Michael P. Doherty, Michael L. Meyer, Susan M. Motil, and Carol A. Ginty

Glenn Research Center, Cleveland, Ohio

Cryogenic Propellant Storage and Transfer (CPST) 

Technology Maturation: Establishing a Foundation 

for a Technology Demonstration Mission (TDM)

NASA/TM—2014-218092

February 2014

AIAA–2013–5458
NASA STI Program . . . in Profi le

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientifi c and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Offi cer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major signifi cant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of signifi cant 

scientifi c and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientifi c 

and technical fi ndings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientifi c and 

technical fi ndings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientifi c and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientifi c, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientifi c and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question to help@sti.nasa.gov

• Fax your question to the NASA STI 

Information Desk at 443–757–5803

• Phone the NASA STI Information Desk at

 443–757–5802

• Write to:

 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

 Hanover, MD 21076–1320
Michael P. Doherty, Michael L. Meyer, Susan M. Motil, and Carol A. Ginty

Glenn Research Center, Cleveland, Ohio

Cryogenic Propellant Storage and Transfer (CPST) 

Technology Maturation: Establishing a Foundation 

for a Technology Demonstration Mission (TDM)

NASA/TM—2014-218092

February 2014

AIAA–2013–5458

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Space 2013 Conference and Exposition

sponsored by the American Institute of Aeronautics and Astronautics

San Diego, California, September 10–12, 2013
Acknowledgments

The authors are deeply appreciative of the efforts of the NASA-wide CPST Technology Maturation Team and supporting 

contractors. Without their passion and dedication, the accomplishments, noted herein, would not have been possible. Their 

contributions have been recognized by the Agency in the form of the 2013 NASA Group Achievement Honor Award.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2014-218092 1 

Cryogenic Propellant Storage and Transfer (CPST) Technology 

Maturation: Establishing a Foundation for a Technology 

Demonstration Mission (TDM)

Michael P. Doherty, Michael L. Meyer, Susan M. Motil, and Carol A. Ginty 

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

Abstract

As part of U.S. National Space Policy, NASA is seeking an innovative path for human space 

exploration, which strengthens the capability to extend human and robotic presence throughout the solar 

system. NASA is laying the groundwork to enable humans to safely reach multiple potential destinations, 

including asteroids, Lagrange points, the Moon and Mars. In support of this, NASA is embarking on the 

Technology Demonstration Mission Cryogenic Propellant Storage and Transfer (TDM CPST) Project to 

test and validate key cryogenic capabilities and technologies required for future exploration elements, 

opening up the architecture for large cryogenic propulsion stages (CPS) and propellant depots. The TDM 

CPST project will provide an on-orbit demonstration of the capability to store, transfer, and measure 

cryogenic propellants for a duration which is relevant to enable long term human space exploration 

missions beyond low Earth orbit (LEO). Recognizing that key cryogenic fluid management technologies

anticipated for on-orbit (flight) demonstration needed to be matured to a readiness level appropriate for 

infusion into the design of the flight demonstration, the NASA Headquarters Space Technology Mission 

Directorate authorized funding for a one-year (FY12) ground-based technology maturation program. The 

strategy, proposed by the CPST Project Manager, focused on maturation through modeling, studies, and 

ground tests of the storage and fluid transfer Cryogenic Fluid Management (CFM) technology subelements and components that were not already at a Technology Readiness Level (TRL) of 5. A 

technology maturation plan (TMP) was subsequently approved which described: the CFM technologies 

selected for maturation, the ground testing approach to be used, quantified success criteria of the 

technologies, hardware and data deliverables, and a deliverable to provide an assessment of the 

technology readiness after completion of the test, study or modeling activity. This paper will present the 

testing, studies, and modeling that occurred in FY12 to mature cryogenic fluid management technologies 

for propellant storage, transfer, and supply, to examine extensibility to full scale, long duration missions, 

and to develop and validate analytical models. Finally, the paper will briefly describe an upcoming test to 

demonstrate Liquid Oxygen (LO2) Zero Boil-Off (ZBO). 

Nomenclature

AREP Atlas Reliability Enhancement Program

BAC Broad Area Cooling 

CAT Cryogenic Analysis Tool

CFD Computational Fluid Dynamics 

CFM Cryogenic Fluid Management 

CPPPO Computational Propellant and Pressurization Program

CPS Cryogenic Propulsion Stage 

CPST Cryogenic Propellant Storage and Transfer 

CryoSim Cryogen Storage Integrated Model
NASA/TM—2014-218092 2 

CTL Cryogenic Test Laboratory 

DAM Double Aluminized Mylar 

EDS Earth Departure Stage

ESTF Exploration Systems Test Facility

GFFSP Generalized Fluid System Simulation Program

GRC Glenn Research Center, NASA 

GUI Graphical user interface

KPP Key Performance Parameters

KSC Kennedy Space Center, NASA

LAD Liquid Acquisition Device 

LEO low Earth orbit 

LH2 Liquid Hydrogen 

LO2 Liquid Oxygen 

MLI Multi-Layer Insulation 

MSFC Marshall Space Flight Center, NASA

NASA National Aeronautics and Space Administration

NVF No-vent-fill

PCB Project Control Board

PMD Propellant Management Device

RBO Reduced boil-off 

SLS Space Launch System 

SMiRF Small Multi-Purpose Research Facility

SOFI Spray-On Foam Insulation 

TankSim Tank System Integrated Model Tank System Integrated Model

TCS Thermal Control System 

TDM Technology Demonstration Mission 

TMP Technology Maturation Plan 

TRL Technology Readiness Level

TVS Thermodynamic Vent System 

VATA Vibro-Acoustic Test Article

VOF volume of fluid 

ZBO Zero boil-off 

1.0 Introduction

In support of the National Aeronautics and Space Administration’s (NASA) Strategic Goal 3.1, 

“Sponsor early stage innovation in space technologies in order to improve the future capabilities of 

NASA, other government agencies, and the aerospace industry,” studies have pointed to the value of 

extended use of cryogenic fluids in space. It was announced during September 2011 that NASA had been 

authorized to proceed with the development of a heavy lift launch vehicle, the Space Launch System 

(SLS). The Cryogenic Propulsion Stage (CPS) of the SLS is conceived to use liquid oxygen (LO2) and 

liquid hydrogen (LH2) as the propellant combination. In addition, Agency mission architecture studies 

include consideration of options for propellant resupply, either via tankers or in-space propellant depots. 
NASA/TM—2014-218092 3 

These mission capability elements have dictated the need for an advanced development program within 

NASA to mature several Cryogenic Fluid Management (CFM) technologies for in-space mission 

operations including the long duration storage of cryogenic fluids using both active and passive thermal 

control and micro-g tank pressure control, tank-to-tank transfer of cryogens, and unsettled propellant 

mass gauging. An in-space flight demonstration of these technologies is critical to the development of all 

cryogenic in-space propulsion system, for example, the CPS (Ref. 1). 

The current CPST baseline mission architecture is to develop, launch and operate a free flying 

satellite in low Earth orbit (LEO) to demonstrate and mature CFM technologies. The design concept 

involves the CPST payload integrated with a spacecraft bus, which will provide attitude control, 

communication, and propulsion functions for the integrated unit, launched aboard a medium class launch 

vehicle that delivers the CPST payload to a circular orbit of sufficient altitude to reduce atmospheric drag 

to acceptable levels. The spacecraft would fly in a solar-inertial attitude with the aft end of the spacecraft 

pointed toward the sun to reduce solar heating of cryogenic tanks and cold structures. 

The CFM technologies included in the planned flight demonstration mission are passive cryogenic 

propellant storage, tank thermal and pressure control, liquid acquisition, transfer, and several methods of 

mass gauging. The mission duration is currently estimated to be two months, which is based upon the 

time needed to complete CFM and spacecraft checkout, passive storage demonstration, and two transfer 

cycles at unsettled conditions. The mission would conclude with a controlled re-entry. After the mission 

is complete, data will be analyzed, and a final mission report will be completed for project closeout. 

In addition to delivering flight data, the project was tasked: to validate performance models suitable 

for analyzing full-scale space vehicle tank systems capable of storing LH2 for an extended duration in 

microgravity with reduced boil-off (RBO) (including active thermal control technology) and to advance a 

suite of technologies that would enable spaceflight systems capable of storing large quantities of LO2 for 

an extended duration in microgravity with zero boil-off (ZBO). 

On June 24, 2011, the CPST Project Control Board (PCB) approved a ground test CFM technology 

maturation strategy for FY12. This strategy focused on maturation through modeling, studies and ground 

tests of the storage and fluid transfer CFM technology sub-elements and components that were not at a 

technology readiness level (TRL) of 5. The CPST Project Manager then directed that a Technology 

Maturation Plan (TMP) be created describing the CFM technologies selected for maturation, the approach 

to be used, quantified success criteria of the technologies Key Performance Parameters (KPP) and planned 

deliverables. The CPST TMP was formally approved by the CPST Project Manager in February 2012. 

In the TMP, the CFM technologies to be matured were identified and described in Section 2.0, Core 

Technologies under the heading of Technology Assessment. Section 3.0 of the TMP and Table 1 

Table 1.—FY12 Ground Tests, Studies, and Modeling of Storage and Fluid Transfer 

CFM Technology Sub-elements and Components

Test/Task Name Objective

LH2 Active Thermal Control Thermal 

Performance (LH2 RBO) 

Demonstration of a flight representative active thermal control system for RBO 

storage of LH2 for extended duration in a simulated space thermal vacuum 

environment

LH2 Active Thermal Control Structural 

Performance (MLI/BAC Vibro-Acoustic 

Test Article (VATA))

Assess the structural performance of an MLI/BAC shield assembly subjected to 

launch vibration loads 

Active Thermal Control Scaling Study Conduct study to show relevancy of CPST-TDM active thermal control flight data 

to full scale CPS or Depot application 

Passive Thermal Control—Penetration 

Heat Leak

Measurement of heat leak due to struts penetration integrated with MLI. 

Passive Thermal Control—Composite Strut 

Thermal Performance in LH2

Measurement of heat leak due to composite struts integrated with MLI with 20 K 

boundary temperature.

Passive Thermal Control—Thick MLI 

Extensibility Study 

Assess optimum approach for attachment of thick (40 to 80 layers) MLI to very 

large tanks 

Liquid Acquisition Device (LAD) Outflow 

and Line Chill 

Quantify the LAD stability (no LAD breakdown) due to transfer line chill down 

transient dynamic pressure perturbations during outflow 

Analytical Tool Development Continue development of tools specific for CPST 
NASA/TM—2014-218092 4 

presented the full set of FY12 technology maturation activities, including tests, studies and analytical 

tool development to ensure that the list of selected flight demonstration technologies was at TRL-5, or as 

mature as possible without flight testing, by the end of FY12.

LO2 ZBO was not considered a technology maturation element but rather a ground demonstration of 

the capability of storing large quantities of LO2 for an extended duration in microgravity with ZBO and 

complimentary to the flight mission. 

2.0 Technology Maturation: Tests, Studies, Model Development

The following sections describe the background and approach, results, and significance of the tests, 

studies, and modeling of the storage and fluid transfer CFM technology sub-elements and components 

that were selected for maturation during FY12. 

2.1 LH2 Active Thermal Control

Long duration, space-based storage of LH2 appears feasible only through the use of active cooling. 

Active cooling can be accomplished using a cryocooler and a closed loop gas as the cryocooler working 

fluid for distributed cooling (Ref. 2). Despite the improving prospects for high capacity 20 K cryocoolers, 

this test focused on available cryocooler technology, applying the much more available 90 K cryocooler 

technology to cool a shield surrounding the LH2 tank for Reduced Boil-Off (RBO) storage of LH2. 

Two test bed tank systems were developed to meet the goal of evaluating thermal and structural 

characteristics of integrated multi-layer insulation (MLI) and broad area cooling (BAC) shield system: the 

RBO was built to thermally evaluate the system at GRC, and the Vibro-Acoustic Test Article (VATA) 

was built to structurally evaluate the system at MSFC. RBO and VATA employed very similar tank, 

thermal control system, and structural penetration configurations. This approach was intended to produce 

thermal and structural data on the same configuration providing a complete characterization of the system 

for the CPST project to consider in the context of a flight test.

The RBO and VATA Thermal Control System (TCS) included Spray-On Foam Insulation (SOFI) 

directly bonded to the tank to satisfy the CPST ground hold requirement. The integrated MLI and BAC 

shield system were positioned over the SOFI and provided both passive and active cooling components 

required for long-duration in-space storage of LH2. The MLI blanket included two primary components: a 

low-density (8-layer/cm) 30-layer blanket between the SOFI and the BAC shield and a standard density 

(20-layer/cm) 30-layer blanket outside the BAC shield. While the majority of the MLI surface area 

maintained specified density, the blankets were compressed at the seams where Velcro was sewn to allow 

for the removal and reinstallation of the blankets. This seam treatment decreased the thermal performance 

of the blanket. Low conductance polymer standoffs spaced the BAC shield at a proper distance off the 

surface of the tank and also constrained movement of the shield to prevent damage during estimated 

dynamic launch loads. The VATA test article employed the appropriate number of standoffs required for 

a flight-like configuration while RBO used a reduced number of standoffs to improve the thermal 

performance of the system.

2.1.1 Thermal Performance (LH2 RBO Test) 

In the RBO test, the shield is installed within the layers of radiation reflector insulation, effectively 

cooling them and reducing the exposure temperature of the insulation under the shield, offering a hot side 

temperature for the LH2 tank of 90 K, instead of the LEO imposed temperature of roughly 200 K. 

Coupled to this shield are straps and collars to cool the plumbing and tank supports. The selected

cryocooler was a 20 W 90 K reverse turbo-Brayton cycle cryocooler, with heat rejection accomplished via 

a heat pipe radiator mounted integral to the test hardware within the vacuum chamber (Ref. 3). A 

predictive system thermal model that was developed prior to testing was utilized for detailed design of the 

experiment as well as to provide insight into the test results.
NASA/TM—2014-218092 5 

Figure 1.—Liquid Hydrogen (LH2) RBO Experiment 

Test Article Being Lowered into SMiRF Vacuum 

Chamber. The white ring above the test tank is 

the heat pipe radiator, behind which is mounted 

the reverse turbo-Brayton cycle cryocooler.

The purpose of the testing was twofold. First it would demonstrate the integration of a BAC shield 

embedded in tank-applied thick MLI to a flight representative cryocooler. Second, it would quantify the 

system thermal performance of a flight representative active thermal control system for RBO storage of 

LH2 in a simulated space environment. The testing was conducted at Glenn Research Center’s (GRC)

Small Multi-Purpose Research Facility (SMiRF), which provided the thermal and vacuum environment to 

simulate space-based conditions. The test article at SMiRF is shown in Figure 1. 

Two essential tests were conducted, a “Cooler Off” passive thermal control test, and a “Cooler On” 

active thermal control test. The Cooler Off test determined the baseline boil-off rate of the system, with 

the active cooling components in place. The Cooler On test determined the boil-off reduction with the 

cryocooler system running. The total heat leak to the LH2 tank was measured by boil-off (a flow meter in 

the vent line) while the tank was kept at constant backpressure. The heat measured is composed of 

radiation (principally through the MLI) and conduction sources. Conduction heat transfer is readily 

determined from the temperature data on the respective conduction paths (e.g., support struts, fill and vent 

lines) into the tank, along with the physical characteristics of those paths. The MLI heat transfer was

through the 60 radiation shields, 30 below and 30 over the BAC, and the 15 layers on the penetrations. 

MLI heat could not be directly measured, but was calculated from the tank overall heat less the 

conduction heat. 

By comparing the Cooler On test to the Cooler Off test, the reduction in heat transfer rates were most 

significant in the struts, where the heat was reduced by 62 percent, while the fill line heat was cut by 

50 percent. The net heat (calculated) to the inner MLI was reduced by 61 percent. Overall, the boil off 

reduction was 48 percent, less than predicted. Yet, the ratio of active storage system mass per watt of heat 

removal (kg/W) and ratio of watts of active storage system input power per watt of heat removal (W/W) 

both exceeded success targets.
NASA/TM—2014-218092 6 

The significance of this test effort is that an integrated LH2 distributed reduced boil off system (test 

tank, MLI/BAC shield, cryocooler, and radiator) has been demonstrated on the ground, offering potential 

for the storage of LH2 for extended durations in space. 

2.1.2 Structural Performance (MLI/BAC Shield Acoustic Test)

The combination of MLI and a BAC Shield is a promising thermal solution to heat leak through the 

tank wall, but the structural properties of the concept must be addressed, as this thermal control system 

must also withstand launch environmental loads (Ref. 4). As indicated above, the VATA was constructed 

to be similar to the RBO test article.

The purpose of this testing was to verify the structural integrity of an MLI/BAC shield assembly 

when subjected to representative launch vehicle vibration and acoustic loads, and to obtain design 

experience for the BAC shield integration to plumbing, tank supports and foam substrates. The testing 

was conducted at Marshall Space Flight Center’s (MSFC) Exploration Systems Test Facility (ESTF) and 

Acoustic Test Facility, which provided the thermal, vacuum, and acoustic load environment to simulate 

launch conditions and perform verifications to determine structural integrity.

The VATA assembly sequence is illustrated in Figure 2. 

A worst-case launch load was chosen to best evaluate the structural integrity of the MLI/BAC shield 

system design. Both random vibration and acoustic tests were considered as possible options to meet this 

requirement. Random vibration tests are best suited for small and heavy components whose local 

environment is governed by the surface to which they are mounted. Acoustic tests are best for low mass, 

large surface area structures whose response is driven by sound pressure. An acoustic test was selected for 

the integrated MLI and BAC shield system. Acoustic test data analysis showed an acoustic load 

consistent with the worst-case envelope of all the launch vehicles currently under consideration by the 

CPST project for the flight payload. The test was successful as: 

• Accelerometer data from the acoustic test yielded no unexpected results. 

• Thermal tests were conducted before and after the acoustic test and no change was found in the 

thermal performance of the system. 

• Leak checks were performed on the BAC tubes after each test in the series and no leaks were

found. 

• A visual inspection of the outside of VATA was conducted after the conclusion of the acoustic test 

with no observation of damage. A small amount of denting was observed on the BAC shield, but 

this did not result in diminished thermal or structural performance of the system.

It is to be noted that the original relationship between the RBO and VATA tests was altered when the 

RBO configuration reduced the number of BAC shield standoffs in order to improve thermal 

performance. Based on this deviation in configuration, the Technology Maturation effort did not yield 

both a structurally and thermally viable configuration for an integrated MLI and BAC shield TCS.

(a) (b) (c) (d) (e)

Figure 2.—VATA build sequence: (a) Primed tank with standoffs, (b) Tank-applied SOFI, trimmed to shape, 

(c) Inner MLI blanket, (d) BAC shield, and (e) Fully assembled VATA, with outer MLI.
NASA/TM—2014-218092 7 

The significance of this test effort is that a structurally viable integrated LH2 distributed reduced boiloff system (test tank and MLI/BAC shield) survived acoustic testing, supporting the potential for the 

storage of LH2 for extended durations in space.

2.1.3 RBO II and VATA II

Recently, the RBO and VATA tests were each repeated, but with the inner MLI and BAC Shield 

standoffs removed and replaced with a self-supporting MLI that is also capable of supporting the BAC 

shield. Both tests were completed successfully and the results appear promising; however, data reduction 

and analysis were not complete in time to be included in this paper. 

2.1.4 Active Thermal Control Scaling Study

Long term, in-space storage of a full scale CPS or cryogenic propellant depot will require both a 

robust insulation system and an active thermal control system to minimize the propellant loss due to 

radiant heat. Active thermal control is being demonstrated for LH2 (RBO with a 90 K cryocooler, using a 

BAC shield) and LO2 (ZBO with a 90 K cryocooler, using distributed cooling applied directly to the test 

tank) utilizing a 1.2 m diameter test tank in ground based testing. Nevertheless, application of such 

technologies to large-scale tanks requires some study.

The purpose of the study was to validate the relevancy of a scaled LH2 RBO active thermal control 

system ground and flight test approach to a full scale CPS or depot application. While investigating RBO, 

scalability for subsystem (BAC tube on shield concept, MLI/BAC shield integration, and support system 

cooling concepts) and component (Turbo-Brayton cryocooler) technologies were considered as well as 

the development of an active thermal control scheme for a “full scale” application in LEO.

A combination of contracted studies for large cryocoolers, large cryogenic tank structures, and large 

tank MLI concepts, and in-house studies on BAC and MLI integration techniques (including sizing 

studies and thermal trades on heat interception straps, BAC shield locations, and radiators) has led to 

augmentation of a comprehensive spreadsheet sizing tool, from which parametric analyses were 

performed in order to evaluate the applicability of active cooling as compared to passive-only thermal 

control for tanks ranging from 2 to 10 m in diameter (Refs. 5 and 6). For each mission architecture, the 

loiter period at which passive, RBO, and ZBO designs result in the lowest cryogenic system mass is 

determined through these parametric analyses. Mass, power, and size relationships were traded 

parametrically to establish the appropriate loiter period where active thermal control reduces mass. The 

projected benefit is compared for passive, boil-off reduction with a 90 K shield, ZBO (20 K cooling 

system only), and ZBO with 20 K cooling and a 90 K shield. The analysis shows: (1) a benefit for active 

thermal control when loiter durations are as little as a few weeks when compared to passive storage, and 

(2) that two stage cooling reduces power and mass when compared to single stage cooling. Furthermore, 

active cooling reduces the significance of varied MLI performance, which historically has large 

performance variability.

The significance of this study effort is that the active thermal control system(s) matured under the 

FY12 TMP can be scaled to full size future space mission architectures, and that components for full scale 

applications such as cryocoolers, gas circulators, recuperators, BAC tubing and cooling attachment straps 

are considered a design issue rather than a technology issue and therefore do not present a scaling risk.

2.2 Passive Thermal Control 

The passive thermal controls utilized for advanced cryogenic propellant storage incorporate 

insulations to prevent heat entering the tank over broad areas and careful design and material selection to 

deal with point conduction sources (structural supports, plumbing, cabling). The CPST TMP addressed 

three aspects of passive thermal control: (1) minimizing the insulation performance degradation due to 

point conduction elements penetrating the envelope; (2) composite materials for structural elements; and 

(3) application challenges of thick MLI to very large scale propellant tanks. 
NASA/TM—2014-218092 8 

2.2.1 Penetration Heat Leak Test

Conductive heat transfer in an in-space operational environment is due to structural, fluid, and 

instrumentation penetrations into the propulsion stage tank, and can be a significant contribution to the 

total system cryogenic tank heat loads. The impact is not limited to conduction through the penetrating 

element itself. The manner in which the penetration is integrated with surrounding insulation can also 

greatly affect heat loads. The state of the art predictive approach combines this effect with multilayer 

insulation scaling factors, which are used to adjust ideal predicted heat load values. These scale factors 

are typically based on the performance of analogous systems and have significant uncertainty.

The purpose of this testing was to characterize the conductive heat leak of a variety of flight 

representative fluid lines, electrical connections, and other penetrations through tank insulation (i.e., 

thermal shorts) and insulation methods on representative thick (>20 layers) cryogenic MLI systems. The 

parameters included: the attachment mechanism, the buffer material (for buffer attachment mechanisms 

only), the thickness of the buffer, and the penetration material. The methods of integration investigated 

were: the use of a buffer to thermally isolate the strut from the MLI and temperature matching the MLI on 

the strut. These were then compared to the case where no integration was performed. Several buffer 

materials were investigated, including: aerogel blankets, aerogel bead packages, Cryo-Lite (Johns 

Manville, Denver, CO) and even an evacuated vacuum space (in essence a no buffer condition). The 

testing was conducted at Kennedy Space Center’s (KSC) Cryogenic Test Laboratory (CTL), in 

specialized test chambers that provided the thermal and vacuum environment to simulate space-based 

conditions (Ref. 7). 

Over 23 tests were run to help characterize the thermal performance impacts of penetrating MLI. 

Testing included the development and fabrication of a new calorimeter and test method for twodimensional thermal performance testing. The testing included null testing of every blanket, no 

integration testing, buffer testing, and temperature matching testing with different size and material 

penetrations. The preferred method of isolating penetrations was shown to be the buffer method with 

Cryo-Lite as the best material to use as a buffer. (This methodology was subsequently employed to 

integrate strut penetrations with the surrounding insulation for the RBO test article, as shown in Fig. 3.) 

The thermal degradation or parasitic heat load was shown to be a function of strut diameter, buffer 

thickness, buffer material, warm boundary temperature, and penetration material. The buffer method was 

shown to be easier to develop, more robust, and less variable over multiple conditions and environments. 

The significance of this test effort is that the effects of penetrations on the thermal performance of 

cryogenic storage tanks has been investigated, showing the advantages of one or more penetration/MLI 

closeout methods. 

Figure 3.—Cryo-Lite buffer integrated around a penetration (within 

RBO Experiment Test Article).
NASA/TM—2014-218092 9 

2.2.2 Composite Strut Thermal Performance With LH2 Test 

One way to reduce the heat leak from a cryogenic support structure is to fabricate the supports out 

of a low conductivity material. However, even if low conductivity supports are employed, the overall 

performance of a tank’s insulation system is dependent on how that structural member integrates with the 

tank MLI. The work discussed above was limited to a cold boundary temperature of 77 K. This work was 

planned for a cold boundary of 20 K. 

The purpose of the testing discussed in this section was to measure the heat leak through a flight 

representative carbon fiber composite strut with one end of the strut at a simulated in-space thermal 

environment and the other end attached to a LH2 calorimeter. The resulting data would quantify 

performance degradation of MLI when the composite strut is integrated with various techniques (collars, 

socks or butt joints) to an MLI blanket. The test was to be performed at GRC’s SMiRF facility and 

employ a flat plate calorimeter from a previous research project. Two differently sized carbon fiber struts 

were planned for testing. Both struts consisted of IM7/8552 tape. The larger strut with a mid-span 

diameter of 15 cm. was originally designed for the Altair lunar lander program and had already undergone 

structural testing at room temperature. The smaller CPST-representative strut had a mid-span diameter of 

5 cm. 

Testing was terminated due to an unexpected vapor leak from the hydrogen calorimeter. The leak 

degraded the vacuum level inside the test chamber to the point where any strut heat leak measurements 

would have been overwhelmed by the increased heat load due to gas conduction in the test rig insulation. 

Thus, a technology gap remains for low conductivity carbon fiber struts especially at LH2 temperatures. 

Limited room temperature and LN2 temperature tests have been performed and the results have been 

promising, but further testing is desirable to support infusion into a cryogenic mission. 

2.2.3 Thick MLI Extensibility Study 

Long term (>2 weeks) in-space storage of large quantities of LH2 (>4 metric t) required for future 

exploration missions, without a significant loss of propellant due to boil off from radiation heat sources, 

will require the application of thick MLI (>7.5 cm) to the outer propellant storage tank walls. Traditional 

MLI systems (alternating layers of aluminized polymer films separated by polyester or silk netting) have 

been used for space missions for over 60 years, but limited thermal and structural knowledge exists for 

the fabrication, installation and venting performance of thick MLI systems applied to very large in-space 

LH2 storage tanks requiring minimal propellant boil off losses. A candidate insulation system is being 

demonstrated for LH2 (a BAC shield sandwiched between 30 layers of low density MLI (inner) and 30 

layers of higher density MLI (outer)) and LO2 (60 layers of MLI) utilizing a 1.2 m diameter test tank in 

ground based testing. Nevertheless, application of such technologies to large-scale tanks requires some 

study. The purpose of the study was to validate the relevancy of a thick MLI (>40 layers) tank attachment 

and MLI blanket fabrication approach from scaled ground and flight test tanks to a full scale CPS or depot 

application. The study also assesses options for attaching thick MLI to very large tanks and addresses 

associated heat loads for each option. 

Lockheed Martin’s Advanced Technology Center in Palo Alto, California was contracted to perform 

a study limited to traditional MLI concepts. Traditional MLI is defined herein as Double Aluminized 

Mylar (DAM) radiation shields separated by one or more layers of netting spacer material such as silk or 

Dacron. Advanced MLI concepts utilizing alternative spacer concepts are at lower TRL and were not 

included in the study. The study was focused on the LH2 tank for an Earth Departure Stage (EDS) as a 

representative large scale CPS. Most results should be extensible to other similar sized cryogenic 

propellant tanks. 

There is relatively little data for thick MLI. For this study, Lockheed Martin defined thick MLI to be 

thicker than 2.5 cm. Their optimized layer density was 14.5 layers/cm (37 layers/in.), which is similar in 

layer count to the CPST team definition above. The existing data suggests that there is an increase in the 

degradation fraction with MLI thickness. The contractor concluded that additional test data is needed 
NASA/TM—2014-218092 10

before thick MLI can be used with confidence for flight applications. Their concerns included the 

following: 

• Traditional MLI blankets are typically fabricated into blankets of widths on the order of 1.2 m.

Due to the large tank size, the relatively small blanket width results in a substantial number of 

seams and total seam length. 

• Available data for MLI performance on relatively large tanks (all 2 to 3 m in diameter) is sparse 

and cannot be clearly explained. There are too many unknowns to reach a clear conclusion. It is 

recommended that a large tank test be conducted with tight controls on layer density variations, 

minimization of seams and numerous measurements of DAM emissivity.

• There is little data on the performance repeatability of multiple builds of MLI systems. Large 

uncertainty may be a reality when using traditional MLI. 

• A review of environmental test data found examples where MLI was able to withstand testing for 

acceleration, acoustic and vibration loads, but a full literature survey was not completed. There 

appears to be a lack of understanding or knowledge on how to structurally model MLI in a 

meaningful manner. 

Lockheed Martin recommends continued use of the well-known “Lockheed Equations” (Ref. 8) for 

predicting thermal performance of MLI. These equations are based on correlation of calorimeter test 

results representing ideal conditions of MLI layup with minimal perturbations. Losses due to seams, 

penetrations, compression and other factors should then be added to predictions. The basis for alternative 

equations is not clear and these equations have a lesser amount of validation. This recommendation may 

not be widely accepted within NASA. 

Finally, the contractor recommended that a large scale (>4 m test tank) ground storage test be 

conducted to demonstrate thick MLI technology specifically developed for large storage tanks in a 

simulated LEO thermal and vacuum environment to reduce the risk of applying unproven thick MLI 

technology to future space missions. 

The significance of this study effort is that a set of factors related to large scale, long duration space-based 

storage of LH2 has been identified.

2.3 Liquid Acquisition Device (LAD) Outflow and Line Chill Test

When transferring propellant in space, it is necessary to transfer single-phase liquid from a propellant 

tank to either an engine or another storage vessel. In Earth’s gravity field or under acceleration during 

“significant” thrusting, propellant transfer is fairly simple: single-phase fluid is transferred by opening a 

valve at the bottom of the propellant tank and installing an anti-vortex baffle over the tank outlet to 

prevent vapor and gas ingestion into the outlet. In low gravity where fluid does not sufficiently cover the 

tank outlet, withdrawing single-phase fluid becomes a challenge. A propellant management device (PMD)

is required to ensure single-phase flow, depending on the gravitational environment. One type of PMD, a 

LAD uses capillary flow and surface tension with a screen/channel device to acquire liquid. Capillary 

flow LADs have been well characterized for storable propellants (propellants that are liquids at room 

temperature) for in-space propulsion needs, and capillary flow LADs have also been characterized over 

a wide range of operating conditions for cryogenic fluids such as LN2, LO2 and LCH4. Some 

characterization has been performed with LH2, but additional work is required to characterize LAD 

performance in LH2 over a range of conditions that will validate their performance for the CPST flight. 

Once bubble free liquid is acquired from the tank, it needs to be transferred through a feedline without creating 

significant vapor. This requires that the feed-line be pre-chilled. Line chill down is commonly accomplished by 

absorbing heat into some sacrificial propellant.

The purpose of the testing was to: (1) measure the static and the dynamic screen bubble point

pressures using LH2 as the test fluid for different LAD screen sizes, while conducting a parametric study 

of LAD screen, LAD channel, frictional and fluid head pressure drops for various LH2 LAD outflow rates

using new, smaller pore size LAD screens, and (2) investigate efficient options for chill down of transfer 
NASA/TM—2014-218092 11

line and quantify the LAD stability (no LAD breakdown) due to transfer line chill down transient 

dynamic pressure perturbations during outflow. Both sets of objectives were accomplished in the same 

test program conducted at GRCs SMiRF, which provided the thermal and vacuum environment to 

simulate space-based conditions. 

2.3.1 LAD Outflow Test

Inverted vertical outflow testing of two 325 x 2300 full-scale LAD channels in LH2 was completed. 

One was a standard LAD screen channel, while the other was thermally flight representative due to the 

presence of a perforated plate and internal cooling from a Thermodynamic Vent System (TVS). The 

LADs were mounted in a tank to simulate 1-g outflow over a wide range of LH2 temperatures (20.3 to 

24.2 K), pressures (100 to 350 kPa), and flow rates 0.010 to 0.055 kg/s). Results indicate that the effects 

on predicted breakdown height (inversely related to bubble point pressure) are dominated by liquid 

temperature, with a second order dependence on outflow rate through the LAD (Ref. 9). The lowest liquid 

level breakdown heights (i.e., bubble breakthrough occurs after more of the screen is exposed, thus is 

indicative of a higher bubble point) are always achieved in the coldest liquid states for both channels, 

consistent with bubble point theory. Higher flow rates cause the standard channel to break down earlier 

than the flight channel, where the presence of the perforated plate is believed to enhance wicking and thus 

screen retention during outflow. Both the heat exchanger and subcooling the liquid are shown to improve 

LAD performance.

Due to fabrication difficulties a 450 x 2750 screen channel was not available for use in the inverted 

outflow test. Testing with 325 x 2300 was judged to be adequate to demonstrate that screen channel 

devices that will be designed for the requirements of the CPST project were capable of handling the 

required flow rates without breaking down and admitting bubbles. Lack of a 450 x 2750 screen channel 

for the CPST payload will likely result in the use of He pressurant (which improves bubble point 

pressure), and probably a LAD TVS system as well. 

The significance of this test effort is that a LAD for LH2 flow rates representative for CPST has been 

demonstrated on the ground, offering potential to demonstrate the transfer of LH2 from tank to tank in space. 

2.3.2 Line Chill Test

Operation of a cryogenic transfer system is complicated by the requirement to chill it down to allow 

single phase liquid transfer. The full flush method has been used to cool engine feed lines for upper stages 

since the 1960s, but is very wasteful in propellant. Pulse and trickle flow also bear a great similarity to the 

engine feed line cooling techniques investigated in the Atlas Reliability Enhancement Program (AREP) 

(Ref. 10). This transient must be well understood so as to accurately determine the amount of cryogen 

required, and the time required to complete the chill down process. Transfer line chill down must also be 

performed in a manner that will not result in unacceptable pressure fluctuations or stresses in the system 

that could damage components. The goal of transfer line chill down is to enable the transfer of singlephase liquid at the required condition from a storage vessel to its destination: either another storage vessel 

or an engine system. Understanding the nature of the transient chill down process, requires an 

understanding of the fluid physics and the important parameters that affect chill down, including: mass 

flux, acceleration, the type of two-phase flow induced, heat transfer mechanisms during chill down, and 

the transfer system physical configuration.

Line chill down testing was conducted to study the chill down of two representative transfer lines 

utilizing two approaches: trickle flow and pulsed. The testing suggests pulsed chill down, with a duty 

cycle chosen to minimize either time or mass, as having the most promise. Chill down of a representative 

line was achievable in less than 90 s while ensuring that vapor free liquid would be available up to the 

inlet of the receiver tank. 

The significance of this test effort is that an effective technique for line chill for LH2 has been 

demonstrated on the ground, offering potential to support the transfer of LH2 from tank to tank in space. 
NASA/TM—2014-218092 12

2.4 Analytical Tool Development 

Analytical tools are critical for the prediction of space flight system performance (Ref. 11). Analytical 

tools matured in FY12 under the CPST project include tools to support overall mission performance 

prediction of CFM system/ subsystems, cryogenic storage thermodynamic and fluid dynamic modeling 

tools to predict fluid behavior, and component tools to guide the design of component hardware. 

The intent of this focus area is to develop and validate analytical tools to be used for the design of the 

CPST-TDM flight hardware and to predict the fluid dynamics and thermodynamics (heat and mass 

transfer) of the CPST systems/subsystems in a relevant environment. The development and validation of 

analytical tools is planned to continue for the life of the project, culminating in the validation and final 

model refinements against CPST-TDM post-mission flight data. Ideally, the tools developed should be 

extensible to larger geometric scales (on the order of 5 to 10 m tank diameter) and longer storage durations 

(on the order of years) compared to the CPST-TDM flight experiment.

A brief description of available codes and their current capabilities are listed in Tables 2 and 3. The 

list below includes the commercial codes for completeness. The description is focused on those 

capabilities relevant to modeling the mission phases of: self-pressurization, pressurization, mixing (with 

or without subcooling), transfer line chill down, tank chill down, and tank filling. 

Table 2.—NASA Developed Tools/Codes for Sizing, Optimization, and Analysis

Name Platform CFM Technology 

Addressed

Nodes Features

Cryogenic Analysis 

Tool (CAT)

Excel/VBA Passive and active thermal 

control

Single Includes single node (homogenous thermodynamic) model 

for self-pressurization

CryoSim

(Cryogen Storage 

Integrated Model)

Fortran Passive and active thermal 

control

Can be coupled with TankSim to predict self-pressurization 

and pressure control

TankSim Fortran Self-pressurization, spray 

bar/axial jet pressure control 

and preliminary model for 

pressurization

MultinodeSingle node represents a zero-thickness interface between 

liquid and ullage and the wetted and dry portion of the tank 

wall. Flat interface or spherical bubble supported for selfpressurization. Creates REFPROP (Ref. 12)-generated 

property tables and interpolates from these tables.

MLI Ascent Venting Fortran MLI performance during 

launch ascent

Single Combines continuum and kinetics theory to predict MLI 

layer transient temperature and pressure

CPPPO (Computational Propellant 

and Pressurization 

Program—One 

dimensional)

Excel/VBA Self-pressurization and 

pressurization

MultinodeSingle node represents a zero-thickness interface between 

liquid and ullage and the wetted and dry portion of the tank 

wall. Flat interface or spherical bubble supported for selfpressurization.

GFSSP (Generalized 

Fluid System 

Simulation Program)

Fortran Fluid and heat transfer 

networks

MultinodeProvides graphical user interface (GUI) for problem setup. 

Includes a large number of fluid network element options. 

Implementation of heat and mass transfer across 

liquid/ullage interface using a zero-thickness node is in 

progress

NVFILL Fortran Tank chill down and novent-fill (NVF)MultinodeIncludes finite-element shell conduction model of heat 

transfer in thin tank walls and a Lagrangian spray model. 

Also supports self-pressurization analysis (locked up tank 

with no inflow).
NASA/TM—2014-218092 13

Table 3.—Commercial Codes Used to Model CFM Fluid and Thermodynamic Processes

Code (Vendor) Type CFM Technology 

Addressed

Standard/Available Options

Thermal Desktop with 

RadCAD and 

SINDA/FLUINT

(C & R Technologies)

Fortran Fluid and heat 

transfer networks

• GUI for problem setup (Thermal Desktop and SINAPS)

• Heat and mass transfer across liquid/ullage interface using 

a zero-thickness node (requires user coding)

• Includes a large number of fluid network elements

FLOW-3D

(Flow Sciences)

CFD

(computational

fluid dynamics)

Two phase flow 

simulations SOA 

for tank sloshing 

and settling

• General non-inertial frame of reference

• Mass transfer between liquid and ullage using a Schrage 

type kinetic equation

• Lagrangian spray coupled with volume of fluid (VOF) (no 

atomization, secondary breakup, turbulence/spray 

interactions, or coalescence models)

• Grid generation by the use of Cartesian cut cells

Fluent

(ANSYS)

CFD Two phase flow 

simulations

• Specialized boundary conditions and physical models

• Fixed interface shape, sharp interface model (not using 

VOF)

• General non-inertial frame of reference

• Mass transfer between liquid and ullage using a Schrage 

type kinetic equation

• Lagrangian spray coupled with VOF (in progress)

• Atomization, secondary breakup, turbulence/spray 

interactions, or coalescence

• Turbulence models

• General unstructured (and polyhedral) grids

• Shell conduction model (tank chill down)

The significance of this analytical tool development effort is that a foundation has been established 

for the development and validation of analytical tools necessary to predict the fluid dynamics and 

thermodynamics (heat and mass transfer) of the CFM systems/subsystems in a relevant environment, for 

the design of hardware for flight demonstrations (CPST) and ideally extensible to future full scale space 

missions (CPS and/or depots) with extended in-space storage durations (> 6 months). 

3.0 LO2 ZBO Test 

LO2 ZBO capability is believed achievable using a 90 K cryocooler with the tubing network located 

on the tank wall, using the wall to distribute the cooling.

The purpose of this test is to control tank pressure and ultimately LO2 temperature with the active 

cooling system in a manner that demonstrates robust ZBO. The aspects to understand and validate this are 

the effect of heat removal rate and its controllability on cryogenic tank pressure. In addition, the 

distributed cooling system’s ability to reduce and control the test tank surface temperature with its 

inherent variations, particularly the anticipated hotter temperatures around the vent tube at the top of the 

tank, will be tested. As such, the ZBO system’s ability to finely control tank pressure will be tested. The 

cryocooler will be tested at 25 percent excess capacity, to determine the tank pressure response, and 

likewise test the system at 25 percent under capacity. In addition, the cooling network’s ability to remove 

heat at two temperatures will be validated: one to demonstrate LO2 ZBO storage at representative 

propulsion system pressures of 172 kPa (25 psi)—corresponding to a LO2 temperature 96 K—and the 

other at 82 K, which is comparable to the BAC shield case in the LH2 RBO testing.

This LO2 ZBO test is an important technology step to demonstrate the ability to control tank pressure 

via a distributed active cooling network, which has not been previously done. Thus, the level of active 

cooling will be coupled with tank pressure and the fluid’s response will be studied. The balancing of the 

tank heat removal with the nominal passive tank heating rate to achieve ZBO and then disturbing that 

balance to understand the tank pressure response to varying heat removal rates is critical knowledge. A 
NASA/TM—2014-218092 14

robust investigation into the fluid’s response to different cooling levels will create curves that show the 

heat removal effect on tank pressure.

The test article is nearly complete with testing to occur by the end of the calendar year.

4.0 Conclusion

The focus of the FY12 CPST technology maturation effort was to mature selected CFM technologies 

(nominally to a TRL of 5) through ground based testing, and to show through studies the relevance of the 

CFM technologies on the CPST flight demonstration to full-scale applications. This effort successfully

mitigated budget and schedule risk anticipated in the development of the cryogenic fluid system payload 

for the CPST flight demonstration. The CPST project is ready to proceed to flight system development 

with many of these technologies. 

References

1. Motil, S.M., Kortes, T.F., Meyer, M.L., Taylor, W.J., and Free, J.M., “Concept Design of Cryogenic 

Propellant Storage and Transfer for Space Exploration,” 63rd International Astronautical Congress; 

Naples; 2-5 Oct. 2012; Italy. 

2. Feller, J.R., Plachta, D.W., Mills, G. and McLean, C., “Demonstration of a Cryogenic Boil-Off 

Reduction System Employing an Actively Cooled Thermal Radiation Shield,” Presented at the 16th 

International Cryocooler Conference, held May 17-20, 2008 in Atlanta, Georgia. 

3. Christie, R.J., Plachta, D.W., and Guzik, M.C., “Integration of a Reverse Turbo-Brayton Cryocooler 

with a Broad Area Cooling Shield and a Heat Pipe Radiator,” Presented at Thermal & Fluids Analysis 

Workshop, July 29-Aug 2, 2013, Daytona Beach, Florida. 

4. Wood, J. and Foster, L., "Acoustic and Thermal Testing of an Integrated Multilayer Insulation and 

Broad Area Cooling Shield System," Presented at the 25th Space Cryogenics, June 23-25, 2013, 

Girdwood, Alaska. 

5. Plachta, D.W., and Guzik, M.C. “Cryogenic Boil-Off Reduction System Scaling Study,” Presented at 

the 25th Space Cryogenics Workshop, June 23-25, 2013, Girdwood, Alaska. 

6. Guzik, M.C., “Development of a Cryogenic Analysis Tool for the Scaling of Cryogenic Boil-Off 

Reduction Systems,” Presented at Thermal & Fluids Analysis Workshop, July 29-Aug 2, 2013, 

Daytona Beach, Florida

7. Johnson. W.L., Kelly, A.O., and Fesmire, J.E., “Thermal Degradation of Multilayer Insulation Due to 

the Presence of Penetrations,” Presented at the Cryogenic Engineering Conference, June 17-21, 2013, 

Anchorage, Alaska.

8. Keller, C.W., Cunnington, G.R., and Glassford, A.P., “Thermal Performance of Multilayer 

Insulations,” NASA CR-134477, April 1974.

9. Hartwig, J., Chato, D., McQuillen, J. Vera, J. Kudlac, M., and Quinn, F., “Screen Channel Liquid 

Acquisition Device Outflow Tests in Liquid Hydrogen,” Presented at the 25th Space Cryogenics 

Workshop, June 23-25, 2013, Girdwood, Alaska. 

10. Schuster, J.R.; Howell, D. J.; Lucas, S. L., Jr.; Haberbusch, M. S.; Gaby, J. D.; Van Dresar, N. T.; and 

Wadel, M. F., “Cold flow testing of revised engine chilldown methods for the Atlas Centaur,” 

Presented at the 32nd AIAA/ ASME/ SAE/ ASEE Joint Propulsion Conference & Exhibit: July 1-3, 

1996, Lake Buena Vista, Florida.

11. Doherty, M.P., Gaby, J.D., Salerno, L.J., and Sutherlin, S.G., “Cryogenic Fluid Management 

Technology For Moon and Mars Missions,” NASA/TM—2010-216070, AIAA–2009–6532 Presented 

at Space 2009, September 14-17, 2009, Pasadena, CA 

12. Lemmon, E.W., Huber, M.L., and McLinden, M.O. NIST Standard Reference Database 23: 

Reference Fluid Thermodynamic and Transport Properties-REFPROP, Version 9.1, National Institute 

of Standards and Technology, Standard Reference Data Program, Gaithersburg, 2013. 

REPORT DOCUMENTATION PAGE Form Approved 

OMB No. 0704-0188 

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the 

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this 

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB 

control number. 

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-02-2014 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Cryogenic Propellant Storage and Transfer (CPST) Technology Maturation: Establishing a 

Foundation for a Technology Demonstration Mission (TDM) 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Doherty, Michael, P.; Meyer, Michael, L.; Motil, Susan, M.; Ginty, Carol, A. 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 645454.01.04.04.01 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION 

 REPORT NUMBER

E-18761-1 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S 

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING 

 REPORT NUMBER

NASA/TM-2014-218092 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 28, 20, and 14 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

As part of U.S. National Space Policy, NASA is seeking an innovative path for human space exploration, which strengthens the capability to 

extend human and robotic presence throughout the solar system. NASA is laying the groundwork to enable humans to safely reach multiple 

potential destinations, including asteroids, Lagrange points, the Moon and Mars. In support of this, NASA is embarking on the Technology 

Demonstration Mission Cryogenic Propellant Storage and Transfer (TDM CPST) Project to test and validate key cryogenic capabilities and 

technologies required for future exploration elements, opening up the architecture for large cryogenic propulsion stages (CPS) and 

propellant depots. The TDM CPST project will provide an on-orbit demonstration of the capability to store, transfer, and measure cryogenic 

propellants for a duration which is relevant to enable long term human space exploration missions beyond low Earth orbit (LEO).

Recognizing that key cryogenic fluid management technologies anticipated for on-orbit (flight) demonstration needed to be matured to a 

readiness level appropriate for infusion into the design of the flight demonstration, the NASA Headquarters Space Technology Mission 

Directorate authorized funding for a one-year (FY12) ground based technology maturation program. The strategy, proposed by the CPST 

Project Manager, focused on maturation through modeling, studies, and ground tests of the storage and fluid transfer Cryogenic Fluid 

Management (CFM) technology sub-elements and components that were not already at a Technology Readiness Level (TRL) of 5. A 

technology maturation plan (TMP) was subsequently approved which described: the CFM technologies selected for maturation, the ground 

testing approach to be used, quantified success criteria of the technologies, hardware and data deliverables, and a deliverable to provide an 

assessment of the technology readiness after completion of the test, study or modeling activity. This paper will present the testing, studies, 

and modeling that occurred in FY12 to mature cryogenic fluid management technologies for propellant storage, transfer, and supply, to 

examine extensibility to full scale, long duration missions, and to develop and validate analytical models. Finally, the paper will briefly 

describe an upcoming test to demonstrate Liquid Oxygen (LO2) Zero Boil-Off (ZBO).

15. SUBJECT TERMS

Cryogenic fluids; Technology readiness level; Space exploration 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

22 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18