W^CA-L"f I ^ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED November 19^5 ae Advance Confidential Report L5G10 EFFECTS OF CCMFKESSIBILITY OK THE MAXIMUM LIFT CHARACTERISTICS AHD SPAHWISE LOAD DISTRIBDTION OF A 12-FOOT-SPAN FIGHIER-TYPE WHJG OF KACA 2^n-SERIES AIRFOIL SECTIOHS By E. 0. Pearson, Jr., A. J. Evans and F. E. West, Jr. 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. DOCUMENTS DEPARTMENT , , 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/effectsofcompresOOIang ^ACA ACR IJo. L5G10 KATIOI.'AL ADVI^OR^ COr.'P-^ITTEE FOR AERO!TAUTIC"> 7 U <.2M^I DVAKCE COI\TFIDEJJTIAL REPORT EFFECTS OF COMPRESSIBILITY ON THE L^AXIMUM LIFT CHARACTERISTICS AND SpANV;iSE LOAD DISTRIBUTION OF A 12 -FOOT -SPAN FIGNTER-TYPE WING OF NaGA 250-3ERIS3 aIRFOIL SECTIONS By E. 0. Poarpon, Jr., A. J. Evans and. F. E. Vs'sst, Jr. SLNSiARY Force and :ores3ure-di strloution me a sur omenta were made on a fighter-type wing model of conventional NACA 230-3eries airfoil sections in the Langley 16-foot high-speed tunnel to dete"^mins the effects of compressi- bility en the rnaximun lift characteristics and the span- v/ise load distrlbxition. The range of sngle of attack investigated was from -10° to 24'-'. The Mach number range was from; 0,20 to 0.70 at small and mediu:n angles of attaclv and from. 0.15 to 0.625 at very large angles of attack. In the Mach numiber range from 0.15 to 0.55, the miaxim_um lift coefficient first increased with increasing Mach number and then decreased rapidly after ha\''ing reached a peak value at a I.Isch nui-foer of 0,50, At Mach nu-mbers higher than 0,55, the rate of decrease of maximiUm lift coefficient with Nach numiber was considerably reduced. At these hifh.er speeds the lift coefficient continued to increase with angle of attack v/ell beyond the angle at which marlced flow separation or stalling occurred, and the miaxirron lift coefficient vvas reached at angles 10° to 12° beyond the stalling angle. No significant changes in the span load distribution were found to occur below the stall at any of the test speeds, '.Tnen the wing stalled at high speeds, the resultant load underwent a m.oderate outboard shift, which resulted in increases in root bending mom.ent up to about 10 percent. CONFID?XTIAL NACA ACR No. L5G10 INTRODUCTIOII y/ind-tunnel te^ts <)f a rectangular wing of NACA 0012 airfoil section (reference 1) shov'^ed that the marririiuia lift coefficient reached a reak valae at the lov/ Mach nu:r!ber of 0.19 and deci'^eased rapidly as the Mach nurnber M v«as increased from this value up to the highest Mach number of the tests (M '^- 0.35) . Althoixgh these tests were necessarily liraited in scope, they indicated the Importance of a knowledge of the effect of compressibility on the maximvjn lift coefficient both In the estimation of the maneuvering performance and loads of high-speed aircraft and in the- interpretation of wind-tunnel maximum lift data as applied to the -prediction of airplsne character- istics at low .speeds. M'^re- recent two-c linens i'>nal. wind-tunnel tests of a number of propeller- type ^irfo.lls over a relatively large Mach number range (reference • 2) showed effects for the thicker airfoils similar :-to, those .o.f reference 1 and in addition .^'.howed large • increases in the maximum lift coef- ficient starting at. Mach -numbei-s of about 0.5. Plight tests of fighter airplanes reported in references 3 and 4 shov.red. large decrea-f,es in the lift coefficient corre- sponding to the stall .up to Mach numbers of about 0.6. A high-speed mn.nd-tunnel investigation of a number of three-dimensional v/ings of different airfoil sections has been undertaken to provide m.ore detailed information on the .high-speed stalling phenomena. Measuremient s to determine the' effect of compressibility on the spanwise load distribution were included in the program because of the related imiportance of the load distribution as a determining factor of the strength requirements of wings. The present report gives the preliminary results of force . and pre.ssure measuremients In the Lanejley 15-foot high- speed tuTJiel on the first of a series of v/ings. The model tested was a fighter-type wing having an aspect ratio of 6, a taper ratio of 2:1, and conventional NACA 230-series airfoil sections. CONFIDENTIAL I:ACA ACK Mo. LSGIO CONFIDENTIAL 3 S^IBOLS V trus airspeed, feet per second a speed of sound In air, feet per second M Mach nuinber (v/a) p air density, slut's per cubic foot q aynarmc pressure, pounas per square loot iTjpV / / — \ \^ / ( pcV 1 R R e y no 1 d s nur.ib e r \ ■' j H- coefficient of viscosity of air, slugs per foot-second The foregoing synobols represent the undisturbed strear^i values. C cross-sectional area of the tunnel at the throat, s^ius.re feet D equlTRlen:, dia:neter oi' the tunnel teet section, feet (M) 3 v;ing area, square feet b. wing span, feet y spanwise distance rueasured from the pltine of 3 AT-Jine try, 1' e e t Cg airfoil chord at plane of symmetx'y, feet mean chord, feet (s/b) c airfoil chord at any sivan'.vise location, feet t Tns.;ci:iur. thicl^'ness of airfoil section correspondinj to the riean chord, feet CONPIDEIITIAL NACA ACR 2Jo. LSGIO --T n Cn ^n C,, wing lift, pounds wlnfT lift coefficient [ — (*:) / section norTrial force {force per unit span), pounds •per foot section normal-force coe load coefficient fficient (^ \ \^y / / c Vising normal-force coefficient 12 -^ / s /■' b 2 \ "(A 'n d;r a Aa LL ^ccsc corrected an.f^le of attack of the root section (section at the plane of s^T.arie try) , degrees angle-of -attack correction due to the jet ■bouncarv-induced uiDVjash at the lifting line, " /_ , - '^^ \ a function of the ratio of wing span to tunnel aiaineter ; — \ 8 \ 1 + b-'4 ^ 5 /h\8 + . " . ■is \d; - 54 ii=; ^ • • -J angle-of -at ':>ack correction due to the jet boundary- indue od streamline curvature , \'/i - ' ,2 APPARATUS AND JIETHODS A dla&ra:-;!matic sketch of the -."rir-g nodelused Irr the tests is given in figure 1. The iDrincipal dimensions given in the figure and other pertinent inf orr-atio.n are given in the following list COrpIDBIZriAL NACA ACR Ko. L5G10 GONFIDE'ITIilL Span, ft , . 12 Aspect ratio ..........<,.<,....... 6 Taper ratio .............. ....o . 2:1 Geometric and s.erod;,T:iaraic tvjist (v.-ashout) , deg , . . 4.2 Ro-t -.ection .,.'............. MAC A 23016 Tip section . . . . , . „ . y.kCk 23009 Dihedral (along the l/4 chord line) , deg Sv/eepcack (alon£^ the l/4 chord line), deg . . . , 5.18 The v/lng was of bi.iilt-up steel constr'i.ction and was machined in such a rrianr.er that surface elements connecting equal percentage-chord points of the root and tip sections ^.^'ere straight lines. Thirty- three pre.'^sure orifices were distributed over each of six v/ing sections, the spanvrise locations of which are given in figure 1. The chordwise distribution of pressure orifices for a typical section is also shown in flg\ire 1. The pressure tubes //ere brought out of the wing to multiple-tube ;rianoxaeters in the test chamber by means of the boom and movable strut arrangement shown in fig- ure 2. ?or the force tests the boora and strut were removed and the boom replaced v;ith a short fairing, which is shown in figure 1. The v;ing was mounted at the tunnel center line on shielded struts having a thickness-chord ratio of 0.15. The thickness-chord ratio of the shields v/as 0.124. "ig- ure 3 is a photograph of the wing mounted upright in the tunnel for the force tests. Most of the te^!-t riuns v^eve made with the angle of attack held constant while the tunnel speed w.e s varied from, about 150 miles per hour to the m.a:'-iirium, speed obtainable, which for v/ing angles of attack between and 4^^ was apx^roxlmately 520 males per hour. The corre- sponding Mach number range was from 0.20 to 0.70, and the corresponding range of ^average Re^rnolds mmiber was from 2.0 x 10'^' to 8.1 x 10*^. Figure 4 shows the varia- tion of average Re;^'T!.olds niomber v/itb Mach nuxnbsr. Vov very large wing angles of attack the m.a.Timu:ri obtainable tunnel speed v.^as about 460 miiles per hour, which corre- sponds to a I'lach number of about 0.625. In the deter- mination of m.axim.'orri lift coefficients additional tests were angl( lift. The geomietric anf"le-of -attack range of the tests was fromi -10° to 24°. CONFIDEIITIAL 6 GONPIDEIjTIAL FACa aCR No. L5G10 In the lo£d.-cli str.llor.tlon tests' the static pressure over the six vrlnq. sections, as indicated, by several multiple-tuhe luanoneters, vere recoi-ded photographuically The chord-ftlse pressure rllstributions deterhiined frori th.ese photographic records xvere integrated ir:echanically to find the '■■•ecti^n normal-force coefficients. '■J >~'ii.ri_i'v^ ... J- wi\i o Force dat a.- The force data have been corrected for strut tares, air- stream r:isalineinent, and tunnel-\7all eff ect=^ . The strut tare forces were determined from tests with the v;ing inverted with and v/ithout image support struts installed. A photograph, of the inverted wing v/ith the ima,c;e struts installed is given in figirre 5. The largest Lncrenents of ].ift coefficient due to the support struts v.-ere between 0.03 and 0.04. The effective inisalineraent angle of the air stream v/as determined from tests of the v/lng upright and inverted ?/ith the image struts installed and was found to be constant at 0,15'-' throughout the speed ra.nge of the tests. In order to prevent air leakage through the strut shields, thin rubber diaphrafTns were fitted around the bases of the shields. An additional correction to the lift v/as necessitated because of a r^ressure differential across the diaphragms. This pressure differential v;as measured biiring the force te?ts 'oj means of a micro- manometer, and a calibration v"as made v/lth the wing removed to determ.ine the variation of lift force v;lth pressure cif f erential. This correction was very small in the region of maxim.um. lift (less than one-half of 1 percent at all speeds) . The effects of the tunnel v/alls vv'ere accounted for by the miethods of references 5, 6, and 7 as follows? The principal part of the angle-of-attack correction given in reference 5 is A.ttt = 5 7.55 ^Ct degrees This equation is strictly valid only for the case of an elliptical spanwise load distribution. A check ca] lAGA ACn ¥.3. LSaiO COKFIDEKTIAL b7 a rnore e.:act but more detailed procedure based on the e:?:periraen tally deterTiined span loading revealed that the error incurred dt the use of the simpler forni was I'legll- gible. At a v^inr, lift coefficient of 1.0 the correction was 0.93^. An additional correction to the anrle of attack due to an induced curvature of the flov/ was calculated from the equation . - 1.05 c, j/l - LF This equation is based on the original incompressible- flow derivs.tlon of reference 5. The modification / o yl - M^ is given in reference 5. This correction amounted to 0.16° at a lift coefficient of 1.0 and a Mach nuirber of 0.6. Corrections to the stream velocity, d7y-naraic pressure, and I'.iach nurfoer^ and to the v;inr lift coefficient due to constriction effects -^A-ere calculated by the method of reference 7. The correction to the velocity is AV _ 0.6bct + ^E£I____ ■vr v/here A7 is the effective incremental velccit; constriction, 3 and H are the breadth and height of a rectangular tur^nel, and Ctjq is the wing profile-drag coefficient. The tv/o terms on the rijht of the equation give the velocity Increments due, respectively, to ''solid'' constriction and "v/ahe" constrictlono Since the magnitude of the wa.l:e constriction effect is a function of the velocit7/ loss in the "rake and the size of the wake, the correction is expressed in term.s of the profile-drag coefficient, v.'hlch is also a function of those quantities. No theoretical treatmient of the problem of constric- tion effects for a finite wing in a circular tunnel e:cists at the present time, and the foregoing relation vvas thought to represent the best available approximation. As m.odified for the care of the circular t^mnel, the equation beca.mie iV _ . 6b ct _^ ^^o^ CONFIDENTIAL JOKPIDEITTI^L NACii ACR ITo . L5G10 where H in this case is tht aver&.go height of the- t-.-'rirel In the region occ;ii:;ied. bj the -jving. ^ The constriction corrections to the dynamic pressure, Hach nunber (refer- ence 5) , and ■■,■_■ — ^ J. -■■'"-' \ Ot Q The correct lor s "A-ere 5ii?iail for lov/ ang?.es of attach over the entire "loch nm:nber range. At a georietric angle of attach of 4*^ ard a Mach nuinb&r of 0.7 the corrections to the lift coefficient and Mach nix-ubcr- v/ere, respectively,. 1.0 percent and 0.6 percent. A.t angles of attack above the stall where the drag becane very large (Indicative of a large v;ake) , the corrections as.'^ujjied f-one importance. At a ^:eor..etrlc angle of attack of 24° and a Mac?n niimber of (-;.6 the corrections to' the lift coefficient and r/^ach nur.ber v.'ere 1-.2 percent nnd 2.2 percent, respectively, Pressur ;^-ll stribuolon data,- The n'ress'-^re-dlstribution data have "o^ber. corrected for the x-'i'^l^dps-l effects of the support struts in that the free- stream values of static pressure and dynamic pressure upon \'vhich the pressure coefficients v^ere based v^ero detsniiinsd frov- a survey of the floviT in the test section with the support struts and shields Installed. Soiree sT;all local effects of the struts on the spanwise load distribution rer:aln. These effects will be discussed in the section entitled ''nesL^.lt! and Di s c u s s ion. " The effect of the tunnel Y.'alls on the spani/vise load distribution \^^as considered and found to be very srnallj consequently, these data as presented are uncorrected for tunnel-v/all interference. RESULT? A^'D DTSCUSSICIT '/.'ing lift cl'.aracterlstlcs (force test resi^lt s) . - The lift cnaracteristlcs o .7 the \-;'lng a.s a function of angle ITACA ACK ¥o. LSCrlC COT'o^IDEKTIAI Ox attack arxd. Mach nuinber are shov/n in figures 6 and 7. Figure 6 is presented to indicate by the scatter of the test points the precision with ■'."rhlch the data -.vere obtained. '^he same data with the te3t-polnt s^nnbols re^noved and with the horizontal Jines dra'vvn for constant and even values of ancle of attack are srlven in figure 7. The variation of inaxinmn lift coefficient Gt^,, ,, with ''ach niuTLber is shcv/n in figu.re 8. The maximu;-n lift coefficient increase-: v/ith increasing Mach nuraber until a M£.ch number of about 0.30 is :'"eached. This increase can probably be attributed to s. combination of Reynolds nujiiber and ''ach nujiVoer effects; hov/ever, the Reynolds nuriiber effect probably predomnates in this rer-'ion. As the Mach nur.iber is increased above J.. 30, the majcirrLUca lift coefficient decreases at an increasingly rapid rate until a •-''ach nu:'aber of abcu.t 0.55 is reached. In the region bet?;een M = 0.55 and I' = 0.625 the rate, of decrease of raaximuj:! lift coef- ficient \7ith Mach nuriber is considerably reduced. The lift-coefficient curves of figure 7 are presented again in fif.-re 9, plotted to a co;rj.".ion anP:le-of-attack scale to Illustrate more dearly the changing character of the stall as the Tach nu:nber is increased. As the Hach nuirfcer increases above 0.30 t?ie angle of attach at which the v/ing stalls progressively decreases; also, at Mach numbers below 0.55 the stalling angle and the angle for r.iaxirnur;. lift are appro:x:iriiately tl-ie same. At Mach nurabers above 0.55j however, the ma:.:imum lift coefficient occurs at an angle of a-:tack 10° to 12^ higher than that at which pronounced separation of the flov^ begins. Thene data indicate that for airplanes ^^/ith wings siniilar to the test vving there exi-'ts at high speeds a range of r-iaximur". obbainabi.e lift coRfficient. This range extends froiii the lift coefficient cor:-esponcin.v to the initial stall (such as sbowr by the lower dashed curve of figu.re 8) to tbat corresponding to the actual iv.a.xinurfi lift coefficient o*^ the wing. At the loi,?er vali-.e oi lift coefficient corresponding to the change in sloj.ie of the lift curves of flgv.re '~', increases in stability dr-.e primarily to decreases in downwash angle are likely to occur. It might be expected therefore that at high speeds the amount of elevator control available vvould be an important factor in determining the maxi.mum lift coefficient obtainable (reference 4). Thus, an. 8.irplc:ne with a limited anovait of elevator control milght be capable of reaching an angle of attnck onl:/ a fev/ 10 C0I^?IDE::TIAL I^aGA ACR :Io, L5G10 de^trees above the 3tall^ and T:be riazimun lift CDefflcients obtainable mi,f";;ht be only sll.^btly gi"eater than thoce represented b?/' the lo^^'er- c.ashed curve of flR"ire S. Tail h'?f fating Is alro likely to occur vrhen the flov/ separates frcr: the v.ing, co that pilotinf;; technique cannot be ov-rr- loohed as a possible d.eterrrilning factor. Pin.^ilj'', at vei-y high Mach numbers it is possible that actual instability Tni,2;ht be encoimter-edj, in v.-hich case hl;;5h angles of attack a-.d high ].irt? inijht be obtained inadvertently regardle?;^ of the control pcv-er. Spanyrjs e load dis tri bution .- Spanwlse load distri- bution curves for a numi::8r of values of v;ing nor":al-f orce coefficient and hac;]i nui.il;-er are Rhov'n in flv;ure 1G« As mentioned previously the principal ef;'^ect of the support struts (v/hich is to increase the effc-ctive r:t2^ea;:-: velocity) was accoi''nted fcx by calibrating the tunnel -//ith the struts installed. fhe locaJ efi'ict of the struts is to reduce tne lift at a given angle cf attach, by a small amount and to produce a >.0-ight distortion zn the span loading. This dls';:r::.on ma:y be seen as £ neair* tne sparv'f:.se s':ui:i::-n ^-u^.^ - o = oc. dne solid ana 0/2 resenz rne xoaa dashed curves of figure "- w^r-v^^c: .=•>-■)- -^--c -.r^r,.-^ ,^^ c-h--.-v.-. bei'ore and after the stall, respeccively. ll'_e crrve^" of firure 11 sur;ariariL.e the ohanres in span loading: t;''at .'ere fomd to occcr at the higher test speeds. Ylo significant chanpes in the s;^an loading were found to occur belov: the stall at an'r of the speeds of the test, even v-'hen shoch v;aves vere ■veil established over the center part cf the v;in£. Chan^^es in the span loading: were observed to take place above bhe high-speed stall, hov^ever, the center of lead beiny shifted outboard. fhese chanf^es at the hiyher speeds were found to be moderate. The largest corresponding increase in bending riO".:ent at the root (for constant lift) '.vs.s fecund to be about 10 percent at a i;Iach nujviber of 0.55 and a wing normal-force coefficient of about 0.95. The chani;^es in load distribution due to stalling at lov; speeds were somewhat larger chan those at hi.'rh speeds but are not of particular significance because the total lift cannot be tnaintained \-ey:>nd the stall. A corvparison of the experirientally determined load distribution curves at M = 0,40 with those calculated by the ]r;ethod of reference 8 is shov/n in figure 12 for values of wing norinal-force coefficient of and 1.0, NACA ACR ITo. lEGlO GOIIFIDEI^TIAL 11 Tlie agreement sho'/m is ty^racal of that e;>:i sting in the uri"tall'^d Dart of the lift curve 3. 'Vind-tiinnel tests of a tapered wing of ITACA 250-sories :il £ect:"ons at Mach n-.i»riber-s rann;in£ fror-i 0,15 to 0.7 1, The rfiaxirroti: lift coefficient fir?t increaf^ed \;ith increasing I'ach number up to a I;I;^ch nirniter of 0.5. As the 'lach nunber vre s increa^'ea above this v.::lue the inaxi- m\}n lift coefficient decreased rapidly. 2, A large reduction In the rate of decrease of naxi- mujii lift coefficient ";ith Mach nuinber occurred in the Mach number ran;;e of 0.55 to 0.625. The tunnel I.'ach nui^iber of 0,625 was the highe-t value that could be obtained at the large angles of at back requisite for riajrimi^iii lift. 3, At Hach nuTTLoers belovv 0.55 the angles of attack at Vs/hlch the nia^rinijin lift coef ■'^icient v^a.s reached and at •which .'^tailing occurred v.-ere ap;oroxiiri.ately the same. At !'ach nuirbers above 0,55 the angle of attach at v/hich the riaxinuni lift coefficient was reached ^^'as 10^ to 12'^ beyond the anr^le at v/hich the v:?in.n: initiallv ocGurred belor; the stall at any of the speeds tested. 5. hoderate chan'-es in the span load distribution occurred ".'hen the v/ing stalled at high speed, the center of load being shifted outboard. The largest corresponding increase in bendinfr r^on^ent at the \:in?? root for constant co!tfidl:i^ial CONFIDENTIAL MCA ACR No. L5C10 lift VB.3 about 10 percent and occurred at a Ifach nurnier of 0,55 a ad Ft a ^^'ing nor-rual-force cb efficient of about o OS Larigle3- llei-norial Aeronautical Laboratory National Advisory Goiranlttee for Aeronautics Langley Field, "--a.. CCNI'IDSI'ITTAI NAvOA ACR No. L^GIO CONPIDENaTAL 15 REFERENCES 1. ?Aise, Thomas C.: Some Effects of Reynolds and Mach Numbers on the Lift of an NACA 0012 Rectangular i'Vlng In the NACA 19-Foot Pressure Tunnel. NACA G3 No. 3E29, l^li^ . 2. Cleary, Harold E.: Effects of Compressibility on Maximum Lift Coefficients for Six Propeller Airfoils. NACA AGE No. Ll+LSla, l^kd ' 5. Rhode, Richard V.: Correlation of Flight Data on Limit Pressure Coefficients and Their Relation to High-3r)eed Burbling and Critical Tall Loads. NACA ACR No. LI1I27,' 19[iJ|. Ij.. Nisson, James M», and Gadeberg, Burnett L.: Effect of Mach and Reynolds Niirribers on the Pow?er-Off Maximum Lift Coefficient Obtainable on a P-39N-1 Airplane as Determined in Flight. NACA ACR l-^o . i|P28, 5. Goldstein, S,, sjid Young, A, D.j The Linear Pertur- bation Theory of Compressible Flow, v/ith Appli- cations to Wind- Tunnel Interference, R. k ¥., No. I909, British A.R.C., I9I1.3 • 6. Lotz, Irmgard; Correction of Lownwash in Wind Tunnels of Circular and Elliptic Sections. NACA TM No. 8OI, 193 6. 7. Thorn, A.: Blockage Corrections and Choking in the R.A.E. High Speed Tunnel, Rep. No. Aero I89I, British R. A. E., Nov. lyiiJ . 8. Anon,: Spanwise Air-Load Distribution. ANC-l(l), Army-Navy-ComiTierce CoFimittee on Aircraft Require- m.ents , U. S. Govt. Printing Office, .iprll I958. CONFIDENTIAL NACA ACR No. L5G10 Fig. 1 < I- n § 5p 1 I I I I I I r / 1 1 i 1 -Q / ' '^ •5; •+~ f .■0 1 t "^ ■^ ' 1 '^ ft' ' i < CO 1-4 G frH •H z Ed ra Q 3 <— 1 Cb CO z o o ■p Ij o c^ • p. M 3 0) M CJ U Q) O bo c>^ te e 0) •H S-i c6 X ■P ■P •H •P s 3 1-1 bo *^ C CO •H s td c c •H •H e V-1 u o 0) +J S 0) u o ■P tM c o u fie. in u 3 bo ■H NACA ACR No. L5G10 Fig. 6 1.6 C, O J .Z .3 .4- .S .6 .7 (y^or ex. = 0°) M -IZ -8 -4 O 4 8 /Z 16 ZO {forM=.z) cx J dea CONFIDENTIAL Figure 6.- Wing lift ooeff/cienf as a function of angle of atfac/<: and Mach number. NACA ACR No. L5G10 Fig. 7 ■■ CO 1 NFIDENTI^ XL / 4 1 / < y 7 7 /I /.2 "/ 7 / /-, 7 \ "/ i7 / /I < /. ? \ \ 1.0 /4 "^z /-- / /__ 7 / n \~~- .^ /p / --'7' ^- ' 7 7 \ ■--. -i^ 5=- ,.__^ .8 7 / / /. / '^ f'i ^t^ 3;;^ ^-r- / '°/ /_ "7 L- --7 7 7 7 -^ s7^,^. f ^ " .6 / / / / /j -7 /.. 7 ^8 ■V / y / / / A / A ,,?^ ^ / L. ~~/ /_. " / /L_ "/ / / / h /\ / / L. --/ L^ 7 .Z "a ^= / d .m / fn / ■s? /. ?// 7 / ^/ r_ "7 /_ "7 /_ '"/ /_ 7/ ^7 ■.7C O i i / / / / / // / A 7 / / "7 7 /" ' , L. '"? /_ "~; /_ / / 7 NATIONAL ADVISOBV COMMITTEE FOt AEBONiUTICS. 1 -.1 / / / / / 7 / -.4- V 7 "7 "7 /__ "7 7 V -^/ /__ /-_ /. /- --7 / /. / -.6 -8 / A 7 z --^ > y -V -/o /_ / /- ^' ^- -.a cc NFID ;nti ^L ./ .7 f/b/-^=^°) .Z .3 .4- .-5 .e M -/Z -8-4- O 4 e IZ /6 20 (for M=.Z) CC y deg FiguLre 7.- Winq lift coefficient as a. function of angie of a.tta.cK and Much numlDer. NACA ACR No. L5G10 Fig. 8 /.6 /^ J.Z W .6 ^L max .Z o CONFIDENTIAL 1 1 -^ ^ ^ \ "\ McLXimum /'iff coe.ff/cienf \ \ / / Li -ft ooeff/c/ent obfcz/nf^jri a.f ^° to 3" above, the. cmqle, of cxtfcuzk of wniah /n/f/CLl ^&pCLrcifi on of the flow from the lY/nq ocourred \ / \ ^ -\ \ \ \ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 1 1 1 1 1 CONFIDENTIAL 1 1 1 .£ .J .s .6 .7 A/I f^/QLcre S. - Veer /at/on of w/nc^ mcDC/'mum //ft coeff/c/ent iA//th /^czch number. NACA ACR No. L5G10 Fig. 9 C ■L CO NFIDENTIAL 1 1 M = a P.O. M JO \~ .8 60 ^ "- -^ " -;v .66— 'i-^ .e .673- /^ f .4- .70^ A '/ f .Z / o / / -z / -4- 1 /i '/// -6 / / NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ^ CONFIC )ENT1AL -/Z -8-4 O 4 6 /Z 16 ^O ^4 £8 /^/qure. 9!- Van at ion of w/ng lift coefficient with angle of attcLcK for several Mucfi numbers. NACA ACR No. L5G10 Fig. 10a, b QJ O UO ^ U Crf /3/('»,^y.1 1 1 1 /"OKiciricwTi Al . , . 1.? ■ ■ ^11'^ ' ~~^ ~~~~~- -^ ^ 1.0 /.2- ^ --^ -\ ~ ~~ -^^ ■ - -^ /.O ^ ^ / -. ■N ^ •*. ./.o^ ^ ( l^abovp 5fa/l) ^ ■^ '^ -~- - - — ■~^ / Q/=-?" le stall) 1 6 " ~-~ =^ ■-^ -- -- 55> t^ =^ i/'c 7bO — - .0 ~— ^ ^ ■^ X ■>j ^ A ■^ \^ s^ \ \ • \ — ■ . "~~~ ^ ^ ^ \, \ n\\ 2 ,?. ^. . , '^"^ ^ >x\V ■ ^ --^ ^ n\ O -— ^ ■ ^ -2 L— -y. 1 .£ .5' ^ .6 y rracflon of semi span , ■#■ if.) M^O.SO. ^/z a w olo /_o Si — - ~-^-^ ^ ,fl ■^■^ "^^^^^ ^ ^ ^ ( approx-2 a, hovt '.St all) 5^ ^ c:^^ ^ \\c^-^s _^ a//) 6 <: -^^5»^ ~=z^^ [^Pl JfOX .b a bove ST ^^ ■^ ■^ "^-\ c:; -^ \ ^ — = A - ^ ~\ ^^ N \ \ . ^ \,\ ? 1 /= " -~~ ^ \ 1 — ^ -^ \}^ 0__ ^ '^ p£ -.a 1 .s Fraction of semispan , X7 (d) M-0.55. .7 .6 1.0 NATIONAL ADVISORY COMMITTEE FOU AERONAUTICS CONFIDENTIAL F/g ure 10.- C onff/J ued . NACA ACR No. L5G10 Fig. 10e,f 0^ o /,0 CONFIDENTIAL .^ --, ^^ C/v= .t> — -^ vx. ,<6 ~~~^ ^^ ^ r, orox S ^ -— . _^ ^ ■ -- . " - (m Shdovt 1 6ta//) 'f 4 ' ^^ "^ \ -^ \ 'i- ^ ^^ \ \ P ^___ .P ^ -_ "->. \ --^ \\ " — . , ^ -- ^s n\ ■ -- ^ y -P — ■ ^^ "^ -.d -3 .5" .6 Fraction of semi s pan , ( ^ u -.a CONFIDENTIAL =^~. ^ ~ ~ — — ^'S'- 7fipf 'OX. a°c ido)/ e SI W/) .o^ -^ ^ ~ - ^ ^ ^ ■" ~ --. ^ ' ■ ~- ~~^ ^- "^ •^ ^ ,d. ~~- s V " ■ — — . . "" ^ -^^ \\ o __ ■ — ^ ^ 7^ — ' —L- ./ :f .S .6 .7 u Fr act ion of sem/span , -?- {cj) M=0.6S. W «5 .<^ _ •K ? ^ Q^ ^ 1 — -^_ ?:; .£ '^^^ ^- ^ .2 — ■ — -- ^^ ■-- ^^ ^ 0_ ■ ■ — ■ ^ ~~i -y -^ — -.2 J .2 .5 ff .6 Fracf/on of semi 5 pan , (h) M=0.7. CONFIDENTIAL /.O' NATIONAL ADVISORY COHHITTEE FOB AERONAUTICS Fi(]ure /0.~ Concluded. NACA ACR No. L5G10 Fig. 11 ^ 5- '^D ' j.L/v/y/jjQoo pt?oj > \~ Qi o% *0 1- .s < -^ s- Z lu v^ O}^ ^ n 8 1 >5 ■< ^ ^1^ ^. •h. •s :i H^ s^ ^ s s ^ s^ ^ t ^ \' \ Q) V. ^ •S^ k. NACA ACR No. L5G10 Fig. 12 ^ c — 5 a y: •>• ^^^ \ ^ / /y // >j > \- a = IS Z UJ o ^ n o / ^ / '^ / / / t~J y / • / 1 '^. 1 ^ 1 1 1 1 1 / > \ fT ^\ h"i _l < 1 1 \ 1 CM z llJ 1 II ■> Q Z i <^ 1 O u 1 1 « 1 1 ^ '^ >0 n-- CVi ^^D '4.US!DU:^ 900 pool ^ cS :? ^S -K ^ c^ -Ci II ^ ^ i? -K ^ Q ta ^ V3 V: Q ^ ■^ >l J^ I) ^ 1^ o 1 ^ ^ > ^ ^ § \ V J5t a) Q) ■K o ^ V) Ij ^