September 2005

NASA/TM-2005-213913

Thermal Structures Technology Development

for Reusable Launch Vehicle Cryogenic

Propellant Tanks

Theodore F. Johnson, Roderick Natividad, H. Kevin Rivers, and Russell W. Smith

Langley Research Center, Hampton, Virginia
The NASA STI Program Office . . . in Profile

Since its founding, NASA has been dedicated to the

advancement of aeronautics and space science. The

NASA Scientific and Technical Information (STI)

Program Office plays a key part in helping NASA

maintain this important role.

The NASA STI Program Office is operated by

Langley Research Center, the lead center for NASA’s

scientific and technical information. The NASA STI

Program Office provides access to the NASA STI

Database, the largest collection of aeronautical and

space science STI in the world. The Program Office is

also NASA’s institutional mechanism for

disseminating the results of its research and

development activities. These results are published by

NASA in the NASA STI Report Series, which

includes the following report types:

• TECHNICAL PUBLICATION. Reports of

completed research or a major significant phase

of research that present the results of NASA

programs and include extensive data or

theoretical analysis. Includes compilations of

significant scientific and technical data and

information deemed to be of continuing

reference value. NASA counterpart of peerreviewed formal professional papers, but having

less stringent limitations on manuscript length

and extent of graphic presentations.

• TECHNICAL MEMORANDUM. Scientific

and technical findings 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. Scientific and

technical findings by NASA-sponsored

contractors and grantees.

• CONFERENCE PUBLICATION. Collected

papers from scientific and technical

conferences, symposia, seminars, or other

meetings sponsored or co-sponsored by NASA.

• SPECIAL PUBLICATION. Scientific,

technical, or historical information from NASA

programs, projects, and missions, often

concerned with subjects having substantial

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific and

technical material pertinent to NASA’s mission.

Specialized services that complement the STI

Program Office’s diverse offerings include creating

custom thesauri, building customized databases,

organizing and publishing research results ... even

providing videos.

For more information about the NASA STI Program

Office, see the following:

• Access the NASA STI Program Home Page at

http://www.sti.nasa.gov

• E-mail your question via the Internet to

help@sti.nasa.gov

• Fax your question to the NASA STI Help Desk

at (301) 621-0134

• Phone the NASA STI Help Desk at

(301) 621-0390

• Write to:

 NASA STI Help Desk

 NASA Center for AeroSpace Information

 7121 Standard Drive

 Hanover, MD 21076-1320
National Aeronautics and

Space Administration

Langley Research Center 

Hampton, Virginia 23681-2199

September 2005

NASA/TM-2005-213913

Thermal Structures Technology Development

for Reusable Launch Vehicle Cryogenic

Propellant Tanks

Theodore F. Johnson, Roderick Natividad, H. Kevin Rivers, and Russell W. Smith

Langley Research Center, Hampton, Virginia
Available from:

NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS)

7121 Standard Drive 5285 Port Royal Road

Hanover, MD 21076-1320 Springfield, VA 22161-2171

(301) 621-0390 (703) 605-6000

The use of trademarks or names of manufacturers in the report is for accurate reporting and does not

constitute an official endorsement, either expressed or implied, of such products or manufacturers by the

National Aeronautics and Space Administration.
Thermal Structures Technology Development for Reusable Launch

Vehicle Cryogenic Propellant Tanks

Theodore F. Johnson∗, Roderick Natividad+, H. Kevin Rivers

Δ

, and Russell Smith∗∗

NASA, Langley Research Center

Abstract

Analytical and experimental studies conducted at the NASA, Langley Research Center

(LaRC) for investigating integrated cryogenic propellant tank systems for a reusable launch

vehicle (RLV) are described. The cryogenic tanks are investigated as an integrated tank system.

An integrated tank system includes the tank wall, cryogenic insulation, thermal protection system

(TPS) attachment sub-structure, and TPS. Analysis codes are used to size the thicknesses of

cryogenic insulation and TPS insulation for thermal loads, and to predict tank buckling strengths

at various ring frame spacings. The unique test facilities developed for the testing of cryogenic

tank components are described. Testing at cryogenic and high-temperatures verifies the integrity 

of materials, design concepts, manufacturing processes, and thermal/structural analyses. Test

specimens ranging from the element level to the subcomponent level are subjected to projected

vehicle operational mechanical loads and temperatures. The analytical and experimental studies

described in this paper provide a portion of the basic information required for the development of

light-weight reusable cryogenic propellant tanks.

INTRODUCTION

One of the goals for the next generation of space launch vehicles is an order of magnitude

reduction in the cost of delivering a payload to orbit. Recent studies on space transportation by

the National Aeronautics and Space Administration (NASA) [1, 2] indicate that a single-stageto-orbit (SSTO) reusable launch vehicle (RLV), fueled by liquid hydrogen (LH2) and liquid 

oxygen (LOx) has the potential to reach this goal. The recent X-33/RLV Program was a

partnership between NASA and industry to create a viable RLV [3]. In this program, current and

emerging technologies were utilized to develop and build the X-33, a 1/2 scale RLV

demonstrator/test-bed vehicle [4]. These technologies were being pursued to develop an RLV

that has efficient, and airline-like operation with 7-day refurbishment cycles between missions to

reduce the operational costs, thereby reducing the cost to place a payload in orbit. One of the

key technology drivers identified by this program was the use of reusable cryogenic tanks.

Large reusable cryogenic tanks will be required to contain the LH2 and LOx for an SSTO

RLV. Cryogenic tank development is critical for an RLV because the tanks may comprise as

much as 70 to 80 percent of the volume of the vehicle, as shown in Figure 1 for two generic

RLVs [1]. The development and fabrication of reusable cryogenic tanks is one of the significant

 ∗ Aerospace Engineer, Mechanics of Structures and Materials Branch, Mail Stop 188E, Hampton, VA 23681-2291,

Theodore.F.Johnson@nasa.gov, (757) 864-5418, Member, AIAA. + Former Aerospace Engineer, Metals and Thermal Structures Branch, Mail Stop 396, Hampton, VA 23681-2291 Δ Aerospace Engineer, Metals and Thermal Structures Branch, Mail Stop 396, Hampton, VA 23681-2291 ∗∗ Aerospace Engineer, Metals and Thermal Structures Branch, Mail Stop 396, Hampton, VA 23681-2291
2

technical challenges to be overcome to develop an operable RLV [4]. Large expendable

cryogenic tanks have been made for launch vehicles, but the durability of flight-weight reusable

cryogenic propellant tanks has not been demonstrated. The cryogenic tanks of an RLV must not

only function as pressure vessels at cryogenic temperatures, but they also must carry primary 

structural loads and support the thermal protection system (TPS). The cryogenic tanks, along

with the TPS, must be easy to maintain, easy to repair, and reusable for the life of the vehicle.

In this paper, candidate reusable cryogenic propellant tank concepts are evaluated as a part of

an integrated tank system that includes TPS, TPS attachment sub-structure, cryogenic insulation,

and integrated vehicle health monitoring (IVHM) [5]. Thermal and structural analyses were used

to compare several candidate combinations of cryogenic tanks and TPS. Combined thermal and

mechanical tests of cryogenic insulation, adhesives, structural elements, and subcomponents

were performed, with the thermal and mechanical loading becoming increasingly complex as the

specimen size increased. Test environments included: temperatures ranging from 20 K (-423 ̊F)

to 810 K (1000 ̊F), pressures ranging from atmospheric to 372 kPa (54 psig), and mechanical

loads of uniaxial tension, compression, and biaxial tension. This paper describes integrated

cryogenic tank analysis, testing, and test facilities development at the NASA, Langley Research

Center (LaRC), which support technology development for a reusable LH2 tank for an RLV.

ANALYTICAL STUDIES OF INTEGRATED CRYOGENIC SYSTEMS

The thermal and structural performance of the tank, cryogenic insulation, and TPS should be

considered as a system to develop the best cryogenic tank for an RLV. Several combinations of

LH2 tank walls and TPS are analyzed to identify lightweight and thermally efficient concepts.

An example of an integrated tank system design is displayed in Figure 2. The titanium (Ti)

sandwich wall acts as a pressure vessel, cryogenic insulation, primary structure, and TPS

support. Extra low interstitial (ELI) titanium is used to reduce hydrogen embrittlement. The

exterior facesheet does not have to be impermeable to LH2 or LOx thus, mechanical fasteners can 

penetrate the external facesheet to attach TPS. Gaskets beneath the TPS panel gaps prevent

subsurface hot gas flow during re-entry.

Figure 1. Generic RLV with an aft- or forward-located LH2 tank.
3

Integrated tank concepts considered in the analytical studies are depicted in Figure 3 and

listed in Table 1. The external foam concepts are based upon a forward-located, IM7/977-2

graphite-epoxy (Gr-Ep) ring and stringer stiffened tank concept from the X-33 Phase I Rockwell

vehicle [6] that used adhesively bonded RohacellTM as the cryogenic foam insulation and either

alumina enhanced thermal barrier (AETB) or tailorable advanced blanket insulation (TABI) as

the TPS. Honeycomb cryogenic insulation concepts consist of a sandwich tank wall with an

evacuated core for insulation and superalloy/honeycomb (SA/HC) metallic panels as TPS.

Several material combinations were considered for the honeycomb sandwich tank wall. The

Aluminum 2219-T87 sandwich tank concept Al/Ti/Al was considered because aluminum is

compatible with LH2 and the titanium core can either be brazed or adhesively bonded to the

aluminum facesheet. The titanium honeycomb core provides sufficient insulation if evacuated 

and enhances the structural stability of the tank. The concept with titanium facesheets and

titanium honeycomb core was investigated because of the high strength and operation 

temperature (645 K, 700 ̊F) of a brazed Ti/Ti/Ti concept. A concept consisting of IM7/5260 

graphite-bismaleimide (Gr-BMI) facesheets with titanium honeycomb core was considered 

because Gr-BMI has a moderately high operating temperature (450 K, 350 ̊F). In the final

concept, Gr-BMI facesheets were combined with a HexcelTM glass reinforced phenolic (HRP).

The Gr-BMI/HRP/Gr-BMI concept has reduced thermal conductivity because of the nonmetallic honeycomb core. Both Gr-BMI sandwich concepts can be adhesively bonded or cocured sandwich structures. Thermal and structural analyses were used to compare the weights

and strengths of these integrated tank systems. In the thermal analysis, the thicknesses of the

cryogenic insulation and TPS insulation were sized to withstand the various operational

temperatures. Thermal loading cases were from the ground-hold filling of the tanks, re-entry,

and soak-through of heat energy after re-entry. In the structural analysis, the structural stability

of the tank wall was investigated for a various concept with various ring frame spacings. The

only mechanical loading used was compression axial line loading.

Figure 2. Example of an all-metallic Ti/Ti/Ti sandwich cryogenic tank with metallic TPS in an

integrated tank system concept for a RLV.
4

Thermal Sizing Study

A thermal sizing study of the designs in Figure 3 was conducted to compare the thermal

performance of each cryogenic tank concept. The cryogenic insulation thicknesses and TPS

were sized to maintain temperatures within the limits shown in Figure 3 using a one-dimensional

non-linear finite element sizing code called TPSSizer [7]. Temperature and pressure dependent

material properties were obtained from industry and in reference [8]. The weight of the TPS and

cryogenic insulation (honeycomb core for sandwich tank walls) was calculated and compared for

each tank/TPS system. The thermal mass of the tank wall skin (sized to withstand the

pressurization load and to limit LH2 permeation in the case of composite walls) was included in

the thermal sizing analysis. Based on reports from industry [9] and National Aerospace Plane

(NASP) Program data, four to eight plies of Gr-Ep and Gr-BMI was required minimum thickness

to resist hydrogen (H2) permeation of a LH2 tank.

TPSSizer Program

The TPSSizer program [7] uses an implicit, one-dimensional transient finite element

formulation. Thermal properties may be a function of temperature and pressure and are updated

at each time step. Radiation to space is assumed at the surface node, and the radiation vector is

converged in an iterative fashion. These one-dimensional models include corrections for such

things as coatings, adhesives, fasteners, and strain isolation pads. Detailed two-dimensional

models have been used to varying degrees to validate the simplifications made in the onedimensional models. Resizing of the TPS is accomplished through a repeated iterative procedure

until all the temperature limits are satisfied within ±2 ̊F. At the beginning of each iteration, new

Figure 3. Concepts used in the thermal sizing and structural buckling studies.
5

TPS and cryogenic insulation thicknesses are calculated based on the maximum temperatures

computed in the previous analysis. Generally, only two to six sizing iterations are required. The

cryogenic insulation is sized to satisfy two possible constraints. The first constraint, a ground

hold condition, assumes that steady-state temperature conditions has been reached while the

vehicle sits on the launch pad. The cryogenic insulation is sized to maintain a specified 

minimum temperature at the TPS to cryogenic insulation interface. By assuming steady-state

temperature conditions, the sizing procedure is reduced to a simple equation based upon the TPS

thickness. There is also iterative sizing of the cryogenic insulation based on the maximum

transient temperature limit of the tank wall during re-entry/soak-through.

Table 1. Honeycomb sandwich cryogenic tank walls.

Facesheet/

core/

facesheet

Abbreviation

aluminum/

titanium/

aluminum

Al/Ti/Al

titanium/

titanium/

titanium

Ti/Ti/Ti

graphite-bismaleimide/

titanium/

graphite-bismaleimide

Gr-BMI/Ti/Gr-BMI

graphite-bismaleimide/

HexcelTM glass reinforced phenolic/

graphite-bismaleimide

Gr-BMI/HRP/Gr-BMI

The criteria used for sizing the thickness of the cryogenic insulation and TPS were: a constant

temperature of 20 K (-423 ̊F) at the tank wall, prevention of frost build-up and air liquefaction on

the cryogenic insulation surface of the tank during ground-hold, and limiting the maximum

operational temperatures (shown in Figure 3) of the various materials used in the structure and

TPS of the vehicle during re-entry. All of the sandwich tanks had a lower limit temperature of

115 K (-250°F) on the outer facesheet during ground-hold to minimize air liquefaction. The

windward side heating load1 was for the vehicles in Figure 1. An example of the heating load is

shown in Figure 4. These cylindrical study vehicles were used because the vehicles did not have

cavities over the cryogenic tanks or an aeroshell, thus the tank wall-insulation-TPS was layered 

as in Figure 2.

The mass of the cryogenic insulation and TPS are plotted in Figure 5 versus the x-location

along the length of the windward centerline of the vehicle as defined in Figure 1. Although the

LH2 tank does not extend the entire length of the vehicle, the tank may be located either forward 

or aft as shown in Figure 1. An all-metallic concept, Ti/Ti/Ti with metallic TPS, is lighter than 

the other organic/ceramic or metallic concepts. The all-metallic titanium concept is lighter

 1 The windward side heating load was provided by Kay Wurster of the Vehicle Analysis Branch (VAB) at

LaRC.
6

because the relatively high use-temperature of titanium requires less TPS for re-entry heating and 

the evacuated titanium honeycomb core acts as an effective insulator, eliminating the need for

cryogenic insulation. The non-metallic core of the Gr-BMI/HRP/Gr-BMI concept has a reduced

thermal conductivity that increased the insulative capacity compared to the metallic core

specimens making this concept lighter than the other sandwich concepts except for the alltitanium concept. The Gr-BMI/HRP/Gr-BMI concept is also lighter than the stiffened Gr-Ep

concepts because the outer facesheet can be allowed to cool to a temperature of 115 K (-250 ̊F)

during ground-hold. The upper use temperature limit and low heat capacity of the RohacellTM

(480 K, 400 ̊F) external cryogenic foam increases the TPS thickness for the stiffened Gr-Ep

structure making the concept heavier than the all titanium concept. The Al/Ti/Al and GrBMI/Ti/Gr-BMI tanks are the heaviest concepts. These two concepts do not effectively use the

higher temperature capability of the titanium core because of the lower maximum operating

temperature of the facesheets during re-entry. The aluminum concept is not included in the

structural buckling study.

Structural Buckling Study

A structural buckling study was performed with the lightest weight tank concepts based on the

thermal insulation weight identified in the thermal sizing study. The feasibility and cost of

fabrication were not factors in the evaluation. The weight of the structural elements of the tank

and the effects of varying ring frame spacing on tank stability were compared in the structural

buckling study. This structural study focused on determining the axial buckling load of an

unpressurized LH2 tank under axial compression line loading, Nx, during the ground-hold and 

soak-through phase of a vehicle mission-profile. Pressure stabilization of the LH2 tank was

ignored to design a tank with sufficient strength to support a full LOx tank in the event of LH2

tank depressurization. A non-linear, finite difference, shell-of-revolution code, BOSOR4 [10],

was used as the analysis tool in the study.

BOSOR4 Program

The BOSOR4 program was based upon the Vlasov-Sanders shell theory and was used to

conduct both non-linear pre-buckling and buckling analyses [11]. Modal iteration was used in

Figure 4. An example of the heat flux on the windward centerline and side of the generic

vehicle for a LOx tank.
7

BOSOR4 to determine the eigenvalues. A representative model of a sandwich tank wall

structure and the boundary conditions used in the analysis are shown in Figure 6. A detailed 

description of the boundary conditions and models can be found in reference [12]. The model

had periodic boundary conditions simulating an infinitely long cylindrical tank. The ring frame

design was kept fixed in this study to reduce the number of variables, but the ring frame spacing 

was varied for all of the models. The core thicknesses from the thermal sizing study were also

used in the buckling models.

The concepts evaluated in the structural buckling study are Ti/Ti/Ti, Gr-BMI/HRP/Gr-BMI,

and Gr-BMI/Ti/Gr-BMI sandwich concepts, and Gr-Ep ring and stringer stiffened tank concepts.

The structural weight of the tank (not including the sandwich core material, cryogenic insulation

or TPS), and the buckling strength of the tank are plotted using the ring frame spacing as the

abscissa in Figure 7. The Gr-BMI/HRP/Gr-BMI concept was found to have the same buckling

response and weight as the Gr-BMI/Ti/Gr-BMI concept and is not shown separately in Figure 7.

The curve in Figure 7 for the Ti/Ti/Ti concept shows a reduction in buckling strength for a ring

frame spacing less than 1.5 m (60 in.) because the buckling mode changes from panel buckling 

to local buckling in the ring frame. The Ti/Ti/Ti tank concept fails in a collapse mode at a low

ring frame spacing, placing a high compression load in the ring frame. The other tank concepts

all buckle between ring frames in a panel buckling mode.

An aft-located LH2 tank, which does not rely on pressure stabilization, must be able to resist a

minimum buckling load of 630 kN/m (3.6 kips/in.) due to the weight of a forward-located LOx 

tank and a payload during launch. A vehicle with a forward-located LH2 tank does not

experience the maximum axial compression load until after landing. The buckling load in this

case is 300 kN/m (1.720 kips/in.) due to bending [13] when the vehicle is horizontal after landing

in an unpressurized state.

Figure 5. Areal masses of TPS and cryogenic insulation along the windward centerline of the

generic vehicle for a LH2 tank.
8

All of the sandwich tank concepts will buckle well above all load requirements for an aft- or

forward-located LH2 tank at all ring frame spacings studied. The Gr-Ep ring and stringer

stiffened tank is designed as a forward-located LH2 tank and requires a 0.75 m (30 in.) ring frame

spacing to resist a 300 kN/m (1.720 kips/in.) buckling load. The Ti/Ti/Ti tank at ring frame

spacings greater than 1.5 m (60 in.) and a Gr-BMI/Ti/Gr-BMI tank at all ring frame spacings is

lighter than a Gr-Ep ring and stiffened structure at a 0.75 m (30 in.) spacing. An aft-located,

non-pressure-stabilized, Gr-Ep ring and stringer stiffened tank would weigh substantially more

than the honeycomb tanks because of the additional structural mass required to resist buckling

and inertial loads due to the weight of a forward-located LOx tank and the payload. Thus, the

results shown in Figure 7 suggest that sandwich structures have a structural and weight

advantage over stiffened Gr-Ep tanks.

Combined TPS, cryogenic insulation, and tank structural areal masses are ranked by their total

weight in Table 2. The studies demonstrate the advantages of sandwich structure in an integrated 

tank system design. Several sandwich tanks with metallic TPS and ring frame spacings of 3.0 m

(120 in.) are lighter than the stiffened forward-located Gr-Ep concepts with ceramic TPS with a

ring frame spacing of 0.75 m (30 in.). Sandwich tanks with larger ring frame spacings are not

considered because of increased pressure pillowing. The lightest tank concept is the GrBMI/HRP/Gr-BMI concept with metallic TPS. An all-metallic Ti/Ti/Ti tank concept is the

second lightest. Both concepts are not only lighter, but also have much higher buckling strengths

than the Gr-Ep stiffened tank and have a weight advantage over Gr-Ep due to their higher

operating temperatures. If the stiffened Gr-Ep forward-located tank is moved to an aft location,

the structural mass of the tank would increase.

The thermal and structural studies indicate that a honeycomb sandwich tank with

mechanically attached metallic TPS is an attractive LH2 tank system for an RLV. However,

more detailed studies are required to corroborate the results from the analytical studies and any

Figure 6. The boundary conditions, model orientation, and model used in BOSOR4 for the

structural buckling study.
9

of the concepts studied will require verification of their insulative capacities, strengths,

reliability, and durability through thermal, and structural testing.

Figure 7. Structural buckling study results for areal masses and critical buckling loads for

sandwich skin tanks and a forward-located.

Table 2. Combined areal masses from the thermal sizing and structural buckling studies with 

variable ring frame spacing.

Tank/

TPS

Ring frame

spacing

(m)

Cryogenic

insulation &

TPS areal mass

(kg/m2)

Tank

structural

areal mass

(kg/m2)

Total

areal mass

(kg/m2)

(Gr-BMI/HRP/Gr-BMI)/

(SA/HC) 3.0 10 6.0 16.0

(Ti/Ti/Ti)/

(SA/HC) 3.0 8.9 8.9 17.8

(Gr-BMI/Ti/Gr-BMI)/

(SA/HC) 3.0 12.8 6.0 18.8

Stiffened Gr-Ep/

AETB 0.75 11.3 10.4 21.7

Stiffened Gr-Ep/

TABI 0.75 11.7 10.4 22.1

TEST PROGRAM

A series of tests and test facilities have been developed during Phase I and Phase II of the

NASA X-33/RLV Program to further evaluate potential reusable cryogenic tank designs for

RLVs. These tests provide information to verify the performance of a concept, to validate

analysis methods, and to demonstrate the scalability of a tank design. The specimens vary in size

from small elements to large subcomponents. The element and panel testing were performed to 

investigate specific aspects of the integrated tank design such as bonding methods, evacuation

0

2

4

6

8

10

12

14

16

0.75 1.50 2.25 3.00 3.75

Distance between ring frames (m)

Areal structural mass (kg/m

2

)

0

200

400

600

800

1000

1200

1400

1600

Nx load (kN/m)

(Ti/Ti/Ti) (Mass)

(Gr-BMI/Ti/Gr-BMI) (Mass)

Stiffened Gr-Ep (Mass)

(Ti/Ti/Ti) (Load)

(Gr-BMI/Ti/Gr-BMI) (Load)

Stiffened Gr-Ep (Load)

Note: Structural areal mass does

 not include TPS or cryogenic 

 insulation areal mass
10

processes, cryogenic insulation integrity, and load carrying capability. Evaluating tank 

performance as an integrated tank system (tank structure, cryogenic insulation, and TPS) at

operational temperatures and load conditions is critical to validate a design. Therefore,

subcomponent testing of full-scale cryogenic tank sections under cyclic thermal and mechanical

loading was performed to investigate additional thermal-structural interactions of the tank design

and to validate performance and fabrication techniques. The cryogenic tank shown in Figure 8,

illustrates the uniaxial tension and compression panel tests, the cryogenic pressure box

subcomponent test, and the mechanical loading each test was designed to simulate. Each of the

tests required the development of new, unique testing facilities and test procedures.

Element Tests

Element tests were developed to investigate specific design features under simplified load 

conditions to provide data that can be incorporated into the design of larger specimens. Two

types of element tests were performed, the flatwise tension test and the sandwich core evacuation

test or ravioli tests. In these tests, element specimens were cycled between temperatures of 20 K

(-423 ̊F) and 395 K (250 ̊F). Flatwise tension tests investigated adhesive strengths after thermal

cycling and sandwich evacuation tests investigated the feasibility of maintaining an evacuated 

core in a sandwich cryogenic tank wall.

Flatwise Tension Test

Adhesives were used to bond honeycomb core to composite facesheets and to bond cryogenic

foam insulation to the tank wall. Flatwise tension tests [14] were used to investigate the effects

of cryogenic and elevated temperatures on the bond line pull-off strengths for sandwich tank 

walls and tank walls with adhesively bonded cryogenic insulation. Examples of flatwise tension

specimens are shown in Figure 9. When an adhesive is subjected to large changes in

temperature, the adhesive may experience a phase transition, becoming brittle at low

temperatures, and could be subject to stress relaxation or creep at elevated temperatures. A

structural system subjected to large changes in temperature, may also develop high stresses

induced by coefficient of thermal expansion mismatch resulting in debonding or core cracking 

without any mechanical load applied.

Figure 8. Panel and subcomponent ground tests (not to scale) and the load conditions the tests

simulated for a cryogenic tank.
11

The adhesive and facesheet/core combinations tested to date are listed in Table 3. Various

types of adhesively bonded or co-cured 5.1 cm by 5.1 cm (2 in. by 2 in.) specimens were cycled

from room temperature to cryogenic or elevated temperatures. Each specimen was then loaded

to failure in tension at room temperature. In Figure 10, the ultimate stress results are displayed 

for HRP honeycomb bonded to Gr-Ep facesheets as an example of the type of information

generated from the flatwise tension tests. The room-temperature specimens are control

specimens which were not thermally cycled, providing a baseline strength for a specimen type.

Full details of the flatwise tension tests results are reported by Glass [14]. The remaining 

specimens were either thermally cycled from room temperature to 20 K (-423 ̊F) 10 times or

from room temperature to 395 K (250 ̊F) 10 times.

Table 3. Adhesives, facesheets, and core materials used in the flatwise tension tests.

Adhesive

Facesheet/Core

EA 9394 PR 1664 Crest

3170 FM 300 HT 435 Co-cure

Gr-Ep/

Ti √ √ √

HRP √ √ √ √ √

RohacellTM WF-71 √ √ √ √

NomexTM √ √ √

Gr-BMI/

Ti √ √ √

HRP √

NomexTM √ √ √

RohacellTM WF-71 √ √ √

Al/

Ti √ √ √ √

Stainless Steel/

RohacellTM WF-71 √ √ √

Figure 9. Flatwise tension specimens: A. Gr-Ep/HRP/Gr-Ep, B. Gr-Ep/RohacellTM/Gr-Ep, C. Ti/Ti/Ti.
12

The three best performing adhesives based on tests performed and reported by Glass [14] and 

as seen in Figure 10, were EA 9394, Crest 3170, and HT 435. The EA 9394 is a roomtemperature-cured adhesive and is widely used as a cryogenic adhesive. The Crest 3170 is also a

room-temperature-cured adhesive and is stronger than EA 9394 after thermal cycling. The HT

435 is a high-temperature-cured adhesive (450 K [350 ̊F]) and had the best overall performance

after being thermally cycled, however, this adhesive was the most sensitive to preparation

procedures such as surface preparation, adequate pressure on the specimen during cure, and the

correct cure temperature. The FM 300, a high-temperature-cured film adhesive (450 K [350 ̊F])

had the lowest strength and the PR 1664, a room-temperature-cured adhesive, also had a

relatively low strength.

Sandwich Core Evacuation (Ravioli) Test

The core of a cryogenic sandwich tank may not only perform its structural function, but if

evacuated, may also act as cryogenic insulation. Evacuating the core enhances its insulation

capacity by reducing the thermal effects of natural convection and gas conduction and may

eliminate the need for additional cryogenic insulation. It is essential to maintain a vacuum in the

core to prevent cryopumping and potential failure of the sandwich.

A test was developed to evaluate the ability of various concepts for a sandwich tank wall

system to resist gas permeation [14]. A series of evacuated core honeycomb specimens, shown

in Figure 11, were fabricated and tested. The specimens were referred to as ravioli specimens

due to their shape. A square of cryogenic foam insulation or perforated honeycomb core

Figure 10. Flatwise tension strengths (tested at room temperature) of HRP honeycomb core

bonded to Gr-Ep with five different adhesives and having three different pre-conditionings. Each bar represents a different specimen tested to failure.
13

material with 45° beveled edges was sealed inside of a 15 cm by 15 cm (6 in. by 6 in.) shell of

Gr-Ep or Gr-BMI. The specimen was sealed by either bonding the two pre-cured halves with an 

adhesive or by co-curing the shell material around the core. The core material in these

specimens is perforated or drilled to allow for evacuation due to air liquefaction/solidification 

from cryogenic temperatures or by mechanical evacuation. An evacuation stem (location shown

in Figure 11) was used to actively evacuate the specimen and allow any trapped gasses to escape

as the specimen was warmed.

All six ravioli specimens tested are listed in Table 4. Specimen 1 had a RohacellTM core, but

did not have an evacuation stem. This specimen was cycled from room temperature to 80 K (-

320 ̊F) 10 times by immersing the specimen in a container filled with LN2, removing the

specimen, then allowing the specimen to warm to room temperature. There were no visible signs

of damage by the naked eye. The specimen was also immersed once in liquid helium (LHe) and 

then removed. The specimen ruptured as it warmed to room temperature. It is believed that the

specimen permeated LHe into the core region and burst because the LHe vaporized faster than

the gaseous helium (GHe) could out-gas. Each specimen listed in Table 4, except for specimen

1, had an evacuation stem attached on the upper portion of the specimen to actively evacuate the

specimen and to allow the specimen to out-gas if LHe permeated to the core region. The

specimens were thermally cycled 10 times to LN2 temperatures (80 K [-320 ̊F]), then immersed 2 

times in LHe (4K [-450 ̊F]) with no visible signs of degradation. The integrity of each specimen

was then investigated after thermal cycling by evacuating the specimen as the temperature of the

specimen was lowered from ambient to 80 K (-320 ̊F). At various temperatures, GHe was

sprayed on the edges of the specimen while a helium (He) mass spectrometer leak detector was

used to actively evacuate the specimen and detect the amount of leakage of GHe into the

specimen. The data from two leak detection tests for specimen 6 are shown in Figure 12. The

plot shows that as the specimen was cooled, the ability of the specimen to maintain a low

pressure diminished. The leaks in specimens 2 through 5 were too large for the He mass

spectrometer leak detector to evacuate the specimen.

a. b.

Figure 11. a.) The components of a Gr-Ep ravioli specimen with HRP honeycomb core and the

b.) top view of assembled specimen (evacuation stem removed).
14

Table 4. Gr-Ep ravioli specimens.

Specimen no. Core Material Adhesive Result

1* RohacellTM WF-110 FM 300 Rupture

2 RohacellTM WF-51 FM 300 Large leaks

3 RohacellTM WF-71 PR 1664 Large leaks

4 NomexTM honeycomb Crest 3170 Large leaks

5 RohacellTM WF-110 Co-Cured Large leaks

6 HRP honeycomb EA 9394 Moderate vacuum

held

Note:

* No evacuation port

The results from the ravioli tests demonstrated that maintaining vacuum in even a small

specimen is difficult, and as the specimen was cooled, its resistance to GHe permeation was

reduced. The helium mass spectrometer leak detector could not localize the origin of GHe

permeation. The GHe permeation into the specimens may have resulted from a coefficient of

thermal expansion mismatch between the adhesive and the two halves of the specimen causing 

microcracks in the bond line. A crease located on the top of the shell at the corners, as indicated 

in Figure 11, may have also been a source of leakage. The co-cured specimen without an 

adhesive layer at the bond line performed only slightly better than the adhesively bonded 

specimens. These results suggest that improvements are needed in the fabrication of leak-free

sandwich structures, and that a vented or an actively evacuated system may be required in an 

evacuated core sandwich structure.

Panel Tests

Panel tests were used to address integration issues between the tank wall and cryogenic

insulation subjected to representative operational thermal and mechanical loads. The panel test

specimens were larger than the element specimens and incorporated “lessons learned” from the

element tests. Two types of panel tests are described: a cyclic uniaxial tension test to simulate

hoop pressurization, and a static compression test to simulate structural and inertial loads, as

0

2

4

6

8

10

75

100125150175200225250275300

Temperature (K)

Test 2

Test 1

Figure 12. Pressure versus temperature for the 6th ravioli specimen.

Pressure

(psi)
15

depicted in Figure 8. Cryogenic insulation was integrated into the specimen for the panel tests to

reduce the complexity of testing. Testing with TPS attached to the specimen would require

additional heating equipment and would have resulted in longer, more complicated testing 

cycles, and more expensive, elaborate test specimens.

Uniaxial Tension Tests

Combined cyclic thermal and mechanical tests of various cryogenic tank wall concepts were

performed on flat 30.5 cm by 61 cm (1 ft by 2 ft) panel specimens. A flat specimen closely

approximates a curved tank wall due to the large radius of the tank. These tests were developed

from earlier tests of a cryogenic insulation tile developed for the Advanced Launch System

(ALS) Program [15]. The purpose of the tests was to simulate both the thermal and mechanical

loads experienced in an RLV mission from launch, to orbit, to re-entry. The cryogenic tanks in 

an RLV must endure biaxial tension loads associated with internal pressurization as well as

maximum thermal and mechanical flight loads. However, for these tests, the only mechanical

load applied was a uniaxial tension load simulating circumferential pressure loading. These

combined, cyclic, thermal and mechanical tests verify: the durability of the cryogenic insulation 

when subjected to cyclic mission-profile conditions, the bond line integrity between cryogenic

insulation and the structure, the performance of cryogenic tank fabrication technologies at a

small scale, and the effectiveness of various IVHM techniques such as fiber optic (strain,

temperature, and H2 sensing) and acoustic sensors.

Specialized test fixtures were developed to operate between a minimum temperature of 20 K

(-423 ̊F) and a maximum temperature of 645 K (700 ̊F). Figure 13 shows a typical specimen

mounted in the fixture with the cryogenic chambers mounted on the surface of the inner tank

wall of the specimen and a convective heating chamber adjacent to the external surface of the

foam insulation. Tension load and temperatures for the cryogenic and high temperature

chambers were independently controlled in a test cycle. A typical cycle lasted 30 to 80 minutes.

An example of a thermal and mechanical load profile for a LH2 tank specimen is shown in Figure

14, which displays the tension loading and temperature profiles on the hot side and cryogenic

side of the panel over a period of time.

All of the panels tested as a part of the NASA X-33 Program Phase I and II are listed in Tables 5

and 6. Fiber-optic or IVHM thermal sensors were used on some of the panels tested (as

indicated in Table 5). A meter of fiber-optic cable was coiled at several locations to obtain a

point-wise thermal reading [5]. A single fiber was used to monitor several locations on a

specimen. The adhesive methods and the ability of the fiber optic thermal sensors to operate

during mission profile conditions were tested in the uniaxial tension tests. The tests of panels

LO-3 and LO-4 support qualification of the SS-1171 spray-on-foam insulation (SOFI) for the Al

LOx tank for the X-33.

Compression Test

The compressive load capability of a tank wall concept under simulated structural and inertial

loads was tested using representative flat specimens in the cryogenic/high temperature

compression test fixture shown in Figure 15. This fixture and the flat specimens was also used to

induce a through-the-thickness temperature gradient in a compression specimen during a test.
16

Table 5. Experimental results for the LOx tank concepts subjected to combined thermal and

mechanical tension load.

LOx 

tank

panel

Panel description

No. of

cycles

Results

†

LO-1 Al-Li* panel, EA 9394 adhesive,

external AirexTM foam insulation

23 No cracks in foam.

No disbonds.

LO-2 Al panel, EA 9394 adhesive,

external AirexTM foam insulation

33 No cracks in foam.

No disbonds.

LO-3 Al panel, external SS-1171, and

PDL-1034 foaminsulation

50

PDL-1034 cracked after 16 cycles.

Insulation thickness reduced due to 

surface charring (no degradation in 

performance noticed).

LO-4 Al-Li* panel, external SS-1171

foam insulation, fiber-optics

50

No cracks in foam.

Insulation thickness reduced due to 

surface charring (no degradation in 

performance noticed).

Note:

* Al-Li - Aluminum Lithium 2195

† Cracks were determined by visual inspection without die or microscope.

In the compression test, a 61 cm by 61 cm (2 ft by 2 ft) specimen with cryogenic insulation

was subjected to a cryogenic temperature load before a compressive load was introduced. Two

temperature conditions were attempted, a uniform temperature, and a constant temperature

gradient through-the-thickness of the specimen. Three identical sandwich panels were tested to 

failure. Each panel experienced one of three temperature load conditions: cryogenic temperature

gradient (with a minimum temperature of 20 K [-423 ̊F] and maximum temperature of 115K [-

Figure 13. Schematic and assembled view of the test fixture for the uniaxial tension test.
17

250 ̊F]), room temperature, and a uniform elevated temperature (maximum temperature of 480 K

[400 ̊F]).

Table 6. Experimental results for LH2 tank concepts subjected to combined thermal and 

mechanical tension load.

LH2

tank

panel

Panel description

No. of

cycles

Results

LH-1 Gr-Ep panel, EA 9394 adhesive,

external AirexTM foam insulation

42 AirexTM foam cracked after 42 cycles

(due to a void in the adhesive layer).

LH-2 K3B/Ti/K3B† co-cured sandwich 

panel 12 Panel failed in the load introductionregion after 12 cycles.

LH-3 K3B/Ti/K3B† co-cured sandwich 

panel with a joint

0

Panel failed in the joint region at 80%

of the design limit load during a pre-test

load check.

LH-4 Gr-Ep/RohacellTM WF-71/Gr-Ep 

foam sandwich panel 25

Facesheet separated from the foam core

after 25 cycles (due to expansion of the

foam with heat).

LH-5

Gr-Ep panel, EA 9394 adhesive,

external AirexTM foam insulation,

fiber-optics

50 No cracks.

No disbonds.

LH-6 Gr-Ep panel, Crest 3170

adhesive, external AirexTM foam

insulation, fiber-optics

50 No cracks.

No disbonds.

Note: 

† K3B - IM7/K3B.

The low-carbon, stainless steel (304 steel) compression load introduction fixtures shown in

Figure 15 were developed to introduce a uniform axial load across the top and bottom edge of

the panel without requiring potting of the ends of thee specimen at cryogenic and elevated

temperatures. These metallic fixtures were designed to control end displacement and to reduce

(0.0)

0.2

0.4

0.6

0.8

1.0

0

100

200

300

400

500

0

200400600800

10001200140016001800200022002400260028003000

Time (Sec.)

Normalized tension load

Cryogenic side of panel

Heater side of panel

Figure 14. Typical combined thermal and mechanical cycle for a LH2 tank specimen in the

uniaxial tension test (load normalized to maximum load).

Tem

perature

(K)

L

oad

(nK

g/nK

g)
18

bending effects from through-the-thickness temperature gradients, thus providing controlled 

boundary conditions at various temperatures. Conventional potting materials soften at

temperatures above 450 K (350 ̊F) so their use was avoided. The metallic load introduction

fixtures were designed to provide consistent and reproducible method of load transfer.

Cryogenic and high temperature platens were developed to heat or cool the compression fixture

to match the temperature of the specimen. Ceramic insulation tiles were be used to reduce heatloss through the platens and thermally isolate the compression fixture and platens from the load

stand. Knife-edged supports (not shown Figure 15) composed of 304 steel were also developed

to impose simple-support boundary conditions on the vertical edges of the test specimen. The

knife-edged supports had a temperature range from 80 K (-320 ̊F) to 480 K (400 ̊F). A Crest

3170 bellows seal was developed to contain the cryogenic fluid on the cold side of the panel and 

resistive heater blankets were be used to heat the faces of the panel.

Only the room-temperature panel test has been completed to date. The room-temperature

honeycomb panel test consisted of co-cured IM7/K3B facesheets with a Ti honeycomb core GrK3B/Ti/Gr-K3B). The K3B resin is a thermoplastic material that has a maximum operational

temperature of 480 K (400 ̊F) and has non-autoclave joint fabrication potential. The upper and

lower 0.06 m (2.5 in.) of the core were filled with a foaming adhesive during the curing process

to facilitate load introduction from the fixture to the specimen and to prevent specimen

“brooming” at the ends. However, the adhesive over-expanded in the core causing a

discontinuity where the foam-filled region ended.

Finite element analysis at LaRC predicted that the room-temperature panel would buckle at a

compressive load of 3,110 kN (699 kips) or 5,080 kN/m (29 kips/in.). Facesheet wrinkling was

determined by the analysis to occur at an axial load of 707 kN (159 kips) or 1,160 kN/m (6.63 

Figure 15. Schematic of the cryogenic/room/elevated temperature compression fixture.
19

kips/in.). The panel failed at a compressive load of only 540 kN (121 kips) or 880 kN/m (5.04 

kips/in.) in the upper section of the panel adjacent to the aforementioned skin discontinuity.

Failure in this region suggested that either the load was not properly introduced into the panel

through the foam-filled honeycomb core region or that the panel was poorly fabricated near the

edge of the foam-filled honeycomb core.

Subcomponent Test

A new test facility, the Cryogenic Pressure Box (CPB) Test Facility, has been developed to

validate full-scale tank subcomponents under realistic static and thermal/mechanical/pressure

cyclical loading [16]. Curved tank panel concepts can be tested in this facility at a relatively low

cost compared to a full-scale or scaled tank tests at cryogenic temperatures. Finite element

analysis performed at LaRC predicts that the stain distributions developed during testing in the

CPB will be similar to the strains seen in a full-sized cylindrical tank can be produced in a 0.75

m by 1.0 m (2.5 ft by 3.5 ft) region in the center of the specimen. The effects of cycling with

mechanical loads, pressure loads, and thermal loads on full-scale assemblies of tank walls with

cryogenic insulation, TPS, and IVHM can be determined with a full-scale subcomponent.

Manufacturing and fabrication details can then be refined before fabrication of a full-scale tank,

thus reducing the risk of premature failure in a tank due to cyclic thermal/mechanical loading.

A schematic section view of the CPB Test Facility is shown in Figure 16. A view of the

chamber is shown in Figure 17. In this subcomponent test, a 1.5 m by 1.8 m (5 ft by 6 ft) curved

(radii of 2.0 m to 6.5 m, 80 in. to 260 in.) panel in Figure 18 is loaded in biaxial tension by

internal pressure and mechanical actuators. In addition, both cryogenic and elevated internal

temperatures and an elevated external temperature can be applied. The biaxial tension load is

introduced into the panel by internal pressure reacted through the load frame and by axial

actuators. The maximum load applied by the axial actuators (not shown in Figure 16) is 2,000

kN (450 kips). Circumferential, or hoop, loads due to pressurization are induced by the reaction

force from the load frame, through load introduction plates, into the test specimen.

Cryogenic Pressure Box Test Facility

Figure 16. Schematic of the CPB Test Facility for subcomponent tests.
20

The vacuum jacketed pressure chamber can withstand internal pressures ranging from

atmospheric to 372 kPa (54 psig) using GHe as the pressurization medium. The test panel and

chamber can also be subjected to internal temperatures ranging from 395 K (250 ̊F) to 20 K (-

423 ̊F) with the aid of twelve copper heat exchange towers that are encircled by copper coils that

contain either LHe, GHe or LN2. The GHe is recirculated by fans through the heat exchange

towers. Helium does not liquefy at temperatures above 15 K (-430 ̊F) and can be used to

convectively cool the specimen and the CPB chamber. The heat exchange towers also have

resistive heaters at their bases, enabling internal heating of the chamber and the internal surface

of the panel to a maximum temperature of 395 K (250 ̊F). A quartz-lamp heater array is used to 

heat the external surface of the specimen to a maximum temperature of 810 K (1,000 ̊F). The

heater array is flat and has eight symmetric zones that can be individually controlled to evenly 

heat specimens of various curvatures.

Cryogenic Pressure Box Chamber

Figure 17. Schematic of the chamber of the CPB Test Facility.

Figure 18. Thermal and mechanical loads applied to a representative specimen in the CPB Test

Facility.
21

SUMMARY

A systematic approach was used in the research for the design of integrated cryogenic tank 

systems for a reusable launch vehicle (RLV). This approach began with thermal and structural

analytical studies followed by testing of specimens ranging from elements to subcomponents at

the NASA, Langley Research Center (LaRC).

The results of the analytical studies identified honeycomb sandwich tank with mechanically

attached metallic thermal protection system (TPS) as a preferred approach for a reusable liquid

hydrogen (LH2) tank system for an RLV. The two most viable honeycomb sandwich concepts

were found to be Gr-BMI/HRP/Gr-BMI and Ti/Ti/Ti, based on structural and thermal efficiency.

Element tests were used to evaluate bonding and fabrication methods as well as the

evacuation process for sandwich tank structures. Adhesives such as Crest 3170 (now Lord 212)

and HT 435 were identified as the most useful for cryogenic tank construction by the flatwise

tension tests. The evacuated honeycomb sandwich (ravioli) tests demonstrated that sealed

sandwich concepts may be problematic, that evacuation of a sandwich is difficult, and that active

evacuation may be a solution to obtain a reliable sandwich tank concept.

The panel and subcomponent tests were developed to investigate structural strength and

durability, the reliability of the fabrication process scale-up, thermal properties, and bond line

integrity of cryogenic tank designs. The uniaxial tension tests provided data for the NASA X-33

Program in support of certifying SS-1171 for the liquid oxygen (LOx) tank and a new cryogenic

foam insulation, AirexTM, for the LH2 tanks. The compression fixture will enable the testing of

specimens at various temperatures or with through-the-thickness temperature gradients.

The unique analytical tools and facilities developed at the LaRC during Phase I and Phase II

of the NASA X-33/RLV Program enable the study and testing of various cryogenic tank

concepts at operational thermal loads, mechanical loads, and pressure loads. The results

obtained from these analytical and experimental cryogenic tank studies will provide vital

information required to develop full-scale, reusable, and integrated cryogenic tanks for future

RLVs.

Acknowledgments

The research conducted at LaRC could not have been accomplished without the efforts of the

members of the cryogenic tank team in the Metals and Thermal Structures Branch (MTSB). The

cryogenic tank team was supported by systems specialists Mr. Joseph Sikora and Mr. Vincent

LeBoffe, designer Mr. Kermit Jensen, structural analysts Mr. Carl Martin and Mr. Jeff Cerro and

all of the technical staff in the MTSB laboratory but especially electrician Mr. Paul McClung for

developing the control algorithms for the uniaxial tension tests. Optimization analyses was

performed by Dr. Satchi Venkataraman and Dr. Raphael Haftka at the University of Florida. The

Cryogenic Pressure Box Facility development was conducted with Dr. Damodar Ambur, Dr.

David Glass, Mr. Marshall Rouse, Mr. Henry Wright, Mr. James Mayhew, and Mr. Carlos Perez.

Many of the original concepts, analytical studies, and tests of the cryogenic tank team were

initiated by Dr. Charles Camarda.
22

References

1. Freeman, Jr., D. C., Stanley, D. O., Camarda, C. J., Lepsch, R. A., Cook, S. A., “SingleStage-To-Orbit-A Step Closer,” Presented at the 45th Congress of the International

Astronautical Federation (IAF), October 1994, IAF 94-V3.534.

2. NASA, “Access to Space Study Final Report,” NASA Headquarters, Washington, D. C., July 

1993.

3. Baumgartner, R. I., “VenturestarTM Single Stage to Orbit Reusable Launch Vehicle Program

Review,” Presented at the Space Technologies and Application International Forum,

Albuquerque, NM, January 1997, Vol. 3: 1033-1039.

4. Cook, S., “The X-33 Advanced Technology Demonstrator,” Presented at the American 

Institute of Aeronautics and Astronautics (AIAA) Dynamics Specialists Conference, Salt

Lake City, UT, April 1996, AIAA-96-1195.

5. Melvin, L., Childers, B., Rogowski, R., Prosser, W., Moore, J., Froggatt, M., Allison S., Wu,

M. C., Bly, J., Aude, C., Bouvier, C., Zisk, E., Enright, E., Cassadaban, Z., Reightler, R.,

Sirkis, J., Tang, I., Peng, T., Wegreich, R., Garbos, R., Mouyos, W., Aibel, D., Bodan, P.,

“Integrated Vehicle Health Monitoring (IVHM) for Aerospace Vehicles,” Presented at the

International Workshop on Structural Health Monitoring, Structural Health Monitoring:

Current Status and Perspectives, Edited by Fu-Kuo Chang, Stanford, CA, September 1997,

pp. 705-714.

6. Anonymous, “Prototype Tank Drawings and Analysis,” Reusable Composite Hydrogen Tank

System TA1 Payment Milestone 14, Part 2 of 5, Cooperative Agreement NCC8-39,

Rockwell International, September 1995, SSD95D0388.

7. Myers, D. E. Martin, C. J., Blosser, M. L., “Parametric Weight Comparison of Advanced 

Metallic, Ceramic Tile, and Ceramic Blanket Thermal Protection Systems (TPS),” NASA

Technical Memorandum 2000-210289, June 2000.

8. Williams, S. D., Curry, D. M., “Thermal Protection Materials: Thermal Property Data,”

NASA Reference Publication 1289, December 1992.

9. Robinson, M. J., "Composite Cryogenic Propellant Tank Development," Proceedings of the

35th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials

Conference, Hilton Head, SC, April 1994.

10. Bushnell, D., “BOSOR4: Program for Stress and Vibration of Shells of Revolution,”

Structural Mechanics Software Series, Edited by N. Perrone and W. Pilkey, University Press

of Virginia, Charlottesville, VA, Vol. 1: 11-143.

11. Bushnell, D., "Stress, Stability and Vibration of Complex Shells of Revolution: Analysis and 

User's Manual for BOSOR4," NASA CR-2115, Oct. 1972.

12. Johnson, T. F., and Card, M. F., “Effects of Stiffening and Mechanical Load on Thermal

Buckling of Stiffened Cylindrical Shells,” AIAA Paper 95-1317, 36th AIAA/ ASME/ ASCE/

AHS/ ASL Structures, Structural Dynamics and Materials Conference, New Orleans, LA,

April 1995.

13. Anonymous, “Test Report of Element/Subcomponent Testing,” Reusable Composite

Hydrogen Tank System TA 1 Payment Milestone 16, Part 1 of 2, Cooperative Agreement

NCC8-39, Rockwell International, November 1995, SSD95D0507.

14. Glass, D. E, “Bonding and Sealing Evaluations for Cryogenic Tank,” NASA Contractor

Report 201734, August 1997.
23

15. McAuliffe, P. S., Davis, R. C., Taylor, A. H., “Development of Reusable, Flight-Weight

Cryogenic Foam Insulation System,” Presented at the American Institute of Aeronautics and 

Astronautics (AIAA) Space Systems Technology Conference, June 1986, AIAA-1189-CP.

16. Ambur, D. R., Sikora, J., Maguire, J. F. Winn, P. M., “Development of Pressure Box to 

Evaluate Reusable Launch-Vehicle Cryogenic Tank Panels,” Presented at the American 

Institute of Aeronautics and Astronautics (AIAA) 37th AIAA/ASME/ASCE/AHS/ASC

Structures, Structural Dynamics, and Materials Conference, Salt Lake City, Utah, April 1996,

AIAA-96-1640.

Nomenclature

304 steel: Stainless Steel

AETB: Alumina Enhanced Thermal

Barrier

Al: Aluminum 2219-T87

Al-Li: Aluminum 2195

ALS: Advanced Launch System

ELI: Extra Low Interstitial

Gr-BMI: IM7/5260 GraphiteBismaleimide

Gr-Ep: IM7/977-2 Graphite-Epoxy

GHe: Gaseous Helium

H2: Hydrogen

He: Helium

HRP: HexcelTM glass Reinforced 

Phenolic

IVHM: Integrated Vehicle Health

Monitoring 

K3B: IM7/K3B

LaRC: NASA, Langley Research 

Center

LH2: Liquid Hydrogen

LHe: Liquid Helium

LOx: Liquid Oxygen 

NASA: National Aeronautics

and 

Space Administration

NASP: National Aerospace Plane

Nx: Axial Load

RLV: Reusable Launch Vehicle

SA/HC: Superalloy/Honeycomb

SOFI: SS-1171 Spray On Foam

Insulation

SSTO: Single-Stage-To-Orbit

TABI: Tailorable Advanced Blanket

Insulation

Ti: Titanium

TPS: Thermal Protection System

MTSB: Metals and Thermal

Structures Branch

VAB: Vehicle Analysis Branch
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

2. REPORT TYPE 

Technical Memorandum

 4. TITLE AND SUBTITLE

Thermal Structures Technology Development for Reusable Launch Vehicle 

Cryogenic Propellant Tanks

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Johnson, Theodore F.; Natividad, Roderick; Rivers, H. Kevin; and 

Smith, Russell W.

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

NASA Langley Research Center

Hampton, VA 23681-2199

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

National Aeronautics and Space Administration

Washington, DC 20546-0001

 8. PERFORMING ORGANIZATION

 REPORT NUMBER

L-19165

10. SPONSOR/MONITOR'S ACRONYM(S)

NASA

13. SUPPLEMENTARY NOTES

An electronic version can be found at http://ntrs.nasa.gov

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category 15

Availability: NASA CASI (301) 621-0390

19a. NAME OF RESPONSIBLE PERSON

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

14. ABSTRACT

Analytical and experimental studies conducted at the NASA, Langley Research Center (LaRC) for investigating integrated cryogenic 

propellant tank systems for a reusable launch vehicle (RLV) are described. The cryogenic tanks are investigated as an integrated tank 

system. An integrated tank system includes the tank wall, cryogenic insulation, thermal protection system (TPS) attachment sub-structure, 

and TPS. Analysis codes are used to size the thicknesses of cryogenic insulation and TPS insulation for thermal loads, and to predict tank 

buckling strengths at various ring frame spacings. The unique test facilities developed for the testing of cryogenic tank components are 

described. Testing at cryogenic and high-temperatures verifies the integrity of materials, design concepts, manufacturing processes, and 

thermal/structural analyses. Test specimens ranging from the element level to the subcomponent level are subjected to projected vehicle 

operational mechanical loads and temperatures. The analytical and experimental studies described in this paper provide a portion of the basic 

information required for the development of light-weight reusable cryogenic propellant tanks. 

15. SUBJECT TERMS

Cryogenic testing; Cryogenics; Sizing studies

18. NUMBER

 OF 

 PAGES

28

19b. TELEPHONE NUMBER (Include area code)

(301) 621-0390

a. REPORT

U

c. THIS PAGE

U

b. ABSTRACT

U

17. LIMITATION OF 

 ABSTRACT

UU

Prescribed by ANSI Std. Z39.18

Standard Form 298 (Rev. 8-98)

3. DATES COVERED (From - To)

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

23-064-20-32

11. SPONSOR/MONITOR'S REPORT

 NUMBER(S)

NASA/TM-2005-213913

16. SECURITY CLASSIFICATION OF:

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- 09 - 2005