SHUTTLE TPS THERMAL PERFORMANCE AND

ANALYSIS METHODOLOGY

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W. E. Neuenschwander, D. U. McBride, and G. A. Armour

Space Transportation and Systems Group

Rockwell !nternationol

Downey, California

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SUMMARY

Thermal performance of the TPS was approximately as predicted. The only

extensive anomalies were filler bar scorching and over-predictions in the high %p

gap heating regions of the orbiter. A technique to predict filler bar scorching

has been developed that can aid in defining a solution. Improvement in high Lp gap

heating methodology is still under study.

Minor anomalies have also been examined for improvements in modeling tech-

niques and prediction capabilities. These include improved definition of low ±p

gap heating, an analytical model for _ convection heat transfer, better _deling

of structure, and inclusion of sneak heating.

The limited number of problems related to penetration items that presented

themselves during OFT have been resolved expeditiously, and designs have been changed

and proved successful within the time frame of that program.

INTRODUCTION

The Space Shuttle orbiter thermal protection s}stem (TPS) is designed to per-

form at least I00 missions before requiring major refurbishment. This is i= con-

trast to the single-use designs for previous manned spacecraft heatshields such as

Apollo. Basically, the protection against severe ascent and entry heating environ-

ments is composed of coated carbon panels, silica insulating tiles, and nylon felt

blankets. The engineering task is further complicated where the system reaches

interfaces with the various mechanical subsystems that are essential for launching,

o,,erating, landing, and servicing the vehicle. These interface areas arf called

ITS penetrations or singularities.

An overview evaluation of TPS thermal performance during the OFT progrzm and

of the analysis methodology is presented in this paper. Results of this evaluation

will have value in the TPS thermal performance certification for more severe entry

missions and TPS design of second-generation Shuttle-type spacecraft. ITS thermal

performance assessments are based on postflight inspection results and measured ITS

structure and component peak flight temperatures. These are compared wit_ _ost-

flight developed analytical temperatures. Assessment of the analysis methodology

is based on detailed comparisons of predicted temperature histories with measured

data and evaluation of differences. Because their performance during the orbital

flight test (OFT) program has not received as much publicity as that given the

acreage tiles, the penetrations are granted some emphasis.

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https://ntrs.nasa.gov/search.jsp?R=19840002093 2019-02-08T00:44:53+00:00Z

Included are the following: a summary description of the Colmnbia TPS design;

thermal performance as evidenced by postflight inspections and development flight

instrmnentation (DFI) temperatures; and evaluation of thermal analysis methodol-

ogy. Results are interpreted ro reflect potential improve_ezts in TPS thermal per-

formance and analysis methodcio_ies.

ABBREVIATIONS

AFRSI

B/F

C/M

DFI

ET

F/B

FRSI

G/F

KRSI

IML

LRSI

OFT

OML

OMS

OV

PLB

RCC

RCS

R/SB

RSI

RTV

SIP

STS

advanced flexible reusable surface insulation

body flap

crew module

development flight instrumentation

external tank

filler bar

flexible reusable surface insulation

gap filler

high temperature reusable sucface insulation

inner mode line

low temperature reusable surface insulation

orbital flight test

outer mold line

orbiter maneuvering system

orbiter vehicle

payload bay

reinforced carbon-carbon

reaction control system

rudder/speed brake

reusable surface insulation

room temperature vulcanized

strain isolation pad

space transportation system

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T/3

TPS

h

h'

P

p,

P

P=

qs

TA

T A '

T o

TSK

TSTR

I

TSK

T t

ts/s

V

V'

w I

Z

a

_p

thermal barrier

thermal math model

thermal protection system

SYMBOLS

heat transfer film coefficient (Btu/ft2-hr-°F)

heat transfer film coefficient at time of vent opening (Btu/ft2.hr.°F)

pressure

compartment pressure at time of vent opening

local pressure (psf)

freestream static pressure (psf)

heat transfer rate (Btu/ft2-sec)

surface heating rate (Btu/ft2.sec)

tile gap heating rate (Btu/ft2-sec)

local air temperature (OF)

local air temperature at time of air vent opening (OF)

initial temperature (OF)

stractural skin temperature (OF)

structure

structural skin temperature at time of air vent opening (OF)

peak surface temperature with turbulent boundary layer (OR)

equivalent thickness of structure skin and stringers (in.)

compartment flow velocity

compartment flow velocity at vent opening

tile-to-tile gap width (in.)

dimension from heated surface (in.)

tile step dimension relative to adjacent tile (in.)

pressure difference across the tile

delta heatsink capability of ts/s due to nearby heavier structure (in.)

tile thickness (in.)

filler bar scorching parameter

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TPS GENERAL DESCRIPTION

The O_T vehicle Columbia, Orbiter 102, employs four basic thermal protection

materials: reinforced carbon-carbon (RCC), high temperature reusable surface

insulation (RRSI) tiles, low temperature reusable surface insulation (LRSI) tiles,

and flexible reusable surface insulation (FRSI) blankets. These have been dis-

cussed in detail in references 1 through 5. Figure 1 shows the location of the

acreage TPS materials on the orbiter and figure 2 shows typical penetrations. The

list of the penetrations given in table I shows that the TPS i_ more complex than

just tiles and FRSI blankets even though the penetrations comprise only a fractionof TPS surface area.

The basic TPS was designed to protect the primary structure from exceeding

350°F (the 100-mission temperature limit). The design was developed employing aminimum--weight engineering approach; that is, there was no intentional conservatism

in the entry trajectory, heating or thermal analysis methodologies. However, the

design was based on a single-orbit type mission that yields higher structure tem-

peratures at the start of entry. This would be the case for a once-around abort,but the TPS was not designed to an abort as is sometimes believed.

Establishment of and design to material temperature limits were of prime impor-

tance. These criteria include those for the basic TPS materials illustrated in

figure 3, and also cover about 50 other materials predominacely used in the pene-

trations. Typical limits are shown in table II. There are a number of exceptions

as illustrated by the following examples. AB312 ceramic fabric, used for gap

filler/thermal barrier covers, has a nominal 2000°F limit, but its use is acceptaole

to 2300°F and 2600°F for ten and one missions, respectively. High-density HRSl(LI-2200) has been certified to 2900°F by arc-jet testing for at least _ne mission

at a worst-case environment predicted in the gap between the elevens and the fuse-

lage. On the other hand, titanium mounting flanges on the reaction control system

(RCS) thrusters are limited to as low as 400°F to protect adjacent components.

The TPS and structure were instrumented with thermocouples and other

temperature-sensing devices to provide data for thermal performance verificatioD of

the TPS for operational missions more severe than OFT missions, Figure 4 shows

representative installations of the flight instrumentation plus number and loca-tions of temperature measurements.

Unique designs in the TPS are called penetrations or singularities. These

include the hatches and actuated doors, windows, aerodynamic control surfaces, RCS

and orbital maneuvering system (OMS) thrusters, main engines, overboard vents and

drains, umbilical connections, service access panels, and structural lifting and

attach points (see table I). Around them the TPS is co_prised of components made

from numerous temperature-resietant metals, ceramics, plastics, and elastomer8,

which have been formed, cast, extruded, woven, and laminated. Penetrations to be

discussed are located on the orbiter, as indicated in figure 2. They include the

external tank (ET) umbilical door, a payload bay door bare hinge, selected rudder/

speed brake components, the aft fuselage stub and body flap, one location in the

wing/eleven seal system, prinmry thrusters in the forward RCS, and the windshield.

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FLIGHT TEST RESULTS

OFT Entry Conditions

Example comparisons of surface temperature responses are shown in figure 5 for

a lower surface mid fuselage location. These comparisons depict trajectory effects

at the time of boundary layer transition to turbulent flow as evidenced by the

rapid increase in surface temperature in the I000- to i300-second time frame. As

shown, the turbulent spike temperatures occur earlier for each successive OFT

flight, and the turbulent spike heating rate is indicated to increase as transition

time decreases. This general trend is typical for the orbiter lower surfaces.

Acreage TP3

For the most part, the OFT heating environments were rather mild relative to

the capability of the TPS. This is illustrated by the STS-2 peak surface tempera-

tures shown by figure 6. STS-4 experienced higher peak turkulent temperatures, es

illustrated by figure 5; however, these are not shown because of limited data (not

available until about I000 seconds after the entry interface.) Peak temperatures

of the filler bars (F/B), strain isolation pad (SIP), and structure from STS-2 DFI

are presented in figure 7. All peak temperatureB are noted to be considerablylower than the 100-mission design allowables. These temperatures are typical of

those for all flights. A comparison of peak structure temperatures for all flights

is presented in figure 8. These temperatures arc about the same, _25_F fo= all

flights, and well below the 350°F, 100-mission allowable.

TPS Penetrations

It is not feasible to discuss flight data from all the instrumented penetra-

tions, so three were chosen for this section as being i%lustrative. They are an ET

door, a payload hay door hinge, and the rudder/speed brake conic seal and structure.

The ET is connected to the orbiter through two umbilical panels located on

eithe: side of the lower centerline just behind the aft fuselage bulkhead. Struc-

tural attachments, various propellant lines, and a number of electrical connections

pass through these panels. During ascent, the ET umbilical doors are held open bylatches located at the centerline. After the ET is jettisoned, the doors are

closed to protect the .umbilical panels. Each door is attached to the fuselage with

two bare hinges, which are flush with the outer mold line (OML). The outer surface

of the door when closed is covered with KRSI. A thermal barrier, which runs around

the periphery of each umbilical well, protects the structure from entry aerodynamic

heating and acco=modates relative movement between the door and fuselage. Pressure

seals between the doors and structure and flow restricters along the door edge

tiles complete the local TPS. The ET door/fuselage interface is depicted in

figure 9. The doors are made from beryllium for stiffness, reduced weight, and toobviate the need for insulation blankets, which would be exposed to destruction

during ascent. Even though the beryllinm could tolerate almost twice the temper-

ature, the thermally critical component for the doors is the 550"F KRSI/SIP bond-

line limit. At this location the thermal barrier consists of an AB312 cover and

sleeve with an internal Inconel (International Nick=l Company) X-750 knitted wire

spring.

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Flight temperatures for three OFI are shown in figure I0. It can be see_ tha:

the thermal barrier cover temperature neared its long term limit on STS-I and may

have been dropping from a peak above 1600°F on STS-4 when post-blackout data were

acquired. The HRSI tile DFI recorded temperatures well below the limit, but this

would be expected because it is located in a tile sidewall below the OHL. _ae doorremained relatively cool.

The payload bay doors are attached to the fuselage by 13 external hinges on

each s_de. The six forwardmost hinges, which experience the most severe he=ring,

are covered to protect them. The covers are made of Inconel and are designed to

permit the doors to open without any restriction on movement. Since the predicted

baseline aerodynamic entry heating was sufficiently lower, the remaining exte.-ual

hinges were left unprotected to reduce complexity. The gaps bet_,een the doors and

fuselage are filled with a si]ica fiber brush thermal barrier. This protects the

door structure and the environmental seal from direct convective heating and alsoaccommodates structural movement (see figure II).

S_S-I through STS-4 DFI temperatures for the most forward unprotected himEe

are presented in figure 12. One DFI is on the forward clevis facing the onccT_ingflow. Recorded maximums were well over the 1200°F criterion for this instrur_ent on

two missions; however, readings were erratic, and poor contact between the sensor

and clevis was suspected. Postflight inspection of the sensor indicated that it

was attached but the amount of contact area with the clevis could not be deter-

mined. Because it was uncertain if the DFI data actually were a true indication of

the clevis temperatura, authority was issued to remove pins on hinge 7 (maximum

temperature) and hinge 9 (lowest strecs margin), inspect the dry lubricant (limit

IO00°F), and run hardness tests on p_n 7. The results Indicated no degradation

caused by excessive temperature exposure and no loss of strength. In order to

eliminate direct stagnation-type heating on the DFI, a piece of filler bar was

placed over the sensor for 9TS-2 and STS-3 and replaced prior to STS-5. The

below-limit readings on these flights are more accurate measurements of the clevis

temperature. TemperatUres at the clevis/fuselage interface remained low and

changed little from flight to flight. The hinge lug is exposed co far less ae_-o-

dynamic heating, and its DFI is located in a cavity between tiles. These facts are

reflected in the much lower lug temperatures (compared to the clevis) plotted infigure 12.

As illustrated in figure 13, the vertical tail consists of a fixed structure,

the fin, and two moveable surfaces that have the capability of acting together as a

rudder or separately as a speed brake. At the hinge line a gap exists to perm/t

movement of the rudder/speed brake with respect to the fin. In addition to exter-

nal LRSl and KRSl, the structure must be thermally protected by closing the gaps

along the hinge line and around the rudder/speed brake perimeter with a aystem of

seals. The most visible element is the Inconel conic seal, which acts as a rub

surface for seals attached to the fin trailing edge and the rudder/speed brake

leading edge_ Other seals are found at the top and bottom of the cone as well as

between rotating sectors. Around the perimeter, there are seal requirements in the

gaps between the fin and the rudder, along the rudder/speed brake trailing edge,

and at the split line between upper and lower panels. The aforementioned seals are

made from several combinations of flat metallic springs, insulation-filled ceramic

sleeving, graphite blocks, and knitted wire springs.

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._here are several DFI on the conic seal. Readings from two of them are

plotted in figure 14. Both sensors are located on the inner face sheets of the

Inconel honeycomb. One is forward near the fin seal, and the other is aft but at

the predic,ed point of maximum interference heating when the rudder/speed brake is

closed during entry. Conic seal temperatures stayed well below the 1250°F limit.

The gentle rise in the aft sensor DFI data to only about 250"F indicates that it is

behind the =udder/speed brake leading edge seal rather than forward of it in the

more severely heated position described above. More rapid increases, although to

lower levels, are seen for the structure instrument located near the rudder/speed

brake trailing edge. Tnis behavior is as expected for metallic surfaces that are

directly exposed to a small amount of wake heating when the rudder/speed brak_ is

opened at Each i0.

TPS Thermal Anomalies

Poscflight inspections have shown every flight to have had some local anoma-

lies. Table IIl identifies the general types of heat protection anomalies that

have been observed during inspections. Figure 15 identifies the locations of those

that are most significant. Most anomalous conditions are easily detectable by

visual inspection of the OML; however, some require removal of access panels and/or

more detailed inspections.

Priority attention has been (and is) given to understanding and preventing

_no_alous conditions that require detailed inspection. Some design changes have

already been implemented to alleviate these problem_. Certain tile gaps near the

wing trailing edge have been filled to reduce air flow into the elevon cove, andthe internal insulation has been changed from FRSI to AFRSI. Designs of the edge

tiles of the gear and iT doors have been modified to restrict subsurface flow. In

some cases, repairs have been made to the existing design. An example of this is

in the bcdy flap stub plate tiles.

One anomaly of concern that requires detailed inspection is filler bar (F/B)

scorching between lower surface tiles. At several hundred randomly distributed

locations, filler bars reached 950°F to greater than 1375°F where 4_0°F to 500°F

were predicted. Filler bar scorching has occurred on every flight. A sugary of

occurrences and scorch severity is presented by table iV. Some filler bar scorch-

ing is known to be caused by structure leakage, e.g., around doors. These can beeliminated by design change or local repairs. Scorching in the general acreage is

a probability event depending on certain combinations of tile steps and gaps,

heating/bomsdary layer conditions, and tile-to-filler bar gaps. Eliminating filler

bar scorching in the acreage regions is a more complex problem.

Since nultiple reflights of scorched filler bars could result in structural

damage, degraded ones are repaired after each flight. They are mostly repaired by

filling the gap with room temperature _ulzanized (RTV) impregnated ceramic fabrics

(Ames gap filler). When severely degraded, the filler bars are replaced, and this

requires tile removal. The repair effort is relatively simple, but the problem is

that many thousands of gaps have to be inspected just to find a few hundredscorches that require repair. This inspection has been required because, uutil

recently, the cause wasn't q,_antitatively defined and step/gap allowables (cri-

teria) that would preclude filler bar scorching could not be developed.

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Continued examination of flight data resulted _ identification of a correla-

tio_ para=e_er that distinguishes be-_ween step/gap combinations that cause scorch-

ing and those that do not. The parameter is

= (Tt/1000)0.5 (A'5J_Wl'83)/-'_2_'5

Thi_ parameter correlated _ statistically representative sample of scorched

filler bar data to within +--20percent. The lowest value was 0.41 and corresponds

to a filler bar temperature of about 950"F which is she threshold of visual detec-

tion of a scorch per inspection criteria and NASA JSC plasma-arc test re_ul:s. The

_alue for 800°F, the 100-mission limit, was identified to t_ 0.38 from a combina-

ticu of flight and plasma-arc data. With these ar_iysis results, tile step/gap

criteria can be defined that will preclude filler ha= scorching.

Un@ortunately, implementation of an absolute no-scorch step/gap criteria has

only theoretical feasibility. Analysis of data indicates that about 20 percent of

scorches were caused (or highly influenced) by aft-facing steps on adjacent tiles.

Step/gap criteri -_ involving multiple tiles would be difficult to apply and execute

with confidence. Other practicality problems exist, l'nese include the fact that

future worst-case trajectory/_ea_-ing environments are required to define criteria,

but are not presently known, and inspection/repair for scorched filler bar is not

included in the planned orbiter t-_rnaround schedule_. At the time of this writing,

the best program solution to this problem is under study.

The body flap is attached to the aft fuselage stub and driven in its trim con-

trol function by rotary actuators. In addition to H]_SI tile_ that cover the exte-

rior of the stub and body flap, there are a number of seals th._t prevent direct

flow and resultant convective heating from passing from the P_igh pressure lower

surface to the low pressure wake region. There is the hinge seal, consisting of

Inconel 718 panels, which covers the entire width fr=u stub to stub, closing off

the lower cove. So-called chain seals block flow around the ends of the stub. The

design for thermal barriers between the body flap e_ tiles az_d the stub consists

of knitted wire springs within AS312 covers. All of the above are Or have ceramic

cloth components that bear on aluminrm heat sinks at:ached to each en_ of the fuse-

lage stub. These are known as the rub plates (see figure 16).

•A temperature sensor on the left-hand body flap rub plate exhibited a rapid

rise to 395°F after communications blackout during 5"/S-I entry (see figure 17).

This behavior was unexpected because the rub plate is heavy. The peak also consti-

tuted an overtemperature since the 100-_ssion al_m£=Bm material lilit is 350°F.

No other overtemperatu=es were recorded in the area, although the output of another

DFI on the upper end of the plate also rose and fell rapidly. Postflight visual

inspection of the exterior revealed no particular evidence of overheating.

Observed aluminum temperatures were used as drivers in the therDal math model (_M)

along with postflight predicted heating. Stress analysis pera_nnei used outputfrom this work and came to the conclusion that no structural degradatiom had

occurred during STS-I. A later visual inspection revealed gaps between the tiles

at the lower, outboard corner of the stub. It was s_-rmised that sneak flow entered

at these points, directly heated portions of the rub plate bounded by the hinge and

cS_in seals, and exited at a lower pressure area--per-haps the upper cove. For

STS-2, the _forementioned gaps, which were Lapered 1_-_-depth, _ere plugged with Ames

fil!e_s to preclude the flow.

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Again on STS-2, the rub plate DFI rose rapidly and went over temperature to

_5°F. Because the shapes of the t_perature-versus-ti_e curves -_ere almost identi-

cal for both flights, it va_ evident that the Ames gap fillers installed after

5T$-I did not cure the anomaly. Carrier panel tiles were removed from the left-

hand stub. As shown in figure 18, a section of charred and missing filler bar was

found. The corner of the rub plate adjacent to this section was discolored indi-

-sting excese[ve temperatures. _nere was also evidence of a forward-facing C_L

step between these tiles. Prior to STS-3, the missing filler bar was replaced, and

nearby pad gap fillers were rebonded. The stepped tile was not refurbished.

F_gure ]7 reveals that the left-hand rub plate temperature followed a repeat

history during STS-3. As noted, a new flow barrier wa_ added along with the pad

gap filler, and two lower surfa_e tiles were remade to remove the forward-facing

step. Finally, on STS-4, the maximum temperature was what ,had been expec:ed for a

heavy aluminum plate; however, the STg 5 temperature history, while not exceeding

the 350°F criteria, looks familiar. More extensive tile changes have been designed

to permit use of full-depth pad i'_ place of A_es gap filler, and these are being

installed during the modificatio,a period.

The elevons provide orbiter flight control in pitch and roll during atmos-

pheric entry flight. They are divided into two segments for each wing, and each

segment is eupported by three hinges (two _ixed and one actuator). TPe seal system

design must permit the required control surface motion ar.d yet pre.ent the destruc-

tive flow of hot plasma from the lower to upper surfaces through the wing/elevon

interface. Basic features are shown in figure 19, a section normal to the hinge

line. The lover surface is protected by sculptured _--_5°Iwith a nominal 0.5-inch

gap between wing trailing edge and el_von leading edge tiles. Yhe lower, internal

cavity is closed out by primary and redundant seal_. Spring-loaded polyimide

blocks bearing on aluminum and _nconel rub tubes make up the priv-_r7 seal while

reinforced silicone elastomer sheets form the redundant seal. Together, the

aforementioned penetration is known as the elevon lower cove. Also depicted in

figure 19 is the closeout to the upper wing-to-elevon cavity, co_noniy known as

flipper doors and rub panels. This TPS is insulated hot structure made from tita-nium and Incom-I on the inboard and outboard elevons, respectively. The wing/

elevon seal system takes on its grcates _- complexity when a chordwise interruption_ccurs at the inboard _ud outboard ends of each elevon and at each hinge and

actuator.

There have been some temperature-related problems with the elevon lower cove,

but none that brought forth safety-of-flight concerns. SI_o-I postflight visual

inspection of the closeout panel FRSI insula:ion revealed evidence of spanwise

and/or leakage flow within the coves. There were areas of scorching and charring,

particularly at the outboard elevons. This is illustrated in figure 20, which is a

Tiew looking up with the wing trailing edge and ele_on leading edge tiles removed.

Physical evidence revealed that flow entered the gap between wing and elevon tilesand exited near the stubs and actuators. After the first flight, the insulation on

the closeout panels was changed from FRSI to higher temperature capability AFRSI.

In addition, Ames gap fillers were placed between the tiles immediately upstre.zm of

_.he trailing edge tiles and thicker SIP was added to improve seating of these tiles.

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Some anomalous behavior of elevon lower cove DFI was notec, but they recordeo

no overtemperatures. T_me-temperature historie_ of the sensor on the primary seal

tube at 90 percent half span ave plotted in figure 21. The shapes have more pro-

nounced peaks than would be expected in pure soakback situations. This is ascribed

Lo the direct, low level convective heating that caused the FRSI degradation. The

heating is thought to have come about from the spaawise flow mentioned above and

from sneak circuit flow under the wing trailing edge tiles. A!ao note in figure 21

that seal leakage measurements in this region have generally declined. All in all,

it would appear that the elevon lower cove seal system on Columbia is currently

functioning as designed.

including the pay!oad bay door bare hinge overtemperatures, which were des-

cribed earlier as instrumentation installation problems, three anomalous penetra-

tions situations have been discussed. This points out that, considering the

hundreds of penetrations and singularities, there were a minimal number of

temperature-related discrepancies during the OFT program.

EVALUATION OF THERMAL ANALYSIS }IET_.ODOLOGIES

Assessment of analysis methodologies is being performed by ccmparing aua!y-

tical results with flight measurements. Discrepancies between flight and calcu-

lated temperatures using design methodologies are examined for cause. When a

discrepancy appears to result from analytical modeling, the technique is examised

for potential improvement. In the case of acreage TPS, the orbiter is surveyed to

assess data consistency of discrepancies and modeling improvements. However, for

penetrations the analysis is mostly unique to a specific penetration.

Acreage TPS

Nodal TMM's are used for flight temperatPre predictions. A design model for a

1.0-inch tile would have about 120 LI-900 nodes or about 240 LI-900 nodes for a

3.0-inch tile. Nominal thermophysical property data are used for all TPS andstructure materials.

Nodal models of tiles are three-dimensional and i_clude simulation of the SIP

and filler bars. Gap heating is applied to the tile sidewalls. For low pressure

gradient regions of the orbiter, the gap heating varies with surface heating rate

and dSmension from the tile surface. For high pressure gradient regions with open

gaps (i.e., wing glove, aft chine, forward fuselage, and elevons), the gap heating

is also a function ef pressure and pressure gradient in direction of the open gap.

High pressure gradients are defiRed as pl/2 dp/ds > 1.06 psf3/2/inch.

The structure is modeled to include the effective thickness of skin, and stringers

(ts/s) and the thermal capacity of nearby'heavier structure (i.e., frames). Honey-

comb structure is modeled as honeycomb. The structure inner mold line (IML) isassumed to be adiabatic.

The FRSI de3ign nodal models are one-dimensional with nominal thermal propertyinputs. Structures are modeled as described.

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Postflight analyses have been performed using these desig_ methodologies to

assess the TPS therm.al performance and the adequacy cf the analytical _Gdels for

flight certification. Although there have been some anomaliec the overall TPS

thermal performance and analysis edequacy appears acceptable. Peak _i_mperatures of

SIP and structure, as measured by '_he DFI, are well within design allowables. This

is illustrated by figure 22, whi._h also slows that calculated temperatures, using

actual flight surface heating, are in reasonable agreement with measured data.

These temperatures, shown for STS-2, are typical of those for all OFT flights.

Detailed examir_tion of the D_i and analysis data shows some discrepancies not

indicated by peak temperatures. Temperature history comparisons are shown by

figurp_ 23 through 27. These are rather typical results. Figure 23 illustrates

Lhat the high pressure g-adient gap heating is quite conservative. Calculated tile

side=,all, filler bax, anl structure temperatures are all well above the measured,

Gsp DFI in high _p r_gions is very limited. '.{owever, at se-:eral high __p ]ocaticns,surface and _tructure skin measurements are available. Analysis at these locations

shows similar conservative disagreement with structure tempe-aturea.

In low pressure gradient regiGns, the trend is for the temperature rise rates

of SIP and structure to be greater than calculated. But the agreement is reason-

ably close and the analytical conservatism of the cool down compensates for ths

rise-rate optimism. Figure= 3_ nd 25 illu=trate this observation.

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Figure 25 implies another anelysis anomaly that also sL_we up at other orbiter

loc_tions: the calculated in-depth reusable surface insulation (RS!)tempera-

turcs are conservative but the SIP and structure t_mperatures are o_timistic. This

could be interpreted to mean that the low gap heating in the analysis is optimis-

tic; however, figure 25 shows this gap heating to be reasonable until late in the

entry (to about 1200 seconds). The Li-900 temperature measurements are _uspect.

It ia speculated thaZ launch vibroacoustics resulted in the thermoeouples cavita-

ring the Li-900 such that they were not in contact with L!-900 at the thermocouple

junction. That this could happen was evideuced in a vibroacoustic tesL performed

ct NASA Langley where thermocouple wires were observed to have cavitated LI-900.

Othe_ explanations (e.g.j LI-900 conductivity variatio_ with temperature, hot air

ingestion at the base of the t_le) were considered improbable following analysis.

Examples of thermal response at upper surface locations Lre presented by

figures 26 and 27. Figure 26 shows the thermal reoponse at an L_Sl location. The

calculated SIP and structure rise rates are slightly lagging _leasured temperatures.

(Note also tLe contradiction by the LI-900 temperatures.) Fi[_re 27 shows the

thermal response at an FRSI location, where again calculated lemperatures are

lagging measured temperatures. Both the examples show good agrJ,_ment with peak

temperatures and cool down rates. This is typical of upper surface locations with

the thinner TPS.

In summary, the acreage _'PS design methodologies produce peak temperatures

that are conservati%e and adequate for Eastern Test Range fligh_ certifications.

However, this conservatism results from compensating effects of optimistic and

conservative thermal modeling techniques.

Modeling improvements are being evaluated to show better agraement with flight

_ata and include ftructure modeling, low pressure gradient gap he_ting, and LMI

con-.ection. Also, the tile models have a finer nod__ network of the SIP and filler

bar.

10..i,5

T

iiw

!

Z_

The design method of structure modeling was to use an effective structure

thickness that included both the skin stringer and At. The new method connects _

to ts/s with a contact resistance to increase the rise rLte of the structure skin.

Values were empirically established. The value for the lower fuselage is 1.0

Btu/ft2-hr'°F. For other locations except lower winE, the coefficient is

0.2 Btu/ft2"hr'°F. At time of this writing, the lower win E modeling is still under

study.

A pressure dependent, low pressure gradient gap beating model was devezoped

that improves agreemen= between flight and analysis data. The model is

q(=)/qs = (i.0 + p/20)l'4_-x (q/_s)des. extended

The lo4_CZexponent is truncated at a value of 1.85. The design gap heating

and design extended distributions are presented by figure 28. The pressure factor

gets to be a large numbe," for higher pressures and thick tiles. Even so, at later

entry times when the gap heating is negative, the pressure factor is not large

enough to match gap cool down rates.

The orbiter interior is repressurized by air vent air, which has a marked

effect on structure temperatures. Air vents open at 2400 fps orbiter ground speed,

which is typically about 85,000 feet altitude. From flight and analysis results,

hot structure (skin) is cooled and cold structure is heated by the ingested air.

A heat transfer convection model was developed from examination of flight and

venting analysis data. The modeling analysis was difficult to accomplish because

of lack of direct data. The heating/cooling is s function nf three parameters than

are dependent on each other and vary ,ith time:

- h (TST R - TA)

Nonetheless, based on a DFI air temperature measurement and a high density of

structure thermocouples at BPIS01, a heat transfer model was developed.

= •Air temperature (TA) is defined to vary linearly with time between compart-

ment air at the time of air vent opening and ambient air temperature at touchdown

at the landing site. Air temperature at air vent opening time is approximated as

!

T_ = TO + 0.I TSK

The film coefficient at air vent opening, h' , was empirically defined to he0.07 Btu/ft2-hr-°F for IHL-insulated structures and 0.10 for uninsulated structures.

Compartment pressure is a function of freestresm static pres=ure, as shown by

figure 29. Velocity ratio (V/V') is defined in the same figure for STS-_. Extrap-

olation of the STS-2 velocity parameter to other trajectories u_es the following

relationship:

v/v' = [_®l(_)szs-2](v/V')szs-2

1036

This extrapolation technique was used to calculate STS-3 structure tempera-

tures. It showed that the agreement was reasonably good (about the same as the

STS-2 data agreements).

Overall, the modeling improvements yield better structure temperature rise

rates and peaks. Also, the sharing of the SIP temperature response _nd peaks are

also better; however, in some cases there is very little improvement over results

from the design models. This is usually the case for the thinner TPS and locations

_here g_ is small compared to the structure skin-stringer thickness.

Assessment of these modeling improvements has been made by comparing analy-

tical and measured temperatures at 47 locations for STS-2 and 25 locations fcr

ST$-3. The agreement is quite good for the fuselage, 0MS pod, and vertical tail

low pressure gradient regions. Modeling of the wing TPS and structure is still

under study.

The agreement between analytical and measured temperatures is illustrated by

figures 30, 31, and 32 for the fuselage lower surface. Figure 30 is for a forward

fuselage location _here the IKL is insulated. Peak temperatures are in good agree-ment and the response shapes for the filler bar, SIP, and structure skin are quite

acceptable The deviation between the filler bar temperatures after the peaks is

caused by the convective gap cooling being greater than calculated even though the

pressure factor in the gap heating model is large. Figures 31 and 32 compare =em-

peratures at an uninsulated location. Again, structure skin and SIP temperaturesare in agreement. The calculated filler bar temperature appears to he conserva-

tive; however, it is not known if the filler bar thermocouple is located in an ope_

gap.

Temperature comparisons at upper surface locations a_e shown by figures 33 and

34. The calculated structure temperature on the vertical tail lags tLe measured

temperature by about 10=F, but the peak is good. The 0MS pod structure temperature

also lags the measured temperature by about 50°F, but the peak is in good agree-ment. On the O_S pod the cal:ulated cool down rate is greater than meacured.

In most cases, the modeling improvements yield better agreement with flight

temperatures than the methodology used for desiEn. This is particularly evident in

comparison of structure temperatures in regions of high heat load where the tilesare thick. In these regions, such as the lower wing, IML cooling helps to offset a

rather significant increase in structure temperature from heat soakback from the

tiles.

More modeling improvement analysis is in process. The structural backfaee of

the three-dimensional tile model could use improvement to better account for radia-

tive interchange between the upper _ud lower wing structure. Even with the m_del-

ing improvements described, the lower skin temperature rise rates are optimistic

and the cool down conservati_e- Another area of modeling improvement is with the

high pressure gradient gap heating since the design method appears to be very

conservative.

1037

Penetrations

The general approach to a penetration thermal analysis required identification

•"cf the physical location and configuration, applicable heating and other boundary

conditions, and subsystem operating _arumeters. To phy_ically represent the pene-

tration, the structural and TPS cross-section of the installation was subdivided

into a sufficiently fine matrix to represent it as a TMM in the computer thermal

analyzer program. Baseline models represented nominal design configurations.

Limit case heating and other boundary conditions were also identified and formatted

for computer input. These included various combinations of the following:

i. Initial temperatures either prelaunch or pre-entry after specified orbital

operation

2. Predicted ascent and entry aerodynamic nesting at all 0ML areas and within

the penetration

3. Ascent radiative heating from the booster and main engine plumes, andrecirculation convection effects from the latter source

4. local RCS and 0MS engine plune heating during system operation

5. External reradiation or r_diation interchange within the penetration and/

or internal to the orbiter

6. Mis< _llaneous system operating parameters (e.g., ablation, internal heat

generation, fluid flow, and cabin conditioning)

The complexities of the installations typically dictated the use of two-

dimensional lMM's based on the worst case location in each individual penetration

rather than more comprehensive models; however, limited three-dimensional models

were forced in a few situations, while other cases were simple enough to representin one dimension.

The basic TPS sizing was accomplished based on minimum insulntion thickness

requirements for each given orbiter location without regard for special situations

such as penetratione and singularities. Further design constraints required the

penetrations to be thermally controlled passively and independent of internal

systems for cooling. All in all, these ingredients imposed challenging design

conditions on these TPS features.

Postflight evaluations consisted of screenings that categorized results as to

their criticality. First priority was given to components that exceeded their

design temperature limit. Two exa_zples were discussed in previous sections: the

payload bay bare hinge and the fuselage stub rub plate and body flap. Second came

those that exhibited anomalous readings that were considered pzoblems even though

they stayed below criteria. Instances include the wlng/elevon lower cove, which was

mentioned earlier and will be discussed further in the following, and the RCS instal-

lation, which is covered in this section. Because of the limited time and resources,

the least attention could be devoted during the flight test program itself to DFI

reading_ that were overpredicted.

1038

|

The canopy _indshields have obvious functional requirements, serving as the

pilots view ports during launch, landing, and other operations. There are a totalof six windshields with each window assembly consisting of thlee high temperature

glass panes. Figure 35 depicts a sample cross-section. The outer pane of each set

is fused silica glass, which serves as a heat shield and is sealed to prevent hot

gas plasma from penetrating into the internal structures. The middle window pane

is also fused silica and serves as a fail-safe redundant member to either the outec

thermal _ane or the inner pressure pane. The inner pane is heat-tempered alumino-

silicate glass and is the primary pressure containing pane for the cabin at=os-

phere. All glass surfaces are coated except for those of the thermal (outer)

panes. Special HRSI (LI-900) tiles that overhang the outer windGw frames protect

the aluminum canopy structure. The HRSI is bonded to 0.09-inch SIP, which is, in

turn, bonded to carrier panels. These assemblies are mechanically attached to the

frames. Captive-type gap fillers are installed between the tiles. AB-312 cloth

filler, bonded only to the carrier panels, supports the overhanging portions of

HRSI. To prevent the inflow of hot plasma during entry and the possible leakage of

cabin pressure, thermal barriers and pressure seals were incorporated into the

window design. The external thermal barriers are made of woven ceramic fiber. For

external pressure seals, fluorocarbon O-rings are used. The crew module windows

(redundant and pressure panes) are contained in frames having steel, aluminmn, and

beryllium components. The crew module _indow frame is actively cooled by w_ter

flowlng in pairs of tubes around each window's inboard perimeter. Environmental

barriers having rigid and flexible portions are located between the canopy and the

crew module. At highly heated windshield io_tions, insulated heat shields are

required to protect the flexible purge barrier, in turn, the crew module wall

requires TCS insulation blankets to protect at f_on radiating surfaces.

Figure 36 shows DFI data and analysis time-temperature histories for a loca-

tion at the downstleam region of the middle windshield. On STS-2, the preflight

work substantially overpredicted the bondline DFI early in entry. A comparison

between prediction and data at the OML surface indicates a probable cause for thebondline error: the experienced aeroheating did not have an early high load period

but did display an unexpected peak. Heating rates were derlved from surface DFI

data and extrapolated to the window pane surface and the overhanging tile sidewall.

These were then used in the baseline TMM. As shown by the solid line in the

figure 36 plot, a much improved correlation at the bondline was achieved.

The situation regarding STS-3 was similar. Just as the preflight prediction

of aeroheating was better, so was the thermal analysis output. Again, an improved

overall match between predicted bondline tem?erature and data came from using

derived heating. Even though the bondline value is slightly undercalculated

(<25°F) at its peak and falls less rapidly than the data, the basic validity of the

TMM has been demonstrated.

The forward RCS provides attitude control and small velocity incremen_ trans-

lation from main engine cutoff during ascent until the entry interface. It

includes 16 radiation-cooled thrusters (14 primary and 2 vernier), made predomi-

nantly from co!umbium (di_ilicide coated except for the injector plate_), and fully

exposed _o aerodynamic heating. Surrounding HRSI tiles of the TPS have been par-

tially replaced by other materials. Insulation-filled metallic plume shields are

placed downstream of long scarf (primary -Z and -X and vernier) thrusters. High

I039

density (22 pcf) RSItiles form the TPS in narrow areas between the -Z, +Y, and -X

thrusters. Thermal barrier= serve to block gaps between nozzle exits and adjacent

TPS from ho: boundary layer gaseS. Other thermal barriers are requi=ed between :he

plume shields and surrounding tiles. Each thruster is housed within a cylindricaltitanium container that seals the internal compartment from the exterior environ-

ment. The design is complex. There are approximately 20 "maximum temperature

limits to be observed, many with several variations acsociated with position,

mission pha3e, or frequency of occurrence. The primary dowefiring (-Z) th=uste=s

experience the most severe entry heating. A section through the do_ustrea= edBe ofthis installation is shown in figure 37.

A DFI is located on the inside of the plume shield adjacent to :he thermal

barrier. The recorded sudden rise in entry temperatures on STS-2 was not predicted

by the baseline TMM employing postflight nominal aeroheating. This can be seen by

comparing the solid curve with the data symbols (filled circles) in figure 38. --he

response is little different than the STS-I experience reported in reference i.

Several rzvisions were made to the TM14 after studying thermal flight data from

the region and the postflight inspection reports. First, the as-built configura-

tion (thermal barrier recessed inside the OMI and rounded) and expansion character-

istics of the plume shield were considered. This exposed additional areas of the

sealing surface to direct _eroheating and altered the perturbation and gap heatiug

factors. Second, interpol_tion between the data from the closest surface DYI was

used to estimate actual reference heating. Third, the plume shield !_I itself was

added to the IMM at its design location. With the revisions, a much better pre-

diction of STS-4 data was obtained, as indicated by comparing the dashed curve to

triangle symbols in figure 38. Peak values are well matched as is the time-

temperature history following the peak. The early overprediction is =hought to 5e_caused by the shape of the derived aeroheating curve. The forward RC3 thruster _S

is considered sound.

Flight data from a DFI in the wing/e!evon lower cove was presented earlier(figure 21) al_Rg with a description of this penetration area. A series of TMM

updates have m_rkedly improved the ability to predict temperatures for future en_-y

environm_t_. Figure 39 shows that the baseline was poor at anticipating both the

peak and _hape of tZe recorded data. One of the first changes included correctinzthe se_o_ locat_on. Since it was obvious from postflight inspections that direct

con_ective F ea_ing was present in the cove, several low percentage values of OML

reference heating were applied to appropriate _ internal sections. This producedF_ea£etions that bracketed the data. Although the assumed levels of sneak flow

b+_=ing were slightly higher than that correlated from ground tests, they seemed

qe£te believable in light of the FLql degradation in the first OFT flight. Fin-

ally, an intermediate level of sneak heating was combined with cooling correlations

developed for air flow following the vent opening. As dc-monstrated by the solidcurve in figure 39, the validity of the current TMM is much better.

CONCLUSIONS

Over_ll, flight data show the ITS design and thermal performance to be quite

good and the thermal analysis methodologies adequate. Examination of physical and

analytical data has identified some elements of the de=ign and methodologies _hat

cuuld have been better if data had been available on which to base i_roveme_uts.

_040

a-

LESSONS LEARNED

Results revealed that some t_chnology inadequacies exist. Their identifica-"

tion could benefit future desig_ and analysis of spacecraft heat protection syste_-s.

Low "_p gap heating is a function of pressure level. This wasn't knc-eu before

STS-2. Ground test facilities did not have the capability to produce datz to ide=-

tify effects of pressure level and theory (engineering methodology) did not iden-

tify pressure level as a significant parameter. This knowledge could have affecte£

ITS dpsigu. Filler bar scorching is caused by abnormally high rates of gap heat-

ing. These high heating rates are pressure-level dependent and the significance of

pressure level is amplified with boundary layer disturbances caused by tile steps.If this had been known before the first flight, filler bar scorching would have

been expected and step/gap criteria for prevention identified; however, it is most

probable that these criteria would have not prevented about 20 percent of the fill_

bar scorches, and would not have el_ _inated detailed postflight inspection for the

scorched filler bars. As such, it is likely that this ignorance was to the advan-

tage of the STS program in its preflight stage. Now, however, gap heating sensi-

tivities to pressure level and tile step/gap dimensions are all recognized and shou!/

be considered in future ground testing of similar TPS configurations. This sug-

gests that consideration should be given to upgrading the flight simulati=n capa-

bility of plasma-arc test facilities.

Other inadequacies relate to the IML convection heat transfer phenozena. Per

interpretation of flight data, the governing heat transfer relationships deviate

considerably from theoretical expectations. The sir temperature and heat transfer

coefficients are predominantly influenced by the local structure temperat-ares (hot

and cold) with little remembrance of prior temperature/boundary layer history.

Also, the variation of heat transfer film coefficient with air density _ flew

velocity is considerably more pronounced than expected. These anomalies are quite

probably unique to the orbiter's air vent system and structural configuration;

therefore, it appears that IML convection cooling i_ more cemplez than theory

indicates but extrapolation of OFT data to o_.her STS entry enviroDmencs can be

justified. Howeverj applying IML cooling for design of new spacecraft warrants

caution.

Thermal analysis prediction capabilities are quite good when the envlroulents

are known. Analysis of flight data demonstrates this claim. Absence of flight

data is evidenced in attempts to understand discrepancies between flight and

analysis results. One example of this is not having adequate instrumentation to

measure pEessure gradients in high Ap regions of the orbiter. Assessment of high

,_p gap heating requires this information. More pressure instrumentation should

have been located in high _p regions of the orbiter. Also, more surface and struc-

ture thern_couples should have been placed in areas where local heating environ-

ments are difficult to predict (e.g., elevon spill regions and penetrations).

1.

2,

3.

4.

5.

REFERENCES

Dotts, R. L.; Battley, R. R.; Hughes, J. T.; and Neuenschwander, W. E-.: 5=ace

Shuttlp Orbiter Reusable Surface Insulation Subsyst_, Thermal Perforzman=_AIAA Paper 82-0005, Jan. 1982.

Curry, D. M.; Cunninghm, J. A.; and Frahm, J. R.: Space Shuttle Orbi=er

Leading Edge Structural Subsystem Thermal Performance. AIAA Paper _--00[=,Jan. 1982.

Dotts, R. L.I Smith, J. A.; and Tilllan, D. J.: Space Shuttle Orbiter Eerie

Surface Insulation Flight Results. Shuttle Performance: Lesscns le____--nec.

NASA CP-2283, Part 2, 1983, pp. 949-966.

Curry, D. k.; Johnson, D. W.; and Kelly, R. E.: S_ace Shuttle Orbiter iea_

Edge Flight Performance Compared to Design Goals. Shuttle Performance:

Lessons Learned, NASA CP-2283, Part 2, 1983, pp. I065-IU82.

Cunningham, J. A.; and Haney, J. W., Jr.: Space Shuttle Wing Leading _-_.dBe

Heating Environment Prediction Derived From Development Flight Data.

Shuttle Performance: Lessons Learned, NASA CP-2283, Part 2, 1983, p_.. I__i-IIi0.

1042

TABLE I.- SUMMARY OF TPS PEN_ETP_tTIONS

ORIGINAL FAGT. 35

OF POOR QUAL_'Y

Ty_e

Actuated door• and hatches

: L_ding gear -Payload ba T (•=Emerita)

s Crew hatches

s Structural vents

• Flipper doors

• External tank

• Star tracker

• Air data probe

Aero surface• as•essbliea

•Elevona

a Body flmp

a Rudder/speed brake

(•e_enta)

Ensine•

a M4in

eOMS

a RCS

Vents and d "ain_

Quantity

3

8

3

18

30

2

2

2

3

2

54

Type _ity

9External window _ne_btie=

Structural eTezen_ interface

area•

Aoteonas

Exposed umbilical connectors

Fixed panel• and door•

MiacellaneouJ s_nSulariticq

* Extez-n_l binges and

latchea

e HOt •tructu=e panels

e &ttacb. boilt, and jack

points

e Vertical tail Is•din8

edge

o _rsency access tele_ea

a Fauive air vents

c External _Jght

&

22

87

295

36

3;

315

TABLE II.- TYPTLAL TPS TEMPERATIFRE CRITERIA

Material

al_imm (2IXI, 6XXX)

Seryllims

_oroailicate &lass

Car•nit fabric (A&-312)

Cer_ic fibers, Iml_

(Dy_f lax)

(Saffil)

Colu=biu=, coated

Fluoroelastomer (Vitae)

Wickel alloy (In¢o_l 6IX)

_ylou (Nc_ex) felt

Polyimide/glaaa laainate

Silica liters, ri&idized

_teet (21-6-9)

(3X_)

Titanium (6AL--_V)

I Te=peratur_e

Li=it

TPS Elements ('Y)*

Structure bracket•, etc.

Heat sinks

HlLqI coating

I_I coating (lov_ a/_)

Theraml berri._r, gap filler, and

inaniat ion covers

Bstt i_su iat inn

Thenul barrier and gap filler

Hub seals, nozzles, and flanges

Seals

_Ot structure, aeal•, etc.

Filler bar

SIP (_or standard •ize tile•)

Isolators, bracket•, etc.

Stm_d_rd (O= 9 pcf) tile•

Higb--de_ait 7 (0 = 22 pcf) tiles

Vents, br&cket•, etc.

Hot at_cture. ]_racket•. etc.

350

I 10OO

2300

1200

2000

2600

20OO

5_!

5_

600

23001300

900

1200

t 8o0

*For I00 " " ; " "axJezone higher values accep:able for limited =xtsxoua cr by apecifi:

test•, lower waluea =pecifie_ in •o_e system applications.

I0_;

1044

ORIGINAL PAGE IS

OF POOR QUALITY

TABLE I!I.- TPS POSTFLIGHT ANOMALY OBSERVATIONS

Item

Tile elumpin$ (overCemperature)*"

Thermal harrier desradati_ e

Scorched FILSZ*

Structural overremperature

Tile e_olion t

Gap filler degradation*

Gap filler leakage

]urned elevou c_ve insulation

Scorched filler bars

Location Code _'_

A

B

C

D

E

Y

G

R

¢madou on

!ower surfaces

*Y_sily detectable by visual inspectiou of the _4L

_e£sr to Figure 15

TABLE IV.- S.CORCHED r'_LLER BAR

STS-2

STS-3

STS-A

Total

360

219

678

i

CAT. 1

113

130

73

238

CAT. 2

269

1%

117

226

fBec_mae of scorche_ filler bars

CAT. I:

CAT. 2:

CAT. 3:

Tiles

CAT. 3 _ed*

232 2_6

36 47

29 34

16 16

950 _ IIO0"Y

1100 - 1375"Y

> 1375"F

_onts

garly Lrana (gO_l_Z)

Yixst £11&ht

Eaxly tr-_s (IR]PO)

OREC._|AL PAGE

OF POOR QUALITY

REINFORCED CARBON-CARBON (RCC)

HIGH TEMPERATURE, REUSABL_SURFACE INSULAT{ON (HRSI)

LOW TEMPERATURE. REUSABLEiaL.¢.£.._-.¢.._1SURFACE INSULATION qLRSI)

CO4,TED NOIWIEX FELT REUSABLE

SURFACE INSULATION _FRSI)

METAL OR GLASS

Figure I.- Thermal protection system (Orbiter I02).

RUDDER/SPEED BRAKE

,PAYLOAD BAY DOOR "_J_/_#./HINGE LINE " ,_7"/ /

_INC_ELEVON ET UMBILICAL DOOR

AFT FUSELAGE/BODY FLAP

Fig_mre 2.- T_pical TPS penetrations.

¢

|

t

1_46

lqLE SYSTEM

/Ts

TSIp

FRSl

Tt

ORIGINAL pA.GE _SOF pOOP, QUAL|I'_'

.'_.LES:

_L = 2300=F HRSI1200°F LRSI

Tt = I050°FTFI B = 800°FTSI P = 550°F TYPICAL

600°F MINI TIL_

TFIB _ FRShSTRUCTURE Ts = 750=; ASCENT

700° F ENTRYTRTV = 550"F

T S

/ //T v

PRIMARY STRUCTURE: 350='F

/INTERMEDIATE BOND STRUCTURE

Figure "3.- 100-mission temperature allowables (acreage).

/_IIEAGE

INIT_IJLilON

lu_IrAcE TEMINERATUREB

SlalIF_F.,JSI(IN _TVREShiTIIUCTUII_. TEIIPID_ATUW.£

tT1EIIIIAL _llU_-'TUI_

TUPERATUIIEQBCTIDI-'f_LE PLUG

LOWIER IURFACEulqqER SURFACE

HEAT W41EL_INDIEWALL T1DIIqERATT/RE

11JE '_.,_l_r MUIAY

TR 4JUNQ EDOI_HEAT

Nm

tNI_'ALLATIO_al

IT[11

I1

2$

1

:fl

T

Bml_

1 I1

S

$4i

2SS

liW

12

1|IUJ

I

TANK ATTAC_IlJ(r_tCTION CONllmL SYSTEMlmeoow_

ANL m/MPPANELS

I_YI.OAO BAY DOOM

2S

73

_2

T

TOTAL m

.: ;_ j

F_UU_

/

5URF/_Ef

Figure 4.- Development flight instrumentation.

71

ORIGINAL PAGE

OF POOR QUALFI'Y

16@0F ---:1_ -+/

_,,ooF1200!-

1000- __STS-2800 - ...---STS-3

_STS-4

6OO

4o0

!O0 500 10001500200025003000

"lIME (SEC_

Figure 5.- Representative lower surface

tempera _-_areresponse.

+ _ - Lo._..-. I I

/_. " "- • '+0 W t_

; o FF_gure 6.- STS-2 pe_ surface temperatures ( ).

ORIGINAL FA,_E IS

OF [<_OR QUAUTY

VVVlWWWIXXYJYYYIZZZ=STS-1, STS.2, STS-3, STS-4, STS.5

w, S:_ww_TARBOARD

TOP

!t

POR1 I STRUCTURE MAXIMUM I BOTTOMI TlUdPERATURES WELL

L BELOW DESIGN UMITS

Figure 7.- STS-2 fltght temperatures (°F).

F_LLER BAR. 220 _ _ - --

I/ -I /s,-_ -- ,, ...r

_,,,-".. I ._',,;.5_//_

SiP _ _$7S,KIN. _lS't

Figure 8.- Structure maximum temperatures (°F) •

048

i

,mm

ORIGI'."I#.Lp_-,.gYi]

OF POOR QUAL |''r'Y

j% L,[

,---- _. J ) .ALUMINUN

/ AB-3 I2 SLEEVE//_-'_INCONEL X'/'_ RETAINER

22 PCF HRSI t / KNII"rED WIRE SPRINGAB312 COVER

Flgure 9.- ET/orbiter umbillcal door.

FORWARD_ _DOOq

TILE f _""_ THERMAL

{S_OEWALL) -- .BARRIER

FLIGHT DATA

ST,S-

LOCA_ON

THERMAL BARRIER

TILE

' DOOR

1 2 3 4

o '-" 0

• • O •

© r © *.

1 ."4)-MISSIG N

TEI_I)ERATU RE LIMITS

" TEMPERATURE (°F'}

1600

®

0

0 []

0 5O0

__I/HRSI TILE

_THERMAL

BARRIER

(COVER)

Qj DOOR(BERYLLIUM}/TILE BOND

1000 1500 20OO 2500

ENTRY "riME (SEC)

F!_ure i0.- ET door temperatures.

"_ _ •Xo=gS6 OF POOR QUALI"I'Y

/ _ ,.CLEVIS/FUSELAGE INTERFACE

DOOR

Figure I].- Unprotected PLB door hinge.

LUG

_',_111 _ \ FORWARD_ CLEVIS

• FUSELAGE

FLIGHT DATA

STS-

LOCATION

CLEVIS

LUG

FUSELAGE

1 2 3 4

o _ <> A

• n • •

IO0-MISSIONTEMPERATJRE LIMITS

TEMPERATURE(=_

1400

1200

1000

80O

6OO

4ool

00

o

-_HINGE (CLEVIS AND LUG)

A "" O

r:,

O o

oO

O

O

FUSELAGE _ _'_II_

STRUCTURE -_=e • i1-'-_

400 80O 1600 200C

ENTRY TIME (SEC)

2=k.'J 280O

Figure 12.- PLB door bare hinge temperatures.

1350

FIN BOX_

LEADING EDGE

INIC SEAL

_RY

ACTUATOR

(TYPICAL)

A--A: CONIC SEAL DETAIL

POWER

DRIVE

UNIT

_F'T ATTACH

FORWARD ATTACH

hER/SPEED

BRAK.--S _R/SB)

FtLLER BAR

B--B: TRAILING EDGE DETAIL

Figure 13.- Vertical tail.

AF _

°""_"'Vf\--fD

_ CONIC SEAL

.,s.s..uc'ru.E\ TEMPERATURE ('_F)

\ 1250 CONIC :; _'AL

80¢-

FLIGHT DAT_

ST_

LOCATION 1

SEAL, FORWARD o

SEAL, AFT •

STRUCTURE

10(TMISSIONTEMPERATURE LIMITS

60¢

2 3 4

o O a

Do

II58 STRUCTURE_

2oc _11"[]

il ,& (D[]

o

r

44, _]--'_ --

160G 2000

Figure 14.- Conic seal and rudder/speed brake structure temperatures.

1051

.

ORIC=iNAL PAG_ 13

__ OF POOR QUALITY' --C- " . ' • _ L L_r__j_ _._ _,j,_:,_.__._.

D / _'i -'- RIGHT-HAND SIDE

_" A AT T-SEALS

X "" _o

'K3_< _ .OWERSURFACE

\

k-\\

x_ iN

\ " /,,

3, F

Figure 15.- Locatiens of significant TPS degradation.

_:_- __ I o_'__ /// //

F_EL.AGE _ _.__ _.._...../BOOYF_LAp I H

LO_ER _ /.._ r_ :_f LOWER SURFACE I II,UR_A_E--..__--_._ t', | Ii

_'_jS- // | ALUMINUM _

SEAL DETAILS JJ / RUBPLATE \\ ._

i _ AFT

TPS TILE DETAILS

Figure 16.- Stub and hod7 flap.

1052

FL;GHT TEMPERATURES]--_ STS-1 |

STS-2 |_ STS.3 |

_-- STS-4 |

..... STS-5 .J

400L

F (LIMIT _C__-"

_- I

P

12t:_ 1600 2000

ENTRY TIME (SEC)

OF POOR QUALITY

POSTFLIGHT CORRECTIVE ACTIONS

• STS-1

• ADDED AMES GAP FILLER BETWEEN STUBTILES ON LOWER SURFACE

• STS-2

• REMOVED CARRIER PANEL TILES (SIDE OFSTUB)

• REPLACED CHARRED/MISSING FILLER 3AR• PAD GAP FILLER REBONDED

• STS-3

• REMOVED CARRIER PANEL TILES• REMADE TWO LOWER SURFACE TILES TO

REMOVE FORWARD-FACIN3 STEP• ADDED AB312 ROPE FLOW BARRIER, NEW

GAP FILLER AND RTV

• STS-4• NONE

• STS-5

• CARRIER PANELS OFF AWAITING TILE ANDGAP FILLER MODIFICATIONS

Figure 17.- Body flap rub plate anomaly.

Y

Figure !8.- STS-2 postfllght cond_tlon of Euselage stub.

1053

OR_C._I4AL p_%_E _;3

OF pOOR QUALITY.f--

_.,Lf \ - " TRAILING EDGE_H$1 ,.-_ \ ", _ SPRING SEAL

"-- --'__"--_ p/' RUB PANEL

L L 7"C'_ .......... )FRIMARY SEAL /

Figure 19.- Elevon seal system: design configuration.

Figure 20.- Elevon lower cove close-oun panels

after STS-I.

1054

J_

,q

C._(_:f_:L _,','.-_ !_; _ iPREFLIGHT LOWER COVE SEAL LEAKAGEIOF PC%.R QUALI'I'_t I/- _ _L L_

• _ l 1/ s_, !I_ _..-_ J 1%_ s'rs-2 I

WING/ELEVON SEALS \SEAL TUBE I (( STS-3 _'

. 250 I- //STS.3 LEAKAGE RATE (_CFIi)o I o o i/o .sxs4_" I o o Zr_'/

/ ° ,, / o/ •pOSTsTs-,ANDSTS2ACTIONS200_- • 4g / o o INVOLVING WING TRAJUNG EDGE

"" / ^ A ^ /o TILES HAVE REDUCED SNEAK FLOW/ o _ - .. - HEATING IN COVE:I - _ ,E' • B • AMES GAP FILLER INCLUDED

_J l v o )m - AHEAD OF TILES

I- 150L ."- • / _, • a _ • SIP ADDED TO IMPROVE FTT| _ / • GENERALLY, LOWER SEALS LEAKAGE

/ o STS.S/ ATI'RIBUTED TO 'SEATING' OF SYSTEM

/ [ FUGHT DATA1 AND MINOR POSTFUGHT TOUCHUPS4_ J • RESULTS OF ABOVI=- LOWER COVE

lOOJ_ • TEMPERATURES

__, , _ ,1so0 2SooENTRY TIME (SEC)

Figure 21.- Trends in the elevon lower case.

F tltl.

YOOLYYY =

SKIP. :_¢/_1SKIN. 113.1 It/

Figure 22.- Comparison of STS-2 flight and postflight

analysis temperatures.

_P

4

A

4056

ORIQINAL PA"_E;_PEAK SURFACETEMPERATURE OF POOR QUALITY"LAMINAR: 1440°FCrURBU LENT: 1010eF

(820 SEC) I1360 SEC}

° // "° ° -

1000 O O I u_m _F_

/- / J-_ / \o\\ ,._o:ou..,E

l /_ /'_. '_'__ _'_ \ o _S,N. FROMOUL0.75 IN." /// \_ \ . FILLERBAR

| I J ,_ _ "_O_"'Z_._...,_ VERY CONSERVATIVE

I t / f :_,_ ,TTHiSLOCATION

"-!')1io KCL_#. , __,..- I - _ ,

0 5OO 1000 150_ 20O0TIME FROM El (SEC}

Figure 23.- Comparison of design method with STS-2 flight

temperatures (high Ap location).

1200 -

I000

ALA.o

-, 800

¢ 600wD.

Ill)--

2OO

00

PEAK SURFACE TEMPERATURE

LAMINAR: 1400°F I TURBULENT 12t50_F _801

(600 SEC) ('t300 SEC)

"_ d_l = 1_5 IN.

500 1000 1500 2000

TIME FROM El (SEC)

PLUG THERMOCOUPLE DATA

o 0.25 IN. DEPTH

A 0.50 IN.

o SIP

O STRUCTURE SKIN

• LI-900 G't_RPREDICTED,

SiP UNI)ERPREDICTTr-9 --ALSO STRUCTURE

• PREDICTED SIP AND

STRUCTURE TEMPERATURESLAGGING MEASURED DATA

Figure 24.- Comparfson of design method with STS-2 flight

temperatures (plug thermocouples).

.

¢

1400

1200

1000

800

_. soo

1-

4O0

2O0

ol

ORIGINAL PAGE TZ

PEAK SURFACE TEMPERATURE OF POOR QUALITYLAMINAR: "I4OOaF TURBULENT: 1260 °F

(600 SEC) (1300 SEC)

o O riP1801 __dRSl = 1.05

DoLOW(R SUfl c

GAP THERMOCOUPLE DATA

-_--0.25 IN. DEPTH

-" 0.50 IN. DEPTH

7 FILLER BAR

STRUCTURE SKIN

• LOW 6P REGIONBETTER AGREEMENT

• IML CONVECTION EFFECTON SKIN TEMPERATUREEVIDENT

500 1000 1500 2000

TIME FROM [] (SECI

Figure 25.- Comparison of design method -_ith STS-2 flight

temperatures (gap thermocouples).

PLUGTHERMOCOUPLE

DATA

400 o 02_5 IN. DEPTH

[] SIP

O STRUCTURE__ o o o o_

2OO _

0

t I I- 20O 0 500 1000 1500

TIME FROM El (sEC)

dRS I = 0.45

-c q o n O

2_O0

Figure 26.- Comparison of deslgn method with STS-2 flight

temperature (LRSI location).

".1357

g

1058

700

600

/#BP734 Jr// _

_J.._L._ OFRSI= u.o._

500 -_,c_ o THERMOCOUPLE DATA

o SIIPFACE400 "" STRUCTURE SKIN

3oo=,I--

100

0 :_I"°'°_'°

o 50o 100o 15OO 20O0

ORIC._iWJ:,LP.'-.3_ ::.OF POCR (_L_,.,_.

TIME FROM El (SEC)

Figure 27.- Comparison of design method with STS-2 flight

temperature (FRSI location).

1.0 ._ /GAP WIDTH

. 0.100

o_o= \\W_ 0"°_" .030

DESIGN0.01

h

o.o01 _

L_ EXTENDED.o.ooolI ......... , To_.R?...;,, ....... ,,

0.01 0.10 10 3.0

z pN.)

Figure 28,- Low pressure gradient gap heating distributlcn.

ti

,#

t

#

n = 1.6 FROM CORRELATION OF STS-2 DATACR:_r_..- - - - : :

OF POOR Q_ALiTf

h' 0.07 INSULATED STRUCTURE J i" STRUCTURE DATA

VIV' FROM INTERPRETATION OF FLIGHT DATA

I STS-2 AND STS-31

,.0F

PIP oo 0.90.8

0 4 8 12 14 16

poo (PSIA)

,.0\0.8

vN'08- ____7,

0.4

0.2

0 0 200 400 600

TIME FROM AN OPENING (SEC)

Figure 29.- Preliminary IML convection data.

W

O.

F,LLER.A_. ANALYS,S-..FILLER BAR. FLIGHT, _/r _ _ /SIP, ANALYSIS

2:0 ;,y' _P, FLIGHT

100 /t, _./_'_STRUCTU RE, FLIGHT

0 _ ""-STRUCTURE. ANALYSIS

_=,o_-_- ,_ ,_ _,;,v_ _____ToME FROM ENTRY !NTERFACE (SEC)

Figure 30.- Comparison of _mproved method with STS-2 flight

temperatures (lower fuselage, LML insulated).

#

!

p_GE _9Om _tAL QU AL_'I'Y

OF FOOR

1600

1400

1200

_- 40(]

2O0

0

GAP TEMPERATURES BP1801 _,//_T_\ .# I_,

0.5 IN DOWN, ANALYSIS

l 0.5 IN. DOWN. FLIGHT

_\ /// ..t- 0.25 IN. DOWN. ANALYS;S

_JL/ FILLER BAR ANALYSIS

__ ;/ /''"..,/// -- ._._._FILLER BAR. FLIGHT

40O 8O0 1200 1600 2000 2400

TIME FROM ENTRY INTERFACE (SEC)

Figure 31.- Comparison of improv=d method with STS-2 fli_ht

temperatures (lower fuselage, IML not insulated).

700

" FT,O0

500-

v

_FILLER BAR, ANALYSIS

'" / FILLER BAR, FUGHT

200 _" / j_/ .... _ prm_,_4.____ __

11111 U

-so l I : _;[_v 'IJVlTD j ,

0 400 800 1200 1600 2000 2400 ,,B30

TIME FROM ENTRY INTERFACE (SEC')

Figure 32.- Additional comparison of improved method with S'_S-2

flight temperatures (lower fuselage, IHL not insulated).

1060

100

90

80

ORIGI;L_,L _:..,iL_T-._

OF PC'CR C_L,;AL:'I-I"

_::fj,,$.40.

STRUCTURE- FLIGHT

_. 70%.'" ;0

!,o4O

30

2O

10

// / 'STRUCTURE AN,.,LY._IS

i///

t , 28000 0 400 _ 1200 1600 2000 2400

TI&IE FROM ENTRY I_TERFACE (SEC}

Figure 33.- Compar _-_='_aof imprcred method with STS-2 flight

temperatures (_ location),

P

l

,j

200r-

i.Lov

telg¢

I,.-

100-

=EUJ

_'_o_

I°I

-200

" //'_ % ./$TRUCI_R_ FUGHT

/I \,,\// \ "-...

/ / \ "--.

// / _QRE" ANALYSIS/

ii / sFr_

T;",',',',',',',',',_-FROM ENT'R', NTERFACE (SEC)

F'_ure 3-'.- Compar _ of impr=-"=-_ method wi._h STS-2 flight

t_atures (.r:SI location).

1OE"

i

i

J

J.i

J.i

J

1062

ORIG;NAL PAGE F_

OF POOR QUALFFY

I HRS4(Ll.gO0)

\

(LI'2_O) | TILEOVERHAN G

,SIP Ii__ _'.LLER

r _I _,.,..- ENVIRON MENTAL

t___r_=_j ) _ARR_ER

/_M "-" v-- ill REDUNDANT PANE

STRUCTURE/''_ . )

'_'_--r_l _ PRESSURE'PA, I'E

ATCS COOUNG TUBES

Figure 35.- Windshield design configuration.

F""_,,,,...._ _.-- S U R F ACE .

_ooo, %. "7".. _SONOUNE

75O

/ n •

, o x-,f soo I- t

i I I o ° _c/ I - ...- -L

|I ,,/ J I*"Ii ..L.... _ _ .-"b _'C.., -''_ ! /°

ok-'w"'_ _' z , ",_- ......J0 500 1000 1500 2000

ENTRY TIME _EC)

, II

SURFACE |

BONOUNE I C'_

50 ]_l--,

PREDICTION 1 | --_--- PREFLIGHT j _ |_WITH DERIVED HEATING -

_ I

o_

a _f---"4 _

/ :

/ := t

; ...-;.---\/o.ENTRY TIME ($EC_

Figure 36.- Windshield modeling improv_r_ents.

.-,.A

SHIELD '"

FWD

1 _.J SIDE PANEL _ pA-i<:L HRSI

[ FRCSMODEL_ " T,LETI':ERMAL

6ARRIER S.IP_, PLUME r

SH,E-_cup -___THERMAL .... _,.--_""_C"_-'_',_,, BARRIER, , _-_----.. _'_,j_ _\

/_ _k_.Z._L_q_. \ - COLU'.._z " _ _ - RING

NOZZLE" _'__ - FLANGE

-- S_. ;I-- VALVF.J\ ' MOUNTING

SECTION A-A INJECTOR / PLATE

°,,

Qu_L/T-

Figure 37,- FRCS-Z thruster design installation.

900

/

7s°I- / ,_ •/ /

j v 'i

_--'"__600- /f

o ./450 •,ez,,,

- /300

/

15C _ _ v

_i _ l I

ENTRY TIME (SEC)

l BASELIN=_']

-- PREDICTION: DESIGNCONFIGURATION (SEEFIGURE 37) TMM

• STS-2 DFI DATA

IIMPROVED I

FLUI_E SH_

EXPOSED SHIELD_ _,_,_'--N'_'_a=ed

NOZZLE /

--- -- PREDICTION: _,.'_BUILTCONFIG UP_,TION. MOREHEATED t mE.A. CORRECTED.01=1LP'C . ")N

"7,STS-_ OF :,'rA

Figure 38.- FRCS thruster modeling improvements.

ORI_NAL p._E :.SOF POOR QU_U,TY

35O

3OO

A 2S0;

WJ

a¢

,¢

uJo.

1sok-

100

50

l'%• -% , 4

0 %I o

I!

O 0 0 O! []

O

,' ,..____°._.._o

I =/f"-'_ ^ ,_o

// -/" o STS-1/_ / o STS.2

_.,./_'. osTs-3"_. o'_ _ STS-4

• • STS-5

I I I i I500 1_."10 1500 2000 2S00

ENTRY TIME (SEC)

SEAL TUBE

WlNGIELE"ON SEALSg0% HALF SPAN}

ANALYTICALPPEDICTIONS

BASELINE TMM, STS-2PREFLIGHT NOMINALHEATING

CORRECTED DFILOCATION, ADDED 0.2%OF OML HEATING TOCAVITY FOR SNEAK FLOW

INCREASED SNEAK FLOWHEATING TO 1.2%

UPDATED TMM, 0.5%SNEAK FLOW HEATING,INCLUDED POST-VENT

OPENING COJDLING

Figure 39.- Elevon lower cove model_ng improvements.

_064