in OD×0.035 in Wall

0.50 in OD×0.035 in Wall

0.50 in OD×0.035 in Wall

0.50 in OD×0.035 in Wall

0.50 in OD×0.035 in Wall

0.50 in OD×0.035 in Wall 0.50 in OD×0.035 in Wall

V-1601

P-1703

V-1600

V-1700

V-1702

P-1701

To/From Cryogenic

Storage System

To/From Cryogenic

Storage System

Figure 25. LH2 transfer subsystem schematic.

2.3.2.5.13 Alternative Configurations. Early in the design study the vehicle contained 

an automated fluid coupling (AFC) that would have been used to demonstrate mating and demating of a coupling in space, then transfer propellants through the coupling with zero leakage. Incorporating this coupling in the current configuration would have resulted in a portion of the transfer 

system with bi-directional flow, since adding a second AFC may not have been desired for cost 

reasons. Adding a bi-directional flow portion would require many more two-way valves to route 

the fluid to the proper tank. 

Therefore, an alternative configuration was created that minimized the total number of 

components to accomplish propellant transfer. The alternative transfer system, which can be seen 

in figure 26, has three-way valves to transfer fluid either by pressurized, pump-fed, or across AFC 

methods. The valve configuration allowed for bi-directional flow in almost the entire system, except 

across the transfer pump, which required unidirectional flow to operate.
44

Cryogenic Storage Tank Cryogenic Transfer Tank

V-1600

V-1601

AFC-1800

V-1602

V-1603

V-1700

V-1701

V-1702

Figure 26. Alternative LH2 transfer system schematic—cryogenic transfer

subsystem (three-way valve configuration; simplified rep).

This system was not chosen for the baseline concept due to the strong desire to use common 

valves throughout the vehicle to minimize DDT&E costs. The use of this three-way valve would 

be a higher risk than a standard two-way valve due to the need to create a new component design 

for this program.

2.3.2.5.14 Payload Instrumentation. Instrumentation is robustly provided in the CPST 

payload to assure thorough measurement of thermal conditioning, transfer, and quantity gauging 

aspects of the flight experiment and employed technologies. Redundancy is provided for the critical sensors needed to prove technology performance advancement in order to validate predictive 

analytical models and validate mission objectives. Specifically, on both storage and transfer tanks, 

the following instrumentation is provided:

• 90 external temperature measurements.

• 16 internal temperature measurements, also used for liquid level sensing, silicon diode devices.

• Four TVS JT/mixing internal temperature measurements.

• 18 external and 6 internal cryocooler/BAC temperature measurements.

• Two tank pressure measurements.

• Settled and unsettled mass gauging (e.g., CryoTracker®, radio frequency mass gauging (RFMG), 

and possibly capacitance probe).

• Video on tank internal and outflow legs for propellant positioning and flow insight. 
45

On the pressurization and transfer systems, the following instrumentation is provided: 

• 42 external temperature measurements.

• Flow rate measurements on all propellant flow paths between tanks and through vent legs.

• Operation measurements on all valves and pumps.

• Up to 20 leak detection devices (electronic devices with capability for direct hydrogen presence 

detection).

• Redundant tri-axial accelerometer measurements for settling/unsettled acceleration rates.

The CPST payload design employs mass gauging techniques for settled and unsettled propellant orientations. Settled mass gauging includes a CryoTracker device using silicon diode devices 

to sense the liquid level within the tanks, with the option to include full tank range capacitance 

probe as backup. Unsettled mass gauging employs an RFMG, which detects tank frequency mode 

responses and matches characteristics to analytically predicted responses to determine propellant 

mass levels (under unsettled and settled acceleration states). For additional data for correlation, 

mass flow rates will be measured on all propellant transfer and vent legs, with positioning of flow 

rates for best results regarding thermal conditioning of propellant passing through these devices.

2.3.2.6 Payload Supporting Analyses. Several conceptual sizing analyses were performed 

for the CPST payload fluids system. These include pressure budget sizing to support pressure level 

definition and component sizing activities, and helium consumable mass sizing to support highpressure helium bottle sizing and envelope definition. Additionally, a conceptual layout assessment 

for tank penetrations and mounted components is provided (in part, supporting the rationale for 

the CPST payload propellant storage tank diameter).

2.3.2.6.1 Fluid System Pressure Budget—Introduction. The CPST technology demonstration will transfer LH2 from the storage tank to the transfer tank and back using pressure and 

pump-fed transfers. The pressure budget for each type of transfer is assumed to be the same flowing from or to the storage tank. In order to mix the propellant in the tank, there is also a recirculation loop on the LH2 storage tank. A preliminary pressure budget was developed to understand 

design limitations required to support the transfer and mixing functions. 

2.3.2.6.2 Propellant Transfer System. Two sets of transfer lines run from the storage tank 

to the transfer tank as shown in figure 27. This allows the propellant to be pulled from the LAD 

in one tank and distributed outside the LAD in the other tank. All of the transfer lines are sized 

at 0.5 in (1.27 cm) outer diameter (OD) with a wall thickness of 0.035 in (0.0889 cm). Each transfer line has a path where the transfer can utilize the pump or bypass the pump. The line length 

was conservatively estimated at 15 ft (4.572 m). In each line, there are three isolation valves. The 

first isolation valve is close coupled with the tank while the other two isolation valves allow flow 

through either the open line for pressure-fed transfer or the line with the pump for pump-fed transfer. For a given transfer, the propellant will only flow through two of the three isolation valves. The 

cryogenic isolation valves are assumed to have an equivalent sharp edge orifice diameter (ESEOD) 

of 0.278 in (0.706 cm). The ESEOD was provided by the MSFC components group, ER33, and is 

representative of a common off-the-shelf cryogenic valve. There is a filter at each tank outlet and 
46

Transfer Tank

Subsystem

Helium

Pressurization

Subsystem

Vent

Subsytem

Storage

Tank

Subsystem

Transfer

Subsystem

Figure 27. CPST fluid system schematic.

a check valve at each tank inlet. Losses for other components and lines were based on nominal 

resistance coefficients. These loss coefficients will need to be updated as better component and 

system information becomes available.

2.3.2.6.3 Propellant Transfer Pressure Budget. The propellant will be transferred by two 

means—pressure fed or pump fed. For each type of transfer, the expected steady-state flow rate 

is about 78 lbm/hr (0.0098 kg/s). This flow rate equates to a velocity of ≈5 ft/s through the transfer 

line. During all transfers, the propellant is assumed to be saturated at 20 psia (137.9 kPa) which 

corresponds to a temperature of –421 oF (21.5 K). 
47

The pressure-fed transfer pressure budget is shown in table 21. The propellant tank is 

assumed to be pressurized to 25 psia (172.4 kPa) with helium pressurant. The total pressure drop 

through the transfer line and transfer line components is ≈0.5 psid. The minimum propellant tank 

maximum design pressure (MDP) is 40 psia (275.8 kPa) for pressurized transfer or 15 psi above 

the propellant tank ullage pressure. 

Table 21. Pressure-fed system pressure budget.

Pressurized Transfer Hydrogen

Mass flow rate 77.9 lbm/hr

Velocity 4.9 ft/s

Estimated transfer line length 15 ft

Storage tank ullage pressure 25 psia

Line ∆P (tank outlet/inlet/bends/ 

 fittings/check valve)

–0.2 psid

Valve ∆P (two valves) –0.3 psid

Transfer tank inlet line pressure 24.5 psia

Propellant tank MEOP/MDP 40 psia

The pump-fed transfer pressure budget is shown in table 22. The propellant tank is assumed 

to be at 20 psia (137.9 kPa) with saturated propellants in the tank. The total pressure loss through 

the transfer line and components is estimated to be 0.5 psid. The pressure increase required by the 

pump was calculated as the pressure loss in the system at the expected flow rate with a 30% margin 

for growth and flexibility. The minimum propellant tank MDP for pump-fed transfer is 35 psia 

(241.3 kPa).

Table 22. Pump-fed system pressure budget.

Pump Transfer Hydrogen

Mass flow rate 77.9 lbm/hr

Velocity 4.9 ft/s

Estimated transfer line length 15 ft

Storage tank ullage pressure 25.0 psia

Line ∆P (tank outlet/inlet/bends/

 fittings/check valve)

–0.2 psid

Valve ∆P (two valves) –0.3 psid

Pump ∆P (P2-P1)

Transfer tank inlet line pressure 20.2 psia

Propellant tank MEOP/MDP 35 psia
48

2.3.2.6.4 Propellant Recirculation System. As shown in figure 27, the propellant tank 

recirculation system pulls propellant out of the tank LAD with the recirculation pump and sprays 

back into the tank to mix the propellant and ullage. The recirculation line is 0.5 in (1.27 cm) OD 

with a wall thickness of 0.035 in (0.0889 cm). The line length was conservatively estimated to be 

15 ft (4.572 m). There is a single isolation valve with an estimated ESEOD of 0.278 in (0.706 cm) 

(representative of a common off-the-shelf cryogenic valve). Other component loss coefficients were 

estimated based on nominal resistance coefficients. These coefficients will need to be updated as 

component and system information becomes available. 

2.3.2.6.5 Propellant Recirculation Pressure Budget. The tank recirculation loop will be 

driven with a pump. The expected steady-state flow rate is about 244 lbm/hr (0.0307 kg/s). During 

the recirculation, it is assumed that the propellant will be saturated at 20 psia (137.9 kPa), which 

corresponds to a temperature of –421 oF (21.5 K). 

The TVS pressure budget is shown in table 23. The total pressure through the transfer line 

and components is estimated to be 2.9 psid. For the pump, the change in total pressure was calculated as the pressure loss in the system at the expected flow rate with a 30% margin for growth and 

flexibility. 

Table 23. TVS system pressure budget.

TVS Pump Mixing Hydrogen

Mass flow rate 243.8 lbm/hr

Velocity 15.5 ft/s

Estimated transfer line length 15 ft

Storage tank ullage pressure 20 psia

Line ∆P (tank outlet/inlet/bends/

 fittings)

–1.4 psid

Valve ∆P (one valves) –1.5 psid

Pump ∆P (P2-P1) 3.8 psid

TVS mixing flow line outlet 

 pressure

20.9 psia

The system pressure budget needs to be reevaluated as the system design matures. The final system 

MDP was set to 50 psia (344.7 kPa) to allow for system changes. 

2.3.2.6.6 Helium Tank Sizing Analysis—Introduction. Helium will be used as the pressurant for the LH2 pressure-fed transfer demonstrations on the CPST technology demonstration. 

The quantity of helium required depends on the ullage volume of the propellant tanks during the 

transfer, the required pressure for the transfer, and the heat transfer between the helium pressurant 

and the fluid in the propellant tank. 
49

2.3.2.6.7 Propellant Transfer. There are two transfer series, each consisting of settled 

pressure-fed transfer, settled pump-fed transfer, unsettled pressure-fed transfer, and unsettled 

pump-fed transfer. For each transfer, the propellant is sent to the transfer tank and then returned 

to the storage tank. The maximum ullage pressure driving the transfer is assumed to be 30 psia. 

For pressure-fed transfers, the storage tank will be pressurized and maintained until the propellant 

transfer is complete, filling the transfer tank ≈90%. The storage tank will then be vented overboard 

to a nominal pressure. The transfer tank will then be pressurized and maintained until the propellant is transferred back to the storage tank. The transfer tank will then be vented overboard. At no 

time during this process will any of the pressurant be recovered. No pressurant will be used during 

the pump-fed transfers.

Figure 28 shows the propellant inventory during the mission timeline based on thermal 

analysis. Figure 29 shows the corresponding storage tank ullage. This ullage provides the storage 

tank volume that must be pressurized during transfers to the transfer tank. The total transfer tank 

volume is required to be pressurized to return propellant to the storage tank. 

600

500

400

300

200

100

0

Propellant Mas (lbm)

0 0.5 1 1.5 2 2.5 3

Mission Time (mo)

3.5 4 4.5 5 5.5 6

Transfer

Demonstration

Transfer

Demonstration

Passive CFM

Storage

Demonstration

Active CFM Storage

Demonstration

Figure 28. Assumed propellant inventory over the mission timeline.
50

100

90

80

70

60

50

40

30

20

10

0

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5 5.5 6

Mission Time (mo)

Storage Tank Ullage (%)

Passive CFM

Storage

Demostration

Transfer

Demostration

Active CFM Storage

Demostration

Transfer

Demostration

Figure 29. Assumed storage tank ullage percentage over the mission timeline.

During pressurization, the helium is assumed to cool to the liquid temperature. For this 

analysis, the LH2 is assumed to be saturated at 20 psia. The cooling of the pressurant is a conservative assumption for determining helium requirements, although it may not be the case for larger 

ullages. Even though the helium is depositing energy into the propellant tank, the distribution 

of this energy is unknown and therefore cannot be assumed to assist in tank pressurization. The 

calculated pressurant required to compete the two transfer series with the assumptions made is 

≈53.5 lbm of useable pressurant.

Helium pressurant bottles—The pressurant bottles are composite overwrap pressurant vessels (COPVs) with a maximum expected operating pressure (MEOP) of 4,500 psia. The tanks are 

assumed to be loaded at 4,050 psia (90% of MEOP) to allow for the relief valve variability during 

loading. The nominal ground-loading temperature is 70 oF. 

In flight, the pressurant tanks will be held at a storage temperature of –20 oF. This temperature was selected due to the soft goods heritage temperature capability in the bang-bang pressurization valves. Heaters on the COPV will be used to keep the pressurant at the storage temperature. 

This temperature corresponds to an initial pressure of 3,370 psia in the COPV. During pressurant 

usage, the helium temperature can fall below the storage temperature (–20 oF) but the assumption is that there is adequate time to thermally recover to the storage temperature prior to the next 

transfer. 
51

The system model is a simplistic, conservative spreadsheet that assumes the fluids are at 

bulk conditions. The storage tank ullage is calculated from the tank volume and the propellant 

inventory for each transfer. The change in ullage volume is then calculated based on the mass 

required for chill-in of the feed system, transfer tank, and the quantity of transferred propellant. 

This provides the ullage volume that must be pressurized. For the transfer back to the storage tank, 

it is conservatively assumed that the entire volume of the transfer tank was required to be pressurized. Knowing the pressure, temperature, and volume required to be pressurized, the helium mass 

required for pressurization was calculated. This mass is then removed from the helium COPV. A 

higher fidelity model is required to adequately predict propellant tank ullage conditions, which may 

allow a reduction of the required helium. 

Figure 30 shows the helium COPV tank pressure over the mission timeline. No helium leakage was assumed. The minimum allowable pressure in the COPV is assumed to be 250 psia. A total 

volume required to provide the necessary useable helium is 24 ft3 with a total helium mass loaded 

at 60.5 lbm. This volume equates to a 3.6-ft-diameter sphere. 

4,500

4,000

3,500

3,000

2,500

2,000

1,500

1,000

500

0

0 0.5 1 1.5 2 2.5 3

Mission Time (mo)

3.5 4 4.5 5 5.5 6

Loaded at 70 °F Active CFM Storage

Demonstration

Transfer

Demonstration

Transfer

Demonstration

Passive CFM

Storage

Demonstration

Stored at –20 °F

Helium Tank Pressure (psia)

Figure 30. Helium pressure over the mission timeline.

Two tank configurations were proposed based on vendor studies; one has six COPVs while 

the other has eight. The information for the two configurations is in table 24. Both configurations 

have the same total volume—the difference is in the packaging and weight.
52

Table 24. Configuration options.

Number of Tanks 6 8

Total length to end to end

 (external)

36.04 in 28.15 in

0.915 in 0.715 in

Outer diameter

 (external)

17.7 in 17.7 in

0.450 m 0.450 m

Total internal volume 

 (per tank)

6,912 in3 5,184 in3

0.113 m3 0.085 m3

Mass (per tank) 46 lbm 35 lbm

20.9 kg 15.9 kg

2.3.2.6.8 Payload Storage Tank Component Layout Assessment. An assessment was performed to evaluate packaging of planned components on the end domes of and internal to the 

propellant storage tank in order to validate the reasonableness of the CPST payload storage tank 

diameter. This conceptual packaging assessment was made under the constraint of minimal maturity associated with these components and also the assumption of providing access to these components for repair or replacement without impacting the assembled payload system. To achieve 

this latter assumption, the conceptual configuration installed the various hardware items on cover 

plates for the top and bottom domes, which then are installed on these domes. Additionally, some 

of the identified hardware items are assumed to be removable without the removal of the integrated 

cover plate assemblies. Also, pending future stress analysis maturation, the guideline was assumed 

to minimize the discontinuity angle between the cover and the tank dome, ideally to an angle no 

more than 10 deg. 

To support these installation operations, there is a 2.5-in, 360-deg clearance around all 

ports/hardware installation covers for structural integrity, manufacturability, and operability. 

Other significant assumptions include:

• A 1-in area around the manhole cover for a flat sealing surface.

• A 4-in camera viewing port.

• A 2-in camera light source port.

• The diffuser is preinstalled on the cover.

• Temperature/wet-dry liquid sensor racks are installed through the 2-in port on the cover.

• The TVS heat exchanger is preinstalled on the cover.

• The spray bar and axial jet mixing devices are installed through the 2-in port.

• The LAD has a 6-in sump/basket and 2-in gallery channels.

• The mixing and transfer pumps are mounted on the outside of the tank cover. 

• The RFMG requires at least a 2.75-in port so a 3-in port is provided.

The capacitance probe is not included in this layout assessment. Also, it is assumed that the BAC 

tubing manifold does not interface with the covers.
53

The conceptual cover to dome interface, as a function of cover size and using the CPST 

payload storage tank diameter of 60 in (5 ft) and elliptical end dome geometry, is shown in 

figure 31.

45

40

35

30

25

20

15

10

5

0

–30 –25 –20 –15 –10 –5 0 5 10 15 20 25 30

Dome Height (in)

Dome Diameter (in)

25-in Cover

Discontinue Angle and Plasticity

Due to Cycles Still in Work

20-in Cover

30-in Cover

Figure 31. CPST payload 5-ft storage tank end dome—cover geometry.

The list of conceptual components to mount to the end domes of the storage tank is 

provided in table 25. Conceptual layouts of these components are shown for the top and bottom 

dome cover plates in figures 32 and 33.

Table 25. CPST payload storage tank end dome and internal components.

No. Component Top Dome Bottom Dome

1 LAD ×

2 Mixing pump ×

3 Transfer pump ×

4 Diffuser ×

5 Temp rack and wet/dry rack ×

6 Camera view port ×

7 Camera light port ×

8 Spraybar and axial jet ×

9 Heat exchanger ×

10 RFMG ×
54

2-in Mixing Pump Port

20-in Diameter

Temp/Wet Dry Racks With

2-in Port

15-in Diameter Cover Placement Violates

the 2.5-in 360° Clearance Around the Ports

3-in Diffuser With 2-in Port

1-in Flat for Seal

3-in RFMG Port

LADs

2-in Press Port

4-in Camera View Port

Port/Fitting

Hardware Outline

Figure 32. CPST payload tank top dome cover conceptual layout.

2-in Transfer Pump Port

25-in-Diameter Cover Violates

the 2.5-in, 360° Clearance

Around the Ports so Would

Need at Least 26-in Cover

6-in Heat Exchanger 

With 2-in Port

15-in-Diameter Cover Placement 

Violates the 2.5-in, 360° Clearance 

Around the Ports

1-in Flat for Seal

2-in Propellant Fill/Outflow Port

1-in Spraybar Axial Jet

With 2-in Port

LAD With 6-in Sump 

and 2-in Channels

2-in Camera Light Port

Port/Fitting

Hardware Outline

Figure 33. CPST payload storage tank bottom dome cover layout.
55

Based on the updated guidelines and assumptions, component list, and if the component 

layout is the deciding factor, the storage tank would need to have a diameter of at least 60 in. This 

allows for the current component list to be accommodated in a 25-in cover. The discontinuity angle 

and plasticity requires further evaluation for structural integrity; it is noted that the 25-in cover and 

60-in-diameter tank does not meet the 10-deg discontinuity goal. 

2.3.2.6.9 Payload Fluid System Forward Work. As part of future maturation of the CPST 

payload fluid system, several key areas of design effort have been identified. These include, but are 

not limited to, propellant transfer tank components and interfaces, and resultant packaging and 

size; conceptual design maturation for the cryocooler circuit (design, component selection, and 

assembly), and for the BAC design (line and manifold configurations, mounting, and assembly, 

active cooling of structural, fluids and instrumentation penetrations, and leakage assessment and 

potential refrigerant resupply approaches); trades regarding transfer operations (flow rate and 

capability for variation, planned sequences, chill-in approaches, and venting operations).

2.3.3 Structural Analysis

2.3.3.1 Ground Rules and Assumptions. The standard ground rules and assumptions for 

NASA in-space vehicles were used per NASA-STD-5001a.3 The prototype verification approach 

was assumed. The metallic factors of safety for yield and ultimate were 1 and 1.4, respectively. The 

composite ultimate design factors for acreage and discontinuities were 1.4 and 2, respectively. The 

proof pressure factor of 1.5 and burst pressure factor of 2 were used for all pressurized vessels. 

A local panel-buckling knockdown of 0.65 was used on all panels. The minimum allowable frequency was assumed to be 25 Hz, which is standard for payloads on the range of launch vehicles 

under consideration during the study.

2.3.3.2 Analysis Methodology. The standard structural analysis procedure from the 

Advanced Concepts Office at MSFC for prephase A studies was employed. This process begins 

by obtaining the CAD configuration model. This model includes all the primary structures, all 

pressure vessels, and any large supporting structures. The CAD model is imported into FEMAP, 

a Windows-based pre- and post-processing program designed for constructing finite element 

models (FEMs). Once the FEM is constructed, the NX Nastran structural solver is employed to 

determine the stresses and deflections in all structural members. These results are then imported 

into HyperSizer, a structural optimization program, to determine the lightest structure possible, 

given the configuration. A subprogram within HyperSizer, HyperFEA®, updates the structural 

properties based on the optimization, and then reruns Nastran to obtain new loads. This process 

is repeated within HyperFEA until the loads and stresses converge, leading to a properly optimized 

structural mass.

Any required spacecraft adapter is included in the FEM to properly constrain the model. 

However, the adapter is purposely oversized, driving more loads into the CPST structure. This 

helps lead to a conservative mass result.

For the current study, the structures were sized for strength, stiffness, and local buckling. 

A global buckling analysis was not performed due to time constraints. However, for payloads of 

the size of CPST, global buckling is typically not a controlling load case.
56

2.3.3.3 Structural Description. Since the primary focus of CPST is based on a cryogenic 

test fluid, it was determined that the storage and transfer tanks would be nonstructural pressure 

vessels. These tanks would be structurally supported via composite struts only, thus limiting the 

heat transferred into the tanks through conduction. This led to an outer panel construction with 

appropriately placed ring frames from which the tanks could be mounted. All structures, other 

than the composite struts, were assumed to be aluminum for cost purposes. Figure 34 shows the 

CPST payload configuration.

Figure 34. CPST payload configuration.

The structural analysis was performed in parallel with the launch vehicle selection. A list 

of candidate launch vehicles was available from which the worst-case constraints were selected. The 

configuration was designed to fit within the dynamic envelope of the Minotaur IV or Taurus XL 

shroud, leading to an OD of roughly 2 m. One-foot spacing between the outer shell and tank wall 

was assumed to allow room for tank insulation, test instrumentation, wiring, etc. These constraints 

were used to calculate the required tank dimensions. 

While a small shroud was used to size the dimensions of the CPST concept, the worst-case 

launch loads from the launch vehicles under consideration were determined to be from the Falcon 9. The maximum axial loads with a simultaneous lateral component were applied to the FEM. 

Due to the symmetric nature of the configuration, the lateral loads were only applied at orientations of 0, 45, and 90 deg. This led to three load cases for each of the load conditions shown in 

table 26.
57

Table 26. Falcon 9 load conditions.

Load Condition Axial g’s Lateral g’s

Maximum axial (compression) 6.5 0.5

Maximum axial (tension) –2 0.5

Maximum lateral (compression) 3.5 2 

Maximum lateral (tension) –1.5 2 

For both the storage and transfer tanks, the test instrumentation was also being determined 

in parallel with the structural analysis. The uncertainty in mounting the test instrumentation led to 

a decision to force one-quarter of both tanks to have a 0.300-in wall thickness. This wall thickness 

was believed to be adequate for any mounting needs of the instrumentation. Applying the structures ground rules and assumptions and methodology described in this section with the worst-case 

loads led to the prephase A FEM shown in figure 35.

Figure 35. CPST finite element model.
58

2.3.3.4 Analysis Results. The optimized results for the primary structures and tank attachment struts were obtained through an advanced prephase, a structural analysis. Secondary structures are estimated to have a mass that is 20% of the primary structure mass based on previous 

studies and as-built masses for other space structures. The same rational is used to estimate the 

mass of the joints and fittings for the struts at 50% of the mass of the struts themselves. These estimates are added to the master equipment list (MEL). The storage and transfer tank are sized using 

the FEM, but they are book-kept under a different portion of the MEL (not within the structures 

section of the MEL). They are included in table 27 for completeness. Due to the preliminary nature 

of the study, and the lack of maturity of full-scale flight requirements, a 30% contingency has been 

added to the basic mass to calculate the predicted mass of the structural components of CPST.

Table 27. Structural analysis mass results.

CPST MEL Qty.

Unit Mass 

(kg)

Basic 

Mass 

(kg)

Contingency

(%)

Contingency 

(kg)

Predicted 

Mass

(kg)

Pathfinder

1.2.3 Storage tank 1 108.78 108.78 30 32.63 141.41

1.2.7 Transfer tank 1 14.30 14.30 30 4.29 18.59

5.0 Structures 150.53 150.53 30 45.16 195.69

5.1 Primary structure 1 112.47 112.47 30 33.74 146.21

5.2 Secondary structure 1 22.49 22.49 30 6.75 29.24

5.3 Tank attachments 1 10.38 10.38 30 3.11 13.49

5.4 Joints and fittings 1 5.19 5.19 30 1.56 6.75

2.3.4 Thermal Analysis

Thermal control is a major issue in the design of CPST. The primary concern is to minimize 

the heat input into the cryogenic system. This section provides details of the thermal design and 

analysis performed for CPST including orbital attitude, propellant storage tank heat loads (tank 

penetrations, support structure, and insulation system), and required heat rejection capacity and 

propellant loss rate.

2.3.4.1 Top-Level Ground Rules and Assumptions. Table 28 contains the ground rules 

and assumptions relating to cryogenic thermal management of the LH2 experiment fluid and 

experiment thermal control. A key design decision to maintain the spacecraft attitude with the bus 

towards the Sun was made on the basis of thermal impact on the cryogenic experiment. CFM is 

critical to the successful completion of the experiment timeline and relies on a number of important ground rules and assumptions, including ground replenishment and gaseous nitrogen (GN2) 

purge for the LH2 tank until ≈T=0, a nonstructural, composite strut supported tank design, and 

active/passive thermal control features on the supply tank.
59

Table 28. Ground rules and assumptions.

3.0 CFM GR&A

CFM GR&A Ground replenishment of cryogens Ground replenishment until ~ T–0

Vehicle structural/thermal design Similar to CPS nonstructural tank concept

Cryogenic fluid management design Similar to COLDSAT plus active system

Target storage conditions Normal boiling point

CFM performance demos Active and passive CFM

Transfer modes for transfer demos Pressure-fed and pump-fed

Thrust accelerations for transfer demos Zero thrust, 0.00003 g axial, 0.001 g axial

Other demos Mass gauging, liquid acquisition

Ground purge Dry GN2 purge provided for LH2 and LO2 tank MLI util ~T–0

Tank venting during ascent Active venting to relieve pressure

Post-ascent thermal transient Passive CFM initiated ASAP to assess thermal transient

Margin on all heat loads 50%

7.0 Spacecraft Thermal Control

Thermal GR&A Thermal control system Thermal control of the spacecraft shall utilize standard, flightproven materials such as MLI, selected surface finishes, foils 

and tapes, conduction (coupled/decoupled mounting details), 

optical solar reflectors, resistance heaters, thermostats, and 

controllers to maintain acceptable spacecraft subsystem 

component temperatures with adequate margins during all 

mission phases

Thermal isolation Spacecraft bus and CFM experiment subsystems electronics 

shall be isolated to the maximum extent possible from the 

cryogenic (experiment) portions of the spacecraft

Heat rejection Deployed radiators, optical properties a=0.1, ε=0.85, areal 

density = 8.5 kg/m2

Attitude control Controlled so that the aft end of the spacecraft is always 

facing the Sun in order to minimize heating of the cryogenic 

systems by direct solar flux

2.3.4.2 Analysis Methodology. A geometric representation of the CPST experiment system 

including vehicle and propellant tank dimensions from the structures and configuration subsystem 

experts was developed using Thermal Desktop®. Environmental radiative effects were analyzed 

using RadCAD, which is integrated into Thermal Desktop. The vehicle orbit is circular about the 

Earth at an altitude of 500 km and a beta angle of 63 deg; this is assumed to be a worst-case hot 

environment. This orbit is shown pictorially in figure 36. The vehicle orientation is solar inertial 

such that the aft end of the spacecraft (bus end) is always facing the Sun in order to minimize heating of the cryogenic systems by direct solar flux.
60

Figure 36. CPST orbital plot.

The thermal analyzer utilized for the analysis is SINDA/FLUINT, also integrated into 

Thermal Desktop. Environmental boundary conditions, radiator temperatures, and heat leak into 

the storage and transfer tanks due to structural attachments were determined from the vehicle 

thermal model.

The active cryogenic storage analyzer tool was used to estimate mass and power for the 

fluid storage tank insulation and other hardware used to maintain in-space cryogenic propellants 

using active cooling. An RBO system was designed for the CFM active system on the supply tank, 

which uses a 90 K cryocooler of 20 W cooling capacities and a BAC network to intercept heat at 

the actively cooled radiation shield imbedded within the hydrogen tank insulation. Helium gas is 

circulated through tubing that is bonded to the low mass foil shield.

The passive cryogenic storage analyzer (PCSA) tool was used to optimize and estimate 

thermal insulation required for the LH2 transfer tank, which will utilize only passive thermal 

control features. 

Thermal control of flight avionics is the responsibility of the spacecraft bus. Experiment 

avionics are located between the cryogenic experiment system and the spacecraft bus and rely on 

active and passive thermal control techniques. Heaters are used in the electronics bay to assure that 

minimum temperature requirements are met when the equipment is powered down. The electronics 

bay is isolated to the maximum extent possible from the cryogenic experiment.
61

2.3.4.3 System Description. The CPST CFM/thermal system is shown in figure 37. Tailoring component layouts, surface finish selection, tank supporting details, and heater sizing to 

maintain experiment and avionics temperatures developed the thermal control system. The thermal 

model incorporated white paint on the exterior of the spacecraft. Supply and transfer tank struts 

were sized for aluminum, and composite properties were used in the model. Pressure tank temperatures are maintained using MLI and heaters.

Spacecraft

Structure

• White Paint • Composite Struts • Isolators

Pressurant

Tanks

• Heaters • Isolators • MLI

Avionics System

• Surface Finish • Isolators

• Heaters • MLI • Coldplates/Pumped Coolant

Transfer Tank

• Heaters • MLI • TVS

Transfer System

• MLI • Heaters • Relief Valves

Supply Tank

• Cryocooler • BAC Shield • Helium System • HEXs

• Radiator/Coolant System • MLI/SOFI/TVS

Figure 37. CFM/thermal system description.

The flight avionics are located on the spacecraft bus and are outside the scope of this 

section. CPST avionics consists of components required to support the payload experiments/demonstrations only and includes an experiment computer and data unit, a video system, and a thermal control system. This equipment is located between the experiment and the spacecraft bus and 

utilizes active cooling (coldplates and pumped fluid system) and passive thermal control (MLI, 

heaters, select surface finishes, and isolators) to maintain temperatures within an acceptable range. 

On-orbit power dissipation levels for the experiment avionics is shown in table 29.
62

Table 29. CPST avionics power dissipation.

Subsystem/Component Quantity

Power

(W)

Total 

Power

(W)

Experiment flight instrumentation and electronics

 Experiment computer and data unit 1 120 120

Video System

 Cameras (four total, only two active at a time) 2 2.4 4.8

 Video power and control 1 36 36

 Video data storage 1 65 65

Thermal control system

 Pump control unit 1 56 56

 Heater controllers 2 7 14

Totals 295.8

The supply tank is outfitted with 90 layers of MLI, SOFI, a BAC shield, helium system, 

heat exchangers, TVSs, and a radiator system for heat rejection. The transfer system and transfer 

tank relies on MLI, heaters, and relief valves for thermal control. Heaters are used to raise the temperature of the transfer system between experiment transfers in order to simulate initial conditions 

in a timely manner.

2.3.4.4 Analysis Results. This section provides analytical results of the CFM system 

thermal design and analysis. 

Table 30 summarizes the mission details that were incorporated into the analysis. Propellant 

loss was calculated for active and passive storage, and each transfer demonstration based on tank 

and transfer system heating rates, as specified in table 30. 
63

Table 30. CFM/thermal summary.

CPST Govt 

POD LH2

Mission timeline

 Propellant load

 Transfer tank vol. (% of storage tank)

 Transfer duration (each)

 Active storage duration

 Passive storage duration

 Number of transfer demo sequences

260 kg

5%

15 min

5 mo

1 mo

2 (3 desired)

Propellant loss

 Propellant loss (active storage)

 Propellant loss (passive storage)

 Propellant loss (transfer demo, each)

 Propellant loss (active storage)

 Propellant loss (passive storage)

 Propellant loss (transfer demo, each)

8.2%/mo

9.3%/mo

20.7%

21.4 kg

24.2 kg 

53.9 kg 

Tank heating rates

 Spacecraft internal ambient temperature

 Storage tank heat load (passive storage)

 Storage tank heat load (active storage)

 BAC heat load (active storage, W)

 Transfer tank heat load (during transfer)

 Transfer system heat load (during transfer)

186 K

4.1 W

3.7 W

2.1 W

2 W

10 W

Tank insulation

 Storage tank No. MLI layers

 Transfer tank No. MLI layers

 Storage tank MLI thickness

 Transfer tank MLI thickness

 Storage tank SOFI thickness

 Transfer tank SOFI thickness

90

90

0.0896 m

0.0877 m

0.0127 m

0.0000 m

CFM subsystem mass

 Storage tank active CFM subsystem mass

 Storage tank passive CFM subsystem mass

 Transfer tank passive CFM subsystem mass

27.9 kg

72.1 kg

30.7 kg

Vehicle thermal control

 Radiator effective temperature

 Total heat rejection

 Total radiating area

 Radiator mass

 Insulation and thermal control materials

244 K

931 W

9.3 m2

79.1 kg

 7.5 kg

Axial thrust for transfers and mass gauging:

Total duration at 0.001 g

Total duration at 0.00003 g

72 min

472 min

Fluid system parameters:

Transfer fluid velocity

Transfer mass flow rate

Transfer system length

Transfer system diameter

Storage tank mixing rate

1.48 m/s

0.0098 kg/s

3.9 m

0.0109 m

7 gpm
64

The thermal analysis temperature results are shown in figure 38 and indicate a spacecraft 

internal ambient temperature of 186 K. The radiator effective temperature is 244 K and required 

radiating area is 9.3 m2. The radiator is sized to accommodate experiment cryocooler and avionics 

heat rejection in a LEO environment. A 50% margin on power/heat dissipation has been included 

to account for growth in the preliminary estimates and a 50% increase in heat rejection area has 

been considered to accommodate reduced radiator capability due to micrometeoroid and orbital 

debris damage.

>292.9

Node

292.9

265.6

238.4

211.1

183.9

156.6

129.3

102.1

74.8

47.53

20.27

<20.27

Temperature (K), Time = 0 s

Figure 38. CPST temperature results.

The current solar array configuration is different from that represented in the thermal 

model, but impact to preliminary results should be minimal. Power requirements for the active 

and passive CFM, transfer tank, and transfer system warm-up, transfer pumps, and avionics heaters are shown in table 31. Thermal input power is shown in table 32.
65

Table 31. Propellant loss and heating rates. .

CPST Government 

POD Propellant 

Losses

Propellant

Loss

(kg)

Steady-State Heating Rates

Tot Q Tank

(W)

Str Q Tank

(W)

Pen Q Tank

(W)

Rad Q tank

(W)

Tot Q BAC

(W)

Rad Q BAC

(W)

Passive storage 24.2 4.13 2.50 0.61 1.02 – –

Active storage 

(reduced boiloff)

107.2 3.66 2.50 0.61 0.55 2.10 2.10

Transfer tank 

(includes chilldown 

and residuals)

 50.4 2.02 1.31 0.53 0.18 – –

Transfer system 

(includes chilldown 

and residuals)

 57.5 10.01 – – 10.01 – –

Storage tank residual 7.8 – – – – – –

Total propellant loss 247 – – – – – –

Table 32. CFM/thermal input power.

LH2 Operation

Power

(W) Duration Comments

Active CFM 51 5 mo Constant

Passive CFM Average 1.8 1 mo

Peak 111 5% duty TVS 5% duty

Transfer tank/transfer system warm-up 22.6 48 hrs After each two-way transfer (8 total)

Transfer Pumps 6.7 0.5 hr Each two-way transfer (8 total)

Thermal control Avionics heaters 25 6 mo Entire CFS avionics subsystem, 50% duty

Additional work is required to incorporate spacecraft bus, experiment avionics, and pressurant tank details into the vehicle thermal model. Evaluation of conductive isolation and heat rejection options for avionics, pressurant tanks, and cryocoolers is needed. Incorporation of active heat 

interception at the tank support structure is also included in future work.

2.3.5 Avionics, Communication, and Data Handling

2.3.5.1 Avionics Approach. This TM focuses on the results of the Pathfinder POD design, 

which has only an LH2 system; no LO2 system is included. The avionics system for the CPST 

demonstrator consists of all necessary command, data handling, and instrumentation required to 

control and monitor the various experiments of the mission. Components of the system include 

an experiment computer/data unit (ECDU), a pump control unit, two heater controllers, a video 

system, and cabling. No redundancy or fault tolerance is required for this mission, and all systems 

are single string. The plan for this POD study was to develop the payload independently of the 

spacecraft bus and integrate the two systems later in the development cycle. The two systems will 

be connected and communicate through a serial data bus, where an RS422 is suggested. Given this 

approach, the only avionics sized are those required to support the payload experiments. No space-
66

craft bus hardware is included. However, the experiment requirements were used to derive spacecraft bus requirements, which are also included in this TM. The spacecraft requirements were used 

to bound potential commerical off-the-shelf bus options. It is realized that a fully integrated payload with the spacecraft could be more efficient in mass, but saving would be in the order of tens 

of kilograms and would not impact the mission in that way. From a program administration perspective, independent experiment development makes more sense.

The spacecraft bus is to perform all stationkeeping operations, maneuvers, and ground 

communications including uplinks and downlinks. The payload ECDU computer will contain the 

experiment software and sequencing control, and order propulsive settling maneuvers to the flight 

computer as required for the various experiments. The payload avionics will store and forward 

experiment data, including video, to the spacecraft flight computer for downlink. The flight computer is to call experiment data from the ECDU based on ground link availability and memory 

capacity. The payload ECDU will contain sufficient memory capacity to store up to 24 hr of peak 

experiment data. It is believed that the data transmission rates of this mission are now well within 

standard capacities of both spacecraft and ground network systems, and that no memory overload 

should occur. 

In order to bound the data rate problem early in the study, a minimum and maximum video 

data rate was determined based on reference missions and video system demands. Assuming two 

out of four cameras (for the Pathfinder POD) would be running, a lower bound of 6 Mbps per 

camera for black and white and 15 Mbps per camera for color was assumed would be running. 

With these lower and upper bounds, it was determined by analysis that no pointing high-gain 

antenna would be necessary if transmission power were highly amplified (from 10 to 20 W) using 

X-band. Using S-band frequencies would require excessive amplification, and was thought impracticable at these data rates. After the video requirements were better defined, it was determined that 

even the lower data bound estimate was too high. A total data rate after compression is only 4.2 

Mbps for Pathfinder. Further link budget analysis showed that by using multiple ground stations, 

S-band could be used. One ground station would probably suffice using X-band. Without knowing 

the cost impact/comparison of the two options, both were carried forward in the documentation as 

ground communication options for the spacecraft bus supplier to decide. 

An experiment instrumentation list was compiled by the CFS definition group and was used 

to establish data rates required for sizing onboard data handling systems and downlink communication capabilities. Instrumentation includes resistance temperature sensors, silicon diode temperature sensors, pressure sensors, voltage and current sensors, accelerometers, wet/dry mass gauging, 

and RFMG sensors. Instrumentation also includes video recordings of fluid transfers from the 

propellant storage tank to the transfer tank and back again, along with monitoring the propellant 

state within the tanks. 

Sampling of temperature sensors was assumed to be at 1 Hz, while most other assumed 

sampling of sensors was at 10 Hz. Adding up the number of sensors at the appropriate sampling 

rates and multiplying by assumed bit counts required per sample gives the total data rate per sensor type. Bit counts were assumed to be either 8 or 12 per sample to include required resolution 

and formatting overhead. With the one fluid system of Pathfinder, there were over 300 temperature 
67

sensors alone. Although this housekeeping instrumentation data rate is <100 Kbps total, it can 

become significant over long periods of accumulation.

2.3.5.2 General Ground Rules and Assumptions. Because of the CPST mission class, payload avionics has no fault tolerance requirement. However, given the criticality of this technology 

development for in-space propulsion systems and for the future of space exploration in general, 

it would make sense to incorporate some fault tolerance if it has little impact on mission cost. Since 

the ECDU mass estimate is small, only 8 kg, some built-in redundancy can be accomplished without affecting mission performance. A backup processor board could increase mission assurance 

significantly while adding only a couple of kilograms to the ECDU mass.

Earlier in the study it was thought desirable to have a real-time uplink capability through 

the TDRSS communication system in order to be able to observe data and video, to alter a test in 

process, or simply shut a test down if there was a problem. After considering both the operational 

cost and spacecraft systems needed to communicate with TDRSS, including a gimbaled highgain antenna, it was decided this real-time feature was not worth the trouble. It was subsequently 

ground ruled that no real-time data link was required and that data would be stored and forwarded 

during ground station passes. From this point, it became a matter of how much time was available 

to downlink the data and video. Uplink commands would be required only when necessary to 

modify operation or troubleshoot. 

To keep operational costs down, an effort to utilize only one ground station was made. Early 

study analysis, with twice the data rate required, showed an X-band link would be required using 

an omni-directional antenna. However, with an extension of ≈7 days between test sets given and 

a reduction in data rates due to program descope, further ground link analysis showed that multiple 

S-band ground stations could accomplish the task. Using at least two S-band ground stations, with 

a minimum of six links per day, all data could be downlink conservatively within 4 days, leaving 

3 days for test data analysis. This plan would give mission controllers time to troubleshoot, derive 

contingency plans, and uplink test alterations before the next test sequence started.

It should be noted that no STK ground track analysis had been done until later in the 

study, by the GSFC team, and all link time assumptions were based on reference missions. Ground 

track analyses are sensitive to altitude and inclination, and link budgets are sensitive to required 

data rates, all of which were fluctuating throughout the study. To do a ground track analysis prematurely would not have been beneficial. The GSFC ground track/link analysis, with a 40-deg 

inclination, shows that 10 links per day are possible using Wallops and Santiago ground stations. 

With this many links, it should be possible to downlink all test data in 1 day using S-band. It was 

assumed that the cost of using multiple ground stations is acceptable. 

The video requirements of the demonstration testing were driven mainly by the need to see 

if bubbles were being generated within the system, either by the LADS or by the transfer process 

itself (pumps, valve turbulence, etc.). The CFM investigators judged that a 0.20-mm, bubble size 

detection capability would be suitable. The 0.20 mm is thought to be detectable using common 

video techniques and equipment. Smaller sizes would require exceptional resolution and drive up 

the data rates, which ultimately drives the spacecraft communication requirements. Without know-
68

ing which spacecraft bus was going to be used, which ground stations or how many, data backlog 

at the spacecraft level was a serious concern.

The video system for Pathfinder POD consists of four black-and-white cameras—one 

attached to each transfer line, one line to the transfer tank, and another line back to the storage 

tank. There are four cameras in the Pathfinder design, with two of them being active at a time. 

It is thought the cameras will view the fluid within the lines through a window viewing port that 

does not protrude into the line itself and would not cause any turbulence. Also, lighting ports will 

be required to illuminate the bubbles. A cold mirror approach with narrow-band, light-emitting 

diodes or multispectrum white lighting was suggested. The transfer lines were kept at 1⁄2-in diameter. Larger diameters were considered (1 to 2 in) to reduce flow rates and frame rates, and to represent more realistic line sizes of an actual CPS. Due to program budget constraints, it was decided 

to keep with the 1⁄2-in line since the 1⁄2-in valves were already qualified. It was thought to be cost 

prohibited to develop and qualify larger valves.

Another camera is mounted to each of the storage and transfer tanks for viewing the propellant state within the tanks. Viewing and lighting ports will also need to be incorporated on the 

tanks. These cameras can operate at a much lower frame rate than the transfer line cameras since 

there are no high-speed fluid movements within the tanks. However, given the much bigger viewing 

angle requirement, higher resolution cameras will probably be required. Since fluid transfers occur 

in only one direction at a time, only two cameras at a time are to be actively recording data. The 

active transfer line camera should be running continuously during fluid transfers. The tank cameras 

can be selected to monitor the acquisition tank or the receiving tank, or alternate between the two 

tanks at any desired duty cycle. With slow activity in the tanks, both cameras do not need to run 

simultaneously. This approach keeps the video data rate burden to a minimum.

Actual cameras have not been selected at this point, but a miniature imaging camera from 

Comtech AeroAstro was suggested. Modifications would have to be made to achieve the 120 fps 

required for the transfer line cameras. 

It was planned to keep the video system a stand-alone system that performs data compression and provides its own data storage buffer. The system will accept commands for configuration 

and operation from the ECDU or flight computer directly. Analysis by the CFM investigators showed that with a frame rate of 120 fps and a data rate of 15.7 Mbps (with a flow rate of 

1.48 m/s) for the transfer line video, and 1 fps at 8.4 Mbps for the tank video assumed sufficient, 

a total camera data rate will be 24.1 Mbps for two active cameras during transfers. Using a 10:1 

compression ratio judged by the MSFC video group to be easily achievable with black-and-white 

video, the after-compression total video data rate is ≈2.4 Mbps. This video rate is based on the 

0.20-mm bubble detection size and the 1⁄2-in transfer lines.

2.3.5.3 Guidance, Navigation, and Control Assumptions. Ground rules and assumptions for 

guidance, navigation, and control (GN&C) were developed to facilitate spacecraft bus selection to 

meet the needs of the experiment/demonstrations. Sun inertial pointing (aft end) was assumed the 

attitude to be maintained as much as possible for thermal control reasons (i.e., maximize the mission lifetime by minimizing boiloff). However, it is conceivable that the investigators might want to 
69

thermally load the demonstrator to challenge the cryogenic coolers, and they might want a different orientation for short periods of time. At one time early in the study, propulsion burns and even 

docking maneuvers were considered. Either way, the spacecraft bus should maintain some amount 

of maneuvering capability.

If reaction wheels (RWs) are used for the attitude control system (ACS) by the spacecraft 

bus, a Sun-pointing attitude should not be a problem in LEO since the momentum accumulation is 

cyclic per orbit, requiring little RCS desaturation of RWs. The microgravity levels of LEO will not 

be an issue for mass gauging experiments. An approximate value of 0.21 μg for orbital acceleration 

is well under the 30 μg settling level, as is the slew rate of 0.1 deg/s generating ≈0.51 μg. However, 

if an RCS is used for ACS, acceleration levels near 1.7 mg may result unless very low-level thrusters are used. Even so, the fidelity of low acceleration—achievable using RWs—will be difficult to 

achieve using only an RCS.

With RWs used for attitude control, it is assumed all major maneuvers will be performed by 

RCS, including propellant settling, tip-off correction, and deorbit operations. Reaction wheels can 

be used for a slow backup slewing capability in a contingency situation. The RCS will be used to 

desaturate RWs as necessary. If a commercial bus does not use RWs for attitude control, the RCS 

propellant mass required may be significant.

The pointing accuracy for the mission was not seen to be critical in any way. Therefore, 

1 deg/axis was chosen to keep the attitude control burden low. Even with RWs, tighter pointing 

accuracy would require more RCS propellant for desaturation of the wheels. The RW attitude 

rate control was chosen to be 0.01 deg/s. Low RW rates will minimize the RW mass required. The 

pointing knowledge of 0.005 deg (18 arcs) was chosen to accommodate the RW rate control, and 

should be easily achievable with a standard low-resolution star tracker. The slew rate of 0.1 deg/s 

minimum (90 deg in 15 min), was judged sufficient since no high slew rate operations were required 

after the elimination of the propulsion and docking requirements.

2.3.5.4 Avionics Data Summary. An uplink of 1 Mbyte per week was assumed sufficient 

for troubleshooting commands, script modifications, and parameter changes. Estimated data 

storage rates for the Pathfinder POD payload are defined in table 33 for a 7-day period:

Table 33. Estimated data storage rates over a 7-day experiment.

Mission Activity Data Storage Rate Period

Fluid transfer 2.4 Mbps Total active cameras during two 15-min transfers

Settled mass gauging 800 kbps Each tank for 60 min, with a 50% duty cycle

Quiescent mass gauging 1.6 Mbps 5-s measuring time, four times per hour, continuous

Wet/dry settling gauging 160 Kbps Both tanks for 60 min

CFM housekeeping (P, T, status) 35.6 Kbps Continuous
70

The data timeline in figure 39 shows the POD data acquisition required for a ‘settled’ demonstration test sequence lasting 306 min. It consists of two fluid transfers—one from the supply 

tank to the transfer tank and another transfer back to the supply tank. There are several operations inbetween transfers, including chilldowns and mass gauging of fluids within the tanks at 

different settling thrust levels. The ‘unsettled’ transfer demonstration test sequence is very similar, 

except without settling thrust being done. These demonstration sequences are repeated about every 

7.5 days, with two settled and two unsettled test sequences being done for a total campaign duration of 30 days. Continuous housekeeping monitoring and quiescent mass gauging is done before, 

after, and for about 14 days between the campaigns. Data rates shown in this graph are before overhead and margin are added, which are variables that will probably change throughout the program. 

However, the general profile of the curve should remain the same, showing the relative proportions 

of data required for each operation. Notice that the video required during the fluid transfers is the 

dominant data rate although the transfers only last for 15 min each.

0

50

100

150

200

250

0 50 100 150 200 250 300 350

Settled Gauging 30 μg 

Settled Gauging 1 mg

Unsettled Steady State

Duration (min)

Data (Mb/min)

ChilldownPrep

Chilldown

Settled Gauging 30 μg 

Settled Gauging 1 mg

Settled Gauging 30 μg 

Settled Gauging 1 mg

Chilldown

Settled Transfer 

From Supply Tank

Settled Transfer

ToSupply Tank 

ChilldownPreparation

Settled Gauging 30 μg Settled Gauging 30 μg 

QuiescentMonitoring

Figure 39. A typical experiment data acquisition settled demo timeline (306 min).

Table 34 shows the data volume acquired over time given the various duty durations. 

Because the last three line items of the housekeeping date are acquired continuously, it becomes 

significant over a 24-hr period, about 30% of the total data volume. The video data rates shown 

are after compression. RFMG is continuous during transfer operations, but reduced with a 50% 

duty cycle (i.e., two units at 800 kbps, 50% duty cycle). Wet/dry mass gauging is done at 1 Hz with 

5 kbps per sensor. During nontesting time, quiescent RFMG is done at 15-min intervals throughout the mission (for 5 s, four times per hour). Only 8 min of these data accumulate in 24 hr. Accumulation of housekeeping instrumentation data will occur while downlinking the transfer test data, 

but any system selected should be able to catch up in under the 4-day maximum downlink limit.
71

The variables in table 34 are the downlink rates, links per day, and link durations. The daily 

data remainder will equal the sum of the day’s volume, plus the previous day’s remainder, minus the 

downlink volume for the day. Several graphs were generated using this matrix with different variables. It can be seen that the resulting slope of the stored date line after downlink is highly sensitive 

to the links per day variable. 

Table 34. Data volumes acquired with duty durations: CCSDS overhead, 

15% and prephase A data margin, 30%.

Measurement

No. Data Type Notes

Collection

rate

(Mbps)

Duration

(hr)

Volume

(Gbits/hr)

Volume

w/CCSDS

(Gbits/hr)

Volume

w/Margin

(Gbits/hr)

1 Periodic

events

Video-line 1 1.570 0.50 2.83 3.25 4.22

2 Video-tank 1 0.840 0.50 1.51 1.74 2.26

3 RFMG-setting 2 0.800 1 2.88 3.31 4.31

4 Wet/dry gauging 3 0.160 1 0.58 0.66 0.86

7.79 8.96 11.65

1 Continuous

data 

collection

RFMG-quiscent 4 1.600 0.0056 0.032 0.037 0.048

2 Pressure/temps/etc.

S/C eng. data

5 0.036 1 0.130 0.149 0.194

3 5 0.005 1 0.018 0.021 0.027

5.01 0.18 0.21 0.27

Peak data rate Totals 7.97 9.17 11.92

Downlink Rate (Mbps)

10 10 10 10 10 10 10

Links per day

2 2 2 2 2 2 2

Link duration (min) each

8 8 8 8 8 8 8

Total downlink time per day (mm)

16 16 16 16 16 16 16

Downlink volume (Gbits)

 9.6 9.6 9.6. 9.6 9.6 9.6 9.6

Measurement

No.

Daily

Multiplier

(hr)

Volume

(Gbits/day)

Day

1 2 3 4 5 6 7

1 1 4.22 4.22

2 1 2.26 2.26

3 1 4.31 4.31

4 1 0.86 0.86

11.65 11.65 – – – – – –

1 24 1.16 1.16 1.16 1.16 1.16 1.16 1.16 1.16

2 24 4.65 4.65 4.65 4.65 4.65 4.65 4.65 4.65

3 24 0.65 0.65 0.65 0.65 0.65 0.65 0.65 0.65

6.45 6.45 6.45 6.45 6.45 6.45 6.45 6.45

18.11 18.11 6.45 6.45 6.45 6.45 6.45 6.45

Remainder 8.51 5.36 2.21 – – – –
72

For example, in figure 40, increasing the number of links to four per day reduces the link 

time to 2 days even with a lower bit rate of 7 Mbps, versus the two links per day and 10 Mbps in 

figure 41.

20

18

16

14

12

10

8

6

4

2

0

0 1 2 3 4

Duration (days)

Data (Gbits/day)

5 6 7 8

Daily Data Acquired

Downlink Volume

Data Stored

Figure 40. Data acquired and stored after downlink with X-band at 7 Mbps, 

four 8-min links per day.

20

18

16

14

12

10

8

6

4

2

0

0 1 2 3 4

Duration (days)

Data (Gbits/day)

5 6 7 8

Daily Data Acquired

Downlink Volume

Data Stored

Figure 41. Data acquired and stored after downlink with X-band at 10 Mbps, 

four 8-min links per day.
73

Increasing the links per day to 10 would reduce the link time required to only 1 day even 

with S-band, as shown in figure 42. 

20

18

16

14

12

10

8

6

4

2

0

0 1 2 3 4

Duration (days)

Data (Gbits/day)

5 6 7 8

Daily Data Acquired

Downlink Volume

Data Stored

Figure 42. Data acquired and stored after downlink with S-band at 3.5 Mbps, 

ten 8-min links per day.

The downlink data rates for X-band and S-band were based on international bandwidth 

limits for the near-Earth network system of 10 MHz for X-band and 5 MHz for S-band. Link budget analysis showed a reasonable link rate for X-band would then be 7 Mbps and 3.5 Mbps. Higher 

or lower rates would depend on factors like modulation techniques and coding strategy. The link 

budget analysis giving these results used quadrature phase shift keying (QPSK) modulation and 

RS/Viterbi coding. In figure 40, an idealized X-band rate of 10 Mbps was used for comparison 

sake. The data rates in these graphs include 15% overhead and 30% margin, as indicated in table 34.

2.3.5.5 Avionics Payload System Description. Figure 43 shows a schematic of the avionics 

system. Note that only the payload experiment avionics system is shown. No spacecraft systems 

such as GN&C, communications and data handling (C&DH), or communications are included. A 

serial data bus will connect the ECDU to the spacecraft flight computer for command and control 

of the experiments, and for uploading experiment software changes. It is expected that the ECDU 

will command the flight computer to perform settling maneuvers as it processes the experiments 

through the experiment timeline and receives go status from experiment sensors. An additional 

dedicated bus will be used to transfer experiment and video data to the flight computer for downlinking to ground. A video control unit is shown that will accumulate and store all video data from 

the four cameras, perform compression, and possibly transfer video to the flight computer directly 

via a third data bus, bypassing the ECDU.
74

Spacecraft

Bus Serial

Comm Link

Serial Link for

Experiment

Data Telemetry

Power From

Spacecraft

Bus

Serial

Command

Bus

Large

Tank

Payload

Computer

Experiment

Data Unit

Power

Unit

Pump

Control

Unit

Video

System Serial Link for

Video Transfer

Heater

Controller

Cryo

Cooler

Transfer

Pump

Heaters

Transfer

Pump

Small

Tank

Cameras

Heater

Controller

Vent

Valve Vent

Valve

Pressure

Tank

Figure 43. Pathfinder POD payload avionics and power system.

A payload power unit distributes power to the various valves, pumps, and controllers. This 

unit is described in more detail in section 2.3.6. Valve commands will be sent to the power unit 

by the ECDU. A dedicated pump controller is used to control the cryogenic coolers and transfer 

pumps. The ECDU will send operational commands to the pump control unit as needed to run the 

transfer test and maintain cryogenic cooling. With a large number of heaters and temperature sensors required, two heater controllers seemed justified for the system. It is thought that smart/active 

thermal control of heaters is preferred for this demonstration testing, giving flexibility to alter temperature parameters during the mission. In an actual CPS system, many of the heater settings can 

be fixed points.

2.3.5.6 Avionics Mass Summary. Table 35 shows the results of the avionics system mass. 

Despite the enormous amount of effort put into determining data rates and ground link strategies, 

the total avionics mass is a small portion of the overall vehicle mass. Since the communication system is to be part of the spacecraft bus, it is not included in this avionics mass rollup. For the most 

part, the masses shown were based on representative components and estimates for such a system. 

As the program develops, particularly CFM testing procedures and hardware, including video, the 

avionics system will also mature.
75

Table 35. Avionics mass summary.

Avionics Quantity

Unit Mass

(kg)

Basic Mass

(kg)

ECDU 1 8 8

Pump control unit 1 5 5

Heater controllers 2 5 10

Video system 1 15.19 15.2

Harnessing 1 9 9

Total – 42.19 47.2

2.3.5.7 Future Work. Future work for avionics systems includes the need to perform transfer video experiments to verify capture of bubbles, adequate frame rates, resolution, etc. A detailed 

design of a digital video system needs to be done, including memory storage, compression software 

development, and a spacecraft bus interface. Detailed design of the ECDU also needs to be done, 

including processor selection, data storage requirements and type, data buses, instrumentation, 

and control input/output. Software functional requirements to be developed include defining 

experiment timeline and control routines, defining contingency and fail-safe modes, and defining 

spacecraft interface protocol.

2.3.6 Power System

The spacecraft bus provides a complete power generation and management system to 

provide all of the power for both itself and the demonstration CFS payload. The power system 

described here accepts an umbilical cable carrying the payload power from the spacecraft bus 

and switches individual payload circuits on/off under the control of the payload computer system. 

The power from the spacecraft bus does not need to be conditioned but should maintain a potential of 22 to 34 V under specified loads.

2.3.6.1 Power Systems Ground Rules and Assumptions. Ground rules and assumptions for 

the power systems are as follows: 

• Off-the-shelf spacecraft bus provides all required power conversion and energy storage.

• The power required for the payload was analytically determined in this study and used 

to configure the spacecraft bus.

• This power is provided via one or two umbilical input circuits to the secondary power-switching 

unit. The switching is controlled by the payload computer system interfaced to the switching unit 

using a MIL-STD-1553B interface.4

• Cabling between the secondary power switching unit and the payload elements was sized 

to 1% power loss using physics-based methods.
76

2.3.6.2 Power System Analysis Methodology. Because the only requirement of the power 

system is to provide secondary switching and distribution, the analysis consists of sizing an appropriate secondary power distribution unit to switch each of the payload demonstration circuits. 

Based on the power levels required and the number of circuits, a representative off-the-shelf power 

distribution unit for which data were available and configured for this application was selected. 

It must be emphasized that the particular unit is selected for sizing purposes only; it is not recommended by this study as being superior to other off-the-shelf units, but is representative of the sort 

of off-the-shelf units available from a variety of makers.

The cabling connecting the secondary power distribution unit to each of the payload 

circuits is sized to incur a 1% power loss. The cables are sized using physics-based methods.

2.3.6.3 Power Requirements. Table 36 details the maximum power requirement for each 

demonstration operation. Note that a 30% design margin has been added to accommodate power 

requirement growth for each operation.

Table 36. Power requirements.

Avionics 

Power 

Valve 

Power

Heater 

Power 

Pump 

Power 

Cryo 

Power 

Total 

Power 

W/30

(%) Comment

Checkout 360 0 12.5 0 5.35 377.85 491 

Passive storage 

only

360 0 12.5 0 5.35 377.85 491 No cryocoolers 

running

0-g mass gauge 395 1 12.5 0 42 450.5 586 

Chilldown 395 3 12.5 0 42 452.5 588 No data transmit

Low-g mass gauge 395 1 12.5 0 42 450.5 586 

Pump transfer 395 1 12.5 7 42 457.5 595 

Pressure transfer 395 106 12.5 0 42 555.5 722 

Warmup 395 1 12.5 0 39 447.5 582 

Active idle 360 1 12.5 0 42 415.5 540 Cryo and heaters on, 

no data transmit

The peak power (exclusive of very short-duration transients) is 722 W. The average power over the 

entire mission is 539 W. Figure 44 shows the power level with respect to time over the 6-mo mission. The power level for each 30-day demo period is shown in figure 45.
77

1,000

900

800

700

600

500

400

300

200

100

0

0 1 2 3

Time (mo)

Passive

Active

Command

Transfer Demo (5-hr Avg.)

Power (W)

4 5 6

Figure 44. Six-month power profile.

1,000

900

800

700

600

500

400

300

200

100

0

0 5 10 15

Time (days)

5-hr Transfer

Demo (Avg.)

5-hr Transfer

Demo (Avg.)

5-hr Transfer

Demo (Avg.)

20 25 30

Power (W)

Figure 45. Propellant transfer power profile.
78

2.3.6.4 Power System Description. The secondary power distribution system consists of 

a minimal VME (Versa Module Eurocard) cage together with a set of VME power boards. Each 

board in the set carries its own external connectors and outer structure so that the card cage itself 

may be kept simple and light. The advantage of this approach is that the designer may configure 

a power unit containing those features (and hence boards) that are needed for a given application 

without carrying extra mass for connectors and structure that enable the features that the application does not need. The disadvantage is that the chassis does not provide anything (power supply, interfacing) other than a bus bar and an enclosure. The power supply, interfacing, and other 

required functions must be implemented on some of the VME boards themselves. The boards used 

for this application include the following types:

• Power distribution module (PDM)—each board switches power from its input up to 20,

4-A circuits.

• Low voltage power supply—provides 3.3, 5, and ±12 V power to the other boards.

• Advanced programmable interrupt controller (APIC)—provides interface (including 

MIL-STD-1553) to payload computer.

Figure 46 illustrates the configuration of the power system.

Flexible Power Unit Bus Controller

20 Circuits

3.3, 5, ±12 V

20 Circuits

To Payloads

MIL-STD 1553B

Commands

C&DH

From

Spacecraft

Bus (22–34 V)

LVPS

PDM

PDM

APIC 3.3, 5, ±12 V

3.3, 5, ±12 V

3.3, 5, ±12 V

Figure 46. Power system schematic.
79

2.3.6.5 Power System Results. Table 37 summarizes the mass breakdown for the power 

system.

Table 37. Power system mass summary.

Item Quantity 

Unit Mass 

(kg) 

Total Mass 

(kg) 

SC-11 VME chassis 1 8.4 8.4 

Power distribution module 2 1.6 3.2 

Low-voltage power supply 1 0.6 0.6 

APIC housekeeping card 1 0.96 0.96 

Power bus control assembly 1 4 4 

Cable/harness 1 1.1 1.1 

Total mass 18.3 

2.3.7 Mass Summary

2.3.7.1 Point-of-Departure Concept Mass Summary. This section provides an indepth mass 

overview for the CPST technology demonstration CPS-Pathfinder POD concept study. The masses 

in this section include the total wet mass/launch mass for the CPST CPS-Pathfinder POD vehicle, 

the CFM system/ vehicle payload, and the CFM subsystems, as well as the spacecraft bus and margin for each CPS-Pathfinder element.

For clarity, a brief description of mass terminology used in this section is included here. 

Basic mass is defined as mass with no margin, contingency, or mass growth allowance (MGA) 

added to it. Margin, contingency, and MGA are basically interchangeable terminologies used in the 

aerospace industry. Predicted mass is the sum of basic mass added to contingency/MGA. Dry mass 

is defined as mass that does not include propellant or gasses/fluids loaded. Wet mass is the spacecraft’s mass before liftoff and is the same as launch mass. Wet mass, also known as launch mass, 

includes dry mass plus the vehicle propellant and any fluids. 

The CPST technology demonstration CPS-Pathfinder POD concept wet mass/launch mass 

is 2,298 kg. This mass includes the CFM system/payload and the spacecraft bus, as well as a 30% 

contingency/MGA applied to the basic dry mass of the CFM payload and the spacecraft bus. An 

image of the CPST CPS-Pathfinder concept is shown in figure 47 with the CFM radiators and 

spacecraft bus solar arrays deployed.
80

Figure 47. CPST POD configuration.

Table 38 provides summary level mass details for each CFM subsystem, as well as spacecraft 

bus mass data for the overall spacecraft bus dry mass and propellant. Table 38 also totals the CPST 

mass for the entire vehicle in the lower right under the ‘Predicted Mass (kg)’ column.
81

Table 38. CPST POD mass summary—CPS Pathfinder.

Basic 

Mass

(kg)

Contingency

(%)

Contingency

(kg)

Predicted

Mass

(kg)

CFM Experiment/Payload

 1.0 CFM demo—hydrogen 

system

605.10 30 181.53 786.63

 2.0 Avionics 47.19 30 14.16 61.35

 3.0 Power 18.16 30 5.45 23.61

 4.0 Thermal 86.60 30 25.98 112.58

 5.0 Structures 150.53 30 45.16 195.69

CFM experiment/payload dry mass 907.58 30 272.28 1,179.86

 6.0 Nonpropellant fluids 27.50 27.50

 7.0 CFM demo fluid (LH2) 260 260

Total CFM experiment/payload 

system mass

1,195.08 1,467.36

Spacecraft Bus

 8.0 Mechanical 173 30 51.90 224.90

 9.0 Attitude control 34 30 10.20 44.20

10.0 Thermal 22 30 6.60 28.60

11.0 Propulsion 65 30 19.50 84.50

12.0 Power 170 30 51 221

13.0 Avionics 10 30 3 13

14.0 Communication 12 30 3.60 15.60

Spacecraft bus dry mass 486 30 145.80 631.80

15.0 Bus propellant 

(hydrazine)

199 199

Total spacecraft bus wet mass 685 830.80

Total CPST spacecraft launch 

mass (CFM experiment and bus)

1,880.08 2,298.16

2.3.7.2 CFM System Mass Summary. The total predicted mass of the CFM system is 

1,467 kg. This mass includes the CFM system basic mass of 1,195 kg and the 30% contingency/

MGA applied to the dry mass, which is 272 kg. The CFM system contains five dry mass subsystems: CFM demonstration, avionics, power, thermal, and structures. 

The CFM demonstration subsystem basic mass is 605. When combined with a 30% contingency/MGA of 182 kg, this yields 787 kg predicted mass for the CFM demonstration subsystem: 

• The predicted mass for the avionics subsystem is 61 kg. This number includes a 30% contingency/

MGA of 14 kg added to the basic mass of 47 kg.
82

• The power subsystem basic mass is 18 kg; the power contingency/MGA of 30% is slightly less 

than 5.5 kg. The total predicted mass of the power subsystem rounds up to 24 kg. 

• The predicted mass for the thermal subsystem is 87 kg. This number includes a 30% contingency/

MGA of 14 kg added to the basic mass of 47 kg. 

• The structure subsystem basic mass is 151 kg. When combined with a 30% contingency/MGA 

of 45 kg, this yields 196 kg predicted mass for the structures subsystem. 

• Nondry CFM mass includes nonpropellant fluids and CFM demonstration fluid. The nonpropellant fluid is a cryogenic-fluid transfer pressurant, which is helium. The mass of the helium 

is 27.5 kg. The CFM demonstration fluid mass is 260 kg; the demonstration fluid is LH2. No 

contingency/MGA was added to the cryogenic fluid transfer pressurant or to the CFM demonstration fluid.

Table 39 provides detailed information for each CFM payload subsystem.
83

Table 39. CFM subsystem mass.

CFM Demo-Hydrogen System

Basic 

Mass

(kg)

Contingency

(%))

Contingency

(kg)

Predicted

Mass

(kg)

1.0 CFM demo-hydrogen system 605.10 30 181.53 786.63

1.1 Storage tank active CFM subsystem 27.90 30 8.37 36.27

1.2 Storage tank passive CFM subsystem 72.10 30 21.63 93.73

1.3 Storage tank 108.78 30 32.63 141.41

1.4 Storage tank subsystem 32.18 30 9.65 41.83

1.5 Helium storage subsystem 142.16 30 42.65 184.81

1.6 Transfer tank passive CFM subsystem 30.70 30 9.21 39.91

1.7 Transfer tank 14.30 30 4.29 18.59

1.8 Transfer tank subsystem 21.16 30 6.35 27.51

1.9 Transfer system 24.86 30 7.46 32.32

1.10 Vent subsystem 22.86 30 6.86 29.72

1.11 Instrumentation 108.10 30 32.43 140.53

2.0 Avionics 47.19 30 14.16 61.35

2.1 Experiment computer/data unit 8.00 30 2.40 10.40

2.2 Pump control unit 5.00 30 1.50 6.50

2.3 Heater controllers 10.00 30 3.00 13.00

2.4 Video system 15.19 30 4.56 19.75

2.5 Harnessing 9.00 30 2.70 11.70

3.0 Power 18.16 30 5.45 23.61

3.1 Cable harness 1.00 30 0.30 1.30

3.2 Secondary power distribution 17.16 30 5.15 22.31

4.0 Thermal 86.60 30 25.98 112.58

4.1 Insulation and thermal control material 7.50 30 2.25 9.75

4.2 Radiators 79.10 30 23.73 102.83

5.0 Structures 150.53 30 45.16 195.69

5.1 Primary structure 112.47 30 33.74 146.21

5.2 Secondary structure 22.49 30 6.75 29.24

5.3 Tank attachments 10.38 30 3.11 13.49

5.4 Joints and fittings 5.19 30 1.56 6.75

CFM experiment/payload dry mass 907.58 30 272.28 1,179.86

6.0 Nonpropellant fluids 27.50 27.50

6.1 Cryo-fluid transfer pressurant (helium) 27.50 27.50

7.0 CFM demo fluid (LH2) 260 260

Total CFM experiment/payload system mass 1,195.08 1,467.36
84

2.3.7.3 Spacecraft Bus Mass Summary. The total spacecraft bus mass for the CPST CPSPathfinder concept is 831 kg. The spacecraft bus basic mass for the dry mass elements comes to 

486 kg. A 30% contingency/MGA of 146 kg is added to arrive at a predicted mass of 632 kg. No 

contingency/MGA is added to the 199 kg of spacecraft bus propellant hydrazine. The propellant 

mass total of 199 kg is added to the spacecraft dry mass (predicted mass) of 632 to arrive at the 

spacecraft bus mass total of 831 kg. Table 40 provides summary level mass detail for each spacecraft bus subsystem, as well as spacecraft bus dry mass and propellant (hydrazine) data and the 

total spacecraft bus wet/launch mass in the lower right corner under the ‘Predicted Mass (kg)’ 

column.

Table 40. Spacecraft bus mass summary.

Basic 

Mass

(kg)

Contingency

(%)

Contingency

(kg)

Predicted

Mass

(kg)

8.0 Mechanical 173 30 51.90 224.9

9.0 Attitude control 34 30 10.20 44.2

10.0 Thermal 22 30 6.60 28.6

11.0 Propulsion 65 30 19.50 84.5

12.0 Power 170 30 51 221

13.0 Avionics 10 30 3 13

14.0 Communications 12 30 3.60 15.6

Spacecraft bus dry mass 486 30 145.80 631.8

15.0 Bus propellant 

(hydrazine)

199 199

Total spacecraft bus wet mass 685 830.8
85

2.4 Study Accomplishments and Forward Work

The CPST Government POD prephase A study successfully accomplished all of the study 

objectives. The study accomplishments are summarized as follows:

• The Government POD concept provides a viable mission concept that provides a balanced 

approach to meeting mission objectives within project constraints.

• The Government POD concept provides a framework for evaluating alternative CPST mission 

concepts.

• The Government POD concept identifies a mission requirement to provide a framework 

for technology development and ground test planning.

• The Government POD concept provides a foundation defining a synthesized reference concept.

The results of the Government POD study provide a basis for preparing for the Mission Concept Review (MCR) project key decision point as well as moving forward with technology development activities and phase A design studies. To conclude this study, a number of areas 

for forward work were identified and are listed below:

• Support phase A level flight system design analysis trade studies.

• Refine planning for integrated technology and flight system component development.

• Refine test and integration plan.

• Refine mission costs estimate:

– Launch vehicle (resolve differences between National Launch Services and WFF costs estimates.

– Cryogenic fluid system (incorporate streamlines development approaches).

– Spacecraft bus (investigate bus procurement options).

– Mission operations.

• Refine project schedule inputs.:

– Launch vehicle, cryogenic fluid system, spacecraft bus.

– Long lead items (define critical path procurements).

• Support continued definition of system concept of operations.

• Build on Government POD study results to develop MCR (key decision point) products 

as required.

• Continue definition of partnership options based on project acquisition strategy.

• Review industry BAA concepts and begin mission concept synthesis.
86

This study culminates over 19 mo of prephase A study activities to define a mission concept 

for the CPST technology demonstration mission. Approximately 20 different mission concepts have 

been defined and evaluated. Over 90 study participants from five NASA Centers have supported 

the prephase A studies in which the study participants have demonstrated an exemplary level of 

enthusiasm, dedication, and collaboration in spite of ever-changing project constraints and mission 

requirements. The efforts of the study participants are greatly appreciated and provide a very solid 

foundation for a successful flight demonstration project.
87

REFERENCES

1. NASA-STD-8719.14, Process for Limiting Orbital Debris, Version 2.0.1, August 2007.

2. NPR 8705.4, Risk Classification for NASA Payloads, June 2004.

3. NASA-STD-5001a, Structural Design and Test Factors of Safety for Spaceflight Hardware, 

August 2008.

4. MIL-STD-1553B, Aircraft Internal Time Division Command/Response Multiplex Data Bus, 

September 1978.
88

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Cryogenic Propellant Storage and Transfer Technology 

Demonstration: Prephase A Government 

Point-of-Departure Concept Study

J.A. Mulqueen, B.M. Addona, D.A. Gwaltney, K.A. Holt, R.C. Hopkins, 

J.A. Matus, P.S. McRight, C.G. Popp, S.G. Sutherlin, H.D. Thomas, 

M.F. Baysinger,* C.D. Maples,* P.D. Capizzo,** L.L. Fabisinski,*** 

L.S. Hornsby,† T.K. Percy,‡ and S.D. Thomas¶

George C. Marshall Space Flight Center

Huntsville, AL 35812

National Aeronautics and Space Administration

Washington, DC 20546–0001

Unclassified-Unlimited

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Availability: NASA CASI (443–757–5802)

Prepared by the Advanced Concepts Office, Engineering Directorate

*Qualis Corporation, **Raytheon, ***International Space Systems, †Jacobs Engineering, 

‡Science Applications International Corporation, ¶

Dynetics

M–1343

Technical Memorandum

NASA/TM—2012–217467

cryogenic fluid management technologies, cryogenic propellant storage and transfer, technology flight 

demonstration

01–09–2012

UU 106

NASA

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The primary purpose of this study was to define a point-of-departure prephase A mission concept for 

the cryogenic propellant storage and transfer technology demonstration mission to be conducted by 

the NASA Office of the Chief Technologist (OCT). The mission concept includes identification of the 

cryogenic propellant management technologies to be demonstrated, definition of a representative mission 

timeline, and definition of a viable flight system design concept. The resulting mission concept will serve as 

a point of departure for evaluating alternative mission concepts and synthesizing the results of industrydefined mission concepts developed under the OCT contracted studies.

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