1、 -. . . . _ I NASA TN D-3580 6, / - NASA TECHNICAL NOTE - -4 -0 - mm mz p= fl w- * 0- fl LOAN COPY: RE %= 2 AFWL (WLdE nb - ”! KIRTLAND AFF = L 0 bn 00 0- f T n 2 + 4 v) 4 z - WIND-TUNNEL-FLIGHT CORRELATlON OF SHOCK-INDUCED SEPARATED FLOW by Donald L, Loving Langley Research Center Langley Station,
2、Hampton? Vu. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION 0 WASHINGTON, D. C. 0 SEPTEMBER 1966 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NY I llllll11111 lllll11111llIl111111 Ill11 1111 Ill1 0330397 NASA TN D-3580 WIND-TUNNE
3、L-FLIGHT CORRELATION OF SHOCK-INDUCED SEPARATED FLOW By Donald L. Loving Langley Research Center Langley Station, Hampton, Va. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientifit and Technical Information Springfield, Virginia 22151 - Price $1.00 Provid
4、ed by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-WIND-TUNNELFLIGHT CORRELATION OF SHOCK-INDUCED SEPARATED FLOW By Donald L. bving Langley Research Center SUMMARY A preliminary study is made of the discrepancies between wind-tunnel predictions and actual fli
5、ght results for conditions of supercritical sepa- rated flow. The limited results obtained for two combinations of Mach num- ber and lift, both involving supercritical-flow separation, suggest that the problem is related to Reynolds number and that an improvement in the correla- tion might be obtain
6、ed by fixing the transition on a model so as to produce a relative boundary-layer thickness at the trailing edge comparable to that calculated to exist in flight. The need for continued research is indicated. INTRODUCTION The purpose of this discussion is to caution experimenters concerning the use
7、of wind-tunnel results in predicting flight loads and moments when supercritical separated flow is present. Whenever separated flow has been observed on wind-tunnel models, the extrapolation of these results to flight conditions has always been subject to question. The discrepancies between aerodyna
8、mic results from flight and wind-tunnel investigations.disclosed herein should not come as a surprise. They are merely additional evidence of the problem associated with separated flows. Two combinations of Mach number and lift, both involving supercritical flow separation, are examined. One is for
9、Mach numbers above cruise at lifting conditions near cruise, and the other is for Mach numbers near cruise at lifting conditions higher than cruise. An example of the difficulty that might be encountered was observed during recent flight tests of a cargo-transport airplane. Mach numbers the wing pre
10、ssures and pitching moments of the airplane were considerably different from those predicted in wind-tunnel tests. No general procedure has been developed for resolving such discrepancies. tions are being conducted, however, to provide a better understanding of the factors involved, and the results
11、herein are presented to report on the progress of these efforts. At supercritical Investiga- lpresented at the classified “Conference on Aircraft Aerodynamics, I Langley Research Center, May 23-25, 1966, and published in NASA SP-124. Provided by IHSNot for ResaleNo reproduction or networking permitt
12、ed without license from IHS-,-,-SYMBOLS drag coefficient, Drag/QS lift coefficient, Lift/%S pitching-moment coefficient, Pitching moment/LSc local pressure coefficient, span of wing, meters chord of wing, meters mean aerodynamic chord of wing, meters free-stream Mach number local static pressure, ne
13、wtons/meter2 free-stream static pressure, newtons/meter2 free-stream mmic pressure, newtons/meter2 total area of wing, meters2 longitudinal distance, measured from wing leading edge, meters angle of attack of fuselage, degrees (p, - pm)/s, DISCUSSION An indication of the differences between wind-tun
14、nel and flight data is shown by the pressure distributions in figures 1 and 2. a coqarison of the chordwise pressure distributions on the upper surface of a cargo-transport wing at a Mach number of 0.73, for a fuselage angle of attack of -0.6, where the lift coefficients for the complete configurati
15、ons are slightly less than 0.3 and the wing pressures are all subcritical. was fixed near the 1,eading edge of the wind-tunnel model by the method dis- cussed in reference 1. The data are for the approximate 40-percent-s and the shock moved farther rearward to the downstream position shown in this f
16、igure. When the flight data points fromfigure 2 are compared with these natural transition model results, the shock positions appear to be, for all practical purposes, the same. For this particular natural transition location, calculations were made and indicated that the relative thickness of the b
17、oundary layer at the trailing edge of the model was the same as that of the full-scale airplane in flight. The test conditions and Visual When the strip was These recent results appear to give evidence that the relative bounda-ry- layer thickness at the trailing edge may be a primary parameter in de
18、termining the shock location and resultant pressure distribution. Additional experimen- tation is necessary, of course, to validate this tentative conclusion. The results thus far obtained, however, do indicate that the discrepancies between wind-tunnel and flight data are a relative boundary-layer
19、thickness effect; that is, a scale effect. The changes in aerodynamic forces that occurred as the transition strip was moved are presented in figure 5 for a near-cruise angle of attack of 2O and a Mach number of 0.85. of lift, drag, and pitching-moment coefficients as a function of the transition- s
20、trip location. The short-dash lines indicate the level of the forces and moment with the transition strip removed. The difference between the lift and drag for the usual forward position of a transition strip and the values obtainedwith natural transition is indicative of an increase in lift-drag ra
21、tio of about 20 percent. variation of pitching moment is representative of a rearward shift in the cen- ter of pressure of 11 percent. Plotted in this figure as solid lines are the variations Of even more importance for the same test conditions, the The results of this wind-tunnel investigation on a
22、 high-aspect-ratio sub- sonic wing at above-cruise Mach numbers, near cruise lift, provide evidence that the discrepancy between wind-tunnel and flight pressure and force data apparently results from a relative boundary-layer-thickness effect on supercritical-flow separation. It would be expected th
23、at the same phenomena also would exist near 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-the cruise Mach number, but at higher-than-cruise lift, since shock-induced sep- aration also occurs for these conditions. pitching-moment coefficients as a
24、 function of lift for the same model just dis- cussed with two extreme boundary-layer test conditions at a Mach number of 0.75. For the configuration with transition fixed near the leading edge (x/c = 0.075) a reduction in stability occurs at lift coefficients slightly above cruise. When the strip i
25、s removed, not only are the pitching-moment coefficients more nega- tive, but the trend toward instability is delayed to a higher lift coefficient. An examination of the wind-tunnel pressure data (which are not presented) indi- cated that this difference is associated with the same separation phenom
26、ena just described for the subsonic wing operating beyond its cruise Mach number; with the transition strip removed, shock-induced separation occurred farther rearward along the chord. As was indicated in the previous discussion, it is probable that the natural-transition configuration more nearly s
27、imulates flight conditions than the fixed-transition configuration. Available flight data do not go up to the point of divergence, so they have not been included in the figure. In figure 6 are plotted the wind-tunnel CONCLUDING REMARKS Because, at supercritical speeds, pressure distributions obtaine
28、d from model and full-scale flight tests may be different, a study has been made for the purpose of improving this correlation. On the basis of this study, a reasonable assumption appears to be that the problem is one of a Reynolds number effect on shock-induced boundary-layer separation. This effec
29、t appears associated with differences between the rela- tive thickness of the boundary layer on models and full-scale airplanes. At the present time no conclusive means are established for exactly simu- lating the supercritical-flow phenomena on models as they exist in flight. On the basis of presen
30、t knowledge, however, it does appear that full-scale charac- teristics may be obtained, at least, on subsonic wings by locating transition on a model so as to produce the same relative boundary-layer thickness at the trailing edge as has been calculated to exist in flight. Until this or other method
31、s can be more definitely established, it is sug- gested as an interim recommendation that wind-tunnel studies be made with transition occurring at various locations. In this manner, at least, the sensitivity of shock-induced separation to modification of the boundary-layer conditions can be determin
32、ed. Langley Research Center, National Aeronautics and Space Administration, Ijangley Station, Hampton, Va., May 23, 1966, 126-13-03-22-23. 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-REFERENCES 1. Braslow, Albert L.; Hicks, Raymond M.; and Harr
33、is, Roy V., Jr.: Use of Grit-Type Boundary-Layer-Transition Trips on Wind-Tunnel Models. NASA D-3579, 1966. 2. Loving, Donald L.; and Katzoff, S.: The Fluorescent-Oil Film Method and Other Techniques for Boundary-Layer Flow Visualization. NASA MEMO 3-17-59L, 1959. 6 Provided by IHSNot for ResaleNo r
34、eproduction or networking permitted without license from IHS-,-,-SUBCRITICAL PRESSURE DISTRIBUTION M20.75; f z-0.6 WIND TUNNEL (TRANSITION FIXED) - FULL-SCALE FLIGHT x /c Figure 1 SUPERCR IT1 GAL PRESSURE DlSTR I BUTlON M.0.85; Qf moo WIND TUNNEL (TRANSITION FIXED) FULL-SCALE FLIGHT - x IC Figure 2
35、7 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EFFECT OF BOUNDARY LAYER ON SHOCK-INDUCED SEPARATION WIND TUNNEL LOW REYNOLDS NUMBER rSHocK FIXED TRANSITION FLIGHT HIGH REYNOLDS NUMBER NATURAL TRANSITION Figure 3 EFFECT OF TRANSITION LOCATION ON PR
36、ESSURE DISTRIBUTION M=0.85; af =Oo - 0.3895 TRANS ITION -STRIP LOCATION, x/c 0.075 .50 NATURAL -_- -_ _ O . FLIGHT 8 x /c Figure 4 I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . _. . . . . . . . . . . EFFECT OF TRANSITION LOCATION ON FORCES
37、M.0.85; at=2 0 I I I I I I I 1 .3 L -.080 II I I I I I 1 1- .I .2 .3 .4 .5 .6 .7 .8 .9 1.0 TRANSITION- STRIP LOCATION, x/c Figure 5 EFFECT OF TRANSITION ON PITCH WIND TUNNEL; Mz0.75 NASA-Langley, 1966 L5227 I m 0 , -NATURAL I I I I I .5 .6 .7 .8 .9 CL Figure 6 9 Provided by IHSNot for ResaleNo repro
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43、sidered Information less broad in scope but nevertheless Details on the availability of these publications may be obtained from: SCIENTIFIC AND TECHNICAL INFORMATION DIVISION NATIONAL AERONAUTICS AND SPACE ADMINISTRATION Washington, D.C. PO546 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
