Thermal Stresses IV. Volume 4 of Thermal Stresses

546 Pages·1996·11.192 MB·English

by  Richard B. Hetnarski (Eds.)

Preface
It is a pleasure for me to present the fourth volume of the Thermal
Stresses handbook, after almost seven years since the third volume
appeared. In 1989, when Thermal Stresses III was published, neither
the Publisher nor the Editor were certain if the follwing volumes would
ever appear in print. But the wealth of important topics, not covered
in previous volumes, has offered an opportunity for the continuation of
this series.
This volume, Thermal Stresses IV, joints the the previously pub-
lished volumes:
Thermal Stresses I, published in 1986, 547 pages;
Thermal Stresses II, published in 1987, 144 pages;
Thermal Stresses III, published in 1989, 573 pages.
The six extensive chapters of this volume are:
(cid:12)9 Earl A. Thornton, Experimental Methods for High-Temperature
Aerospace Structures.
(cid:12)9 Shane A. Dunn, Non-Linear Effects in Stress Measurement yb
Thermoelastic Techniques.
(cid:12)9 George L. England and Chiu M. Tsang, Thermally Induced Prob-
lems in Civil Engineering Structures.
(cid:12)9 Kumar K. Tamma, An Overview of Non-Classical~Classical Ther-
mal-Structural Models and Computational Methods for Analysis of
Engineering Structures.
(cid:12)9 Liviu Librescu and Weiqing Lin, Thermomechanical Postbuckling
of Plates and Shells Incorporating Non-Classical Effects.
(cid:12)9 Louis G. Hector, Jr. and Richard B. Hetnarski, Thermal Stresses
in Materials Due to Laser Heating.
As the Editor, I am proud of making this collection of interesting
topics available in one volume. I will be happy to receive readers'
comments.
Richard B. Hetnarski
Rochester, New York
October, 1995
lamrehT ,sessertS VI
.B.R iksranteH )rotidE(
(cid:14)9 6991 reiveslE ecneicS .V.B llA sthgir .deJreser
Experimental Methods for High-Temperature Aerospace Structures
Earl A. Thornton
Mechanical, Aerospace and Nuclear Engineering, University of
Virginia, Charlottesville, Virginia 22901, USA
1. INTRODUCTION
The need for experimental methods for high-temperature aerospace
structures began with the advent of supersonic flight in the 1940s. Over the last
four decades the design of flight vehicles for high speed flight in the
atmosphere, either to and from space or in sustained flight, have posed a
formidable challenge to structural engineers. Major aspects of the challenges
are to select materials and design structures that can withstand the
aerothermal heating of high speed flight. Over this same time period,
engineers have also addressed challenges offered by design of spacecraft for
earth orbit, flight to the moon, and more recently for flight to Mars. For these
latter missions, material selection and structural design are influenced by
radiation heating and cooling that may cause temperatures to vary transiently
through extremes of hot or cold in the low vacuum environment of space. A
1992 paper 1 by the author describes the evolution of thermal structures for
vehicles subjected to aerothermal loads, and a 1985:paper 2 assesses thermal-
structural analysis methods for space structures.
Although a significant body of literature exists for thermal structural
theory and analysis, the literature available on high temperature structural
testing is far more limited. There are, for example, virtually no survey articles
on thermal structural testing in the open literature; most publications cited
here will be government documents. A 1958 AGARD paper 3 by Taylor
describes experimental methods in high temperature research in Great
Britain. A NASA report 4 published by Heldenfels in 1982 provides historical
perspectives on NACA thermal structural research from 1948 to 1958. An Air
Force technical report 5 by Boggs in 1979 provides a history of structures
testing, including elevated temperature testing, at the Air Force Wright
Laboratory. Research on high temperature structures testing for aircraft
waned in the 1970s after the major effort devoted to the manned space program
during the 1960s. However, research for the National Aerospace Plane in the
1980s stimulated renewed interest in thermal structural testing, and surveys
conducted at that time provide excellent evaluations of testing capabilities in
the 1980s. For exmnple, Hanson and Casey, in a report 6 prepared for the Air
Force in 1987, present a comprehensive evaluation of test technology for full
scale vehicles. The proceedings 7 of a 1988 NASA Dryden workshop on hot
structures provide a collection of papers that cover contemporary work and
2 E.A. Thornton
research over the preceding 25 years. A 1990 report 8 by NASA researchers
DeAngelis and Fields describe techniques for hot structures testing at NASA
Dryden. Since then there have been a number of specialized meetings with
papers on thermal structural testing, and later sections of this chapter will
refer to these at appropriate points. Note that the preceding description of
thermal structures research focuses primarily on aircraft structures
subjected to aerodynamic heating. Over the same period, thermal structural
testing capabilities for spacecraft were also developed. Literature on these
testing methods will be cited in later sections where this technology is
described.
This chapter describes experimental methods for high temperature
aerospace structures. The emphasis is placed on airframe and spacecraft
structures; the closely related problems of propulsion system structures are
not addressed. Experimental methods for an important propulsion system
problem, low cycle thermal fatigue, is described in a chapter of a companion
volume in this series 9.
The chapter will be of value to current researchers engaged in
experimental, analytical or computational studies of aerospace structures
under elevated temperature conditions. For the experimenter, the chapter
provides an overview of thermal-structural test technology as well as
numerous examples for approaches employed and issues encountered by
engineers in past tests. For the analyst, the tests cited provide sources for data
that may be used for validation studies for new and improved analytical and
computational techniques. For future researchers, the chapter provides basic
background for new studies.
The chapter begins with a historical description of thermal structural
problems encountered in early supersonic flights. Then thermal structural
tests are described, first for heating in the atmosphere and then for the space
environment. Next, test technology is reviewed with discussions of heating,
cooling, instrumentation and related test procedures. The chapter concludes
with brief closing comments.
2. EARLY SUPERSONIC FLIGH'I~
The development of thermal structural test capabilities after World War II
was motivated strongly by problems disclosed by early supersonic flights. This
section highlights a few of these flights that provided impetus for both
theoretical and experimental thermal structural research over the last forty
years.
2.1 The German V-2
Although the WW II German V-2 rocket development program has been
documented extensively 10-14, the fact that these missiles ushered in the
modern era of supersonic flight has not always been emphasized. The most
famous of the rockets, the German A-4 (V-2), had its first successful flight on
October 3, 1942. The references cited above describe the fascinating history of
the V-2's developmentat Peenemtinde, its use as a war weapon, and its post-
war flights at White Sands Proving Grounds in New Mexico. At launch, the 14
m (46 ft) long V-2 weighed 12,900 kg (28,400 lb), Figure .1 In a typical flight the
High-temperature aerospace structures 3
i
FUZE
DAEHRAW
SHELL 155 LB
AMATOL 1654 (cid:12)9
I TOTAL 2ZOS LB
INSTRUMENT
4' T" SECTION
105e LB
F EU L TANK
SECTION
EMPTY
I 63 6 LB
20'5"
OXYGEN 10,940 LB
ALCOHOL _8~452 -
TOTAL 19,392 LB
AA__.LG S S LOOW
INSULATION
~..l ..
TAIL SECTION
3938 LB
14' 5"
EXTERNAL
"--'VANES
FIRING TABLE
,~ . . ~ . ~ ECTOR Figure 1. Cross section of V-2 on firing table.
ETERCNOC
rocket reached a maximum altitude of 80 km (50 mi), achieved a velocity of 1600
m/s (5300 mph), and had a range of 275 km (170 mi). The rocket achieved sonic
speed in 25s and attained a maximum Mach number of 4.5 during its 5 minute
flight. On reentering the earth's atmosphere the rocket experienced
significant aerodynamic heating, and German engineers estimated
maximum skin temperatures to be 600 ~ (1100 ~ The aerodynamic skin was
fabricated from 0.63 mm (0.025 in) thick sheet steel.
During an extensive test program numerous failures occurred often due to
propulsion or control system problems. These were solved systematically by
trial and error. Finally, one perplexing problem remained - the explosion and
disintegration of the rocket after reentry at two or three miles above the target.
Without telemetry data, the full explanation for these failures could not be
resolved definitively, but in retrospect the failures were quite likely the first
major thermal structural failures due to aerodynamic heating. General
Dornberger 12 describes the event as ... "When the missile re-entered the
4 E.A. Thornton
atmosphere, some fluttering of the skin, already weakened by air friction heat
which increased the temperature to approximately 600 ~ occurred. The skin
burst, air rushed in, and the missile blew apart. However, this was found out
only in the last months of the war. A rivetted cuff around this section improved
the situation noticeably." Whether Dornberger used the word flutter in the
modern aeroelasticity sense we do not know; but, certainly he recognized the
degradation of structural strength due to aerodynamic heating; and he thus
identified a fundamental thermal structural design issue for subsequent
supersonic~ypersonic missiles. Interestingly, in the 1960s the United States'
X-15 did encounter supersonic panel flutter.
2.2 The X Plonos
In the United States the need to understand aerothermal loads and the
design of thermal structures had their origins in the late 1940s. In WW II,
airplane speeds had become high enough for compressibility phenomena to
have a significant role in performance. Transonic phenomena were not
understood very well, and over a period of time the phrase "sound barrier"
came into use. The need for a transonic research airplane was recognized
during the war, and in 1944 the design development of the Bell X-1 program
was initiated 15. The X-1 proved enormously successful, and the flight of
Captain Charles E. Yeager on October 14, 1947 proved beyond doubt that
manned aircraft could fly faster than the speed of sound. An advanced version
of the aircraft, the X-IB, flew several research missions for NACA to study
aerodynamic heating effects, Figure 2. The original X-1 aircraft as well as the
advanced version used aluminum construction throughout. Measured skin
temperatures for a NACA mission flown in January 1957 at Mach 1.94 showed
that skin temperatures were low, less than 200 ~ (90 ~ , 4. Thereafter,
supersonic flight speeds increased rapidly, and the need for considering
aerodynamic heating became evident.
Mmax = 1.94 T T = FO022
* ELPUOCOMREHT SNOITACOL
)
~051.
FO351 FO49
.r 136~
...~I l~~
'"i.JJ~"~ "I 159~
,J - "'" .'~. r 185~
FO551
13~ FO64IL
FO221
Figure 2. Maximum measured temperatures on X-1B airplane,
Mach 1.94, 1957, 1.
High-temperature aerospace structures 5
After the first supersonic flight, research and development of high speed
aircraft intensified. A contract for the design, development and construction of
two X-2 swept wing supersonic research aircraft was awarded to the Bell
Aerospace Corporation in 1947. The X-2 was the first aircraft structure
designed for aerodynamic heating 16. Until the X-2, speeds had not been high
enough for the structure to be affected adversely by aerodynamic heating. For
increased strength at elevated temperatures, the fuselage was constructed
from K-Monel, and the aerodynamic skin used stainless steel. The X-2 became
the first research airplane to achieve speeds above Mach 2.5. On September 27,
1956 the X-2 achieved its maximum speed of Mach 3.2; unfortunately the plane
went out of control, and the pilot was killed.
The next major flight program that stimulated thermal structures
research was the X-15. The X-15 had complex snigiro including the prewar
and postwar work of German stsitneics Eugene Sanger and Irene Bredt who in
1944 outlined a hypersonic, rocket-propelled .tfarcria The evolution of rieht
ideas which contributed ot the development fo the X-15 si described by Hallion
.71 Further snoitpircsed fo the X-15 program are given by HaUion ,81 NASA
Langley researcher Becker 91 and tolip Thompson .02
A thick-skinned heat-sink approach was adopted ot suit the short duration
missions of the X-15. A typical research mission lasted 10-12 minutes .12
Surfaces exposed ot aerodynamic heating were made of Inconel X, a nickel
.yolla Internal structures not exposed ot high temperatures were made fo
titanium. Skin temperatures were designed rof a maximum of 1200 ~ (650
~
On October ,4 1957 the Soviet Union orbited Sputnik ,1 the world's tsrif
laicifitra .etilletas This event changed the nation's seitiroirp rof high speed,
high altitude thgilf making the X-15 program vital ot America's national
.egitserp Between 1959 and 1968 the X-15 accomplished 199 missions, and ti
was the only manned vehicle capable of gniylf atmospheric missions ta Mach 5
rof altitudes of 100,000 feet or higher. tI made many contributions ot the
understanding of hypersonic flight including the identification of several
fundamental thermal structural problems.
The X-15 was the tsrif manned tfarcria rof which aerodynamic heating was
the dominant problem of structural design. Becker 19 notes that great
ecnailer was placed on laboratory stset ni which heat was applied yllacirtcele
and loads mechanically ot represent the thgilf environment. Measurements fo
the behavior of the primary structure in flight verified the ground simulations
confirming that complex high temperature structures could be developed
reliably with ground-based tests, Figure 3.
r940~ /-I,IlO~
061,1 ~
- L960o F
730~
890 ~
FO519
Figure 3. Measured temperatures on X-15, 1.
o22,~ 820~
6 E.A. Thornton
Although the primary airframe performed very well in flight, several
unanticipated problems were encountered in the secondary structures.
Among these problems were panel flutter, thermal buckling and shock
interference heating. Becker reports that early in the program the pilot
reported a rumbling noise at high dynamic pressures that turned out to be
panel flutter of large areas of the skin on the side fairings and tails. The
recognition of the panel flutter problem provided motivation for an extensive
NASA aerothermoelasticity research program. During the flight test program
the X-15 was exposed to surface temperatures as high as 1350 ~ (730 ~ All
three X-15s experienced thermal buckling of the external skin. During the
rocket boost when the aircraft was accelerating and heating rapidly, the pilot
could hear the skin buckling 22. Test pilot Joseph Walker was quoted as
observing, " The a~rplane crackled like a hot stove ." In 1967 NASA conducted
a series of X-15 flights with a dummy ramjet engine mounted on a pylon under
the rear of the fuselage. On the third flight with the dumm_y engine on October
3, 1967 the X-15 reached a maximum 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 23. Considerable heating-induced damage occurred on the engine
pylon showing that local temperatures exceeded the Inconel X melting
temperature of 2600 ~ (1400 ~ Since then shock interference heating has
been recognized as a critical problem for high speed vehicles because extreme
pressure and heat transfer rates can occur in highly localized regions where
the interference pattern impinges on the surface. Shock interference heating
is an important consideration for the engine structure of the National
Aerospace Plane. The problem was strong motivation for 1980s studies of shock
interference heating on engine leading edges.
3. TESTS IN THE ATMOSPHERE
The preceding section described early supersonic flight tests that
demonstrated aerodynamic heating phenomena and established the need for
laboratory testing. Over the last thirty years extensive hot structures test
technology has been developed to address these needs. This section describes a
variety of laboratory tests that study heating effects on aircraft materials and
structures. A representative selection will be presented to illustrate the
diversity of tests conducted.
3.1 Small Components
Testing of small components accounts for the majority of laboratory
research. Such test programs may range from tests to characterize material
behavior to combined mechanical-thermal tests of representative segments of
larger built-up structures.
3.1.1 Isothermal Materials Testing
Among the most widely conducted tests are tensile/compression tests using
universal testing machines. These tests are used frequently to characterize
material behavior over a range of loads with controlled temperature and strain
High-temperature aerospace structures 7
rates. In a typical test (Figure 4) a small specimen is mounted in grips
between crossheads in a load frame. The console of the testing machine
controls the crosshead motion and displays the applied load as well as other
test data. The test specimen with grips is enclosed in a "clam-shell" split
furnace of either a cylindrical or box shape. The temperature within the
furnace is controlled at a specified value so that the test specimen and grips
experience an isothermal environment. Test machines with a wide range of
mechanical loads and programmable control systems are available
commercially. Furnaces, temperature control systems and other accessories
are available for a broad range of temperatures. Similar test machines and
furnaces are used for isothermal creep and fatigue tests.
, Crosshead
Furna Grip
I remieeps
I__
Load Frame
1
Figure 4. Isothermal tensile test with furnace.
3.1.2 Plate with Temperature Gradient
In an early NACA study Heldenfels and Roberts 24 investigated the plane
stress problem for a rectangular plate with a temperature gradient. Simple
"tentlike" temperature distributions were introduced by heating an aluminum
plate along a centerline with a heating wire and maintaining constant
temperature along parallel edges by water flowing through coolant tubes,
Figure 5. Top and bottom surfaces of the plate were insulated to produce
uniform, one-dimensional, linear temperature variations between the heated
centerline and cooled parallel edges. In-plane displacements were permitted
to occur freely, but out-of-plane displacements were prevented by restraints
that forced the plate to remain flat. Thermocouples and strain gages were used
to measure temperatures and strains. The results showed important
characteristics of the stress distribution. The tentlike temperature distribution
8 E.A. Thornton
- - j
2b
W
=q
2a
_S
Insulated
Insulated
J
Coolant
Coolant
Figure 5. Centrally heated plane stress plate, 24.
causes the central portion of the plate to be in compression, Figure 6. For an
unrestrained plate, these compressive stresses must be equilibrated by tensile
stresses along the outer regions of the plate. The most important point
demonstrated by the experiment is that the plate may experience thermal
buckling due to the compressive stresses induced by the spatial temperature
gradients. /k
T 1 ~ ~ _
/
/
/
I"o
/
/
J a -Y
Figure 6. Tentlike temperatures for
t~ x plane stress plate, 24.
High-temperature aerospace structures 9
Shortly after the Heldenfels and Roberts paper appeared, a closely related
paper by Gossard, Seide and Roberts 25 described the buckling and
postbuckling of the same plate. The plate was tested with simple supports for
bending displacements, and in-plane displacements were unrestrained. The
plate was tested under steady conditions with the tentlike temperature
distribution used earlier. The maximum temperature rise during the test was
about 150 ~ (66 ~ The experimental results showed that the effect of the
initial plate deflection was appreciable. The tests also showed that the plate
deflection varied nonlinearly with the temperature rise even for deflections less
than one-half the plate thickness.
3.1.3 Buckling of Ring- Stiffened Cylinders
In the late 1950s and 1960s researchers began to investigate thermal
buckling of shells. When used as structural components in a missile or
launch vehicle, shells may undergo aerodynamic heating which is
nonuniform around the shell circumference and may vary along the length.
Several experimental studies of thermal buckling of cylindrical and conical
shell were conducted; see the author's survey paper 26 on thermal buckling
which summarizes these tests.
Most of these thermal buckling studies concerned monocoque shells, but
Anderson and Card 27 studied ring-stiffened cylinders. Ring-stiffened
cylinders were loaded by a pure bending moment and then heated
non-uniformly until buckling occurred. Figure 7 shows a typical test
configuration and test specimen. The test specimens were stainless steel with
19 inch diameters having a wall thickness of 0.030 inch with a R/t value of 300.
All specimens had an overall length of 45.75 inches, but there were two
different ring spacings L. Ten specimens had nine rings, and three specimens
had five rings with a resultant I JR of 2/1 and ,1 respectively.
The cylinders were heated rapidly by a 25 inch.long quartz lamp heater. In
most of the tests the heater covered approximately one-third of the
circtmfference and was symmetrically located about the bottom of the cylinders
as shown in Figure .7 The procedure for each rapid heating test was to apply a
bending moment less than the room- temperature bending strength, and then
heat the cylinder at a rate of approximately 20 ~ (10 ~ until buckling
occurred. During each test, temperatures at several ring and skin locations
were measured with thermocouples, and in some tests strains were measured
with strain gages. Strain gage data was generally valid for temperatures less
than 175 ~ (80 ~ The behavior of the specimens at buckling was typical of
cylinders loaded in pure bending. Diamond-shaped buckles between rings
which extended to the vicinity of the neutral axis snapped-in suddenly at the
buckling temperature.
A typical experimental temperature distribution along the bottom heated
portion of a cylinder is shown in Figure 8. Note the heat sink effect of the rings
for temperature variations in the longitudinal direction as indicated by the
temperature dips at each ring. Such longitudinal variations in temperature
induce longitudinal variations in the circumferential membrane stresses. In
the circumferential direction, the temperature variations shown occurred
midway between rings. Such circumferential temperature variations induce
significant circumferential variations in the axial membrane stresses.