bli\Cf{L'l(<^ </ 
 
 ACF. No. L^31a 
 
 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
 
 WARTIME REPORT 
 
 ORIGINALLY ISSUED 
 
 January 19^6 as 
 Advance Confidential Report L5G31a 
 
 EFFECT OF CC^ffKESSIBILnY ON TEE PEESSORES AKD FORCES 
 ACTIHG on A MODIFIED NACA 65,3-019 AIRFOIL 
 HAVUJG A 0.20-CHORD FLAP 
 By W. F. Lindsey 
 
 Langley Memorial Aeronautical Laboratory 
 Langley Field, Ta. 
 
 nxcx 
 
 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. 
 
 I. - 76 DOCUMFKnrS DEPARTMEHJ 
 
Digitized by tlie Internet Arcliive 
 
 in 2011 witli funding from 
 
 University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation 
 
 http://www.archive.org/details/effectofcompressOOIa 
 
ITACA ACR No. L5G31a ?^- 
 
 NATIONAL ADVISORY COI.MITTEE FOR AERONAUTICS 
 
 ADVANCE CONFIDENTIAL REPORT 
 
 EFFECT OF COMPRESSIBILITY ON THE PRESSURES AND FORCES 
 ACTING ON A ?;^ODIFIED NACA 65,^-019 i^IRFOIL 
 HAVING A 0.20-CKORD FLAP 
 By W, F. Li nosey 
 
 SUMMARY 
 
 An Investigation has been conducted in the Langley 
 rectangular high-speed tunnel to determine the effect of 
 compressibility on the pressure distribution for a modi- 
 fied NACA 65,5-019 airfoil having a 0.20-chord flap. The 
 investigation vvas made for an angle-of-attack range 
 extending from -2° to 12^ at flap deflections from 0^ 
 to -12^. Test data were obtained for Mach numbers 
 from 0.28 to approximately 0.^14.. 
 
 The results shov; that the effectiveness of the 
 tralling-edge-type control surface rapidly decreased and 
 approached zero as the Mach number increased above the 
 critical value. 
 
 INTRODUCTION 
 
 The available information on the aerodynamic charac- 
 teristics of airfoils, v/ith and without flaps, at low 
 speeds Is quite extensive and previous investigations 
 have shown the general effects of compressibility on air- 
 foils without flaps. For airfoils with flaps, however, 
 the available information at high speeds is United. 
 
 The earlier investigations at high speeds demon- 
 strated the limitations of the theoretical methods in 
 extrapolating low- speed data to high speeds (reference 1) 
 In addition, the investigations illustrated the Inappli- 
 cability of the theoretical methods in the supercritical- 
 speed range, in which pressures and forces change radi- 
 cally. These radical changes in pressures and forces are 
 the adverse effects of compressibility, a knowledge of 
 
2 CONFIDENTIAL NACA aCR No. L5G31a 
 
 which Is necessary in the design of high-speed airplanes. 
 Because of the inadequacy of the theoretical method in 
 applications to the supercritical region, recourse to 
 exoeriirent is necessary to detenriine the adverse effects 
 of compressibilitjr. 
 
 Data at high speeds were required in connection with 
 the design of a specific airplane; accordingly, an inves- 
 tigation was conducted in the Langley rectangular high- 
 speed tunnel to determine the effect of compressibility 
 on the r:ii- assures and forces acting on a modified NACA 65,5-019 
 airfoil having a 0.20-chord flap. Pressure-distribution 
 measurements were made at Mach numbers between 0.26 
 and ."{k- for angles of attack from -2° to 12^ cind flan 
 deflections from 0^ to -12°. 
 
 APFaRATUS AiVTD TESTS 
 
 The tests were conducted in the Langley rectangular 
 high-speed tunnel, which is an induction- type tunnel 
 without return passages and has an l8-lnch by i^-inch test 
 section. The variation in the Mach number in the test 
 section along the tunnel axis without a model installed 
 in the tunnel is tO.ii. percent of the stream Mach number. 
 In a plane norm.al to the tunnel axis the variitinn is 
 to. 8 percent of the test-section Mach number at a stream 
 Mach number of 0.60. The direction of the air flow 
 appears to be misalined by -O.l'"'' with a possible variation 
 of ±0 . 1*^ . No correction for misalinem.ent has been made 
 for the data presented herein. 
 
 The model completely spanned the test section along 
 the k-inch dimension and was supported by large circular 
 end plates which v;ere fitted into the tunnel walls in 
 such a way as to rotate with the model and to retain 
 continuity of the surface of the tunnel walls. The 
 juncture between the model and the end plate .vas sealed. 
 
 The profile of the 5-' i^ch-chord model having a 
 0.20-chord flap differed from the modified NACA 65,3-019 
 airfoil section in that the profile from approximately 
 the Rl-percent station to the trailing edge was fonaed 
 by straight lines having an included angle between 20° 
 and 21'^. Porty pressure orifices were installed in the 
 
 model surface in two chordwise rows f- inch from and on 
 
 CONPIDSNTIAL 
 
NACA ACR No. L5G51a CONFIDENTIAL 
 
 either side of the model center line. The model profile 
 and pressure-orifice locations are shown in figure 1. 
 The airfoil ordinates are given in table I. 
 
 Pressure-distribution measurements were made for a 
 range of Mach numbers from 0.28 to approximately O.yli. at 
 angles of attack from -2° to 12° and flap deflections 
 from 0° to -12*^ (up). Additional pressure-distribution 
 measurements were made for positive (down) deflections of 
 the flap at angles of attack of 2° and 1].°. These tests 
 were supplemented by schlieren photographs of the flow in 
 the supercritical speed region for a few of the lower 
 angle-of -attack configurations. These photographs show 
 density gradients in the flow by changes in light 
 intensity. (For details, see reference 1.) 
 
 S'i^ilBOLS 
 
 M Mach number 
 
 M^p Mach number at which sonic velocity was obtained 
 locally within the flov; field (as at the model 
 surface ) 
 
 q dynamic pressure 
 
 Pq free-stream static pressure 
 
 p local static pressure (as at model surface) 
 
 a angle of attack 
 
 P pressure coefficient 
 
 / 
 
 Vr.yy crltlcal pressure coefficient, corresoonding to 
 
 /0.528E - Po\ 
 local Mach nmnber of 1.0 ( -j 
 
 K total pressure 
 
 P-g upper-balance-chamber pressure coefficient 
 
 Pjj lov/er-balance- chamber pressure coefficient 
 
 On section normal-force coefficient 
 
 COl^IDENTIAL 
 
k COflFIDENTIAL NACA ACR No. L5G51a 
 
 -m 
 
 ^/i section pitching -moment coefficient of normal 
 force about quarter-chord location 
 
 Cn-^ flap normal-force coefficient 
 
 "x 
 
 Ch flap hinge-monent coefficient (determined by the 
 pressure distribution from flap hinge axis to 
 trailing edge) 
 
 5 flap deflection measured with reference to the 
 model chore?; negative deflection is up 
 
 RESULTS Airo DISCUSSION 
 Tunnel-lVall Effects 
 
 The results of these tests have not been corrected 
 for constriction or tunnel-wall effects. The most 
 important constriction effect on these data in the super- 
 critical region is the change in the luach numbers given 
 herein to higher effective stream Mach numbers. (See 
 reference 2.) The difference between the tv;o Mach numbers 
 increases rapidly as the maximum tunnel speed is approached. 
 It is further shown in reference 2 that very near or at 
 the maximum speed attainable for a given model-tunnel 
 configuration large gradients in velocity occurred at the 
 v/alls. The maxinura- speed test points given herein for 
 each angular configuration are therefore considered to be 
 of questionable value. 
 
 Pressure Distribution 
 
 The pressure distributions along the chord of the 
 model are presented in figures 2 to Jl, inclusive. Each 
 figure shows the effect of compressibility on the pressure 
 distribution for a given angular configuration. The 
 effect of angle of attack and flap deflection can be 
 obtained from a comparison of the various figures. 
 
 Subcritlc^l region .- The figures for the lower angles 
 of attack and small normal-force coefficients show that 
 increases in Mach number in the subcrltical ranges are 
 accompanied by increases in the maximum negative pressure 
 coefficient, which is in agreement with theory. In the 
 high angle-of-attack and large normal-force-coefficient 
 
 CONFIDENTIAL 
 
NACA ACR No. L5G31a CONFIDENTIAL 5 
 
 range the change in the majcimum negative pressure coef- 
 ficient with Mach number is approximately zero and not in 
 agreement with theory, probably as a result of the exist- 
 ence of separated flow. 
 
 A comparison of part (a) of figures 2 to 51 fo^ S- 
 constant angle of attack and various flap deflections 
 shows that the increment in load produced by a deflection 
 of the flap is distributed approximately uniformly along 
 the chord as could be expected from low-speed tests. It 
 can be seen, hov/ever, at this Mach number ( approx . O.iiJ) 
 that the max;lm^im relative change in loading on the main 
 part of the wing occurs near the leading edge. This 
 change has an appreciable effect in increasing or relieving 
 the pressure peaks that occur near the leading edge for 
 some angle-of -attack conditions. (See part (a) of figs. 2 
 to 5 and 18 to 21 . ) 
 
 Supercri tic al region .- Although it has been shown 
 that in tne subcritical Mach nuraber range the action of 
 the flap in changing the loading along the chord was 
 similar to that shown by the low-speed tests, in the 
 supercritical range, when the region of supersonic flow 
 is relatively large, the loading in and ahead of the 
 supersonic region is a function only of angle of attack. 
 The extent of flow affected by the flap is limited to the 
 region of subsonic flow behind the supersonic region and 
 to the flow over the flap Itself. (See parts (d) and (e) 
 of figs. 6 to 9-) The chordv;ise influence of flap deflec- 
 tion on the flov; over the main part of the v;ing could be 
 expected to be limited to the subsonic flow region ahead 
 of the flap, since pressures are propagated at the speed 
 of sound. This effect of compressibility in producing a 
 marked change in the flov/ over the forward portion of the 
 airfoil at supercritical Mach numbers is comparable to 
 that which occurs for cambered airfoils, as evidenced by 
 the change In the angle of zero lift. The similarity can 
 be more clearly seen when the cambered airfoil is con- 
 sidered to be a multiple flapped airfoil. 
 
 Comparison of the pressure distributions for the 
 flap deflected and neutral shows that, for this model, 
 the chordwlse extent of the influence of the flap 
 decreases gradually as the speed is increased above the 
 critical speed. 
 
 If the fundam.ental aspects of the flow are con- 
 sidered, deflections of the flap could produce changes 
 
 CONFIDENTIAL 
 
CONFIDENTIAL NACA kCR No. L5G51a 
 
 in the pressures in the subsonic flov; region, thereby- 
 changing the shock location. This effect, however, is 
 not apparent in these results because of seoarated flow 
 and the resulting absence of the usual discontinuities in 
 the pressure-distribution diagrams which indicate the 
 shock location. 
 
 Schlieren Photographs 
 
 Schlieren photographs of the flow in the higher Mach 
 number range are shown in figures 52 to 57 > inclusive, 
 for the model at angles of attack of 0^ and IxP with 
 various flap deflections. (Note compressions are white 
 on fig. 52, black on figs. 55 to 57') These photograohs 
 show that for angles of attack from O'^ to I|.° "the flow 
 separates at approximately the 0.60-chord location for 
 Mach numbers near the critical values, and with increasing 
 Mach number the separation point laoves forward. The 
 forward movement of the separation point is accompanied 
 by a rearward movement of the shock along the separation 
 boundary. 
 
 A comparison of the schlieren photographs with the 
 corresponding pressure-distribution diagrams shows that 
 the existence of separated flow has a serious effect on 
 the pressure distribution in the vicinity of compression 
 shocks. (See, for exajiiple, figs. 9(d) and ^^(d).) It 
 will be noted that for the condition of a well-defined 
 shock and separated flow the pressure-distribution diagram 
 indicates smooth compression. In addition, the location 
 of the pressure corresponding to the critical pressure 
 coefficient generally occurs from 5 to 10 percent of the 
 chord downstream from the siiock location. This phenomenon 
 is probably the result of the existence of large static 
 pressure gradients in the separated-f low region between 
 the boundary and the model surface where pressures were 
 measured. The extent of the separated flow would be 
 reduced for an airfoil having a smaller thickness-to- 
 chord ratio . 
 
 Critical Mach Number 
 
 The variation of the critical Mach number of each 
 surface v/i th flap deflection for constant angles of attack 
 is presented in figure 56. The large decrease in critical 
 Mach number that occurs at angles of attack between 6*^ 
 
 CONFIDENTIAL 
 
NACA ACR No. L5a31a COITPIDBITTIAL 7 
 
 and 8° for the upper surface is a result of a rapid 
 increase in the magnitude of the negative pressure coef- 
 ficients near the leading edge. The difference hetween 
 the critical Mach numbers for the upper and lower surfaces 
 for corresponding conditions can be attributed to air- 
 flow misalinenent . The highest critical Mach number for 
 this airfoil is approximately O.65 j^nd is obtained for 
 angular configurations corresponding to a normal force of 
 approximately zero. 
 
 ComDressibility Effects on Force and Moment Coefficients 
 
 and 
 integrations of 
 
 The normal-force, moment, flap normal-force, 
 hinge-moment coefficients obtained from integration.^ ^^ 
 pressure-distribution diagrams are presented in figures 39 
 and 'I4.O . Each figure shows the variation in the coeffi- 
 cient with Mach number at a constant angle of attack for 
 each flap deflection. These figures are cross-plotted in 
 figure kl to shov/ the variation in the coefficients at a 
 constant Mach number. 
 
 Normal-force coefficients . -Figure 59 shows that, with 
 the flao at 0", the effect of compressibility on the 
 normal-force coefficient at subcritical l.iach numbers is 
 in accord with previous experimental and theoretical 
 results. The variation for the other flap deflections at 
 a fixed angle of attack, however, appears to follov\f, to 
 some extent, both in direction and magnitude the variation 
 for the condition of the flap at 0"-', as indicated by a 
 lesser divergence of the curves than would have been 
 expected from theoretical estimations. (See parts (c), 
 (d), and (e), fig. 39-) A variation of this type shows 
 
 dc^ 
 tnat the effect of comoressibility on — — is small in 
 
 dS 
 the subcritical Mach number range. 
 
 In the supercritical Mach number range above the 
 value at which the peak normal-force coefficient occurs, 
 the convergence of the curves for the various flap 
 deflections at a given angle of attack indicates a rapidly 
 
 dc>^ 
 decreasing — ^ with increasing Mach number. The effect 
 
 H ft 
 
 of com-oressibility on — —, presented in figure 42, is 
 
 d5 
 in accord with the variations indicated in figure 39' 
 
 CONFIDENTIAL 
 
CONFIDEOTIAL NACA ACR IIo . L5G31a 
 
 In the supercritical f'ach range for angles of attack 
 not greater than 6'-' it can he seen in figure 59 that the 
 peak normal-force coefficient for a given angular configu- 
 ration occurred at a Mach nu^nber of approxiraately O.65, 
 which indicates that this airfoil section should be 
 restricted to designs v.herein the rr.aximum Mach number 
 does not exceed O.65. 
 
 A comparison of the various parts of figure kl in 
 which is given the variation of the norr.al-f orce coeffi- 
 cient with angle of attack shov/s that as the Ma.ch number 
 dCv, 
 
 Increases increases and reaches a majcimurri value at 
 
 da 
 a Mach number of O.65 as could have been expected from 
 the preceding discussion. 
 
 A further examination of figure kl shows that 
 
 deflections of tlie flap produce a change in the angle of 
 
 dCv, 
 zero normal force, but have no appreciable effect on — ==• 
 
 da 
 
 for the more linear parts of the curves. The effect of 
 
 dcp 
 compressibility on — — for the more linear parts of 
 
 da 
 these curves is presented in figure J4.2. 
 
 The flap effectiveness — , obtained from the ratio 
 
 dc dc 
 of — — to -k^ (fiG. H-2) and presented in figure L.5 , 
 
 d 5 da 
 decreased with increased Idach ni.u;iber and rapidly approached 
 zero as the Mach number was increased above the critical 
 value . 
 
 Moment coefficients .- The variation of the moment 
 coefficient with ?/:ach number in the subcritical range as 
 shown in figure 39 i-3 small. In the supercriv^ical region 
 the moment coefficients generally increase witl'i increasing 
 Mach munber. This increase is followed by a rapid 
 decrease, which occurs at a Mach number above the value 
 at which the decrease in normal-force coefficient occurs. 
 The pressure-distribution diagrams show that the decrease 
 in noniial-f oi'ce results from a general decrease in the 
 magnitude of the loading, and at higher Mach numbers the 
 continued decrease in loading is accompanied by a change 
 in distribution whJ.ch results in a decreased moment coef- 
 ficient. 
 
 COMFIDENTIAL 
 
NACA ACR No. L5G51a CONFIDENTIAL 
 
 At the supercritical Mach numbers at which the normal 
 forces lor the various flap deflections tend to he approxi- 
 mately the same, the direction of the change in the moment 
 coefficients is similar to the direction of the change 
 in Cn- A variation of similar magnitude, however, could 
 not be expected in these coefficients because of the large 
 effect that a small change in load at the rear of the model 
 has on the moment coefficient. 
 
 A comparison of the effects of angle of attack and 
 flap deflections on the variation in m.oment coefficient 
 with normal-force coefficient can be seen in figure l+l. 
 
 For constant flap deflection there is a small positive 
 increase In moment coefficient v/ith an increase in normal- 
 force coefficient. This slope remains approximately con- 
 stant for Mach numbers to approximately 0.65. For Mach 
 numbers above O.65, the variation depends on the flap 
 deflection. 
 
 At a constant angle of attack between -2° and 6°, 
 large changes in moment coefficient occur in a negative 
 direction with an increase in the nonnal-f orce coeffi- 
 cient. These slopes, which are approximately constant to 
 a Mach nuinber of 0.6, increase as the Mach number is 
 further increased to O.7. 
 
 Flap normal force .- The flap normal-force coeffi- 
 cients are presented in fi^Tare I4.O . In the supercritical 
 region, the variations in the coefficients are large and 
 irregular, probably being influenced by the effect of flow 
 separation. 
 
 Hinge-moment coefficients .- The variation in the 
 hinge-moment coefficients with I.'ach number, also shown in 
 figure J4.O , are very similar to the variation in flap 
 normal force. 
 
 At Mach numbers from 0.01 to 0.02 below the maximum 
 test value, a flap deflection range is indicated in which 
 the flap tends to become or is overbalanced. At the same 
 Mach number and for the same angular configuration, in 
 figure 39 J ^o change occurs in the value of normal-force 
 coefficient, and the control v/ould therefore be unresisting 
 and ineffective. Although the Mach number is near the 
 maximuim test value, for v/hich the data are of questionable 
 value, the possibility of this condition of the control 
 should not be overlooked. 
 
 COWFIDSNTIAL 
 
10 CONFIDENTIAL 2'TACA ACR No. L5G31a 
 
 The hinge-moment coefficients are presented in 
 figure iil in the sane manner as the moment coefficients. 
 It can he seen that for a constant flap deflection the 
 magnitude and direction of change of hinge-raoment coeffi- 
 cient with increase in nori.-ial-f orce coefficient depends 
 on the absolute value of the flap deflection. The slope 
 for a given flap deflection increases positively v.-ith 
 increases in Kach number to 0.675 5 further increases in 
 y.ach number are accompanied by increases in the slope in 
 the negative direction. 
 
 At a "ach number of 0.1x0 the changes in hinge moment 
 with flap deflection at a constant angle of attack are 
 generally uniform over the flap deflection range. At a 
 ?.!ach number of 0.60 and above in the negative normal- 
 force-coefficient range the changes in hinge-moment coef- 
 ficient for flap deflections between 0'^ and -UP is very 
 small compared v/ith the changes at larger deflections. 
 This effect is possibly a result of the reversals in the 
 load over ■che flap produced by the very thick boundary 
 layer or separated flow. (See pressure-distribution 
 diagrams and schlieren photographs.) 
 
 Balance- chamber pressure-coefficient differential 
 ( Pj^ - Pp ) . - The difference oetween the pressure coeffi- 
 cients for the lower-surface and the upper-surface balance 
 chambers is presented in figure i^4. This figure shows 
 the effect of Kach number, flap deflection, and angle of 
 attack on the pressure-coefficient differential. 
 
 The effect of compressibility on the pressure- 
 coefficient differential is generally small at Mach 
 numbers below O.65. Tliis small effect could be expected 
 when the magxiitudes of the individual pressure coeffi- 
 cients are sm.all and are measured at a station in rear of 
 the position of the maximum negative pressu-.-e coefficient. 
 The decrease in the magnitude of the pressure coefficient 
 at the high Mach numbers is primarily a result of the 
 effects of flow separation. 
 
 The balanced hinge-moment coefficient for this model 
 can be obtained by adding the hinge -moment coefficient of 
 figure li.0 to the product of the balance- chamber prcssure- 
 coufficient differential ( Pl - ^) ^""^ ^ constant, the 
 constant depending on the length of the balance tab. A 
 brief comparison of figures 1^.0 and kh indicates that at 
 
 COHFIDEIITIAL 
 
NACA ACR No. L5G51a CONFIDENTIAL 11 
 
 dcv, 
 low speeds the effect of the balance would be to reduce — — 
 
 d5 
 
 as compared with the unbalanced condition. With increasing 
 
 Mach number the effect of the balance decreases for the 
 
 lower angle-of -attack range. 
 
 CONCLUDING R 3,1 ARE 
 
 The results of the investigation on the modified 
 NACA 65,3-019 airfoil having a 0,20-chord flap indicate 
 that the effectiveness of a traillng-edge control surface 
 of small chord rapidly decreases and approaches zero as 
 the Mach number increases above the critical value. 
 
 Langley Memorial Aeronautical Laboratory 
 
 National Advisory Co:rmiittee for Aeronautics 
 Langley Field, Va. 
 
 REFERENCES 
 
 Stack, John, Lindsay, W. P,, and Littell, Robert E.: 
 The Compressibility Burble and the Effect of Com- 
 pressibility on pressures and Forces Acting on an 
 Airfoil. NACA Rep. No. 61^.6, I938 . 
 
 Byrne, Robert W.: Experimental Constriction Effects 
 in High-Speed V.'ind-Tunnels . NACA ACR No. LiiLOya, 
 
NACA ACR No. L5G51a 
 
 CONPIDEITTIAL 
 
 12 
 
 TABLE I.- BASIC SECTION ORDINATES FOR MODIFIED 
 HACA 65,5-019 AIRFOIL SECTION 
 j Stations and ordlnates are in percent chord J 
 
 Station 
 
 — — — - — — — 1 
 
 Ordinate 
 
 Upper 
 
 Lower 1 
 
 
 surface 
 
 surface ! 
 
 
 
 
 
 
 
 .5 
 
 1.108 
 
 -1.108 
 
 1 
 
 1.921 
 
 -1.921 
 
 2 
 
 2.59S 
 
 -2.598 
 
 ^ 
 
 ^.620 
 
 H.I)-37 
 
 -^.620 
 
 -4.457 
 
 8 
 
 5.127 
 
 -5.127 
 
 10 
 
 5.725 
 
 -5-725 
 
 ^ 
 
 6.715 
 
 -6.715 
 
 18 
 
 7.525 
 
 -7.525 
 
 22 
 
 6. 192 
 
 -8 . 192 
 
 26 
 
 8.721 
 
 -8.721 
 
 50 
 
 9.115 
 
 -9.115 
 
 1 
 
 9.371 
 9.II90 
 
 -9.371 
 -9.490 
 
 ¥ 
 
 9.500 
 9.471 
 
 -9.500 
 -9.471 
 
 k6 
 
 9.515 
 
 9.02/1. 
 
 -9.315 
 -9.024 
 
 50 
 
 5^ 
 
 8.597 
 
 -8.597 
 
 5^ 
 
 8.059 
 
 -8.059 
 
 62 
 
 7.570 
 
 -7.570 
 
 66 
 
 6.612 
 
 -6.612 
 
 70 
 
 5.791 
 4.922 
 
 -5.791 
 
 -4-922 
 
 ^^ 
 
 78 
 
 ii..029 
 
 -ii.o029 
 
 82 
 
 8.128 
 
 -8 . 128 
 
 86 
 
 2.2II7 
 
 -2-247 
 
 90 
 9k 
 
 l.Zllo 
 
 .683 
 
 -1.I116 
 -.688 
 
 98 
 
 .158 
 
 -.158 
 
 100 
 
 ° 
 
 
 
 L.E. radius: 
 
 2.159 
 
 
 CONFIDENTIAL 
 
NACA ACR No. L5G31a 
 
 Fig. 1 
 
 O 
 
 to 
 
 > »- 
 a. => 
 
 O 2 
 55 o 
 
 > « 
 
 I' 
 
 < "^ 
 
 Z UJ 
 
 O 
 
 Vj 
 
 < 
 
 z 
 
 UJ 
 
 O 
 U 
 
 (U 
 
 o 
 
 o 
 
 0) 
 
 u 
 
 bo 
 
 X 
 
Fig. 2a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Lower surface 
 
 (a) /V= 0.434. 
 
 (ti M= O. 6/8. 
 
 % 
 o 
 
 i: 
 
 o 
 
 
 
 \ 
 
 vA 
 
 -^. 
 
 / 
 
 
 
 
 '^^^^^-^^'^ft^, 
 
 
 (c) 
 
 M-O.t 
 
 t7S. 
 
 
 (dJ M=0.70J. 
 
 (e) M^0.7£7. 
 
 20 40 60 60 
 
 
 
 
 
 ^. 
 
 
 A 
 
 r^' 
 
 
 
 -^cr 
 
 r 
 
 
 
 
 
 /oo 
 
 (f) M= 0.745. 
 ao 40 60 60 
 
 /OO 
 
 Perccnf chord 
 
 CONFIDENTIAL 
 
 NATIONAL AOVISOQV 
 COMMITTU m UKMUUTICS 
 
 Figure 2.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a - -2°; S = 0°. 
 
NACA ACR No. L5G31a 
 
 Fig. 3a-f 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Lower surface 
 
 (a) M ^0.4 J 4. 
 
 r,^ 
 
 ^,-0-0' 
 
 ^x- 
 
 — — 
 
 ^rr 
 
 ( 
 
 ■^ 
 
 N 
 
 
 ^-^ 
 
 f 
 
 
 
 
 
 {b) M^O.6/4- 
 
 I 
 
 
 / <r 
 
 
 
 
 1 
 
 N 
 
 
 ( 
 
 
 
 \ 
 
 ,^-^®**~. 
 
 
 (c) 
 
 M=Oa 
 
 '57:5. 
 
 
 M) fl^O. 706. 
 
 -/ 
 
 r 
 
 f?r 
 
 (e) M -0.7J6. 
 O so 40 60 60 /CO O 20 
 
 Percenf chore/ 
 
 CONFIDENTIAL 
 
 
 A 
 
 ^ 
 
 \ 
 
 X3^ 
 
 {^ 
 
 r^ 
 
 
 
 ^r 
 
 1/ 
 
 
 
 
 
 [ 
 
 (f) 
 
 M-O. 
 
 746. 
 
 
 4^ 
 
 60 do 
 
 NATrONAL ADVISORY 
 COMMITTEE FOB AERONAUTICS 
 
 /GO 
 
 Figure 3.- Pressure distribution for a modified NACA 65,3-019 
 
 airfoil with 0.20— chord flap. a 
 
 oO . 
 
 S = -4' 
 
Fig. 4a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 -/ 
 
 O 
 
 X Upper surface 
 G Lower surface 
 
 ,too-^ 
 
 
 ^ 
 
 >\ 
 
 ^. 
 
 '-"^ 
 
 "~^ 
 
 f 
 
 
 
 \ 
 
 
 
 (a) 
 
 /•^-O.^J^Z. 
 
 
 (b) /^=0.6/J. 
 
 Ss. 
 
 
 (C) M-0.6 7/. 
 
 icf) n-o.70/. 
 
 (e) M-O. 73^, 
 
 (t) n-O. 74-5. 
 
 ZO ^O 60 GO lOO O ZO 40 60 QO lOO 
 
 Percent chord 
 
 NATIONAL AOVISbQV 
 COMNITTti rn AlKMAUTKi 
 
 CONnoCNTIAL 
 
 Figure 4.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = -2°; h - -8°. 
 
NACA ACR No. L5G31a 
 
 Fig. 5a-f 
 
 CONFIDENTIAL 
 
 X Upper sur/ace 
 O Lower surface 
 
 (a) n^0.4J6. 
 
 r^ 
 
 
 \ 
 
 --P.r 
 
 ^ 
 
 
 - K 
 
 
 r 
 
 
 
 
 ^ 
 
 li 
 
 lb) 
 
 /7--a 6/0. 
 
 
 
 o 
 
 0; 
 •n 
 
 Ic) M^ 0.674. 
 
 id) n^o.yoo. 
 
 
 r^. 
 
 ^ 
 
 
 ■"-o-o-^j 
 
 
 r^jy 
 
 
 N 
 
 -J 
 
 r 
 
 
 
 
 
 leTn^aTJ/. * (f) n^o.7^5. 
 
 £0 40 60 80 /OO 20 40 60 SO 
 
 '^r 
 
 /oo 
 
 Percent chotd 
 
 CONFIDENTIAL 
 
 NATIONAL ADVrSORY 
 COMMITTEE FOB AERONAUTICS. 
 
 Figure 5.- Pressure distribution fur a modified NACA 65,3-019 
 airfoil with 0.20-chord flap, a = -2°; S = -12°. 
 
Fig. 6a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIOENTJAL 
 
 X Upper surface 
 O Loiver surface 
 
 (a) n=0.442. 
 
 (D) n -0.646. 
 
 .<b -/ 
 .o 
 
 o 
 
 % 
 
 
 .-ts 
 
 >~^ 
 
 
 
 v 
 
 r^ 
 
 '\ 
 
 ^ 
 
 - ^r 
 
 ^ 
 
 
 
 ^ 
 
 — ^ 
 
 i 
 
 (c) 
 
 n=o.t 
 
 "JZ. 
 
 
 
 u^ 
 
 ^^ 
 
 
 
 ,/ 
 
 K 
 
 \ 
 
 ^ 
 
 - ^r 
 
 r 
 
 
 
 ^ 
 
 
 id] 
 
 n-0.706. 
 
 
 -2 
 
 
 
 (e) n =0.716. 
 
 (f) M -0.143. 
 
 O 20 40 60 80 /GO 20 40 60 60 /OO 
 
 fierce nf chord 
 
 CONFIDENTIAL 
 
 Figure 6.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 0°; 5 = 0°. 
 
NACA ACR No. L5G31a 
 
 Fig. 7a-f 
 
 CONFIDENTIAL 
 
 X Upper surface 
 G Lower surface 
 
 (a) A7= 0.455. 
 
 it) M -0.650. 
 
 a 
 
 .0) 
 
 O 
 
 Q) 
 
 (C) /^= 0.677. 
 
 / 
 
 y 
 
 % 
 
 s 
 
 N 
 
 -f^cr 
 
 f 
 
 
 
 \ 
 
 
 1 
 
 id) 
 
 /V- 0.709. 
 
 
 -I 
 
 o 
 
 
 
 
 ^r 
 
 (Q) /^= 0.753. 
 
 £0 40 60 60 
 
 /GO 
 
 O 
 
 if) n^O.746. 
 ZO 40 60 SO 
 
 /-^Qrc enf cAorc/ 
 
 CONFIDENTIAL 
 
 NATIONAL ADVISORY 
 COMMITTEE FM AERONAUTICS 
 
 /OO 
 
 Figure 7.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with .20-chord flap. a = 0°; 8 = -4°. 
 
Fig. 8a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Lower surface 
 
 
 
 I ^ 
 
 I 
 
 (QJ M^0.4D7. 
 
 i£>J M-0.6^Z. 
 
 (W f7=0.708. 
 
 -/ 
 O 
 
 / 
 
 ^4P 
 
 (9) M =0.7 J 6. 
 O 20 40 60 SO /OO 
 
 O'J M -0.744. 
 ZO 40 60 do /OO 
 
 Perc enf ch or d 
 
 NATIONAL A0VI5ORV 
 COMMITTU rot AiKMAUTKS 
 
 CONFIDENTIAL 
 
 Figure 8.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap, a - 0°; S - -8°. 
 
NACA ACR No. L5G3la 
 
 Fig. 9a-f 
 
 CONFIDENTIAL 
 
 X Upper surface 
 Q Loi\'er surface 
 
 (a) M-0.43 /. 
 
 
 y^ 
 
 \ 
 
 
 ' — 1 
 
 o 
 
 
 'ij 
 
 
 cr 
 
 ( 
 
 
 
 \ 
 
 
 4 
 
 (i^\ 
 
 /^ - rs t 
 
 ^^ Q 
 
 
 a 
 
 
 s 
 
 -I 
 
 (c; fi^o.c76 
 
 (Of J fl^O.lOJ. 
 
 -i 
 
 o- 
 
 
 ^ 
 
 Y -^, 
 
 
 
 A 
 
 y 
 
 s 
 
 ^-^ 
 
 H 
 
 ^ 
 
 
 
 
 
 \, 
 
 (e) 
 
 M -O. 
 
 734. 
 
 
 
 (f) n^O.743. 
 
 f?r 
 
 ^o 40 e>o 60 /oo o zo 
 
 Perc ent chores 
 
 CONFIDENTIAL 
 
 40 to 80 
 
 NATIONAL ADVISORY 
 COMMITTEE FOB AERONAUTICS. 
 
 /OO 
 
 Figure 9,- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap . a = 0°; S = -12°. 
 
Fig. lOa-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Loivpr surface 
 
 (a) M^ 0.^-56. 
 
 (t) M=0 643. 
 
 c 
 
 O 
 
 ^ o 
 
 "-0 
 
 ^ 
 ^ 
 
 
 ^ 
 
 \ 
 
 
 
 f? 
 
 
 
 ^ 
 
 -^r 
 
 i 
 
 
 
 
 ^^"^^ 
 
 
 (C) 
 
 n^ac 
 
 5/(5. 
 
 
 
 y^ 
 
 --^ 
 
 
 
 ^ 
 
 -y 
 
 1 
 
 k 
 
 -^cr 
 
 r 
 
 /_/! 
 
 
 
 pC>-«fc=^ 
 
 /■<pj M--0.725. * (f) t^^ 0.744 
 
 O ZO 40 60 QO lOO O 20 40 60 80 
 
 f^QrcQnf cf)oro/ 
 
 /OO 
 
 NATIONAL AOVISOnv 
 COMMITTII rOI UKWAUTK& 
 
 CONFIDENTIAL 
 
 Figure 10.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a - 2°; S - 0°. 
 
NACA ACR No. L5G31a 
 
 Fig. lla-f 
 
 CONFIDENTIAL 
 
 -/ 
 
 o 
 
 / ®- 
 
 X Upper surface 
 O Loiver surface 
 
 (F 
 
 M 
 
 -^ 
 
 ^ 
 
 
 f 
 
 
 
 \ 
 
 r^ 
 
 (a) M=0.43/ 
 
 m M- 0.647 
 
 a 
 
 (d) M =0.7/0 
 
 -2 
 
 -/ 
 
 n 
 
 
 .^-^ 
 
 ^ ^^^T"^ 
 
 ^ 
 
 
 A 
 
 
 
 ^^^. 
 
 ^-=*^ 
 
 
 
 
 
 
 cr 
 
 (e) M = 0.75/ 
 
 
 ^ 
 
 
 l^ 
 
 
 //" 
 
 ^ 
 
 
 
 
 
 r 
 
 
 
 
 
 
 if) 
 
 M-O.-y 
 
 '42 
 
 
 /?. 
 
 O ZO 40 60 60 /DO O 20 40 60 60 100 
 
 f-^6 re 6n f chord national »dvisoov 
 
 OHHITTEE FOfi AERONAUTICS 
 CONFIDENTIAL 
 
 Figure 11.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 2°; S = -4°. 
 
Fig. 12a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 X Upper .'iurfcjce 
 O Loiver surface 
 
 \7^^*^' 
 
 (a) n=o.4ZZ. 
 
 (t) M- 0.644 
 
 \ 
 
 
 
 
 ■^ 
 
 /* 
 
 ^ 
 
 ^1 
 
 k-1 
 
 
 f 
 
 
 
 \ 
 
 ■^^i 
 
 ^ 
 
 fc; n=o.675. 
 
 (d) M= 0.707. 
 
 tw.- 
 
 ^$r^ 
 
 -^^ 
 
 ^. 
 
 ^ej M= 0.754. 
 20 40 60 SO fOO 
 
 (f) n=0.74/. 
 20 40 60 60 
 
 f-^ercen^ chord n»tionai «ovisooy 
 
 COmiTTII r» UKMiuTICS 
 
 CONFIDENTIAL 
 
 /GO 
 
 Figure 12.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 2° ; 8 = -8°. 
 
NACA ACR No. L5G31a 
 
 Fig. 13a-f 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Loive>r surface 
 
 (a) /^ =0.433. 
 
 
 ^ 
 
 ^^^^N 
 
 ^ 
 
 
 
 
 
 N 
 
 .^ 
 
 
 M 
 
 n=0.64.8. 
 
 
 (1) 
 Q 
 
 o 
 
 
 (c) >i^0.677. 
 
 Id) fi=0.707. 
 
 -2 
 
 O 10 do 60 eo /oo 
 
 (f) M =0.74 2.. 
 10 40 60 aO 100 
 
 Fierce n^ chord i«»noN*L *ovrsoRY 
 
 COMMITTEE FW *£IK)»I4UTICS 
 
 CONFIDENTIAL 
 
 Figure 13.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 2°; S = -12°. 
 
Fig. 14a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Low£>r surface 
 
 (a) M = 0.^33. 
 
 ^,_^ 
 
 
 "^<r~ 
 
 - f?r 
 
 
 
 \^ 
 
 '"l 
 
 N 
 
 
 Y 
 
 
 
 \ 
 
 ^u^ 
 
 
 fb) 
 
 M^O.6/^ . 
 
 
 
 03 
 
 
 
 
 
 . 
 
 "x 
 
 
 
 
 
 
 
 
 
 
 y^ 
 
 — ^ 
 
 k 
 
 ^cr 
 
 (/ 
 
 \ 
 
 N 
 
 s 
 
 
 ( 
 
 
 
 
 
 f 
 
 
 
 
 
 
 (c) 
 
 M=0.(. 
 
 580. 
 
 
 
 
 fd) 
 
 M=0.70^. 
 
 
 o 
 
 
 le)/^=0.729. ' ff)M=0.74Z. 
 
 ZO 40 60 60 /OO iO 40 60 60 
 
 Percent chord 
 
 NATtONAL A0VI5ODY 
 COMMITTIE FM AIKMAUTK& 
 
 100 
 
 CONFIDENTIAL 
 
 Figure 14.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-^chord flap. a = 4°; S = 0°. 
 
NACA ACR No. L5G31a 
 
 Fig. 15a-f 
 
 CONFIDENTIAL 
 
 -/ 
 
 o 
 
 X Upper surface 
 O Loi/i/er surface 
 
 (a) M-0.4Z6. 
 
 (6) M= 0.644. 
 
 a 
 
 C 
 $ 
 
 .0 
 
 0) 
 
 o 
 ^ 
 
 X 
 
 f 
 
 A 
 
 v\ 
 
 -i 
 
 H 
 
 ■^r 
 
 f 
 
 
 
 sj 
 
 r^ 
 
 (C) M=0.675. 
 
 
 .-■^ "i 
 
 ^<r^-^ 
 
 
 c\ 
 
 Y 
 
 \ 
 
 ^ 
 
 ^r 
 
 ( 
 
 
 
 
 
 
 (m 
 
 n= 0.703. 
 
 
 ^ 
 
 -z 
 
 
 -— ^ 
 
 
 1^ 
 
 
 r) 
 
 X 
 
 ^ 
 
 
 ^^^^^-^ 
 
 • j^ 
 
 
 
 
 
 F^r 
 
 ' "* (e)M=0.72S. ^ ff) A1 = 0.739. 
 
 O 20 40 60 ao lOO O 20 40 60 80 100 
 
 Percent chord nation»l advisory 
 
 COMMITTEE FM AERONAUTICS 
 CONFIDENTIAL 
 
 Figure 15.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 4°; S = -4°. 
 
Fig. 16a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 X Upper sur/ace 
 O Loiver surface 
 
 (a) n^ 0.43 7. 
 
 (? 
 
 ^ 
 
 ^ 
 
 r^'r- 
 
 ^^ 
 
 f 
 
 
 
 \ 
 
 
 (d) M =0.646. 
 
 c 
 
 o 
 
 Q) 
 o 
 o 
 
 S 
 
 / 
 
 
 ^^^ "^ V 
 
 
 
 
 M 
 
 
 ^. 
 
 ^. 
 
 
 
 \ .^ 
 
 
 
 "^Ti 
 
 k^ 
 
 / 
 
 
 
 \ 
 
 
 / 1 
 
 \ 
 
 
 
 
 
 (C) r-f=0.675. 
 
 
 ^ — 7 
 
 ;><"~^. 
 
 
 (? 
 
 X 
 
 \ 
 
 H 
 
 
 f 
 
 
 
 
 
 
 
 {d] n^O.706. 
 
 (e) n-0 75S. '^ W) TT^074r. 
 
 O ^0 40 60 so too ZO 40 60 80 100 
 
 Perce nf chord nation»i «Dvi$oiiy 
 
 COMNITIU rot UIONMJTKS 
 
 CONFIDENTIAL 
 
 Figure 16. - Pressure distribution for a modified NACA 65,3-01! 
 airfoil with C.20-chord flap. a = 4° ; S « -8°. 
 
NACA ACR No. L5G31a 
 
 Fig. 17a-f 
 
 CONFIDENTIAL 
 
 -/ 
 
 
 
 X Upper surface 
 O Lower surface 
 
 {a) f7^ 0.455. 
 
 (b) n^0650 
 
 (A 
 
 y 
 
 \ 
 
 \ 
 
 
 f 
 
 
 
 N 
 
 
 
 id) 
 
 n=o.7 
 
 06. 
 
 
 
 
 (e) M^0.7J5. 
 10 40 60 dD 
 
 
 
 -^ 
 
 -^ 
 
 
 (A 
 
 \ 
 
 
 
 
 f 
 
 
 
 
 
 100 
 
 if) M=0.740. 
 dO 40 60 SO 
 
 P'ercenf chord 
 
 CONFIDENTIAL 
 
 NATIONAL ADVISORY 
 COMMITTEE FM AERONAUTICS 
 
 ^cr 
 
 100 
 
 Figure 17. - Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 4°; S = -12°. 
 
Fig. 18a-f 
 
 NACA ACR No. L5G51a 
 
 K^ 
 
 n 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Lower surface 
 
 (a) A7= 0.^55. 
 
 *i?«>^-=*q 
 
 (b) ri^ 0.566 
 
 Q> 
 
 O 
 O 
 
 
 -I 
 
 
 
 y 
 
 ^ 
 
 1 — ^"^ — 1 
 
 h^ 
 
 ^r 
 
 / 
 
 
 
 
 
 
 (0 
 
 n^ot 
 
 57^. 
 
 
 id) n= 703. 
 
 (n>i= 0.73a. 
 
 (e) r/= 0.724. 
 
 CONFIDENTIAL 
 
 O 20 40 60 80 KX) O £0 40 60 80 100 
 
 Percent chord 
 
 Figure 18.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with O.PO-chord flap. a = 6°; 3 = 0°. 
 
NACA ACR No. L5G31a 
 
 Fig. 19a-f 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Loi^er surface 
 
 (O) M^O.44/. 
 
 ib) f1=Q6/S>. 
 
 a 
 
 
 o 
 
 u 
 
 
 -/ I 
 
 (c) M =0.6 16. 
 
 (d) M =0.703 
 
 O 
 
 IdFhOJJI. 
 20 40 60 60 
 
 /OO 
 
 If J M = 0.736. 
 20 40 60 80 /OO 
 
 f^&rc£-n^ chord 
 
 CONFIDENTIAL 
 
 NATIONAL ADVISORY 
 COMMITTEE FOfl AERONiUTICS 
 
 Figure 19.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 6°; 8 = -4°. 
 
Fig. 20a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Lower surface 
 
 v_^ 
 
 
 
 
 
 
 j^ — '■ ■ 
 
 \ 
 
 -^r 
 
 
 ^^ 
 
 ^ 
 
 M 
 
 ^ 
 
 ^:^^>^ 
 
 i 
 
 y 
 
 ^ 
 
 ^ 
 
 M 
 
 ^^:^ 
 
 { 
 
 
 
 
 — ^ 
 
 r 
 
 
 
 s 
 
 ^-^^ 
 
 
 (a) 
 
 M=0.440. 
 
 
 
 
 (b) 
 
 M=0.625. 
 
 
 8 ^ 
 
 0) 
 
 
 V\ 
 
 ^ 
 
 y 
 
 
 ( 
 
 
 
 
 
 
 (C) 
 
 A;-<?.d 
 
 )82. 
 
 
 roV M^O.7/4 
 
 (e) M-0J34. 
 
 (f) M-- 0.136. 
 
 
 
 20 40 60 
 
 GO 100 20 
 Percent chord 
 
 40 
 
 60 dO 
 
 NATIONAL ADVISQQV 
 COMNITTII rot AIKMIUTICS 
 
 CONF(DENTI.*L 
 
 /oo 
 
 Figure 20.- Pressure distribution for a modified NACA 65,3-019 
 
 airfoil with 0.20-chord flap. 
 
 -8^ 
 
NACA ACR No. L5G31a 
 
 Fig. 21a-f 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Lower surface 
 
 (a) r^= 0.435. 
 
 (b) n-0.6SI 
 
 
 ^ 
 ^ 
 
 M 
 
 r^ 
 
 ^ 
 
 
 n 
 
 
 
 \ 
 
 
 u 
 
 (c) 
 
 n^o.L 
 
 574. 
 
 
 (d) M^ 0.704. 
 
 a 
 
 O 
 
 (e) M=0.727. " (f)M = 0.734. 
 
 £0 40 60 aO 100 O 20 40 60 80 
 
 Percent chord 
 
 100 
 
 NATrONM ADVISORY 
 COHHITTEE FOB AERONAUTICS 
 
 CONFIDENTIAL 
 
 Figure 21.- Pressure distribution for a modified NACA 65,3-019 
 
 airfoil with 0.20-chord flap. a = 6 
 
 -12^ 
 
Fig. 22a-f 
 
 NACA ACR No. L5G31a 
 
 X Upper surface 
 O .Lower surface 
 
 CONFIDENTIAL 
 
 {a) /^=0.4.36. 
 
 r-.. 
 
 \ 
 
 
 
 
 
 \ 
 
 - ^cr 
 
 
 1 
 
 ,'^ 
 
 
 "^ 
 
 ~ 
 
 / 
 
 Y 
 
 
 
 
 V 
 
 (b) /^=OSS<5. 
 
 
 ■C -Z 
 c: 
 
 .0) 
 
 I-' 
 
 Q) 
 O 
 
 ^ o 
 
 
 ; -- 
 
 \ 
 
 
 - Per 
 
 M 
 
 
 
 ^ 
 
 
 p^ 
 
 V 
 
 
 
 
 
 b" 
 
 ^c; 
 
 /v=c.^ 
 
 j'i'/. 
 
 
 r-r^" ^ 
 
 
 
 
 
 
 H 
 
 
 H 
 
 '^^ 
 
 1 ^^ 
 
 
 
 
 r 
 
 ^ 
 
 ro'; 
 
 M~0.674. 
 
 
 -/ 
 
 ^ 
 
 ^ — 
 
 V 
 
 ji — , 
 
 
 
 > 
 
 H 
 
 /-- 
 
 \^ 
 
 
 V 
 
 
 
 
 \ 
 
 p 
 
 reJ 
 
 M=0.? 
 
 '05. 
 
 
 ■Per 
 
 (f) /^=0720. 
 
 ^ 40 60 60 /OO 
 
 20 40 GO 80 /OO 
 
 fierce nf chord 
 
 CONFIDENTIAL 
 
 NATIONAL AOVISOQV 
 CONH'TTfl rO* A|>0N4(;TKS 
 
 Figure 22.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 10°; S = 0°. 
 
NACA ACR No. L5G31a 
 
 Fig. 23a-f 
 
 Upper surface 
 Lower surface 
 
 CONFIDENT 
 
 (a) M = 434. 
 
 (b) AI=0.J66. 
 
 \ 
 
 
 
 
 
 s._^ ^ 
 
 
 
 
 
 
 
 \ 
 \ 
 \ 
 
 
 ^r 
 
 
 
 ^ 
 
 >- 
 
 '^^^ 
 
 -^-*=^ 
 
 A 
 
 
 
 
 
 V 
 
 (c) 
 
 A7 = 
 
 0. 
 
 6/5. 
 
 
 p- — -- 
 
 X 
 
 
 
 
 I 
 
 H 
 
 :>^tzz 
 
 1 
 
 
 / 
 
 
 
 
 
 c 
 
 (d) 
 
 Ai=a< 
 
 J6/. 
 
 
 (e) M = C.701. 
 
 O ZO 40 60 60 /GO 20 
 
 Perceni chore/ 
 
 CONFIDENTIAL 
 
 r^ 
 
 
 
 P--i3^ 
 
 \-^^^ 
 ■*^ 
 
 
 ^ 
 
 / 
 
 y 
 
 Y 
 
 
 ^cr 
 
 f 
 
 
 
 
 
 p 
 
 (f) 
 
 fA^ 0. 
 
 7/9. 
 
 
 40 
 
 60 60 
 
 /GO 
 
 NATIONAL ADVISORY 
 COMMITTEt FOB AEftONiUTrCS 
 
 Figure 23.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0. 20-chord flap, a = 10^; S = -4°. 
 
Fig. 24a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O LoiA/er surface' 
 
 (a) n-O.453. 
 
 
 
 
 
 
 \ 
 
 
 
 
 
 \ 
 
 p 
 
 
 
 V 
 
 
 
 
 
 
 ^ 
 
 X- 
 
 
 
 
 y^ 
 
 
 
 ^^^*"=Si^ 
 
 / 
 
 
 
 
 
 r- 
 
 X 
 
 \ 
 
 
 
 
 j\ 
 
 H 
 
 f<f~V' 
 
 h 
 
 V 
 
 
 
 
 
 B 
 
 (d) 
 
 r^-O.679. 
 
 
 -/ 
 
 / \ 
 
 r' 
 
 .^^- 
 
 .^ 
 
 r- ff==«>^ 
 
 -\ 
 
 ^ 
 
 H 
 
 y 
 
 
 
 ^r 
 
 / 
 
 
 
 
 
 
 ^/; 
 
 Ai~0. 
 
 7ZO. 
 
 
 (eJ >/-0.7/Z. 
 ZO 4^ 60 eo lOO O ZO 40 GO SO lOO 
 
 Percenf chord 
 
 NATIONAL AOVISOQV 
 COMNITIK F(M AfKMlUTtCS 
 
 CONFIDENTIAL 
 
 Figure ?4. - Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a » 10°; S = -8°. 
 
NACA ACR No. L5G31a 
 
 Fig. 25a-f 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Lower surface 
 
 (a) M =0.459. 
 
 (b) /^= 0.55 7. 
 
 (d) /^= 0.696 
 
 (e) r/=0.7JS. 
 
 (f) n=0.72/ 
 
 ^ 
 
 -^ — 
 
 
 Z:z^ 
 
 
 A 
 
 y 
 
 
 
 
 V 
 
 
 
 
 
 -e^ 
 
 O ZO 40 60 60 /OO O ZO 40 60 SO /OO 
 
 Percent cf>ord 
 
 CONFIDENTIAL 
 
 NATIONAL ADVISORY 
 COMMITTEE FM AERONAUTICS 
 
 Figure 25. - Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. o = 10°; S = -12°. 
 
Fig. 26a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 o 
 
 X Upper surface 
 O Lower surface 
 
 r 
 
 (aJ n^O.43/. 
 
 
 ^^ — 
 
 ^■v 
 
 — — 
 
 cr 
 
 ( .^ 
 
 ^xr-o-^ 
 
 ^ 
 
 f^ 
 
 
 ( 
 
 
 
 -\ 
 
 p>^ 
 
 
 
 n^o.eao. 
 
 
 M=o.lO^. 
 
 o 
 
 le) M -0.13 5. 
 ao 40 60 60 /OO 
 
 (f) M-O. 145. 
 ^0 40 60 60 100 
 
 Percent cford 
 
 NATIONAL ADViSOav 
 COMMlTTtt FM AltOHAUllCS 
 
 CONFIDENTIAL 
 
 Figure 26.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 2°; S • 4°. 
 
NACA ACR No. L5G31a 
 
 Fig. 27a-f 
 
 CONFIDENTIAL 
 
 X Upper surface^ 
 O Lower surface 
 
 (a) n =0.434. 
 
 ^ 
 
 .^^^ 
 
 V 
 
 
 
 -^r 
 
 
 ^ 
 
 \ 
 
 ^ 
 
 
 ( 
 
 
 
 1 
 
 ^-^ 
 
 
 (b) 
 
 n =0.6/8. 
 
 
 \ 
 
 a 
 
 
 !1J 
 I. 
 
 / 
 
 /^' 
 
 .-.r^ 
 
 
 
 — ^. 
 
 
 
 L>^ 
 
 
 
 N: 
 
 ^^-^ 
 
 / 
 
 
 
 'N 
 
 l^^-^r^ 
 
 1 • 
 
 I — 
 
 
 
 
 
 T" 
 
 — /?: 
 
 i::iQ, 
 
 {C/J M= 0.703. 
 
 O 
 
 (e) M =0.7 32. 
 iO 40 60 80 ZOO 
 
 (f) M =0.744. 
 20 40 60 80 
 
 Percent chord 
 
 CONFIDENTIAL 
 
 NATIONAL ADVISORY 
 COMMITTEE FOB AERONAUTICS 
 
 100 
 
 Figure 27.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 2° ; 8 = 8". 
 
Fig. 28a-f 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
 X Upper 5urfac£> 
 O Loiver surface- 
 
 (o) /^-0.-f-3/. 
 
 (t>) M-O.6/5. 
 
 .C5 -/ 
 
 
 o 
 
 \ y 
 
 ^ 
 
 
 ^ 
 
 ^. 
 
 / 
 
 
 
 ^ 
 
 '"''o ■ 
 
 % 
 
 (c) 
 
 M=0.<. 
 
 ^^73. 
 
 
 (cf) M-0.70Z. 
 
 (e) M-0.13/. * (t) M-07^5. 
 
 zo 4-0 60 eo /oo o 20 40 60 ao /OO 
 
 Percerit chord 
 
 N&TION&l ADVISOBV 
 
 oxmrTTii rn tuoaiuiKS 
 
 CONFIDENTIAL 
 
 Figure 28.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a - 2°; S - 12°. 
 
NACA ACR No. L5G31a 
 
 Fig. 29a-f 
 
 CONFIDENTIAL 
 
 X Up/XT surfacp 
 O Lower surface 
 
 (a) /^=0.430. 
 
 JV— 
 
 --- — ■ 
 
 -\ 
 
 — ^r 
 
 J/- 
 
 
 \ 
 
 
 
 
 ' ° — J 
 
 r 
 
 
 
 
 
 a 
 
 -2 
 
 
 ^ 
 
 
 I 
 
 /^ 
 
 -^ 
 
 \ 
 \ 
 
 ^ 
 
 
 r 
 
 ^ 
 
 \ 
 
 r 
 
 
 
 1 
 
 >-Cv_C?»~' 
 
 \ 
 
 (c) 
 
 /^= Ot 
 
 ^47. 
 
 
 (d) M-a 104. 
 
 (e) n- 0.1 3 3 
 £0 40 60 QO /CO 
 
 (f) M = 0.74/. 
 20 4i> 60 80 
 
 /OO 
 
 Perce/?/ c/)orc/ 
 
 CONFIDENTIAL 
 
 NATIONAL AOVISOQV 
 COMMITTEE rot AERONAUTICS 
 
 Figure 29.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 4°; S = 4°. 
 
Fig. 30a-f 
 
 NACA ACR No. L5G31a- 
 
 CONFIDFNTIAL 
 
 X Upper surface 
 O Lower surface 
 
 (a) M- 0.434. 
 
 ft) Af- OSS) a. 
 
 fcf) >f=0.705. 
 
 -/ 
 
 
 — 1 
 
 
 a 
 
 
 fy 
 
 r 
 
 
 . \ 
 
 
 
 
 1 
 
 L? 
 
 r 
 
 
 
 
 
 
 ii 
 
 fA 
 
 M- n 7 
 
 ^5f 
 
 
 o 
 
 30 
 
 -^r 
 
 /oo 
 
 o 
 
 (f) M =0.742. 
 20 40 60 60 
 
 f^erce.nt chore/ 
 
 NATIONAL AOVlSOnv 
 COMHITTEl rOI AIMWAUTlCS 
 
 /GO 
 
 CONFIDENTIAL 
 
 Figure 30.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with .20-chord flap. a = 4°; S - 8°. 
 
NACA ACR No. L5G31a 
 
 Fig. 31a-f 
 
 CONFIDENTIAL 
 
 X Upper surface 
 O Lower surface- 
 
 ta) fi-O.^bi. 
 
 (b) Af-0.je4. 
 
 (d) n-O. 70/ . 
 
 ■Per 
 
 
 
 ie) A7-0. 127. 
 20 40 60 60 
 
 if) MzO. 14/. 
 /OO 20 40 60 60 /OO 
 
 /^ercent c/iorc^ 
 
 CONFIDENTIAL 
 
 Figure 31.- Pressure distribution for a modified NACA 65,3-019 
 airfoil with 0.20-chord flap. a = 4°; S = 12°. 
 
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Fig. .39a 
 
 NACA ACR No. L5G31a 
 
 NATIONAL ADVISORY 
 COMMITTCE FOB AEDONtUTICS 
 
 J .^ ^ .6 
 
 Mach number, M 
 
 (a: 
 
 ■2°. 
 
 Figure 39.- Effect of compressibility on section normal -fo rce 
 
 and moment coefficients. 
 
NACA ACR No. L5G31a 
 
 Fig. 39b 
 
 .2 
 
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 Figure 39.- Continued. 
 
Fig. 39c 
 
 NACA ACR No. L5G31a 
 
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NACA ACR No. L5G31a 
 
 Fig. 39d 
 
 ./ 
 
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 Figure 39.- Continued. 
 
Fig. 39e 
 
 NACA ACR No. L5G31a 
 
 ^ 
 
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 NATIONAL ADVISORY 
 COMMITTEE FM AEBOHAUTICS 
 
 .J A .5 .a 
 
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 Figure 39.- Continued. 
 
NACA ACR No. L5G31a 
 
 Fig. 39f 
 
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 .7 
 
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Fig. 39g 
 
 NACA ACR No. L5G31a 
 
 .S .6 
 
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NACA ACR No. L5G31a 
 
 Fig. 39h 
 
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 NATIONAL ADVISORY 
 COMMITTEE FOB AERONtUTICS 
 
 O ./ .2 .3 .4 S 6 
 
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 Figure 39.- Continued. 
 
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Fig. 39i ■ 
 
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 Figure 39.- Continued. 
 
NACA ACR No. L5G31a 
 
 Fig. 39j 
 
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 Figure 39.- Concluded. 
 
NACA ACR No. L5G31a 
 
 .3 4 .S .6 
 
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 Figure 40.- Effect of compressibility on section hinge-moment 
 
 and flap normal-force coefficients. 
 
NACA ACR No. L5G31a 
 
 Fig. 40b 
 
 CONFIDENTIAL 
 
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 Figure 40.- Continued. 
 
Fig. 40c 
 
 NACA ACR No. L5G31a 
 
 ^/, 
 
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NACA ACR No. L5G31a 
 
 Fig. 40d 
 
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Fig. 40e 
 
 NACA ACR No. L5G31a 
 
 CONFIDENTIAL 
 
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NACA ACR No. L5G31a 
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 Fig. 40f 
 
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 COMMITTEE FOt AERONAUTICS 
 
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 Figure 40.- Continued. 
 
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Fig. 40g 
 
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 COMMITTEE FM AERONAUTICS 
 
 ^A O 
 
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 Figure 40.- Continued. 
 
NACA ACR No. L5G31a 
 
 Fig. 40h 
 
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 NATIONAL ADVISORY 
 COMMITTEE -FM AERONAUTICS 
 
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 Figure 40.- Continued. 
 
Fig. 40i 
 
 NACA ACR No. L5G31a 
 
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NACA ACR No. L5G31a 
 
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 .3 .4 .5 
 
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 NATIONAL advisory' 
 COMMITTEE FM AERONAUTICS. 
 
 .7 
 
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 pressure-coefficient differential, Pl~Pu* 
 
Fig. 44d-f 
 
 NACA ACR No. L5G31a 
 
 
 c 
 
 
 0) 
 
 o 
 
 I 
 
 t, 
 
 3 
 
 <o 
 
 (D 
 ■Q 
 
 Q 
 
 -C 
 (J 
 
 O 
 
 C 
 CQ 
 
 I 
 
 -a. 
 
 
 
 (of) a-- 4 ". 
 
 I 
 
 Ce; ex ^ 6° 
 
 .3 rh .S 
 
 CONFIDENTIAL 
 
 ffj cc = a: 
 
 Alach number, M 
 
 .6 
 
 .7 
 
 Figure 44.- Continued. 
 
NACA ACR No. L5G31a 
 
 Fig. 44g,h 
 
 I- 
 
 o 
 
 CL 
 
 ./_ 
 
 1/7) <x = /^r 
 
 CONFIDENTIAL 
 
 ..4- 
 
 ..3_ 
 Alach number, /^ 
 
 NATIONAL ADVISORY 
 COMMITTEE EM AERONAUTICS. 
 
 I I 
 7 8 
 
 Figure 44.- Continued. 
 
Fig. 44i,j 
 
 NACA ACR No. L5G31a 
 
 I 
 
 I 
 
 
 >5) 
 
 c 
 
 0) 
 
 u 
 
 C 
 
 .&x 
 
 4- 
 
 O- 
 
 -4- 
 
 0_ 
 
 ii) oc^a' 
 
 NATIONAL ADVISORY 
 COMMITTEt FM *f«ON«UTICS 
 
 .a. 
 
 CONFIDENTIAL 
 
 (jJ 0:^4". 
 .4 5_ 
 
 XI 
 
 ..3_ 
 Mach number, M 
 
 .6. 
 
 .7_ 
 
 .Q 
 
 Figure 44.- Concluded. 
 
UNIVERSITY OF FLORIDA 
 
 262 08 
 
 05 004 8 
 
 UNIVERSITY OF FLORIDA 
 DOCUMENTS DEPARTMENT 
 120 MARSTON SCIENCE UBRARY 
 RO. BOX II 7011 
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