Flight measurements of airplane structural temperatures at supersonic speeds

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NACA RM H57D18b

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM


FLIGHT MEASUREMENTS OF AIRPLANE STRUCTURAL

TEMPERATURES AT SUPERSONIC SPEEDS

By Richard D. Banner


SUMMARY


Skin and structural temperatures have been obtained on the X-1B
and X-1E research airplanes under transient aerodynamic heating condi-
tions at speeds up to Mach numbers near 2.0. Extensive temperature
measurements were obtained throughout the X-1B airplane, and temperature
distributions are shown on the nose cone, the wing, and the vertical
tail. Temperatures for the X-1E wing leading edge and internal wing
structure were compared with similar data for the X-lB.

No critical skin and structural temperatures were obtained on the
two airplanes over the range of these tests.

Simplified calculations of the skin temperatures in the laminar-
flow regions of the nose cone and the leading edges agreed favorably
with the general trends in the measured data. The flat-plate skin-
temperature calculations in the turbulent-flow regions agreed favorably
with the measured data on the nose cone and at the midsemispan station
of the wing but overestimated the vertical-tail skin temperatures and
also the upper wing skin temperature near the wing tip. The relatively
low values of the upper skin temperatures that were measured at the wing
tip were believed to be caused by separated-flow effects in this region.


INTRODUCTION


In the design of supersonic aircraft, aerodynamic heating is
becoming increasingly important. Analytical studies and controlled
testing represent the basic methods utilized in the design of complex
structures to withstand the effects of aerodynamic heating.

Concurrent with the basic research studies, the National Advisory
Committee for Aeronautics is conducting a program at the IIACA High-Speed
Flight Station at Edwards, Calif., to investigate the skin and structural
temperatures actually experienced by airplanes during flight at super-
sonic speeds. The purpose of this paper is to summarize the results of







NACA RM H57Dl8b


this program to the present time. The results of simplified calculations
of the skin temperatures are compared with the measured data in the
regions of the fuselage nose cone and the wing and vertical-tail skins
and leading edges.


SYMBOLS


b/2 wing semispan

c chord length

hay average heat-transfer coefficient, Btu/sq ft-hr-F

hp pressure altitude, ft

M Mach number

P local surface pressure, lb/sq ft

Po free-stream static pressure, lb/sq ft

qC free-stream dynamic pressure, lb/sq ft

r recovery factor

T skin temperature, OF

Taw adiabatic wall temperature, OF

TT free-stream stagnation temperature, To(1 + 0.2M2), OF

T. free-stream ambient air temperature, OF

t time, sec

a angle of attack, deg

7 thickness, in.


TESTS


Data have recently been obtained on the X-IE airplane at Mach num-
bers up to 2.10 and on the X-lB at Mach numbers up to 1.94. (See fig. 1.)







NACA RM H57D18b


These speeds do not necessarily represent the maximum speed capabilities
of the airplanes. Both airplanes are constructed primarily of aluminum.
Temperature measurements were made at approximately 60 locations on the
X-1E, and approximately 500 temperatures were measured on the X-lB.

Figure 2 shows the flight conditions for both airplanes under which
the temperatures were obtained. The Mach number, pressure altitude,
ambient air temperature, and angle of attack are shown as time histories.
As can be seen, the two flights are generally similar.

Transient heating conditions were experienced by both airplanes
during the flights. The maximum heating rate was experienced on the
thinner skinned X-1B and was on the order of 30 per second. This heating
rate is relatively low in comparison with the rates being obtained on
missiles and rocket models and in controlled wind-tunnel tests; however,
it is believed to be representative of the heating rates which are being
experienced or which will be experienced in the near future by fighter
and interceptor aircraft.


CALCUIATIOIS


For laminar flow in the stagnation regions, approximate heat-transfer
coefficients were calculated from expressions relating the Nusselt number
and the Reynolds number given by Stine and Wanlass. (See ref. 1.) In
the regions of flat-plate turbulent flow, approximate heat-transfer coef-
ficients were calculated by using Colburn's expression. These calcula-
tions were greatly simplified by the use of free-stream conditions.
This procedure was considered justified in that only the overall effects
were desired. A detailed theoretical analysis is not considered to be
within the scope of this report.

The results of the calculations just described indicated relatively
small variations in the heat-transfer coefficients with time; therefore,
average values were determined and used in calculating the skin
temperatures.

Newton's law of heat flow to the skin, which considers the heat capac-
ity of the material and neglects the effects of conduction, was assumed.


RESULTS AND DISCUSSION


The X-1B installation provides a rather detailed case history of the
temperatures that exist throughout an actual airplane; and, in view of
this fact, more attention is given to the X-lB data in the present paper.







NACA RM H57Dl8b


The X-IE data are used to point out any differences that exist in the
measured temperatures due to either the configuration or the construction
or both.

Some of the maximum temperatures that were measured on the X-1B and
an indication of the higher temperature areas on the airplane are shown
in figure 5. The nose of the airplane and the leading edges of the wing
and tail surfaces are shaded in figure 3 to indicate the higher tempera-
ture areas. The approximate thermocouple locations are indicated by the
dark points. The maximum stagnation temperature for the flight was 2200 F.

A maximum skin temperature of 1850 F was measured at the forward point
of the nose cone. On the wing leading edge the maximum temperature was
1680 F. A maximum temperature on this order was also measured on the
leading edge of the vertical tail. The maximum temperatures in other
general areas are also shown. Some of the temperatures shown at these
locations are appreciably affected by internal heat sources and heat
sinks; for instance, the fuselage skin temperatures adjacent to the
liquid oxygen tank are relatively low (150 F); whereas, just ahead of
that area on the fuselage skin a maximum of 1220 F was measured. Other
areas of interest are the windshield and canopy and the rearward part
of the fuselage in the region of the rocket engine.

It should be pointed out that the maximum temperatures shown in fig-
ure 5 did not all occur at the same time. In areas where internal con-
duction is negligible, skin thickness, boundary-layer temperature, and
the heat-transfer coefficient are the primary factors affecting the skin
temperature rise.

Figure 5 gives an overall picture of the measured temperatures.
In subsequent figures, the temperature distributions in the areas of the
nose cone, the wing skins, the leading edges and the vertical tail are
considered in greater detail; and comparisons are made with calculated
results. Attention is given only to the supersonic portion of the flights,
since at the subsonic speeds the combined effects of increasing Mach num-
ber and decreasing ambient air temperature produced only small changes in
the skin temperatures.

Figure 4 shows the nose of the X-1B, where both skin-temperature and
pressure measurements were made at intervals along the nose between sta-
tion 0 and station 55.0. The pressure data are plotted as pressure coef-
ficients at the bottom of the figure. The measured skin temperatures are
shown in the upper part of the figure as the open symbols. The skin tem-
perature decreases over the forward part of the nose cone, the laminar
flow calculations showing fairly good agreement. This variation is fol-
lowed by a region of no change in the skin temperature and then an increase
toward the rear of the nose cone where the calculated turbulent flat-plate
and cone temperatures are approached. In this region the nose-cone







NACA RM H57Dl8b


shape asymptotically approaches the cylindrical shape of the fuselage,
and the pressure drops to near the free-stream value. The overall tem-
perature variations that are shown suggest that transition from laminar
to turbulent flow takes place along the nose at about 25 to 50 percent
of the nose-cone length, fully turbulent flow developing at the rear of
the nose cone. A summary of various wind-tunnel data on cones and bodies
of revolution indicated this general transitional area.

The skin on the nose cone was relatively thin in comparison with
that on other areas on the airplane. The effects of varying skin thick-
ness on the skin temperatures can be seen in figure 5. Figure 5 shows
the spanwise distributions of maximum temperatures on the upper wing
skin at the 66-percent-chord line and at the leading edge. The span-
wise variation in the skin thickness is shown below. The temperature
variations indicate an inverse relationship between the skin temperature
and the skin thickness at both chordwise positions. The trend which
would give rise to thermoelastic considerations at higher temperature
levels is clearly seen in the data here.

The calculated temperatures, shown by the dashed lines in figure 5,
were based on a constant spanwise heating input, laminar for the leading
edge and turbulent for the 66-percent-chord line; and the same variations
due to thickness are seen in the calculated temperatures.

The data in figure 5 also illustrate the differences in the skin
temperature with chordwise position. Both top and bottom skin tempera-
tures were measured at several chordwise positions at about midsemispan
and near the tip of the wing. These data are presented in figure 6.
In this figure the calculated temperatures were estimated on the basis
of zero angle of attack. No detailed consideration is given to the
effects of angle of attack on the measured temperatures; however, from
an overall standpoint it should be recalled from figure 2 that the angles
of attack were positive during the flight and were on the order of' 2
to 100.

The data of figure 6 illustrate several interesting trends. First,
notice the chordwise temperature gradients that are shown at the mid-
semispan station. The higher temperatures are experienced at the leading
edge and the trailing edge (which at this location is the outboard tip of
the flap), and the lower temperatures are experienced on the thicker
skinned wing box section.

Secondly, note the differences in temperature between the top and
bottom skins at the two span stations, the bottom skin temperature being
higher in both cases. At the tip station, the fairly large differences
seen between the top and bottom skin suggest that the flow might be partly
separated in this area.







NACA RM H57D18b


The approximate calculations, which were based on the assumption
of laminar flow over the leading-edge section and turbulent flow over
the remainder of the chord, give a fairly good overall estimate of the
skin temperatures at the midsemispan station. At the tip station the
calculations agree fairly well with the bottom skin temperature; however,
they indicate an overestimate of the top skin temperature, probably
because of the flow effects previously mentioned.

An example of the internal temperatures that were measured through
the wing is seen in figure 7. Shown in the figure are the front wing
spars at about midsemispan on the X-lB and the X-1E. The temperatures
were measured at the locations shown by the black dots on the structures,
and the values are given on the right-hand side of the figure. The higher
temperatures shown for the X-1B were measured on the skin a slight dis-
tance from the spar.

In order to give an indication of the temperature rise that has taken
place in the internal structure, it is worthwhile to mention that the
ambient air temperatures were between -70 F and -900 F and that the
assumed turbulent adiabatic wall temperatures were on the order of 2000 F
for the highest Mach numbers shown here. The measured internal tempera-
tures are relatively low and show only slight temperature gradients across
the thickness of the X-1B wing, the lower temperature being obtained on
the spar center line. Essentially no differences are seen on the thick
spar construction of the X-1E wing, the heavier type of construction of
the X-1E having a temperature-neutralizing tendency due to the higher
heat capacity. The thermal lag effect is also seen in figure 6 by the
increase in the measured temperatures as the Mach number decreases in
the later portions of the flights.

Effects of material thickness differences are also seen in the meas-
ured leading-edge temperatures. These data are shown in figure 8 in
time-history form together with time histories of the assumed laminar
adiabatic wall temperatures. The locations at which the temperatures
were measured are shown by the dark points on the leading-edge sketches,
and the material thicknesses are given at these locations. Values of
the average heat-transfer coefficients utilized for the calculated tem-
peratures are shown below the sketches. The calculated temperatures are
seen to agree very well with the measured data.

Maximum temperatures on the same order of magnitude were measured
on the wing and vertical-tail leading edges of the X-1B. For comparison,
the temperatures that were measured at the rear of the solid leading
edge of the X-1E are seen in the middle of the figure. It will be
noticed that the maximum measured temperature was on the order of 200 F.
The high heat capacity of the solid leading edge is the contributing
factor to the small rise in the temperature measured at this location.
The calculated skin temperature is shown to agree very well with the







NACA RM H57D18b


measured data; however, this result is not considered significant because
the measured temperatures are relatively low. (For example, a 50-percent
reduction in the assumed heat-transfer coefficient at this point would
produce a decrease in the calculated maximum temperature of 70 F.)

The chordwise variation in the vertical-tail temperatures is shown
in figure 9 for a time near maximum Mach number at near midspan. The
temperatures were measured on the skin and spar center lines at the loca-
tions shown by the black dots on the sketch. No appreciable gradients
are seen in the chordwise variation of the measured skin temperature.

Transition from laminar to turbulent flow was assumed to take place
at the point where the leading-edge section attaches to the front spar
because inspection revealed a relatively large discontinuity in the skin
at this point. Skin temperatures calculated from the average heat-
transfer coefficients shown and based on these assumptions agree fairly
well with the measured trends in the leading-edge region but deviate
somewhat over the remainder of the chord and give a conservative overall
estimate in this region.


CONCLUDING REMARKS


Skin and structural temperatures have been obtained on the X-1B
and X-1E research airplanes under transient aerodynamic heating condi-
tions at speeds up to Mach numbers near 2.0. Extensive temperature
measurements were obtained throughout the X-1B airplane, and temperature
distributions are shown on the nose cone, the wing, and the vertical
tail. Temperatures for the X-1E wing leading edge and internal wing
structure were compared with similar data for the X-lB.

No critical skin and structural temperatures were obtained on the
two airplanes over the range of these tests.

Simplified calculations of the skin temperatures in the laminar-
flow regions of the nose cone and the leading edges agreed favorably
with the general trends in the measured data. The flat-plate skin-
temperature calculations in the turbulent-flow regions agreed favorably
with the measured data on the nose cone and at the midsemispan station
of the wing but overestimated the vertical-tail skin temperatures and
also the upper wing skin temperature near the wing tip. The relatively
low values of the upper skin temperatures that were measured at the wing
tip were believed to be caused by separated-flow effects in this region.


High-Speed Flight Station,
National Advisory Committee for Aeronautics,
Edwards, Calif., March 6, 1957.







8 NACA RM H57D18b


REFERENCE


1. Stine, Howard A., and Wanlass, Kent: Theoretical and Experimental
Investigation of Aerodynamic-Heating and Isothermal Heat-Transfer
Parameters on a Hemispherical Nose With Laminar Boundary Layer at
Supersonic Mach Numbers. NACA TN 5544~, 1954.








NACA RM H57D18b


RESEARCH

X-IE
M-2.10
60 TEMP. GAGES


AIRPLANES

X-IB
Ms 1.94
300 THERMOCOUPLES


cZ~


Figure 1


FLIGHT CONDITIONS


X-IE X-IB


M
0.5
65103

hp, FT 25

25 -F_'------- 0s :------

-100


a, DEG 12
0 100 200 300 4000 100 200 300 400
TIME, SEC TIME, SEC


Figure 2


c3








NACA RM H57D18b


MAXIMUM MEASURED TEMPERATURES,


Mmax = I 94


X-IB


TT = 2200 F


* THERMOCOUPLE LOCATIONS


1500F


1540F


Figure 5


NOSE CONE TEMPERATURES AND PRESSURES, X-IB
M 1.94, t =270 SEC
o MEAS.
--- LAMINAR
TURBULENT CALC.


CONE
o -- FLAT PLATE
0 0 0


Figure 4


II I I I I I
0 10 20 30 40 50 60
DISTANCE, INCHES







NACA RM H57Dl8b



MAXIMUM SPANWISE SKIN TEMPERATURES, X-IB WING


LE., M= 1.90, t=290 SEC


66%c, M= I 32, f = 325 SEC

o MEAS
----CALC fLE (r=0.85)
66% c (r=090)


F, -66/oc
, IN.
0


Figure 5


CHORDWISE SKIN TEMPERATURES, X-IB WING
Mc-=1.90, TT=212F


---- BOTTOM SKIN I
--o-- TOP SKIN MEANS
* CALC (0=0)


.3 TOP
r, in. 1-OTTOM


-TOP AND BOTTOM


100 0


Figure 6


200
54% b/2

T, OF s o
1000

50-







NACA RM H57D18b



MEASURED INTERNAL WING TEMPERATURES

X-IB


-f9 D,.395-

.125

.3655


0 40 80 120
T, F


X-IE


0 40 80 120
T' F


Figure 7


LEADING-EDGE TEMPERATURES
o MEAS ---- CALC Taw(r=085)


X-IB WING
54% b/2
T =.081 in,


X-IE WING
63.3% b/2
=.62 in


\,


X-IB VERTICALTAIL
MID-SPAN
r =.051 in


BTU BTU BTU
hov = 25 S FT-HR-F h FT-HR-F FT-R-F ha =25S FT-HR-F
200 r r A r


250
t, SEC


350 150 230 310 150
t, SEC


250
t, SEC


Figure 8


100
T'F
0
I!







NACA RM H57Dl8b


CHORDWISE TEMPERATURES-X-IB VERTICAL TAIL
M =190, t = 290 SEC
a SKIN M ---CALCULATED
SPAR MEASURE (SKIN)


TRANSITION
200 TRANSIT ADIABATIC WALL

150 N %-
T,OF 1 <' o
100 -
I3

50

0
50 LAMINAR
hv' 25- r-TURBULENT
BTU/SQ FT -HR-"F .

0 20 40 60 80 100
% CHORD


Figure 9


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