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DEPARTMENT OF MECHANICAL ENGINEERING AND MECHANICS

COLLEGE OF ENGINEERING AND TECHNOLOGY f 13_< i

OLD DOMINION UNIVERSITY / C/ J

NORFOLK, VIRGINIA 23529

THERMAL STRUCTURES: FOUR DECADES OF PROGRESS

By

Earl A. Thornton, Principal Investigator

Final Report

For the period ended December 31, 1989

Prepared for

National Aeronautics and Space Administration

Langley Research Center

Hampton, Virginia 23665

Under

Research Grant NSG-1321

Allan R. Wieting, Technical Monitor

SMD-Aerothermal Loads Branch

DEC^Dr2, L__ PI,:_'_'LSL tin_l [_,2ort, period

enGin:r] _i Oe_.:. !:_9 (LJl-J _ominion Univ.) CSCL 20K

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N90-2_I05

GJ/ q

August 1990

https://ntrs.nasa.gov/search.jsp?R=19900018789 2019-01-28T13:39:24+00:00Z
DEPARTMENTOF MECHANICALENGINEERING AND MECHANICS

COLLEGEOF ENGINEERING AND TECHNOLOGY

OLD DOMINION UNIVERSITY

NORFOLK, VIRGINIA 23529

THERMAL STRUCTURES: FOUR DECADES OF PROGRESS

By

Earl A. Thornton, Principal Investigator

Final Report

For the period ended December 31, 1989

Prepared for

National Aeronautics and Space Administration

Langley Research Center

Hampton, Virginia 23665

Under

Research Grant NSG-1321

Allan R. Wieting, Technical Monitor

SMD-Aerothermal Loads Branch

Submitted by the

Old Dominion University Research Foundation

P.O. Box 6369

Norfolk, Virginia 23508-0369

August 1990
A_OWLE_MENTS

m

This AIAA paper which was presented at the AIAA/ASME/ASCE/AHS/ASC, 31st

Structures, Structural Dynamics and Materials Conference, Long Beach, CA is

being submitted in lieu of a final report for the research project "Finite

Element Methodology for Integrated Flow-Thermal-Structural Analysis." The

work was supported by National Aeronautics and Space Administration, research

grant NSG-1321. It was monitored by Mr. Allan R. Wieting, SMD-Aerothermal

Loads Branch.

iii
THERMALSTRUCTURES: FOUR DECADES OF PROGRESS

EarlA. Thornton,Professor

Departmentof Mechanicaland AerospaceEngineering

Universityof Vi_'ginia

Charlottesville, Virg.,nia22901

Abstract

Since the first supersonic flight in October 1947, the

United States has designed, developed and flown flight

vehicles within increasingly severe aerothermal

environments. Over this period, major advances in

engineering capabilities have occurred that will enable the

design of thermal structures for high speed flight vehicles

in the twenty-first century. This paper surveys progress

in thermal-structures for the last four decades to provide a

historical perspective for future efforts.

Introduction

The design of structures for winged flight vehicles

that fly through the earth's atmosphere, either to and from

space or in sustained flight, poses severe challenges to

structural designers. Major components of the challenge

are to select materials and design structures that can

withstand the aerothetmal loads of high speed flight.

Aerothermal loads exerted on the external surfaces of the

flight vehicle consist of l)re.-sute, skin friction (shearing

stress), and aerodynamic heating iheat flux). Pressure and

skin friction have important roles in aerodynamic lift aatd

drag, but aerodynamic heating is the predominant

structural load. Aerod.vnamic heating is extremel.v

important because induced elevated temperatures can

affect the structural behavior in several detrimental ways.

First of all, elevated temperatures degrade a material's

ability to withstand loads because elastic properties such

as Young's modulus are significantly reduced. Moreover,

allowable stresses are reduced, and time-dependent

material behavior such as creep come into play'. In

addition, of course, thermal stresses are introduced because

of restrained local or global thermal expansions or

contractions. Such stresses increase deformation, alter

buckling loads and change aerothermodynamic behavior.

Today, the advent of the National Aerospace Plane

offers structural engineers new challenges for the design of

thermal structures for high speed flight. A theme of this

paper is that the progress in research and development of

thermal structures and related technologies over the last

four decades provides the fouadations to meet the new

challenges.

One justification for a survey paper is expressed

very well by the opening of the "House divided" speech

{given by Abraham Lincoln in 1858, "If we could first

znow where we are, and whither we are tending we could

better judge what to do, and how to do it". The speech,

iven to a political convention in Springfield, Illinois in

une of that year, referred to an alltogether different

subject, but surely the words apply in the present context.

As our nation looks to high speed flight vehicles for the

next century, we should review past efforts to provide a

historical perspective for the future.

Aerospace technology has advanced far in a very

short time. These advances are particularly notable for

high speed flight vehicles. This paper attempts to provide

a historical perspective for thermal structures by

describing the evolution of thermal--structures technology

from the early 1950's to the technology of the early 1990's.

The early 1950's was selected as the reference point

because the first manned supersonic flight in 1947

stimulated a period of intense research on high

temperature structures. We are fortunate today to have

access to the paNts, reports and books that describe

research of that era. Unfortunately, although many

researchers of that period are alive today, much of their

personal experience is not available to younger researchers

because of the "generation _ap" that currently exists in the

aerospace establishment. Thus for all of us interested in

thermal structures, there are lessons to be learned from

surveying the progress of the last forty years.

Evolution of Thermal Structures

From World War II to Soutnik

The need to understand aerothermal loads ,_nd the

design of thermal structures have their origins in the late

1940's. In World War II. airplane speeds had become high

enough for compressibility phenomena to have a significant

role in performance. Tr,m-,)nic phendmena were not welt

understood, and over a period of years the phrase "sound

barrier" came into use. The need for a transonic research

airplane was recognized during the war, and in 1944 the

desi__n develooment of the Bell X-I program was initiated

(ref_). The "X-1 proved to be enormously successful, and

the flight of Captain Charles E. Yeager on October i4.

19,t7 proved beyond doubt that manned aircral, could I]_

faster than the speed of sound. An advanced vet-iotL ul th'e

aarcrat't, the X-1B shown in Fig. i, flew several re,-earch

missions for NACA to stud)' aerodynamic heating effects

The original X-1 aircraft as well as the advanced version

used aluminum construction throughout. Measured skin

temperatures axe shown for a NACA research mission

flo,vn in January 1957 at Mach 1.94. Note that skin

temperatures are low, less than 200 °F. Thereafter,

supersonic flight speeds increased rapidly, and the need lot

considering aerodynamic heating became evident. The

difficulties presented bv high temperatures accomp,_nvinz

flight at supersonic speeds became known as the "rhek:ma-I

ba:rier". For about ten ),ears, the "thermal barrier" ot

"thermal thicket" caused concern that large structural

weight increases would be required to keep material

temperatures within allowable values. Subsequently

researchers (refs. 2-3) found that these concerns did not

materialize because the problems were overcome through

research and development of effective thermal struct ores.

After the first supersonic flight, research and

development of high speed aircraft intensified. A contlact

for the design, development and construction of two X-2

swept wing, supersonic research aircraft was awarded to

the Bell Aircraft Corporation in 1947. The X-2 was the

first aircraft structure designed for aerodynamic heatinz

(ref.4). Until the X-2, speeds had not been high enougfi

for the structure to be affected adversely by aerody'nan_ic

heating. For increased strength at elevated temperatures,

the fuselage was constructed from K-Monel, and

aerodynamic skins used stainless steel. A drawing shown

in Fig. 2 illustrates the X-2 structure. The X-2 became

the first research aircraft to explore the "thermal thicket"

with speeds above Mach 2.5 On September 27. 1956 the

X-2 achieved its maximum speed of Mach 3.2. hut
unfortunatelytheaircraftwentout of control and the

pilot, Captain Milburn G. Apt. was killed.

During this period theoretical and experimental

studies were in progress to develop the technologies needed

for high speed aircraft. Considerable research was

underway in compressible flow, thermal structures and

materials. Excellent descriptions of these efforts are

available through survey articles and collections of papers

in books describing thermal--structures conferences. The

Applied Mechanics Review article (ref.5) of November

1955 by Professor N. J. Hoff of Polytechnic Institute of

Brooklyn describes high temperature effects in aircraft

structures with emphasis on the effects of temperature on

buckling and creep. In December, 1955 Professor R. L.

Bisplinghoff presented the Nineteenth Wright Brothers

Lecture to the Institute of Aeronautical Sciences in

Washington, D. C. The paper (ref 61. including discussion

by several prominent researchers, was published in April

1956. Professor Bisplinghoff's paper zives a comprehensive

review of structural considerations _or high speed flight

and an excellent account of design and analysis practices.

A book (ref.7) published by the Advisory Group for

Aeronautical Research and Develonment (AGARD)of the

North Atlantic Treaty Ozganization (NATO) in 1958

describes effects of high temperatures on aircraft structures

caused by aerodvnanfic heating. The book. edited by

Professor Hoff, contains 1(; articles written by U. S. and

European authors.

The next major fligllt program that stimulated

thermal structural research was the X-15. The X-15 had

complex origins including tile preheat and postwar work of

German scientists Eugene Sanger and Irene Bredt who, in

1944, outlined a hypersonic, rocket-propelled aircraft.

The evolution of their ideas which contributed to the

development of the X-15 is doscribed in the award

winning papel by Richaid P. Hallion (ref. S). Further

descriptions of the X-15 program a_e given b.v Hallion in

Ref. 9. and by NASA Langley research scientist John V.

Becker in Ref. l0 The paper written by John Becker was

pre.-ented in Bonn, \Vest German:' m IDe;. 196S. At this

meeting he accepted tile Eugene S,mger medal awarded b.v

the German Society of Aeronaut_c_ and Astronautics to

honor the success of'the X-15 program

On June 24. 1952 the prestigious NACA Committee

on Aeronautics charged the agency to study problems of

ma,med and umnanned flight at altitudes between 12 and

50 miles and speeds of Mach 4 to Mach 10. Bv 1954 the

NACA Langley Laboratory (now NASA's" Langley

Research Center) had formal a hypersonic study team

with John Becker as the chairmaa_. According to Becker

(ref.10) the unprecedented problems of aerodynamic

heating and high-temperature structures appeared to be so

formidable that they were viewed as "barriers" to

hypersonic flight. Nevertheless, the NASA group evolved

a baseline design that closely resembled the ultimate X-15

configuration. Three X-15 aircraft were built by North

American Rockwell.

A thick--skinned, heat-sink approach was adopted

to suit the short duration missions of the X-15. A typical

research mission lasted 10-12 minutes (ref.ll). Surfaces

exposed to aerodynamic heating were made of Inconel X, a

nickel alloy. Internal structures not subject to high

_emperatures were made of titanium. Skin temperatures

were designed for a maximum of 1200 °F. Figure 3 shows

maximum temperatures experienced m an X-15 flight.

On October 4. 1957 the Soviet Union orbited

Sputnik 1, the world's first artificial satellite.

This event changed the nation's priorities for high speed,

high ait!tude flight making the X-15 program vital to

Americas national prestige. The X-15 program served

the nation well accomplishing 199 missions between 1959

and 1968. The X-15 was the first, and to date, the only

manne:l vehicle capable of flying atmospheric missions at

Mach 5 for altitudes of 100.000 feet or higher. It made

many contributions to the understandiug of hypersonic

fligh't including the design of thermal structures. Later

sections will refer to some of these contributions.

From World War lI to Sputuik, the nation's

research programs were focused on high speed flight of

alrcra.fc. In thermal structures, the NACA Langley

laboratory made significant contributions. Many of the

Langley research efforts were led by Richard R. Heldenfels.

After "his retirement, he describes (ref.12) research

conducted at Langley from 1948 to 1958 on structural

problems caused by a_rodynan_ic heating.

Aftqr Sp_ltpik

After Sputnik, the nation's high speed flight

researth prograan was broadened to include a major

emphasis on manned space flight. The National

Aeronautics and Space Act of 1958 created NASA. aud

NACA becaxne NASA on October 1, 1958.

The expanded scope of the research is reflected in

the 1960 paper by Heldenfels (tel.31. The paper describes

proposed re-entry structure- as well as space vehicles and

spate ;tructures. The protc_ di._s itef.13) of a (onfelence

held n Camblidge. .Ma,_,,_im._etts uu July 25. 1961

conta_as 13 papers desc_ii)ing research on thermal

structures for manned and _lHi]anhed i'e--elltrv vehicles.

Tile conference was concerued primazily wit'h thermal

protec'ion systems for lifting vechicles. Two types of

therm,d protection systems are described:

"cool-structure" and "hot--structure" approaches. In the

cool-structure approach, the load-bearing structure is

insulated from high temperatures by an external heat

shield. In the hot-structure approach, the load-bearing

structure operates at nearly skill temperatures.

A NASA- University Conference on the Science

and Technology of Space Explu, ation was held in Chicago.

Illinois on November 1-3, 1962. The conference

proceedings describes research on structules tbr laund_

vehicles, winged aerospace vehicles and planetary entry

vehicles. The paper by Mathauser (ref. 14) gives a _,oo_1

account of Langley research on thernlal--s-truc(ural

problems for winged vehicles.

Types of winged vehicles under consideration I)_

NASA are shown in Fig. 4. These include a rese,uci_

airplane (the X-15), a re--entry glider, and a large

hypersonic aircraft. The re-entry glider is representative

of a X-20 Dyna-Soar vehicle that, was to be launched witi_

a ground-based booster. The hypersonic aircraft was to

possess horizontal t,ake---off capability and be capable ot

suscaiaed hypersonic flight. All three vehicles utilize

radiation--cooled hot structures. Radiation equilibrium

temperatures (See Eq. (13)) estimated for,a re--entry glider

are shown in Fig. 5.

The re--entry glider and hypersonic aircraft of the Mathauser paper are hypothetical'vehicles used in N ASA's

fundamental research studies. However, the Dyna-Soar

project sponsored by the Air Force did lead to the final

design of the Boeing X-20.. The events leading to the

Dyna--Soar program, its evolution, and subsequent demise

are described by Miller (ref4) and Hallion (ref,S) The

.....,''_" P_d3.E IS

07 pOOR QUALITY
Boeing development program began with a contract award

in November 1959. The X-20 was designed to provide a

piloted,maneuverable vehicle for conducting experiments

in the hypersonic and orbitalflightregime. However, the

X-20 program was cancelled inDecember 1963 before the

firstvehicle was completed. The Dyna-Soar program

acceleratedprogress in severaltechnologiesthat ultimately

were applicable to the space shuttle. A review of the

Dyna-Soar winged spacecraft technology appears in the

1961 paper by Yoler (ref.15). The structural design

utilizeda Rene 41 nickelsuper alloy primary structure,a

columbian alloy heat shield,a graphite and zirconia nose

cap and molybdenum alloyleading edges.

Concurrent with these activities,the U. S.'s

manned space flight program accelerated rapidly. The

effortbegan with the Mercury program including the Alan

B. Shepard, Jr. suborbital fligat on .May 5, 1961 and

America's first orbital fligi:t by John H. Glenn. Jr. on

February 20, 1962. The,effort continued _ith tile Gemini

program and the first two-man tlight by' John W. Young

and Virgil I. Grissom on March 23, 196.5. The Gemini

program achieved the first rendezvous and docking in

space and the first American "space walk". The Apollo

pro,zram began with the October 11-22. 1968 tlight bv

Wa_'ter M. Shirra, Jr., Donn F. Eisele and R. \Valte'r

Cunningham The program reached its zenith with the

historic flight of Apollo 11 by Nell A. Armstrong, Edwin

E. Aldrin, Jr. and .Michael Collins and the first lunar

landing on July 20. 1969. The hmar i)zogram concluded

with .Apollo 17 making the sixth and last hmar landing in

December 1972. America's first E_rti_-orbiting space

station. Skylab, was launched atop a Saturn \: booster on

.May 14, 1973. Three, three-manned crews visited the

space station with the last mission returning to Earth in

February 1974. The rendezvous and dockino_ of an Apollo

spacecraft with a Russian Soyuz cr_tt iu l_arth olbit on

July iS. 1975 closed our the Apollo I)rog_am.

All ot the manned sp,_cec_aft missions used

blunt-body re-entry _ei_icle_ arid ablative heat shields to

dissipate aerodynamic heating. These ballistic or

sen:i-ballistic vehicles h_d sznatl lift to drag ratios

permitting only limited maneuverability.

The lack of a follow-on winged research flight

vehicle to succeed the X-15 forced thermal-structures

research in the 1960's into a period of more fundamental

rather than applied research. The 196:2 paper (ref.14) by

Mathauser indicates that research had begun at Langley on

hypersonic structures. A 1966 paper by Heldeufels (rei.16)

discusses in more detail structund prospects for hypersonic

vehicles. Figure 6 shows isotherms on a hypersonic vehicle

assumed to be cruising at Math S at 88,000 ft.

Temperatures shown are radiation equilibrium

temperatures expected during a typical flight of one to two

hours duration. The paper discusses thermal--structural

designs for fuselage tanks for liquid hydrogen, wings and

engine structures. Hydrogen fuel---cooled structures for

engines and passive hot structures of high temperature

materials for airframes are described. Tile proceedings of a

conference held at Langley in November, 1971 contains

several papers describing basic hypersonic vehicle r_earch.

A paper by Anderson aJld Kelly (ref. 17) reviews the

technology base for propulsion structures, primary

structures and liquid hydrogen tanks. A recent paper by

Shore (ref. 18) reviews research on convectively cooled

structuresin the 1960's and 1970's.

The Soace Shuttle

The Space Shuttle resulted from a perceived need in

the 1960's for a logistical spacecraft to support orbital

space stations. However, after the lunar landing in 1969.

NASA recognized that funds would not be available to

support both the Shuttle and a space station. Justification

for the Shuttle shifted from space station support to its use

as a substitute for expendable launch systems. With

strong support from the Department of Defense, the

preliminary analysis phase began in February 1969 with

contracts to Lockheed, General Dynamics, McDonnell

Douglas and North-American Rockwell. Rockwell

eventually became the prime contractor, and construction

of the actual Shuttle orbiter began in June 1974. Rockwell

completed this vehicle, the Enterprise. m Septe|nber'1976.

Difficulties with the main engines and thermal protection

system delayed the second shuttle, the Columbia until

1981. Piloted by astronauts John W. Young and Robert

L. Crippen, Columbia completed the Shuttle's first orbital

flight on April 14, 1981. Further details of its history are

given by Hallion (ref.8).

The orbiter basically has a conventional

skin--stringer alumninum aircraft structure. The design ot

the thermal protection system (TPS) had the requirement

of k_el)ing structural temperatures less than 350 °F. The

thermal protection is compo.,ed ot twu type> iJ[ teu-,bh,

surface insulation (RSI) tiles. The RSI tiles cove|iw_, ih,_

orbiter are made of coated sih:a tibet. The two type- ditl,'t

only in surface coating to provide i)roteLtion for dilfe:eut

tenlperature envhOlllgleu is. The lox_.-telnl)et a t tl! e

insulation (LRSI) consists of S-inch square silc,t tiles ,rod

covers the top of the vehici, where temperatures are less

than 1200 °F. The high 1,mperature insulation (HRSI)

cons sts of 6-inch square tiles that cover the bottom and

some leading edges of the o_l)iter where temperatures ate

below 2300 OF. Reinforced carbon-carbon iR(.'C) is used

for the nose cap and _viug leading edges where

teml,eratures are above 2300 °F. Flexible teus,d)le stuface

insulation (FRSI) is used at locations whe,e tempe_a',tueare l_s than 700 °F. The distribution of tile tiles is show:_

in Fig. 7. and orbiter isotherms for a normal flight ate

sho_n in Fig. 8. Further details of the orbiter are given in

Ref 19. The launch of the first shuttle flight was del,_yed

by problern.s with the TPS, and the article ¢ref. 201 bx

Langley researchers Paul A. Cooper and Paul F. Hollowai

explain_ the problems and theil re,olmiuu

The National Aerosl)a_e Plane, X-101

In his State of the Union add,ess in February 1'_$6.

President Reagan made tile initial announcement o{

research on development of the National Aerosl),tce Plaue

(NASP). The November 1986 article by Robert L.

Wilharns (ref. '2_1) describes the program objective "to

develop, and then demonstrate in an experinaental flight

vehicle (the X-30), the requisite technologies to permit the

nation to develop both military and civil vehicles capable

of operating at sustained hypersonic speeds within the

atmosphere and/or operating as space launch vehicles for

delivering payloads into orbit". Under joint sponsorship

by NASA and DUD, technology development began iu

1986, and by 1988 McDonnell Douglas, General Dynamics

and Rockwell had won separate contracts to proceed with

development of the airframe. Pratt and Whitney aa:d

Rocketdvne had won contracts for propulsion system

development. A consortium of these contractozs is

parucipating in a 30-month materials and structures

augraentation project (ref. 22).

The competitive nature of the development

program and security issues have limited l)ubliLa::uns

describing thermal structures lo, the X-.J0. Hosteler. at
the 30th SDM conference,David A. Ellisof the McDonnell

Douglas Corporation presented an overview paper (ref.23)

describing the concepts under consideration as well as the

types of trade studies being conducted. Figure 9 shows

representative surface temperatures for a typical NASP

hypersonic condition. Candidate primary structure design

concepts include hot structures, activel.v cooled structures

and thermally protected structures. The design of the

center fuselage is particularly challenging because the

external skin is subjected to aerodynamic heating, and the

internal structure is exposed to cryogenic temperature due

to the hydrogen fuel. Candidate structural design concepts

are shown in Fig. 10. The paper stresses the importance of

a highly interactive, multi--disciplinary design/analysis

procedure.

Aerodynamic Henri ng

The prediction of aerodynamic heating is a problem

that has challenged thermal analysts since before World

War If. The problem involves soiving the equations that

describe the conservation of mass. momentum and energy

for the fluidflow. For altitudesless than 300.000 ft.tile

atmosphere may be assumed to )e a contimmm, and the

set of conservation equation_ are known as the

Navier-Stokes equations. For t'vu--dimensional flow the

Navier-Stokes equations governing viscous, compte_->ible

flow are a set of tbur non-linear, partial differential

equations with mixed hyperbolic, parabolic and elliptic

behavior. A solution of these equation, witi_ appropriate

boundary conditions aud con_tilutive equations for the

fluid pr()vides distributions for tie den_,ity p. the velocity

components u and _. the l)les,,,urt , 13. ,rod l]le temperature

T throughout the fluid AI the tlIJi,l.'aerodyIlalnic sin[ace

interface, tile aeiodxl_alnil hl'a_iH[_ I- tui]lputed !l(Jll/

FouriPr's la_ ushlg _he fiui(! _]:_._ .,taJ lunduc_i_it_ aIld the

derivative ot the temperature hor_:,,d _u the wall.

Prior to the digital compute_, there was no hope for

solving the complete .Xavier-Stokes equation,,, htstead.

analysts devoted substantial effort to obtaimng

approximate solutions, and with the help of experimeutal

data they developed engineering relations to predict

aerodynamic heating as well as skin friction. The

engineering relations typically are based on boundary layer

or other approximations. The relations are restricted by a

numbe_ of assumptions, and tile application of computers

to the problem of aerodynamic heating has contributed

significantly to a better uuderstanding uf the phenon_ena

and improved the reliability of predictions.

There is substantial literature describing

aerodynaanic heating, and the esuteric details will be left

for thermal analysts. However. because of the

interdisciplinaxy nature of thermal---structural problems for

high speed vehicles, some fundamental aspects will be

presented. Earl)" approaches for the prediction of

aerodynamic heating are described i,a the 1956 paper b_

Van Driest (ref.24), and in the t960 text by Robert W.

Truitt (ref. :25). A recent discussiou appears in the 1989

text by John D. Anderson, Jr. (ref.26).

The preceding section mentioned that aerodynamic

heating effects become significant at Mach numbers above

2.5. For Mach numbers up to about five, the flow is

considered supersonic, and above Mach five, the flow is

termed hypersonic. The NASP will experience

aerodynaxnic heating through the supersonic and

hypersonic regimes with a peak Mach number of about 25.

As the Mach number increases above five. new physical

phenomena become progressivel3 of greater importance

making hypersonic flow more complex than supersonic

flow. These phenomena include: (lj fluid dynamics effects

that limit the validity of boundary layer approximations,

and (2) high-temperature effects that introduce chemical

reactions. In addition, at altitudes above 300,000 ft. low

density effects restrict the application of continuum

models.

Classical Boundary-Laver Predictions

The concept of the boundary layer was introduced

by Prandtl in 1903 for incompressible flows. The

extension of the concept to compressible flows began ill the

1930's, and analytical work continued until the mid 1950's.

Thereafter, most research turned to the development of

computational methods for solving aerodynamic heating

problems.

The boundary layer equations are obtained from the

Navier-Stokes equations by an order--of-magnitude

aaatysis for the thin fluid layer next to the I)ody where

viscous effects dominate. Outside of the t)oundarv-la_e_

the fluid is a.,,<_umed invi_,cid, and the iluw is descril)e,[ b_

the Euler equations. Fur _tead 3. t_o---dimell.,iu_,,i.

laminar flow, the boundary-layer equatiun_ are:

-_(,_+_-_,v =0

&÷pv _=& 7:, 0 ___ "' . )

OP

,37 = 0

(1}

el :'.)T) O(c.>T)

. • ,-TT.,- u_-ut FK)

In the above, Pe denotes tile p,essme at the edge of the

boundary layer. The fluid specific heat is c . the thermal

P

condtctivitv is k. and the viscosity Js p. In oen_al, the>,-,

fluid aroper:ties are temperature dependent. "l'he I)oundar_

layer equations are nonline,_ but exhibit parabdi_

behavior. Tile perfect gas la_ is u.,ed _u rel,tte ple.-:-ute.

den>ilv _nd tenlpe,-atute Fu_ lmlhe_ det(til:, includizt_

bounda.r? conditions see Ref. 2t_

For some simple but impoltant cases, analytk,d

solutions to the boundary layer equations have been

obtained. The approach" involves making changes of

independent and dependent variables to yield simpler.

nonlinear, nondimensional equations These

transtormations lead to the concept of self-similar

solutions for problems such as a flat plate. Typically. the

nondimensional equations can be solved numerically tor

various values of the Mach number M. the Reynolds

number Re, and the Prandtl number Pr.

The solution for the flow over a flat plate obtained

by _,an Driest in 1952 (ref.27) demonstrates typical

featmes. The problem consists of a fiat plate with

specified uniform temperature T w, Fig. 11. The inviscid

flow outside of the boundary layer has coustant velocity

u_c and temperature T x. fVithin the boundary layer the

viscosity varies with temperature accordin_ to

Sutherland's law (see ref. 26), and the Prandtl numbe_ Pr

= pCp/k is assumed constant. Pr = 0.7. The ratio of the

wall ;o free stream temperature is taken as Tw/Too =

0.25. Figure (lla) shows typical u velocity profiles, and
Fig. (lib) shows typical temperature profiles. In these

figure, the Reynolds number, P,e x = px%cx/_*_c. The

fiat plate boundary layer solution illustrates results

representative of high speed flows. The velocit.v profiles

show that the boundary layer thickness increases rapidly

with increa-¢ing Mach number. In fact, Au(lerson {,'el.26)

shows that a hypersonic boundat3 laser thickness increases

appproximately as Mach n_mlber squared. One

consequence of a large hypersonic I_ounrlary bLxcr thickness

is that the viscous boundary laver alters the outer inviscid

flow limiting the accuracy of boundary layer predictions.

This behavior is called viscous interaction.

The temperature profiles show tile fluid behavior

that is of great importance to the structure. The

important point is that the peak temperature within the

I)oundarv layer is higher than the Ireo stream temperature

Moreover, thi_ peak temlJetatt, re incre_e_ rapidly with

increasing Mach immbet. The hill1 fluid temperature

within the boundary laver is (_ue to viscous dissipation.

Viscous dissipation is tl'_e prote.-.s where the large kinetic

energy of the high speed flo_ is _.onvetTe(I to thermal

euergy by the boundar.v layel shearing stresses. The

aerodynamic heating rate is proportional, hv Fourier's law.

to the slope of the tempelatme profile at the wall (y=oi.

Fie_ore llb shows that _he fluid re:nporaTutes withi_ the

I)ounldaly layer rapidly mtrt,_-e _i_h imleosilto_ .\lat:il

illllll})er. Tiff,, i., tile I,:,,it le._-_,_ hJ_ lb, it_,.,[ of -I,'_i,tl

thermal stru(_lmes tot hi-h ,])e,'_i '.[i-}H.

Classical. seif-_.hm]ar t)oltndal3-OhVi_mr ha_e ,_lSo

been de:wed for tile .',l,t_th.lritHI U"it)ll ol ,I I;]hhl [*otiv. All

impolzant result of -thi, ,tm_](si- wa, rv show 1o_

hypersonic flows that ti,e sl,v_,rhvio_ ll,e,,tm,., _,tre valies

inverseiv with the squaie :oot ,o! ;Iw b[tlnt l..,odv ladius.

Thus to" reduce aerody:_,_mic i,eating the _,,hi, Iv m)s, ,_nd

leading edge regions need to be ,,s bhml as possible

The ('onv_ctive Bounda_v {"ondi;ion

For der_erlllillil_ .,,tlll(ttl:al l¢*I,,tpe_,:ttlle>, aelo---

dv_amic heating is olte_ rel;_osenled as ,, convective

b_.,undarv condition. Boundary lave_ analysi: el compressibie flo[v shows that tile "heat rral>le" hetx_een the

boundary la)e_ and tile wall cat: i)e exp_e_,seti a_

qw = h (Taw- Twi (2)

where qw is the local heating _a_e _e.,-. BTLsft 2 - s) at

the wail. h is the t.on,.etlio_ _o,,lfhient. T i., the

aw

adiabatic v,al| temperature, and fw is the wall

temperature. Tile adiabatk wall temperature is tile

temperature the fluid attains for no heat transfer between

the fluid and wall. Temperature profiles illustrating the

heat transfer possibilities for high speed flow are shown in

Fig. 12. The figure shows that for wall 'temperatures less

than Taw, the fluid flow heats tile wall: when tile

temperature at the wall is Taw. there is zero heat transfer:

and. for wall temperatures greater than Taw, the wall

heats the fluid. Eq. (2) shows that aerodynamic heating is

directly proportional to the differeni:e between the

adiabatic wall temperature and the actual wall

temperature. The adiabatic w,dl temperature is always

ifigher than the free stream temperature T x aud "drives"

the heat transfer.

The heat transfer coeffitient is often written in

non--diu_ensional folm.

h (3)

CH = Pe Ue Cp

or using Eq. ('2_)

qw

CH= Pe Ue Cp (Taw - Tw) (4)

where C H is the Stanton number. Another ahernative is

to define the Stanton number in terms of fluid enthalpies.

i.e

q W

CH=Pe ue (Haw - Hw_ ,3,

where Haw is the adiabatic wall enthaipy, and tt._ > th,-

wail enthalpy. In this apploach, entl_alp,. _,epl,tetemperature as an unknown ill lhe tqow anai',si_ Fo; a

catcli(.ally perfect gas (cunstam spet_ifi( heats., h = _pl.

othe_vise, dh = Cp (T) dT.

Thus. to use a cow,,-_ tive buurlxlarv condition it i:

necessary to know lie St_:t_on numbe, and tile adiat)a_it

wall temperature. Both. ;;; £ellelai. are [o(,_[ (ill,rarities:

tha is. they vary with ;.JS[riol* dlUll£ the wail. The

cJas...it.al [lat plate bound,-t: ', laver solutigan will serve ,_: an

illu:_uation, gxter_sive sti_tiies ut ti_e flat 1)!ale pzoblem

/e-,, ref. 24/ show _ha_ ,he adi,_b,_ric wall _empetalfue _at:

he _aicuiated with o..uud a(<u:'at3 h_

r - - _ 2 _U!

Taw = T x i I -.- I -----r--" Xl x

(

',vii-re " iS lhe latiu ot fit,hi -W'tili* h,,,_l> = l' .

t

\

is the recoverx tactol. .-kn,iy_ff,_l ,ulld ex,pelil/wlllid lest:it>

aho_ tltat the recw.el._ t,_t.tul c,m be tumputed a.,.

r = P 1/2 laminar flow

r

r = P 1/3 turbuleut lloyd

r

The boundary layer analysis for tile flat pla_e t'hm q_ows

that the Stanton nunlber has the lorm

CH = f(Me' Pr' 6' Tw' Te) (7)

v--aT7.

where f denotes a functional relationship. Ecke,t ill l%t,

(ref.2S) showed that I)v using the _uncept of ,t reletel_o.

teraperature, a simple i'ornnnla for ('1t could tw developed

frOlrt tile correspondin_ r_uh fur hm ._l>eed. in,. Uml)l"'"i el,

Ilow. Tills approach, Lalled tile letete_iLe tenli_t,ttutt'

method, computes the Stanton ntunflger from

t,_" FCOR QUALII"_
"2

C H = 0.33:2 (Pr')

X

where Re_ aJ_d P," are

temperature T . That is,

evaluated at a

Is)

refel ence

s

p ue x

Re x =

#

(9)

_ C"

P; = ___

k

where P', Cp = and k = are evaluated at a reference

temperature T . The reference temperature is computed

= f ,I _

I =Teil+O.O:_2.',t_-,-o..5< (I -1 -El , ]o)

where as before tile sub-,o{pt e ([_nute.- qualllities at the

edge of the boundary {ayel. l:urthet d('taiI:, int.luding

eXl)te...-ions tot ski,, l"tictiu,,. ,q,d ¢'([I.NtliUll:' [OI' tuxl)ulent

flow are given ill tile Eckert l)al)et !rei. 25). the text by

Anderson Iref. 26). and other ct.llxettive heat tt,_uster

texts.

[,,imiTations of ('la_si(a{ Boundary [avel l heorv

Classic,d bounda',) la) er theory piovide.- basic

insight to understand fundamentals of aerodynamic

heating. Tile boundary, laAer ._oiutio,_ in some ca.,e.- .Yields

very practical eno.inee, i,,o_ results th,_t a_e fiequen_i.v used.

Bu{ boundar)layer theory has limit,,tious that deserve

mentioning.

Classical self-similar solutions are limited to a few

problems with simple geometries (e.g. the fiat piatej and

simple boundary conditions (e._. constant ,,'all

tempe, ature). However. i,,the ag,ior tile tonlputer, this

limitation i.', not vet 3 :,t.liOub be_;._d:,e tht. [KJulldaly ldVel

equations. Eqs. (l). ale :,oWed numeticall 5. hi fact. siilce

the equ,ttions are palabolic, mar(hing methods may be

used al_d very efficient computel pro-_ralns have been

developed. When combiued with inviscid'flow programs to

determine variables at the edge of the boundary layer, very

effective analysis procedure, have been developed. Some

of these methods are described in a 1987 paper (ref.29) by

Fred R. DeJarnette et al. which reviews approximate

methods foraerodynaanic heating aaalysis.

There are several flow situations where the

boundary layer approximation of uncoupling the inviscid

and viscous analysisdoes not apply. Two such situations

are where viscous interactionor flow separationoccur. As

mentioned earlier,viscous interactionrefersto flows where

the viscous boundary layer interactswith the inviscidflow.

Such a situation may occur for hypersonic flows with

relatively low Reynolds numbers. Boundacy layer

thicknesses vary inversely with the square root of the

Reynolds numbers, so low Reynolds number flows have

thick boundary layers. Flow separation refersto the case

where the flow turns away from the wall into itself

introducing a local recirculation region next to the wall.

Flow separation does ,tot occur when the pressure

decreases in the flow direction. However, when the

pressure increases in the flow direction (calledan adverse

pressure gradient) flow separation occurs, and boundary

layer theory does not apply'. Flow separationmay occur in

severalproblems of importance in high speed flight.X-15

experie:tce provides two such examples: (1) flow over

protuberances, and (2) shock-boundary layerinteractions.

In both examples an adverse pressure gradient occurs, and

flow separation develops. A consequence of flow

separat,on in both instances is to inclease significantly the

local aerodynamic heating. For these ptol)lems, bouudary

layer theory is not applicable, and the pheliOt,en,_ i.,

goverut.d by the Nay,el-Stokes e(luat)(lns.

Ni_vl_,,---c;;t,okes Solutions

One of the major advances in computational

mechanics in the last four decades is the development of

capability to solve the Navier-Stokes equations

computationall.v. The survey article (ref. 30) by' Dougl,ts

L. Dwoyer et al. describes current capability ot

con)t)utational fluid d>u,mli(s (CFD) for hyl)e,_t)ni(

aircralt The Naviet-'stokes equ,_tious writreu 1:1

conse, x ation form are

o {[.}. _{E}- o __ {F} =_, ,L:>

where :I'} is a vector of co,><vation variables: {El and

{F} a,_ vectors of the flux ,omonents in the x ,_ud v

direc:ichs. These vectors are

[.IT= rI* H {; I)_ t j_'

-IO

k

i" 2.p _ l"

O'x 7xy UO'x+ V 7X_ "_lx i

{F}T= [pv

- [o rxy

.)

puv pv-+P /)v &'-Px i

O' v Lit,y* vo'x-_1 v ;

where $ is the total internal eue,g). Each of ti,e flux

vectors contains t_'o vectors of components representing

inviscid and viscous flux co,ttributions. Tile conservation

equations are supplemented by a thermodynamic equation

of state relating pressure, teml)erature and densitx

(typically the perfec_ gas la_v), h, ,he viscous llu_

{:OlltpOltents, the stre:,..,es e . e ,_m} 7 AlL' le[,lted It) )lie

X x xy

_elocit 2 gradients a.,_suniing St,Jke'a ll)t_othesiz ]lie heat

fluxe, qx and qy are related to temlretature gradients t)_

Fourier's law. For air, the temperature--delx_,,de,u

viscosity is computed from Suthe,'land's law. and tile

thermal conductivity is computed assunuug a coJl.,taut

Prandtl number of 0.72. Considerable success has beeu

achieved in numerical sohttions for the two-dimensional

Navier-Stokes equations with supereomputers. Numerical

solution schemes include finite difference, finite volume

and, more recently, finite element methods.

Three-dimensional llows ,emain a challenge alti_ough a

viscous solution around an X-24c lifting body [re,. 31) has

a,' ,',,_<,_ :eUALrF_
been obtained. Such solutio,s, eve, with tile hu-gest

Sul_rcutnputers, are ,or t:ap,d)le of resolving

three-dimensional flow details, particularly aerodynamic

heating rates. An important inoblem for which

Navier-Stokes solutions are Woviding valuable insight is

two---dimensional sho<'k interaction phenomena for the

NASP.

In 1967 NASA conducted a series of X-15 flights

with a dummy hypersonic ramjet engine mounted on a

pylon under the rear of the fuselage. On the third flight

with the dummy engine on October 3, 1967 the X-15

reached a ma.ximum Mach number of 6.7 at an altitude of

99,000 feet. During the flight severe structural damage

was experienced due to complex shock impingement and

interference effects on local aerodynamic heating (ref. 39.).

Considerable heating-induced damage occurred on the

engine pylon indicating that local temperatures exceeded

the |nco,le[ X melting temperature of 2600 °F. Since then,

shock interference heating has been letugnized as a critical

problem for high speed _ehicles because extreme pressure

avd heat transfer rates can o(cm in highl_ localized

re_ions where tile illterlerelKe p,_lletit illapil_ges Oil the

_urtaLe.

Shock interference heating i_ an important

consideration in the design of the cowl leading edge of the

engine structure of the NASP. Tile problem has been

strong motivation for recont studie> uf shock interference

heatino_ on leadino, edges. Figure i_ shows the overall flow

field tol the NAS-P. The inte,action I,u_een the aircraft

bow shock and the cowl bow shock can cause severe local

aerothermal loads on the co,vl leading edge Shock

interference heating on leading edzes has been studied

experimentally by Wieting and Holden !ref. 331 and

computatioi_ail3 b3 Kloplel and "xv,, tel. 3.1-35) a_- well as

Thaxeja et al. (ref. 36) Dechaumphai et al (lef. 37)

studied tile flow-thermal-stru¢.t_nal behaxio_ of leading

edges. Figure 14 shows the l)robiem statement tor the

Navier-Stokes solution, an adap:i_e fiMte element mesh

and a comparison of experimental ,tnd predicted heating

rate distributions. Adaptive. ut_structmed meshes, a

recent advancement il, fi,nte eieme,I ,rod finite volume

flow analysis, provide high re:.ulutioJl of Iho_ details such

_t_ shocks and boundar5 h_3et." _ he _OlnlJala_ive .-,urface

heating rates sl_own m Flu.. l-t_ .,how _etv good

agreement, but the paper p'oint., out there wa's some

diffe,euce in prediction uf the un,Jisturbed heating rate qo

that normalizes the curves. Note that the results show the

interference heating ,'ate is almost 15 times the

undisturbed level. Peak interference heating rates for

NASP flight conditions and geometries can lye as high as

70.000 Btu/ft 2- s. To put this value in perspective, a

typical 1 kW portable hair dryer produces a heating rate of

.)

about 30 Btu/ft" - s. Such high local heating presents a

formidable challenge to the structural designer.

A discussion of aerothermodyumnics of

transatmospheric vehicles from the design perspective

appears in the papers of Tauber et al. (refs. 38-39).

Aerothermal Load Effects on Flight Structures

In his classic 1956 paper (ref. 6), Professor R. L.

Bisplinghoff identified the basic stlucttnal and aeroelastic

considerations for high slmed flight. Flotn a structural

perspective the conside_ation_ include: (1) deterioration of

mechanical properties at elevated temperatures. (2)

thermal stresses introduced by l elnperature gradients. (3)

modification of stiffness and vibration p,-uperties, and (4)

a_ruelastic instabilities. "_rhe_ considerations remain

important today. The general problem consist-q of

determ ning the behavior of the flight structure under

aerothermal loads, primarily aerodynam c pressures and

heating. The behavior of the flight structure can be

classified according to whether the response is quasi-static

or dynamic. In a quasi--static response inertia forces are

neglighle, and the structure responds to the aerothermal

loads slowly. In a dyna,nic re._ponse iuertia forces have a

signifi_:ant role. and tile structure responds with vibratory'

motions. The determination of the quasi-static response

of structures has traditionally been called thermal stress

analysis. The determination of ti,e structural dynamic

response considering the interaction of the deforming

structure and the aerodynamic pressure in the presence of

aerodynamic heating is called aerothermoelasticity. An

area of recent concern is the response of structures to

acoustic loads. The determination of the dynamic response

of a structure to random fluctuating pressures in the

preser_ce of aerodynatnic heating ,nay be called

aerot t-.ermoacousti cs.

Ot, asi-$_atic Interactions

lo determine the quase-static response, a logitM

al)ptoacll is to separate the ae,othermal-structmal

prol)l,m into distinct uncoupled problems by assuming

weak coupling between the external aerodynamic flow and

the s ructural response. This a_umption is pernnssible

when structural deformations, ,_ze ioo ._mall to alter the

exter,_M flow. Durin_ N-_', flights i,efs. 9-10) several

unexl,ected thermal ijrohler,_- wet,' ,,ntountered due to

i,iten:.e local aerodvt,amtc I:,.,I mg ],etau:,e ol qud.si-static

inter.-ctions. \Vinilshield tl,,_,_,,°.e occurred when therluai

buckling of the letdinel _ta;_._' t,;t>ed il,te,,-e local heating

in th. _ o_lass. These plD})lellts _,t,l:'l_," identified and sol_ed

durin_ flight tests. However. according to Ret. 10. "the

reaii_ important iesson here is that what are minor and

uuinipor_ant features ot a sub:,onic ut supe, sonic ailctaft

must be dealt with as prime design problems in a

hype:sonic airplane".

A more recent example of a quasi--static

flow/structural interaction is tl_ermal protection systems

testeJ in the Math 7 eigM-foot Higl_ Temperature Tunnel

at the NASA Langley Research Center ref. 40). The tests

sho_t that panels "bowed-up" into the flow to produce

heating rates that are up to i.5 _imes greater than fiat

plat{ p,edictions. Thornton aud Dechaumphai (ref. 41)

used a finite element approadx to study coupled flow.

thermal and structural behavior of aerodynanucally heated

panels. Some typical results tur concave and convex

deformations are shown in Fi,.,. iS. On13 _ery modest

deformations occurred, but flow features were altered

significaaitly. For the .klach 6.0 conditions studied, panel

deformations introduce shocks, expansions and

recirculation regions in the flow. The effect of convex

panel deformation is to increase local heating rates on

windward surfaces and decrease them on leeward surfaces.

In the overview paper (,'ef. 23) describing the design

of au airframe structure for NASP, the interaction between

the external flow and the thermoelastic deflections of a

mo_able wing is described. Figure 16 shows the results of

the aerothermoelastic load analysis. The wing deflected

shape is shown for several different iterations in the

aataysis cycle. The paper notes the consideration of the

interaction resulted in the prediction of lower deflections

and stresses which translated into reduced structural

weight.

The recent anah.ses of qua_i--static flow/structural

interactions illustrate some of tile impo,tant effects of this

'_ ,"'"_' ' }:'AGE IS

OF poOR QUALITY
behavior on flight structures. However. because of their

multi-disciplinary nature, the analyses are difficult and

expensive. In addition, the analyses are unvalidated by

experimental data.

Dynamic Interactions

In a 1963 paper (ref. 42) I. E. Garrick of NASA

Langley surveyed developments in aerothermoelasticity.

The classical aeroelastic triangle representing interactions

between the fields of aerodynamics, elasticity and inertia

was extended to include thermal effects. The new figure,

the aerothermoelastic tetrahedron, illustrates the

interdisciplinary aspects of aerothermoelasticity. The

paper also discusses aeroelastic consideration of the X-15,

effects of transient heating on vibration frequencies, and

panel flutter. Earb it] the X-15 flight program the pilol

reported a rmnblin_ noise at high dxnau)ic px'es>ure _,. The

turued out to be pauel tlut ter oi lalge atea, s ol Ihe skill ol

the side faivings and tail, This ptM)lem was strong

motivation for NASA studies in tile 1960's and 1970's of

panel flutter at elevated temperatures. A recent

assesslnent (ref. 43) of flutter model testing relating to

NASP provides an excellent summary ot flutter literature

for supersonic and hypelsonic flight.

The response of aerodynamically heated structures

to acoustic loads is an important consideration because of

aCoIl_IiC [atioue. Retvteuce 44 note.', that operational

experience for a variety ot airclalt has demonstrated that

intense acoustic p_essme> can cause fatigue failures of

lightweight structures Preiimin,_r) estimates of acoustic

loadings for the NASP indicate pressme levels that are

_ell into tile range where ace _stic fatigue failures have

occurred in the past. Thermal effects are important

because: (1) there is little acousti( data for high

temperature structure., a_d _atetial.,. (2i thelmal

prestram and buckling cau aflet >)r,tln Ietels _ignificantl),

alld (3) temperatme can chauo.e ilialellal p[ul)erties aud

tile latio.ue lifetime ,t., exlnes_ed iu tile S-N diagram.

tle_ei;l elloll., t_) duxelop _u_E_l)u'.,_i(>iml me,bud, Iur

ae_othelunoacoustit., ale de,(_il,ed in _hc dis.,ertation by

Locke [ref. 45) and the paper I)_ Locke and .Mei tref. 46).

De,¢ign of Thermal St tuc[ ur_;s

The design of thermal structures is a complex

process that involves consideration of the tlight vehicle

trajectory, aerothermal loads, tilermal structural concepts

and materials. A near optimum design involves tradeoffs

among these and other factors. The design of thermal

structures is too complex to be discussed in detail in this

paper, but fundamentals _ill be cited.

Flight Regime

To determine aerothermal loads and heat transfer

to the vehicle, the flight trajectory must be determined.

The trajectory is determined by mission requirements.

The basic equations of flight mechanics, hypersonic

aerodynaznics and re-entry heating appear in the text by

Wilbur L. Hankev (re. 47). Recent design studies for

hypervetocity vehicles (ref 48'. a l( for NASP (ref. 23)

describe relationships between mission requirements and

thermal-structural design concepts.

A parameter used in the design of thermal

structures is ra/,.liation equilibrium temperature. The

radiation equilibrium temperature i_ the upper level that

the surface of a structure can reach. An energy balance at

the surface of a structure at radiation equilibrium states

that the aerodynamic heat flux given by Eq. (2) is equal to

the heat flux emitted by radiation, i.e..

qw = h (Taw-Tr) = at Tr4 (13)

where T r is the radiation equilibrium temperature, < is the

surface emissivity, and a is the Stefan-Boltzman constant.

Eq. (13) states that all of the incident aerodynamic heat

flux is emitted by radiation; none of the incident flux is

conducted into the structure. Actual surface temperatures

mat be lower than radiation equilibrium temperatures

because of conduction heat transfer into the structure.

Two key factors of the flight regime for thermal

structural design are radiation equilibrium temperatures

and exposure times. Radiation equilibrium temperatures

for the X-15, the Spate Shuttle, aJ_d the X-30 can be

compared in Figs. :l. S and 9. respectively. The flight

regime for NASP vis-a-vis other flight vehicles is shown

in Fig. 17. The importaalt point is the design of hypersonic

cruise vehicles like NASP differ from previous designs

because of much longer flight durations. A third

important factor of hypersonic fligtlt is use of li(luid

hydrogen fnel. .-_mong other (onsiderations. cr)ouenit

hy(l:o_en introdutes ,i_nifi(ant slructttra] Ietll[._,ta_u¢

Rlit(]i(.'lllS. _otlt IIh.'lXill/lllll IPtll'_Xqdttll+es &lid gt.uti(..t-

_ntluence structural design concept • and material selection.

lh_;rmM Prgrec)ion _v<Vtms

Hypersonic flight vehicles require thermal

protection systems to will,stand sustained aerothermal

louts. The therm,d prot(,c.tion system, the supporting

aitfl ame and the eugme sut,t ture ale examples of thermal

Stlll(lllres. file desio.u ot,_ thernla[ protection system is

ba_-ed Ul_Jn the l)_iucil)le that )he energy transmitted by

Ihe hot boundary la)e_ flow ll'lllbl l)e absorbed or rejected

b_ the therm,d p_(,tettion s_steu_. Figme 17 identifiethermal protection s_stems used b) hypersonic flight

veh ctes. For _elatixely short missions the .Mercury and

Al)dlo spacecraft., absorbed tile thermal energy through

abl;,tive heat shields. The Space Shuttle absorbs re-entry

hexing and thermally insulates the airframe by a very

effe,:tive, but fragile, tile system. For relatively short

ntis.dons, tile X-15 absorbed the aerodynamic heating b5

usiLg .-kin and ait:tiame of high temperature met,dli¢

materials. Fo_ lone. (lu_ation but lower temperature tligh_s

tile YF-12 [or SR-71) u•ed ,_ hi_h-temperature tilamtml

structuLe _ith high eHti_ivity ,at late __oatiug at radiation

equilibrium to _eiect tile aeto(Ivnamic heating. For

sustained, very high t_.mperatnre flight_ a combination el

several thermal protection concepts is likely to be used.

Fizure l0 shows candidate thennal protection systems i6t

N._SP which include hot structures, insulated structures

and convectivety cooled structures. Convectively cooled

airframe structures are relatively new with no previous

fl!ght experience. Design and experimental studies (ref.

1_) conducted over the last twenty },ears indicate that

these systems will be effective in absorbing thermal energy

in regions of intense local heating, e.g. shock interlerence

heaung on engine structures.

_D_er_als

The evolution of thermal structures for high speed

flight vehicles has placed ever-increasing demands on

material performance. Structural designs typically require

high-stiffness, thin-gauge materials that can be fabricated

inw complex, built-up structures. Thermal smtctural

designs typically retluire high strength, low density

tna_etials that _etaiu their desirable ptoperti_ at ete_ated
temper_.ures. There is a substantial research effort

currentlyunderway to developnew aerospacemateti_s to

meet these challenges. References 49-56 are recent papers

describing progress in material development for thermal

structures.

The performanceof the NASP willdepend on the

development of new, lightweightmaterialsthat can

performat temperatureshigherthan today'smaterialscan

staJld. Today's high-temperaturematerialssuch as

nickel-basedalloyscannot be used because structural

weight limitsmandate ligh.ter-weightmaterials. The

structuralweightlimitsrequlrethe useof thingauge,low

densityalloysand composites.FigureIS compares specific

strength and stiffnessof severaladvanced n_terials.

Below 2000 °F the structural designer has several

ahernatives, but a_ve 2000 °F tJ_oic_ are limited. The

NASP materials and structure augnientatiou project (ref.

22J has tl_e specific goal ot advancing the

readiness-for-use data of several advanced materials that

ate needed for the vehicle. The (,ontractor and materials

under development in this program aae listed in Table 1.

Table l

N ASP Materials and Structures Augmentation Program

Coni ractor

General

Dynamics

McDonald

Douglas

Pratt

and Whitney

Rockwell

North American

Rockwell

Rocketdyne

Reflactor.v composites

carbon-carl)on composites

ceramic composites

Titanium metal matrix composites

High creep strength materials

Titanium aluminide alloys

Titanitim aluminide composites

Titanium aluminide alloys

High thermal conductivity composites

Coppe_ matrix composites

Be,yllium alloys

Heat Transfer and Thermal Stresses

The design of thermal structuresrequiresthe

determinationof temperatures,displacements,stressesand

strains throughout the structure. A structural heat

transfer analysis is needed to determine maximum

operating temperatures that guide material selection.

Temperature distributions axe also needed to compute the

"thermal-loads" for a structural thermal---stress analysis.

The heat transfer and structural analyses rest on the

conservation equations of continuum mechanics,

constitutive models of material behavior, and

computational methods implemented on modern

computers.

_onser vat ion Eo flat i_lllS

For continuum formulations of solid mechanics, the

conser.ation equations that must be considered ate

consercation of linear momentum, conservation of angular

momentum and conservation of energy. The conservation

of linear momentum produces three equations of motion

.-}

O a.. a "u.

/---LI+ Bi = p2----@

Ox. atJ

(14)

where aij are components of the stress tensor, Bi denote

body :orce components per unit volume, p is the material

deusity and ui are the displacement components.

Conservation of angular momentum shows that the stress

tensor is symmetric. Conservation of energy considers the

work done by the stresses, thermal energy transported

a('lO,_:, sulfacl_,s I)v col|dnttitHl, lhell'nal and ulecllaliicctl

I.qll, l_.2it'r- >ft,ll(_J xvilhiil Ihe lii.ilcl ial. ,lllll kilie'l i¢ elit'lU- _, ttlt,:'

Io llt,llt'iiit] IfiOliOII. [lie llJll_'t'l\dli _jll t)l enel_. _, eqtlatiOli

for a ((llltillilillll lllodol (ff a clohlrnlM)lo Iw)dy is

0 qi O cii " O tl__

-aij_-"-e p _-_'- ,4

tlS)

whele qi denotes COITIPOrH'IIISof heat flux, _ij are

components of the strain ten,sot, u is ti_e internal energy

per unit mass, and Q is the z,_te of internal heat generation

per emit volume. The interh,d energy of the solid depends

_,,lll lilt' ,"_,l,tili:' dUd [CIl;])O:.,:tlll', Ih,ll. i:'. U : tl ('ij' T)

1-he 5ii,tii_ coml)OllellI- dt a point ,,le related to the

,[i.i;l,, elne_at (_Oilll)Ollt'lii 3 I)>

l.fO ui.o ui,o Ukd Uk';

. - " j l i 1"

_1(,)

Ofter displacement ..radients .le assumed small, and the

last term ii, Eq. (ibi-tllat iil\olx_'., pit)ducts and powel_ of

(li:pl;,_.enient gradients is neglected in conlpatison to the

first two lerlll$. ['ho toslill is the lille,U

<t i;dl,_tisplatenwnl ipl,tt ioni17)

Equations (16) or (17) show that the strain tensor is

svmr_etric. Thernial--_tructural problems are often

formalated in terms of the linear strain-displacement

relations. However, under severe conditions, structures

may experience large deformations, and the nonlinear

str_tin--displacement relations may be required In these

circumstances, the definitions of the stress components

have to be-interpreted relative to the undeformed and

defutlned configurations ,f the structure. The

conservation equations, in the context of thermal stresses,

are derived in the classic text by Boley and Weiner (ref.

57). and in the recent series of volumes on thermal stresses

edited by Hetnatski (refs. 58-59).

Conservation of energy. Eq. 115}, states that there

is a relationship between stresses, strains and temperature

in a deformable body. In one interpretation, conservation

of energy indicates that variation.- ul stresses and strains

within the solid alter the heat flow and thermal energy.

The equation expresses a conversion of mechanical to

ORILii:,_L PAGE IS

OF POOR QUALITY
thermal energy. The relationship between stresses, strains

and temperature in the energy equation is known as

thermaJ-mechanical coupling. The conversion of

mechanical energy to thermal energy according to

con_rvation of energy is a well-known phenomena which

has been studied exten,_ively for ela_qk behavior 0el. 60).

Studies of coupled thermoel_titit) have e,hu_n that for

metallic materials within tile elastic range, the conversion

of mechanical energy to thermal energy may be neglected

for aerospace applications. More recent studie_ (refs.

61-'-/03) of thermaJ-lnechanical COUl)lhlg tot non-linear.

inelastic behavior suggest that coupling may be importaa_t

under some circumstances. However, in analyses of flight

structures, thermal-mechanical coupling is usually

neglected. One argument for this assumption is that the

external energy supplied to the structure by aerodynamic

heating is so ]'arge that the thermal energy _:onverted from

nlechanical energy is negligble ill comparison. This means

that the energy equation can be simplified by assuming tile

solid is undetormable, that is _.=0 which is the usual

_J

approach in heat transfer texts, l'mle_ this circumstance

tile inte, nal eumg? is ,ega_,ted as a functio,_ ol

teuaperature alolle, aild it is customary to take

0 u ,._,0 T

T-_-=o I,F-T

i_ the mat,,rial'> -pt,_ilh hoau which i,

l,'!llp('l,iltll,' d,,pt.'n_i,'lll \\itl_ i,,>, -il,_i,;i',ita,Tiou_. _]le

t.OIl.'._l \dl F()II Of _1!_.'i '2_ ,'tiLl,iT il)ll I,'(i:l_ ,'- I_}

+.)ql+.(O I=C_ rL$) Ox. ' O t "_

I

which is the equation customaiii', found in hear rzansfer

lexts. The heat flux compouent¢ ,_e normally _elated to

im_pe_atule gradients by Fouli(,T's law. For an

alfiSOtIOpic, material Fuu_iei"., 1,_\ .-t,,re-.

aT

qi= -kij_7--- 7 (19)

J

where kij are components, of ,-tthezmad tonductivity tensor

In general, the mate,'ied the_mal couductivities are

temperature dependent.

Structural Heat Tr_n_f_'r

[-'or COllth]( tiOlJ [IU&l I I,iIJMt'l ill ,l_'t O_l).4(P

.-tttttt tu'es, the tl,_.-it al [w.i Comit_cl ioil t,qllatiOll i:_ tl_e£l.

Substitutiug Eq. { 19) into Eq. _15! yiehts

0-1-

0 "k ] O 2

-0--xi [ ij "0"_'j +/x_--_- = Q (20)

which is a parabolic partial differentialequation. This

means that thermal disturbances propagate at infinite

speeds through the body. To address this a,mmoly, some

authors (refs., 60, 64,-.G5} have used a modified form of

Fourier's law to derive a hyperbolic energy equation. Ill

the hyperbolic energy equation, thermal disturbances

propagate at a very high. fi _te _ave speed w _ch is called

"second sound". The application of the hyperbolic energy

equation for solids is controversial because the phenomena

of finite wave speeds of thermal disturbances has never

been demonstrated for structural materials although it has

been for gases. The parabolic energy equation is

customarily used for st,'uctural heat tlausfer.

The heat conduction equation is solved subject to

an initial condition lind boundary conditions on all

portions of the surface. The initial condition specifies the

temperature distribution at time zero. The boundary

conditions may consist of specified surface temperature,

specified heat flow, convective heat exchange and radiation

heat exchange. These may be written

Ts=Tl(Xl,X2 x3 t) on S1

qi ni="qs on S2 (21)

qini=h(Ts-Te ) on S3

qi ni=a_T4s"aqr on S4

where ni denote components of a unit outward normal, and

Si (i=1.4) denotes portions of the surface. The specified

surface temperature is T1. and the specified surface heat

flux ,s qs (positive into the surface). In the convective

boundary condition the convective exchange temperature

Te is Taw (see Eq. (2j) for aerodynamic heating. In the

radiation boundary condition c, is the surface absorptivity.

and qr is the incident radiation heat flux. For structural

heat transfer with surfaces at significantly different.

elevated temperatures, radiation exchanges between

surfa.:es can occur. The determination of radiation

exch_,nges between surfacos is complit.ated because:(1)

radiation emitted by a 13pical su,face depends on its

surface tempexatu,e which is unknown, and I2} the

geometrical relatiouship t_etween each surfate must l)e

consiJered. Ofteu. the radiation problem is handled by

discritizing the radiation boundary' into N discrete smtaces

which are assumed isothermal. If a radiation heat flux on

the ith surface is called H i, a set of N simultaneous

equations may be developed to determine H i in terms of

the temperatures T i. In matrix form these equations may

be wzitten as

[ - Iv] I]{H}= IF]{,oT4}

(.22)

where the components of the matrix [F] are the

viewtactors Fij. [I] denotes the identity matrix, and Fij is

the fraction of the radiation energy leaving surface i that

arrives at surface j. The determination of viewfactors for

complex three---dimensional structures is a formidable

computational task.

An excellent exanaple of heat transfer for a complex

struc;ure with aerodynamic heating is the space shuttle

wing. William L. I_o and co-authors have documented

several studies (e.g. refs. 66--68) of heat transfer analyses

with comparisons to flight-measured temperatures. The

wing geometry, a three-dimensional model of a wing

segment, and comparative histories of structural

temperatures are shown in Fig. 19. The calculated

alumtnum structural temperatures agree reasonably well

with the flight data from re--entry to touchdown. The

effects of internal radiation and internal convection were

found to be significant.
For convectively cooled structures lhe heat lransfer

in coolant pa_sa_,es must be considele(I. "I'he dotumant

mode of heat u'anMer in the coohu_t tlow is forced

convection. An engineeril_gmodel of llow in the coolant

passage is typically used. Tile engineering formulation is

based on assumpi, ions that produce a one-dimensional

energy equation with tile bulk temperature Tf(x,t) of the

coolant as the fundamental unknown. The coolant ener_'

equation takes the form

_ _x(kfAf_xI+ri_cf "6"7" 0Tf _hp(Tw_Tf)+pf cf_c)Tf =o^ (23)

where the subscript f denotes fluid quantities, In the

above -kf is the cross--sectional area of the coolant passage.

rh is fi_e coolant mass flow rate. h is a convective

coefficient describing heat exchange between tl_e wail of

_he coolant passage and the coolant, aud I) is the coolant

passage perimeter. Structural Lempetatures and coolant

temperatures are dete,mined b) .soivin_ tl_e energy

equations. Eqs. 120) aud 123)..,mmhaneouMy Further

detail_ of the formulation and mm_vlual exauqflP', are

d¢_cLilx'd b) l'hul utuJL ,rod \\ ietill_ Ii1 I{elL-. tiU-T i.

Over the la..t Ibm (h-,cades there hay(' t)eon

_i_nifi(ant advances in {o_nputaliunal methods fur

structural heat tra_sfcr. [he use ul t-omputez_ and

computer graphics has made analyses of complex thermal

s_ructures a routine s_ep in the desizn process. The

proceedings !ref. T2i ul a conference on computational

aspetrs o! heat _ansler in s_ructtues held at NASA

Lanzle) in 19_1 d_,scribes (apahiiity ro (ompute

temi)elatures of flio.ht vehicles. Cumpmel hard_are and

soft_a]e have ma_-le aud contiuuu to make ;iglHlicallt

advautes, bolt_a]e is based on the fimte

diffeleu(e/lumpP,[-p,_ramP'.ot u_--t hocl ,n ,be finirp elemeu!

metitu(i Pro_:au]., ba_ed Oil lit! '_orFllel ineti_uti iudu(le

TRA$'fS. XII_AS and SINDA. TRASY5 has been used

exte]L-i_elv fol U. 5..q)acecra:t tadiatiou heat transfer

anal_.,is s_nce i972. There a_e seve,al wide!y-available

finite element programs with heat transfer capability

i,_cluding ABAQL S. A.X 6'_ S. EAL. MARC.

MSC/XASTRAN and PAq-RkX. (.'H_tent computer

[iald_r_te [ttmds inclu(lt' iitlpiu, Hl',g'cu_ul)Ht,_tiunal :peeds

I))_ec_ur _n(I/ozp,,taIM l)lO(e._,il'_

Thcrm,d St resses

Thermal str_'ses in a s_ructure are determined by

solving the equations of motion, gq. (14) aa_d the

strain--displacement relations, gq. (16) simultaneously

with constitutive equaoons relating the stresses and

strains. One of the basic assumptions in thermal stress

analysis of flight structures is that the inertia forces in Eq.

(14) can be neglected, In most practical appLications the

thermal response of a structure is relatively long in

duration compared to characteristic times of the'structural

response. Under these circumstances, the structure

responds I,o a time--v_ying therrrufl lo_d in a quasi--static

n_nner proceeding through, a su¢ce_ion of equilibrium

positions without oscillations. Thermally induced

oscillario_ occur only when the thermal response time and

the structural response time are about the same. Suddenly

applied thermal loads of high intensity to thin beams,

plates or shells can induce oscillations (._ee ref. 57), but

typical flight ._tructure_ r_pond tu aerodynamic heating

quasi---star call)'. Neglecting inertia forces h_eans that in a

quasi---static thermal stress anal._sis the equations of

motion reduce to the equilibrium equat ions.

Bi= 0 (24)

The equilibrium equations are _lved subject to boundary

condi:ions specifying either the displacement compot_ents

or surface tractions on all external surfaces of the

strut_ute. The initial (t)nditio, s include specifying value:,

of the structure's displacements, and if the structure has

I_-,en ,ubject to previou-', load',, initial value_ ot _ttm_._e.- and

strair,_ may be required

For linear, elastic behavior the stress components

are related to the strain components by generalized

Hooke's law. For a homogeneous, isotropic material this

constitutive relation may be written as

ai j=: ,\ _ij(kk +2G _ij-q 3A+2G )tSij(l(T-To) (251

wllet_ <'.. i.,, the Krone_ker deha: ,\ and G are the L,,r_,'

U

const ants: fl is the coefficient of thermal expansion, and ] ,,

is the reference temperature for zero thermal stress. The

Lame' constants are related to more familiar engiueerit_'-'

_uu_' attts by

uE

\=( t--, c)(l-2r) ,:26,

I.:

_',=

whe e L i._ the modulo-- ul t'la._tltit3, and t' i_ Poi.--Onlaliu FuI .-luall leu,pc_atult {i,,ugt'.- _he ei,_tic i)topt'rtw,

are :onsta.nt. but for large teml)erature changes the el,_,

properties are temperature--dependent. Thus il the

tem3erature varies throughout the structure, tlao

plopc_ies vary from point to point. For small st:e.,se>.

stratus and/or temperature changes, the behavior ul

structural members is elastic, and the solntioll oi

bou _(I,_ .v/ nitial value p_oblems makes up the field _)I

thet nueiasticitv (ref. 51'-o01.

One of the most significant developments oL the la._

four d_:cades is the fiuite eiemeut methud. \Vith tinit_

eler_aents, virtually any complex structure can be modeled

and analyzed with a hich degree of accuracy. Literall_

dozen_ of finite element_books exist and a vast literature

comprising thousands of papers describe finite element

me_hodolo.gv. Although the method originated for aircraft

structural analysis, finite elements enjoy success i,l related

thermal--structural disciplines including heat transtel ,rod

COml)rt_ssible flow analysis. Previous sections l_a_e

lllt':ltiOllLqJ these aplJlitatitJll dtitt ._i\t.'li =,elected tettqc, ucc'-

l:iuite element therl_lal stie:_s analysis capabilit_ exist., i, ,,

varie! _ of commercial software including the _ i)d,.,

inetltioned in the preceding section on structural heat

transfer.

Under high rates of loading with the material a_

elevated temperatures, flight structures will experience

inelastic behavior that includes rate--dependent plastk'

(viscoplastic) deformations. Most metals exhibit

viscoptastic behavior at temperatures above 40% of thei;

m_ltiug temperature. One of the significant advances in

the last twenty years is the development of constitutive

models that include both plasticity aid creep in a sing, e

sel of equations called unified constitutive equations.

Several investigators have developed unified constitutive

model.,. The models are ba-_ed on microphysical behavior

Uk,_: ....... PAGE IS

OF POOR QUALITY
of materials, are guided by phenomenological

considerations, and employ concepts of continuum

mechanics. Typically. the equations use the concept of

internal state variables to repre_nt the evolution of

material behavior. The equations also involve a number of

material paa-ameters (some of which are temperature

dependent) that must be determined experimentally.

Reference 73 contains _rticles describing several

constitutive models.

One of the most well known of the constitutive

models was introduced by S. R. Bodner in 1968. A survey

article in Ref. 73 by Bodner describes the evolution of his

unified constitutive model. The approach has become

known as the Bodner-Parton constitutive model. To

consider the time dependent character of the behavior, the

equations are _vritten in rate Iotm The _otal _train rate

tensor is separated into ela>tic and iHeia_tit coIHI)ol_ellL_.

E 1

i..=¢ . ._t..

U I j L!

_27)

where a dot denotes differentation with lespect to time.

The elAstiC component indicated by the _uperscript E is

related to _trecse_ by the time derivative of Hooke's law.

Eq. _25t. The inelastic compouent is derived beginning

with the Prandtl-Reuss flow rule of plasticity. The

ilwla.-ric straiH coinponeur i_ repte..ented by Bodner and

Parton as

i= 14i _2S)

where S.. is the deviamri¢ st_'s_ lensot, and .J.> is the

tj

-etUlld devidtorl( sliest- ill\alial_t ]'tie co,:fficiem D is ,

O

[JlatellHl [)alan,s:tar. 11i, il tel:lE)t,Icllllle (le])olldellt material

])drdllltqel. a[Id Z is dJ[ /illl._llU_l] ]UO([ Jii>rot_ (lel)el_dent

i)atalnc,,or _alled whe iHretn_l _t,_t_, x,Liial)],. The exOiLltiOll

t'tlU,-tl loll tot ] ll_ ?-Idle _, dlinlJi_-' hH ]-_ll tlJplt :ldl ([(?Itlll ° iS

{Z-Z._[rl

2=,nI(ZI-Z)\VI-A 1Zl [-_1 -j (29)

with the initial condition Z(o)=Z o In this equation, m I.

'Z1. and Z.-_ are specified parameters: .h I and r 1 are

tempe,ature dependent pa,-ametets: aud W 1 i._ the inelastic.

work late,

Wl=aij_i] (30)

In contrast to classical elasticity and plasticity, the

Bodner-Parton constitutive model assumes that for all

load levels there is inelastic strain. In ',ddition. the model

does llot employ a yield criteliou, hi applications, tile

inelastic strain compouent is snlall ill compari.van to the

elastic conlpollellt al lOW load levels and becomes

significant only when inelastic phenomena become

prominent. A NASA-Lewis spo,lsored research program

(HOST) conducted by the Southwest Re_earch lnstitutie

recently concluded a foul year effort (ref. 74) to furthe_

develop unified constitutive models for isotropic materials

and to demonstrate their usefulness for analysis of high

temperature gas turbine engines. One result of this study

is material property data for nickel-ba-_ed alloys over a

wide temperature range. .M_terial property data for the

Boduer-Parton model is available for Ti, Cu. AI. Rene' 9.5.

IN-100, Inconel 718. Hastelloy-X, BI900+Hf and

Mar-M247.

Unified constitutive models have been implemented

into finite element analysis by a number of researchers and

iu tile commercially available MARC pro[tam. The paper

(ref 7.5) by Thornton et al. contains relerences to these

finite element viscoplastic analyses and describes

viscoplastic analysis of hypersonic structures subjected to

severe aerodynamic heating. A thermoviscoplastic analysis

was performed for a convectively cooled segment of the

scramjet engine structure for the NASP, The quasi-static

thermal--structural az_alysis was performed with a finite

element mesh (Fig. 20a) for a nickel super alloy

,ep, e.-ented by the Bodner-Parton constitutive modei.

Fig.Jre 20b presents of the temperature history on the

ae_odvnaJnic surface indicating the very high temperatme

the niaterial experienced. Figure 20c presents the history

ot t iw normal suess component at tile aerodynamic

surtae_. Note the material yields early in the response.

After the temperature drops there is a rapid decay ut ti_e

high compressive stresses, and the analysis predict,

re._idual tensile slre.s in the i)l_,titall.v deformed _utlat,,

lie_en_ thernlal--_.ttuciulal analv.,es of hvdrogen_,ded

h,adiu_ edge concepts fur h)'lVe_'>om{ fligilt xehkt,,- ate

de., ribed ill ref.,. 7U-77.

1 I,'rmat--$rr_Uutat Test m,.:

Tl_ermal--structurai testing rem_.ins all important

step in the development of ,tructures for high slx_ed flight.

Tel facilities exist at .,everal NASA and Air Force

ins:,alIatiuns as well as in private industry, but the

priwipal 1_'. S. eovernment facilities are at the Air Force

[-ii.,.i_t Dynamic'., Laboratuly _ref. TS) and the NASA

[)l/den l_light Re,,ealch F,_ci]iD (ref. 79). Two recent

.<udie.,. related to the NASP are good sources of

i_tl0rmation about past test programs and current

c,q:abilities. In Ref. SO. Inger P. Friedman et al. provide

an a_,sessment of thermoelastic analysis and testing

,tlHHic,tble to NASP: in Ref. St. H. A. Hanson and .l J.

(.'a.ex describe ,_ stud_, to determine high temperature

(1CO0-:]O00 OF) tal)ability tor testing full-scale aetu:pace

_ehitlestructures. Over the vcars NASA Dryden it,t: bet, H

inxol_od with exten,ive ilioht and lat_ratu,y test

I)tt,graill_ for the X-IS. the YF-I2. tile Spate Shuttle atl(]

h_ )er-o_lic structtlral cOihl._HentS

An excellent review of hot structules test and

analysis technology is the proceedings (ref. $2) of a

,o_kshop I_eld at NASA D,yden in Novembe,'. 19S8 In

.,Hmmarv V. Michael DeAngelis, the workshop cooidinator.

made th'e following observations: ( 1) hot structures testing

is expensive requiring sophisticated computer control

systems, Large amounts of instrumentation, and high power

requirements, (2) it is time consuming because of the need

for extensive instrumentation checkout and calibration.

and (3) test procedures axe in an early stage of

development for new, high temperature structures. He

noted that correlation of test data with analysis is

becoming more difficult as Mach number increases

because: (1) structures axe becoming more complex with

¢_ew materials and active cooling, (2) computational

complexity is increasing with the need for finer models to

captme thermal gradients and high local stresses. (3) rest

requirements are increasing, e.g. more instrumentation is

needed. (4) measuren,ent, capability decreases with

illCtedMllg temperature: e.g. accurate strain llleastirelllellt_

at.' particularly critical, and (5) test capability decrease-

............ /' t. _"_[" IS

0_, r Pi20R QUALITY
.significantlyathightemperatures.

A conclusionthat maybedrawnfromreviewin_

thisrecentliteratureisthatthe,'eisasignificant, realneea

todevelopnewhightemperaturetesttechnology.This

technoloKvis neededto assesstheperformanceof new

materialsanddesignconceptsaswellasto validatetile

analysistoolsrequiredto developthernmlstructuresfor

highspeedflight.

_@n(.hl(litl _, Remar_

This paper .,urveys l)rogres_ iu _ht'rnml sttuctutes

from the early dav_ ul supersomc flight to the current

research and development for the ._a.tiona] Aerospace

Plane. Fundamental concepts of aerodynamic heating,

aerothermal load effects un flight struciures, design of

thermal structures _ well as heat transfer and thermal

stress analysis and testing are rewewed. Major advances

in technology have occurred that provide tile foundations

for the design of thermal structmes Ibr flight vehicles in

the twenty-first century. Much progress has been

actomlflishTed, yet there are a nun,bet of research needs

vhat .rest be adche,.,vd:

• fhe X-15 wa.- tile iitst. ,u,I to date. the unh

wta]ll]ed _ehicle capable ot ll:,]i_g atmosl)heri_'

missions at .Math .5 at altitudes of 100.000 feet or

hio_her. The last X-l_ fl_ht was in i')68. The

XZI5 lliuhts were elmrm.ilqy successful making

many si_l-dth_aut (out[il,uti.,_t- to the understatldhlg

of hipeptonic fli°J_t tu(lu<li_ _he (L<:sigI_ ot thermal

<,' L'tlCl 1.11es .\It ht,H_la t!_t, _l)a(e SJltttt[o [1&5

_,olIcllblllt'd >i,_,Itili_a_i1tlx I,, the ll.41iUI£'; sl)a_e

])ID_I dill. rhl't I" i_ ,, _OI', -1 1_112 !IV'I 'd ]Ill all

,,:..petm+e::t.t. h'.i,'t-.+It fli._,t_l _,,.lllt]('. [tJ lt'llidill

Ihe x_utid Jv,-,tlet ih hi..:,h -i,',',i t]iv,ill. tt_e IJdLIOll

mu.',_ not t_itet iu etfutt> tu _h_x.:iOl., the SASP

• Significant advanLes have occurred in

computational fluid dynanfic.- and computer

hardware that pernfit high qualiu solutions for the

Navier-$tokes equations. Accurate prediction of

aerodynamic heating lof two---dinmnsional flows is

i)u_ailSle, but the prediction oI aeludvnamic heating

in tl_ree-.dimensionaI flows _emains a challenge.

• The impottan(:e ot imerac_ions between high speed

flows and hot. deforming strttctmes i_as been

recognized, but analyses of coupled

flow-thermal--structural interactions is in an early

stage of development. Computatio,ml studies of

interactio,_ effects should coutint,e. There is also a

need tur experin_ental data to SOl)port

computational st udies ot interaction effects.

• Substantial progress ha_ tx_en made in the design

and develop,ne,_t of convettively cooled structures

for high speed flight, but there is a clea, need lur a

flight test program to validate the designs under

realistic conditions.

• The need for new, lightweight materials with

well-understood behavior for structural

applications above 2000 °F is critical. High

material costs currently limit experimental studies

particularly in university research programs.

Substantial basic research ._tudies of new materials

such as metal matrix composites a_e needed

Thcrmal-_tructural analysis capability with the

finite element method h_ reached am advanced

,tage of development. New developments in

¢_nstitutive modeling permit aaa.lvsis of highly

nonlinear material behavior. For _:omplex fl_ght

structures, analysis capability has exceeded the

available experimental data base. There is a need

for high temperature experimental data to validate

analysis capability-

• New high temperature test techno[o_' is needed to

support experimental studies of new material_and

design concepts at elevated temperatures.

Development of methods for accurate measurement

of strains at elevated temperatures is a high

priority.

Many of these i-econm_endations should and are

I_in_ l)urstied to sUpl)ort the develol)meut of _he N-kSP

In a-<hlition, the nation mttst tnainta_t_ a l)_oad-b,_sed

lunclal_ental research I)l(Jglat_ flit'hard R Helciv_te!,

Looking, back in 1982 on NAL'A t i_ermal-sttut_tu_l

re>earth from 1948 to 1938 made the argument for basic

research very well: "A healthy research program must

provide freedom to explore new ideas that have no

obvious applications at the time. These ideas may

generate the technolo_,v that makes important.

t,tmnticipated flight or veh_}:le opportu,fities possible"

.At knt)_l,'d_emeut .

The authol i:. xe_ i)leased to acknux_ledge the

gllbl ill I()d lesea.lLh _.tlppOtl ol 1 he Aelot ilellllal Loads

Bta_!_ at NASA L,u,o_le_ Tire technical advice and

encuu agement of mv te(hnical L'nunite,. Allan R. \Vietmg.

over these years is g'_eatly apprecmted. 5tudents from Old

Dotal,ion {.'niversitv. Texas A and .M University and the

t.niversitv of Virginia contributed to the l)aper. Their

enthtt.ias]n fo_ the I)ronuse of uu, nation's space prog,am

make:, the authm's modest e!fo_ts _xorthwhile. The p.l)e_

\Xd-, I_'Oall w[lih + lh(" ,ILll[lOl _rt" d _l_.ilIII_-,(ho];tl cll tht'

I tLi',,_-it'+ Ot le,_as ,,it Atl_.,i.t lh, ,,ulh.: +, '-'tatll] '+,{ I¢)

,It };lit;'.', te_l_L' I he _t,iwI t)$['., htJ-lnl,,'.+l _ ,,l+(t -t IUtl_ !),t'[ -,Itid]

,Ul)ptlt oflqote'-o_ .I l i+>le) Ode+;

I_efet ent.ef

1. Hallion. Richard: _ul_rsonic Fli-_ht. Tile $tot,+ ot

the Bell X-1 and Douglas D-558. The .Mac.Mfllan

Company, 1972.

o Huff, N. J. "The Thermal-Barrier-Structures"

- Transactions of the .-_nwi-itail Suciet.v ol Mec}_atli_-,d

Engineer.'.. \ul. 77. No. 5..July L935. pp. 759--70_

.1. Heldenfels. Richard R.: "Frontiers of FLight

Structural Design", A¢lqnautics and Astronaut tcs.

Proceedings of the Durand Centennial Conference

Held at Stanford [:,fiversity. August 5-8, 1959.

edited by N. J. Huff and W. G. Vincenti, Pergamon

Press, 1960.

4. Miller. Jay: The X-plaaes. X-1 to X-:_9.

Specialty F;ress Publishets and \.Vholesalers, 19S.3.

5. Huff. N..l.: "lli_,h letltl_.'lalure Elie(.ts in-\ill tall

Structure.,". Apl)lied Mtxh,mic_ Review>. \ul. 5.

No. 11, Novemix_r. ts155, pp..153---156.
6.

i.

8.

Bisplinghuff. R. L.: "Some btlu(xural and

Aeroelastic Cunsiderations ol High Speed Flight",

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of the Ae)onautical Sciences. Vol. 2:]. No 4. April,

]956. pp. 25_)-32'L

Hoff. N J (editor): lligh 'relill)ela_ule Ffl'e(T- ill

Aizcrafl _I_u(uHies. Publ_she(l l¢)J lhc .\dviso_5

Group for lle_-ealth and Devclupmenl .Xolth

Atlantic Txcaty Oigan_z,_tion by P,.qgamun Pze.,s,

19,58.

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0 Hallion, Rich,tld P.: "X-L5 tfi_h,sl ,rod F(_stesl

of Thent Y,II". Flight lwei]l,_liuu;d. Dot 21 197&

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IO. Becket. John V.: "lht' X-J5 I_'lu_iam in

llelluspecl". R<_uu_fahurfulbclSun,_,,. .MalclL-April.

i909. pp 45-3:3

i..

i2

11.

.=

1 I_).

1-

lS.

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__Sol-(red Bll)ii(bVai)h_. 'g:\",-_ bP_,0. IHh3

Gl,set. Pcqe_ E tedilul_ A,':_,i'.,,,_,,i_._i]_ 14,,,,ne(I

.'_IH(illIes. Plwoedin_s ol ,_ tOllIelt'll(i_ si)um, oled

bv tlie Air Foist' Olfite I)l "_,CI:Iil'R ke'.calth ,'old

--k_thur D. Li':ie. Ila n,.hI ,< (_'a:nl)lid_e.

XIassachu:e_l....lul._ 25. ;961 l)lentite Hall. hic.

1962

\ l<_th,_e,,::. Ei,h,l E: "II,.,c,,,_ h. De.-i.gn

£'on.-idcq<<iu]_.-. <u_d I_ _ h_lu_a,_i P!oblem., of

S_ructule,, lul \\ill',Lt'd A_:IO:'I),&L' \ohiLleS' NASA

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H_l)el._Olm Ail \elmle-'. \v_i_._m_(. lquo--ediilo.,>

I%6 \,i 1 c_llt_.d I> .l(fl,, IJl,albluuke. ]u<m

Biuce ,<m ilubcll I_ l)..:dcl I lle plucet.dil_os ul

the Ruy<d Ae_uii,<iti(,i '.,oclei_ Cenlenarv Congres,,

in cunjunctlui_ -ith lia' thLh ('ung_e.-., ul the

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Seventies and Beyond. A ('onl'erei_(e held at

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2---4, 1971..NASA SP-292, pp. 179-192.

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.\[ay 12-14. 1986, XA_A T,M-_77,10,

19.

20

21,

'2"2.

21.

21.

23

#

.)-

29

}0

31

.J2.

3:}.

Staff of Johnson Space ('enter. Sl)ace Shuttle,

NASA SP---q0r. 1976.

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II,',,I i_.._o '1 Ix" lh,l,,Id 1'1('.-.- _ *):,ii,,lI } . lUoU

AmJe:.-o,. h,i.i. I) .Jl [l',11,_l..,)llil an_l lli,.i_

lo!_.l,-!_t,ll_- (i<_<. ll.n,_mi¢< Xlt(;_ax,,-ltill Book

Comport',. 19."0

\an Dlie::. E R "In_e,'ig,_:iun of L,_illal

Boundary L,_'.e: in ton,!):e:,ible Fluid.- L's_ tl,_

C:u(tu _'h.:l,_,d" \'\_-_ IN 2;' 7 Jail i_52

Ecke:! E 14, (i "ti_.,,(._,::_ R_,i,_t',oI_ lul lle,_:

]lall_tei a;,d }:lntmul lu ]l*.'ii-\_lu(_l_ l(_,_t:,_:

al]d ] U_:,',"_l,_ Bo,il,l_l '.-L_'I Flu- O,rl

_.tll_dtt'_ ',', I'..1, (.,/iI2tdli_ IJ[U_:'_I;L _ ,_:l(.i i _llii)_'l,_l illL'".

rlans, of the" -_XIE. \'oi 7-. Nu 6. Aug. 1%6. pp

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rhermophy.-ic_. \'ul 1. No. 1. ]an 1957, pp. 3-12.

Dwoyer. Douglas L. l(atlel. Paul and PoviHelll.

Louis \ "Re_uotit;g t'FD fot ItSi)e:-ol_t

Aiic:aft". Aeru.-pate A_lern.,,. Octobel !t_7. pl )

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Shang. J. S. and _(he_r. 5..l.: "Xavie_-S,okes

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Dvnami_.s ("unfeie_(,., Cimi)mati. Ohio. ,Jul_

15-17, 1985. AI-_A P._l)Cl Nu ,"5-1509

Watts, Joe D. "l]ight Experience with bhock

lmpingen_ent and lntcltelelae Iteatln_ oil Ihe

X-1.5-2 Research Aiiphme'. NA%_ TX] X-It,6CL

October 196_',.

V¢ieting. A. R. and Hohlen. M. S.: "Exp{_imental

Study of Shock \\axe lnterlerence Heatin_ on a

Cylindrical Leading Edge at Mach 6 and 8"7 .a,IAA

24nd Thermophysics Conference. Honolulu. Hiwaii.

June 8-10. 1957. A1AA Pape_ No. 57-1511

OR!GiN,_L PAGE IS

OF POOR QUALITY
H. Klopfm. G 11. ,_u(I Yee. 11. I.: "Vis(ous

Hypersonic SI,o_k-Ou-Shu_ k llli,'I,,c l ilHI 011 Bh.lJ]t

Cowl Lips". AIAA 2f, th .\elu.-pate Scie,ces

,Meeth_, Reuo, Nevada, Jauuar_ 11-14, 10SS,

AIAA Paper Nu. SS---0233

:35. Klopfer, G. H., Yee, H. C. and Kutler, P.:

"Numerical Study of L;l_steady Viscous Hypersonic

Blunt Body Flows whh au Impinging Shock",

NASA TM-100096, April 198S

16. Thareja. R. R. Stewart. J. R., Hassan. O., Morgan.

l(.and Peraile, J.: "A Point Implich Unstructured

Grid Solver for the Euler and Navier-Stokes

Equatml, , AIAA 26th Aeto:,l)ace Sciences

Meedug, Renu. .Xe_ada, Jauua[_ 11-14, 19SS,

AIAA Paper No. SS-003b

]. Dechaumphai. P.. Thornlon. E A ,rod \Veitin °. A.

R.: "Flou-Thel _n,,l-S Iu( _uI,,I bludv of

-_erod>uaHmall) He.ted l._',,ii, _- [.d"c,_"..]oulnal

u[ blJ,_teulall alld l_t_l,_'- \,,, 2(,. Xu, -I.

.lul)-\u_ I')"q pp 2_)l-VtJ!'

.IS. Tauber. 71. E <_ml _ I.::.,'_'< G. P.:

".-_el ot ht.'l lllut [) i m J :ilt .- _I I i ,, h.','l lilu'l)]Wl hVehicle.-". Jom_;al ul \iii l,dl. \ ot G4. .\u 9.

Septembe_ I987. !)p. 5'j4_.,1/2

I'_ Tauber. M E ,,lLd \dol_l,,;_. II (.;.: "Thelui,d

Ellvirolimelli iJ[ ] I a h.-nl L:,<J-ph,q it \c'hic It>-<''.

.Jouinal ot Aiilialt. \el 7] \,) t.-\p_it i'JSS. I)P

155-.JO]

40 _gii]clele_. I L.. \\eld_. (; L .... l_d l_i!iu_Ai,. (' M.:

"\ e!i!iodioi_ I,.'.-l.. el D:_:,_Si0 ]l:,u ::l<d I)',electiol_

_ \ .- I. t:'l li ('olI(.C])I £". .]IJ]Z !]_[ O[ _,l),1 t l't rdlt illld

R'ock_ts. \ul 22..No_-IJvc lq.<'j. 10 39S_i04.

4; rhorlitOU. E. A and Dech<,_m!)h<_!. P: "Coupled

Flow. Ihei m,,l. <<M <.,!i ;t! tl',,_i -_li<l[\ 5e._ Ot

.-_erudyllamicall) Ht,aied P,<;c'{-' .iou]nai ut

-khcraft. \%1 23. X<J I1. X_,x ['iS-. Ig; lti59-1059

i2 (;,_ii](k I l;: ".\ %i!_,'_ ul \,!_,".,!n_l,'Li,(i_[iv".

-_elo:lJ,.ttt, Lll_llit'eLlil_. \ol 72 _<J i. ,Jail iU0;.I,

pp. 140-147

4..I. Reed. Ill. \Vilmer H.. H,m,un. Pet;) H. and Alford

Jr., \\. J.: "A, sses.,nient ul Flutter .Model Te.-lin¢

Relating to the .Naliuilal .-_elu-pate Plane", NASP

(.'ontractor Rel_vt 1002 Dvn,mfic Engineering

lncorpor,_ted. July 19_-7.

.14. .Mixson. John S. and llous:o_, Luui- A.: "Acoustic

Fatioue: Oxelvieu- ul .-%ttlvi'ie:, at NASA

Lal_gley . NASA T31 _914:L-kl)ril lOST.

45 Locke. J. E.: "A Finite Element Purmulatiol_ for

the Large Defter tion rla ndoi_l Rest)onse of

Thermally Buckled $1rtu.tules". Ph.D Dissertation,

Old Domhiion Univer.-ily. Xultolk. \'A 1985.

-it,. Locke, J. E. <rod Xlei. t'.: "A Iiuite Element

Formulation for the I.,_ _,. D_fl+.tiiull Ralidum

Response of [Iwlnh_l ( llll( i,h.d Bt'aills".

AIAAIAS.MEtA$CEt.klI'_'-\<5( ' .l_lth _11tl(t u ff:'$.

$iltiCtIJlai [)_il<lliiic> ,tll(i \l,ill'l:.l.- ('Oil[t'leliLe.

Mobile, Alal2ania. AI)_iel .i-'J. 19_!J. I)p. it,94-1700.

AIAA Paper No D.q-I VJI-('P

47 Hanker. \\d!,:,: L.: Ib_l:]_llrv 4ol,}dvHaliH('_.

.\[A.-%Edlital;,_;. %clips. l!jb_

48. Shih. Peter K.. Prulity, Jack aild Mueiler. Ricliaid

N.: "Theil,u-':mtul<d ('uucepls tot H.vpervelocity

Vehitles". 2'-":, $tructu_es. ._lructul_[ Dvn,_dcs

and 31ateri<ll- (lill[_.lt.lllt', \\ illiam_l,ul_. \'A. Al)ril

IS-20, 195.'. pp btl-4,$_. :\I\A Papt'i No

$$-2?95. Fo: ,_ i_,l,,I<'(I ,t:i_lp by I>l,ler Ix. _hll, el

_1., see "7',-" ._ lille.', lot H) pel_ehxity Flight".

.Aerospace .America, Xla) 195'J, I)P. 7_-.]1.

49. Pritchett, T R.: "Advanced Technulog3

Aluminum .M,_terials fo_ Ae_uspace Applications".

Light Metal -%,,0 Vol. 44. No., T-S..A.u-,. 1986. pp

tO-14.

50. Steinberg. Morris A.: "Materials for .Aelosp<ice".

Scientific Amerit,_il. Vol 235. NO. 4. Oct 19S(,. pp

66-79

31 Rul)inson, Jai:.e_.J.: "(;et_iug the _l)ace Pl,tne O1[

the Ground". Juurnal ol .Metals, Vol. ]'). _u. T

July 19ST. pp. S-9.

",..' J,ltkSolh L I;<,1,'i_. lJix_Ji,. Bidut.) C'.. Te:.:.v_

D<_rlel R.. (',.:,'_ Al,,u 1. <rod '_iel/,_en.-. J,_.-,'p!: I_

"Ilypel:u:m " : i_lhl,'.', ,_i:,l Maleli,l]', -_ l)ilJu:_'/>

H(.I)()II 'l " g\t,.(* ,,lie .%llliql(,l. {Jl : }')'_t i.qJ J4--111

"*l ('tlllllllll. I"':": "+\, ",)_,1_,1(_" \l<_lelh_]-". hstllt,,_! _JI

kle',,d;. \u: ;'2. \u ", XI,t', !,l-_, Pl; [U-14

I Ramel. RI< !.,_:+I \ ,li_d Ek_,dl. ]cJiln C

"Eleva ted-T_'r.;:_er,_ _v,re .-\llu_ s fol :\il(ialt

Structure". Jo :::_<d el XIc"<d.- Vol. 40. Xo 5. hl,_)

19$$. pp. i6-;--

-,_l Destegani. .J,:;,,_e, 10 "Adx, u_o.- lu

Intel-3let,_ ZT_ '. AU, ,_ncc.d Matel h_ [': ,,nd

PlOCe.%es. \u. ;15. X_>. 2. Feb 19_¢'). pp _?--li

% Ihmald. Tew_:.,c' \1 F: "-kd,,,uited \l,ite_t<_i- !u

Fly High 11; \ASP". .Mi_,<lced ).lat('iials ,llid

P(ucesses, \u; 115..Xo 5. XI,_ 19S'7. I)P .Jg-lT

5T Bolev. Bruno A and \Vuiilei. Jelome H: Tl_e,J:_ .I

Thet:mal S:Ip_: _-. Juhn \Vilv_ ,_nd Sons. I%0

"j- Hetlidr.-ki. II!t;.,_id .13 {,'d,!,,n ': ] h,.in_,l', ":I,'--_"- I.

_t.ll'lh 14ul',,_:,.. i'i';I,

"r' tletll,_l.-ki. H,,.i._d It ,._iTi,,i. li,,'rl .,:! "_: .... '-

11. _ulth I1_ ....... i. I'I',T

t>O. Nowirlski, .J.L.: lli,'_llV o1' Tholmoel<_-li_il_ ,_i,i_

Apl)ticalions. Sijhoil alld .Xoidholl ]iiteilia' ,_]

Publishers. 19-S.

61. Argyris. J. H. \'az. L E. aim \Villain. I( J.:

"lnteo.ri_ted F,i_ne-Ii',ement .-\hal)sis tit COUi)lp',l

Therl_lovi.-cup._,lic Pioblt'lll,'". jotlill&l of fh,.':lll<_l

Stresse.,, Vul. 4. 19$1. pp 121-153.

i,2. Allen. D. t1.. "Pit, ditlc'd Axbd [u'mi_elature

Oiadient it, ,_ \iSlol)l,_-'lic t'niaxial [3ai due to

T hernlonleL ha::it ai (-'tllll)lill_". ]lltel li,q iuual

.Journal for Nu:-:m-i(.al .Methods in [ngh_eerin_. \oi

21, 1986, pp. W.I-917.

Og :7t ?:; (,L R,._,'r_
63.

64 •

bS.

o6.

f,7

hS

hll

T0.

71

7:L

73.

Ghoneim. H.: Thern_uv:S_Ul)lasticit.v by Finite

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4, 1986. pp :]-15- J5S.

Tamma, l<u.mt K. and Railkar, Sudhir, B.:

"Evaluation uf Thetmall.,, luduced Non-Fourier

Stress \Vave Disturbances via Tailored H)blid

Transfinite Element Formulations". Ptoceediugs of

AIAA/ASME/ASCE/AHS/ASC 30th Structures,

Structural DyLiarnics and Material Conference,

Mobile, Alabama..April 3-5. 19$9, pp. 285-295,

AIAA Paper No. S9-1190-CP

Tamma. K I<, Xambtnu. FI R, and Glass. D E.:

"E,,'althltiu. ul Non-l.<mI,. llcal \\a_+,,. Itdluemed

I)', Nunlinle_l/I.iueat I]_ml.l,,;', l-_fll.ul'. Empl,.r, inl_

all EXl._li¢i t .\l_Jlilut I lit' dHd ('urn tutled

Stabilizatiu,". 2>lh .-\,..lU-l_,,,.v :5(ien(e> Mevti,g.

Reno. Nm,d,. l._Im,n', S--t!. Its')0. AIAA P,-,pet

No. 90-0.545.

Ko, \Villiam L. Qumn, Robert D. and Gong.

Leslie: "Finhe EIeme,t Ruentlv Heat-Transfer

Anab:gs of 5+.,a_,-, _,hm_lc, Orhiner", N..\SA

Techni_.al P,tp.el 2637. D,..',.. I'L',0

Ko. William I.... n,] Ohm,,. Tin,oth_: "Effect of

glel]wnt Size uil lh,- _tJltltltJJ{ .-_{(UtdLieN LIt Finite

Element H(,,_; "[l,_.,.-lel ,rod Thelmal S_ess

.Ala._l',ses ol _.p,,,.e Shuttl, + Otbilm". N:\'5.'A TM

>S29_2..ALl'.5' lUST

l<u. \Viilimn I.. an,l F'h..hJ:. Rugel A.: "The,mal

Stress Anaiv.-.is of _p,_{e _.],,tt]e Otl)ite, :_ubjected

to R(v,tl_ \(.t,d,rt.,la_it ih.,ril_". .XA>A TM

<S2St,. O_ t" i'r',,

lhutnlu,. E \ ,,:,d \\i,+:.:,_. A. I{ "tinite

EJelli_qll \]' [ [it_lh_i()_ + l(*l I ;rill'fell{

(..'onduttiotL"Pul< ,.(I (oil', _'( _t, lliet real Anal',sis".

Plo_ie, ¢, +i} .A-t:_nJ,_u_I¢, ,,+,,J A,._ur,,_,ei,.-: Heat

ltaaslet. ]het_n,_i ( outt+,i an,I Heat Pil_;, Vol. T0.

edited I.)3 \\ _lte_ B Ol.-tad AI.-\A. \Vek, hi,gtOn. D.

C. 1989. pp. 77-10:3

Thotntou. Earl A. _n(I \Vieting. Allan R.:

"Evaluation of Finite EleJnent lrulLmulations for

Transient Co,ductiun Fur(ed Convection

.-_nai,',.,,is". NULII,'Ii(_I Heat Transfer. \oi..J. J_O$0.

Pl). 281-293

\Vieting. Allan 1{ ,rod fllurl+tUn. Earl A.: "Finite

Element Thermal :\nah>is ol Convectivelv-Cuoled

Aircraft Structures".._Ulnelical Methods" in H_t

Transfer. edited I)v R. W. Lewis. K..Morgan and O.

C. Zienkiewicz. John \Viley IU$1. Pl:'. 4al.-44a.

Adehnan, Ho_,_,d ._1. (5._ml)osium Chairman):

"('oln_putatiuu,_l A:pe<l, ut Heat ftansfet in

_tlULIIIIU>", JllUtUt+tJill=: _ U[ d _)llllgg_illlll held _t

NASA l.,_ugk'._ l¢c.,e,-,t,.h Ceute_. H,_l.i)tun.

Virginia. Novembe_ ,:I-5. lq._l. NASA ('onlelence

Public.,tion 2o16.

.Miller, A K. (editor): l'nifi,"dConctitutive Models

for Creed and Plasticit;._, Elsevier Science

Publishing Co.. Inc. New '_u_k. 1987.

74. Chart. K. S., Li,tdhohn. U. S, and Bodner, S. R..

"Constitutive .Modeling for Isot_opi( Materia-ls

(HOST)", Fi,u,I Report, Southwest Research

Institute, San Antuniu. Texas. June 1985, NASA

CR IS2132

7,5. Thornton, Ea,t A., Oden, J Tinsley. Tworzvdlo.

W. Wo)tek and Youn, Sung-I<ie:

"Thermo-Viscupl&qic Analysis of Hypetsonic

Structures SubJected to Severe Aerodynamic

Heating", Proceedings of

AIAA/-ASMA/ASCE/AHS/ASC 30th Structures.

Structural Dynamics and .Materials Conference.

April 3-.-3. Xiohile, Alab,,ma. pp 604--620. AIAA

Papes No. $9-1"_'26. "lu appear in the .lou]nal oI

Airt-r_fl.

:t,. .\leli.-. Marcher,. E.. (.;Iddd(.,L lleil.,e.tt .J. 5(hulxq_.

John, F.. _,Vestf_ll. l.e,onard. J. and Trin:aJchi, P,_,I

A.: "A Unique l,ter(liscildilm_. ', Re:.eaJ,.h Elfv_I lu

Support ('owl Lip Technology Development Io_

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Technical Paper 2S76.

77. Gladden. Herl×-rt J., .Melis..Matthew. E. ?,lock',es.

Theodore. f. and Tono. .Mike

"'] hei mal/Sxx Itt _+ut,d A!,"i'. -es ut S,e', ,-t _1

[I \ t]: U_Oll--I 'ooJ(',[ l.v,,dm--J_dte Col,t',.'l)t - ful

I-t).p,.'l._utfi( f'li,:nt \ehi(k,.,'. "Presented at ,h,.

AIAA '25th .\&u."i_.;,, .Me,..tiug. ReHu, Ne',,,,i,,.

],mu_ry S-ll. Ig',u AI-_A 90-0053. .Also

a_aiiable a.'. NASA l ,.clinical .Melnuiandum 1023u1

7>. Boo_o_.s. Bernard C: "The History of S_atic "les_

a,ld_Ail Fulc.e 5tlu(luI,eS T,.,stmg".'Air Force Fli,.,h_

D_ nl,_mic._ l.,,!_o_,t_ u; _ I-eclm_cal Rel.xX t.

AFFDL-fR-Tu-.]071. l'u,e !'079.

T', _,'li_. \\,,h,'l I X \'_\ t)l(hl)l.\ I11_]_; I .... _*-

I(_.'-e,_u,. h I ,,, li_ It'- ' _, A_.X 1,'_ !,::._ ,,:

\lemorandtHn ',I 16".. l)t.(t,ud),,i t9'_1

S0. Friedman, lu._e_ P.. Vosteen. Louis F., (.uope_.

Micho.el. J.. Gul_. l{unald R. _nd Reed. Ill. \\ iiJne_

H.: "Assessrne:_t of Thermoelast_c Anal,.sis and

Testing Applicable to the National Aeio-Space

Plane". Dynamic Engineering Inc.. Newp,.n_. News.

Virginia, .'qASA CR 1017. July 1988.

-I. Hanson. H. .'\ ,,nd Ca>%. j 1

"Hi,gh-feml>el,t_,.e lest ledmulug.v". Flu}d',]w

EnoineeriI_e ( orp.. ),li m_e,q_oli:. _lil;i:e-,,; _

AFVCAL-'flI-.',I,-]IUS. F,'btu,_ny 195T

82. De Angelis. V. Michael (\\urkshup couldin+H_u_l:

"Workshop on ('onrelation of Hot Structutes Test

Data With Analysis". Vols I and I[. NASA Ame.,.

Research ('enlet Dt,.(len Flight Re.-ea_ch Facilit',.

November 15-17. Iu_,S.

OR!CL:NAL PAGE IS

OF POOR QUALITY
%. • t._ T_• m_ -I,000 TO 1,500

. • THERMOGOUPLE LOCATIONS

lJ_'_[, _o°F TEMP.,'F

4.000

15) 16oi:

Fig. I Maximum me_ured temper,uures on X-IB 2,2OO

airplane.Mach 1.94,1957 (ref.12). Fie,.5 P_diation equilibrium Lemperatures for a

" re--entry glider computed by NASA in 1962

(ref. 14).

l!

__ \4_800 ,-16OO .... r3tO0

Fig 2 X-2, K Monel int_nai ._truc_u_e and stainless __00

1400

steel skin. Mach 3._, L9.56, (ref 4). :_1__ '-14C

Fig 6 Equilibrium hypersonic vehicle surface concept temperatures for sustained for a .NASA flight

at Mach 8 at 88.000 feet. 1966 (ref. t6). 11_' r

I(E__ %100 ° F

lZZO'_ _ _-13250F

"_ _ "_101_ F

Fig. 3 Measured temperatures on an X-L5 flight, ,,,,.

circa 1.965, Math 5.0, (ref. 11). '"',,

REENTRY GUOER

RESEARCH AIRPLANE

Fig. 4 Winl_ed aerospace vehicles under

eon_lder_ion by NASA in 1962, (ref. 14).

t-'i_ -

Space shttt_ le thelmal plotection's.v_tem. L970

(ref. tg).
Representative surfa(e temperatures for the

NASP, 1989 (ref.231.

Fig. 9

Candidate HOt $tructunl

J_

f ,_ Ci_(lidlll

SPFDB Honeycomb SPFDB 1"1_irrl_ally Prolected Structure

Sandwich Su_wich Corna0ation x ,

_ _ "SPFOll Mot_ycomO

Sandwl¢l_ Sanaw,ch

InlMlllliOfl

Clnllllllt,ll

ActwWy Cooled

Structure

St nnllM.Sl|f fened HoneycOmO

Skln SlnOwtcl'_

%

Inluliltl¢

Cooled Panel

U/U_

(a) Velocity profile

0:

32

>,l x

16

8

I I I I

_ M=,=20

16

2

0 4 8 12 16 20

T/T_

(b) Temperature profiles

Fig, I0 Candidate structural design concepts for the

X-30, 1989 (ref. 23).

Fig. I Compressible boundary layer flow. 1952 (ref.

27).

ORiG!fl!h'-':_ncF#',,..i _- IS

OF POOR QUALITY
Temperature

Velocity T. - - -

u. ..... BounOary Layer Edge

Tw T=, Tw

qw O_ = 0 qw

Fluid heating wall

Wall heating fluid

speed llow

,"----mow

i domain

I

!

o,

i .......... •

(a) Shock interaction.

E

SOU_OARV/ ..... / /---- \ _ 15

lOW SHOC_ COWl. lOW I_OC_ 10

5

Fig. 13 Overall Flow field for the NASP, (ref. 34). 0

\d,_l)tt_¢ zlte.',h v, ilil t.|tl,.tl_'d :,t'ttiUlL.

--F.E.

0 Experiment

I I I I

"s-_o .._ o _ 9o

e, Oetlr_S

_t_ Cuutpar(ttive ,_utla.ce [lea,ting_ra.te disLributio,

Fig 14 Shock interaction problem for cylinder. Iref.

37).
fG i3

p • _4"e mt-elnL _

.aN

14yor

-OOUMBy A--0.051 I_ Ememwn

W_

i / "-.

/ ",

,t

0 _' 4 O

X.m.

[nte:,¢ctlun ui a paue _icLh ._l,tcl_ b tlu_, ilel.

411

Fig. 16 NASP wing deflected shape considering flow

interactions, (ref. 23 t.

T,_

W

t.i_, l>

Iq._lTf IqEGIM(S OF _NIC VtHICLES

._'-i_y ILAT°IS

INSULATION X-m

I I

1 2

ExPOSUIttTIMF. hr

Flight regimes for hypersonic vehicles. (aftel

ref. 17}.

¢

OS

UN 2. am 1NO _ NO

"ii

i

_.-._k ---_ -...

I I I * J t _ *

_.mm 2.110 Sm m 4 I1_

Specific strength and

material._. _tef. 52).

_tiffness of adva,ced
illl 1 | i ii I I I 1 I_ I_/_ | I I | I | ill I I I I _

_,w (a) Finiteelement model

(a) Geometry of wing segment

-t ---- pRll --

t ,.._ _-'- i i _ -;_ 2

m

E

0

O0

I I I I I J

Ol 08 12

T_* (sic)

{bl Temperature history

(b) Finite element model

_ mP

,- ,= .-4:,

t" ! ....

(c) Time histories of structural temperatures (t:)

dU_

8aJn

I , /,

OJI _2

TWIn (me.)

S,ress history

Fig, t9 Space Shuttle Wing l"hermal :\lialysis (ref

66).

Fig 20 A Viscoplastic Analysis for the SASP {ref

75).

*-J,IdG:D,'AL PAGE IS

OF i_Lq_DRQUALITY