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r r I '.9 June 1945 as Advance Confidential Report L5E21 COMPLETE TABULATION IN TE UNITED STATES OF TESTS OF 24 AIRFOILS AT HIGH MACH NUMBERS (Derived from Interrupted Work at Guidonia, Italy in the 1.31- by 1.74-Foot High-Speed Tunnel) By Antonio Ferri Langley Memorial Aeronautical Laboratory Langley Fielc, Va. WASHINGTON 'ACA 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. L 143 DOCUMENTS DEPARTMENT I. VAt4 3 .L AACR No. L5E21 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS N WARllTIME REPORT ORIGINALLY ISSUED S' hi r fc^. ., Digitized by the Iilernei Archive in 2011 with lunding Irom University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation hlip: www.archive.org details compleledlabulal00unil /3' ?# '' NACA ACR No. L5E21 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE COM'-,.TDEUTI AL REPORT COMPLETED TABULATION IN THE UNITED STATES OF TESTS OF 24 AIRFOILS AT HIGH MACH NUMBERS (Derived from Interrupted Work at Guidonia, Italy in the 1.51- by 1.7L-Foot High-Speed Tunnel) By Antonio Ferri SUMMARY Two-dimensional data for 24 airfoil sections tested in the 1.51- by 1.74-foot high-speed tunnel at Guidonia, Italy, are presented. The test Mach numbers ranged from 0.40 to 0.94 and the test Reynolds numbers from 530,000 to 420,000. The results indicate that thickness ratio is the dominating shape parameter at very high Mach numbers and that important aerodynamic advantages are to be gained by using the thinnest possible sections. The results of preliminary tests made to investigate the effects of jet boundaries, Reynolds number, and humidity at very high speeds are also presented. It was found that the jet-boundary effects became very large at high Mach numbers when models large with respect to the tunnel height were used. In the absence of suitable correction factors for large models it was considered essential to use models small enough to make the jet- boundary effects negligible. It was indicated that the data presented for the 2L airfoils tested are essentially free from jet-boundary and humidity effects. INTRODUCTION The rapid increase in airplane speeds during the past 5 years has greatly accentuated the need for experimental data in the subsonic Mach number range above 0.7. Experimental aerodynamic data in this speed range, however, are still very scarce. There are two principal reasons for the lack of data. First, the experimental equipment required to obtain data at high NACA ACR No. L5E21 sr-es on models of significant size is extremely costly '., ojnctruct and operate. Second, the problems of tech- nique involved in obtaining data at these speeds are very complex and are not yet fully understood. The tunnel- wall-effect phenomena occurring at very high Mach numbers with the presence of shock waves become so complex that there see:ns little hope at present of obtaining correc- tions for these effects by analytical methods. The principal purpose of this report is to present aerodynamic data for 19 related airfoils and for 5 miscellaneous airfoils at Mach numbers in the range 0.40 to 0.9L. The data were obtained on models of 1.575-inch and 1.565-inch chord in the 1.51- by 1.74-foot high-speed tunnel at Guidonia, Italy. Before the presentation of the test results, a description is given of the equipment used and the findings of preliminary tests made in an attempt to develop a suitable testing technique and to determine the isolated effects of such experimental varia- bles as Reynolds number, ratio of the size of the model to the size of the tunnel, and humidity. The results presented herein represent the completed part of a broad high-speed research program at Guidonia, which was interrupted by the war. I. EFFECTS OF REYNOLDS INURPBER, JET BOUNDARIES, AND HUMIDITY IN TESTS OF AIRFOILS AT HIGH SPEEDS A systematic study of the effects of Reynolds number, air-stream boundaries, and humidity at high speeds was made prior to the main part of the present investigation. It is not certain, of course, that these are the only factors affecting the results, but they are considered the most important. WIND TUNNEL All the tests were made in the high-speed tunnel at Guidonia (reference 1), a single-return tunnel that could function at a pressure below atmospheric. The pressure in the test section of the tunnel could be varied from 1.0 atmosphere to 0.1 atmosphere. The tunnel had a CONFIDENTIAL CONFIDENTIAL NACA ACR No. L5E21 system of refrigeration by which the temperature at low speeds could be held constant at as low a value as 150 Centigrade. The temperature of the air as it left the compressor was very variable, depending on the velocity and the pressure of the jet. The tunnel was Dowered by a 3000-horsepower fourteen- stage axial-flow compressor, which could produce a velocity ranging from 0.4 to 2.9 times the speed of sound when one minimum retangular section of the jet 1.31 by 1.74 feet in size was used. In tests at subsonic speeds the test section of the jet was kept constant at these dimensions. The jet was enclosed between two straight, parallel side walls, which were perpendicular to the axis of the model. The jet was not restrained by top and bottom walls. (See fig. 1.) The effuser A-A was shaped in such a way as to give a uniform flow at the plane a-a. This uniform flow was attained in a series of preliminary tests by increasing the length of the parallel-sided effuser until satisfactory flow distribution was obtained. The diffuser 3--L ," placed in a position to give uniform flow arnd to e '..i--.i e the vibrations that tended to occur. With the d'iffu3.- -'pe and location finally determined, the veloc::ty W.aSs constant along the plane b-b in the test' se,:;-L:.n of the tunnel even at the highest speeds. By varying the position and the dimensions of the diffuser a stable .end uificrm flow could be obtained even in the Mach nin.itEr rd I.& approaching and exceeding the speed of sound (I'.ach n... t-:-s of 0.9 to 1.2). The present test program incluo 3.d lie.itsusreients made at Mach numbers up to 0.94. Informr'- _'on on the shape and location of the diffuser has been lost; therefore, the exact dimensions of this setup are not available. The velocity and the TMach number were determined from a tunnel calibration basid on measurements of the total pressure in the larce section of the tunnel ahead of the entrance cone sn.d .n 'r.-rs..c-emtLents of the static pressure at the wali neor tL ..-it of the entrance cone. In order to check the e''.c-rity ensureded in this manner, pitot-static cubes vfere ii?,t.i.lled at the top and bottom of the jet just downrt-~enz of the exit of the entrance cone. These tube.? -. q,:-lli-tative indication of the jet-bcun-dary interf- .:wre L.'t "s. s hen the velocities measured by these tubes WLtr appreciably different from the velocity indicated by the entrance-cone pressure CONFIDENTIAL CONFIDENTIAL NACA ACR No. L5E21 calibration, it was usually fund that the interference effects were so large that they appreciably altered the aerodyrna.Tic characteristics of the test models. No .ata were taken when this condition existed. EFFECT OF REYNGOLDS NTJIBER AND AIR-STREM BOUNDkARIES Experimental methods.- In the study of Reynolds number effects at high speed, preliminary tests were made first on cylinders and spheres of various dimensions (reference 2). An analogous series of preliminary tests was then made for airfoils. l:,odels of airfoils of con- stant profile bit of varying chord were tested. For the study of the effect of the air-stream boundaries, tests were made with varying ratio of model chord to tunnel height over a range of Miach numbers. The ratios used were: 0.0755, 0.09~2, 0.115, and 0.151. The Reynolds number at each M1ach number was held approxi- mately constant by varying the density. Test models.- A profile was chosen having an arc for the upper surface and a straight line for the lower surface because this profile could be exactly reproduced in various sizes. The rnoer surface could be made by use of a lathe and the lower surface could be formed by use of a shaper. The leading edge and the trailing edge were sharp. The maximum thickness chosen was 8 percent, Lnd the profile was designated C-8 (fig. 2). Four models were constructed with such a profile; three with chords of 1.575, 1.969, and 2.362 inches (L, 5, and 6 cm) for force tests and one with a chord of 3.15 inches (8 cm) for detenrining the pressure distribution along the pro- file. Tests and results.- At Mach numbers of 0.4, 0.5, 0.6, 0.7, 0.F, and 0.9, the lift coefficient, the drag coefficient, and the pitching-moment coefficient about the quarter-chord point of the airfoil were determined for the three profiles having chords of 1.575, 1.969, and 2.562 inches. All the models were tested at two Reynolds numbers: approximately 250,000 and 840,000. The model with the 1.575-inch chord was also tested at a Reynolds number of 150,000. For the profile having a chord of 3.15 inches, pressure readings were made at angles of attack between -5.5 and 5..50 for Mach numbers CONFIDENTIAL CONIDENTIAL NACA ACR No. L5E21 of approximately 0.7, 0.8, and 0.9. Values of lift and of pitching moment were obtained from the pressure distri- butions. Force-test results are shown in figures 3 to S. In figures 9 to 11 the results of pressure measurements are presented. Figure 12 shows the results obtained from integration of the pressure diagrams compared with the results obtained by use of the balance. Reynolds number effects.- The results of the pre- liminary tests of cylinders and spheres showed that for the range of Reynolds numbers covered in the tests the effect of Reynolds number decreased as the velocity increased. At Mach numbers close to 1.0 there was vir- tually no Reynolds number effect. In the airfoil tests the importance of Reynolds number was considerable at low Mach numbers and the effect of Reynolds number was noted up to the critical Mach numbers at which the phe- nomenon of shock began to appear (figs. 5 to 6). For supercritical Mach numbers, the effect of Reynolds number became less until it virtually disappeared for Mach num- bers very near 1.0. In this range the formation of shock waves seems to control the aerodynamic phenomena and the development of the boundary layer. The boundary-layer thickness probably depends to a large extent on the angle of deviation of the air as it passes through the shock wave. The friction drag is a reduced nart of the total drag and, therefore, the Reynolds number effect is small. The Reynolds number, however, could have an effect on the characteristics of the shock wave itself through its action on the boundary layer, but such an effect is not indicated. In general, these airfoil test results con- firmed the results of the sphere tests. Large-scale tunnel tests made at the Deutsche Versuchsanstalt fur Luftfahrt (the DVL) in Germany and flight tests made at various times showed similar results. Effect of air-stream boundaries.- The jet-boundary effects for the ratios of chord to jet height of 0.0755 to 0.115 covered in these tests appear to be negligible. Essentially equivalent results were obtained at a given Reynolds number for all values of the ratios employed in the tests. For a larger jet-boundary effect, a test was made of the model with a chord of 3.15 inches for which the ratio of the chord of the model to the height of the air stream (0.151) is twice that normally used in the tests. From the results of integration (fig. 12) the values obtained for CL and Cmc/A are seen to CONFIDENTIAL CONFIDENTIAL NACA ACR C No. L5E21 coincide at high Mach numbers with the values found by the force tests. This agreement indicates that the boundaries of the air stream probably aid not interfere appreciably with the distribution of the pressures. For a I'ch number of 0.9' the effect of the air-stream boundaries is important for the model of 3.15-inch chord but is not important for the models of 1.575- and 1.969-inch chord. For higher Mach numbers the boundaries also affected the results obtained with the two smaller models. It is interesting to note that the phenomenon of choking of the air strewn, which occurs in closed-throat wind tunnels at high speeds (reference 5). did not occur in the tunnel in which the present tests were made. For example, for model C-6, which had a chord of 3.15 inches, it is estimated that choking in a closed-throat tunnel would occur at a Mach number of 0.88 or lower. The choking Mach number for the 2.552-inch-chord model is estimated to be 0.90 or lower. These choking Moach num- bers were calculated from one-dimensional theory for the zero-lift condition. They are therefore somewhat higher than the choking Ma-ch numbers that would actually be obtained, especially for angles of attack other than that for zero lift. In the present tests it was possible to obtain data for these models at Mach numbers as high as 0.94, and the results of the jet-boundary-effect tests indicate that the data are essentially free from tunnel- wall effects at this Mach number. EFFECT OF HUMIDITY ArD CONDEIISATION The air becomes very cold in the expansion that occurs in the tunnel at high speeds. (The process is very nearly aciaoatic.) Total condensation may occur in the whole jet at high seeds if the dew point is passed. Even if condensation does not occur in the jet, there is a possibility of its occurring in the low pressure regions over the test model where an additional expansion and temperature drop occur. Very low local temperatures, which are usually smaller than the local dew point, are found at high subsonic speeds; local condensation there- fore could occur and could produce a "condensation shock" or a localized region in which condensation occurs. CONFIDENT AL CONFIDENTIAL NACA ACR No. L5E21 Condensation complicates and modifies the flow over the body because it alters the values of the temperature, the pressure, and the speed in the air stream and, hence, modifies the values of the resultant aerodynamic forces. A complete examination of the effects of the phenomenon of condensation shock is very complicated. The variables involved include the value of the local humidity, the speed of the condensation, the possibility of the exist- ence of supersaturated air, and the scale of the model. The condensation process is not instantaneous but requires a finite time and its beginning may depend on such factors as the nuclei of condensation. (The super- saturated air may sometimes exist for a time at a tem- Derature much lower than the critical.) If the tests are made at small scale, the air can pass through the low temperature region in so short a time that appreciable condensation does not occur. Condensation is therefore less likely to occur in small-scale tests than in large- scale tests. In flight, for example, when appreciable relative humidity is present, condensation normally occurs and is easily seen on propellers and wings in high-speed dives. Since the characteristics of the con- densation vary with scale, it would appear to be practi- cally impossible to simulate full-scale conditions in tests in which small models are employed. The problem is further complicated because the degrees of supersaturation existing in the tests in a wind tunnel may be different from in flight and the beginning of the condensation depends on certain variable conditions of the air. The condensation characteristics of different wind tunnels, even with the same setuo, have in several instances been noted to be widely different. In the subsonic tunnel of the Aerodynamische Versuchsanstalt (the AVA) at Gottingen, for example, it is normally necessary to dry the air before it converges in the test section to prevent con- densation; however, in the Langley 24-inch high-spted tunnel, which has a comparable entrance-cone shape and which operates under similar conditions, it is not nec- essary to dry the air, and complete condensation seldom occurs for relative humidities below 60 percent. All the test data obtained up to the present time tend to indicate that even for large-scale models the effects of humidity are of secondary importance provided that the percentage of humidity is low. In the Guidonia high-speed tunnel previously described, it was very dif- ficult to study humidity effects because of an automatic CONFIDENTIAL CONFIDENTIAL 8 CONFTDENTTAL IThCA ACR No. L5E21 drying up of the air which took place. A small quantity of water '.as removed from the tunnel air by the pump which was used to evacuate the tunnel to the low initial ores- sure. The condensation that occurred when the tunnel was started was believed to cause water to collect behind the test section and to adhere to the tunnel walls. As a result of this automatic water removal, fog did not occur in lhe test section even at supersonic velocities and no air-crying equipment was necessary. Because the humidity be'rame less during the progress of a test in this tunnel, it was impossible to give precise results as to the effect of humidity, but the general indication of the data that have been obtained was that Ehe humidity effects were not appreciable, at least not for the small-scale models tested. Tests to study the effects of humidity have been conducted in the 8.?6-fjot high-speed tunnel of the DVL in Gernany using an N.CA 0015-64 airfoil section with a 1.64-foot chord. In this wind tunnel the amount of condensation existing in the test section can be con- trolled by varying the cooling of the tunnel and thus regulating the temperature of the air in the test section. For very high values of relative humidity, it is necessary to eliminate the cooling entirely in order to raise the temperature enough to avoid condensation. The results of the humidity-effect investigation in the DVL tunnel dem- onstrated that, even for the relatively large-scale model employed, the humidity effects were of secondary impor- tance when the relative humidity was small. In order to indicate the conditions under which con- densation might occur in flight, figure 13 is presented showing the local Mach number as a function of the flight Mach number for which the conditions required for satura- tion are reached. (hdiabatic expansion of the air from its static condition to the conditions corresponding to local Mach number is assumed.) Also shown in figure 13 are the values of maximum local Mach number that are attained locally on two typical airfoils. The data cal- culated for the NACA 25015 airfoil (unpublished) were obtained from tests made in the Langley 2L-inch high- speed tunnel. The data for uhe NACA 0015-64 airfoil were obtained from the DVL tests mentioned previously. Fig- ure 15 indicates that, even for very low values of the relative humidity, local ,ach numbers are obtained at which condensation is possible when the flight Mach num- ber is 0.6 or greater. CONFIDENTIAL NACA ACR No. L5E21 The discussion in the preceding paragraphs has shown that humidity effects are likely to be most pronounced under large-scale conditions. Systematic tests to deter- mine humidity effects could best be made in a large-scale wind tunnel in which the temperature of the circulating air could be varied by regulating the cooling. The tests in such a wind tunnel could be made at various periods in order to cover a wide range of relative humidities. Fig- ure 14 has been prepared to indicate the conditions for saturation in the test section of a wind tunnel for three values of relative humidity and for various temperatures of the air in the entrance cone of the wind tunnel where the airspeed is low. Also shown in figure 14 is a com- parison of the maximum local Mach numbers of the NACA 23015 and 0015-64 airfoils as functions of the stream Mach num- ber to determine at wnat Mach number the conditions for saturation are locally reached. The figure shows that, for high relative humidity, it is necessary to have a high temperature of the tunnel air stream in order to eliminate condensation in the test section. It is also shown that, even if condensation is eliminated in the test section, the necessary conditions for the formation of local condensation over the test model will normally be attained. COMPARISON OF TEST RESULTS FROM VARIOUS* 'IND TUNNELS AND FROM FLIGHT Airfoil tests.- For a thorough examination of the accuracy .nd significance of the test results obtained in a given wind tunnel, it is essential that the results be compared with those obtained in other wind tunnels and in free flight on models of similar orofile. As a step in this direction, tests were conducted on the NACA 0015-64 airfoil in both the Guidonia 1.51- by 1.74-foot rectangular high-speed tunnel and in the DVL 8.86-foot-diameter high-speed tunnel, which has closed circular walls. The model used had a rectangular plan form enclosed between two end plates. The chord of the model was 1.658 feet (50 cm), the span was L.5 feet, and the end plates were 25.6 by 43.2 inches. The ratio of the model chord to the tunnel diameter was 0.185. With this setup, the choking Mach number was about 0.86, which is considerably higher than the choking Mach number that would have been obtained with the model completely CONFIDENTIAL CONFIDENTIAL NACA ACR No. L5E21 spanning the tunnel jet. The data obtained in these tests consisted of pressure distributions and wake sur- ve ys. The test conditions were adjusted to produce an equivalent relative humidity of the air of 20 percent at sea level. The Reynolds number varied with the Mach num- ber from about 5,800,000 to 6, .00,000 in the high-speed range of the tests. The model tested at Guidonia had the same profile but was of much smaller scale, the model chord being 1.575 inches (4 cm) and the ratio of model chord to tun- nel depth being 0.0755. The relative humidity in the Guidonia tests was always very low. The Reynolds num- bers were, of course, very much lbwer than those of. the DVL tests and varied around a value of about 500,000. Force measurements of lift, drag, and moment were made in the Guidonia tests; pressure-distribi.tion and wake- drag measurements were made in the DVL tests. The results obtained are compared in figures 15 to 17. Figure 18 shows pressure-distribution measurements made at the DVL for one angle of attack, a = -0.250. It may be noted that the results from the tvo tunnels are at variance, especially at high speeds. This lack of agree- n,-nt indicates that the testing technique and the pro- portions of the testing system are of great importance in high-speed wind-tunnel work. The differences in the drag-coefficient values at low Mach numbers are probably due to the difference in Reynolds numbers. The largest differences between the results from the two tunnels ure in the crag and pitching- moment coefficients at high f.ach numbers. The abrupt changes in the coefficients from the DVL tests at Mach numbers in the vicinity of 0.8 arc probably associated with the phenomenon of choking, and the results obtained in this range are therefore considered extremely ques- tionable. Because of the much smaller relative size of the model in the Guidonia tests and also because of the fact that the jet was not restrained by top and bottom walls, similar effects did not 3ccur. Further tests were made at the DVL tunnel of a smaller model of the same profile having a chord of 1.148 feet, the model-chord to tunnel-dianeter ratio being 0.15. The results obtained with the smaller model are shown in figure 16. It will be noted that the rate of drag rise past the critical speed is appreciably less than with the larger model and CONM'IDENTIAL CONFIDENTIAL NACA ACR ITo. L5E21 thus is in better agreement with the results of the Guidonia tests. Free-flight tesus were in the general research pro- gram at Guidonia, but they were interrupted by the w-vr. The few flight tests made, however, indicated that the drag-coefficient curves had about the same slpe's at supercritical speeds as wsre obtained in the Guidonia wind-tunnel tests. Bomb tests.- additional comparisons between high- speed winTd-5unnel and flight data were obtained in tests of an airplane bomb of conventional shap_. The approxi- mate shape of the bo.mb is indicated in figures 1i and 20, v,hicli show the results of the tests. The bomb .'as launched in flight at an altitude of 59.5.00 feet, and its trajectory as a function of ti'me w-;s recorded v.ith a phototheodolite. The sceed, the [Ki.ch tLnum/er, the accel- eration, and the drag. coefficient :were obtained from the trajectory data. A one-third scale rmodel ft this bomb was tested in the PVL C .6-foot-d.iLmeter high-speed tun- nel (the ratio of bomb diameter to tunnel diameter was 0.01'55, much lower than that normally; used). A one- tenth scale model cf the sane bomb was tested in the 3Giidonia 1.31- by 1.74-foot rCectangular high-speed tun- nel using a ratio of model diameter to air-streamn height of abbut 0.071L. Sii;ilar tests were .mcie in a .jiid tun- nel at the AV,\~ in Coctir.gen, which has a partly free air stream similar to that at SuiLonia but 47 inches high. The size of the model used in these tests is int known, but it is believed that the ratio of model diameter to tunnel air-stre- i height was considerably higher than that used in the tests in the other wind tun..Iels. The results shown in figI Lr 19 indicate reasonably ,Dod agreement in the frnn of the drag curves obtained. As might be expected. how.Jever, the drag-coefficiert valLes obtained at very high i:ach numbers in the closed DVL tun- nel are higher than those found at Guio-nia in the relatively unrestricted jet. The results obtained in a subsequent launching Df the bomb, with reinforcements to the tail structure, in flight: tests at the DVL are shovn in figure 20. C':jPFIDE:N'IAL CONr'FI .ETNTIAL NCCA aCR No. L5E21 CONCLUSIONS The following conclusions were drawn from the investi- gati-on of the effects of ieynolds number, air-stream boundaries, and humidity in tests of airfoils at high speeds: 1. It has been shown that the ratio of tunnel height to model size, the form of the test section, and the testing technique have a very great bearing on the results obtained at subsonic Mach numbers above 0.7. 2. Reynolds number effects were of secondary -mpor- tance at very high Mach numbers for the range investi- gated. 3. In the absence of suitable correction factors, the only safe experimental technique consists in keeping the scale of the model small enough so tnat the correc- tions required are negligible. I. In a closed air stream, the model must be small enough that the highest desired test M1acn number is below the choking Mach number ?f the tunnel, at which the effects of the tunnel walls on the flow over the model become extremely large. 5. By use of a jet which is not restrained by top an. bottom walls the maximum rMdch number that can be used for a given value of the ratio of jet height to model chord is appreciably higher than the value that can be obtained in a closed jet. 6. The considerations of condensation phenomena that have been discussed have brought out the fact that the conditions under which condensation occurs depend on many variables and that only with great difficulty could flight conditions be simulated in wind-tunnel tests in which small-scale models are used. Vyind-tunnel tests should be conducted with low values of the relative humidity, because under such conditions the effects of condensation are known t) be negligibly small. CONFIDENTIAL CONFIDENTIAL TACA ACR No. L5E21 rT7. T r"7' .ErT ,''c IP. ?!L ATPFf .1IG ; T1 T T T'A'T-. il"E'qP" 'AM~IF'F 03T 1 O1.0 0 c 1. .KPF 4 .i ..S ;.'T' l'ET nD? Tr.enty-four profiles vwc-re tested' in th.e 1 .51 .y 1 .7!.-foo t hi 'ih-s peed tun.tel at ui.'._crna with the ,-rtly free te"t sect ion rr-vious'ly ScrTibd. T-r ever-r r r C- file the ift, the Crs. ,, nm. ch ,s ,tchic n- ,i"c:t 5t t the uairter-c!,ord jcintc. re v .neasuAred bL,'c Le CI th tree- como:r eni .sent- l l.Lic b al ni; 'escrl e in r. fer.en Tn extre: -ities cf tie models wIere Et":.ed at t-e balance suppoCrts 3!and tihe models "ere .: Lcl:ed durinr tihe tcsts to verify that the aerod;*,nrmo: c lo r s .'.*:LC r. c bend thi.em appreciably. AT tih tests ,e,.. r-e- pc r.td vi th t.e :no'.cel inverted. Fr.r some models, the tests were rer, _,ted later when the static atmospheric con.'i-ti.ns v..re c.cio-lEtely different and '.:ith different hu.-iditli s in the test sec- tion. (The values of tha -elat '.ve humidity v.ere sli'-.'vs lo";.) The differences i: L-e rE.su'.lts o-ts r.ed '-Ere not cppreciabl. . All the models v.'ere mnde i.f -roll. -rl' Ced steel end had chord:..3 cf 1 .75 inch.-.s for th -ic.: 'e.r "s tic~ c rcf r- c- nt or .r,:-oter Tn order 'o r:'~'ert exc:s-s i v: ten'ii. , the rmoccls vi th t-ic-.:nCss rating of l'sE th-a3 per.nt hvd chlords of I .ar.' In-ches Tl-e 'crefil L-P of -nmall model, s.lori ,L rre.s rnd exactly W, :ith the profile d.sir:ed. Fr u.-.-sc c. e.cu- racv, ther -roi'cre an onpical device E'3 constu:td iNt": t per'.nitttd photo,'ap;in,, v' t exLtrc.re "i cisi cr the true sccti.or of ea.Sch mo.,del ( n 3 .re-a-' ., ncr aCscid sc'l- e L ;-'or each model two end sections v'cre U. -Oap.ed and Th3 trie n;rofile v's projected on che :h' o tc.rjph to nrov'dc the desired cno:-orison. B'cu.s -the ar -f.oil were con- structed by riachine, -he .ronf"il.e s'-pe _did not v.ry a ;ross the model soan. This fact ."as corfi -r-f d br su:permir:-;oosir.p drwi'-.'ins of the t:'.o cr,: sections. It v es 'e-r filed that the surLi'ace V.w s c.dequ".tcly cnc).t Ly r mL, v r.ir there t n- gential illLuminated surface r'.nder .'reat ma.'nification. Figure 21 shows tlje specifid .:hc.pes of the profiles tested. In figure 22 the cattua.l shp.-E cf the proIiles tested are compared with the specific. e.] shapes. In crder CO 1.r'TDEINTIiL *n 71 T 'rrAT ~ ACA CR ;:o. L5Z21 that the diffeo:enca between the actual aid thi specified prt?1 les say be cleq4lz seen, the ordin.- te scale used in ft -ure 2d ha3 been eiilar.:ed. Trvble T sh)ws the ordinates of t-Phe orfiles tested. 1.11 the terts wcie o-erformed at an sa.roiximately constant ReynDldcs nui:ber varying in the rare from 5LC,000 to a2C,032'. i'he density, and consequently the eynmolds n'u7her, i-d to be kept low for the thinner airfuils in nrder to oreve:-t excessive loadc. .ATRFl'rL TESTED The profiles listed in the fcllcvwing table vere tested: I -- -- -- 5 :'--- iiACA 0012-6L 2 . IA A I 1OC-6L' 5 "N: PT ' :'ACA 'o,5-6L 5 .I. (L,. Cer: -AA 00o3-gL 5 I '.s. (c, per :hcA C0012-. L U.S.I 10 pe I' 6 lFer( ;IACA 0009-t.6 5 1; TO Q 1 per ;ACA 2506 1 2A ,I (,' perc t;- CA 2512 6 E I rf -'" i I "I-ference A. 515 6 III 2L 1 67 ::.F. 1 7 cent thick) U. P.3 2 cent thick) recent thick) iCA Ml cent thick) :n C006T A 2509 cent thick) rTH69? (a) aDevelcoped at Zurich '.i-versity. Fr T T :T7 : T~I C:0C.) ID-I'ETIAL NACA ACR ;o. L5E21 RESULTS In figures 23 to 46 the results of the tests of the 2l. airfoils are shown in the form of the usual coefii- cients: CL and CD are plotted against the test i,.scl number at ths same angles of attack, and Cmc. /L is plotted against the P.!ach number at r.slues cf CL -orre- spanding to the given angles of attack. Figures lh7 to 70 show a, CD, and Cm-/. plotted against the corre- sponding CL for ecch airfoil rt the- samr ;."Szh n,.Lmbers. In figure 71 the anrle of zero lifL is pitted against Mach number for representative i.irfoils of t-ie group. Figure 72 gives th3 r.:eximum lift-irag ratio (L/D)mnax a for (L/E)max, and CL for (L 'max as funztio.ns of Mach niurber for all the airfcils tested. Figures 75 and 7 present CDmi. and (L/D)max as functions of the maximum percentage thickness for all the airfoils at various Mach numbers, and figures 75 and 76 show C.in and (L/D)jmx plotted against .n.ach number for several groups of airfoils havin- the same maximum thickness. It can be observed from the test results that: The lift-coefficient curve as a function of Mach number presents a ma:-irrmun and later a minirmum value. The Mach numbers at these values can be defined as the first and the second critical iM.ach nur.bers f-r the lift. The M:ach number at .which the drag-2oeffic'rent curve abruptly bends upward is defined herein Gs the critIcal Mach number for the drag. It will be noted that the critical Mach numbers as defined herein ire different for the lift and for the drag drt.. The critical ach nu-mbers used, furthorrore, do not n-cessarily correspond to the streak '.:sch number Et wviich local sonic velocity is reached. The rate of drag rise past the crnit'-cil '.':h nurbcr increases as the lift coefficient, th; anr-le jf ottclk, and the thickness ratio are increased. The first critic-. :Mach n.u.ber for CL and the critical Mach n-umber for CD for each airfoil is lowered with the increase in angle of attack. CONFIDENTIAL CO:i-IDEi]TIAL NACA ACR No. L5E2l For each series of airfoils at the same angles of attack, these critical Mach numbers decrease as the thickness increases. The critical Mach numbers at the same thinness and the same angle of attack are much lower for the cambered profiles than for the symmetrical profiles tested at the same angle of attack. At equal thickness and equal camber, the critical Mach numbers are higher where the maximum thickness was at the LO-percent-chord station than where it was at the 50-percent-chord station. Above the critical Mach numbers, the drag increases and the lift decreases very rapidly; for a profile with a larger thickness and sharper curvature, the increase in drag and the decrease in lift is sharper. These general phenomena agree vjith results of other laboratories. (See, for example, reference L.) Lift.- At subsonic Mach numbers the increase in lift coefficient with Mach number follows approximately the theoretical relation 1 especially for the low thickn-jess ratios. After the first critical Mach number is reached, the lift coefficient decreases very rapidly until it reaches a minimum at the second critical Mach number when it again starts to increase. This second critical Mach number is lowered with the decrease in the first critical Mach number. Airfoils with larger camber had greater decreases in lift. For these airfoils, generally, the angle of attack fr zero lift changed greatly and tended toward positive values (fig. 71). At the hi* hest test Mach numbers all the wings functioned in a manner very similar to symrr.etrical profiles. This phenomenon agrees with the fact th&t the value of the angle of zero lift for an unsy.rnrLrical profile changes sign and becomes considerably reduced in magnitude in passing from a subsonic to a supersonic velocity. The lift-curve slope dCL/da increases up to the first critical Mach number after which there is a con- siderable decrease up tn the second critical Mach number. (See figs. 57 to 70.) The second critical Mach number is greatly affected by the value of the maximum per- contage thickness. For the lover thicknesses tested, CONFIDENTIAL COF 0 TF DEFTI AL NACA ACR V:o. L5E21 the second critical '.Mach number was reached only at the maxinumr speed of th_ test. 'Morent.- The carve of pitching-.:noment coeffic-i.nt against Alach number has a fairly regular form (figs. 35 to 46). Generally, the value of Cm/i remains constant up to the first critical Mach nurrber and then tends to decrease. For the larger thickness ratios there is Tn increase in Crnci, at the first critical ; Pech number, and it appears thet the center of pressure rm.oves forward. Vihen the Mach number is increased beyond the critical value, C,. decreases antil it re.cehes a n.ini:mu and then tcnds to increase. The center of pressure moves ap-reciably at low values of Cy for profiles of large csrber. For sym- metrical profiles the -artations of C c/1 t low values of CL are small, especially if the maximum thickness is about .40 percent of the chord. Drsg and lift-draF ratio.- The value of (L/D)max decreased rapidly beyond the first critical Mach number for the lift and continued to decrease until the second critical Mach number was reached. It then varied very slo'wlyv with further increase in Msch nurrber (fiT 72). The larger thicknesses suffer more pronounced relative changes in (L/D)max. The ongle of attack and the lift coefficient corresponding to the (L'D)ra. (fig. 72) decrease as far as the second critical point and then begin to increase rapidly. The variations are -aprecisbly influenced b', the value of maxi'Tmum thickness, ratio asnd by the mean camber. .n order to empha-ize the importance of rraximum per- centage thickness on the values of CD3n and (T./D),my, figures 75 and 7 v.ere prepared to show the values of these factors as f-nction of m].ximum rers;ntaseo t..i-2kness at constant Mo.ch number f)r eech series considered. These figures sn.o, that the? offe't cf the .ma.irmun percento..e thickness beccmes greater Es the iTL:h nu-rber increases for all profiles tested. For ?.nch numbers around 0.W, the effects of thickness ratio are very larce. At lo'v test speeds, for example, when the m3xirium thickness is COtF I DEFTIAL CONF0'I DEITIAL NACA ACR No. L5E21 varied from 6 to 12 percent, CD increases about 30 percent; at Mach numbers around 0.8 or greater, the increase becomes 200 to 00 percent. The r-tio of the values of (L/D)nax for the ;ACA 0006-31 airfoil to those for the NACA 0012-34 airfoil changes from about 1.18 at a Mach number of 0.65 to 4.5 at a Mach nur.ber of 0.85 and to 2.2 at a Mach number of 0.94. It is also interesting to compare the aerodynamic character- istics of various profiles at equal maximum percentage thicknesses. (See figs. 75 and 76.) The larger increments in C in occur for the pro- files with larger camber, for which the critical Mach number is lower. l.ith increase in the value of the maxi- mum percentage thickness, the value of dCDin /dM increases and even at very high Mach numbers this differ- ence Oetween various profiles is considerable. The profile shape has considerable effect on (L/D)max (fig. 76); the unsymmetrical profiles have larger (L/D)max values at low Mach numbers. ;:t nirher speeds, the symmet- rical profiles with the maximum thickness at about 40 per- cent of the chord had higher efficiencies than those with the maximum thickness at 30 percent of the chord. The difference between the various profile types is consid- erable for low Mach numbers: however, it decreases with increase in Mach number and is small for Mach numbers around 0.94. CO NC LUS TONS The following conclusions may be drawn from the results obtained from tests of 24 small-scale airfoils in the Guid-nia high-speed tunnel: 1. At subsonic Mach numbers both the profile shape and the thickness ratio had a large effect on the minimum drag coefficient. 2. Reducing the thickness ratio, moving the point of maximum thickness from 50 to 40 percent of the chord, and reducing the camber all tended to increase the critical Mach number. COnlIDET TIAL CONFIDENTIAL NACA ACR No. L5E21 5. Airfoils of lsrge percentage thickness shcild n3t be used at high I'ach nu.nbers because of the radical adverse changes in their characteristics at supercritical speeds. 1... "Then the critical speed was exceeded, the drug coefficients increased rapidly. Abrupt decreases in lift and changes in moment occurred at somewhat higher critical Mach numbers. c. The lift coefficient continued to decrease as the speed was advanced beyond the first critical M3ach number until a second critical .ch number vwas reached, bIeond which the lift coefficient inorcessd in v.lue. 6. At very high s upercritical Yach n'iumrbers the thickness rat o is the dr~inating variable, the drag coefficient being almost directly proportional to the thickness at a Tisch number of 0.9.4 rational Advisory Commnittee for Aeroneutiks Langley Memorial Aeronautical Laboratory Lanrgly Field, '.'s C DNFIDE7TI AL CO NFIDENTI A L NACA ACR No. L5E21 REFERENCES 1. Ferri, Antonio: La galleria ultrasonora di Guidonia. Atti di Guidonia No. 15, 1939. (Available in Air- craft Engineering, vol.XII, no. 140, Oct. 1940, pp 502-305.) 2. Ferri, Antonio: Influenza del numero di Reynolds ai grand numeri di Mach. Atti di Guidonia No. 67-68-69, 1942. (Available as R.T.P. Tr. No.1933, British Ministry of Aircraft Production.) 3. Byrne, Robert W.: Experimental Constriction Effects in High-Speed Wind Tunnels, NACA ACR No. L4L07a, 1914. 4. Ferri, Antonio: Investigations and Experiments in the Guidonia Supersonic Wind Tunnel. NACA TM No. 901, 1939. 5. Stack, John, and von Doenhoff, Albert E.: Tests of 16 Related Airfoils at High Speeds. NACA Rep. No. 492, 1934. 6. Jacobs,Eastman N., Ward, Kenneth E., and Pinkerton, Robert M.: The Characteristics of 78 Related Air- foil Sections from Tests.in the Variable-Density Wind Tunnel. NACA Rep. No. 460, 1933. 7. Jacobs, Eastman N., and Anderson, Raymond F.: Large- Scale Aerodynamic Characteristics of Airfoils as Tested in the Variable Density Wind Tunnel. NACA Rep. No. 352, 1930. CONFIDENTIAL CONFIDENTIAL NACA ACR No. L5E21 F--i 4.30.0. C 00 Cd "."l - >04403 .3I.0.3..0,-i .- 7 gfvcrjf\>nninr-r^^l-'u''" 00. 4^oou3o"j0.40.F-04.."o 7047L_3.1.04.10..-C.F..1U.040,f -i * a J.y r3-.4 0 _30- 0- *|a OO-.4,Poi.0.4'S L5 I rr-ccl--E r~ mm 0 ( -. o -. 1" T, .. , a.. o .- a --o .'IJ .... ....... -- 3 OUX 0 .-M^- '.''..... r .0-e-- o o 3N: . * S ,a .- .+ S o 074.-, 7.7 4,4.1. 0 .wr- Su o>.,l u .- -.. ..,:.... I ,. d OU o g .7 3! 0 3 oi Y..^^ .=0 ,, .- ,j. .-r-0o 4.... ..4 "r.,x.-.++ r +o '=,o U) 4 1-1 43-4 -4-. 3i7o'3 O r- 1 M403.C..i,. .ao .. 4 .U i. . _. 3.. .... a...h-c.0 ....00 0... ..3.. -- ----- -14 -------- 0 4- F l *j j 3 +.j --*j *i d7- 0 oT4 0 I: -r------s ---------- 3 . ..^~ -_-------------------- _---------------- I3. a 0 0. 3 7 O'3 -4 O-3i. C 4,-->l' 01 0.U 0 J1''_________________ .4 4 4 -i-- I- 4 0.4. 0 35 0.4 3"'* 41 ..0104 7'4F4443 7.01.04%*- OU. * 0 < 040'4.0.44. .30404.7 '; ._ .-.^ -^ -.. ^ L **'''***- ;0 33JC~ -S~ < 3U' ~- Ci~i^\Bti C=3 a ^i J 3T'Sr"""7--*-/'3" kj aiT O' ir-- Ii ~ c ^ l? T rr IT' rf< o *M" f P- 0 a "-ij'r^ ^-' /_- a Oi 3 ;------- -^ -- ^ --- L* l *l-" 0^ ^. 30 31 -3 3 '- -*-3J u* -- ** ~ I-3 (^ -. . S a -I I- I < 3 - L. ;1 a a '- j. z 0 U< I 1 n r 1 NACA ACR No. L5E21 I S*,'I u I z: 0 . Q I) (r)) K p~ Q 't, .^ ^ $ A) > ., h G< 9, ^ ^ I .^^ rt ` Fig. 1 NACA ACR No. L5E21 -, Jcc CL: 0 rE 8 I. LL u0 Q% I. I Fig. 2 -0 3E09----- NACA ACR No. L5E21 CONFIDENTIAL Air fo// Reynolds number, R o 575- ,, c. or/d z.o, 0oo 0 .' 96 C/, cord,- 25O, 000 S / 969' -, / c/7 c/0orC 480 000 L 362 inch chord 495 000 - --------------- o- ..J ic .QJ '4 -8 -6 -4 -2 0 2 4 6 8 6------i- -- - 4- --------- --- -P---- --- ------ --- --- --- ------ 2 nTiOAL ADVISORY COMMITTEE FOR AERONAUTICS CONFIDENTIAL -----------------4 I -.3 -2. -./ 0 ./ .3 C, F-0 ure 3. ;/I--, 0'rag--fO a;?~/-r e/-coe /5/e/-7 Sc/?aroc/e-/-sics o/ 0 C-a arfe// .ectio7 a A-= O 50 . I.' qI. B O---- ----------,-- .04 0 ===^========= ./2 ,cfq/e of o//oc/4, OC, doey ~airr7e~f coe f~C/;C~e~f/ Fig. 3 NACA ACR No. L5E21 .9?'/e of a 7fock C OC, )ce9 .4----- - .6----- ------ NAIIOIAL ADYMISORY OMM-ITTEE FOP AEROA utiK 74 CONFIDENTIAL --- -2 -/ -/ .2 .3 F 4u.e 4.--- -- d-o--d ,,,me-- coe//cten/ chorac/er5t/;cs ofa C-8 airo/-0 secI/or 0f /M= 0. 60. Fi g. 4 NACA ACR No. L5E21 S--- - - -6 -4- -2 o 2 4- 6 ^ 97 a/e of atrchce o'ecy A4,r .od/ R o / 57"- ,c/h c,7ord 245, o0 o /.9-9 /r c chor-d Z3J, 000 o / 6 -ch ch7 d c0" oo'o CONFIDENTIW A 2 32-1'4 c-h/-_70,- 000 5 -------- C- e O-- ---- l------ -.3 -2 -/ C ./ .2 .3 A7orv)eV'/ coef/'/c/e'7, C,/7c ; ,,ure .5. -- //6'-, o'raq-, o ,d orne4f-c'e fA-/.,en -'c/orac/erI-cs of d C-8 aorfo// Sectfon oft 7= 0 C70. ./2 c 08 .04 O Fig. 5 NACA ACR No. L5E21 - -4 -2 0 2 4- 6 8 c/e "* a CC, e'e 9 -3 -2 -/ 0 / .2 .3 chorocfrlst of oo C-6 o'rfo,/ jc~cf/oi 0* /1= o7eo. Fig. 6 NACA ACR No. L5E21 ,y9/e of" aC/aoc/, cc, dae 6-.- I"- UNAL AI A' -COMMITTE OR ARONAUTS - 4 _ LLL CONfRDENTIALj -.3 -2 -/ 0 .2 .3 / ,gure 7 //rt-- / 7-,O7/ 'nrre'-...cr = 'r>? ccoryctcer/,f*c5 e. c C-S Oc2.'^O,/ / sec /V/o a/ AL- C 90. Fig. 7 NACA ACR No. L5E21 -- .8 QJ \ \j NATIONAL ADVISORY COMMITTEE FOR AERONAuTICS C-K a-= CONFI ENTAL .3 .4 5 .7 9 /o (o) /f/ v-or/ous anF/es o.fo racr. F' a re 8. a/o Con/ of ///i a'/7d rCr coe/Ac/en~s w.:th /f'ac/7 ,-mber for a C- 8 o, -/i/ Jechon . Fig. 8a NACA ACR No. L5E21 S6 .7 .8 .9 /0 tIc7 ane07ber1 A-1 Ie! rarv/ os -P/ties of //i/ coeff/ci/'en F--ure 8 Conc /udecd. Fig. 8b -6 - -4 - -2 4 6 . o -. ol ll | | -4 - 0 .2 - 4 6 40 60 .f 0s5- e'rf CJr .& 40 60 -'ffff')^' c^Of Fig. 9a-e o .'00 lArMUM ll.Ow., Ilfnn run aIgrr .miifr NACA ACR No. L5E21 .4 'k -I, I-/5 6 .u -4 .8 0 to. --6 ,/o,' se c. .O 7' .-.rfo., .. c.ho.o C 315 .2 Es NACA ACR Ho. L5E21 I I i I -- t 1 AI I |||| III ^ :: L'uM tL I -- i I i-- U I--.- I L2 --i *-4-. I 4 --- -1 I- 0' ~ ~ 'JI- 4 *b -9.5( l l ll l l11 1 l cl.5 . H I - -s r e -cris/rhe/0co 72 'ro ch'rs/ed7e, /ron. A' =00o ; rfo// cho70rd, 'IN ,.- .. L. ' . -C~1~l-t~ .-I .. Fig. lOa-e ~''"'C -s~ 1 NACA ACR No. L5E21 Fig. lla-e LONII[l[NTML - -6 -4 4 a ,O 'C. -6'c -6 -_ -.a 4 -/.0 -.4 -6 =2 O .2 10 .4- .6 .6 .8 Z0 /- 4-.4-1--- -- ------ ----- iiiziimmm:'.i I II 2. 7* .....t- o2 40 60 80 1 7rclen chord 49 60 O8 .-.* ceT7' ch o/ r .-.e .' _-?e 3, 5 ^0re d/s7/ru/o meosure/7en - .. *. 1 n, =- 0. 90 o/fo,/ ch/?ord, _ 5 s.-,. l /0o x ,-' J "' 'e , I NACA ACR No. CONFIDENTIAL 6------------ -----^--- 9 a "A .9-/4cbt7 Chor I iTI iPV1 l 1Ti '-d /i'=5a~O -C~ C----------------- -8 -6 -4 -2 o 2 4 6 5 91/7/1e OtfC7TOacA c/ e -.2 I ---- -------------------- C-- -- --- ---- -- ----- --- -- -- -- -- -- --- NATIAL M i D4 -72-- COMMITTEE FOR AERONAUTJIS CONFIDENTIAL -J 0 ./ .2 5 /formeni coeff c/en7', Cr7c, (C//A= o. 70. /Pcqure /2 Corpaor/'o. OE of eu// oSa"/z'eq' o0 C-8 o/t f/b sec A- -= for nmede/. of /we sizes c1/ /t2? A. fch 77berss. . . L5E21 Fig. 12a I NACA ACR No. L5E21 NI I.I I| ,,,,I I- S ./. o-9,-,, c o, 4=,0,oOo CONFIDENTIAL -0-0 4 __- -- 4------------ --- -. -4 -2 -- --4- zqrl/e cr/ Z%3c,.C /c ec? 'U 6 ---------------------- S- .2 3 -/~ r ~ ~~- cCONFIDENTC ALM 3 -2 0 ,, .2 .3 7c,-2',?! / c 'e,//c e Cm (2) //- o. 0 o.. F/%U/6F. / - Fig. 12b NACA ACR No. L5E21 Qj b I I I CONFIlr ENuAl I I |I / 969 -/-7c Co.ho.r'. r-= 46o0,0 ./50- /, 7 c/y:v I /./ O, - 2 2------------------- -- _ _ _- - --- --- --- --- --- -- ^/ /' ,n/e o/'a/,acA, cc, o'ey \U iii --- l-i--i-i---- ---- ---- ---- .4-------------------- NATIONAL ADVISORY I MMIFT FORI AERONAUTlCS 4 CONFIDENTIAL -5 -.2 -/ 0. 2- .0 //o-/ren coe //,c/e-, C, c/ (C)/ A= o. So. F/-ure /2 Conc/c/deo' 7_ . 0 "' Fig. 12c Oj rj NACA ACR No. L5E21 1 F .. 0 t C S (JU ao.q a- 1- c ^ ^, ^ _- iS _* ^ \ _S __ _u ___ .g , ^A ^ S ^ __ ^ __ _g -^-^^^^^S.--"s ^ \ "'*< < '' \ ^ 's1- > \F W T *? '? \ In - -- -- -^--^- -- -- -- o~, la w _--S,-- -- Gt^ 0 ^SJI rr \ar~r *~3 L3 *^ 's '' \ < Fig. 13 NACA ACR No. L5E21 -II- - ua - ^ -- -- -- ^'- - -Q -- -- -- = -- -- -- -- -- --- -- -- -- -- -- -- T h N-,1 -N-F- U) 0 h-A--------------------------------------- --------^-K - -- -- --- -- -- --- ,-, ^ ' ^ oo r- SZ __ __ l Q 9 :--N^-^^- L^ a ,T .wr ra^ o a u ^ c o agea /oI f^/ Crt/j.as-e'3 '///o^^o3 jy.^ cftur~/ 0-)^ I P J|^ M< Fig. 14 NACA ACR No. L5E21 14 f tj .St j10 Sq 0 h j m ^c` o al >Q ir eos t^ I ^ Fig. 15 NACA ACR No. L5E21 Fig. 1E -------------------------------[-- - . -1 1c , siFJ < _ - -^ -I o , __-U _ ' _____ __ _^ ---------- ^ z1 Ez~~rB r c75 roJ,,, eb-67 NACA ACR No. L5E21 0 C~$ eU~d- aLeQD~60 ,~wOac/Qks (hii (to N *^ j 9)' Ff tB c o~s. O0 ?1U I; Fi g. 17a, b r- NACA ACR No. L5E21 Fig. 18a-g CONFIDEhTIAL O' O 760. 2T I- -1--1 E I = SI I I [ 11 1I I 11 1 1 I I 1 1 0 i (' 0 6,0II W 1- -4 "a^": Percent chord I I I I I I I I .L_I L .L_ I IJ _ 'a e 8 essu.e ^.* r- VC/'/ ?-- *^ecascr'ee /or- c -,, '.' C >-- 7 or -. C = -. '5-, r-.7f //'S L-er,"?c 'V/L L-.^-'v .",.-r,, ,e'". f. Q /0 Ge-r~/Ev- NACA ACR No. L5E21 o 0 0 o o c r \1 VSt C7 - 1e - Q) -4% o j (b -NN 2 o dI -O . -' cS^ ^ .< i ^ i f I u Fig. 19 NACA ACR No. L5E21 Q3 'fo'-wo paco de va *n < s4 qL - t -o Oz - s.r -oW - < i t .0'Lpp Fig. 20 NACA ACR No. L5E21 Fig. 21a-x r 3 mkrluranu 1 C. ' L L AWL 9. r 200.6a ;4 /0, ,3,' 14'C aL.5 -64 /0- fJ 44OC,9OX-34 S0 .(9 nci 34- 3 SC O6.63 ( %so o ,'o i /O I, ,) C i OW, 9.6J (I' NIAC 3q ? 9 0 2') 6"'R 23 M J F0 0 40 b 5 60 .1 F,9tre 2/.- jpec.. /ecd booe. ,c .0 -. /0 10 I- ,i O 1o r 0 C --------- - ,0 - ^u(i\ac ' .C, I ' -, 0 - C ,0 j 3 w e0 f 0'- CT~ 4 O 6'S 7 *erce.',r croro o' the ao. s. v 'I ~II .1 .4~yr; ~5~~t~ f$ .11 ih NACA ACR No. L5E21 I- 4(o1 1" 4 0 - 4- Fig. 22a-h A- ;C o'/ o./ '0. 0 cec# ea a '& 4- 06 64 re,' ,4C 0006-34 -a (6l AA'A5 0009-64 -a-- ( N''*, / -7 ,\ (cJ ,VAC94 002-64 N I q ---- -_---At 'A a, ^' it .4lA 0008S-34 (/ q 0 2- 4 (\-V90 GA 0 p -0 --- ---- --- - 4 COfrIt riru t ____ i \6' '4CW5-64 _(_ .1[kJC06-63 0 0 4J 60 40 0 -- t 3o -'e, c ,'* ,- .,-.o tPercen t c.oyd -.AJi' C.!.- -'c.taa .' *.'r5.'S eC.l'e. 35 cO-nOcicreo Dc 5,c.'e'/ .' r,-C'e- . *.-a r.i. ;e. j.sc < 'e.'- Tres j-c :c.. e . 4 0 4 t A, i ,,,M..,u ,--i jiOU. . ur~ua NACA ACR No. L5E21 -- ,9c /ua.' o'r-,,// S oec, 'ed ,a ob/ kiOhl'lihTl 4 / O __ --- 4 -" 1 ,CtA 0009-63 4 , 4 4 ..4 44 / 0 - 4 (K) ---. -309 - 4 - ,4 . ,.-y .- i^.r- or ~ ~C ~ 23 CftL'hr.I ' Per'c, r,'_. --, -C -'.0 ,, --- -- - ,, r ,. S4 - - 60 : A5 Fig. 22i-p NACA ACR No. L5E21 Fig. 22q-x joec.. /leo oro 4 / -I- ... 0 Cz v J.-' 07 / 4 4 cerce-.- t,,cA 6i t'ce'irf t~..r Cv) A/RCA 2509 C 4 (sj /..y` 5 3., /O percent t, c o -- ---- 4 aC' cv',. l nvrcenT~ crcl' (6l Davis, 9percent- 8 4 '9 4hielhil 4 (x) ET- 3609 o 2O 40 60 0o /VO Perc entf cord 9 ''^.- cd. tC7C L,. CO0 `.' '4 1' K' 4 NACA ACF No. L5E21 8 7 NI L.6 2 4 /7- F/qure 23 .- /ec- of cc-r/6bss//,/ on /he 'erdO/-/C 6Char4CC' c-S ,c'f A/) 006-64 oa,.o//. Fi g. 23 NACA ACR No. L5E21 .C9-- .05 07 j 06-- 3\ kj S.04 .03 O (--- - Fig. 23 Cont. Fw as23 Conh,'e .' /Adch n7umber /I 7 NACA ACR No. L5E21 O ,0 ,, S0 0- 0 S. Qtj Fig. 23 Cone. CONFIDENTIAL C_ = 0./ C 0.2 C =0.4 I i NATIONAl ADVISORY CONFIDENTIAL COMMITIM FOR AEPONAUTC ~ =05 [ , ./ 3 .- .5 .6 .7 .8 .5 /.C /LX-ach 'fd'*7 ber /.-7 /oure 2-3 Coc/2ufe*&. NACA ACR No. L5E21 Fig. 24 ~7 CONFIDENTIAL 6 / ---- .NATIONAL ADVISORY CONFIbEN'IAL COMMITTEE FOR AERONAUTICS S I 7 I Q I \I /Aach nern/her, /, F/Cure 24, efe ofo ccn- ressb//// cJ e V xerc/uri/t croc/5s of7- 7e IV19C9z-OOC9 -64- o;/'-r7Ebv NACA ACR'No. L5E21 C qi 4J q- 0 Fig. 24 Cont. II, 0 :ONF1DENTIA .0--- _--- 506 --- 4 ----~-. O c / 2 L / nI/ 3 -- -__deq / w/ J NATIONAL& ADVBOVM 0 1 1 I I r 0 ./ .2 .- 6 6 7 .8 .V /O /v/ach nc/mfeber, /17 F/re 24- .- Coc- Kn^ee NACA ACR No. L5E21 I 1 i~ HZ{rIIIEILIK I I I III1~(~f I I I ~ I I -I 0 0 U 0 " Q , F~y/~ Q Cecy~/dec1. .-- .3 .5 .6 7 1 1 , t C-0, CONFIDENTIA COMMnrEE Frm AEIOmKNA -- 0.4 r 4 .2 .3 d 4 .5 .6 -7 *Q *9 /4 / occh noc/wnher, 'Ad I -FT -- 1-' Fig. 24 Cone. - I NACA ACR No. L5E21 " .5 4 33 ,J.a ./ o Adach nrcber, A-1 F/yUre 25 .- Efect C co0mpres/b/ on 6e qerodq nor vc cahoqc rs/fcs c ,/VC/' /z -S 64 oair >:/ Fig. 25 NACA ACR No. L5E21 Qj S7------.------------------------ ---------___-------------------- 41 ^ ^-----_----------------------------- J o j NA CONFIDENTIAL --comin 0o / .3 .4 .5 .6- .7 /woch /72c5-Y /7 _T/f /e 25 CGO-vAijUec- Fig. 25 Cont. NACA ACR No. L5E21 'U .1 V U QJ 0 i .1 QJ Fig. 25 Cone. C = 0. 2 --' :CL--Lo.2 ~ = -3 NAIIONAL ADVISORY COUMMITTLE IOR AEONAlmCS(S CONFIDENTIAL C, = 0.4 ~ ./ .3 .4; 5 C ..' .8 .3 /0 / 'och r-, ,"' 'e, ", F,-ure 25-. Ccnc/cdcd. NACA ACR No. L5E21 6 .5 A) /Vlcch r/orber, A- four/e z26 .- fecw' of p2pres;/// cM o/e Verod/n-/ ic chaac/erJ/c.s of 7he /VNc c 00/5-64 o/,r-o//. Fig. 26 NACA ACR No. L5E21 Af 26 n C rneA -7d. F/yure 26.- Coboveo'. Fig. 26 Cont. NACA ACR No. L5E21 r ./ r S./ -, Fig. 26 Cone. - Q = o'./ ~ - i- I I C1 = 0.2 C_ = O. 3 -- -- -C^ L 0: :^ h^ - NATIONAL ADVISORY COMMIfEE fOR AERONAUTICS CONFIDENTIAL 0 ./ .2 .5 .4 .5- .7 A/fach /umber, /ld. Fiqcure as .- Ccr/uce'- .8 .9 /0 NACA ACR No. L5E21 S 3 4 .6 .7 Ml7ac17 nc^^^ber /V1 .6 / 0 UF/nu- 27 E fec of ccopreassib, on 0f denodqnaormc chor-oc7'rAshc o C Ile Acx o 04; oir 74c/A/. .7 .6 .5 .4 Qj .3 0 -.1 CONFIDE TIAL dey) 5 3, 3 --- ; -"- -o NATIONALL AM SORY CONFIDENTIAL COMMITJE FOR AERONUTI Fig. 27 NACA ACR No. L5E21 c te 2 7 r. 'on/ 7;vu F/l~c./re 27 .- Cont'n/uect Fig. 27 Cont. Fig. 27 Cone. NACA ACR No. L5E21 / -CONFIDENTIA II . J ./. , .-4 -.. _, 3 G/ --- COMMITTEE FOR ARONA114ICS --. / S---/ -2 .3 .4 .6 7 8 .9 <~\a/ c.uh rZ Z^. Il^ZZ*7^ F~re 27. Ccv7c/A/ced. .NACA ACR No. L5E21 Fig. 28 CONFIDENTIAL (ciej.F----------__- --- 4-- _ 6 4 -- _- 5 ------------- ---------------------^y--------------^--------L--^-------- -NK, __- ------- --^------^\ - ij 3 -- -- ..-- ^ 3- -- - o *- NATIONAL ADVISORY -, I- \ -/ -i ----Z-- I--' - -M CAONOIDFNTIAL 0J 11- -- l- -- ) 1 Ha -. -^ ^ ^ ^ -- ">> >a,10^i- - /V/oc/7 7curnyber /V7 ..'Q' 2 f/ec/- ofr ccpn~S6sib///J/ on /he CrC/'^q-orn7aC ch/r,-oc ~ta .'/. ,of /VACti coc6 -3sf y/'- '/ NACA ACR No. L5E21 .10 --- .0---- .07 Qj D7------ .Q .06---- 05 ---- .03 \ ^ Fig. 28 Cont. /A- c' cL, nber, A-'7 Fuare 26 Con/f Ated. NACA ACR No. L5E21 CONFIDENTIAL I I ----------- -- ------- Tcr-- .-'- ------ CL =O. 0- c, o. Q~ --NAl--NAL AbvSi -- - COMMITTEE FOR AERONAUTICS CONFIDENTIAL F -1 ./ .5A .9 1 28 .fE -B Crc~c/udede QJ 0-/ o 0 .--/ Fig. 28 Cone. c> 0 -w v ~7UC~7 NACA ACR No. L5E21 I- '5 Qj 3 '2 ,14'c~ h3 -. ~6~ r F/Vure 29. /Eec" of co/rEresssb/'&' n 4/Ae werc//ram /c c1hroc-raics7ca o h/e /V9C/9 C/2-34 air^//. Fig. 29 NACA ACR No. L5E21 .4---- .O - .0 S-.o---- D7---- .06 - .05 ----- .0; Fig. 29 Cont. 2 ^7ach /7o7nber, A7- 'y/ure 29 -orw ec. NACA ACR No. L5E21 -,o ---------- --I JL <2' F_ .J . -- _ QJ .----/ II I 02 = 0. 3 ./ -- "/Lde i. Fig. 29 Cone. NACA ACR No. L5E21 7 '5 S.a rJ G; ^1 I I S(deq 5 3 \ I INTIONM AVISORY -- COMMITrEE FOR AERONATImS - i i i 1i SI 1 L CONFIDENTIAL I I I 4 5 .6 .7 / --foch no/tber, /1#7 .8 .3 /o / dfure .- ff/ec3 o0 cY0 ?-I- -es)/7 /,,y e Oercyromic6.. c-harocer/sf/cs 5of 7he A44CA co006 -63 or,;6-/. C(ONFlEM flA // -- r Fig. 30 NACA ACR No. L5E21 Y\ b b C, P F/fE .30 Cc./,uec.3 Fig. 30 Cont. NACA ACR No. \U QI 'J I, Q) 'j - L II I / I I i I 3 4 .5 .6 .7 /L7ach nLrnmber /' .Y /0 F/;7~-~- -- Ccrc/eded. CONFIDENTIAL I J I -. 1i 1 F 1 11i r i T1T~FFTT~ Y~ I - I I I I I C* =o./ C~ = o. 1 C_ = o.3 -- ---------- .1-- - NATIONAL ADVISORY I t 5 II- CONFIDENTIAL COMITEE FOR ABSE ND R CS C_ = 0.5 I I .* C1 Fig. 30 Cone. L5E21 NACA ACR No. L5E21 6 .5 V4 Q .3 0 -. a oc1 27/7ber, /V7 Fi/qure 3/. ec-/- of c p.res~v/&y,' O//e ,erogerv-c9 c/racer/scs ohe /VdC/7 0009 -63 q/i/7. Fig. 31 NACA ACR No. L5E21 .// --- 19 .0 --- .' .07-- 05----- S0-- ,.6 -- - .--- - Fig. 31 Cont. f/qure 3/ .- cn//ueo. /AOch I-2> Aber A- NACA ACR No. L5E21 S----/ U _ JI -I ( Fig. 31 Cone. CL_ 0.3 S I I I NATIONAL ADVISORY COQN FIDENTIAL OMMITE FOR AERONAUT --------------__ ___ CL = 0.4 3t * -~ .-> .7* .0 / * /Ifc/c or7c/n67er, / 7 A~9cr-e 3/.- C7Q,,c/ulced. - .r Fif~7UTe 3~ NACA ACR No. L5E21 *\r ' .\- \N /7CtC h r7.r/t6er /47 F/Ique 32.- f/ect O7 c071cire ,/ on Ce- aero'nurrOic Ch/qrc7r- /fer. -5Cs :f he /V?/C* 2306 q/r7//. Fig. 32 NACA ACR No. L5E21 I0 OS0 ------- .0--- (7---- .0 5-- ^- 0.5---- rN0I- - Fig. 32 Cont. .A7cch n/7/efber /v7 F/cyre 32 -- Co /iunea/. NACA ACR No. L5E21 CONFIDENTIAL E -/ - . - /~ ,-I C I I f .'P C .V1 U, or rF CL- FCN=0OF I CL = 0 4r--^^--^ c- -- -r-H,-4J |--- C 5 I CONFIDENTIAL conMNII ALiRCMI C CONFIDENTIAL cjMlirht loF sAONAUICS F/C-/re 352. Co.-c/ e, Yeo. I I I H- cr .7 PH 0- -./ - * 1 I I I I I I - - " ' 1 1 11---t--- Fig. 32 Conc. .," Fig. 33 NACA ACR No. L5E21 CONFIDENTIAL (de 9 - 1 4 .7 .8 . -*e -e -.3 -c'------- o -> ---- ^--'-- -- -- C-4,T- \ 3 __ _ -_ 3 4 .5 6 7 .8 .9 ,,ooq" 'e ch/r7c4j'r/r' s A of IC -' 2309 oir_/_/c ".4 \j %.J .. S* \ ', 0, NACA ACR No. L5E21 .10 ---- .09---- 0(j .^ .^7 -- -- Q ,. Fig. 33 Cont. l/-Ic/ 17c1-43er~^ /- F~gL~re ~~.-- Cc~f/ix/ed. Fig. 33 Conc. - tI I| -- .- "-I- _ Co F -0.1L 2 0 1- "./ ICL-.' I ltr~ ut 'It - U-) 00 -NAilONAL ADi,-,,r. - CONFIDENTIAL i:.MMinE FOR AmLROAulTC o C' .' .e ., 5 6 .7 .6 .' , ,4/4&och n' /7 -f er- /1r7 '= .,,r- 3.3 Co,-,c/c//ec/. NACA ACR No. L5E21 N I C.. o. 2 , - ,. S-., JAACA ACP No. L.5E21 7 -- 6 - 1.5 .j .33 Y^ --t-z 14 /q'/ue. 34-.- / O comr./ores/'/'/" on /-he m ,/- C;/e'-. ;C5 OA" C . oerJocm.qic c/roac'ers/c o W90=3 *23/2 c iy,'r / Fig. 34 n~7c~ r~~/-~bef, /L~ .iACA ACR No. I.FE21 .7 5 6. t1. S - CONFIDENTIAL ( -) \- - 0 ; ^ E : -,d ---- ------------^ ------ P ..L I ' I NAiiONAL ADil UI - --_ CONFIDENTIAL COMMmE FOR ALRONAUrI( o .2 3 5. 7 8 .9 -C mic c23oc1r2i c-- /' ==Z/2 -,i~r^e/ ? Fi g. 34 NACA ACR No. L5E21 c22 o I. \-I-i I I1 1 -- I-- -- /3 - CONFIDENTIAL I 7671 09- 0I8 06-- -- - 04-4 '3- -1------- - 3 3 _---- 1- ----- --, - ^ -^ ;- --'-f -o-- "1,, I .t 7.. . 92 3- -= -, i --iz iir i nECIi /-'tyra, 34-. Con//uef. l.' i raiiONAL ADVISORY OMMITIFE FOlR AERONAuTICS ---- -------I CONFIDENTIAL COMMITTEE F '- : . Fig. 34 Cont. /Ljc7C~:. /;~/r~;zber, fl-? t^,' J NACA ACR No. L5E21 SCOFI NIlAL I/ - C,-- = 0. / I I I4 -/ -- / - F - r>__ __ C --0.2 I I I ~CL -0.5 NATIONAL ADVfSORY COMMIflLE FOR AERONALJCS CONFIDENTIAL I I I -- =S c .o .r /,Co.4 \ S ... .2 .3 .- 5 .c .7 6 .2 / /AqcAc Ar-6er 1 A/ C' L Fig. 34 Cone. /tis 34, 34.- Cb-ic/odec. NACA ACR No. L5E21 'Z N /'j C /L-1~c/ ,2C~? br, A Fi'cje 35 .- 6ffec t ccPrssi'// on h6a At~fr-o' ic coaric/es/i/c-s of ,e /VN/C9 23/-5 o/-ro/7. Fig. 35 NACA ACR No. L5E21 .,7--- ./6 - ./4---- ./3 ---- \ 0a- - Oj .IJ QN, 0 - Cu -. -- d - o P ~,~ ~ Fig. 35 Cont. .Z 3 4 5 6 7 /-lac/7h .7/ber. /7 ^'qurfe J5'.- Cont/heo'. SNACA ACR i-" i I j % ^ -/ .''. ; No. L5E21 Fig. 35 Cone. CONFIDENTIAL CLO.2 I 11 I I I I I I I IV I I 'A-4c/7 /?SM/bes/^ /17 C, = 0-4 NATIONAL ADVISORY CONFIDENTIAL COMMITTEE FOR AERONAUTICS S / j .4 7 .A .9 F~7ure 35.- COT)C/L~CJEICJI' 0 O . / NACA ACR No. L5E21 .4z .) 4 . ..9 .' .A - /V/och nc *be r, / / F /ure 36.- E-fsde o' ccm rerss/bi/b orn e e-oc/nc7,c ciqr7Ce's" /c ofe A/^c2 2o4& o,?-to//. .7 .6 J5 X I, '\ Q0 0-2 U 0 _OFID NT AL 3 2 .IO F NAON AISORY COMMITTEE FOR AERONAUTICS CONFIDENTIAL _ _7 _1_ Fig. 36 1 NACA ACR No. L5E21 0 .i r 0* /A-lach /rer76bte r, /1- F>wmre 36 .- CoA/,i ed. Fig. 36 Cont. NACA ACR No. L5E21 CONFIDENTIAL NATIONAL ADVISORY- - COMMITTEE FOR AERONAIICS I 7- m ---------------------------------------------C. 5 I -_^^ C = 5. CONFIDENT AL .5 .6 .7 .8 . / icEh FRAOAbeUI /tS A~,/re 36 .- Cc-c/deead . Li U -I Fig. 36 Cone. NACA ACR No. L5E21 .7 .6 .5 \N /^oc7h nrewnber, /V7 F/dcre 37 .- Effec9 o7f ccv/9,resseb///'^ on M c 1e-ocyno1r9c c2Oradcer9/s/C 0I7/ & 9ca 's409 0/r8b/'/ Fig. 37 NACA ACR No. L5E21 ./! 08 U - .07 -- O .05 --- t .06 05--- Fig. 37 Cont. /lac/7 ncvber, A/V 7equrtr 37 Cbp* /A>?c ed'. 0NACA ACR No. L5E21 CONFIDENTIAL c-I I I l I' I I I I I I I I I I I~r41t1 1Io I 3-I 0 J\0 1.V S-I - 1 44- 4 L[ -4 / 're J7 C &7c/cded. .!:; '.E *ii" .-i: f". t.: r c,=-o.. / C = 0. ZZZZZZZ^^^^1111 ^i CL =0.4 iNAIOlNA ADVISORY CONFIDENTIAL COMMITLE FOP AERONAUnCS R 4 I I I I, I I I 1 i 3T Fig. 37 Cone. . -- *- .^ rr J rr .I u v A-o2CA ~7/r~r A- uMre 30 .- i41 c/2 of coy.r4essi/ or */s. aerco'qami'c chCr/ocfrCsf/cC/ of V/C7f 24/ 2 cOr6//. NACA ACR No. L5E21 Fig. 38 7 .6 N5 .3 0 NACA ACR No. L5E21 // ---- 14 - ./0 U o----- 0 Q7-- L .o*---- 05-- Fig. 38 Cont. A-12 Ct? ,'76--X69P A-i' F17Ldre Xi C'o,,-~/,>dje NACA ACR No. L5E21 \./ k --0 , tU\ - QJj 0 ~~ Fig. 38 Cone. CO Fl EN IAL III-I--. {J^-t II I- I- 9- / CL = 0.2 I 1 I II, S =o. 3 NATIONAL ADVORY COMMITTEE FOR AERONAUTICS CONFIDENTIAL --, 1 /01 1- N -/ ------- --C = 0.4 I I I C .2 .j 4 .5 .6 .7 .8 .9 .0 7'cfch no ^77bher~^ /717 A7u"ee M3 Ccv-c-/kedec/. -- |
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