CB Ho. L5E22 '))ttk i^l NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED June 19^5 as Confidential Bulletin L5E22 FLIGHT IHVESTIGATIOH AT HIGH SPEEDS OF FLOW CONDITIONS OVER AH AIRPLANE WING AS INDICATED BY SURFACE TUFTS By Clotaire Wood and John A. Zelovcik Langley Memorial Aeronautical Laboratory Langley Field, Va. NACA WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were not tech- nically edited. All have been reproduced without change in order to expedite general distribution. 91 DOCUMENTS DEPARTMEN J J Digitized by the Internet Archive in 2011 with funding from University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation http://www.archive.org/details/flightinvestigOOIang NACA CB No. L5E22 CONFIDENTIAL : ATIONAL ADVISORY COMMITTEE FOR AERONAUTICS CONFIDENTIAL BULLETIN PLIGHT INVESTIGATION 47 HIGH SPEEDS )F PLOW CONDITIONS AN AIRPLANE RING A3 INDIC BY SURFACE TUj ! By Clotaire Rood and John A. Zalovcik SUMMARY plight tests were made at high speeds with a P-I4.7D airplane to determine the flow characteristics, as indicated by wool tufts, en a section of the upper surface of the wing. The behavior of the tufts, which were distributed ever a section cf the wing from y$ .5 to 52. S percent semispan, was determined from motion pictures, The tests were made in straight flight and in turns under conditions in which airplene lift coef- ficients from 0.10 to O.5I+ and airplane Mach numbers from O.58 to O.78 were obtained. The results of t- t 3ts indicated that the flow remained smooth over the test panel until the critical Mach number of the panel was exceeded by 0.08 at a lift coefficient of 0.10 and by 0.05 at a lift coeffi- cient of 0.50- Beyond these Mach numbers, the tufts indicated unsteadiness of flow end, finally, local separation when the Mach number exceeded the critical value by 0.1J at a lift coefficient of 0.10 and by 0.10 at s lift coefficient of 0.50. the region of separated flow originated in the neighborhood of JO percent chord at high lift coefficients and R5 percent chord at low lift coefficients. Separation appeared to extend over not xuov^ than 15 percent chord. INTRODUCE In the course of flight tests rrade to determine the profile-drag characteristics of the ' Ln of a F-47'R sir- plane, a few wool tufts were fastened to the wing surface to permit visual observation of the direction of flow In the boundary layer. The behavior 01 the tufts at high CONFIDENTIAL 2 CONFIDENTIAL NACA GB No. L5E22 speeds indicated disturbances in the flow over the wing, apparently associated with compressibility effect?, and suggested that tuft observations might provide interesting information on flow phenomena at high speeds. A more complete tuft installation was therefore made over a section of the wing surface between 39*5 an< ^ 52-5 percent semispan from the plane of symmetry. The tufts were photo- graphed during flight at high speeds. The tests were made in straight flight and in turns under conditions in which airplane lift coefficients from 0.10 to O.5I4. and airplane Macb numbers from O.^S to C.7O were obtained. The flow conditions indicated by the behavior of the tufts are presented graphically herein for a few typical flight conditions and are correlated with the flight conditions . APPARATUS :\ND TESTS Tufts were located en the upper surface of the right wing of a P-I4.7D airplane (fig. 1) at four span- wise stations: 39. 5> ^-3-5> l; +°> an d 52-5 percent semi- span from the plane of symmetry (fig. 2). The tufts consisted of strands of white wool yarn arranged in chordwise rows with each row attached to the surface by a continuous strip of black "Scotch" cellulose tape. Spanwise chalk lines were drawn on the surface of the wing at intervals of 10 percent chord, and each line was identified by a number beginning with 1 at the 10-percent -chord station and continuing through 7 a ^ the 70~0 ercen t-chord station (fig. 5 ) ■> The wing of the P-.Lf.7D airplane incorporates Republic S-3 airfoil sections, which have pressure distributions similar to those of the NACA 230-series sections . The average chord of the test panel was about 96 inches and the average thickness was about 12.6 percent chord. The behavior o p the tufts during the tests was photographed with a l6 -millimeter motion-picture camera operating at a speed of approximately 32 frames per second. Measurements of normal accelervitlon and free- stream impact pressure were recorded by means of NACA recordi'ig instruments. The altitude of the tests, indicated by an altimeter in the cockpit, was noted by the pilot. CONFIDENTIAL NACA CB No. L5E22 CONFIDENTIAL 5 The testa were made in strai ;ht flight and in turns (l|g to lj.igj st an altitude of 20,000 feet and at indicated airspeeds from 315 to L|_08 miles psr hour. The airplane Mach numbers ranged from O.58 to O.78, and the airplane lift coefficients ranged from 0.10 to . 5 -4- . VR" N [ STILTS An enlarge! 3nt of one frame of the motion-picture film taken "] ;ht n ' s s] own as figure I4.. The quality of the Dhoto rraphs was, in general, too poor to permit satisfactory renrod iction 1 this form; in fig- ures 5 to 8, therefore, sketches based on the original photographs are used to illustrate, for a few typical flight conditions, the flow c icated by the tufts. The flow sonditions for various airplane lift coefficients at constant airplj ach numbers of O.69 and 0.7-1 are shown in figures 5 -^ 6, respectively; the flow conditions )r various a :h r rs at constant airplane lift coefficients of 0.13 and 0.1l,3 are shown in figures 7 an( 3 3, respectively. Inasmuch as the field of the camera covered only the forward 70 to 80 percent chord, the flow conditions downstream of this region are not known. flow r; tions Lnd jated ; the tufts at various lift coefficients and ach Lumbers are summarized in figure ' . The interpretation of tl havior of tr 3 tufts is as follows: Tufts I Lng straight back and motionless indicate smooth flow, tufts oscillatii lateral] icate unsl fl ' , tufts "flopping" around leisurely or lying curved on the sui ' ce indicate flow separation. ( '.' sare figs. I4. and 6(d).) e critical Mach n * h cr c M ig sections at 25 and 63 percent a ipan and the Mach number at which shock was first evident Ln the wake at 63 percent semispan were determined from the results (unpublished) of other tests of the P-lnO i .rplane and are compared in figure 10 r/itt the Mach numbers 'low dis- turbance and flow separation vere first indicated in the ^resent tests. (Wing stations in figs. 10 and 11 are designated 2y/b, where is [stance of the wing station from the plane o' letry and b is the wing span.) The determination of the critical Mach 00. ?T T - k CONFIDENTIAL NAG A CB No. L5E22 numbers involved extrapolation, by the von Karman method, of pressure-distribution data obtained at Mach numbers 0.02 to 0.06 less than the critical value. The sirplane lift coefficients were correspondingly modified ty means of the Prandtl-Glauert relation. The pressure- distribution measurements were obtained with static- pressure tubes and therefore, according to the results of reference 1, the critical Mach numbers may be as much as 0.01 higher than would have been obtained from pressure measurements with flush orifices. The critical Mach number at 46 percent semispan, which is the center line of the teat panel, was obtained by linear interpolation between the critical Mach numbers at 25 and 63 percent semispan. A comparison is made in figure 11 of the flow behavior and the critical Mach number obtained in flight and in the Ames l6-i'cot high-speed tunnel on a O.J-scale model of the P-ii7D airplane (reference 2). The comparison of flow characteristics is made on the assumption that tuft behavior is interpreted in the same way in the wind tunnel and in flight. The critical Mach number shown for the wind-tunnel tests was determined from pressure-distribution measurements made with flush orifices at hi percent semispan. DISCU3SI . i RESULTS The sketches of figures 5 to it show that, as Mach number or lift coefficient increased, the flow first became unsteady over a small chordwise region; this region then became more extensive, and finally local separation occurred. The region of separated flow originated in the neighborhood of yO percent chord at high lift coefficient and I4.5 percent chord at low lift coefficients. The region of separation appeared to extend over not more than 15 percent chord and was followed by a region of unsteady flow beyond which the flow again was steady. Three distinct regimes of flow are evident in figure 9. At a lift coefficient of 0.10, the flow ained smooth up bo a Mach number of 0.73; beyond this Mach number, the flow was unste : h flow separation occurred at a Mach nui ber . .'"?• At a CC\ SNTIAL NACA CB No. L5E22 CONFIDENTIAL lift coefficient of 0.^0, the flow remained smooth up to a Mach number of 0.b2 and local flow separation occurred at a Mach number cf O.67. Comparison of these results with the critical Mach number in figure 10 indicates that the flow remained smooth until the critical Mach number was exceeded by 0.05 to 0.08, depending on lift coefficient. Local separation of flow occurred when the critical Mach number was exceeded by 0.10 to 0.15. Tie ?' T ach number at which compressibility shock was first evident in the wake at 63 percent semi- span was apparently exceeded by 0,05 to 0.08 before local flow seoaration occurred. The comparison in figure 11 o° flight and wind- tunnel results indicates that the critical Mach number was 0.03 to . 0J4. higher, depending on lift coefficient, in the flight tests than in the tests of the Q.J-seale model of the P-I1.7D airplane in the Ames l6-foot high- speed tunnel. The Mach number at which local separation occurred was 0.02 higher .in flight than in the tunnel. The flight and tunnel remits are therefore in good agreement. C0NCL T L0IN3r REMAR Flight tests made at high speeds with a P-K7D air- plane to determine the flow characteristics, as indicated by wool tufts attached to a section of the upper surface of the wing, showed that- the flow remained smooth until the critical Mach number of the wing section was exceeded by 0.08 at a li p t coefficient of 0.10 and by . 05 at a lift coefficient cf 0.50. Beyond these Mach numbers, the tufts indicated unsteadiness of flow and, finally, local separation when the Mach number exceeded the critical value by 0.13 at a lift coefficient of 0.10 and by 0.10 at a lift coefficient of 0.S0. The region of separated flow originated in the neighborhood of 30 -percent chord at high lift coefficients and l±*j percent chord at low lift coefficients. The region of separation appeared to extend over not more than 15 percent chord. Com- parison of these results with results obtained in the Ames l6-foot high-speed tunnel on a 0.3-scale model of CONFIDENTIAL CONFIDENTIAL NACA CB No. LSE22 the P-Lj_7D airplane indicated good agreement between the flight and tunnel results. Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va. IEPER CNCES 1. Zalovcik, John A., and Daum, Fred L«: Plight Investigation at High I.lach Numbers of Several Methods of Measuring Static Pressure on an Air- plane Wing. NACA RB No. li+HlOs, l-ii^. 2. Hamilton, William T. , and Boddy, Lee E. : High-Speed Wind-Tunnel Tests of a 0.3-Scale Model of the P-47D airplane. NACA ACR No. 5D20, I9I4.5 . CONFIDENTIAL NACA CB No. L5E22 Fig. 1 O O CO >> CD > u 3 CO *-> 3 t-i o -a CD CO 3 ^ td T " ex, .0 cu 3 bo •H NACA CB No. L5E22 CONFIDENTIAL Fig. 2 •Spanwise location of tufts CONFIDENTIAL NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Figure 2. - Plan view of Republic P-i^D airplane showing spanwise location of tufts. NACA CB No. L5E22 CONFIDENTIAL Fig. Figure 3.- Test panel on right wing of Republic P-47D airplane, showing rows of tufts at 39.5, 43.5, 48, and 52.5 percent semispan. Numbers identify span- wise lines at intervals of 10 percent chord. CONFIDENTIAL NACA CB No. L5E22 CONFIDENTIAL Fig. 4 Figure 4.- Photograph showing tuft behavior at an airplane lift coefficient of 0.49 and at an airplane Mach number of 0.71. CONFIDENTIAL NACA CB No. L5E22 Fig. 5a-c CONFIDENTIAL Steady flow Unsteady flow Separated flow NATIONAL ADVISORY COMMOTE FOR AERONAUTICS CONFIDENTIAL Figure S~ F~lov/ oonditlons oi test section for an airplane Mach number of 0.69. Fig. 5d-f NACA CB No. L5E22 Ftqure 5~ CONFIDENTIAL Concluded. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA CB No. L5E22 Fig. 6a-c CONFIDENTIAL Steady flow Unsteady flow Separated flow C L -0.43 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS CONFIDENTIAL figure 6~ flow conditions at test section for an airplane Mach number of 0. 77. Fig. 6d NACA CB No. L5E22 CONFIDENTIAL NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS (d) CONFIDENTIAL Figure 6.~ Concluded. Steady flow Unsteady flow Separated flow NACA CB No. L5E22 Fig. 7a-c Steady flow Unsteady flow Separated flow NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS CONFIDENTIAL Figure 7T~ Flow conditions at test sect ion for on airplane lift coefficient of 0./3. Fig. 8a-c NACA CB No. L5E22 CONFIDENTIAL H <5tesdy flow Unsieaay -flow Separated flow NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS CONFIDENTIAL Figure & — f-~/ow conditions ai test secilon for <3n airplane lift coe-fficierii of 0.-4-3. NACA CB No. L5E22 Fig. 9 c I .8 S^» CONhDENfTIAL q Smoo fh fio w + Unsteady flow x Separated flow beginning of separation Separated flow .2 .4 .6 L /ft coefficient, C L Unsteady flow ■Smooth flow NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Figure 9.- Flow conditions indicated by tuft behavior. Fig. 10 NACA CB No. L5E22 5 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS C ONFIDENTIAL ■beginning of se para fion , Q<395< •$ <0.5Z5 Beginning of unsteady f/ow^ 0.39S< %