^tcmf<\\\SM'iL 
 
 RM H57D18b 
 
 RESEARCH MEMORANDUM 
 
 FLIGHT MEASUREMENTS OF AIRPLANE STRUCTURAL 
 
 TEMPERATURES AT SUPERSONIC SPEEDS 
 
 By Richard D. Banner 
 
 High-Speed Flight Station 
 Edwards, Calif. 
 
 UNIVERSrrv OF FLORIDA 
 DOCUMENTS DEPARTMENT 
 
 GAINESV/lXE,FL326J,.70noSA 
 
 NATIONAL ADVISORY COMMITTEE 
 FOR AERONAUTICS 
 
 WASHINGTON 
 
 June 7, 1957 
 Declassified July 22, 1959 
 
1^0 V33 ^^'f 57J:»v.^; 
 
 NACA RM H57Dl8b 
 
 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
 
 RESEARCH MEMORANDUM 
 
 FLIGHT MEASUREMENTS OF AIRPIANE STRUCTURAL 
 TEMPERATURES AT SUPERSONIC SPEEDS 
 By Richard D. Banner 
 
 SUMMARY 
 
 Skin and structural temperatures have been obtained on the X-IB 
 and X-IE 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-IB airplane, and temperature 
 distributions are shown on the nose cone, the wing, and the vertical 
 tail. Temperatures for the X-lE wing leading edge and internal wing 
 structure were compared with similar data for the X-IB. 
 
 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 temperatijres 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. 
 
 IfJTRODUCTION 
 
 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 
 structiores 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 NACA 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 resiilts of simplified calculations 
 of the skin temperatijres 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 semi span 
 
 c chord length 
 
 hg^^ average heat-transfer coefficient, Btu/sq ft-hr-^ 
 
 h^ pressure altitude, ft 
 
 M Mach number 
 
 P local s\irface ^ress^xce , Ib/sq ft 
 
 Poo free-stream static pressure, Ib/sq ft 
 
 q^^ free-stream dynamic pressure, Ib/sq ft 
 
 r recovery factor 
 
 T skin temperature, 'r 
 
 T adiabatic wall temperature, °F 
 
 Tm free-stream stagnation temperature, T<x,(l + 0.2M2), '^ 
 
 Too free-stream ambient air temperat\ire, '^ 
 
 t time, sec 
 
 a angle of attack, deg 
 
 T 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-IB at Mach numbers up to 1.9^. (See fig. 1.) 
 
NACA RM H5TDl8b 
 
 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-IE, and approximately 5OO temperatures were measured on the X-IB. 
 
 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-lB and was on the order of 5° pei" 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 futiore by fighter 
 and interceptor aircraft . 
 
 CALCULATIONS 
 
 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 resiolts 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 
 temperat\ires. 
 
 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-IB 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-IB data in the present paper. 
 
NACA RM H57Dl8b 
 
 The X-LE 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-IB 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 220° F. 
 
 A maximum skin temperature of 185° F was measiored at the forward point 
 of the nose cone. On the wing leading edge the maximum temperature was 
 l68° F. A maximum temperature on this order was also measiored 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 (15° F)} whereas, just ahead of 
 that area on the fuselage skin a maximum of 122° 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 maximiim temperatures shown in fig- 
 ure 3 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 3 gives an overall pictixre of the measured temperatiires. 
 In subsequent figures, the temperatiore 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 h shows the nose of the X-IB, where both skin-temperature and 
 pressure measurements were made at Intervals along the nose between sta- 
 tion 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 measiired 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 10°. 
 
 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 siiggest that the flow might be partly 
 separated in this area. 
 
NACA RM H57Dl8b 
 
 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 temperatiires 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 temperatiires that were measiired thro\agh 
 the wing is seen in figure 7» Shown in the figure are the front wing 
 spars at about midsemispan on the X-IB and the X-IE. 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-IB 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 -90° F and that the 
 assumed turbulent adiabatic wall temperatures were on the order of 200° 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-IB wing, the lower temperature being obtained on 
 the spar center line. Essentially no differences are seen on the thick 
 spar construction of the X-IE wing, the heavier type of construction of 
 the X-IE 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-IB. For comparison, 
 the temperatures that were measured at the rear of the solid leading 
 edge of the X-IE are seen in the middle of the figure. It will be 
 noticed that the maximum measured temperature was on the order of 20° 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 H57Dl8b 
 
 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 7° F.) 
 
 The chordwise variation in the vertical-tail temperatures is shown 
 in figure 9 for a time near maximum Mach nimber 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 measiured 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-IB 
 and X-IH research airplanes under transient aerodynamic heating condi- 
 tions at speeds up to Mach n\mibers near 2.0. Extensive temperature 
 measurements were obtained throughout the X-IB airplane, and temperature 
 distributions are shown on the nose cone, the wing, and the vertical 
 tail. Temperatures for the X-IE wing leading edge and internal wing 
 structure were compared with similar data for the X-IB. 
 
 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 measixred 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. 
 
MCA RM H57Dl8b 
 
 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 ^ikk, 195^. 
 
NACA RM H57Dl8b 
 
 RESEARCH AIRPLANES 
 
 X-IE 
 
 MfsZ.IO 
 60 TEMP. GAGES 
 
 X-IB 
 
 M~\.9A 
 300 THERMOCOUPLES 
 
 Figure 1 
 
 M 
 
 hp.FT 
 
 Tm "F 
 
 a, DEG 
 
 FLIGHT CONDITIONS 
 X-IE X-IB 
 
 100 200 300 400 100 200 300 400 
 
 TIME, SEC TIME, SEC 
 
 Figure 2 
 
10 
 
 NACA RM H57Dl8b 
 
 MAXIMUM MEASURED TEMPERATURES, X- 1 B 
 
 153° F 
 
 lAS-'F 
 
 I22°F 
 
 Figure 3 
 
 NOSE CONE TEMPERATURES AND PRESSURES, X-IB 
 
 M = 1.94, t =270 SEC 
 
 o MEAS. 
 \ LAMINAR I 
 
 TURBULENT 
 
 CALC. 
 
 STA 55.0 
 
 THERMOCOUPLES 
 PRESSURE ORIFICES 
 
 10 20 30 40 50 60 
 
 DISTANCE, INCHES 
 
 Figure k 
 
NACA RM H57Dl8b 
 
 11 
 
 MAXIMUM SPANWISESKIN TEMPERATURES, X-IB WING 
 L.E., M=l.90, t = 290 SEC 66%c, M = l.32, t=325 SEC 
 
 o MEAS 
 
 CALC |LE(r=0.85) 
 
 ^^^'^ ^66% c (r = 0.90) 
 
 Figure 5 
 
 CHORDWISE SKIN TEMPERATURES, X-IB WING 
 Mco=l.90, T-p = 2l2°F 
 
 -- 0~ BOTTOM SKIN 
 200 
 I50i- 
 
 TOP SKIN 
 
 MEAS 
 
 T/F 
 
 100 
 
 50 
 
 
 
 • CALC (a=o) 
 957ob/2 
 
 irzEzzcEr*- <=x=izji:r:jz^ 
 
 T. m. 
 
 Jl 
 
 TOP 
 
 'Y 
 
 BOTTOM 
 %c 
 
 100 
 
 
 TOP AND BOTTOM 
 
 %c 
 
 100 
 
 Figure 6 
 
12 
 
 MCA RM H57Dl8b 
 
 MEASURED INTERNAL WING TEMPERATURES 
 
 X-IB 
 
 ^T.- 
 
 .081 
 
 365J 
 
 X-IE 
 
 
 
 1 
 
 1 
 
 
 1 
 
 
 
 1.40 
 
 .60 -» 
 
 
 -^ 
 
 1 
 
 
 
 40 80 120 
 Tf F 
 
 M 
 2.04 
 
 /^l.30 
 
 r 
 J 
 
 40 80 120 
 TtF 
 
 Flgixre 7 
 
 T =.081 in 
 
 LEADING-EDGE TEMPERATURES 
 o MEAS CALC Tg^(r = 085) 
 
 X-IB WING X-IE WING X-IB VERTICALTAIL 
 
 54% b/2 63.3% b/2 MID-SPAN 
 
 r =.051 in 
 
 '■ =,62 in 
 
 BTU 
 
 BTU 
 
 BTU 
 
 150 250 350 150 230 310 150 250 350 
 
 t, SEC t, SEC t, SEC 
 
 Figure 8 
 
3B 
 
 NACA RM H57Dl8b 
 
 15 
 
 CHORDWISE TEMPERATURES -X- IB VERTICAL TAIL 
 M = 1.90. t = 290 SEC 
 
 o SKIN 
 o SPAR 
 
 MEASURED 
 
 — CALCULATED 
 
 (SKIN) 
 
 « — I — • r^-* 1 — •- 
 
 200r 
 
 TRANSITION 
 
 150 
 
 T.'-F 
 
 ADIABATIC WALL 
 
 
 50 
 
 av 
 
 BTU/SQFT-HR- 
 
 aur r-LAI 
 
 LAMINAR 
 
 
 20 
 
 o o 
 
 i 
 
 TURBULENT 
 
 40 60 
 
 % CHORD 
 
 80 
 
 100 
 
 Figure 9 
 
 NACA - La.igley Field, V« 
 
I 
 
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 UNIVERSITY OF FLORIDA 
 
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