MCft-T/^-/Mi^^fc fcfc ,K. LEADON O CO H NATIONAL ADVISORY COMMITTEE < i FOR AERONAUTICS TECHNICAL MEMORANDUM 1240 AIRFOIL MEASUREMENTS IN THE DVL fflGH-SPEED WIND TUNNEL (2.7 -METER DIAMETER) By B. G'othert Translation of "Profilmessungen im DVL-Hochgeschwindigkeitswindkanal (2,7 -meter Durchmesser)." Lilienthal-Gesellschaft fiir Luftfahrtforschung, Bericht 156 Washington June 1949 feocuME^rrs department -l(eZ "6^(^17 NKTIGML ADVISQFY COI^fl^rTEE FOE AERONAUTICS TECBNICAL MEMSRANIUM 12ii0 AIRFOIL MEASUREMEOT'S IN THE DVL HIGH-SPEED WIND TUNNEL (2.7-ME:rER DIAMETER)* By B. Gbthert SUbflARY The present report is a 'brief summary of the investigations on symmetrical and cambered airfoils in the DVl high-speed wind tunnel- In the light of measurements of wings of different aspect ratio with equal profile it is shown that the effect of the aspect ratio on the increeise in lift with rising Mach number expected on the basis of calculations is closely confirmed by the measurements- Several recent experiments on a wing with the small aspect ratio of b^/p = 1-15 are discussed, where the lift diaclosed no distitrbances within the test range, that is, up to angles of attnckr of a - S° and Mach nvmbere of M = 0-9, and the drag (at c^ = O) starts to increase at a much higher Mach nvmiber than for a wing with the same profile but gi«o«U;er aspect ratio. I. INTRODUCTION The airfoils tested 1© the DVL high-speed tunnel are represented In figure I-*' The tests priiaarily involved a series of symmetrical * 'Prof ilmessungen im IVL-Hochgeschwlndlgkeitswinclkanal (2,7-nieter Durchaesser) . Lillenthal-Gesellschaft flir Luf tfahi'tforschung, Berlcht I56, pp. 5-I6. ^The notation for the profile used in figure 1, which is normal practice in the DVL, is exemplified on the.NACA airfoil 1 30 12 - 1-1 Ij-O: Depth of camber, percent f/l = 1 Distance of maximxm camber from the leading edge, percent Xf /I = 30 Thlclcness ratio, percent d/i = 12 J^ose radius, percent -i-i — y. «= 1-1 (d/:)2 Jackward position of maximum thickness from the nose, percent x^n = kO The added lettergNACA indicate that the contour of the NACA systemlga- tion is maintelnicd- *^ MCA TM 12i*0 standard NACA airfoils with thickness ratios ranging "between d/l = 0.06 and d/l = O.18 (reference l) • In addition three sjanmetrical profiles having thickness ratios of 9^ '^■'^■f s^^d 15 pe;ocent with zaaiimim thickness moved "backward were investigated; the tests, however, have not yet "been correlated because of their us© for different purposes and the long delay between tests (reference 2) . So far only one profile with 2-percent camber and normal thickness distribution has been investigated at the request of the Messer- schmitt Co. (reference 3) . A greater number of profiles of various camber are ready to-be tested. .(In the meant-'me three more cambered profiles were investigated up to January 19^3') On the previously enianerated profiles the pressure distribution was measured in the center section of the rectangular wings fitted with end plates and the drag was computed from measurements of momentum loss behind the model wings. The wing chord of all these models was 500 and 350 millimeters, respectively, so that the Reynolds number even at the lowest Mach number of the tests, (M = 0.30), with EQjjj^j^ s2.2 X 10^ was still oonsidarably above the critical value for the boundary -layer reversal on flat plates. In addition a number of profiles with thickness ratio up to d/l = 0.50 and chords of 1 = 60 miUimetere and loom vere tested so as to obtain a preliminary insight into the effects of such profiles on the drag at high Mach numbers, especially with respect to propellers . II. RESULTS OF TESTS ON PROFILES AT HIGH SUBSONIC SPEEDS 1. Lift of Symmetrical Profiles The value Bcq/^ governing the lift at high airspeeds was determined for symmetrical profiles of varying thickness from DVL tests and plotted against the Mach number in figure 2. The theoretical curve for t he lift increase according to Prandtl's approximation ratio l/ Vl-J?, is included for comparison. In the p^jre subnonic range, that is, up to the critical Mach number, M*, at which the velooity of sound Is reached or exceeded locally, the tost curves of thin profiles up to d/Z = 12 percent manifest an increase in the Bca/^ value which is in good agreement with Prandtl's calculations- Only on the thickest profile with d/Z = 18 percent does the lift increase disappear with rising Mach number, obviously as the result of the more unfavorable devBlopnent of the boundary layer on thick profiles. NACA TM 121+0 After exceeding the critical Mach numter M*', that is, as soon as the sonic velocity is locally exceeded at the profile and compressibility shock occurs, the premises of the Prandtl law are no longer complied with, hence no agreement is to be expected between the experimental and theoretical curve. V/hile in this range the lift increase for thin profiles up to d/Z = 12 percent increases at first, and then shows no marked decrease until after substantially exceeding the critical Mach number, the lift on the thick profiles of d/Z = 15 percent and l8 percent decreases as soon as the critical Mach number is exceeded and finally drops to zero at about M = 0.83. This dissimilar behavior of the lift on thin and thick profiles in the supercritical range of Mach numbers can be explained frcan the pressure-distribution records by the dissimilar developcient of the compressibility shocks. On studying the range of Mach niimbers on a profile with compressibility shock occurring only on the suction side (fig. 3) > it is seen that this shock has already: moved much farther toward the trailing edge on the thin profile than on the thick profile J this causes the earlier manifested lift increase on the thin profile and the lift decrease on the thick profile, as seen by a comparison with the pressure distribution curve at Bmnll Mach number indicated by thin lines. If at further rise in Mach niunber the pressure side also happens to ccme into the range of the ccmpressibility shock (fig. k) , these shocks develop on suction and pressure side near the trailing edge of thin profiles at sufficiently high subsonic Mach numbers, so that a large lifting surface remains, which however is substantially less compared to the pressure distrilsution without pressure-side compressi- bility shock. On the thick profile, on the other hand, the suction- side compressibility shock lags behind in its rearward movement, so that for the lifting surface at the forward part of the profile there is a corresponding negative lifting surface at the rear part of the profile and the lift therefore drops to zero or even negative values in spite of the positive angle of attack of the profile. The formation of the cited pressure distributions also affords a possibility of estimating the lift distribution on profiles at higher Mach numbers than correspond to the measured range. As the pressure distribution of the thin profile in figure k indicates, at M = 0.88 the compressibility shocks on suction and pressure side have already travelled close to the trailing edge so that on this profile no further fundamental change in lift is to be expected. This is readily seen by a comparison with the pressure distribution of an airf»il at a supersonic speed, as represented in figure 5 according to a measurement by Ferri (reference k) . The pressure distribution NACA TM 1240 at supersonic velocity resembles in the low-pressure Tariation as well as in the position of the compreseihility shock the pressure distribution on the thin profile. Accordingly it is to he presumed that the curves for the lift increase in figure 2 may not he extrapolated to zero for the thin profiles, hut that the lift is actually maintained throughout even at further approach to sonic velocity. Several tests in the UVL high-speed tvinnel are availahle \*ich show that the lift at the very high Mach numher of M = 0.90 does not continue to drop steeply hut rather starts again to rise to higher lift coefficients. For the same reason it is to he expected that on the thick airfoils a rise in Mach numher heyond the test range will he accompanied hy a lift increase to rational values, hecause the suction- side compressihility shock will travel to the trailing edge even on the thick airfoils" at sufficiently high suhsonic Mach numhers and so cancel the overlap in' the pressure- distrihution plots. 2. Lift of Camhered Airfoils Only one camhered airfoil with 2 -percent depth of camber and 13-percent thickness ratio has been tested, althou^^i a large number of camhered model wings had been prepared for testing. Aside frcm the lift increase hc^^ and the drag which are similar to those on symmetrical airfoils of equal thickness ratio, an unusual result of these measurements was the sudden displacement of the angle of attack for zero lift at high Mach numhers (fig. 6) • While up to M = 0.8 the angle for zero lift of the cambered airfoil waa located at a^ c -1.5°, it rose to a^ = 2°, or changed by ahout 3.5° when the Mach number increased to M = 0.86. Thus a wing must be given a 3.50 higher setting in this range, if lift conditions similar to those at small Mach numbers are to he reached- This displacement of zero angle of attack is also attributable to a particular formation on the pressure side of the conpressibillty shock that suddenly moves in direction of the wing trailing edge in the 0.8 to 0.86 Mach number range (figs. 7 and o) . 3. Neutral Stability Point of Symmetrical Profiles The neutral stability point which owing to the symmetry of the* profile in this instance also indicates the position of the applied moment was computed for the series of NACA airfoils of varying thick- ness from the pressure- distribution measurements over a range of small lift coefficients. According to figure 9 the curves for the varyingly thick profiles are so staggered at small and medium velocities that for thick profiles the neutral stability point lies nearer to NACA TM 12i+0 the leading edge- This is due to the boundary layer which thickens more on the suction than on the pressure side and therefore reduces the lift at the trailing edge in accord with the effective Jet bovmdary at the edge of the hovindary layer. Thin decrease in lift caused hy the one-sided thickening of the "boundary layer is naturally BO much greater as the profile is thicker. At Increasing Mach number all neutral-point curves shift toward the wing leading edge, the shift being greater for the thicker profiles. This also is explainable by means of the boundary layer, since with rising Mach number the pressures governing the boundary -layer development increase and thus the same effect is produced as by a further thickening of the profile. After exceeding the critical Mach number, that is, after locally exceeding the velocity of sound the curves of the thin airfoils with 6* to 12- percent thickness ratio bend sharply in direction of the tall-heavy neutral stability -point positions, while the curves of the thick profiles continue to rise in the old direction- This dissimilar behavior In the range of the canpreBBlbility shocks is already evident from the previously discussed pressure-distribution curves in this speed range (figs. 3 and k) ■ Owing to the dicslmilar compressibility- shock position on the pressure and suction side of thick airfoils it results in overlap in the pressure distribution which as a result of the downwash shifts the applied moment forward. This overlap cancels out on the thin Birfoils especially at high Mach numbers. On the contrary, the pressure distributions disclose that the applied moment travels in direction of the wing center as in pure subsonic flow. k- Drag of Symmetrical Profiles The drag of the symmetrical profiles with different thickness ratios explored in the high-speed tunnel is plotted eigainst the Mach number in figure 10 for symmetrical air flow. There is no change in drag coefficient up to the critical Mach number, that la, in the subsonic range, but after a certain Increase in Mach number beyond the critical, the well known steep-drag rise occurs. The amount of the drag rise covered on an average by these tests is up to drag coefficients (referred to frontal area) of around 0.1+. In the pure subsonic range the drag coefficient of the profiles tested is practically constant, although the Reynolds number of the tests increases with increasing Mach number. This observation was made on the symmetrical NACA airfoils as well as on other tests on different models. In the eYaluation of these curves NACA m 1240 It should te remembered that the Reynolda number effect leading to a drag decrement is opposed by the effect of compressibility on increasing the velocities which acts In the sense of a drag Increase. Calculations for predicting the drag curve at eubcritical Mach nvunibers have up to the present not given the complete equality of these two mutually counteracting Influences; it must not he forgotten that these approximation processes considered only the Beynolds number and the apparently Increased profile thickness, but not the con^resai- blllty in the formation of the boundary layer. By this oranission no provision is made, for exainple, for the fact that substantial temperature differences occur within the boundary layer which, for Instance, at the point where locally the velocity of sound is exactly reached, already amounts to about 1*5° C. So when the bo\JDdary-J.ayer laws obtained for incompressible flow are applied to compressible flow the difference between measurement and preliminary calculation is, in view of this omitted temperature gradient, not aurprlsing. Incidental to the steep drag Increase the question arises as to ^fAl&t values the drag will attain at further Increasing Mach number. Frcxn ntmerous measurements in the high-speed tuiuiel it was found that the drag coefficient based on frontal area appears to ^nd toward a terminal value of the order of magnitude of o^ = 1.2 at high Mach numbers, so that this value may serve as reference point for an extrapolation of the drag coefficients beyond the test range. This value is confirmed satisfactorily by the measurements on thick syimietrical airfoils of small chord represented in figure 11, and whose thickness ratio had been raised to 50 pei'cont in view of propeller root profiles. The profile chord of these wings rangsd between 36 and 60 millimeters, the span was about 5OO mllltmsters. The wings were mounted on an adjustable central spindle, the drag at zero setting was detormlned by means of a thin wire extending forward at the model. These measurements also show within measuring accuracy the nearly constant drag coefflciont \xp to the critical Mach ntmber. m. CORREIATION OF THE PROFILE MEASUREMENTS IN THE WL HIGB-SEEED TUNNEL WITH THE MEASUREMEWTS IN THE LARGE WIND TUNNEL OF THE WL The measurement B in the DVL high-speed tunnel generally cover a speed range of from 100 meters per second to about 90 percent of the NACA TM 12li0 velocity of sound. At the lowest airspeed compresal'blllty visually has such a small effect that the test data in this range are directly comparahle with the measia'einents in low-speed wind tunnels. The series of symmetrical NACA airfoil sections, for which an abstract from the high-speed measurement is given in the present report, wao also tested in the large tunnel of the DVl (5x7 met-sre) at low speeds (reference 5). The data from "both tescs are compared in figures 12 and Ik' For the comparison the tests from the high-speed tunnel wore extrapolated to the Mach nioaber of M = 0.2 corresponding to the speed in the Inrce tvinnel which involved no difficulties in view of the flatness of the curves in this range of Mach numhers. The Reynolds nianber in "both wind-tunnel tests was Ee ~ 2.7 X 10°. The agreement "between the measurements in "both tunnels is unusually good. The discrepancies in drag and in mcanent about the quarter -chord point range within measuring accuracy. Only for the lift increase ^Cg^/oa do the hfgh-speed mfiaouremente systematically exceed those of the larce tunnol "by about 5 percent. This difference is probably due to the fact that normal wings with aspect ratio \r/F = 5 were used in the lar:_,e tunnel,, while the -i^ints in the high-speed tunnel were fitted with end plates. In the conversion of the angle of attack with the end- plate measurements to two-dimensional flow, it is concel'^ble that the simple equations employed for talcing the end plate Inxo account, according to 0- Sclirenk, still contain certain inaccurax:ies. IV. EFFECT OF ASPECT BATIO ON THE AERODYNAMIC COEFFICIENrS OF RECTANGULAR WINGS The previously repoi'ted test data were obtained, as a rule, from pressure-distribution measurements In the center section of rectangular wings, hence reproduce the profile characteristics for two-dimensional flow in good approximation. The drag measurements on the sma ll wings of large thickness with their comparatively large aspect ratios, b^/F ~ 10 or more, form the only exception. 1. Effect of Aspect Ratio In the pure subsonic range it is of Importance to know whether the aspect ratio acts in accord with the theoretical calculations (reference 6) on the rise in the lift increase ^Cgjha with increasing Mach number. According to these calculations the lift 8 NACA TM 12»+0 increases at constant angle of atta ck of the vrlng only In two- dimensional flow In the ratio l/\/l — M^^ while for finite aspect ratio the rise is smaller and disappears altogether in the extreme case of very small aspect ratios. On the "basis of later measuroments on wings with the eaine profile NACA 00012 - 1.130 with three different aspect ratios (hS/F = « from pressure-dlstrlTDvction moasurementq, and h^/S" = 6 and I.I5 from force measuroments) this question regarding the lift increase could "be answered to the oxtent that the tost data are in satisfactory agreement with the calculations. On the wing with very small aspect ratio in particular, h^/F = 1,15, the rise of the Scg/Sa values with increasing Mach number is oxtromely small (fig. 15). According to the cited theoretical investigations this diminished rise in 3ca/^ for small aspect rati? is attrihutahle to the fact that, while the lift "belonging to the effective angle of attack of the wing increases in accord with t he c alculations for two-dimensional flow with the Prandtl factor l/Vl -Ir, the induced dowawash angle at the wing also increases as a result of the increased lift and so reduces the lift "by way of the effective angle of attack- This reduced lift increase is therefore so much greater as the downwash effect is greater, that is, as the aspect ratio "becomes smaller. For the very seme reason this reduction disappears completely when ae for infinite aspect ratio the downwash is Infinitely small. The measurements for h^/F = » and \r/F = 6 show no substantial differences in their fundamental distribution even in the supercritical range of Mach numbers, ae for instance, the temporary marked rise in lift after exceeding the velocity of sound followed by a marked drop as a result of the pressvire-distribution overlap. On the wing with aspect ratio b^/F = 6, which was measured up to M = O.90, it is noted again thaji the lift characteristic ^0^/^ at very high Mach numbers of M = 0-90 no longer decreases but begins to rise again with further increase in Mach number. From this general ^.rend for b^/F — > ^ and t^/F = 6 the test curve of the wing with b^/F = I.I5 differs in noticeable manner, to the extent that this wing neither exhibits a greater increase nor a marked decrease in the ScgAx curve in the supercritical range of Mach numbers- NACA TM 12U0 9 2. WllSfrWlth Aspect Ratio \^t^ = 1.15 (KACA airfoil section 00012 - 1-130)2 The wing designed with the NACA airfoil section 00012 - I.I30 had a chord of X = O.35 metarg- It vas supported at virv? centar "by a single halance support vhich transmitted the lift and drag of t'J© model wing at a k'jP angle obliquely to the balance. The moment was' meeisured on a separate balance over a symmetrical uoraent arm. Lift; Figure I6 shows the lift coefficient Cq^ plotted against the angle of attack a for several Mach numbers. In contrast to all profile measurements loiown- so far, no disturbances of any kind were observed throughout the entire range of Mach numbers up to M = 0.91» notwithstanding the fact that the angle of attack was raised up to a = 8°, save that the highest Mach number M = O.91 at high a no longer fields the slight deflection of the lift curve in direction of higher Ca/ but on the contrary, a slight deflection In the opposite sense, that is, toward lower c^^. However, it is to be noted that the lift coefficient as a result of the small aspect ratio at a = 8° mounts to little more than c^ = 0.2. Schlieren observations made contemporary with the force measirre- ments indicated that at M » O.82 severe compressibility shocks occurred even at zero angle of attack- According to the Bchlieren photographs the location of the caapreeelblllty shocks on the pressure side was a little nearer to the trailing edge than on the suction side, so that the overlap obsei*ved on the profile in two-dimensional flow and the loss in lift produced by It, was largely canceled out. The small effect of the I-Iach number on the rise of the lift curves Ca B f(a) is more accurately represented in fifure 17 on the variation of the lift increase dc^/^ with respect to the Mach number. 3ca/^ Increases very little, in fact, a little lees even than the theory stipulates- Mcment: In figure 18 the moment coefficient referred to the \/h chord of the wings is plotted against the lift coefficient Ca for Boveral Mach numbers- According to it the mcanent curves are not rectilinear any more even at low speeds, and this oscillating variation is considerably simplified at increasing Mach number - ^These meaaiurementa are to be published as a separate report In the near futvire- XO NACA TM 12U0 But dleoontinuouB mcjnent Jumps as on wings In tvo-dlmsnsional flow ore completely absent. Another unusual fact is that the moment curves with rising Mach number turn considerably in the unstable sense, which corresponds to a shift of the neutral stability point tofward the leadiivg edge (fig. 19) • Even at low speeds the neutral stability point, which owing to the symmetrical \ring profile is also identical with the center of pressure, lies about l8 percent ahead of the l/k point, that is, only about 7 percent aft of the wing leading edge. This forward position is explained by the fact that owing to the small aspect ratio of the wing, the stream lines are considerably curved compared to the wing chord which at the tip ©specially leads to greater negative lift and even loss of lift. Since at high Mach numbers for the model wing with b^/F = 1-15 t!icre corr&eponds a ccmperablo wing in incompressible flow with a smaller aspect ratio, the further forward shift of the neutral stability point with increasing Mach number is readily understood, even though the magnitude of this travel was, at first, surprising. For eiangple, the neutral point lies already at the nose when M « 0.76 is reached; at still higher M it and with it the center of pressure lies even ahead of the wing leading edge. A ccmparipon with Wlnter*B testa on wings of Hrn^n aspect ratios (reference 7) indicated that a wing with b^/F = 1 and 12. 7- percent thickneBS ratio corresponding to the previously reported measurements likewise ejchibltod a very forwardly located neutral stability point, which was determined as being less than 10 percent behind the leading edge. Drag: The drag of the wing with b^/F = 1.15 in gynnetrical flow is plotted against the Mach number in figure 20, along with the corresponding drag cvirves of the rectangular wing with b^/F « 6 and of a wing with 35° positive sweepback and b^/F = 6. All wings had the same NACA 00012 - I.130 profile- The drag increase on the wing with b2/F = 1-15 is considerables delayed relative to the rectangular wing and starts only a little earlier than on the wing with 35° positive angle of sweepback. The reduction from b^/F = 6 to b^/F 3 1.15 therefore causes the sesae delay in drag increase at high Mach numbers, as obtained for b^/F «= 6 with a 30° positive sweopback. The cause of this beneficial drag action of the wing with small aspect ratio is likely to be found in a transition to three- dimensional flow as a result of the small aspect ratio. At the wing tips, whose range of influence in this particular case amounts to a considerable portion of the total wing area, a flow occurs which NACA TM 12U0 11 is rather ccjapara"ble to an axlally eynoietrlcal than to a two-dimenalonal flow, ai^ as a conseqiience exhibits the more favorahio characteriatics of a three-dimensional flow< V. EECAPITULATION 1. A "brief survey of the airfoil maasurements In the DTL high- speed tunnel is given in the light of several dia^arns. 2. By meesurements on wings of different aspect ratios it is proved that in the pure subsonic range the aspect ratio affects the rise of the lift as the theory stipulates. Thus the lift for win gs of infinite aspect ratio, for example, increases with l/i/l - >r, while in the extreme case of vanishingly small aspect ratio the lift remains constant in spite of increasiiig Moch number. 3- High-speed measurements on a rectangulai' wing with b^/F « 1.15 disclosed that in the entire tost range, that is, up to M = 0.90 and angles of attack up to a = 8°, no rough discontinuities in lift and moment were observed. The drag rise at zero lift on this wing compared to the wing of b^/F = 6 was shifted considerably toward higher Mach numbers. DISCUSSION Khappe: Gothert's drag curves showed no rise at BmalT and medium Mach numbers. But according to theoretical considerations a linear increase above the Prondtl factor should occui', as definitely observed in the Hoinkel tunnel- The reason that this effect did not occur on Gothert's curves might be found in a superposition of mutually opposing effects of Reynolds and Mach number. In the measurements in the Heinkel high-speed tunnel the effect of the Reynolds number was suppressed by turbulonco edges." This is also plainly observed on the position of the test points on port of Gothert ' s curves . Helmbold. According to Gothert (JB. l^^+l, DLF, p. I, 684) it is possible to correct the airspeed in the tunnel, even when supersonic areas already occur at the profile. Pi'omise is, of course, that they do not yet reach the tunnel wall. But this condition exists long before reaching the velocity of sound. In this respect the free Jet shows more favorable characteristics, since on it the supersonic areas extending out from the profile 12 NACA TM 1240 ore able to reach the Jet toundary only when the Jot leaves the noizle with sonic velocity. The corrective possit^lity which in the tunnel rests on the measurement of the minimum pressixre at maximum speed at the tunnel wall and provides reference points for the prediction of the intensity of the reflected dipole grid "by means of the Prandtl-Busemann law. is, on the free Jet, given "by the measurement of the greatest Jet expansion and the subsequent determination of the reflected dipole grid. The correction possi- "bility therefore ends only at sonic velocity on the fi'ee Jet. Trai^latcd by J. Vanler National Advisory Committee for Aeronautics NACA TM 121^0 13 REFERENCES 1. Gothert, B.: ProfilmesBungen tm UVX-Hochgeschwindlgkeltswindlcanal (2,7 m ^). FB 1U90. Gothert, B. : Druclrverteilunge- und ImpulsverluBtBchaubilder fur dae Profil NACA 00 06 - 1,1 30 tei hohen Unterechallgeechwin- dlgkciten. FB 1505/1- FB 1505/2: dasselbe fiir dae Profil NACA 00 09 - 1,1 30, FB 1505/3: daaselbe fUr daa Profil NACA 00 12 - 1,1 30, FB 1505/U: daseelbc fur das Profil NACA 00 15 - 1,1 30, FB 1505/5: daeselbc fto das Profil NACA 00 I8 - 1,1 30. G^hert, B.: Hochgeschwindigkeitsunterauchungen an symmetrischen Profilen mit verechiedenen Dickenverhaltnissen im DVX-Hochgeecliwin- digkeitsvindlcanal (2,7 m ^) und Vergleich mit Meseungen in anderen Windkanalen. FB I506. 2. Gothert, B./Richtor, G- : Messungen am Profil NACA OOI5-6U Im Hochgeschwindigkeitsvindkanfll der DVX (2,7 m ^) , Druckverteilungs- auBwertung. FB 12U7. DVX-Industriebericht: Hochgeschwindlgkeitamessurigen am Heinkel- Profil 00 12 - 0,715 36,6 im WL-Hochgeachvindigkeltswindkacal (2,7 m ^) . Editor: B- Gothert. Commissioned by Firma Ernst Heinkel Flugzeugwerke G-m.b.H., Rostock. 3. DVX-Industriebericht: Hochgeschwindlgkeitsmessungen am Messerschmitt- Profile Me 2 35 I3 - 1,1 30 im DVX -Hoobgeschvindigkeitgyindk anal (2,7 m ^) . Editor: B. Gothert. Commissioned by Messerschmitt- A. G. , Augsberg. k. Ferri, A.: Alcuni rlsultati sperimentali riguardantl profili alari provatl alia galleria ultrasonora di Guidonia. 5- Doetsch, H: Untersuch\ing der Profilreibe NACA 00 im 5 x 7 m Vindkanal der DVX. FB 91^. 6. Gothert, B-: Plane and Three Dimensional Flow at High Subsonic Speeds. NACA TM 1105, I9I+6. 7. Winter, H. : Stromungsvorgange an Flatten und profilierten Korpern bei kleinen Spannweiten. Forechung euf dem Gebiete des IngenletirweBenB, Bd. 6(1935), Ausg. A, p. U5, figure I6. NACA TM 121+0 15 ^ ^ ^ >- >- i: ^ ^ ^ <^ «:i «::i ^ ^ ^ I 16 NACA TiM 12kO /f \ MA 00 OS-/./ S8 > \ ,1 /J- 1 'A dCa / A \PrandU da^ t_fi_j j-o- ^ ^ - Crltical Mach nimber for c„ r M* * _ 1 1 1 KUm/B-I.IJS 1 fWi / A PrandiL Y A — a-o- ^^ V 1 » i m^ 000 /£'/./ 30 1 ' dCf 1 1 f- / iniUL A 1 i Jt i^ r ^^ ^ ^ 1 I 4 \m^ \ 0.Z 0.4 I^AU 00 /5 - // 30 ^ r-^ -^Pra nd(l J \ td \ •/ \ 0.6 f ^^ ^ f.o (Ifi" ^^ f /fcSf^^ 00/8-// 30 A _H ' 1 ^j?^ 2/7^// U-fl — ; — ~^ r *y° \ /7 A J tf^) 1 1 Mach number M = 0.2 6 / 1 ^ ► ' :^-4= =4= ^ 4 1 \ 1 1 1 1 1 |^*-j. ^ Large wind tunnel DVL (5 x 7ni) (FB 914) 2 ■ 1 1 1 _ Thickness ratio d/t _ ' 1 0.04 O.fi 0.f6 0.20 0.24 Figure 13,- Lift increase Sc„ /^ in two-dimensional flow on profiles of different d/1 (NACA series 000 d/1 - 1.1 30) according to measurements in the high-speed tunnel (pressure distribution measurements) and in the big tunnel (weighing). 0.08 0.06 I / • DVL - High-speed wind tunne o I (2.7m P) / / 0.04 Mach nuniber M = C >.2 « / /o "m,, > (tall heav t/4 y) ^ ) 0.02 _^y ^ - - ""^ < 1 ^ ^ f Thlc kness rat. if* c\ /f ^ 0.04 0$ f2 f6 20 0.24 Figure 14.- Neutral stability point position >)Q. /^c for profiles of different d/l (NACA series 000 d/1 - 1.1 30) according to pressure distribution measurements in the high-speed timnel and weighing in the big tunnel. NACA TM 1214-0 23 da — -1 1 1 1 ' increase according 1 to t /> 1 / /' 1 ■% ^ / \\ / / A-e L=^ Ail.lL — \ - i^-<^-«-— tsi::! ' Mach number u « 2 0. * Qf as to Figure 15.- Effect of aspect ratio A on the lift increase SCg^/^ at high Mach numbers in the zero lift range (NACA series 00012 - 1.1 30). Figure 16.- Lift coefficient c^^ plotted against a for a rectangular wing of A= 1.15 and NACA airfoil section 00 012 - 1.30 at different Mach numbers M. 2k NACA TM 22kO ?Jl 0.U m m &- - ir t crease 1 1 1 1 1 1 accordiirr to theorv A = 1 1.15 - ^ . ' __\- ■ — "■ ° Uach nuni)er m 93 n* 0.5 0.6 07 o.t 0.9 1.9 Figure 17.- Lift increase sc„ th 0; tail heavy, referred to I /4 line). NACA TM 1240 25 «« er y V t* I.S «* «/ en CI t* Figure 20.- Drag coefficient c^ at zero lift plotted against Mach number for several model wings of different aspect ratios A and angle (p ; profiles of all models measured in flight direction are NACA 00012 - 1.1 30. 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