Patrick H. Dunlap, Jr. and Bruce M. Steinetz

Glenn Research Center, Cleveland, Ohio

Donald M. Curry

Johnson Space Center, Houston, Texas

Jeffrey J. DeMange

Ohio Aerospace Institute, Brook Park, Ohio

H. Kevin Rivers

Langley Research Center, Hampton, Virginia

Su-Yuen Hsu

Lockheed Martin Space Operations, Hampton, Virginia

Investigations of Control Surface Seals

for Re-Entry Vehicles

NASA/TM—2002-211708

July 2002

AIAA–2002–3941
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Patrick H. Dunlap, Jr. and Bruce M. Steinetz

Glenn Research Center, Cleveland, Ohio

Donald M. Curry

Johnson Space Center, Houston, Texas

Jeffrey J. DeMange

Ohio Aerospace Institute, Brook Park, Ohio

H. Kevin Rivers

Langley Research Center, Hampton, Virginia

Su-Yuen Hsu

Lockheed Martin Space Operations, Hampton, Virginia

Investigations of Control Surface Seals

for Re-Entry Vehicles

NASA/TM—2002-211708

July 2002

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

38th Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Indianapolis, Indiana, July 7–10, 2002

AIAA–2002–3941
Acknowledgments

The authors gratefully acknowledge Jeff Hagen and Prakash Chaudha (Lockheed-Martin) for their contributions

in designing the scrub test fixture; Larry Zielke (NASA Johnson Space Center) for his assistance in conducting the

scrub tests; Ron Lewis and Winston Goodrich (NASA Johnson Space Center), and Ignacio Norman (Boeing)

for their work on the heat transfer analyses; Tom Doeberling (NASA Glenn Research Center) and

Dick Tashjian (QSS) for their assistance in test support.

Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Trade names or manufacturers’ names are used in this report for

identification only. This usage does not constitute an official

endorsement, either expressed or implied, by the National

Aeronautics and Space Administration.

Available electronically at http://gltrs.grc.nasa.gov/GLTRS
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 1

INVESTIGATIONS OF CONTROL SURFACE SEALS FOR RE-ENTRY VEHICLES

Patrick H. Dunlap, Jr.*

Bruce M. Steinetz†

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio

Donald M. Curry‡

National Aeronautics and Space Administration

Johnson Space Center

Houston, Texas

Jeffrey J. DeMange§

Ohio Aerospace Institute

Brook Park, Ohio

H. Kevin Rivers

National Aeronautics and Space Administration

Langley Research Center

Hampton, Virginia

Su-Yuen Hsu

Lockheed Martin Space Operations

Hampton, Virginia

as a case study to evaluate a currently available control

surface seal design for applications in future re-entry

vehicles. A thermal analysis of the rudder/fin seal assembly

based on representative heating rates predicted a peak seal

temperature of 1900 °F. Seals were heated in a compressed

state at this peak temperature to evaluate the effects of

temperature exposure. Room temperature compression

tests were performed to determine load versus linear

compression, preload, contact area, stiffness, and resiliency

characteristics for as-received and temperature-exposed

seals. For all compression levels that were tested, unit

loads and contact pressures for the seals were below the

5 lb/in. and 10 psi limits required to limit the loads on the

Shuttle tiles that form the adjacent sealing surface for

these seals. The seals survived an ambient temperature

1000 cycle scrub test over sanded Shuttle tile surfaces and

were able to disengage and re-engage the edges of the

rub surface tiles during scrubbing. Arc jet tests were

performed to experimentally determine anticipated seal

temperatures for representative flow boundary conditions

(pressures and temperatures) under simulated vehicle

re-entry conditions. Installation of a single seal in the

gap of the test fixture caused a large temperature drop

(∆T = 1710 °F) across the seal location confirming the

need for seals in the rudder/fin gap location. The seal acted

Abstract

Re-entry vehicles generally require some combination of

control surfaces (e.g., rudders, body flaps, elevons) to

steer or guide the vehicles during re-entry into and through

the Earth’s atmosphere. Control surface seals are installed

between these movable surfaces and stationary portions of

the vehicle both along hinge lines and in areas where

control surface edges are actuated close to the vehicle

body. These seals must operate in high-temperature

environments and limit hot gas ingestion and transfer of

heat to underlying low-temperature structures to prevent

over-temperature of these structures and possible loss of

vehicle structural integrity.

This paper presents results for thermal analyses and

mechanical testing conducted on the baseline seal design

for the X–38 rudder/fin. This seal application was chosen

* Mechanical Engineer, Mechanical Components Branch,

Member AIAA.

† Senior Research Engineer, Mechanical Components Branch,

Member AIAA.

‡ Aerospace Technologist, Thermal Branch.

§ Senior Research Associate, OAI, Mechanical Components Branch,

Member AIAA.
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 2

as an effective thermal barrier limiting heat convection

through the seal gap and minimizing temperature increases

Introduction

Since the Space Shuttle went into regular service in 1981,

NASA has been examining vehicle concepts both to

complement and replace this launch system. Vehicles

capable of putting payloads and personnel into orbit with

faster turnaround times, higher safety margins, and at

lower costs are especially in demand now that the

International Space Station (ISS) is in orbit. New vehicle

concepts are being pursued for crew transfer vehicles

(CTV’s, such as X–37) and crew return vehicles (CRV’s,

such as X–38) that would be used to transport personnel

while in orbit or to leave the Space Station during a

medical emergency or an evacuation.

Although each vehicle concept is unique, they all have

some combination of control surfaces that steer or guide

the vehicles during re-entry into and through the Earth’s

atmosphere. These control surfaces include rudders, body

flaps, elevons, and other surfaces that move with respect

to the body of the vehicle. Seal interfaces exist between

these movable surfaces and stationary portions of the

vehicle both along hinge lines and where control surface

edges seal against the vehicle body. These seals must

operate in high-temperature environments and limit hot

gas ingestion and transfer of heat to underlying lowtemperature structures to prevent over-temperature of

these structures and possible loss of the vehicle. Hinge line

seals are especially important because they ensure that the

actuators that move the control surfaces remain cool

enough to operate and control the flight of the vehicle.

One recent vehicle concept is the X–38 vehicle that NASA

designed to demonstrate technologies that would be

required for a potential CRV for the International Space

Station (fig. 1). The vehicle is designed to glide from orbit

in an unpowered freefall that is controlled by two movable

rudders, two body flaps located at the aft end of the

vehicle, and a steerable parafoil deployed after re-entry.

Seal interfaces exist between the movable body flaps and

the bottom surface of the vehicle and between the rudders

and their respective fins (figs. 1(a) and (b)). Wong et al.1

performed a series of two-dimensional computational fluid

dynamics studies that modeled the gap between the rudder

and fin during re-entry of the X–38 vehicle and concluded

that a seal is required along this interface to prevent

excessive local heat fluxes on these structures.

The objectives of the current study were to summarize the

tests and analyses that have been performed on the baseline

seal design for the X–38 rudder/fin seal and assess the

seal's capabilities relative to future reusable launch vehicle

requirements. The specific goals of this effort were to:

1. Measure seal flow rates, resiliency, and unit loads

both in an as-received condition and after temperature

exposure.

2. Examine seal durability and wear resistance in

order to recommend rub-surface treatments required to

maximize seal wear life.

3. Determine the effects of seal damage incurred

during scrubbing on flow rates through the seals.

4. Experimentally determine anticipated seal

temperatures for representative external flow boundary

conditions under arc jet test conditions simulating vehicle

re-entry.

Case Study: Design Requirements for

X–38 Rudder/Fin Seal System

The design of the X–38 rudder/fin seal assembly consists

of a double seal attached to the rudder that seals the

vertical hinge line and the fin shelf line (figs. 1 and 2). The

vertical seal loop surrounds and protects the rudder drive

motor and attachments between the rudder and the fin

(fig. 2). The seal assembly must allow the rudder to

rotate during the entire mission and must accommodate a

rotation range of ±12° (fig. 2).

Temperature Limits/Temperature Drop

The rudder/fin seal assembly will be expected to endure

high temperatures caused by convective heating in an

oxidative environment with radiation exchange in the seal

gap. A thermal analysis predicted that peak temperatures

for the exposed seal could reach approximately 1900 °F

(with laminar boundary layer assumption) to 2100 °F

(with turbulent boundary layer assumption) with seal

attachment temperatures of 1500 °F (fig. 3). Peak

temperatures occur about 1200 sec (20 min.) into re-entry

with a subsequent decrease in temperatures for the

remainder of the re-entry. Materials used in the seals must

be able to withstand these high temperatures. Because the

predicted attachment temperature exceeds current adhesive

temperature limits, the seals must be mechanically attached

to the seal carrier plate and rudder. In addition to

withstanding these extreme temperatures, the seals must

act as a thermal barrier to minimize temperature increases

downstream of the seal in order to protect underlying lowdownstream of the seal to acceptable (< 200 °F) levels.
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 3

Rudder structure

Rudder/

fin

Vertical seal

Shelf seal

Mechanical seal attachment

Shelf/sealing surface:

AETB–8 tile on aluminum

structure

Rudder drive motor

Fin structure

Forward

Body flaps

Figure 1.—(a) X–38 vehicle. (b) Rudder/fin structure and seal locations.

(a) (b)

Fin structure

Outboard vertical seals

Inboard vertical seals

Top rub surface

Lower rub surface

Outboard

shelf seal Inboard shelf seal

12° rotation

21.98 in.

7.5 in.

Top seals

Bottom seals

Inner

Outer

Figure 2.—Computer model depicting rudder/fin seal rotated to full outboard position with

 seal dimensions.

Drive motor location

(removed for clarity)

36.35 in.
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 4

(a)

q

.

Seals

Locations

Titanium

attachment

1

6

5

4

3

2

AETB–8 tile AETB–8 tile

Figure 3.—Thermal analysis of rudder/fin seal. (a) Rudder/fin gap area Thermal Math Model (TMM).

 (b) Rudder/fin seal temperature and pressure predictions.

(b)

2500

2250

2000

1750

1500

1250

1000

750

500

250

0

Temperature, °F

0 200 400 600 800 1000 1200 1400 1600 1800 2000 2200

Pressure, psf

10

20

30

40

50

60

0

Time, sec

Seal

temperature

Effect of

turbulence

assumption

Pressure

1

2

3

6

Location

Tile with RCG coating

Tile with RCG coating

Exposed seal

Titanium attachment

Structure

temperature structures. In particular, the electromechanical

actuators used to move the rudders must be kept cool

enough for them to operate properly. A detailed discussion

of the thermal analysis will be given later in the paper in

the section entitled “Thermal Analysis.”

Pressure Drop

The maximum predicted pressure drop across the seal

during vehicle re-entry is about 56 lbf/ft2 (psf, outboard:

high pressure) and occurs about 450 sec after the peak

heating (fig. 3). To be conservative, flow tests were

conducted up to the peak pressure. The pressure across the

seal during peak heating is 35 psf and occurs at about

1200 sec into the re-entry mission (fig. 3).

Resiliency

No specific design requirement was established for seal

resiliency. A main requirement for the seals is that they

remain in contact with the sealing surface while the vehicle

goes through the maximum re-entry heating cycle. The
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 5

seals must be able to accommodate differences in thermal

expansion between parts adjacent to them that cause the

seal gap to change size. Subsequent to the re-entry heating

cycle, any small thermally induced gap opening is of no

consequence as the convective heating rate drops off

sharply.

Seal Loads/Gap

The seals are to be installed at approximately 20 percent

compression to ensure good sealing contact with the

rudder/fin surfaces (fig. 4). The seals will seal against

Shuttle-derived tile that limits the maximum seal unit

or contact load. The tiles used for the rudder/fin

horizontal shelf sealing surface are Alumina Enhanced

Thermal Barrier—8 lb/ft3 density (AETB–8) with Reaction

Cured Glass (RCG)/Toughened Uni-Piece Fibrous

Insulation (TUFI) coating. Designers have set a unit load

limit of less than 5 lb/in. of seal to prevent tile damage

during installation or actuation. The seals are required

to seal a nominal 0.25-in. gap between the surfaces of the

X–38 rudder and fin.

Life/Wear Resistance

The X–38 rudder/fin seals are only required to last for one

mission and will subsequently be replaced after each

mission. The seals must be robust enough to endure the

scrubbing that they will experience during preflight

checkouts and during the mission. The vertical hinge line

seals are scrubbed over an Inconel sealing surface as the

rudder actuates, but they are designed to remain in contact

with that surface throughout the mission. The rudder/fin

shelf seals experience both scrubbing against the

AETB–8 tiles and a “scissoring” action as they are moved

onto and off of the shelf sealing surface. When the seals are

moved off of the fin shelf they will tend to return to an

uncompressed shape. As they are moved back onto the

surface and compressed again, they must be able to endure

the shear forces that they will be subjected to without

causing excessive loads on the rudder drive motor. Because

the rudder/fin shelf seals will be exposed to more severe

loading conditions, the wear resistance of these seals is

examined as part of the current study.

Figure 4.—Cross section of rudder/fin shelf seal location (standing aft looking forward) showing seal

 components.

Seal bracket

Blanket

Fin structure

Inboard

shelf seal

1.3 in.

0.25 in.

1.0 in.

Lower rub surface

(Shuttle tile)

Outboard

shelf seal

Rudder TPS tiles

Rudder structure

0.5 in.

Flush inboard

position

Flush outboard position

Spring

tube

Core Sheath
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 6

Test Apparatus and Procedures

Seal Specimens

The seal design examined in this study had a nominal

diameter of 0.62 in. (table 1 and fig. 4). It consisted of an

Inconel X–750 spring tube stuffed with Saffil batting and

overbraided with two layers of Nextel 312 ceramic fibers.

The Saffil batting stuffed into the Inconel spring tube had

a density of 6 lb/ft3. This seal design will hereafter be

referred to as the 6 pcf design. The seal is currently used

in several places on the Space Shuttle orbiters including

the main landing gear doors, the orbiter external tank

umbilical door, and the payload bay door vents. It was

selected as the baseline seal design for the rudder/fin

location of the X–38.

Thermal Analysis

An analysis was performed on the X–38 rudder/fin seal

location to predict maximum seal and attachment

temperatures during vehicle re-entry. Figure 3 shows a

schematic of the rudder/fin gap area Thermal Math Model

(TMM) and predicted temperatures for the exposed seal

and surrounding hardware. The TMM was a quasi-twodimensional representation built using the System

Improved Numerical Differencing Analyzer (SINDA)

V3.1. It consisted of approximately 150 nodes that

represented the Thermal Protection System (TPS) tiles,

the dual seals, and the titanium attachment structure

(fig. 3(a)). The TPS material on both the rudder and the fin

was modeled as RCG/TUFI coated AETB–8 tile. The seal

was modeled as Nextel 312 fabric over Saffil batting

(6 lb/ft3). The seal attachment was modeled as a solid

titanium structure. For each material, temperature

dependent and pressure dependent (where required)

properties were used. All connections between dissimilar

materials were assumed to be perfect, i.e., no contact

resistance was modeled. The gap was modeled as being

1.5 in. deep and 0.25 in. wide. All modes of heat transfer

(i.e., conduction, convection, and radiation) were accounted

for in the model. The model did not include the effects of

flow through the seal. The Thermal Synthesizer System

(TSS) was used to resolve the radiation exchange between

all exposed surfaces inside and outside of the gap including

radiation to space. Results from the TSS analysis were

coupled to the SINDA analysis program.

The surface heating used to drive the TMM was based on

Cycle 8 reference heating supplied by the Johnson Space

Center Aeroscience and Flight Mechanics Division

(EG/W. Goodrich). The heating supplied was

representative of the heating predicted on the windward

surface of the rudder/fin area. The heating distribution

within the gap was determined using the gap heating

relationship presented by Nestler.2 Nestler’s empirical

relationship provides for the heat flux to a certain gap

depth (e.g., to the seal) and assumes no flow through the

floor of the gap (e.g., an impermeable seal).

Temperature Exposure Tests

The thermal analysis predicted that the rudder/fin seals

would be exposed to temperatures at or above 1900 °F

during X–38 re-entry. To simulate exposure to such extreme

temperatures and to determine the effects that this exposure

has on the seals, specimens were placed into a tube furnace

in a compressed state and heated at 1900 °F for 7 min.

This 7 min. temperature exposure closely simulated the

amount of time that the seals would spend at the peak

temperature during re-entry (fig. 3(b)).

Seal Type Size

Diameter

in.a

Material Density

lb/ft3

Measured

percent of

seal by 

mass (%)

Material Measured

percent of

seal by

mass (%)

Material Number 

of

layers

Measured

percent of

seal by

mass (%)

6 pcf 0.620 Saffilb 6 12.5 Inco X–750c 33 NX 312d 2 54.5

a

1×10–3 in. = 25 μm b

Saffil insulation, density of individual filament = 0.0975 lbm/in.3 (2.70 g/cm3)

c

Inco X–750 = Inconel X–750: 70% Ni, 15% Cr, 7% Fe, 2.5% Ti, 1% Cb, 0.7% Al, 

density of individual wire = 0.298 lbm/in.3

 (8.25 g/cm3)

d

NX 312 = Nextel 312 fabric, 3M product: 62% Al2O3, 24% SiO2, 14% B2O3,

density of individual filament = 0.123 lbm/in.3

 (3.4 g/cm3)

Core Sheath Spring Tube

TABLE 1.—X–38 SEAL CONSTRUCTION MATRIX
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 7

In these tests, 1-ft-long seal specimens were clamped into

a fixture between two flat stainless steel plates and subjected

to a linear compression of 20, 25, or 30 percent of their

overall diameter of 0.62 in. After temperature exposure,

the fixture and specimen were removed from the furnace

and allowed to cool at room temperature. Specimens were

then removed from the test fixture and subjected to

compression tests to examine the effects that the

temperature exposure had on the stiffness and resiliency

of these seals. Further details of the setup and procedure

for these temperature exposures can be found in the paper

by Dunlap et al.3

Porosity Measurements

Seal porosity was measured using two different approaches.

Archimedes Approach—The first approach was based

on Archimedes’ theory of volume displacement. This

principle states that the volume of liquid that is displaced

when a solid object is dropped into the liquid is equal to

the volume of the solid. Seal porosity was determined by

putting covered and uncovered seal specimens into a

100 ml graduated cylinder partially filled with isopropyl

alcohol and comparing the volume displaced by the

specimen in each condition. A total of four seal specimens

of approximately 2.5 to 3 in. in length were tested. For

each measurement, a seal specimen was wrapped in plastic

wrap and dropped into the graduated cylinder. Covering

the specimens in plastic wrap prevented alcohol from

penetrating into the porous structure of the seal and allowed

the cylindrical volume of the exterior of the seal to be

determined. The volume level of the alcohol in the cylinder

was measured before and after the seal was inserted so that

the difference in volume was the amount displaced by the

seal (Vcovered seal). The specimen was then taken out of the

cylinder, and the plastic wrap was removed from it. The

wrap was then placed into the cylinder by itself to determine

how much volume it displaced (Vplastic wrap). This amount

was subtracted from the volume measured for the wrapped

seal to determine the actual volume that only the exterior

of the seal would have displaced. Finally the unwrapped

specimen was placed into the cylinder. Alcohol easily

soaked into the uncovered seals and filled the voids inside

of them. After allowing the alcohol to absorb into the seal,

the volume displaced by the material in the seal was

recorded (Vuncovered seal). Porosity (ε) was then calculated

using the following equation, where Vf is the fiber volume

ratio

ε= −

= − −













1

1

V

V

V

f

uncovered seal

Vcovered seal plastic wrap

Mass/Volume Approach—A mass/volume technique

for determining seal porosity was used to corroborate the

results of the Archimedes approach. First, overall specimen

volume was determined by making precision

measurements of the cross-sectional dimensions of the

ends of 1-in. long seal specimens compressed in a groove

machined in a test block. A flat plate compressed the seal

in the groove to the 20 percent design compression. The

total volume (VTotal) was determined by multiplying the

measured seal elliptical cross-sectional area times the seal

length. Next, the seal was cut open and separated into its

individual components: the Saffil core, Inconel spring

tube, and Nextel sheath. Each of these constituents was

weighed on a precision (1 mg) mass balance to determine

its respective mass. (Note: Table 1 includes measured

constituent masses as percentages of the total seal mass,

for reference purposes.) Volumes of each of the constituents

were then determined by dividing their masses by their

respective individual fiber densities as shown in table 1.

Finally the fiber volume ratio (Vf

) was found by summing

the individual constituent volumes (Vi) and dividing by

the total measured volume (VTotal). Porosity (ε) was

determined for the two specimens by the following

equation:

ε= − = − 1 1 V V

V f

Σ i

Total

Compression Tests

Compression tests were performed to determine seal

preload and resiliency behavior at room temperature using

a precision linear-slide compression test fixture. A

specimen was loaded into a stationary grooved specimen

holder, and an opposing plate was compressed against the

specimen. The groove was rectangular in shape with a

width of 0.62 in. and a depth of 0.37 in. Stainless steel

shims were placed in the groove behind the specimen to

vary the amount of compression on the seal. The amount

of compressive load on the specimen was measured versus

the amount of linear compression for several load cycles.

Multiple load cycles were applied to the specimen before

the preload data point was recorded to remove the effects

of hysteresis and permanent set that accumulated with

load cycling of the specimens. Most permanent set occurred

within the first four load cycles, so four load cycles were

performed for each test. A pressure sensitive film mounted

on the opposing plate was used to determine the contact

width of the specimen as it was compressively loaded. The

footprint length and width at the end of the fourth load

cycle were used to calculate seal preload in pounds per

square inch. The measured load versus compression data

was used to determine residual interference corresponding

to a given linear crush value.4 Residual interference is
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 8

defined as the distance that the specimen will spring back

while maintaining a load of at least 0.01 lb/in. of specimen.

The hardware and procedure used to perform these tests

are described in detail by Steinetz et al.4 Overall accuracy

of the preload values measured using this method was

calculated to be ±3.4 percent.5

Test Matrix—In previous studies by Dunlap et al.3,6

seal compression tests were performed at compression

levels of 10, 20, 25, and 30 percent of the specimen’s

as-received overall diameter. Control surface seals are

typically installed at approximately 20 percent

compression, including the seals used in the X–38 rudder/

fin seal application. Tests carried out at the low compression

level of 10 percent compression were performed to

determine seal resiliency and preload under minimal

loading conditions. The low compression tests were

performed to simulate conditions in which seals may

become unloaded during use or take on a large permanent

set due to temperature exposure. Tests were also performed

on seal specimens after temperature exposure. Primary

and repeat compression tests were performed for all test

conditions.

Scrub/Wear Tests

To test seal wear resistance, a series of tests were performed

in which the seals were scrubbed over a representative

sealing surface for repeated cycles. These tests also

examined the structural integrity of the seal attachment

technique and the ability of the seals to engage and

disengage the shelf sealing surface while the rudder pivoted

through its ±12° of rotation.

Test Fixture—The seals were evaluated in a test

fixture that simulated the motion of the X–38 rudder with

respect to a stationary sealing surface such as the fin shelf

(fig. 5). In these tests, two seals were mechanically attached

side-by-side to a representative carrier plate by threading

0.020-in. diameter Inconel 600 annealed (MS20995N20)

wires through the outer layers of the bottom of the seal,

passing the ends of the wires through holes in the plate,

and twisting the wire ends to secure the seals (fig. 6).

These loops of wire were located every 0.25 in. along the

carrier plate. Two different seal lengths were tested due to

a limited supply of seal material: a full-scale seal that was

27.5 in. long and a shorter 6.5 in. seal. A coating of

vermiculite (Microlite Vermiculite Dispersions, 923 from

W.R. Grace & Co.) was brushed on the seals. This treatment

has been shown to prevent fraying and reduce abrasion of

fabrics. For each test, the seal/carrier plate assembly was

attached to the top surface of a movable arm that simulated

the rudder in the rudder/fin seal assembly (figs. 2 and 4).

The movable arm was pivoted on one end and mounted on

the underside of a table (fig. 5). At the other end of the arm

was a roller that allowed the arm to be suspended below

the table and to rotate freely with respect to the pivot. Two

stops fixed to the edge of the table limited the amount of

rotation of the arm to ±12°. The seals were scrubbed

against a simulated seal rub surface that was suspended

through a rectangular hole cut in the center of the table.

Four bolts and adjustment nuts were used to compress the

rub surface against the test seals to the design point of 20

percent compression.

The rub surface used for these tests was composed of three

RCG/TUFI coated AETB–8 tiles. The three tiles were

lined up side-by-side opposite the seals along the length of

the movable arm. The total length of these tiles was

identical to the length of the fin shelf sealing surface on the

X–38 rudder/fin. Gap fillers consisting of four-ply planar

Nextel 312 sleeving about 0.10-in. thick were bonded into

the joints between the tiles. Assembly of the scrub test

fixture revealed the need for the gap fillers to be installed

flush with the outer surface of the tiles to minimize sneak

flows between adjacent tiles above the gap fillers. The

27.5-in. seal was scrubbed against all three tiles with gap

fillers between the tiles. Based on the results of this first

test in which damage was observed in areas where the seal

was scrubbed over joints between the tiles, the 6.5-in. seal

was installed for the second test such that the center

portion of the seal passed over a tile joint. In addition,

Kapton tape was placed over the gap between the tiles to

see if smoothing that area would reduce the amount of

damage the seals incurred during scrubbing. Using a

profilometer, the surface roughness of the sanded tiles in

areas away from the joints was measured at 80 to 90 RMS

before testing.

During a test, the arm was manually rotated back and forth

from one stop to the other over the full range of ±12°. Each

rub surface was sized so that the seals would move off of

the edges of the surface and out of contact with it as the arm

was moved between the stops in both directions. This

forced the seals to re-engage the rub surface as the arm

rotated back toward the centerline of the rub surface

(see view A-A in fig. 5). A 0.25-in. radius was fashioned

into the tile edges to ease re-engagement of the seals back

onto the tile surface. A torque meter located at the pivot

point measured the amount of torque required to rotate the

arm and scrub the seals over the rub surface. The torque

meter was also used to measure the amount of torque

required to engage and disengage the seals as they were

moved off of and back onto the rub surface. Both test seals

were subjected to 1000 scrub cycles. This cycle count was

selected as a conservative estimate of the number of cycles

the seals would experience during preflight checkout and

during the single mission.
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 9

RJC

Roller

Stop

(two places)

Stationary fin

A

A

Test seals attached

to movable arm

View A–A

Movable arm:

simulates rudder motion

Figure 5.—Isometric of scrub/wear test fixture showing seals re-engaging the tile rub surface.

Movable

arm

Test seals

Tile rub

surface

Seal compression

adjustment nuts

(four places, two shown)

Pivot point

Tile rub surface

(AETB–8 tile) 

Figure 6.—Photo of double seals mechanically attached 

 to carrier plate with Inconel wires for scrub testing.

Carrier

plate

Double

seals

Inconel

wires

0.25-in.

spacing
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 10

Flow Tests

Flow tests were performed on the seals in an ambient

temperature linear flow fixture shown schematically in

figure 7. The flow fixture was designed so that either

single or double seals of different diameters could be

tested in removable cartridges that are inserted into the

main body of the test fixture. Seals can be tested in this

fixture with different seal gaps and under different amounts

of linear compression.

Flow Path/Instrumentation—During flow testing

pressurized air entered through an opening in the base of

the fixture and passed through a plenum chamber before

reaching the test seal. Air flowed through the gap between

the cartridge and the cover plate, passed through the seal

and its interface with the cover plate, and then flowed out

of the top of the fixture (fig. 7(a)). A flow meter upstream

of the flow fixture measured the amount of flow that

passed through the test seal. The flow meter had a range of

0 to 100 standard liters per minute (0 to 4.5×10–3 lbm/sec)

and an accuracy of 1 percent of full scale. A pressure

transducer (0 to 5 psid, 0.07 percent accuracy) upstream of

the test seal measured the differential pressure across the

seal with respect to ambient conditions, and a thermocouple

measured the upstream temperature.

Test Fixture—Test seals were mounted in the groove

of a test cartridge (fig. 7(b)). Individual cartridges were

machined with a groove width for a single seal of 0.62 in.

and a groove width for a double seal of 1.30 in. Two

different seal lengths were used in this series of tests. Seals

that had been scrub tested were approximately 6.5 in. long,

while as-received seals that had not been scrubbed were

tested in either 6.5- or 12-in. lengths. For the 12-in. seals,

the amount of preload, or linear compression, was varied

by placing steel shims in the cartridge groove behind the

seal. A slightly different setup was used for the 6.5-in.

seals (fig. 7(c)). These seals were mechanically attached

to carrier plates, and the entire assembly of the seal and

carrier plate was installed in the center of the 12-in.-long

groove of the test cartridge. Aluminum blocks with O-ring

grooves in them were installed on either side of the seal

assembly to seal the outboard seal ends.

After a seal specimen was installed in the cartridge, the

cartridge was inserted into the test fixture. An O-ring

sealed the perimeter of the cartridge chamber to prevent

flow from passing behind the cartridge during testing.

Pairs of spacer blocks secured to the cartridge at the ends

of the test specimen controlled the gap width between the

cartridge and the cover plate that the seals sealed against

(figs. 7(b) and (c)). Blocks of different thicknesses were

used to vary the gap width. A small amount of RTV was

placed between the spacer blocks, filler blocks, and the

cartridge to prevent flow from passing through these gaps.

Another O-ring was placed in a groove on the surface of

the test fixture and into a groove in the spacer blocks to seal

the plenum chamber upstream of the test seal. For the tests

performed on the shorter 6.5-in. seals, grooves in the filler

blocks on either side of the seal specimen extended the

O-ring path from the spacer blocks to the ends of the seal

specimen (fig. 7(c)). For both seal lengths, the ends of this

O-ring were pressed up against the ends of the test seal to

prevent flow from passing around the ends of the seal. For

the 12-in. seals, end effect leakage was minimized by

exposing only the center 10 in. of the seal to the prescribed

gap. One inch at each end of the 12-in. test specimen was

embedded into the fixture (i.e., gap width is zero) to reduce

the effects of flow passing between the seal ends and the

O-ring. The ends of the 6.5-in. test seals were coated with

RTV to form a better seal between the seal ends and the

O-ring to minimize end effects. Preload was applied to the

test seal through an interference fit between the seal and

the cover plate.

Test Matrix/Non-Scrubbed Seals—A series of flow

tests was conducted using 12-in. seal specimens. Single

seal flow tests were conducted on non-scrubbed seals at

compression levels of 0, 10, 20, and 25 percent of the

specimen’s overall diameter. Double seal flow tests were

conducted on as-received seals at 0, 10, and 20 percent

compression. The 12-in. seals were not coated with

vermiculite, while the 6.5-in. seals were coated for the

scrub tests.

Test Matrix/Scrubbed Seals—Flow tests were

performed at 20 percent compression on the 6.5-in. double

seals that were attached to carrier plates. Tests were

performed on two 6.5-in. segments that were cut out of the

27.5-in. seal that was used in the first scrub test. One

segment was chosen from an area with the worst amount

of wear, while the other segment was selected as the area

with the least amount of damage. Flow tests were also

done on the seal specimen from the second scrub test that

had been scrubbed against the tile joint that was covered

with Kapton tape. Additional tests were performed on a

baseline non-scrubbed double seal attached to a carrier

plate to determine the effects of scrubbing and carrier plate

attachment on seal flow rates.

To simulate conditions in which a gap filler had fallen out of

the joint between adjacent tiles on the sealing surface, flow

tests were performed in which a groove was cut into the

cover plate. This 0.050-in. wide by 0.060-in. deep groove

simulated a gap between the tiles through which air could

flow past the seals. Tests were performed using the worstworn seal segment from the scrub tests and the non-scrubbed

seal segment sealing against this grooved cover plate.
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 11

Shims Gap

Test seal

Cover

plate

Seal

cartridge

(a) Metered air supply

Plenum

chamber

Figure 7.—Schematic of flow fixture. (a) Cross section. (b) Isometric. (c) Front view of 6.5-in. seal specimen

 installed in flow fixture showing grooved filler blocks that join O-ring to both ends of seal.

Cover

plate

Spacer

block

Test seal

Spacer block

(b)

(c)

Seal cartridge

Spacer 

blocks

(both sides)

Filler blocks with extended 

O-ring path (both sides)

Flow

RTV on seal ends (both sides)

O-ring

Test seal
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 12

To determine the best possible flow rates through these

seals, tests were performed in which RTV was placed in

between a seal and carrier plate and on the backside of the

carrier plate to eliminate these leak paths around the seal.

The seal used for these tests was the least damaged

segment that was cut out of 27.5-in. scrub tested seal.

These tests were performed against the flat cover plate.

All flow tests were performed with a 0.25-in. seal gap.

Both primary and repeat flow tests were conducted for all

of the above test conditions.

Arc Jet Tests

A series of tests were performed on the 6 pcf seal design

in the 20 Megawatt Panel Test Facility at NASA’s Ames

Research Center to simulate exposure of the seals to the

extreme thermal conditions that they would experience

during atmospheric re-entry. The seals were installed in a

test fixture that was positioned in the test chamber such

that high-temperature exhaust flow passed out of the semielliptical nozzle (17 in. wide nozzle) of the arc jet heater

and over the top surface of the test fixture. A gap in the test

fixture allowed hot air to flow down to the seals. During

testing the test chamber was evacuated down to a pressure

of 5×10–2 torr (9.67×10–4 psi) to draw flow out of the arc

jet nozzle, over the test article, and through the seals. The

arc jet facility has the capability of producing heat fluxes

on the order of 0.5 to 75 BTU/ft2 sec with temperatures on

the top of the test article as high as 2200 °F.7

Test Fixture—The test fixture was based on the

geometry of the X–38 body flap but had many similarities

to the rudder/fin hinge-line seal configuration (fig. 8).

It consisted of a stationary upstream structure and a

movable control surface. Test seals were placed into the

gap along the hinge-line between the stationary structure

and the control surface. The depth of the gap from the free

stream to the seal of 1.5 in. was in the range of the depths

from the free stream to the rudder/fin hinge-line seal

depending on rudder position. The 0.25 in. seal gap size is

used in both the X–38 rudder/fin and body flap sealing

applications. The seals were installed in the test fixture so

that they were under 20 percent compression for these

tests. The control surface could be actuated over a range of

0° to 10° with respect to the stationary surface, and the

table upon which the test fixture sat could be adjusted to

vary the angle of attack of the entire test fixture over a

range of –4° (out of the flow) to +6° (into the flow). The

greatest angle that the control surface was raised into the

arc jet flow equaled the sum of both angles (e.g., 6° table

angle +10° control surface angle = 16° control surface

deflection into flow). Raising the control surface into the

arc jet flow raised the static pressure above the sealed

hinge-line, deflected high-temperature flow into the seal

gap, and increased seal and gap temperatures.

The test fixture was composed of a water-cooled copper

box and a movable stainless steel control surface section

that were covered with AETB–8 tiles to simulate the

rudder/fin thermal protection system in the X–38. For

each test a single 19-in. long seal was used with a “tail” of

Nextel fabric sewn on to it. This tail was clamped in

between tiles in the stationary structure of the test fixture

to secure the seal in place (fig. 8). The test fixture was

instrumented with 34 thermocouples and seven pressure

taps to record temperatures and pressures upstream and

downstream of the test specimens and monitor the health

of the test fixture. A more detailed description of the test

fixture, instrumentation, and procedure used to

perform these tests can be found in the final report by

Newquist et al.8

Results and Discussion

Temperature Exposure Test Results

The temperature exposure tests conducted on the seals

caused a distinct change in their shape and properties.

After temperature exposure at 1900 °F for 7 min. while

compressed between two steel plates, the seal specimens

did not return to their original circular cross section. They

took on an elliptical cross section that was quite flat in the

areas that were in contact with the plates (fig. 9). The

specimens were stiffer and much less flexible than they

were before the temperature exposure.

Most of these changes are believed to be due to changes

that occurred in the Inconel X–750 spring tube. The

Inconel X–750 spring tube contributes significantly to the

Figure 8.—Cross section of arc jet test fixture.

Movable

control

surface

Test seal

Seal attachment tail

1.50-in.

gap

depth

Stationary

structure

Flow

0.25-in.

seal

gap
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 13

resiliency of the seals and appeared to have taken on a

large permanent set during these tests. This is because the

yield strength of Inconel X–750 at 1900 °F is less than

5 percent of its room temperature strength.9 This prevented

the seal from returning to its original circular cross section

and caused it to take on the new elliptical cross section.

The Nextel 312 ceramic fabric that formed the outermost

layers of the seals and the Saffil batting in the seal core did

not undergo any noticeable changes during these tests.

Further discussions of the specific changes to seal resiliency

and stiffness due to 1900 °F exposure will be addressed in

the following sections.

Porosity Measurement Results

Porosity measurements were made on four specimens of

the 6 pcf seal design using the Archimedes approach. The

porosity values measured for these specimens ranged

from 83.3 to 85.2 percent with an average porosity of

84.4 percent. This means that almost 85 percent of the

volume of the seal was composed of air, and only

15 percent was actual material. This value was corroborated

by making porosity measurements on two additional

specimens using the mass/volume approach. The porosity

values for these two specimens were 81.3 and 82.6 percent

for an average porosity of 82 percent. The mass/volume

porosity measurement was done in a compressed state that

could account in part for the slightly lower measured

porosity.

The high seal porosity level is attributed primarily to the

loose packing of the 6 pcf Saffil batting. For comparison

purposes, thermal barriers braided from continuous fibers

have porosities in the range of 37 to 50 percent.10 Porosity

is important for understanding the thermal and flow

response characteristics of these seals and is an important

variable used in seal thermal analyses.

Compression Test Results

Table 2 summarizes the results of the compression tests

performed on the 6 pcf seal design in the as-received

condition and after 1900 °F temperature exposure. Values

listed in this table are for single seals and include the

measured residual interference, contact width, unit load,

preload, and seal stiffness per unit inch of seal for each

amount of linear compression at which the tests were

Figure 9.—Photo of 6-pcf X–38 seals before and after performed.

 1900 °F temperature exposure. Side-by-side photo

 of seals before (left) and after (right) temperature

 exposure.

Seal Condition Diameter, Nominal Linear Residual Contact Unit Preload, Stiffness (k)

Type in. linear crush, interference width, load, psi at % linear

crush, in. (springback), in. lbf /in. a

% in. lbf/in.

6 pcf As-received 0.620 10 0.062 0.046 0.330 0.54 1.7 14

20 0.124 0.084 0.455 2.01 4.4 39

25 0.155 0.115 0.581 2.98 5.1 51

30 0.186 0.118 0.692 4.47 6.4 66

6 pcf After 0.620 20 0.124 0.018 0.379 0.91 2.4 58

1900 oF 25 0.155 0.036 0.452 1.77 3.9 76

exposure 30 0.186 0.029 0.489 1.90 3.9 106

a

Seal stiffness is calculated as the slope through the final two data points at the maximum amount 

of compression.

TABLE 2.—X–38 SEAL RESIDUAL INTERFERENCE, CONTACT WIDTH, UNIT LOAD,

PRELOAD, AND STIFFNESS FOR SEVERAL LINEAR CRUSH CONDITIONS

crush,
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 14

Residual Interference/Resiliency—The residual

interference, or springback, of the seals generally increased

as percent linear compression increased (table 2). However,

exposure of the seals in a compressed state at 1900 °F for

7 min. caused a large permanent set and loss of resiliency.

At each compression level, the residual interference, or

springback, of the temperature-exposed seals was only

20 to 30 percent of that for the as-received seals. As

discussed previously, most of this loss of resiliency is

believed to be due to permanent set that occurred in the

Inconel X–750 spring tube during temperature exposure

testing. In the as-received seal, the spring tube contributes

significantly to the resiliency of the seals.

The main requirement for the X–38 rudder/fin seals, and

control surface seals in general, is that they remain in

contact with the sealing surface while the vehicle goes

through the maximum re-entry heating cycle. Good seal

contact is required to prevent hot gases from leaking past

the seals and into cavities behind them in which lowtemperature structures reside. An additional requirement

is that the seals are able to accommodate differences in

thermal expansion between parts adjacent to them that

cause the seal gap to change size. Because these seals

experienced a large permanent set and loss of resiliency

upon temperature exposure, designers were forced to

change the X–38 rudder/fin vertical rub surface design to

ensure that the seals remained in contact with the sealing

surface during re-entry. Finite element analyses were

performed to determine how different Inconel rub surface

designs would deform and deflect under anticipated

mechanical and thermal loads during flight. Different rub

surface thicknesses, constraint schemes, and attachment

methods were evaluated. Early designs yielded rub surfaces

that moved away from the seals such that they were not in

contact with the seals during maximum heating conditions

(fig. 10). Some locations on the rub surface moved as

much as 0.163 in. away from the seals on the outboard side

of the rudder/fin. After several iterations, a final Inconel

rub surface design was selected that was 0.05 in. thick with

gussets distributed on its outboard side to restrict deflection

based on the predicted temperature distribution. This

design provided engagement between the seals and the rub

surface during maximum heating conditions (fig. 10).

Unit Load (Load per Unit Inch)/Preload/Seal Stiffness—

The amount of unit load (or load per unit inch), seal preload

(or footprint contact pressure), and seal stiffness per unit

inch of seal increased as the amount of linear crush was

increased on both the as-received and temperature-exposed

seals (table 2). Although the temperature-exposed seals

were noticeably stiffer and less flexible to the touch than

the as-received seals, unit loads and preloads were lower

for the temperature-exposed seals. This can be explained

by the differences in stiffness between seals in the two

conditions. Seal stiffness per unit inch of seal was calculated

as the slope through the final two data points on a load versus

compression plot at the maximum amount of compression.3

Seal stiffness after 1900 °F exposure was 1.5 times higher

than for the as-received seals at 20 percent compression.

Therefore, even though unit loads and preloads were lower

for the temperature-exposed seals, they were stiffer than the

as-received seals at the same amount of compression. This

increased stiffness is again believed to be due to changes that

occurred to the Inconel X–750 spring tube during the

1900 °F exposure.

Sealing surfaces that control surface seals mate against

include Shuttle-type thermal tiles, ceramic matrix

composite structures, and metallic structures. Of these

Figure 10.—Schematic of X–38 rudder/fin cross section showing vertical Inconel rub 

 surface redesigned to accommodate lack of seal resiliency (note: not drawn to scale).

0.035 in.

maximum

0.163 in. 

maximum

Inboard

side

Outboard

side

Actuator

location

Seals

Rudder

Early rub surface design: 

moves away from seals 

during re-entry heating

Redesigned rub surface: 

stays in contact with seals

during re-entry heating
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 15

surfaces, Shuttle thermal tiles are the most damage-prone

in terms of the unit loads that they are able to resist. In the

X–38 rudder/fin seal application, for instance, the seals

should not put a load of more than 5 lb/in. of seal on the

thermal tiles that make up the sealing surfaces. For this

application, the seals are to be installed at approximately

20 percent compression with a nominal 0.25-in. gap

between the surfaces of the rudder and fin. The results in

table 2 show that unit loads were below 5 lb/in. of seal for

all compression levels tested. The maximum seal preload,

or contact pressure, that was measured was 6.4 psi for the

as-received seal at 30 percent compression. Even at this

high level of compression, the pressure that would be

applied to the tiles would be seven times lower than the

flatwise tensile strength of 46 psi for the tiles and nine

times lower than the compression strength.3 The results of

these compression tests indicate that the 6 pcf seals meet

the seal load requirements established for the X–38 rudder/

fin seal application.

Scrub/Wear Test Results

In the first scrub test, the 27.5-in. seal was scrubbed for

1000 cycles over the three RCG/TUFI coated AETB–8

tile rub surfaces. After the test, there were several areas of

the seal that were clearly more damaged than other areas.

At seal locations approximately 8, 17, and 21 in. from the

pivot point of the test fixture the outer Nextel sheath layers

were so worn away that the Inconel spring tube was

exposed. Figures 11(a) and (b) show these worn areas of

the seal. Away from these specific locations, the seal was

in much better condition. The outer Nextel sheath layers

showed signs of wear and some fiber breakage, but they

were still intact. After 1000 scrub cycles, broken Nextel

fibers were spread over the area in which the seals were

swept over the tiles, but the tile surfaces were still in good

condition. There was some minor wear along one tile edge

17 in. from the pivot point.

Figure 11.—Photos of seals after 1000 scrub cycles against RCG/TUFI-coated AETB–8 rub sur-

 face. (a) 27.5-in. seal with damage indicated at locations 17 and 21 in. from pivot of scrub

 test fixture. (b) 27.5-in. seal with damage indicated at location 8 in. from pivot of test fixture. 

 (c) 6.5-in. seal with minimal damage after scrubbing against tiles with Kapton tape covering

 gap between adjacent tiles.

Damage at 21 in.

Damage at 8 in.

Damage at 17 in.

Pivot point (a)

(b)

(c)

Pivot point

Location of Kapton covered tile gap
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 16

The major cause of seal wear and degradation in this test

was roughness associated with the joints between adjacent

tiles. The seal locations that were more heavily damaged

at 8 and 17 in. from the test fixture pivot corresponded to

locations of joints between tiles. Three tiles were used to

form the rub surface, and gap fillers were stuffed into the

joints between the tiles. The gap fillers were installed flush

with the outer surface of the tiles as they would be in a

vehicle to minimize sneak flows between adjacent tiles

above the gap fillers. To make the tiles smoother than they

were in their as-fabricated condition, the tiles were sanded

by hand to a surface roughness of 80 to 90 RMS. The

process of hand sanding the tiles produced very smooth

surfaces in the centers of the tiles but left the tile edge

surfaces slightly rougher. The rougher tile edges and the

gap fillers combined to wear away the outer Nextel sheath

layers of the seal during the 1000 cycle scrub test. Areas

of the seal away from tile joints showed only superficial

Nextel fiber damage.

Based on the results of the first scrub test, a second test was

performed in which a 6.5-in. seal specimen was scrubbed

over an area of the rub surface in which the joint between

two tiles was covered with a layer of Kapton tape. This

was done to remove the tile edges and gap fillers as sources

of roughness for seal degradation. The condition of this

seal after 1000 scrub cycles is shown in figure 11(c). The

seal showed few signs of damage with the Nextel sheath

layers remaining intact throughout the test. Clearly, the

Kapton tape was very effective in reducing seal degradation

at a tile intersection. The tape was evaluated in this study

as a means of mitigating seal damage during considerable

preflight cycling of the rudders and seals for checkout

purposes. It is recognized that the tape would not survive

re-entry temperatures.

Another test goal was to evaluate how well the seals were

able to move off of and then re-engage (e.g., scissor) the

edge of the tile rub surface. This is a unique feature of the

X–38 rudder/fin seal application in that control surface

seals generally remain in contact with their sealing surface

during use. Observation during both scrub tests revealed

that the seals were able to disengage and re-engage the

edges of the tiles satisfactorily. They remained securely

attached to the movable arm, qualifying the Inconel wire

attachment technique. The amount of torque required to

disengage and re-engage the seals was recorded

periodically during each test. These torque measurements

are shown in figure 12. For both seals the amount of torque

needed to disengage and re-engage the seals dropped off

after the first cycle, leveled off at about half of the initial

torque value around the 100th cycle, and remained

relatively constant for the remainder of the 1000 scrub

cycles. The drop in torque over the first 100 cycles was

caused by the seals settling into the groove into which they

were installed and experiencing some wear. This reduction

in torque with cycling is beneficial in that it puts less stress

on the rub surface tiles over time. The measured peak

torque of 115 ft-lb is less than 13 percent of the rated

actuator torque of 875 ft-lb (at a rotation speed of

30 degrees per second) ensuring that the X–38 rudders can

be actuated with the seals installed.

Figure 12.—Torque results for scrub tests of seals against RCG/TUFI-coated AETB–8 rub surface. Range and 

 midpoint of torque values shown for a given scrub cycle. Torque to rotate compressed seal, ft-lb

27.5-in. seal 

6.5-in. seal scrubbed 

 over Kapton tape 

0 100 200 300 400 500 600 700 800 900 1000

Number of scrub cycles

0

20

40

60

80

100

120

115

105

95

82

80

77

70 68

67 59

55

60

70

60

55

24

22

27

22

23

20

27

25

27

25

30

28

52

40

35

32 30

73
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 17

Results of these tests also showed that reducing the

roughness of the sealing surface by sanding the tiles and

covering joints between tiles enabled the seal to endure a

1000 cycle scrub test. Similar tests in which the same seal

design was scrubbed against unsanded tiles revealed

significant damage to the seals after only 100 scrub

cycles.6 For applications in which this type of seal would

be required to withstand 1000 or more scrub cycles over

multiple missions, making the seal rub surface as smooth

and continuous as possible is critical.

Flow Test Results

Seal flow rates are summarized in figures 13 and 14.

Figure 13 presents flow rates at compression levels of 0,

10, 20, and 25 percent for both single and double seals that

had not been scrub tested. Figure 14 shows flow versus

pressure curves for the variety of test conditions at which

the 6.5-in. double seals were tested. The flow rates shown

in these figures are presented as the measured flow rate at

room temperature divided by the length of seal exposed to

flow in the test fixture (either 6.5 or 10 in.).

Effect of Compression Level—As shown by the flow

results in figure 13, flow rates decreased with higher

compression levels. As the amount of compression on the

seals was increased from 0 to 25 percent the amount of

flow through the seals decreased. This is to be expected as

the act of compressing these seals closed the gaps and flow

paths in their porous structures and allowed less flow to

pass through them. Note in figure 13 that the flow for a

double seal at 0 percent compression was nearly twice that

for a double seal at 20 percent compression. Flow rates

measured at 0 percent compression represent seals in an

unloaded condition.

Effect of Single Versus Double Seals—Flow rates

through a double seal configuration were lower than those

for a single seal at 0, 10, and 20 percent compression

(fig. 13). Adding a second seal into the flow path reduced

flow through the seals by roughly 17 to 26 percent as

compared to flow rates through single seals at the same

compression level. Although the second seal caused a

drop in flow rates, it did not cut the flow in half. This type

of behavior in multiple seal flow tests was observed

previously by Steinetz et al.5 and will be discussed further

in the section on seal permeability.

Note also that the flow rate for the 6.5-in. double seal was

about 10 percent higher than the flow rate through the

12-in. seal (with a 10-in. flow path) at the same compression

level. This could be due partly to end effects for the

6.5-in. seals that were not present for the longer seals.

As discussed previously, the 12-in. seals were installed

such that 1 in. at each end of the seal was embedded into

the fixture to reduce the effects of flow passing between

the seal ends and the O-ring. Because the 6.5-in. seals were

installed in the center of the 10-in. flow path, the butt joint

between the seal ends and the O-ring was in the flow path.

Another possible reason for the higher flow rate through

the 6.5-in. seals was that these seals were mechanically

attached to a carrier plate, whereas the 12-in. seals were

installed as is into the seal cartridge. The shorter seals

could have incurred some minor damage in the process of

attaching them to the carrier plate that could have led to

slightly higher flow rates. It is also possible that the

vermiculite coating on the 6.5-in. seals reduced the

compliance of the outer Nextel sheath and led to higher

flow rates.

Effect of Scrubbing—Figure 14 presents flow results

for the 6.5-in. seals both in non-scrubbed and scrub-tested

conditions. Flow rates through each of the scrubbed seal

specimens were actually lower than the flow rates through

the non-scrubbed seal. This type of behavior has been

observed before by Steinetz et al.11 Flow rates decreased

after scrub cycling most likely because the seals and the

nature of the braid surface compacted during scrub testing.

Of the seal specimens that were scrub tested, the segment

of the 27.5-in. scrub tested seal with the worst damage had

the highest flow rates while the segment of that seal with

the least scrub damage had the lowest flow rates. The seal

segment with the worst damage most likely had the

highest flow rates out of the scrubbed seals because its

outer sheath layers were quite damaged. This seal segment

included both of the damaged areas shown in figure 11(a)

in which the seal sheath was worn away so that the Inconel

spring tube was showing. These damaged areas likely

allowed more flow to pass through the interface between

the seal surface and the surface of the cover plate, resulting

in higher flow rates. The seal that was scrubbed over the

tile joint with the Kapton tape over it had flow rates that

were only slightly higher than those for the least damaged

segment of the 27.5-in. scrub tested seal. For example, the

flow rate through the seal scrubbed against Kapton tape

was only 5 percent higher than the flow rate for the least

damaged seal at 56 psf (fig. 14). This difference in flow

rates is basically negligible given that both seal specimens

showed little wear after 1000 scrub cycles.

The results of these tests show that some amount of seal

scrubbing while under compression actually improves the

flow resistance of these seals. This drop in flow rates

through scrubbed seals occurs until the seals reach a point

at which they become overly damaged. At this point the

damaged areas of the seal cause seal flow rates to increase.
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 18

4.5x10–4

Figure 13.—The effect of number of seals, compression level, and seal length on seal flow. Seals 

 are in as-received condition (nonscrubbed), P = 56 psf, gap = 0.25 in. Flow per inch of seal, lbm/sec/in. 3.0

3.5

4.0

4.27

3.55 3.67

2.72

2.41

1.89

2.18

2.5

2.0

1.5

1.0

0.5

0% 10% 20% 25%

0.0

Single seal, 10-in. flow path 

Double seal, 10-in. flow path 

Double seal, 6.5-in. flow path 

2.10

Compression level

Figure 14.—Flow versus pressure data for 6.5-in. double seals at 20 percent compression under

 different test conditions with a gap of 0.25 in. Flow per inch of seal, lbm/sec/in. 1.0

0.5

0 60

56 psf

X–38 maximum

10 90 110 20 100 120 30 40 50 70 80

Pressure drop across seal, psf

0.0

3.0

2.5

2.0

1.5

3.5

4.0x10–4

Nonscrubbed baseline 6.5-in. seal

Worst scrub damage

Scrubbed against Kapton tape

Least scrub damage

Least scrub damage, with RTV

Grooved Face Plate

Nonscrubbed

Worst scrub 

damage

Flat Face Plate
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 19

Effect of Groove in Cover Plate—To simulate flow

conditions in which a gap filler had fallen out of the joint

between adjacent tiles on the sealing surface, flow

tests were performed in which a 0.050-in. wide by

0.060-in. deep groove was cut into the cover plate. Tests

were performed on both the worst worn seal segment from

the scrub tests and on the non-scrubbed seals. Flow rates

using the grooved cover plate increased for both seals as

compared to flow rates using the flat cover plate (fig. 14).

At a 56 psf pressure drop across the seals, flow through the

non-scrubbed seal increased by more than 9 percent, and

flow rates for the scrubbed seal increased by over

16 percent. These results are to be expected because the

groove provided an additional flow path through which air

could flow past the seals.

Effect of RTV Behind Seals—A set of flow tests was

performed using the flat cover plate and the least damaged

segment from the 27.5-in. scrub tested seal in which leak

paths around the backside of the seal were eliminated by

placing RTV on the backside of the carrier plate and in the

interface between the back of the seal and the carrier plate.

As shown in figure 14, the flow rate for the seal with RTV

behind it was reduced by over 10 percent at a pressure of

56 psf as compared to the flow rate through the same seal

without RTV. Flow rates such as these represent a bestcase condition in which the seals had been compacted but

not damaged by scrub cycling and had not yet gone

through a re-entry heating cycle that would destroy the

RTV.

Seal Permeability—In addition to porosity and flow

rates through the seals, seal permeability is often used to

represent the resistance to flow through a seal. In this study

permeability was defined as follows:

K = ( m ̇ * μ * L)/(ρ * A * ∆P)

where

K permeability (ft2)

m ̇ flow through the seal (lbm/sec)

μ dynamic viscosity of air (lbf-sec/ft2)

L length of flow path through seal (ft)

ρ density of air at 300K (lbm/ft3)

A frontal area of seal exposed to flow (ft2)

∆P pressure drop across seal (lbf/ft2)

The dynamic viscosity of air (μ) used for these calculations

was 3.86×10–7 lbf-sec/ft2. The density of air (ρ) was

0.0768 lbm/ft3. The length of the flow path through the

seal (L) was either 0.62 in. (0.0517 ft) for a single seal or

1.30 in. (0.108 ft) for a double seal. The frontal area of the

seal exposed to flow (A) was calculated as the length of the

seal exposed to flow (either 6.5 or 10 in.) multiplied by the

gap size of 0.25 in. Thus, values for the frontal area were

either 2.5 in.2 (0.0174 ft2) for the longer seals or 1.63 in.2

(0.0113 ft2) for the shorter seals. The remaining variables

of flow through the seal ( m ̇ ) and pressure drop across the

seal (∆P) came from the flow test results for each seal.

Figure 15 presents seal permeability for non-scrubbed

single and double seals as a function of compression level.

The permeability values plotted in this figure were

determined using the flow rate for each seal at a pressure

drop across the seal of about 56 psf. For both the single and

double seal configurations, seal permeability decreased as

the amount of compression on the seals increased. Because

the permeability calculation is based on measured flow

rates, this result is to be expected as the flow rates also

decreased with increases in compression level.

Seal permeability for a double seal configuration was

higher than it was for a single seal (fig. 15). This is due to

the way that the permeability equation is defined.

Permeability is a direct function of the flow through the

seal ( m ̇ ) and the length of the flow path through the seal

(L). As discussed previously, addition of a second seal

into the flow path reduced the amount of flow past the

seals, but it did not cut the flow rate in half. Adding a

second seal doubled the length of the flow path through

the seals, though. Because both of these terms are in the

numerator of the permeability equation, doubling the flow

path through the seal while reducing the flow rate by less

than half caused an increase in permeability for the double

seals as compared to the single seal configuration.

Seal permeability values for the 6.5-in. double seals at

20 percent compression with a 56 psf pressure drop across

the seals are presented in table 3. These permeability

values are referred to as “apparent” permeability because

the flow rates used in these calculations were a function of

the test conditions. For example, flow tests were performed

using a grooved cover plate that allowed air to flow around

the seals. The flow rates measured in those tests were a

function of both flow through the seals and flow around

the seals. The permeability values shown in table 3

assumed the same ranking as the flow rates plotted in

figure 14 such that the seal with the highest flow rates also

had the highest permeability. This is to be expected

because the permeability calculation is based on the

measured flow rates.
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 20

Figure 15.—Seal permeability versus compression level for nonscrubbed single and 

 double seals.

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2x10–9

0 5 10 15 20 25

Compression level, %

Permeability, ft2 

Single seal, 10-in. flow path 

Double seal, 10-in. flow path 

Double seal, 6.5-in. flow path 

Face Plate Condition Seal Condition Permeability

(ft2×10–9)

Flat Non-scrubbed, baseline 6.5-in. seal 1.15

Flat Worst scrub damage 0.86

Flat Scrubbed against Kapton tape 0.81

Flat Least scrub damage 0.77

Flat Least scrub damage, with RTV 0.70

Grooved Non-scrubbed 1.26

Grooved Worst scrub damage 0.99

AT 20 PERCENT COMPRESSION WITH 56 PSF PRESSURE DROP

ACROSS SEALS

TABLE 3.—APPARENT PERMEABILITY OF 6.5-IN. DOUBLE SEALS 
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 21

Arc Jet Test Results

The data presented in this section is sample data from a

series of 12 test runs performed on several different seal

designs under a variety of test conditions. Details of the

complete test program funded by NASA Glenn are in a

final report by Newquist et al.8

Open Gap Test—The results of an arc jet test

performed with no seal installed in the test fixture are

shown in figure 16. This test (Test no. 12 in the series)

served as a baseline to determine how hot the open gap in

the fixture would get with no seal installed. The open gap

was designed to be nominally 0.25 in. wide but ended up

ranging from 0.288 in. at a control surface angle of 0° to

0.260 in. at 10°. The test was performed with the overall

test fixture (table) angled up into the arc jet flow at a 6°

angle. The control surface was initially fixed at its baseline

0° position when the test started, and it was held there until

the top surface temperature stabilized at about 2200 °F.

The control surface was held at 0° for 39 sec under these

maximum heating conditions and was then rotated upward

into the stream an additional 2° and held in that position for

the remainder of the test. Including the 6° angle of the

overall test fixture, the control surface total angle into the

flow was 8°. After 23 sec in this configuration, portions of

the test fixture reached their temperature limit, and the test

was ended.

Figure 16 shows that the temperature inside the seal gap at

a position 0.5 in. above the usual seal location reached

2230 °F while the control surface was at its 0° position and

peaked at 2240 °F at the 2° position. The gap temperature

0.5 in. below where the seal would have been located

reached 2010 °F at the 0° control surface position and

2100 °F by the end of the 62 sec test (i.e., 2° control surface

angle). Although no seal was installed in the gap for this test,

the temperature drop across the seal location can be evaluated

as the difference between the temperatures recorded

0.5 in. above and below where the seal would have been.

Before the control surface was rotated, the temperature drop

across the seal location was about 220 °F. After the control

surface was rotated, the temperature drop decreased to

140 °F by the end of the test. The temperature 1.5 in. below

the seal position rose to 1620 °F by the end of the test. These

results confirm that a seal is required in the X–38 rudder/fin

gap and in control surface gaps in general to reduce heat

fluxes into the gap and protect underlying structures.

The study by Wong et al.1 also predicted high temperatures

in an open rudder/fin gap and emphasized the need for a

seal in the gap.

Figure 16.—Temperatures and pressure differential measured during arc jet test with no seal installed

 (test no. 12), 6° table angle, 0° and 2° control surface angles, and 0.25-in. nominal gap. (Note that the

 symbols on the graph are given for identification only; data were recorded every 1 sec.) Temperature, °F

Differential pressure across seal, psf

800

400

0 400 600

Time, sec

200

Control surface angle

800

18

16

14

12

10

8

6

4

2

0

1000

0

2000

1600

1200

2400

0° 2°

1.5 in. below seal, TC19

0.5 in. below seal, TC16

0.5 in. above seal, TC10

Top surface, TC4

Differential pressure
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 22

Arc Jet Test with Seal Installed—Figure 17 shows the

results of an arc jet test (Test no. 5 in the series) with a seal

installed in the gap. The 6 pcf seal was installed at

20 percent compression in a nominal 0.25 in. gap. As in the

open gap test, this test was performed with the overall test

fixture (table) rotated up into the arc jet flow at a 6° angle.

The control surface was again fixed at its baseline 0°

position when the test began and held there until the top

surface temperature reached a steady state temperature

condition of about 2200 °F. The control surface was held

at 0° for 38 sec and then rotated upward in 2° increments

approximately every 45 sec until it was angled 10° with

respect to the upstream stationary portion of the test

fixture. It was held in this final position for an additional

41 sec before the test was ended. Including the 6° table

angle, the final control surface position was rotated upward

16° into the arc jet flow. The total time spent at maximum

heating conditions (2200 °F on the top surface) was

263 sec. This is comparable to the 250 sec of peak heating

for the rudder/fin seals predicted between 1100 and

1350 sec of the X–38 re-entry mission (fig. 3). During the

arc jet test the average pressure differential across the seal

during maximum heating conditions was 15.6 psf,

indicating that flow passed through the seal (fig. 17). This

pressure level was about 44 percent of the 35 psf pressure

predicted at the 1200 sec maximum heating point during

X–38 vehicle re-entry (fig. 3).

The average temperature on the top surface of the test

fixture during maximum heating conditions was 2220 °F.

With a seal installed in the gap, the temperature 0.5 in.

above the seal reached 1610 °F for a 0° control surface

angle with the table angle at 6°. The temperature 0.5 in.

above the seal reached 1920 °F by the end of the test with

the control surface at 10° and a 6° table angle. Clearly,

installing a seal in the gap created a flow block that limited

the amount of heat convected into the gap under these

extreme test conditions.

Temperature Drop Across Seal—Temperatures

recorded 0.5 in. below the installed seal were much lower

than those recorded in the open gap test. During maximum

heating conditions, the peak temperature only reached

207 °F, resulting in a temperature drop across the seal of

about 1710 °F (fig. 17). Temperatures recorded 1.5 in.

below the installed seal barely increased during this test,

reaching a peak temperature of 101 °F. Installing a seal in

the gap reduced gas temperatures to a level at which an

electromechanical actuator behind the seal would survive

re-entry.

The seal specimen that was used for this test survived the

arc jet exposure. A limited amount of damage was caused

to the outer Nextel sheath layers of the seal due to limited

actuation (less than 10 cycles) of the control surface

during the test. Some broken fibers were seen spread over

the surface of the control surface in areas where the seal

was wiped over the surface, but the seal was generally in

good condition after the test.

0

Figure 17.—Temperatures and pressure differential measured for arc jet test with seal installed at 20

 percent compression (test no. 5), 6° table angle, 0°, 2°, 4°, 6°, 8°, and 10° control surface angles, 

 and 0.25-in. nominal gap. (Note that the symbols on the graph are given for identification only; data

 were recorded every 1 sec.) Temperature, °F

Differential pressure across seal, psf

800

400

400 600

Time, sec

200 800

18

16

14

12

10

8

6

4

2

0

1000

0

2000

1600

1200

2400

1.5 in. below seal, TC19

0.5 in. below seal, TC16

0.5 in. above seal, TC10

Top surface, TC4

Differential pressure

0° 2° 4° 6° 8° 10° Control surface angle
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 23

It is clear from the results of these tests that installation of

a seal in the gap of the test fixture caused a large temperature

and pressure drop across the seal location as compared to

an open gap condition. The seal acted as an effective

thermal barrier limiting heat fluxes through the seal gap

and minimizing temperature increases downstream of the

seal during maximum heating conditions. The pressure

differential measured across the seal was 44 percent of the

35 psf maximum pressure predicted at the 1200 sec

maximum heating point during X–38 vehicle re-entry (fig.

3). Larger pressure drops during re-entry could potentially

cause more flow through the seal with higher temperatures

downstream of the seal. However, only one seal was used

in these tests whereas two seals will be installed side-byside in the X–38 rudder/fin seal application. This will drop

the amount of flow through the gap as shown by the results

of the flow tests presented earlier in this paper.

Seal Requirements for X–38 Versus

Future Reusable Re-Entry Vehicles

Seal requirements for the X–38 vehicle are different from

those of future reusable re-entry vehicles in several ways.

The most obvious difference is that the X–38 seals are

expected to be replaced after each mission while seals for

future reusable re-entry vehicles would be required to last

for hundreds of missions. Reusable seals will need to be

wear resistant to endure many scrub cycles against sealing

surfaces to allow for multiple missions without being

replaced. They will also need to be resilient after repeated

temperature exposures.

Control surface seals for the X–38 are expected to be used

at much higher temperatures (1900 °F) than similar seals

that are used as elevon and body flap seals on the Space

Shuttle. These Shuttle seals are generally used at

temperatures less than 1500 °F. Seals for future reusable

re-entry vehicles will likely be placed closer to the vehicle

surface and closer to the extremely hot gases that pass over

the outside of the vehicle during re-entry. Thus, future

control surface seals will need to endure temperatures

similar to those predicted for the X–38 seals, if not higher.

The seal designs in this study took on a large permanent set

after one temperature exposure due mostly to permanent

deformation of the Inconel X–750 spring tube. This loss of

resiliency after temperature exposure most likely would

not be acceptable for reusable applications. Most metals

cannot endure the high temperatures that the seals will

experience in high-temperature applications without

causing the seals to lose resiliency. For future applications

the seals will most likely have to be composed entirely of

ceramic components. Preloading devices can also be placed

behind the seals to improve resiliency. Requirements for

higher temperatures and reusability in future re-entry

vehicles will necessitate that novel seal designs are

developed that exhibit low flow rates and remain resilient

and wear resistant for multiple missions.

Summary and Conclusions

Re-entry vehicles generally require some combination of

control surfaces (e.g., rudders, body flaps, elevons) to

steer or guide the vehicles during re-entry into and through

the Earth’s atmosphere. Control surface seals are installed

between these movable surfaces and stationary portions of

the vehicle both along hinge lines and in areas where

control surface edges seal against the vehicle body. These

seals must operate at high temperatures and limit hot gas

ingestion and transfer of heat to underlying lowtemperature structures to prevent over-temperature of

these structures and possible loss of vehicle structural

integrity. The main objective of this study was to evaluate

a currently available control surface seal design for

applications in future re-entry vehicles based on the design

requirements of the X–38 rudder/fin seal system as a case

study.

The baseline seal examined in this study was a thermal

barrier used in several locations on the Space Shuttle.

A thermal analysis of the rudder/fin seal assembly based

on representative heating rates predicted a peak seal

temperature of 1900 °F. Seals were heated in a compressed

state at this peak temperature to evaluate the effects of

temperature exposure. Room temperature compression

tests were performed to determine load versus linear

compression, preload, contact area, stiffness, and resiliency

characteristics for as-received and temperature-exposed

seals. Seal scrub tests were performed to examine durability

and wear resistance and to recommend surface treatments

required to maximize seal wear life. Flow tests were

conducted at ambient temperature to examine seal leakage

characteristics both before and after scrubbing. Arc jet

tests were performed to experimentally determine

anticipated seal temperatures for representative flow

boundary conditions (pressures and temperatures) under

simulated vehicle re-entry conditions. Based on the results

of the these tests, the following conclusions are made:

1. Exposure of the seals in a compressed state in a tube

furnace at 1900 °F for 7 min. resulted in permanent set and

loss of seal resiliency presumably due to yielding of the

Inconel X–750 spring tube. This loss of seal resiliency was

accounted for in the re-design of the vertical Inconel

rudder/fin rub surface. The re-designed Inconel rub surface

was made stiffer to limit thermally induced movements

away from the seals so that seal contact would be
American Institute of Aeronautics and Astronautics

NASA/TM—2002-211708 24

maintained. From a resiliency standpoint the double seal

should follow the Inconel rub surface for a single use but

should be replaced before subsequent flights.

2. Unit loads and contact pressures for the as-received

6 pcf seal were below the 5 lb/in. and 10 psi Shuttle

thermal tile limits for all compression levels that were

tested.

3. The seals survived a 1000 cycle ambient temperature

scrub test against sanded RCG/TUFI coated AETB–8 tile

surfaces. They were able to disengage and re-engage the

edges of the rub surface tiles while remaining securely

attached to their carrier plate, thereby qualifying the

Inconel wire mechanical attachment method. Furthermore,

the seals did not damage the Shuttle tiles that they were

scrubbed against. Finally, making the seal rub surface as

smooth and continuous as possible greatly improved seal

wear resistance.

4. Arc jet test results confirmed the need for seals in

the rudder/fin gap location. Installation of a single seal in

the gap of the test fixture caused a large temperature drop

(1710 °F) across the seal. The seal acted as an effective

thermal barrier to limit heat convection through the seal

gap and minimize temperature increases downstream of

the seal (e.g., to the actuator cavity) to acceptable

(<200 °F) levels.

The results of these tests have verified that this seal is

satisfactory for the X–38 application. However,

requirements for higher temperature limits and 100 to

1000 cycle reusability in future reusable launch vehicles

necessitate the development of high-temperature seal

designs that remain resilient for multiple missions while

still exhibiting low flow rates and good wear resistance.

References

1. Wong, H. and Kremer, F.G.J., “Numerical

Assessment on the Heating of the Rudder/Fin Gap in

X–38 Space Vehicle,” International Symposium on

Atmospheric Reentry Vehicles and Systems,

Arcachon, France, March 1999, pp. 1–10.

2. Nestler, D.E., “An Engineering Analysis of

Reattaching Shear Layer Heat Transfer,” AIAA

Paper 72–717, June 1972.

3. Dunlap, P.H., Steinetz, B.M., and Curry, D.M.,

“Rudder/Fin Seal Investigations for the X–38 ReEntry Vehicle,” NASA/TM—2000-210338/REV1,

AIAA–2000–3508, November 2000.

4. Steinetz, B.M., Adams, M.L., Bartolotta, P.A.,

Darolia, R., and Olsen, A., “High Temperature

Braided Rope Seals for Static Sealing Applications,”

NASA TM–107233 rev., July 1996, also Journal of

Propulsion and Power, Vol. 13, No. 5, 1997,

pp. 675–682.

5. Steinetz, B.M. and Adams, M.L., “Effects of

Compression, Staging, and Braid Angle on Braided

Rope Seal Performance,” NASA TM–107504,

July 1997, also Journal of Propulsion and Power,

Vol. 14, No. 6, 1998, pp. 934–940.

6. Dunlap, P.H., Steinetz, B.M., Curry, D.M.,

Newquist, C.W., and Verzemnieks, J., “Further

Investigations of Control Surface Seals for the X–38

Re-Entry Vehicle,” NASA TM–210980, AIAA–

2001–3628, July 2001.

7. Newquist, C.W., Verzemnieks, J., Keller, P.C., and

Shorey, M.W., “Advanced High Temperature

Structural Seals,” NASA CR–210522, November

2000.

8. Newquist, C.W., Verzemnieks, J., Keller, P.C., and

Shorey, M.W., Final Report for Contract Number

F33615–95–D–3203, Delivery Order Number 0014,

“Advanced High Temperature Structural Seals,”

May 2002.

9. Kattus, J.R., “Inconel X–750,” Aerospace Structural

Metals Handbook, Vol. 4, Article 4105, Metals and

Ceramics Information Center, Battelle Labs,

Columbus, OH, 1990.

10. Steinetz, B.M. and Dunlap, P.H., “Development of

Thermal Barriers for Solid Rocket Motor Nozzle

Joints,” NASA TM–209278, AIAA–99–2823,

June 1999, also Journal of Propulsion and Power,

Vol. 17, No. 5, 2001, pp. 1023–1034.

11. Steinetz, B.M., Kren, L.A., and Cai, Z., “HighTemperature Flow and Sliding Durability

Assessments of Hypersonic Engine Seals,” NASA

TP–3483, December 1994.
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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NASA TM—2002-211708

AIAA–2002–3941

E–13456

WU–706–85–33–00

30

Investigations of Control Surface Seals for Re-Entry Vehicles

Patrick H. Dunlap, Jr., Bruce M. Steinetz, Donald M. Curry, Jeffrey J. DeMange,

H. Kevin Rivers, and Su-Yuen Hsu

Seals; Flow; Design; Test; Re-entry vehicle

Unclassified -Unlimited

Subject Category: 37 Distribution: Nonstandard

Prepared for the 38th Joint Propulsion Conference and Exhibit cosponsored by the AIAA, ASME, SAE, and ASEE,

Indianapolis, Indiana, July 7–10, 2002. Patrick H. Dunlap, Jr. and Bruce M. Steinetz, NASA Glenn Research Center;

Donald M. Curry, NASA Johnson Space Center; Jeffrey J. DeMange, Ohio Aerospace Institute, Brook Park, Ohio;

H. Kevin Rivers, NASA Langley Research Center; and Su-Yuen Hsu, Lockheed Martin Space Operations, Hampton,

Virginia. Responsible person, Patrick H. Dunlap, Jr., organization code 5950, 216–433–3017.

Re-entry vehicles generally require control surfaces (e.g., rudders, body flaps) to steer them during flight. Control surface

seals are installed along hinge lines and where control surface edges move close to the vehicle body. These seals must

operate at high temperatures and limit heat transfer to underlying structures to prevent them from overheating and

causing possible loss of vehicle structural integrity. This paper presents results for thermal analyses and mechanical

testing conducted on the baseline rudder/fin seal design for the X–38 re-entry vehicle. Exposure of the seals in a compressed state at the predicted peak seal temperature of 1900 °F resulted in loss of seal resiliency. The vertical Inconel

rudder/fin rub surface was re-designed to account for this loss of resiliency. Room temperature compression tests revealed

that seal unit loads and contact pressures were below limits set to protect Shuttle thermal tiles on the horizontal sealing

surface. The seals survived an ambient temperature 1000 cycle scrub test over sanded Shuttle tiles and were able to

disengage and re-engage the tile edges during testing. Arc jet tests confirmed the need for seals in the rudder/fin gap

location because a single seal caused a large temperature drop (∆T = 1710 °F) in the gap.