'l\tl\L-l^^ ^ ACB No. L'5H08 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED October 191^5 as Advance Confidential Report L5HO8 SUPERSONIC -TOHKEL TESTS OF PROJECTHES IN (33RMAHY AND ITALY By Antcaalo Ferrl Langley Memorial Aeronautical Laboratory Langley Field, Va. 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. 152 DOCUMENTS DEPARTMENJ 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/supersonictunnelOOIang NACA ACR No. L^HOS NATIONAL ADVISORY COmfllTTSS FOR AERONAUTICS ADVANCE CONFIDENTIAL REPORT SUPERSONIC-TUNNEL TESTS 0^ PROJECTILES IN GERMANY AND ITALY By Antonio Ferri SUMMARY Tests were performed in ths Gottlngen (Germany) and Guidonia (Italy) supersonic turinels in order to determine the aerodynamic characteristics of projectiles of various shapes. The Mach numbers ranged from about ^1. J to 5«2 for the Gottingen tests and from 1.1|Il to 2.66 for the Guidonia tests. The results sho«v that incr^easing the relative length of the nose causes the drag coefficient to decrease and the center of pressure to move forv.'ard. For a given length, the nose having minimum drag has a curved profile; the curvature is greatest at ths tip and decreases to a very small value tov/ard the rear of the nose, v/here the shape becomes approximate I7/ conical. As the Mach niamber increases, the drag coefficient decreases and the center of pressure moves toward the tail. For the higher Mach numbers the variation of the drag coef- ficient and the movement of the center of pressure are small. Existing aerod^mamic theory gives values of the aerodynamic characteristics close to those determined experiFientally for small flov; deviations. INTRODUCTION Research on nrojoctiles was started at approximately the saiae time ( 19L.2 ) at the Gottingen Laboratory in Germany and the Guidonia Laboratory in Italy, The original data were brought to the United States in 19^)- snd vifere tabu- lated and analyzed at the Langley Memorial Aeronautical Laboratory of th3 National Advisory Corrmitt'^e for Aero- nautics, The aim of the German research wgs to determine the variation of aerodynamic characteristics with Mach number for various fundamental geometrical shapes for projectiles. 2 CONFIDENTIAL liACA ACR No. L5H08 Force tests of small models 0.593 ir.ch in diaraetei^ were conducted in a supersonic tuniiel having a test section ap-oroximately 2.36 by 2.8I4. inches. The shape of the models was systematically varied to determine the effects on the aerod^mamic characteristics of: (1) Nose profile shape for a typical fineness ratio (2) Nose length for noses with circular-arc profiles (5) Small taper of the tail of the projectile The program carried out at the Guidonia Laboratory had as Its aim the development of an optimum shape for a 1. 8l2-inch-caliber antitank projectile. Because of the relatively large size of the test model, it was possible to obtain pressure-distribution studies as well as precise aerodynamic data for com.pari3on v/ith results derived by existing theory. The theories presented in references 1 to 6 for sharp-nose orojectlles at zero angle of attack and the theories of references 7 snd 8 for bodies of revolution at an angle of yavi? were used to compute the theoretical characteristics of the various conical lioses for com- parison with the experimental results. The aerodynamic theory of m. in im^ urn- drag projectiles presented in refer- ences 3 and 5 to 7 was used as a guide in the design of the various nose shapes of the projectiles tested. SYMBOLS The symbols used for defining the aerody.mamic coef- ficients and the geometric characteristics of the pro- jectiles are given in figure 1. Slq speed of so^ond in free stream Vq free-stream velocity Mq free-stream. Mach number ( Vo/aQ ) q^ free-stre?m dynamic pressure p local static pressure CONFIDENTIAL MAC A ACH Fo. L^KOS COrFIDSNTIAL pQ free-stream static pressure P pressure CGelTicient \^ — 5 .' d diameter of body of projectile R resultant force on projectile D drag L lift M pitching moment about rear face of projectile Cq drag coefficient of model n n N f- ^ MoV" f)/ Ct lift coefficient of model / ; rr- \ L . i / d2^ ■ C-fjj olt ch in p; -moment coefficient ^ibout rear face of proje / ectile / ■ I lengtn of projectile l^T length of nose of projectile r radius of nose of projectile f center-of-pressure position, measured frcm^ resr face of projectile a angle of attack of projectile fineness ratio cf projectile (l/d) fineness ratio of nose (^nA^) € angle betv/een intersection of tengent on nose cf projectile end generatrix of cylinder COJ^FIDSKTIAL k C0NPID3CTIAL 'MCA ACR No. L5HO0 EXT'ERir.'EKTS AT GvO^TINOTI^ #lnd-Tunnel ai-.d Experimental Methods Experiments were carried out by Geriran technicians of the Aerodynainische Versuchs enstalt (AVA) in the small supersonic tunnel pt the Gbttingen Laboratory. The layout of the wind tunnel is shomi in figure 2. The t'onnel has a rect^ngul?,r section vvith s throat about 2.56 by 2,814- inches. As Indicated in figure 2, a semiopen throat arrangement was used. The side v/a.lls of the tuxinel were straight and parallel, but the jet was not restrained 'o^T top and bottom walls. It has been fouund that -chis arrangement makes it possible to obtain reliable aero- d7vTnsmlc data at Mach ri umbers only slightly greater than 1.0 (reference 9) ^i^d that the choking condition which vi/ould exist if the jet ware completely restrained does not occur. Some trouble was encountered during the tests because of condensation phenomena, in spite of the fact that the hijLiidity of the am^olent t-ir had been reduced to a low value by preliminary drying. The tests consisced of the i.ie asurement of lift, drag, and pitching m.oment with a semiautomatic balance. Each model was tested through an angle-of-attack range from 9° to -3° and a Mach number range from 1,5 to 5.2. For the longer models, it was not possible to perform the tests at the low velocities because the front shock v;ave reflecting from the jet boundaries interfered with the flovi' on the rear face of the projectile. Test Models The models each had s diaireter of 0.593 inch and were supported by s. sting attached to the rear face. The dimensions of the sting and the tare system adopted are not knov.Ti, The models can be separated into three distinct groups to determine the effect of: Nose shape .- Four projectiles having over-all fineness ratios n of J>.0 with nose fineness ratios of 2.5 (moaels 1, 2, 5» snd L\., fig. 5) were tested to determine the effect of the nose shape. Because the models were small, they had no tail taper or bourrelet ring. All the noses CONFIDENTIAL NACA ACR No. L5HO8 C0WPID3NTIAL 5 were of circular profile with the radii r varying from 6.5d to co (conical nose).. For the nose v/ith 6,5'! radius (model 1, fip;. 3), the end of the ncse v/as tangent to the cylinder at their juncture bat, for the other models, the profile of the nose terminated with its tangent inclined at an angle € with respect to the generatrix of the cylinder. Length of the nose .- Five models having nose lengths varying from C.5d to 3.5d (models S* 6, 1, 7> snd 8, fig. I4.) were tested to determine the effect of the length of the nose. The noses of all the models had circular profiles tangent to the cylinder forming the body of the projectile. Tail taper .- Three models derived from m.odel 1 ai'id having three different tail tapers (models 9j 10> snd 11, fig. 5) were tested to determine the effect of tail taper. Results W ind-tunnel tests .- The results of the experiments at the Gottlngen Laboratory are given in figures 6 to 21, Figures 6 to 11 shov/ the results of tests to 'determine the effect of the nose shspe. In figures 6 to 9 the variation of the aerodynamic coefficients for models 1, 2, 3, and ij. is shovim for several angles of attack and a rai'ge of Mach n'ombers. Figures 10 and 11 show the aero- dynamic coefficients of each projectile at equal Mach numbers as a function of the ratio d/r. As shown in figures 10 and 11, the m.inim.'um drag coefficient was obtained for a nose intermediate to the noses for which r = 03 and r = 12.5d> The differences in the minimum drag coefficients were not large. The results of the tests to determine the effect of the length of the nose are given in figures 12 to I7. Increasing the fineness ratio of the nose nv caused the drag coefficient to drop noticeably (fig. I6 ) . The slope dCD/dnjj decreased as nj^ increased. The center of pressure of the projectiles moved toward the nose as the fineness ratio of the nose increased (fig. 1?). The results of the tests to evaluate the effect of the tail taper are given in figures I8 to 21. The vari- ation of the aerodynamic coefficients vjith Mach number is CONFIDE] TIAL CONFIDENTIAL HAG A ACR No. L5H08 shov/n in figures l8 to 20, and data for comparison of the tapered models with model 1 are shovm in figure 21. I'h'? drag of the projectile ■-vas lowest for the longest tapered tail, especially at the lower Mach numbers (fig. 21). The differences in drag coefficients were not large. The other aerc'dynemic characteristics v/ere not appreciably altered. The drag coe decreased as the was most pronounc the variation at slight. The posi change appreciabl supersonic veloci toward the rear o increased but ten higher Mach number fficient for a given projectile shape Mach number was Increased. This effect ed at the lower supersonic velocities; Mach numbers of the order of J.O was tion o± the center of pressure did not y with angle of attack. In the lower ty range the center of pressure moved f the projectile as the speed vi/as ded to approach a fixed location at the Firing tests .- Actual firing tests were performed to verify the experimental values, and the following results vifere obtained: CD Model Firing tests (a ^ 5°) Tu;nnel tests (a = 3'') 1 2 o.Lo .56 .40 0.29 .26 .28 These I'esults were for a r.'iach number of 2.2. The projectiles used in the firing tests had an angle of attack of nearly 3° ^■^^ S- bourrelet ring. Drag coefficients obtained from firing tests at angles of attack near 3° gave drag coefficients that were equiva- lent to a wind-tunnel angle of attack of about 7«5°' The differences therefore cannot be entirely due to the presence of the bourrelet nor to the error in angle of attack. The differences may probably be attributed in Dsrt to the difference of surface finish between the tunnel model and the fired projectiles, the rotation of the fired projectile about its axis, and the difference in Reynolds number for the firing and the tunnel tests. CONFIDSIITIAL NACA ACR l^o. L5H08 CONFIBd]NTTAL EXT=ERI?/IEKTS AT GUTDOKIA Wind -Tunnel and Exoorimental IJethcds Projectiles v/ith nine diTfereiit nose shapes -were tested in the closed-throgt high-speed tijnnel at Guidonia (refsrence 10) st M&ch numbers ranging from l.Iili- to 2. 06, The system of the psrtially open str-eai'a v;ss not used because it i-'equired a larger Piiiouiit of pcv;er and there- fore lirrdted the maxim-om velocity. The test section was large enough not to require special attention to prevent choking of the air stream, with the model in the tunnel when the Mach number was greater than LIiJa.* The nozzles were of rectangular section of the two-dimensional type with a mdnimijm section I3. ?■'-!- by 15* 74 inches. Tho forces on each modjl were determined by use of a three -component balance (r-exerence 10). The model was attached to the balances by a sting on the projectile axit on the rear face of the orojectlle. The sting, although of smsll aiameter, affected the experimental results scmewhst since it iricr-eased the pressure on the rear face of the projectile. It was necessary, therefore, to rra^B rn accurate' tare measursinent by suspending the model' on a faired strut attached to the side of the projectile . Pressure distributions and optical observations of the flov/ were also cbtafr.ed for 3oji;e of the projectiles tested. It w.?3 difficult to obtain good flow photographs becfuso the phenomena were conical and the density of the air v/as extremely low. Some of the observations were made with a schlieren Toiparatus, and in some cases data were obtained by means of a shadowgraph apparatus. Before the sy?;tematlc experi;rients were started, the results obtained in the tunnel v/ere compared v/ith these obtained by firiu']; tests. A sphere tested at two velocities (I.i^^ = 2.C6 and Mq = 2.62) hrd a constant drag coefficient (Cp = C.95). These wind-tijnnel rosulcc en the sphere at Guidonia agreed with the firing data (Cj) = Os96) ^J^d kvere close to the results obtained at Gottlngen, for which tiie dreg roe '7ficlent in the l^Iach number range bet".vern I.5 and 3.I v-;a£ almicst constant and equ.al to 1,01. In an earlier v>:ind- tunnel exneriment (reference 11), somewhat lo'v-sr drag coefficients were C0KPID5KTIAL 8 C0NFID3NTIAL N,,CA ACR No. L5KO8 found: O.87 for a Mach nuir.ber cf I.85 and 0.36 for a Ivlach nua-n.ber of 2,13. These earlier results, Lov/ever, are questionable bec?use the effect of the suppoi'-t strut, which increases the pressure at the rear somewhat, v/as nec'lected. Test Projectiles A 1. 8l2-inch-csliber antitank projectile was used for the body of the nrojectile in the Guldonia tssts, the detf^ils cf which are shown in figure "22. Nine different nose shapes {fig. 25) were tested with this boay in order to deternine the nose for minim-ura drag. Conical noses of varying fineness ratio (projec- tiles 1, 2, 5j ^'nd I4. of fig. 25) vi'ere bested first in order to estimate the importance of fineness ratio. The simple conical nose form was chosen to permit comparison with existing theoretical data for conical noses. Both the approximate projectile theories (refer- ences li and 5) snd the com.plex but mere exact theory of Ferrari (references 6, part II, and 7) show thrt, in order to minimize the drag of the projectile, it is necessary to concentrate the pressure at the vertex of the nose snd then to carry out the most rapid expansion possible, 'i/i/lth this criterion as a guide, five noses of fineness rstio n^ =2.0 were designed (projectiles 5> 6, 7, 3, and 9 of fig. 23). (See table I.) Two of the noses tested, projectiles 8 and 9> v;ere blunt-ended but were otherwise sirrdlar to the nose of projectile 5" The nose of projectile 5 approaches that theoretically derived by Ferrari (reference 6). A nose exactly corresponding with the optimum nose described by the theory of Ferrari v;as about to be tested v;hen the tests were suspended. Results The results of the experiments at Guldonia are shown in figures 2lx to 35 • The variation of Cj) and f/d with fineness rptio of the nose at a Mach number of 2. Ob and en angle of attack of 0° is sho^vn in figure 2l+; the variation of Go and f/d with Mach number for the nose having a fineness ratio cf 2.0 (r.-ro jectile 5) st an angle of attack of 0° is shovvn in figure 25. The drag coefficient C0WID3NTTAL KACA ACR No. L5H08 C0NFID3NTIAL for a givv3n projectile sh^pe decreased as the Mach number incressed. The position of the center of pressure did not chsnge appreciably -.vith angle of e.ttack. The values obtained from firing tests for two noses having fineness ratios of 2,0 and 2.5 at a Mach number of 2.IJ4. are also shown in figure 2I4.. The angle of atteclv in the firing tests varied bt^tweeii 2° and 3°» The drag v&lues given by the firing tests are some- what higher than those determined in the v/ind tunnel. The differencG can probably be attributed to the fact that the models in the tunnel were perfectly finished but the firod projectiles had s rough machine fiiiish. The fired pro- jectile also had a rotational motion that v/as not repro- duced in the tunnel tests and that undoubtedly altered the phenomenon of the bouiidary layer. The angle of ure 26 f grab! on The pre 3 determ.in the pro j seven st The pres sho^ATi in grarhs f pro jectl variation ol' the aerodynamic coefficients with attack for projectiles 1 and 5 is s'hovm in flg- or ■Mr - 2.06. The values obtained from. Inte- of the pressure distributions are also shov^n. sure distributions over projectiles 1 and 3 were ed at angles of attack of 0^, \\^ , and 8°. IVhen octile was yawed, the pressure v/as determiined at atlons around the projectile from C*-* "to l80°. sure distributions for projectiles 1 and 5 are figures 27 and 28. In figure 29 flow photo- or zero anf'"le of attack are shovi/n for these les. The following tables show calculated values of the initial shock-wave angle and the pressure coefficient on the nose for projectiles 1 and J at a Mach number of 2,06, These quantities wei-e calculated by the method of refer- ences 7 snd 8. Experim.ental values of these quantities are shown for comxO arisen. Kose Angle of (de shock g) 1 wave ! The ore tic a 1 Experimental j a = 0° 1 5 35.0 . a.6. C 53.2 CO'TPIDENTIAL 10 COWPIDSNTIAL^ I'TxvCA .iCR No. L5HO8 Nose 1 3 1 1 3 Stf'tion (deg) I P - Po I Pressure coefflcienc, — : 180 180 Theoretical a = 0° 0.53 .18 = I^° 0.b7 .1+0 .27 .11 a - 6" Exoerlmental 0.50 .19 0.71 .28 1 1 0.83 o.SIj. 1 180 1 .29 .1+; 3 .37 •5° 3 l3o .06 .03 The follovving observations can be made from these test results: (1) W'nen the flow deviations sr-e small, th3 nose phenomena aro close to those 10 re dieted by the thaory. (2) At the higher angles of yaw appreciable differences exist between the theoretical and experimental pressure coefficients, particulsrly if "che nose is short. (5) Prossura on the rear face of the projectile is only sllRhtly affected by the nose sliapo but is appreciably decreased with an increase of angle of attack. Figure 50 shows the variation of the aerodynamic coefficients vd.th angle of attack for projectile S, and figure 51 gives the pressure distributions for this projectile at a Mach number of 2. Go. The following C0ITPID3NTIAL NAGA 4CR No. L5H08 ' GONFIDS^^TIAL 11 table com>oares the dreg coof f Icien ts for a = 0° and the CGnter-of-prGssuro positions for the various projec- tiles for ?.!o = 2.C6: Projec ti le i Cd f/d 3 0.568 1.98 5 .352 2.05 6 . 592 7 . .376 8 .562 2.19 9 1 .558 1 2.09 It will be observed that projectiles 5» 8j ^^d 9 had the lovi/est drag. These three shapes ere closer to the optimum profile predicted the Ox^e tic ally than any of the other noses tested. The pressure-distribution disgrar-.s (fig, 51) si'-'d the flov; photographs (figs. yl+ sua 55) shov/ that, when the front part of the projectile is fist as for projectiles 8 and 9j ^ normal shock wave occurs and the pressure at the nose .npproacbes ^he stream total pressure in value. The shock vi-ave is detached from the projectile. Immediately behind the blunt face, of the nose a rapid expansion occurs, and the pressures a short distance from the nose become lower than for the conical nose. These lower pressures act over a relatively large part of the frontal area of the projectile; consequently, a lovver drag coefficient is obtained for the blunt nose than for the conical nose. The pressure on the rear face of the projectile Is about the sairie for both types of nose. The lift at the same angle of attack for the blunt noses is slightly greater than for the conical noses, and the center of pressure is therefore i'arther forv;ard. These differences are very sniall, however. It rnay be mentioned that the blunt type of nose is more practical than the sharp-Dointed nose from the standpoints of con- struction and maintenance. CONCLUSIONS Tests were perform.ed in the Gottinger. (Germany) and Guidonia (Italy) supersonic tunnels in order to determine the aerodynamxic characteristicrj of ;'jro jectiles of various CONFIDENTIAL 12 COKFIDEIITIAL N.-CA /ICR No. L^HOo shapes. The following conclusions sre besed en the results of both the German and the Itsliaii experiments: 1, The fineness ratio of the nose is of primary impor- tance In determining the aerodynamic characteristics of supersonic ■pi^ojectj les . As the fineness ratio increases, the drag coefficient decreases and the center of pressure moves forward, 2, The drag coefficient for a given projectile shape decreases as the Llach number is increased. This effect is m.ost pronounced at the lower supersonic velocities; the variation at Mfch xu-jubers of the order of 3.0 is slight, 5» The position of the center of pressure does not change appreciably v^dth angle of attack. In the lower supersonic velocity ranpe the center of pressure moves toward the rear of the projectile as the speed is increased but tends to approach a fixed location at the higher Mach numbers , ij-. The pressure en the rear face of the projectile varies appreciably with angle of attack but is only slightly affected by the form of the nose. 5» For a given fineness ratio th^ optimum nose profile has p rslativ^ly blunt end, which is faired to the cylindrical pf.rt of the projectile. The theoretical criterions for the design of the optimum nose profile have been v .^-Ified, 6, The existing aerod^/namic theory for the calcu- lation CI the pressure distribution about pi-ojectlles is adequately precise for small flo':/ deviations. 7. The addition of a small taper to the tail of the projectile diminishes the drag slip-htly, particularly at the lov/or Mach numbers, v/ithout p.ltering the other aero-- dynamic characteristics, Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va. COMFIDEITTIAL .iCR ITo. L3H08 CCl^IDEFTI^L I5 REF51-S^CE3 1. Taylor, G. I.., snd Msccoll, J. '■iu : The Air Pressure on a Cone Moving et High Speeds, Proc. Roy. Soc. (London), ser. A, vol, 159> no. SjS, Feb. 1, 1955* pp. 278-511. 2, 3usemann_, A.: Drucks auf ksgelformige Spitzen bel Eewegitng mi t ijoers chalices cbwindigkelt . Z . f . a. ?'. M. 5 Ed. 9, Heft 6, Dec. 1929, pp. ksS-k^o. 5. Ferrari, C.: Campo serodlnaiaico a velocita iper- acustica attorno a im solido dl rivoluzione a prora acuminata. L» A-'^rctecnica, vol. XVI, no. 2, Feb. 1956, pp. ].21-150. 4-, von Ksrrri&n, Theodor, and Moore, Morton 3.: Resistance of Slender Bodies Moving vd,th Supersonic Velocities, with Special Reference to Projoctiles. Trans. A.S.M.E.^ vol. SL;., no. 23, Le-i. I5, 1952, pp. 505-510. 5. von Karinsn, Th. : The Px'-'oblera of Resistance in Compressible Fluids. GALGIT Pub. No. 75, I936. ■ (From. R. Accad. d'Ttalls, cl. sci . fis., .mat. e nat. , vol. XIV, 1956. ) 6. Ferrari, C: The Deterriilnation of the Projectile of Ivlinirum 7i/ave-Resis tajrice . R.T.P. Translation No. 1180, British Ministry cf Aircraft Production. (Part I frora Atti R. Acc?d. Sci. Torino, vol. 'JU, July-Oct. 1959, pp. 675-693; Part II from Atti R. Accfed. Sci. Torino, vol. 75, Nov. -Dec. 1959, pp. 61-96. ) 7. Fei'-reri; C; Campi di corrente Ipersonors attcrno a solidi dl rivoli.Lsione. L' Aerotocnlca, vol. XVII, no. 6, J^one I957, pp. 507-516. 8. Ferrar ■ , C: Determinatlcn of the Pressure Exerted on Solid Bodies of Revolution vd.th Pointed Noses Placed Obliquely in a Strepm of Compressible Fluid at Sunersonic Velocity. R.T.P. Translation No. 1105, British Ministry of Aircraft Production. (Fro.m Atti R. Accsd. Sci. acrino, vol. 72, Nov. -Dec 1936, po. lLC-165. ) CO^•FIDeNTIAL ik COWIDEIITIAL N.Xh ACR No. L5H08 9. Ferri, Antonio: Completed Tabulation in the United States of Tests of 2i| Airfoils st High Mach Numbers (Derived from Interruptad Work at Guidonia, Italy, in the I.5I- by 1.7l+-Poot High-Speed Tunnel). NACA ACR No. L5S21, 19[;5. 10. Perri, Antonio: La galleria ultrasonor? di Guidonia. Atti -.li Guidonia No. I5 , 1959. (Available in Aircraft Engineering, vol. XII, no. IJ4-O, Oct. 19^0, pp. 502-505.) 11. Ferri, Antonio: Influenza del numero di Reynolds ai grandi nutneri di Mach. Atti di Guidonia No. 67-68-69, I9I12. (Available as R.T.P. Trans- lation No. 1988, British Ministry of Aircraft Production. ) CONFIDENTIAL NACA ACR No. L5H08 CONFIDENTIAL TABLE I NOSE CRDE^ATES FOR PROJECTILES :TSD AT GUIDONIA y 1 1 '~--. ->^ Projectile ""-^ 5 6 7 8 9 1 ! X "^^-...^ ^ .25 .1 .6 .2 2. -5 1.5 i .50 .2 .9 2.7 1-5 1.00 .i^ 1.6 .6 3.2 1.6 2.00 1.0 2.9 2.0 L.p 2.1 3.00 i.L. U.i 2.1^ 2.5 i 5.00 2.2 "^'^ 3.8 ^■.^ 5.3 ! 10.00 5.9 8.9 6.7 5.1 20.00 7.5 lii.O 11. ii 9.2 8.9 30.00 liO.OO 10.9 17.9 I4.9 12.3 12.^ lU.i 20.7 17.0 15.1 so. 00 17.1 22.5 19.9 17.7 17.9 60.00 19.7 2,3.8 21.7 19.8 19.9 70.00 21.5 2L.5 23.0 21.5 21.5 i 30.00 2^.0 2I+.2 2I..9 2^.9 24.5 2^.0 24.2 23.0 90.00 25.0 2I4.2 1 100.00 25.0 25.0 25.0 25.0 25.0 NATIONAL ADVISORY C0MMITT33 FOR AERONAUTICS COKFIDSl^ITIAL NACA ACR No. L5H08 Fig. < \- Z UJ o o u CO o ,Q E >i CO d-i O a o •H +3 •H □ n 3 D ■ D tJ- □ "fe I' ^.6 aa — a -6 1.4- i.o OL (deq) y- O ^-^ o Q □ CI CO NATIONAL AD MMITTEE F0« A VISORY ERONAUTICS cc NFID FNTI ^L /.O /.^ /.6 ^.2 2.e 3.0 J>.^ Figure 6.- Concluded. Fig. 7 NACA ACR No. L5H08 0" <;: Q; I) .6 — NATIONAL ADVISORY COMMITTEE FM AERONAUTICS o-^ so -> 9 ^— — ' '6- o < /• > o 3 P r r c ONFII JENT AL /.O /.^ /.c9 ^.<:? ^.e d.O u'.^ Figure 7.- Lift, drag, and moment coefficients and position of center of pressure as functions of Mach number for various angles of attack. Model 2. (Gottingen). NACA ACR No. L5H08 Fig. 7 Cone. .Qi V. Of o o ■k ^o i.e /.^ .<3 .4- (c/eQj CONFIDENTIAL 1 1 -9- ^ ii £ r ^ ^ e ^ <:> — — i Cj ^ 3 3 a c 3 Q J. 6 3.1 — 5^ Qj 0. q: ^ 2.Q 2 4 a.o ■ oc (c/e<^^ 1 0^^ 5 ] 3 E c CO NAT 10 MMITTE NAL AC E FOR A )VISOR ERONAU i rics cc )NFID 1 1 :nti> \L /.o 1.4- 1.6 a.e ^.6 /^och nunobe r- , M, 3.0 ^.4 Figure 7.- Concluded. Fig. 8 NACA ACR No. L5H08 Q .0* Q .6 .£ o <> o s □ D D □ o (b I' ^3.6 y3.a. ^.(3 a.4- 2.0 ) 0/S3 D D D co> NATIONAL ADVISORY (MITTEE FOB AERONAUT ICS cc 5NFIC ENTI AL hO /.^ /.6 a.2 2.6 a O 3.4- Figure 8.- Concluded. Fig. NACA ACR No. L5H08 Q Qj f^ 0^ .0 .«t £ .■4- .<£ ex. {O^eg J Q A A I \ A A I. Y' 6 / — /\ 3 [ Q — — D \ D CO NFIDI ENTIi L /.o /.^ 1.(3 ^.^ Mac h number ^ a. 6 iS.O k3.4- Figure 9. - Lift, drag, and moment coefficients and position of center of pressure as functions of Mach number for various angles of attack. Model 4. (Gottingen) NACA ACR No. L5H08 Fig. 9 Cone. ^« Qj /.6 l.i .6 /I- oc 1 1 1 CONFIDENTIAL 1 1 (c/e<^J A A A '' . 9 i L G < > ^ O k ,=^ □ □ u ■□ t p NATIONAL ADVISORY COMMITTEE FM AERONAUTICS «k o ' s. Q; "b ^ Q. (.1 X «. 1< S s^ s C5 0, % a. 6 2.4- a.o cc (o^e^J /9 7^ T n ( 1 □ { 1 cc )NFID ENTI AL /.o J. 4- 1.6 2. a a. 6 /^ach number , Mo ao vi.^ Figure 9.- Concluded. Fig. 10 NACA ACR No. L5H08 ^^ .6 i .4 % 8 ^ §^ ^ Model 4 O "^ % 8 (DC ( <^eo) 3 ' \ Moo A 'el ^ ) iel 1 —A .? — ^\ oalel k 4 ^^ A Model 3 1 Moc \ \ ■ \^ \ — -0 Y \ \ O \ A — Q CO NFIDE :nti/ iL NATIONAL ADVISORY COMMITTEE FM AERONAUTICS . .. . a . . 1 -J 1 — .05 d/i ./O ./S r Figure 10.- Lift, drag, and moment coefficients and position of center of pressure as functions of the ratio of the diameter of the body of the projectile to the radius of the nose of the projectile for various angles of attack. M, 2 ,0. (Gbttingen) NACA ACR No. L5H08 Fig. 10 Cone. 5 ^^ \ e I 2.0 oc CO 1 1 NFIDENTIAL 1 ( 9H ) LhA/^r 'el 1 Mc (c - \ \ Model S Mode/ Z ^del 4 3 _-\ c^ a \ \ \ \ o — — -o \\ .8 \ \ , \ □■ — e ^ ' 3 n 1, " Model 4- 6 O ,05 /o ./s Figure 10.- Concluded. Fig. 11 NACA ACR No. L5H08 ^ Moder 4- .6 ■4 .Z Model 4 ■ 05 djr Figure 11.- Lift, drag, and moment coefficients and position of center of pressure as functions of the ratio of the diameter of the body of the projectile to the radius of the nose of the projectile for various angles of attack. M, 3.0. { Go t tingen ) NACA ACR No. L5H08 Fig. 11 Cone. F 11 So 8 /6 /.Z .8 ^ O OC c ONFIDENTIAL 1 1 c ^9J Model 1 -) \ ode/ 3 — ;) Model 3 Model 2 — —A 4 ' \\ \\ V \ ^ er ^— _-o > \ \ \ B- T 3.Z Model 4 Figure 11.- Concluded. Fig. 12 NACA ACR No. L5H08 s .0 /z /o 5 .a o /.o y^ctc/l r)cy^^S er. /^o Figure 12.- Lift, drag, and moment coefficients and position of center of pressure as functions of Mach number for various angles of attack. Model 5. (Gbttingen) NACA ACR No. L5H08 Fig. 12 Cone. ^ •5! U /.z 3 ■4 o NATrONAL ADVISORY COMMITTEE FOft AERONAUTICS ^ 1 '^c ) 1o 5 20 \ ~^ □ r } D /^ /7 CO NFID iNTI/ VL /a /^ /e ^.^ ^.e 3.0 s<^ Figure 12.- Concluded. Fig. 13 NACA ACR No. L5H08 (J I .6 .4 .^ o .1) N N \1 .6 .-4 .^ OC ^ A ^^^^ --^ -^ — A 9 A <^ ^ — r> O 3 n Q- n -v3- -□ — — Q ■ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS . — 1 1 1 \ ze ^^ ^.o ex. {o^3 ) ofo 5 ^ \, ' V ■~>— E -■LJ 1 1 CONFIDENTIAL 1 1 1 /O /^ /.& £.Z ^.(S 30 3^ Figure 15.- Concluded. Fig. 14 NACA ACR No. L5H08 U^ Q) I) /f- NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS G- z /.O /.^ /.6 £.2 2.6 3.0 3.^ Figure 14.- Lift, drag, and moment coefficients and position of center of pressure as functions of Mach number for various angles of attack. Model 7. (Gottingen) NACA ACR No. L5H08 Fig. 14 Cone. ^ 1.0 ■■^ ■K Q) 1.6 /.a 6 . \ 4 o 1- Q I t 1 ■ CONFIDENTIAL 1 1 (c ) ^ - A— — A , c. ■zs -t\ o 3 n — D «^ y O -q u ^ G — =^— ^ O 1 1 ' NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 0* 10 I (deqj 9 A —A -J i A 6 o— -o — o. o -0 A 3 D— t S j3 -D d cc )NFID ENTI AL /.O 1.4 /.a 32 E.6 3.0 3.4 Figure 15.- Lift, drag, and moment coefficients and position of center of pressure as functions of Mach number for various angles of attack. Model 8. (Gbttingen) NACA ACR No. L5H08 Fig. 15 Cone. ^ C 9} I o I ^.4 l.O L€ 1.2. .8 oc cc 1 1 JNFIDENTIAL 1 ( ) A A 3 A A A A o O A -O >^ < A — -o \ o -6 3 i y F f □ Q CON FIDE^ TIAL NATIONAL ADVISORY COMMITTEE FOt AERONAUT I I . . L .J /O ^^ /& ^.Z ^(S 3.0 3.^ Figure 18.- Lift, drag, and moment coefficients and position of center of pressure as functions of Mach number for various angles of attack. Model 9. (Gbttingen) Fig. 18 Cone. NACA ACR No. L5H08 8 I z.o 1.6 1.2. .8 <4 CONFIDENTIAL 3.6 3.2 ^.& ^.^ OC ~Cok^ ) ~0 to 5 i. NATIONAL ADVISORY CONFIDENTIAL committee F0« aeronautics 1.0 /<4 /.& ^Z ^.6 3.^ 3.4 Figure 18.- Concluded. NACA ACR No. L5H08 Fig. 19 ^ 9 .6 ^- «^ ■4 ^ ^ Q\ <:^ .^ U {^ ^ o >J so .6 - .4 .Z (P oc J — 6 3 AKi- ^__i2.^- CONFIDENTIAL NATIONAL ADVISORY COMMITTEE FOB AERONAUTICS /o /<^ /e ^d ^. 6 O >- -^ « > ^ a J D D ' a ■D 1 D- □ 3.6 32 ^.8 ^.4 OC {de<^ ) O 7 'o 3 I □ \ \ \J D a D ■o c ONFI[ )ENT lAL NATIONAL ADVISORY COMMITTEE FM AERON«UTI CS 1.0 14 /& '^.^ ^e ■ O 3.4 Figure 19.- Concluded. NACA ACR No. L5H08 Pig. 20 s A) 5 .L _ /" lAL 5: C^ .ONFI = C DENT /.o /.^ /.a az ze ^.o ^.^ Figure 21.- Drag coefficient and relative center of pressure as functions of Mach number. Models 1, 9, 10, and 11. ( Gottingen ) Fig. 22 NACA ACR No. L5H08 a) z z o o '^!^l9'l=p■ —P'^QC — ■P^WI CTN T3 o P^QOl Vi «i Z lU O ji ii 8 I- z o o O to b3 C o •a ■H o cd T3 0) ♦3 CO 03 O > O (U x: C\3 0) (30 NACA ACR No. L5H08 ?ig. 23 < o o •H c o T3 •H 3 O (d (U CO 0) o o O, 0) ♦A ^°) \. i. Qj '\ Q) V Q; :i N> s \ I ^ a ^.^ /.(3 /.4- /.o ^ -^ r- c ONFI 5ENT lAL 2 c? Figure 24.- Variation of drag coefficient and relative center of pressure with fineness ratio of nose. a = 0°. (Guidonia) Fig. 25 NACA ACR No. L5H08 u' '\ U CN /.o .6 .6 f4 CONFIDENTIAL 1 1 1 Lo-^ --- -^ -o NATIONAL ADVISORY COMMITTEE FM AERONAUTICS 1 \ 1 1 2.2 /.(5 /.^ /.O (X. Cd&g) ^ y / ^Z^'? o o e ^ c DNFIC )ENT AL 2 /^ac/) number-^ A/q ^ Figure 25.- Variation of drag coefficient and relative center of pressure with Mach number for projectile 3. ( Guidonia) NACA ACR No. L5H08 Fig. 26a V.,^ Q O .1) 1<^ 1 s ^ 1^ >i VJ o V p^ 5 \l <^ .6 ° C£3 (fofe grated) ^ C/_ {/hi-Ggrai'eo/) £. f/dC/nfe grated) NATIONAL ADVISORY COMMITTEE FOft AERONAUTICS (a) Projec-h'/e. I- CONFIDENTIAL .8^ ^.^ !b ^ ^ <^ 6 3 , /O /9ng/e of a/tac/(^ ^c/e^ .8 A X-N ^ Vn Q) ^ s. 4) A) N .0 Vi ^ ^ ^ ^ ^ o D o A V Co ^ D ^^7*£> qrofetJ) C j_ (/nfeqraf