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ACR Ho. LÖB21

NATIONAL ADVISORY COMMITTEE FQR AERONAUTICS

All NO.

WARTIME REPORT ORKJINALLy ISSUED

March 19U6 as Advance Confidential Report L6B21

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FLIGHT INVESTIGATION AT HIGH SPEEDS OF PROFILE DRAG

OF WUB OF A P-47D AIRPLANE HAVING PRODUCTION

SURFACES COVERED WITH CAMOUFLAGE PAINT

By John A. Zalovclk and Fred L. Saun

Langley Memorial Aeronautical Laboratory LangLey Field, Va.

OocuiPRnts

FILE COP

*nswf«i»,i).>iH5; ' tiT WASHINGTON

NACA WARTIME REPORTS are reprints o< papers originally issued tu provide rapid distribution of advance research results to an authorized group requiring them lor the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were nut tech- nically edited. All have been reproduced without cnangeTn order to expedite general distribution.

L - 98

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NACA ACR Ho. L6B21

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

ADVANCE CONFIDENTIAL REPORT

PLIGHT INVESTIGATION AT HIGH SPEEDS 0? PROFILE DRAG

0? WING OK A P-1|7D AIRPLAT1E HAVING PRODUCTION

SURFACES COVERED WITH CAMOUFLAGE PAINT

By John A. Zalovcilc and Fred L. Daum

SUMMARY

A flight investigation was made at high speeds to determine the profile drag of a P-J|7D airplane •..inj having production surfaces covered with camouflage paint. T-'-e profile drag of a v/lng section somewhat out- board of the flap was determined by means of walte surveys in testa made aver a range of airplane lift coefficients from 0.06 to O.rj'.i mid airplane Mach numbers i'?or.\ O.iL'y to 0."3.

The results or the tests indicated thet a minimum profile-drag coefficient of O.OO97 v"-c attained for Lift coefficients from 0.'l6 to 0.^5 at Mach numbers lucss than 0.67. Below the . :ach number at which compressi- bility shock occurred, variations in I lach number of as much as 0.2 appeared to have no effect on profile-drag coefficient. The variation in Reynolds number corre- sponding to this variation in ': i-.ch number, however, was appreciable **nii may have had .'jov.it effect on the results obtained. Comparison of the FJrtch number at which 3hocl losses '.vo:'e first evident in the v/ahe with the cnticaJ Mach nu.iibor indicated tli.'it oliock waa not evident until the critical Mach number was exceeded by fit least ü.025«

INTRODUCTION

A flight investigation ivas made to determine the • • L Le-drag characteristics ^r a P-I4.7D airplane

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C0JIFIDE1TTIAL NACA „a He L6B21

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•wing with various surface finishes. T»o 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 boanr'.ary-layer-transitlon locations and profile drag of a wing section with faired and smoothed surfaces. In reference 2 results were reported of tests made to de- termine the effect of surface roughness on the profile drag of the faired surfaces with transition fixed far forward. Iha results reported herein are of tests made to determine the profile drag of a wing section having unfaired production surfaces covered with camouflage paint. TA<3 present tests arid those of references 1 and 2 included Mach numbers through the critical values; in the present tests, however, the fcaeh number range was extended to somewhat higher supercritical values than those of ref- erences 1 and 2.

Profile drag was determined by means of wake surveys. Trie tests were made for conditions In which airplane lift coefficients from 0.06 to 0.69, Reynolds numbers from C.q. x 10^ to 23.1 x ic6, and Mach numbers from 0.25 to O.78 were obtained.

APPaRATOS AND TE3TS

The investigation wag conducted on a right wing sec- tion of a T-I4.7D airplane (fig. 1). This wing section, a Republic S-3 section, had a chord of 36.05 inches, a thickness of 11 percent of the chord, and was located at 63 uercent of the semlspan from tho plane of symmetry, or about 2 feet outboard of the flap. At this sp&riwise sta- tion the teat section included the aileron but was out- board of the propeller slipstream, the gun ports In the learling edge, and the shell ejector slots in the lower surface. The measured orainates of the test section art* given in fractions of tho chord in table I, The Republic 3-3 section tested has pressure-distribution characteristics similar to those of the NACA 23C11 airfoil.

The surfaces of the test section were prepared by covering the production surfaces with one coat of zinc chromate primer, one coat of gray surfacer, and two coats of olive-drab camouflage paint. Moacuraments of surface roughness made by means of a ahop microscope (described In reference 2) Indicated that tr>e surface roughness con-

CONPIDENTIAL

W3A AC?.. No. L6B?.I CONFIDENTIAL

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sisted of particles of about 0.0012 Inch In height and numbering roughly 10,000 per square inch.

An indication of surface wavlness was obtained by •.leans of a curvature gage (fig. 2) with legs spaced i; percent of the test section chord. The «avlnesa 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 I4. by the plot of the v/avines.- inisx d/c against s/c, where d i3 the de- flection of the c-rvature 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 «ith a wake-survey rake (fig. 3) 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 wtr-e added to each end of the rake (mak- ing a total width of 2?.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, frefe--strea a imp i-t pressure, and the position of the right aileron wore mea-urel with NACA recording instruments. The section prof'.le-.lrag coefficients cdo were determined by the integrating mathod of reference jj that is, the total-press are loss was integrated across the wake and then multi->lied by factors depending on free-stream im- pact pressure, maximum total-pressure loss, static pressure in the wake, and flight Mach number.

The teats wore made in lev-3l flight, dives, and turns at 20,000 feet and over a range of calibrated airspeeds from 1'jO to Ü15 miles per hour. The airplane lift co- efficient CL obtained in the tests ranged from O.06 to O.69, the Reynolds number R from O.L;. x 10^ to 23.1 x 10^ and the Mach number M from 0.25 to O.78.

RESULTS AND DISCUSSION

The investigation of flow conditions indicated by surface tufts located over a portion of the upper sur- face of the P-VfD airplane wing, reported in reference I|, showed that norr.ewhat inboard of the test section, at 63 percent semispan, cross flow was present at Mach numbers greater than 0.70 at a lift coefficient of O.I4.O and

CONFIDENTIAL

CONFIDENTIAL NACA ACR )!' . I.C.B21

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greater than O.76 at a lift coefficient of 0.15. Be- cause of thl3 flow condition and the fact that the wake at Mach numbors greater than 0.66 at a lift coefficient of 0.1)0 and greater than O.72 at a lift coefficient of 0.15 extended beyond the limits of the wake-survey rake, the -«alte surveys for these flight conditions were not evaluated«

Profile-drag coefficients 3olected for several lift coefficients for which the data were most complete are plotted against Mach number in figure 5» Tne correspond- ing Reynolds numbers are plotted above the profile-drag curves.

Figure 5 shows that the profile-drag coefficient de- creased with lift coefficient and attained a minimum value of O.OO97 over a range of lift coefficients from at least O.lo to 0.25 &t Mach numbers below O.67. The minimum value of the profile-drag coefficient of the faired and smoothed surfaces reported in reference 1 was O.OO62. At Mach numbers below that at which compressibil- ity shock was evlJent, as indicated by the rapid in- crease in profile-drag coefficient, variation in Mach num- ber of aa much a3 0,2 appeared to have no effect on the profile-dray coefficient. This variation in Mach number, however, was accompanied, by an appreciable variation in Reynolds number, which may have had some effoct on the re- sults obtained. In the testa of referencea 1 and 2 var- iations in Mach number of as much as 0.16, with negligible variation in Reynolds number, had no effect on the profile- drag coefficient for the wing section with smooth surfaces and for the wing section with smooth and rough surfaces v/lth transition fixed far forward.

The flight Mach number and airplane lift coefficient at which compressitillty shock losses became evident in the .va'.ie ara shown by figure 6. The rapidly increasing width of wake with Mach number is shown in figure 7 as un indication of the presence oi compressibility shock losses In the wake. In this figure the total-pre3sure loss across the wake is presented for several Mach numbers at a lift coefficient of about 0.16 aa a plot of AH/qc against y/c, where \t[ is the lots in total pressure at position y in the waks, q0 Is the free-stroam impact pressure, and c is tho chord of the wing section. (Position y/c = 0 cor- responds to the top tuba of the rake.) ,Vake profiles for Mach numbers 0.6!(. and O.67 showed no evidence of shock, but profilos for Mach numbers 0.68, 0.70, and 0.7G Indicated shock of increasing intensity on the upper surface.

CONFIDENTIAL

If AC A ACH Ho. L6B21 CONFIDENTIAL

In figure 6 the demarcation of flight conditions with rospecb to tho presence or absence of 3hock losses in the wake is well def:ned. At lift coefficients of 0.10 and O.50, shock was first indicated at Mach numbers O.65 an? 0.62. respectively. The first indications of shock in the wake as shown by figure 6 correspond to the beginning of the rapid increase in profile-drag coef- ficients in figure 5«

The critical "ach number for the .vlng section having pro'iucüion surfaces covered with camouflage ~<alnt was not determined« The critical Mach number, for the- corresponding left siriiif section with faired an;! smoothed ."jurfactin,how- ever, -vac determined in tests reported in reference 1. This critical lach number, shorn plotted in fjfure C, nay be as touch as 0.0J too high Xor the -rasant tost3 because of the method of measuring tho chordwise pressure distri- bution and because the left aileron was deflected upward about 5° and the right aileron was deflected downward about 10. Comparison of the Mach number at which shock was first evident in the wake with this critical Mach number indicates that shock losses were not evident in the wake until the critical Mach number was exceeded by at least 0.025. A similar result was obtained in refer- ences 1 and 2 except that, in reference 1 for the faired and smooth wing section, the critical Mach number was exceeded by at leapt O.Ojj. The appearance of shod: in the wake of the unfaired and roughened production sur- faces at a lower Mach number than for the faired and smooth surfaces may be assoiiated with a lower critical Mach number for the unfaired and roughened nroduefcion surfaces than for the faired and smooth surfaces.

C0NCLÜSI0KS

Tho flight investigation cf the profile drag of a P-!.;73 airplane wing having production surfaces covered with camouflage paint indicated the following results:

3. A minimum profile-drag coefficient of O.OO97 was attained for airplane lift coefficients from 0.16 to O.25 at Mach numbers below O.67.

2. Below the Mach number at ,vhich compressibility shock ./as evident, as indicated by re.pld rise in profile- drag coefficient, variation in ?.*ar:'i number of cs much 33

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IDKWTIAL HACA ACR Ho. L6321

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0.2 appeared to have ao effect on profile-drag coeffi- cient. The variation iu Reynolds nuiaber bh A ccrrearxaiAed to taia variation in Mach nunber v;e.3 appreciabla and r.iay have hRd some 9fffcot on the results o"ütt>i:ieü.

3« Comparison of tr<e "..;üC> iiuivber at which shod; IOSEC;: were first evident in Li^o .;ako with tins critical Kaeh nuiiber of the win• section with faired :md snoot!; 3iirfacs3 i.:C.c ited that ca.pressibilit shod: losses were not evident until the ci'ib:.cal Uach nunber was exceeded by at least 0.025-

Langley Memorial Aeronautical Laboratory national Advisory Coi".nnittee for Aeronautics

Langley Field, Va.

REFSREITCSS

1. &alovcik, John A.: Plight Investigation of Boundary- Layer and profile- iractcjristics of Ssiooth Wing Sections of a P-I;.7D Airplane. II AC A MR Mo. LSIMia, 15^5.

2. Zalcvclic, John ".., and Wood, Qlotaire: A plight Investigation of the iCffoct of Surface Rougl r.sss on '.Ving Profile Drag with Transition Fixe .. IIACA AP.R So. Ll|i2^, i?l+U-.

'•j, 3ilver3tein, A., anC Katzoff, ,-...; A Simplified '.vt id for Determining '.. ; Profile ira Ln Flight. Jour. Aero. Sei., vol. 7» no. (, I.'.ay L ' , ;;%. 255-30-»

I4.« Wood, Clotaire, and Zalovclk, John A.: Flight Inves- tigation s.t High Spaods Conditions over Airplane "*.'...-" as I dicated V", jurface Tofts. IIACA

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Figure 4.- Surface-wavinees index of unfaired and roughened production surfaces and of faired and smoothed surfaces.

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Figure 6.- Variation with airplane lift coefficient of critical Mach number and Mach number at which shook became evident in wake.

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AUTHCHHS)

DIVISION: Aerodynasdce (2) SECTION

CROSS REFEI Änga - Profil» (31230)

IATI- 6,77 Mug« and Airfoil» (») _ ^.^ EtENCEs/Praasur« 0<lairlbutlar'-glnga?f74500); Lift distribution IW-TJUJ; Dr»g7"rodjna"le

CMNG. AGENCY NUMM«

ACB-L6B21

REVISION

AMEt. TITLE,

FO«G'N. TITLE,

Plight investigation at high speeds of profile drag of wing of a P-47D alrplana having production surfaces covered with eaaoufläge paint

0«K3INAT1NG AGENCY: Rational Advisory Conwdttee for Aeronautics, Washington, D. C. TRANSLATION, a COUNTRY

D.S. LANGUAGE NSwSRZEOf u. SJCLASS.

One lass. DATE

•ar'46 MOB

14 uus. 8 photo»

FEA TIMES tablaa, graph»

ABSTRACT Wing •action outboard of flap was tested by wake surreys in kach rang» of 0.25 - 0.78

and lift coefficient rang» of 0.06 - 0.69. Results Indicated that arlnimuw. profile-drag eoafficinat of 0.0097 was attained for lift coefficients froa 0.16 to 0.25 at kach less than 0.67. Below Mach number at which compressibility shock occurred, variations In Mach of 0.2 had negligible effect on profile drag coefficient. Shock was not evident until critical kach was exceeded by 0.025.

NOTE: Bequests for conies of this report Must be addressed to: I.A.C.A., Washington, D. 6.

T-J, HQ, AK MATEKIEL COMMAND TGTT ECHMCAl INDEX WRIGHT FIELD, OH». USAAF JOS wfLO-ai aua « MM

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lEKIC'i i;ULLT; tiATED 31 —J- 1 V7.

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ZslcTlels, John Denn, Fred L.

AUTKoars) AK£U. Trrtfj

FOOCN. TTTUi

'wvisONi Aorodynorieo (3) IssCTtON, Dingo and Airfollo (6) I COOS. CffEEaects. Prossuro distribution - Kings (74500); 'tangs ~ Lift distribution (99170); Drag, Aorodynanie - Profile (31220)

,70° 6j>77 COtG. AG04CY NUMBEQ ACB-I/>B?1

BEVtSK»)

Plight investigation at high opseds of profile drag of wing of a P-47D airplam having production surfaces eovorod uith c acsuflags point

oaKuNATlXG AC1 iCYi national Advisory Conolttoo fbr Aeronautics, nachington, 0. C. T0ANSIAT1ON,

c COUNJBY U.S.

lANGVAGl lOG'NUASSl "• SJCLASS. (tec lass.

DATt Bar146

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FEATU&S tables, graphs

ßOOTDflCT Qing section outboard of flap ess tooted by nake ourvoys In üch rango of 0.2$ -0.78

and lift coefficient reogo of 0.06 - 0.69. Hooulto indicated that ajnlaa profilo-drag eceffieinot of 0.0097 aas attoinad for lift coefficients froa 0.16 to 0.25 *t Each looo than 0.67. Belon Bach nuobar at nhich eo^prcaolbility dxoeH occurred, variations in lisch of 0.2 had negligible effect on profile drag eoaffieicnt. Eioch cas not ovident until critical täeh cas ozcesded by 0.025.

Crar: Haqueste forcopies of this roport oust be eddrooscd to: O.A.C.A.,

~w YSTJGHT R£lO. OHX), USAAf CMMI COD a rra