hJhCf^L'ff K ACR No. L6B21 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED March 19^6 as Advaoce Confidential Eeport L6B21 FLIGET INVESTIGATION AT HIGH SPEEDS OF lEOFILE IHAG OF WING OF A P-1^7Il AIRPLANE HATING PEOIOCTION SORFACES COVERED WITH CAMOUFLAGE PAIHT 4 "Bj John A. ZaloTcik and Fred L. Daum Langley Memorial Aeronautical Laboratory Langley Field, Va. NACA WASHINGTON NACA WARTIME REPORTS are reprints of papers originaUy 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. 98 DOCUMENTS DEPARTMENT 7 r '' NACA ACR NO. l6;321 3W ^^^^ NATIONAL ADVISORY COr-^/II'TTEE FOR AERONAUTICS ADVANCE COI-JEIDSiTTIAL REPORT .IGilT INVESTIGATION AT HIGH SPEEDS 0? PROFILE DRAG OF WING OF A P-It-TD AIRPLAIFE HAVING PRODUCTION SURFACES COVERED V'/ITH CiUlO^JPLAGE PAINT By John A. Zal.ovcik and. Fred L. Daum SU?vMi\RY A flight investigation was made at high sp>eeds to determine the profile drag of a P-1l7D airiDlane wing having production surfaces covered r/ith camouflage paint. The profile drag of a v/ing section soinev/hat out- board of the flap v;as determined by meaa-'.s of v;a]ie surveys in tests made over a range of airplane lift coefficients fro;ri 0,06 to O.69 and airplane I-Iach nujnbers from 0.25 to 0,73. The results of the tests indicated that a minimum profile-drag coefficient of 0.0097 ^''--s attained for lift coefficients from 0,'l6 to 0,25 ^'^ Mach numbers less than 0.67. Below the liach nuiiiber at which compressi- bility shock occurred^ variations in Mach niamber of as much as 0.2 appeared to have no effect on profile-drag coefficient. 'I!1\q variation in Reynolds niimber corre- sponding to this vai'iation in i'.tach number, however, v/as appreciable and may have had some effect on the results obtained. Comparison of the Ivlach niuiiber at which shock losses were first evident in the waice v/ith the critical Iilach. nuiiiber indicated that shock was not evident until the critical Mach nutiaber was exceeded by at least 0.025 . INTRODUCTION A flight investigation v/as made to determine the profile-drag characteristics of a p-[i.7D airplane 2 CONFIDii.WTIAL NACA : iR ITc . l6s21 v^ing with various sixrface finishes, T-vo phases of this investigation were reported in references 1 and 2, and the third and last phase is reported herein. In refer- ence 1 results were reported of tests made to determine bomidary-layer-transition locations and profile drag of a wing section v.'ith faired and smoothed surfaces. In reference 2 results v;ere reported of tests made to de- termine the effect of surface roughjiess on the profile drag of the faired surfaces with transition fixed far fojTward, The results reported herein are of tests made to determine the profile drag of a wing section having unfaired production surfaces covered -with camouflage paint. The present tests and those of references 1 and 2 included JJach numbers through the critical values; in the present tests, however, the ivlach number range w;as extended to somev;hat higher supercritical values than those of ref- erences 1 and 2, Profile drag was determined by means of wake surveys. Tlie tests were made for conditions in which airplane lift coefficients from 0.06 to 0,69, RejTiolds numb6rs fromi 8.1i X 10^ to 23.1 X 106, and Mach numbers from O.25 to 0,76 were obtained. APPARATUS AND TESTS The investigation v/as conducted on a right wing sec- tion of a F-i47D airplane (fig, 1). This wing section, a Republic S-5 section, had a chord of 36,05 inches, a thickness of 11 percent of the chord, and was located at 63 percent of the semxispan from the plane of symmetry, or about 2 feet outboard of the flap. At this spanwise sta- tion the test section included the aileron but was out- board of the propeller slipstream, the gun ports in the leading edge, and the shell ejector slots in the lower surface. The measured orainates of the test section are given in fractions of the chord in table I, The Republic 3-5 section test 3d has r.'ressure-dlstribution characteristics similar to those of' the NACA 2JC11 airfoil. The surfaces of the test section were prepared by covering the production surfaces v«fith one coat of zinc chromiate primer, one coat of gray surfacsr, and two coats of olive-drab camouflage paint. Measurements of surface rouglmess made by means of a -ihop microscope (described in reference 2) indicated that the surface roughness con- con?ide:^tial I .\CA AC:i HO. LcB21 CONFIDENTIAL sisted of particles of about 0.0012 inch in height and numbering roughly 10,000 per square inch. An indication of surface waviness was obtained by means of a curvature gage (fig. 2) with legs spaced I4. percent of the test section chord. The waviness con- dition of the unfaired and roughened production surfaces and also of the faired and smoothed surfaces of refer- ence 1 is indicated in figure l\. by the plot of the v;aviness index d/c against s/c, where d is the de- flection of the curvature gage, s is the distance along the surface from the leading edge, and c is the test sec- tion chord. Profile-drag measurements -were made with a wake-survey rake (fig. 5) located 19 percent of the chord behind the trailing edge of the test section. The rake was the same as that used in references 1 and 2 except that two tubes spaced four inches were added to each end of the rake (m.ak- ing a total width of 25.9 inches) in order to permit a survey'- of more of the wake at supercritical speeds than in references 1 and 2. Wake total and static pressures, free-stream impact pressure, and the position of the right aileron were measured v/ith NaCA recording instruments. The section profile-drag coefficients Cjjo wsi''© determined by the integrating method of reference 3>* that Is, the total-pressure loss was integrated across the wake and then multiolied bv factors dependin,?; on free-stream im- pact pressure, maximum total-pressure loss, static pressure in the v/ake, and flight Mach nujnber. The tests were made in level flight, dives, and turns at 20,000 feet and over a range of calibrated airspeeds from 150 to I4-I5 rfdles per hour. The airplane lift co- efficient Cl obtalne.(3. in the tests ranged from O.06 to O.69, the Reynolds niimber R from G.Li, x 10^ to 25. 1 x 10^ and the Mach number M from 0.25 to O.78. RESULTS .-iND DISCUSSION The investigation of flow conditions indicated by surface tufts located over a portion of the upper sur- face of the P-l^-YD airplane wing,, reported in reference i; shovj'ed that some-.vhat inboard of the test section, at 63 percent semispan, cross flow was present at I.iach numbers greater than . 70 at a lift coefficient of 0.[|.C and CONFIDENTIAL ij. CONFIDENTIAL NAG A ACR No. l6b21 greater than O.76 at a Ij.ft coefficient of O.I5. Be- cause of this flov/ condition and the fact that the v;ake at Mach members greater than 0,66 at a lift coefficient of O.Ij.O and greater than O.72 at a lift coefficient of 0.15 extended beyond the limits of the wake-survo-- rake, the wake surveys for these flight conditions were not evaluated. Froflle-drag coefficients selected for several lift coefficients for \;hich the data v/ere most coir.plete are plotted against Mach nuciber in figure 5* 'T^^^s correspond- ing Reynolds numbers are plotted above the ::rofile-drag curves . Figure 5 shows that the profile-drag coefficient de- creased 'With lift coefficient and attained a minimuTii value of 0,0097 over a range of lift coefficients from at least 0,l6 to 0.25 ^t Mach numbers below 0,67. The minimum value of the profile-drag coefficient of the faired and smoothed surfaces reported in reference 1 was 0,0062. At Mach numbers below that at which compressibil- ity shock v/as evident, as indicated by the rapid in- crease in profile-drag coefficient, variation in Mach num- ber of as much as 0.2 appeared to have no effect on the px'of lie-drag coefficient. This variation in Mach number, however, was accompanLed. by an appreciable variation in Reynolds number, which may have had some effect on the re- sults obtained. In the tests of references 1 and 2 var- iations in Mach number of as much as O.16, with negligible variation in Reynolds niomber, had no effect on the profile- drag coefficient for the wing section with smooth surfaces and for the v/ing section with smooth and rough surfaces with transition fixed far forward. The flight Mach number and airplane lift coefficient at which compressibility shock losses became evident in the wake are shown by figure 6, The rapidly increasing width of v/ake vvith Mach number is shown in figure 7 ^s ^^ indication of the presence of compressibility shock losses in the wake. In this figure the total-pressure loss across the wake is presented for several Mach numbers at a lift coefficient of about 0,l6 as a plot of AK/qc against y/c, v/here ^H is the loss in total pressure at position y In the wake, qc is the free-stream impact pressure, and c is the chord of the wing section, (Position y/c = cor- responds to the top tube of the rake.) ,Vake profiles for Mach numbers O.oJ; and O.67 showed no evidence of shock, but profiles for Mach numbers 0,6S, 0,70, and 0,78 indicated shock of increasing intensity on the upper surface, CONFIDENTIAL i>;ACA AC:t no. l6b21 COIfPID"E:i^^TlAL 5 la figure 6 the deriarcation of flight conditions with respect to the presence or absence of shock losses in the wake is well defined. At lift coefficients of 0.10 and O.5O, shock -.vas first indicated at Mach nuribers 0.69 and 0.52. respectively,. The first indications of shock in the wake as shov;n by figure 6 correspond to the beginning of the rapid increase in profile-drag coef- ficients in figure 5« The critical '.lach number for the wing section having production surfaces covered vjith caraouflage paint -vvas not determined. Tt-e critical Mach nuinber, for the corresponding left wing section v/ith faired and smoothed surf aces, how- ever, v;as determined in tests reported in reference 1. Tnis critical Mach number, shown plotted in figure 6, may be as much as O.O3 too high for the present tests because of the method of measuring the chordwise pressure distri- bution and because the left aileron v;as deflected upward about 53 and the right aileron was deflected downv/ard about 1°. Comparison of the Mach number at -which shock was first evident in the wake with this critical Mach n-umber indicates that shock losses were not evident in the wal<:e until the critical Mach number v/as exceeded by at least = 025. --^ similar result was obtained in refer- ences 1 and 2 e?:cept that, in reference 1 for the faired and sm.ooth wing section, the critical Mach number was exceeded by at lea?t O.Oii. The appearance of shock in the wake of the unf aired and roughened production sur- faces at a lower Mach number than for the faired and smooth surfaces m^ay be associated v/ith a lower critical Mach number for the unfaired and roughened 'oroduction surfaces than for the faired and smooth surfaces. coiiclusion; The flight investigation of the profile drag of a P-I17D airolane wing having production surfaces covered with camouflage paint Indicated the following results: 1. A mdnir.um profile-drag coefficient of C.OO97 was attained for airplane lift coefficients from O.16 to 0.25 at Mach numbers belov/ O.67. 2. Belo'.v the Mach number at .vhich coirpressibility shock v^as evident, as indicated by rapid rise in profile- drag coefficient, variation in Mach number of as much as CO:iFIESTT]:i.L CONFIDENT I. AI. NACA ACR IIo . l63?-1 0.2 appeared to iiave no effect on profile-drag coeffi- cient, Tlie variation in ;.ie\\nolds nuitiber that corresronded to this variation in M^^-ch nirrr.ber was appreclabls and liiav have had sor.ie effect on the results obtained. 5. Coriparison of the Mach nui.iber at '.vhlch sh.ocl-: losses v/ere first evldeiit in the valie v/ith the critical Mach nui.iber of the v/ing section v/ith faired and snooth surfaces indicated that co-..iprsssibllit7 shook losses x:eve not evident until the critical Mach nuauber was e::ceeded by at least G,02;3. Langley Mencrial Aeronautical Laboratory ITationai Advisory Co:;ii7iittee for Aeronautics Langley Field, Va. RSFHRErlCES 1. Zalovcik, John A.: Flight Investigation of Boundary- Layer and Profile-Drag Characteristics of Snooth/ Wing Sections of a P~1|.7D Airplane. IT^^CA ACR No. L5Klla, ISkj. 2. Zalcvcik, John L. , and Wood, Glotaire: A Plight Investigation of the Effect of Surface Roughness on Wing profile Drag v/ith Transition Fixed. ITACA .'iRR No. lIiL2'3, l^lik- y . Silverstein, A., and Katzoff , S.: A Simplified Method for Determining V/ing Profile Drag in Flight. Jour. Aero. Sci., vol. 7, no. 7, T.-ay l^hD , pp. 255-5OI. .'■-. V.'ood, Clot aire, and Zalovcik, John A. : Flight Inves- tigation at IJigh Speeds of Flov/ Conditions over an iUrplajTie Wing as Indicated by surface Tufts. I-IACA CB No. L5E22, 19k3 . CONFIDENTIAL II AC A ACR No. L632I COIIx^IDENTI.'iL 7 TABLE I OEDii-T/iTics or r:];pitj-3Lic s-3 v/iitg section tested OH" P-il.70 AIR?L/:^i? I /ill volues given in fraccions ei' chox-dj oi'^dlnates neacured relativs to an arbi- trary chord and v/ith inboard trailing edr~e ol' ailijron in line witn trailin;/; eel 1 lap] Staticn Ordinate Upper surfc'-co Lo\,er SLirraco .0125 .0250 .0500 .0750 .10 :i6 .25 .50 -7 1— ,[•0 .45 .50 .bO •z° .90 1,00 .0192 .OplO .oU^L • Ojy ; .O6O.V .067;.. .0706 .0711 .Oob5 .0656 .0627 .OS09 cC354 .0250 .0105 -.009p -.018IL -.0221 -.0255 -.o^oG -.035i.L -.0581 -.0595 -.0400 -.0595 -.0570 -.0550 -,0296 -.0221 -.0156 -,0060 L.E, rai:_i::s: 0.0037 1 KATION.Ui ADVISORY COLIMITTEE FOR A]]ROK.AuTICS COIfflEGNTIA] Digitized by the Internet Arclnive in 2011 witln funding from University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation http://www.archive.org/details/flightinvestigatOO NACA ACR No. L6B21 Fig. 1 < 1— I Eh C3 O O +3 CO ? (U CJ CO en U 3 CO do C ■H Li CO cd e •a U 3 I CM 0) 3 bo < E-< w o I— I o L NACA ACR No. L6B21 Fig. < da Z o o Q t— I Ee. o >> > u 3 (Q I 0) (0 » I to 3 bo NACA ACR No. L6B21 Fig. 4 /O^IO CONFIDENTIAL d/c Unf aired and roughened production surfaces Faired and smoothed surfaces (reference 1) .40 .SO NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Figure 4.- Surface-waviness index of unfaired and roughened production surfaces and of faired and smoothed surfaces. CONFIDENTIAL Fig. 5 NACA ACR No. L6B21 CONFIDENTIAL +- -X- O.l(o .Z5 .40 ■:> n V 1 A Q, r ^ ii.U* tu R 10 E X + ' NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS .0\A-0 OIZO ,0100 .O08O X r f ^J- ' [ \5 X y ? X (^ C yc ^'^ i^ 5»^ d ><' O- ^ L 't^ <^Co >~ Si. X ^ 5-+^^ n ^* X ^' ^\ -1 < Z Hi 9 It z 8 1, 1 6 ■ ■ ■ ' o + X a o X 1 ,i « 1 o • 1 ■ 3 Oi ' I i 1.,-* - 1- 1st lo iS 2S I- t K) (\J < u. g 8 o o o o o o • « o ^ o XX • • 4J *» -^ n c « O c a « e So V H •ho O 4» a n : 9 9 S o 31 o UNIVERSITY OF FLORIDA 3 1262 08104 99 UNIvfeRSIP/ OF FLORIDA DOCUMENTS DEPARTT^IENT 120 MARSTON SCIENCE LIBRARY RO. BOX 117011 GAINESVILLE. FL 32611-7011 USA ll i