Wing section outboard of flap was tested by wake surveys in Mach range of 0.25 - 0.78 and lift coefficient range 0.06 - 0.69. Results indicated that minimum profile-drag coefficient of 0.0097 was attained for lift coefficients from 0.16 to 0.25 at Mach 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 Mach was exceeded by 0.025

Daum, Fred L.

Zalovcik, John A.

English

NASA Technical Reports Server

L 
ACR No. L6B21 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
ORIGINALLY ISSUED 
March 1946 as 
Advance Confidential Report L6B21 
FLIGHT INVESTIGATION AT HIGH SPEEDS OF PROFn.E IlRAG 
OF WING OF A p-47D AlID'LANE HAvmG PRODUCTION 
SURFACES COVERED WITH CAMOUFLAGE P.AINr 
By John A. Zal.o\'"(~1k and Fred L. Daum. 
Langley Memorial. Aer onautical. Laboratory 
Langley Field, Va. 
PROPERTY OF JET PROPULSION LABORATORY U 
CALIFORNIA INSTITUTE Of TECHNOLOGY 
WASHINGTON 
NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of 
advance research results to an authorized gr oup requir ing 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 r eproduced without change in order to expedite general distribution. 
L - 98 
J 
https://ntrs.nasa.gov/search.jsp?R=19930092751 2020-06-17T00:58:42+00:00Z
l 
NATIONAL ADVISOEY Cm .. TI':iITrrEE FOR AERONAU;rICS 
ADVANCE CONFIDEjJTI AL REPORT 
:2'LIGlIT IN~STIGkrIOl'T AT HIGH SPEEDS OF PROFILE DHAG 
OF WING 0:8' A P- l.((D AIRPLALTE HAVIHG PrlC'DUCTIOJ'J 
SlJRF ACES COVERED WI'fII C At,\WUFLAGE PAINT 
By John A. Zalovcil\: and F::,ed L . Dawn 
SlJ1VIMARY 
A flight invest~gation v'Jas Ii1<lde at high speeds to 
deterr.line the profile draG of a P -J~7D airplane ;::ing 
having p roduction s'..lrf aces covered \',· l t:~l c a1ilouf~age 
paint . Th e p rofile drag of a v/ing sect";"on somewhat out -
board of the flap wc:..s determln8d by meaJ.'s of ·:la~{e surveys 
in tests made o ver a 1'lE..l1<:se of aIrplane ll.ft coefficients 
from 0 . 06 to 0 . 69 and airplane l:ach numbers fror.l 0 . 25 
to 0 . 78 . 
The resu lts of the tests indlcated that a ni::1:L1':lU.n1 
profile - drag coefficient .of 0 . 0097 wc:..s attained for 
lift coefficients from a . 16 t o 0 . 25 at .. !ach nUlilbers l ess 
than 0 . G7 . Below the :.~ach nU::lber at whicJ.l conpressl-
bility shock occurred , variatJ.ons iLl. Lach nwnber of as 
nn-wL as 0 . 2 appeared to have no effect 0~1 l'r ol'ile -d::''lag 
coefficient . The variation in Reynolds m.U1lber corre -
sponding to this vaI'lation in j',.c.ch nLU~lbe r, however, Vfas 
aI)]) reclable ane. ma::' r.ave had SOlHe effect on the results 
obt aii1.ed . Comp arlson of t}:e i,felch number at w~'..ich shocl-
l osses were first evident in the TIlli.e ~ith the cr~tical 
mach nu::aber inc.icated tl1 nt shocL ';'!a3 not evider~t unt il 
the critical r,IaC~ l nu.mber wc:..s exceedt:;<.'!. b:, at least 0 . 025 . 
L \:TFlO DUCT IO::J 
A flight investigat~on was made to determine the 
pl'ofile - d2ag cl-:.aracteristics of a P - L,.7D airpla...'1.e 
l 
2 CONFIDENTIAL N AC 4. . v rt j'fc . L61321 
, j ng wi th var lous surface fini shes . Trw pb.8.ses 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 
bow1dary- layer- transit i on locations and profile drag of 
a wing section with faired and smoothed surfaces . In 
r eference 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 
forvvard . The results reported herein are of tests made 
to de termine the profi le drag of a wing section having 
unfaired production surfaces covered with camouflage paint . 
Tile p res ent tests and those of references 1 al!d 2 included 
Mach nu..rnbers through the cri tical values; in t he pr'esent 
tes ts , hov.,rever, the IVlach number range was extended to 
som~what higher 3upercritical v a lues than those of ref-
erences 1 and 2 . 
Profile drag was determined by means of wake surveys . 
The tests were made for condit i ons in which airplane lift 
coefficients from 0 .06 to 0 . 69 , Re~Tnolds nUl'nbers from 
8.4 x 106 to 23 . 1 x 106, and Mach numbers from 0 . 25 to 0 . 78 
were obtained . 
APP liRA 'l'US AN D TES TS 
The investigation was conducted on a ri ght wing sec -
tion of a P-47D airplane (fi g . 1) . This wine section , 
a Republic S- 3 section , had a chord of 86.05 inches, a 
thick ness of 11 pe rcent of the chord , and was located at 
63 "pe rcent of the semis!)an from the plane of synunetry , vI' 
about 2 feet outboard of the flap . At this spanwise sta-
tion the test section included t he ai leron but was out -
board of the propeller slipstream, the gun ports in the 
lead ing edge , and the shell ejector slots in the lower 
surface . The measured ordinates of the test section are 
given in fY'2.~tions of the chord in table 1. The Republic S- 3 
section test3d has p ressure - distribution characteristics 
similar to those of the NACA 23011 airfoil. 
The surfaces of the test section were prepared by 
covering the producti on surfaces with one coat of zinc 
chromate primer , one coat; of c.,ray surface r , and two coats 
of olive - drab carnouflaCe p~ln t. Measurements of surface 
roughness mGtde by 'TlE;anS of a shop mIcroscope (described 
in reference 2) Indicat ed that the surface roughness con-
CONFI DENTIAL 
CONFIDENTIAL 3 
sisted of particles of about 0 . 0012 inch in height and 
nwmbering roushly 10 , 000 per square inch. 
An indication of surface waviness was obtained by 
:neans of a curvature gage (fj.g . 2) 'VlTi th legs spaced 
1+ percent of the test section chord. The Naviness con-
dition of the unfaired ~nd rou£hened production surfaces 
and also of the faired and smoothed surfaces of refer-
ence 1 is indicated in figure 4 by the plot of the 
w'aviness i~riex dlc aeainst sic, where d is the de-
flec tion of the ct::.rvature gaE.';e , 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 
ral-::e (fig . 3) located 19 percent of the chord behind the 
troiling 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 (mak-
ing a total width of 23 . 9 inches) in order to permit a 
survey of more of the wake at sunercritical speeds t~an 
in references 1 and 2 . Nake total and static- pressures, 
fr e-stream impact pressure, and the position of the rig1:lt 
aileron were measured with ~ACi recording instruments. 
'rhe section prof . Ie - drat:; coefficients cdo were determined 
by the integrating method of reference 3; that is, the 
total - pressure loss was intecrated across the wake and 
then multi~lled oJ factors dependin8 on free-stream im-
pact pressure, mazimum total - pressure loss, static pressure 
in the wake , and fli gh t Mach number. . 
The tests were made in level flight, dives, and turns 
at 20 , 000 feet and over a range of calibrated airspeeds 
from 150 to 415 miles p er hour . The airplane lift co-
efficient CL obtained in the tests ranged from 0.06 to 0.69 , 
the Re~Tnolds number R from [3 . L!. x 106 to 23.1 x 106 and the 
Mach number M from 0.25 to 0. 78. 
RESULTS i·\.ND DISCUSSION 
The investigation of flow conditions indicatec by 
surface tufts located over a portion of the upp er sur-
face of the P-47D airplane wire., reported in reference 4, 
showed that somewhat 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 0.40 and 
CONFIDENTIAL 
--, 
4 
Breatsr than 0.76 at a lift coefficient of 0.15 . Be -
cause of this flow cond ition and the fact that the wake 
at !Viach numbers greater than 0.66 at a lift coefficient 
of 0 .40 a nd greater than 0 . 72 at a lift coefficient of 
0 . 15 extended beyond the limits o~ the wake - s urvey r ake , 
the ·Naka surveys for thea e flight c ondi tions were n o t 
eva l uated. 
profile - drag coeffi cients s elected for several lift 
coefficients for which the dat a were most comp lete are 
p lotted against Mach number in figure 5. The correspond -
ing Reynolds numbers are p lotted above t h e p rofile- drag 
curves . 
Figure 5 shows that the profile - drae coefficient de -
creased wi th lift coefficient . and attained a minimum 
value of 0 ~ 009 7 over a range of lift coeff icients from 
at least 0 . 16 to 0 . 25 at Mach numbe rs below 0 . 67 . The 
minimum va lue of the p rofile-drag coefficient o f the 
faired and smoothed surfaces re ported in reference 1 was 
0 . 0062 . At Mach numbers below that a t wh ich compreSSibil-
ity snock vas evi dent , as indicated by the rapid in -
crease in profile-crae; coefficient, v ariati on in Mach num-
ber of as much as 0 . 2 a pp eared to have no effect on the 
p roflle- drag coefficient . This variation in Mach number , 
howev'3r , was accompani..ed. by an a ppreciab le variation in 
Reynolds number, which may have had some effect on the re-
sults obtained . :n the tests of reference s 1 and 2 var -
iations in Hach number of as niuch as 0 . 16 , 'Nith ne s ligi!J l e 
variation in Reynolds number , had no effect on t he profi l e -
drag coefficient for the wing section with smooth surfaces 
and for the wing s ection with smooth and rough surface s 
with t r ans ition fixed far f orward . 
'l'he fli ght Mach number a nd airplane lift coefficient 
a t which compressibility shock losses became evident in 
tr-_e wake are shown by figure 6 . The r ap idly increl::l.sing 
wi d t h of wake with Mach number is shown in figure 7 as a n 
ind ic at ion of t _'le presence of compressi b ili ty shock loss e s 
in the wake . In this figure the total-pressure lo ss across 
the wake is Dre sented for several Mach numbers at a lift 
c oe f l'icient of about 0 . 16 as a plot of t;H/qc against y /c, 
where ~H is the lo s s in total p r essure at p osit i on y in 
the wake , qc is the fr ee - stre am imract pressure , and c 
is t he chord of the wing s e ction. (Position y/c = 0 cor-
responds to the top tube of the rake .) Wake profiles for 
Mach numbers 0 . 61~. and 0 . 67 showed no evidence of shock , but 
p rofi l es for 1!.ach num!Jers 0 . 68 , 0 . 70 , and o. 78 ind icated 
shoc k of increasinc intensity on the uppe r surface . 
CONFIDZNTIAL 
I ______J 
CmJFI D~WJ.'IAL 5 
In fig 'lre 6 thE; denarcat i on of flight conditions 
with respe c t to the preser.ce or a~sence of shock losses 
in the wake is well de~lned> At lift coefficients of 
0 . 10 a.1d 0 . 50 , shock -,vas first i'1d5cated at :'luch numbers 
0 . 69 and 0 . 62: respective l y. 1he ~irst indications of 
shock in the wake as s:lOvm by figure 6 corres:-,ond to the 
b0ginnlnr of the rapid increase in profile - drag coef-
fic i ents in figure 5 . 
~le critical Mach nunilier for ~~e win[ section h&v~ng 
pror~uction surfaces covered wIth. c:.il1\oufl&ge :?8.int was not 
determined . '::.1:.e critical l'vlach nUJ:l;Jer, for the cOr'responc:ing 
l eftwill[: section VIi th [air ed an:] .sn:o:Jthed 3urfac0s ,hON -
ever, was de~ermined in testa reDorted in r3ference 1 . 
TlliE c ritic&l Fb.ch numD E; r , s: o-.'<;n .~,lctted 1n flrure G, Y'1ay 
be as HlUch as 0.03 too hl€;D. for t~.Le ~resent tests becduse 
of the netho r1 of :'1easu:c>ing t he c:lo::,dwise r-ressvre distri -
bution 2nd beCE:.use the left ai10rol1 was deflected u~'warcl 
about 30 and tr_e r:cht aileron was deflected dONDdard 
about 10 . Co::-'pari son of the ~i!ac~l n"mb~r [.. t -,ir.ic!l shock 
W3.S first evident :n tne wa..'Lce Wit;~1 this critical :dach 
niJ.r1ber indicates that; S~10C',{: losses ;ere not evicent in 
the wake :.:ntil the critical r,;ach number -t/D..s exceedec by 
at l east 0.025 . ; sirrilar result was obt~inGd in refer -
ence s 1 and 2 exce,t t~at , in re~e~ence 1 for the f~ired 
a~d s~~oth wing sect~8n , the critical Mach number w&s 
exceeded by at lea.:'!'.; 0 . 04 . The a;lpE;arance of sh~clr in 
t~1e wake of the -,.mfaired and. rouel18 iled Drod-.1.c ti Oil sur-
f aces at [l lower ~_ach number than ~or the fd _~l'ed and 
s lTJo oth 3urfaces may be associated vith a lo 've r crl.tical 
Mach numb:,r for the unfaired cmd -rou~henec: 'Jr'ot3uc bion 
su:::,f&ces t:tan for :'he [aired &r:d srJ~~oth :::,ur1'clces . 
C OHCL l!S IOEE: 
T.1G flight investi:;8.tio11 of t:18 [:::,ofil'3 dr'S: of 
a P-47D airDl~ne w!ng h2vinc rro~~ctl~~ Eurface3 covEred 
with c a?'1ouflage paint Ll0.icatec. t~1e fol2.')·Ni~1:::: rE.sults: 
1. A minirUl!l ::;rofile - dri.l[ coe~f:'cient of 0 . 0'J97 
wa~ a t tai~ed for airDl~ne lift coe~f.l.c~ ents f:::'om 0 . 16 
t o 0 .25 at Mach nu:nbers belo,'oj 0.67 . 
2 . BeLYN the if13.ch nur:iber at :fhler:,. cOi""j:ressioili.ty 
shock NaS evident , as inilcated by rLpi~ rise in ~rof'ile ­
drag coeff'icient , v:J.riat5.on in N::3.ch number' of as r.1Uch as 
6 C Ol:lTFIDE1JTI AL N /l.C A A.C RHo . L6 1321 
O. 2 a~:;~-'8 ar3d to have l10 ef:i.' e c t on pI'ofile - drag c oei'f i -
cient . The v&ri a~l.on in ;~8 ~Tno lds number th Llt corresl-:on,·l.ed 
to t his varj_E~tion in iv1<'.. c ll nur:tber' Vl2S appreci ab18 EUl<l :11O:Y 
h_ve ~ad some effGct on the results obtained . 
3 . CO:':1parison oi' the :.;ach l1Ul'I'.ber at \·,ih1. ch 8~lOC]~ 
103;:'e8 "vere fi,rst e v ident L'1 t ho ,.-aka w~~t!:1 the cl."' j.t ic al 
T{ach lTLu:lber cf t1'1e \:rinz ~;Gctio:J. 1'1i th fa:_:!.~ec1 m:d snoot?:. 
8u:rf a c es L1dicatecl t~la'c CO:.lr r 8 S fll.bilit~T sb.ocl: losses 
v.'e1 e n o t eviden-i:. untI l the cj:,i t .i.cal LI2.Gh nUJ':lbe r '.'f8.S 
e~CeGd8d by at least G. 025 . 
r an(J'lp~T nC,e'I10r'j ~Il j\ol' O':1,,;·-t 1 caJ T abor",o:-t"'y,y 
.L..J <.:) 'J J r,~ I __ c:. -~ 1. e. .. \e.._ . _ .LJ...... CI. v .... 
National Advlso:i.:'y CO;'.lD,i ttee for Aeronautics 
Lan g ley Fie~d , v ~ . 
1 . Zalovcik, John A.: FliGht Investigation of BounJary -
Layer Emd Profile - Dr 3."s CharacterL:-::tics 0:; S}'.100t~1· 
,.,. S " , p L '7D l ' 1 ' J"r'! \ " 
" lng e c G.".. ;J2'13 ~l a .I. - 1_ r 1.7.l"' J:.i ane . .i h',J Ii "lCl. 
No . L5Ella , 19~5 . 
2 . Za:l.ovcll{ , J O:ll1 L , and '.'';ood , Clotaire ~ j~ Fl<:.ht 
Investlgatio~ of the Effect of Surface nouCl~G8s8 
on ~i~g Profile Drag nith Tra~sition Fixed . 
"T" 0A ;,\,p o ilJO TI, r 2r.: lC1I"I" 
.l'i J:-~v ... .4 i. .... {_ • ..LJ~·r.LI ,j , .... '--tL+ • 
3. 311 verstein , A., anL:. Eatzoff ) S .: A S hl;) 1 ified Ee t ho d 
for Determinin.; -·il5.;''12; Profile Droa.C i!1 FllS:-lt . J our . 
1'e2"'0 . Sc i., vol. 7, no . 7, ~,=aJ- 19L.o , pp . 295 - 302 . • 
4. Wood, Cl.otaire , aYld Zaloycik , J o::--.ll"l A.: i?light Inves -
t :;,gat ion at Eigh Speeds of :5"lo·. ~- Cond:Ltior.s OV01: all. 
~ l r plarle ',nn£:; as I nd :'.ca.ted by Surface Tufts . iJACl~ 
CB No . L5E22 , 1945 · 
CO N11' ID:::;;I'·:T I AL 
~- - -- - - --,~-- ~-- --- - -- ---- , - --
--~--- - -.--~~~~---~~-~---
ON P _ I,'7D / T iJ "D .. , L'" - '~J:;' _ . . .... -1-/ ... _ ..... .;.t .... -. . ..1 
[t.Il v Dl ues gi 'len i -"1 fr a c ~:,.ons of c~lnl d; 
o7dinates noacurcd r e~ativ J ~o an ~~bi ­
t~il['_r t{ G~-Icr(i L1.:::d ", ':'CX1 i rl·;)C[~J.'")(l -cralli ::1g 
edce o~ ail~ron in l i0~ ~~c 1 tra~~ing 
)c1e.;e of flap] 
Cl 
. OJ2 15 
. 0?5 Cl 
. 0500 
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r, 
( 
NACA ACR No. L6B21 Fig. 1 
• Ul 
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CON F.I DEN T I AL 
Figure 2.- Curvature gage used in measuring surface waviness. 
CONFIDENTIAL 
z 
» 
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o 
~ 
z 
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r 
0'1 
OJ 
(\) 
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(\) 
NACA ACR No. L6B21 Fig . 3 
.....l 
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NACA ACR No. L6B21 Fig. 4 
/ox/o-4 
CONFIDENTIAL 
-
I I I I I \ 
lJ Unfaired and roughened 
8 
\ 1 production surfaces \ 
\/ 
~~ 
",r- Faired and smoothed surfaces 
\~ (reference 1) 
\ 
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die \ 
4-
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IV \- { 
III \ 
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.40 .50 
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 
20x /0" 
R 10 
o 
.0/40 
.0/20 
.0100 
.0080 
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C L 
-0- O./fo 
-+- .2.5 
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--- -----~--..-., 
NACA ACR No. L6B21 
CONFIDENTIAL 
+ 
+ 
~ WJiJ o . 00 
NATIONAL ADVISORY 
COMMITTEE FOA AERONAUTICS 
I 
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-0 ./ )( 
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CONFIDENTIAL 
.4 ,5 ,0 .7 ,8 
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Figure 5.- Profile-drag coeffic1ent of p-47D wing section having 
production surfaces covered with camouflage paint. Reynolds 
number is plotted above drag curves. 
NACA ACR No. L6B21 Fig. 6 
.8 
,7 
M 
,b 
CONFIDENTIAL 
o No evidence of shock in wake 
+ Evidence of shock in 'Wake 
--First indication of shock in wake 
---Critical Mach number for section with faired 
and smoothed surfaces (reference l) 
I :t++ 
++/ + 
~-± -t.\-+-T-~~ + + 
--r-~~ ~ +Jl 1-
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.£l.. t--
(:) ~ - t--_ 
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0 
.4 .5 .0 
NATIONAL ADVISORY 
COMMITTEE FOR AERONAUTICS 
Figure 6.- Variation with airplane lift coefficient of 
critical Mach number and Mach number at which shock 
became evident in wake. 
CONFIDENTIAL 
,5 
.4 
·.3 
.f1I1 
?,c. 
.2. 
.1 
M CONFIDENTIAL 
r-,--'--~-r-.--.--r--'-4-~ 0 O. ~4 9P + G 7 r.V _ -,---,--,---r--,---,----,---.---.---
t--t--t--f--t--+-+--+-l--4.----i x : 68 ,;# ~~ I~ •. %g ~ ~!~\\ ~~1 --t-~~~1'~-+--+-~~~4--+--~ 
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o .01 .02.. ,03 
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.04 .05 .06 .07 .08 .09 .10 
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CONFIDENTIAL 
,/I .12 ./3 ./4- ./S .1 to 
NATIONAL ADViSORY 
COMMITTEE FDI AEAOIIAUTICS 
Figure 7.- Wake prof tIe. at an eirplane lift ooeffioient of 0.16 for Beveral Maoh 
nuaberB. (Positton ~ z 0 oorreBponds to top tube in wake-survey rake.) 
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Flight Investigation at High Speeds of Profile Drag of Wing of a P-47D Airplane Having Production Surfaces Covered with Camouflage Paint

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