Published on: **Mar 3, 2016**

- 1. cLAs9;: T:3:iTioN C_'HA1"G. !‘-. D igiiiiiiiiiii. iiiiii; iilIii ii‘ I. 1176 0O‘I10 97 for the A11-' Matériel Comna. nd. , Army Air Forces HIC£ii—SPEED Anaoommc cHA1=ui. c:i. 'ERIsrIcs 01' A 1/7—SCAI. E MODE} OF ‘IE3 NORSREBOP YB—'li-9 AIRPLANE _By Robert G. Robinson Amos Aeronautical Laboratory Moffett Field, Calif. CLASSIFIED DOCUMENT Thic documgnt curtains cjassified‘ inform§tio_n_ affectirig thl ‘ ‘ 'N3ﬁonai Defensg of the United States within the moaning of - the Espionage Act. USO 50:31'and 32. its i‘. ransrr1Ission'or thi roveiation of its contents in any manner to an unauthorized person is prohibited by law. information so classiﬁed may in imparted only to persons In the miiitary and naval Servlcis of the United states. appropriate civilian officers and empifyaoo of the Federal Govemrnant who have a legitimate interest therein. and to United States citizens of known loyalty and discretion who of necessity must be informed tharoof. ' NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON
- 2. W. NACA RM No. A7013 MICCIFED NATIONAL ADVISORY C FOR AERONAUTICS _ MEMORANDUM for the Air Materiel Command, Army Air Forces EIGE—SmED AERODYNAMIC CHARACTERISTICS OF A 1/7LSCALE MODEL OF THE NOR'. i1'EiROP Y2B—ll~9 AIRELW By Robert C. Robinson ‘ SUMMARY Tests were conducted to find the effects of cciupressibility on the longitudinal stability and control of a 1/ T—sca_'I. e seimispan model of the Northrop YB-ll-9 airplane. Lift, drag, pitching moment, and elevon hinge moments were measured and are presented in graphi- cal form. The results Show that, due to a. loss of lift on the out-— board portion of the wing, the longitudinal static stability decreased rapidly as the Mach number increased above 0.725 and for Mach numbers above 0.735 the model experienced a climbing moment at positive Lift coefficients. Also, longitudinal—control effectiveness began to decrease at a Mach number of about OJT25. INTRODUCTION The Northrop YZB-5+9 airplane is s. 3et—pz-opelled modification of the 3128-35 airplane and is powered by eight ‘I'G—l80 engines. Extensive low—speed w: Lnd-tunnel tests of models of the B-35 airplane have been reported in references 1 and 2, but, since the , jet—powered YB—1i9 will attain appreciably higher speeds, it was considered desirable to test a model at high Mach numbers in order to find the effects of compressibility on its longi- tudinal stability and control. Accordingly, at the request of the Air Materiel Command, U. 8. Army Air Forces, a l/ T—sca. 'I. e senﬁspan model of the ‘$13-49 airplane was tested in the Ames 16-foot high- speed wind tunnel. The ha1f—span model was mounted with its plane of symmetry at the w: i_nd—tunnel wall and with no supporting members inside the test section. Lift, drag, pitching-moment, and eleven hinge-moments were measured at several angles of attack and Mach numbers. This report describes the effects of canpressibility on the above
- 3. : ii. 2 NACA am No. A7013 characteristics and on the effectiveness of the longitudinal-control surfaces. ' MODEL The model, furnished by Northrop Aircraft, Inc. , was a modifi- cation of the B-35 model used forthe tests reported in reference 2. The structure of the model consisted of a steel box spar and plywood ribs covered with a plywood skin. All supporting members were outside the tunnel. Two control surfaces were provided, an eleven which extended from 38.2 to 75.9 percent of the semispan from the plane of symmetry, and a. longitudinal trim flap which extended from 75.9 percent of the semispan to the wing tip. The eleven, which was equipped with an electrical strain gauge and a remote-—control positioner, had an internally sealed. bs. lance, the chord of which was approximately 1+0 percent of the eleven chord aft of the hinge line. The airplane duct system was represented with the entrance built to scale and the total exit area to scale but made up of three Jet tubes instead of four tubes as on the airplane. During these tests the ratio of duct entrance velocity to free- stream velocity was approximately 0.31. There were two vertical fine on the model, one Just outboard of the Jet tubes and the other on the inboard side of the Jet tubes. Dorsal fins extended forward on the upper surface to the l0—percent—chord line. The fine can be seen in figure 1, which shows the model mounted in the l6-foot wind. tunnel, and in figure 2 which is a drawing of the model. The more important dimensions of the half—spa. n model are as follows: Wing areaofnuadel, square feet . . . . . . . . . . . . . . 1+0.82 Spanoi‘model, feet. ... ... ... ... ... ... .12.30 Wingchordatroot, feet. ... ... ... ... ... ..5.36 Wingchordattip, feet. ... ... ... ... ... .. l.33 Mea. naorodynamicchord, feet. ... ... ... ... .. 3.75 Sweepback of 25—percent—chord line, degrees . . . . . . . . 23.12 Dihedral of 25-percent—-chord line, degrees . . . . . . . . . 1.00 Twist about 25—pex-cent-chord line, degrees washout . . . . . 1+. O0 Airfoil section at root . . . . . . . . . . . . NACA 65(318)—. o19 Airfoil section at tip . . . . . . . . . . . . . . NACA 65,3—0l8 Eleven chord, percent wing chord (s. pprox. ) . . . . . .. . . . . l8
- 4. KACA RM No. A7013 3 Elevenspe. n,feet. .;. ..'. ... ... ... ... ... it-.75 Eleven ‘balance, percent of eleven chord aft of hinge 1ine(a, proI. ).. ... ... ... ... ... ... ... ho Mean of squared eleven chords, square feet . . . . . . . . . 0.309 SYMBOLS The symbols used in this report" are defined as follows: V’ free-stream velocity, feet per second V1 indicated airspeed, lmets _ 9 mass density of air, slugs per cubic foot q free—strea: m dynamic pressure Q-%pv2 >, pounds per sqrmre foot M Mach number corrected for blockage due to the model 1 V ‘II Kspeed of sound/ ' Mo uncorrected Mach number S wing area, square feet M. A.C. mean aerodynamic chord, feet (:92 mean of the squared eleven chords, square feet be span of eleven, feet _ lift CL lift coefficient -63- CD drag coefficient iiggﬁ / pitching moment about Cmmg’ pitching—mement coefficient K the center °f gravity / q_S M. .A. C. / I Che eleven hinge-moment coefficient ‘e3-97011 1111189 mﬂmﬂnt) 9. °e2 be or. angle of attack of rent chord corrected for ; ]et—boundary effects , degrees
- 5. .. ,.. i. :1 1+ NACA RM No. A7013 an uncorrected angle of attack of root chord, degrees Ac. a. n., 'le of attac‘: increment, degrees ACD drag coefficient increment 89 eleven deflection, degrees (Positive when trailing edge is down. ') St longitudinal trim flap deflection, degrees (Positive when trailing edge is down. ) c. g. ' longitudinal center-of—gravity location for neutral stability, stick fixed, percent of M. A.C. 7 specific heat of air at constant pressure divided by specific heat of air at constant volume CALIBRATION AND CORRECTIONS T0 DATA The dynamic pressure calibration used in these tests was obtained from a static pressure survey of the test section with the tunnel empty except for the survey apparatus, and the cali-— bration was corrected for the blockage d. ue to the model. As this correction was applied to the Mach number and to the dynamic pressure before the model tests, it was not necessary to correct the poefficients for a. change in dynamic pressure as was done in reference 3, which discusses the method of calibration and the blockage corrections. The blockage correction was 1:}. 1 + 2 M02 M = Me Me K -————————~--- + - 9/iaefls where K = 0.00637 + o. o5h5 CD Moments were computed about a center-—of—gravity at 25 percent of the mean aerodynamic chord and on the root chord line. Correc- tions to the angle of attack and drag coefficient due to the jet boundary were calculated from the charts of reference 14- and applied to the data. The corrections worm‘ A0. = 1.125 CL, degrees AC1) = 0.0197 Q?
- 6. mm. RM No. A7013 5 RESULTS AND DISCUSSION Lift and Longitudinal Stability The effects of compressibility on the aerodynamic character- istics of the YB-—1l-9 model, as shown by the tests in the 16-foot high—speed wind tunnel, began to appear at a Mach number of about 0.70 and were marked at a. Mach number of 0.775. Figure 3 shows curves of angle of attack, pitching-moment coefficient, and eleven hinge—momont coefficient against lift coefficient at four eleven angles for Mach numbers from O. ’-L0 to 0.725. Vibration of the eleven prevented the measurement of hinge moments at higher Mach numbers, but by rigidly restraining the eleven at ‘both ends it was possible to take force data up to a Mach number of 0.775. Lift and pitching-moment data with the eleven restrained are presented in figure 14. The variation of tching—mcment coefficient with lift coefficient in figure 1:. (a shows longitudinal instability beginning at a lift coefficient of about 0.35 and a Mach number of 0.70. At higher Mach numbers the instability progresses to lower lift coefficients, being strong at a Mach number of 0.775 and zero lift. The l: Lft—curve slope was also much reduced at a Mach number of 0.775. Tuft studies showed that the instability resulted from stalling of the outboard portion of the wing, while lift was maintained over a large area near the root. Figures 5 and 6 present tuft pictures for uncorrected angles of attack of 2° and ll-°. It is apparent that at 20 the tip stall occurred between Mach numbers of 0.75 and 0.775, while at 1:9 it occurred between 0.725 and 0.75. The variation with Mach number of lift coefficient, pitching- mement coefficient, lift-curve slope, and longitudinal stability is shown more clearly in figure 7. Curves of lift coefficient against Mach number are presented for constant angles of attack from -20 to 6° and pitch1ng—moment coefficient against Mach number is presented for constant lift coefficients from -0.10 to 0.1+0. In the range of model attitudes covered, the Mach numbers of lift and pitching—moment divergence varied from about 0.72 to 0.675. The curves of pitch. ing—momsnt coefficient against Mach number in figure 7 show a climbing moment at all positive lift coefficients for Mach numbers above 0.735 due to the loss of lift on the outboard part of the wing at the higher Mach numbers. Lift-curve slope and center-. of—gravity position for neutral stability are shown for level-—flight lift coefficients (160,000 lb gross weight) and zero pitching moment at altitudes of sea level, 25,000 feet, and 35,000 feet. At sea level the lift-—curv'e slope reached a maximum of 0.091-L per degree at a Mach number of 0.725 and then decreased rapidly with increasing Mach number as the outer portion of the wing stalled. The stick-fixed neutral point reached its most rearward position (314- percent of the M-A-C‘-) at a Nhch number of 0.725 and then moved rapidly forward with increasing Mach number. At 25,000 feet altitude the variation
- 7. 6 NACA RM No. A7013 of lift—ourvo slope and neutral point with Mach number was little different from that at sea level, but_at 35,000 feet the breaks in the curves occurred at a somewhat lower Mach number. The variation of eleven angle and eleven hinge—-moment coef- ficient for balance with indicated airspeed is shown in figure 8 for three altitudes and three model conditions chosen to show the effects of trim—flap deflection and boundary—layor transition on tho stability of the model. The curves show the presence of stick- fixed stability over most of the speed range for all the conditions shown except at the higher speeds where there is a small region of instability followed by extreme stability. The extreme stick—fi1od stability at the higher Mach numbers is evidently due to the loss of eleven effectiveness which is shown in a later figure. The effects of trim—flap deflection on the variation of eleven angle for bal- ance with airspeed varied with altitude. At sea level the main effect was a reduction in stick-fixed stability, while at the higher altitudes the stability was increased at the lower speeds, followed by a region of reduced stability. At the highest speeds the effects of the trim flap were small. Fixing the boundary—layer transition on the upper surface at 15 percent of the chord had an effect similar to that of the small deflection of the trim flap. ‘B10 variation of eleven hinge—moment with airspeed indicated stick—froe stability over the speed range covered in the test, and the stability increased with increasing altitude. Deflection of the trim flap affected the hinge——moment coefficient in much the same manner as it did the eleven angle for balance. Fixing the boundary-—layor transi- tion at l5 percent of the chord on the upper surface decreased the stick—free stability slightly over most of the speed range. Longitudinal Control Figure 9 presents the variation of pitching-moment coefficient and eleven hinge-moment coefficient with eleven angle at constant lift coefficients for Mach numibers from 0.140 to 0.725. Hinge moments were not measured at Mach numbers high enough to show any pronounced effects due to compressibility, however, for negative eleven angles the curves show an increased negative value of 50110/650 at a Mach number of 0.725, indicating the beginning of a decrease in balance effectiveness. The variation of pitching- moment coefficient with eleven angle for Mach numbers of 0.70 to 0.775, with the eleven rigidly restrained, isprosentod in figure 10; and in figure 11 the effects of longitudinal--trim—fla. p deflection on pitching-moment coefficient are shown. Eleven and trim—flap effectiveness were measured from figures 9, 10, and ll and are plotted against Mach number in figure l2. The data indi- cate that at zero lift coefficient a rather rapid decrease in eleven effectiveness began at a Mach number 0.725, while at a lift coefficient of 0.1+ an increase in effectiveness began between Mach numbers of 0.65 and 0.70. However, these values are for zero
- 8. NACA RM No. A7013 7 eleven deflection, and the curve of pitching—moment coefficient against elevon angle for a lift coefficient of 0.1+ at 0.725 Mach number in figure 10 shows that at small elevon deflections the effectiveness was reduced and it may be expected that the eleven effectiveness will decrease rapidly as the Mach number increases above M = 0.725. The trim flap also began to lose effectiveness at a Mach number of 0.725. Drag Characteristics The variation of drag coefficient with lift coefficient for several Mach numbers is presented in figure 13. The drag coeffi- cient at zero lift and 0.1410 Mach number (Reynolds number about 8.5 X 105) was 0.011 compared with 0.012 reported. in reference 2 for a Magh number of about 0.12 and 15:. Reynolds number of about 7.5 x 10 . comcmsxons, The results of tests of a 1/7-scale model of the 23-119 air- plane in the Ames 16-foot high-speed wind tunnel led. to the following conclusions: 1. Static longitudinal instability began to appear at 0.70 Mach number and a. lift coefficient of 0.35. At level—flight lift coefficients the stability decreased rapidly as the Mach number increased above 0. 725. 2. The loss of lift on the outboard portion of the wing resulted in a climbing moment for all -positive lift coefficients at Mach numbers above 0.735. 3. The effectiveness of both the eleven and the longitudinal trim flap began to decrease with increasing Mach number at Mach numbers between 0.70 and 0.725. Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics, Moffett F 1815., Calif. . A1’P’°”°‘1‘ . Robert c. Robinson, _, Aeronautical Engineer. . , ’ 5*-: f!< Donald E. Wood Aeronautical Engineer.
- 9. NADA E4 No. A7013 CES Sivelle, James 0., and Burgess, Jack‘: Tests in the NACA l9-—Foet Pressure Tunnel of a l/10.75-Scale Model of the Northrop XIB-35 Tailless Airplane. IIAGA CMR, Feb. 19113. Teplitz, Jerome, Kayten, Gerald G. , and Cancro, Patrick A. : Tests of a l/7—Sca. le Semispan Model of the 13-35 Airplane in the Langley l9—Foot Pressure Tunnel. NACA OMB No. L5L27, 19%. Nissan, James M. , Gadeberg, Burnett L. , and Hamilton, William T. : Correlation of the" Drag Characteristics of a P——51JB Airplane Obtained from Bigh—Speed Wind—Tu. nnel and Flight Tests. NACA ACR No. J-LKO2, 1916. Silverstoin, Abe, and White, James A. : Wind-Tunnel Inter- ference with Particular Reference to Off-Center Positions of the Wing and to the Downwash at the Tail. NACA Rep. No. 5h7, 1935.
- 10. EAOA H4 No. A7013 FIGURE LEGENDS Figure l. — The l/7-scale semispen model of the EB-1:9 airplane mounted in the Ames 16-foot high--speed wind tunnel. (a) Rear view. (b) Front view. Figure 2.- The J. /7—soa1e seniispan model of the Northrop YB-1:-9 airplane. Figure 3.- The effects of eleven deflection on lift coefficient, pitching-moment coefficient, ‘ and eleven hinge--moment coefficient. 1/7—sce.1e rs-1+9 model. (a) M = o.1+o. Figure 3.- continued. ('b) M = 0.55. 0.65. Figure 3.- Continued. (d) M = 0.70. Figure 3.- Continued. (c) M Figure 3.- concluded. (a) M = 0.725. Figure II-. — he effects of eleven deflection on lift coefficient and pitching-moment coefficient with the eleven rigidly restrained. l/7—ecale YB—’-+9 model. (a) M = 0.70. Figure 34-. — continued. (13) M = 0.725. Figure h». — Continued. (c) M = 0.75. Figure 1I». - Concluded. (d) M = 0.775. Figure 5.— Tufts on the 2349 semispan model. a. = 20. Figure 6.—- Tufts on the '13-’-I-9 semiepan model. a. = ll-°. Figure 7.-- The effects of compressibility on lift coefficient, pitching—moment coefficient, lift-curve slope, and longitudinal stability. 1/7—scale ns—u9 model. Figure 3.- The variation of eleven deflection and eleven hinge- moment coefficient with indicated airspeed for balance. l/7—sce. 'Le IBJI-9 model. Figure 9.— ‘lbs variation of pitching—moment coefficient and eleven hinge-moment coefficient with eleven deflection. l/7—scale IB—1I-9 model. (9.) M = 0.1+. Figure 9.— Continued. Cb) M = 0.55. 3
- 11. NACA RM No. A7013 Figure 9.— Continued. (c) M 0.65. Figure 9.— Continued. (6.) M 0.70. Figure 9.— Concluded. (e) M = 0.725. Figure lO. — The variation of pitching-moment coefficient with eleven deflection with the eleven rigidly restrained. l/7-scale YZB—1I-9 model. ' Figure ll. — The variation of pitching—-moment coefficient with trim—fle. p deflection. l/7-scale YB«—1¢-9 model. (5.) M = 0.1+O. Figure ll. — Continued. (b) M . . o_.65. Figure ll. — Continued. (c) M = 0.725. Figure ll. — Concluded. (d) M = 0.75. Figure l2.— The effects of compressibility on the effectiveness of the eleven and the longitudinal trim flap. 1/ 7—sca. le YB—1+9 model. be = 0°, 5., = 0°. Figure l3.— The variation of drag coefficient with lift coefficient at several Mach numbers. 1/ 7—ecale YB—1L9 model.
- 12. NACA HM No. A’ ' '. n'u'l‘. .mr. ;r -5 - (a) Rear view. (13) Front view. Figure l. — The l/7—eca. le semiapan model of the Y. B—lI»9 airplane mounted in the Ames 16-foot high—epeed wind tunnel. Nu-noun. muse‘ I0 zwrrnzz ARES LEONAITHCAL LLEKATOBY — KOFI"l'l'l' YELD, CALI)‘.
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- 22. "ax-vi . ’ "E ’ IlLlJ! §J. . III III “I I III Ii". : I ﬁg] . ;.. "II; '. 3; . . IIIIIIIIIIIII ‘’ II II IIIIIII III III , -III - __ _IIIII_ ‘; ‘iIIIIIIIiI. . IIIIIIIIIIIIIIIII IIIIIIII-II I . .IIIIIIIIIIIIIIIIII IIIIIIIIIIII IIIIIIIIIIIIIIIIIIIIIIII ‘III III. —@IIIIIIII I III -I. , III IIIIIIIIIIIIIIIIIIIII , _ I IIIIIIIIIIIII . IIIIII II ' °. I _ _ '§ gagggéﬁﬁﬁg E W "3: E . ... = x="_= § mill '. .'I ﬂ%- . IIIIiIiIIII§IIIIIIIIIIII, IIIIIIIIII IIIIIIIIIIIIIIIIIIIIIIIIIEIIIIIIIIEIIIIIIIIIIIIIIIIIIIIIIIEIIIIIII ﬁﬁl ' IIIIIIIIIIIIIIIIIIIIlﬁ IﬁIIIIIIIIE{I%IIIIIIIIIIIIII%%ﬁIIIII§I§IIII! IIII§II IE L§}I! iIIII§iIIIIIIIIIIIIIi I F I IIIIIIII . ..HI§IIIﬂIIIII IIIE IIIIJIIIIIIIILIII _ ’ IIIIIIIIII II% = ' "IIII‘IIIIIII§IIIII"‘IIIIIIIIIIIIIII'IIII IIII'IIIIII ‘I I I I I I . ':'I}‘I"' IIII IIIIIIIIIIIEIIQW H ﬁ%%Wﬁw%qH%m¥; mww§%g§ III ' _ : _ .615” ‘I'i__I _§IiI‘. I_iI. . .;! I II IIIIIIIIII'IIIIIIIIIIIIIIIIIIIIIIIII Iﬁ%m%%mmﬂmm%%%ﬁ%mmﬁmﬁﬁ%mIwmnWWII mmﬁmﬁm . n—i = .I. .-'‘§ " I‘ :1.‘ III I I. iégi IIIIII* BII! . III. I x- M‘ II'II ' ‘ ‘ 3| E-3% Egagagﬁ E §%= ggga . gaaaag ' 3% 5% "3 E E E I‘ II II Ei-ﬁg E I I III"II Li IIIIIIIIII EEE E §_ Ia » I IIIII I I IIIIIIIEIIIIIIIITIW . .‘ I -I" IEIIIIII . .. mm IIEIIIIII _! III' IIIIIIIII. {III IM 3% EE . .— . .‘. .:. gaaggaga “ L" EEEE 12-‘ a ' II = — IIIIIIIIIIIIIII IIIIIIIIII IIIIIIIII II I IIIIIIII II IIIIIIIIIFIIIII IIIIIIIIII ﬁE%%%ﬂ%%mmmmmImmImmﬁm%%ﬂﬂm%ﬂmmmmImmmm m IIIIIII I III IIIII| I‘III'IIIiIIIIIIIiIIIIiIIiIIIIIIII1IIIII! III| IIIIIIIIIIII"IiIiII1II! ! IIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIII . IIIIIII ' III: 'IIIIIIiIIIIiiIIifg%IIIii§IIIIlIIIiIIIIIIIlIII. §IIIIIIIII' IIIIIIIIII IIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIII IIIIIIIIIIIIIIIIIIIII | IiIEHﬁ| IHIIIIﬂE: .ﬂ§IIIIIi§§‘I ‘IIIEIIIIIIIIIlIIIIIIIE: I§I! IiIII" IIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIII I w%ﬁWFW%ﬁ%%ﬁ%@@I-“%%%%W%%%%%%@%%ﬁW%%$%“ . . . I a: .:. . E. I : : ! .I I : . II I . II I . .: IIIII IIIIIIIIJIIIIIIIIIIIIIIIIII gﬁiﬁgglim I IIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIEIIIIIIIIIMIIIIIIIII III I IIIIII III IIII II I ﬂI? "I%ﬂ%ﬁ%ﬁ , {I mé%%%%%%WhI' III 3ﬁ? ?%'%£%'%%§%I%a” I I_ Iﬂmw I . III I I . _ K . ’ - II I- I‘ - ‘ ‘ 7 A .4 I - I w _r, . - WIIIII '5 I‘: I *I''''-? I - F I‘ -, - . I ' , 7,; ‘ I I_ _. ._5. I: I" . am 3 < W‘ . .I'' . - _ , I « I II U”. H. “I §% E; EE :55: . 'I . II I. - 5% 925 mg §§a§ II 'I ‘ﬂ IEII IIIII Iﬁ . EIII 5%? 2%; -*7.’ EE E
- 23. ﬁm- mm. 4 . .u -F . .. . ..ua. u_. “Fm nr E‘ . ; _ ‘ﬁIr. a:| ::'a: nn: .mﬁ‘ . ﬂu .1- '. .'! i%! a» ; '!i! !| . _ liiiigigl ﬂu ' I! l§I: ilii! §!§Eill! !éiﬁé! lli: l'li*’. i![ae'ii ' Ellliiilﬁili I HM. r. u. ... ... .. mm. - . - um . ... :. InLHu. M.: .x W“ I “H s . . I» . H. - . .. ..W. ... - -H. w. .mm. .m. . W% . W . . ,.. .., . ..u. ..H. .W. .W%. I¢ mm -. . . . Imm. .. «water. .1.. ... . . ..: W4 . . . . w. - - -. . --. .- W J. mmﬂﬂ. ... .._. m “. ... HWW. .. WWW . n E 4. . .,. ... ._. ..m. .., ... . . !. ... ,.. .._ . . mxmﬂm undﬂmm .1. w. ..4.. .. ”. ... ..W. G.wm. ._. .»r. ... .m. mWw . ... ... : . ... _,. ..m. ”.. .. . ... -
- 24. NACA RM NO. A7013 . .(. ..o; ... ... l.. .. .. .. lnI. iau . . ! Il. .|: .l. u a la _. ..I. rIlII . (a) M = 0.725. (c) M = 0.775. Figure 5.— Tufts on the IB—-M-9 aemispan model. 0:. 2°. VKTIONAL ADVIKRY COMi(I'FIEE FOR AﬂO. '4'Al"l'N'S ‘VI-A‘ l§0§Al1'ICI. Ia I-LEIKHJII — KCFTZTT FIELD. CHI, -
- 25. IIACA RM N0. A7013 0.70. M (a) ('b) M = 0.725. ’+°. C6 Figure 6.— Tufts on the YZB—J+9 semispan model. NATIONAL ADVISORY COIﬂ£l. '.| 'l"ll FOR Aa0N. |l"l'H"S -Ibi | F|(l). ‘4Al‘I‘1CA. |. LAIOKAWII-Y — MDPFPET FIELD. CALIF.
- 26. A . nwa; : I . u-nwmllm mm . null . ruler. .m . I' . u . an nae vn-. : -mu nu J-unnn .1 . c. u we MAI
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- 34. Fl-CPI .0!‘ . V H . 00 I139?‘ I JIWIIII «vs»! «-1.: u. . , )nl-Inns null and I -ul-nun: I nu ll 1-ul NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
- 35. Otuu. .1»! .1 M .024 I-Jet! .v l<'I'IaI'| 44.41 4-um . u.; .IaoIar-nu -um mm c I-I'l-J‘-IIIJIIN : .z. I. KI mu E9. ~. &'75.»'l= §—-. --— , 5 4 ' ""_L _
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