RESEARCH MEMORANDUM. NATIONAL ADVl SORY COMMITTEE FOR AERONAUTICS. AMD TRBrnG-EDGE FUR WASHINGTON. Lang1ey Aeronautical Laboratory LmgLey Field, Va.

Transcription

1 RESEARCH MEMORANDUM LOW-SUBSONIC INVESTIGATION TO DETERMINE THE CHORDWISE PRESSURE DISTRIBUTION AND EFFECTIVENESS OF SPOILERS ON A THIN, LOW-ASPECT-RATIO, UNSWEPT, UN'TAPERED, SEWISPAN WIMG PlND OM THE WIMG WITH LEADmG- AMD TRBrnG-EDGE FUR By Wwin R. Croom Lang1ey Aeronautical Laboratory LmgLey Field, Va. I NATIONAL ADVl SORY COMMITTEE FOR AERONAUTICS WASHINGTON April 23, 1958 t?+

2 1J TtACA RIvi ~58~05 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RESEARCH WORANDUM LOW-SUBSONIC INVESTIGATION TO DETERMINE THE CHORDWISE PRESSURE DISTRIBUTION AND EFFECTIVENESS OF SPOILERS ON A THIN, LOW-ASPECT-RATIO, UNSWEPT, UNTAPERE=D, SEMISPAN WING AND ON THE WING WITH LEADING- AND TRAILING-EDGE FLAPS By Delwin R. Croom SUMMARY An investigation was made in the Langley 300-MPH 7- by 10-foot tunnel to determine the effect of spoilers on the static longitudinal and lateral aerodynamic characteristics and the chordwise pressure distribution on a thin, untapered, unswept, semispan wing having an aspect ratio of 3.33 and NACA 65~004 airfoil sections. The wing was equipped with leading-edge flaps and trailing-edge flaps. Results of the investigation, without discussion, are presented in the form of static longitudinal and lateral aerodynamic characteristics and tabulated pressure coefficients, section normal-force coefficients, and section pitching-moment coefficients. Considerable interest has been shown in the use of spoilers for lateral control and gust alleviation. The spoiler is more effective as a lateral control when it is located in a more rearward position, and, as pointed out in references 1 and 2, the spoiler is more effective as a gust alleviator (reduces the lift-curve slope) when it is located in a more forward position.

3 2 NACA RM L58~05 The present investigation was made in the Langley 300-MPH 7- by 10-foot tunnel to obtain the static longitudinal and lateral aerodynamic characteristics and the chordwise pressure distribution at the 60-percentsemispan station over a thin, low-aspect-ratio, unswept, untapered, semispan wing with a 10-percent-chord projection spoiler of various spans located at various chordwise locations on the wing with leading- and trailing-edge flaps. The results of the investigation are presented without discussion. Pressure coefficients, section normal-force coefficients, and section pitching-moment coefficients are presented in tabular form. SYMBOLS chord, ft plain-wing chord, ft trailing-edge-flap chord, ft leading-edge-flap chord, ft wing span, ft wing area, sq ft lift coefficient, Twice semispan lift qoos slope of lift curve of basic model (measured at a = 0 ) per degree slope of lift curve of model with spoiler (measured at a, = 0') per degree drag coefficient, Twice semispan drag qoos jet-boundary correction applied to drag coefficient pitching-moment coefficient of wing referred to wing quarterchord, Twice semispan pitching moment qwscw

4 NACA RM L58BOFJ 3 rolling-moment coefficient, Rolling moment caused by spoiler qwsb Yawing moment caused by spoiler yawing-moment coefficient, - %Sb pressure coefficient, ptlw P (where subscripts u and qm 2 denote upper and lower surfaces, respectively) 2f distance from wing quarter-chord to hinge line of trailingedge flap measured parallel to trailing-edge-flap chord, ft distance from wing quarter-chord to hinge line of leading-edge flap measured parallel to leading-edge-flap chord, ft x X f X N Y 't,m P '4, P longitudinal distance, ft distance from hinge line of trailing-edge flap to center of load on trailing-edge flap, ft distance from hinge line of leading-edge flap to center of load on leading-edge flap, ft lateral distance, ft free-stream total pressure, lb/sq ft local static pressure, lbtsq ft 2 f ree-stream dynamic pressure, - 2 mass density of air, slu@;s/cu ft ", lb/sq ft v~ free-stream velocity, ft/sec 6f trailing-edge-f lap deflection (positive direction, trailing edge down), deg % leading-edge-flap deflection (positive direction, nose of flap down), deg a angle of attack of chord plane set in tunnel, deg jet-boundary correction applied to angle of attack, deg

5 NACA RM L58BO5 ac c NJ f corrected angle of attack, deg section normal-force coefficient of trailing-edge flap based on trailing-edge-flap chord c section normal-force coefficient of leading-edge flap based 37 N on leading-edge-flap chord 'N,w' C N,w section normal-force coefficient of that portion of wing between leading-edge and trailing-edge flaps based on plain-wing chord wing section normal-force coefficient based on plain-wing chord (chord force of leading-edge flap and trailing-edge flap neglected), Cm, f 'm, N Cm,w' C m,w section pitching-moment coefficient of trailing-edge flap based on trailing-edge-f lap chord (moments taken about trailingedge-flap hinge line) section pitching-moment coefficient of leading-edge flap based on leading-edge-flap chord (moments taken about leading-edgeflap hinge line) section pitching-moment coefficient of that portion of wing between leading-edge and trailing-edge flaps based on plainwing chord (moments taken about wing quarter-chord) wing section pitching-moment coefficient based on plain-wing chord (moments taken about wing quarter-chord; chord force of leading- and trailing-edge flaps neglected), MODEX AND APPARATUS The model was tested in the Langley 300-MPR 7- by 10-foot tunnel by means of the semispan technique with the ceiling of the tunnel as the reflection plane. The geometric characteristics of the semispan wing used in this investigation are given in figure 1. The wing had O0 of sweep, a taper ratio of 1, an aspect ratio of 3.33, and

6 NACA RM ~58~05 5 NACA 65~004 airfoil sect ions parallel to the f ree-stream direction. The wing was equipped with leading-edge and trailing-edge flaps. The leadingedge-flap chord was 15 percent of the wing chord, and the leading-edge flap pivoted about the lower surface along the 15-percent-chord line. (see fig. 2. ) For the deflected condition the break in the upper surface was faired to an arc of a circle. The sealed trailing-edge-flap chord was 25 percent of the wing chord. (see fig. 2. ) The leading-edge flap, the trailing-edge flap, and the wing were constructed with flush-surface pressure orifices located at the 60-percent-semispan station. The 10-percent-chord projection spoiler used in this investigation was made of wood to the dimensions shown in figure 2. The spans, spanwise locations, and chordwise locations of the various spoiler arrangements are indicated in figure 2. TESTS The tests were made in the Langley 300-MPH 7- by 10-foot tunnel at a dynamic pressure of approximately 25 pounds per square foot. For these tests, the leading-edge flap was deflected at 0' or 30 and the trailing-edge flap was deflected at 0' or l5o. The spoiler had a projection of 10 percent chord and was tested as a full-semispan control, a half-semispan inboard control, and a half-semispan outboard control hinged at the lo-, 30-, 50-, TO-, and 90-percent-chord locations. (see fig. 2. ) The maximum angle-of-attack range for this investigation was from about -12' to 24'. Chordwise pressure distributions were obtained at the 60-percent-semispan station. CORRECTIONS The jet-boundary corrections applied to the data of this paper were obtained by the method outlined in reference 3 and are as follows: The blockage correction as applied to the dynamic pressure was obtained by the method outlined in reference 4.

7 RESULTS The results of this investigation are presented without discussion. The order of presentation in the figures and tables is as follows: Static longitudinal and lateral aerodynamic characteristics: Figure ~alf-span inboard spoiler; SN = 0'; S~ = o Half-span outboard spoiler; % = 0'; Sf = 0'... 4 Nl-span spoiler; % = 0'; Sf = 0'... 5 Half-span outboard spoiler; SN = 0' and 30'; 6f = 0 and l'jo. 6 Nl-span spoiler; SN = 0' and 30'; Sf = 0'... 7 Lift-curve-slope comparison... 8 Chordwise pressure distribution: Basic model; GN = 0'; tjf = o Half-span inboard spoiler; SN = 0'; Sf = to 14 ~alf-span outboard spoiler; SN = 0'; 6f = o to 19 ~1-span spoiler; SN = 0'; S~ = o to 24 Half-span outboard spoiler; SN = 30'; Gf = 0' and 15' to 28 Full-span spoiler; % = 30'; Sf = 0' to 32 Tabulated integrated section data: Table Half-span inboard spoiler; % = 0'; Sf = 0'... I Half-span outboard spoiler; FN = 0'; Sf = o0... II Nl-span spoiler; % = 0'; Sf = 0'... I11 Half -span outboard spoiler; SN = 30'; Sf = 0' and 15'... IV ~ull-span spoiler; 6~ = 30'; tif = o0... v Since the location of the pressure orifices is at the 60-percentsemispan station, the section data for the full-span and half-span outboard spoilers show the effect of the spoilers on the section loading; whereas the section data for the half-span inboard spoiler give an indication of the carryover of the load outboard of the spoiler. It should also be pointed out that these data are for only one value of spoiler projection and, although they are useful in determining the

8 NACA RM L58BO5 7 effects sf the several variables, they are not necessarily applicable to the design of a control surface that uses small deflections. A method has been developed by the Data Reduction Branch of the Langley Instrument Research Division whereby the IEM type 407 accounting machine can be used to obtain approximate plots of data as they are tabulated. This tabulator was used to obtain the plots of figures 9 to 32, and even though they are not as accurate as may be desired, they do show trends of the pressure distribution. An accurate tabulation of these data is provided at the left of these figures. Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, Va., January 14, Croom, Delwin R., Shufflebarger, C. C., and Iluffman, Jarrett K.: An Investigation of Forward-Located Fixed Spoilers and Deflectors as Gust Alleviators on an Unswept-Wing Model. NACA TN 3705, Croom, Delwin R., and Ruffman, Jarrett K.: Investigation at Low Speeds of Deflectors and Spoilers as Gust Alleviators on a Model of the Bell X-5 Airplane With 35' Swept Wings and on a High-Aspect-Ratio 35' Swept-Wing-Fuselage Model. NACA TN 4057, Polhamus, Edward C. : Jet-Boundary-Induced-Upwash Velocities for Swept Reflection-Plane Models Mounted Vertically in 7- by 10-Foot, Closed, Rectangular Wind Tunnels. NACA TN 1752, Herriot, John G. : Blockage Corrections for Three-Dimensional Flow Closed-Throat Wind Tunnels, With Consideration of the Effect of Compressibility. NACA Rep. 995, (supersedes NACA RM A7B28. )

9 NACA RM L58BO5 WBLE I.- UPPtGNLT'Z SEl'IVIl DAli\ pf = 00; 6~ = oo; has-span inboard spoiler I, ,8353, ,0512-8, ,8107,1447,0574- a5769, ,7538,1509,0547-,7309, ,2729,0989-,8171, ,1147, Spaller hlnged at 0.7~ so ,3547,0840-,3469,2254 a0200,9781,6458,0596-,2806, ,8493 $0376-,2227 a5520, ,0054,1112 ~7508, ,7253,0792,0646- Spoiler hrnged at 0.9~, ,3811-,2351-,2312- a ,0546,5271,3489 $1751-,1224,1785, ab569,1494-,1139,3390, ,9289,1090-,0994,5120 a ,0403-,0717,6764, r7836,0604,0370,7855, ,8444,1332 a0133,8423 a ,8191,2423,0222-, ,7091,3386 a0587-

10 NACA RM L58BO5 [6* = ' 0 m m 11.- nmmm SECTS011 QAm ; rji1 = 0'; half-span outboard spoiler I 0,0361, a0826, st443,1149-,5915-,7422-,2345,4067-6,3569 a1310-, ,3231,6211-8,4301, ,3640, ,5429 a1293-,4409,5191, ,7164, ,3913, a7736 a , ,7560, , , ,0708.I022, , ,3196 a3326, ,0849-,9231,6538,3834, ~6044, a9476 a4045, , ,6724,3390, , ,6778,3456,1325- Spoiler hinged at 0.X 4 e0117-, ,2398 r1072-6,1084,0093- a ,2524 ~ , ,8436 ~3120, , ,8991,3895, , ,4250 ~ , ,7227,4024, , ,7231,3902 a15b3-

11 NACA RM L58~05 TABLE XI.- INTxwArn smmoii DAm Cf oo; % = oo; full-span spoiler I -, deg C ~, Cm,w ~ 'N,N ',N '~,f 'm,f Spoiler hinged at O.lc 4- ~ ,2106,3387-0,0365-, ,1067-,2656, ~1342 a1257-, ,3056,5675- b,1932,1240- a7737-,9776-,3213 a6061-8,2520, ,3262, , ,3260 a a4637,1108-,3825 a3800,3612, ,5736,1084-,8497,8720,3826 a ,7145, ,4277, ,8571, ,4708,9941- Spoiler hinged at 0.3~ 4-,4900-, ,9840-,1379 a0555-0,2769-, ,7151-,1809,0757-4,0428-,0847-,3905- al88o-,2658,1136-6,0468,0823- m0745- ~0200,2836, e'1561 a0791- a2797,2498 a2976,126k- 10,3268,0805-,7309 a4916,3422, ,6310 a a6600,3744, ,7551, ,6618,3725, a ,4090, a8360, a1244,4524,1803- Spoiler hinged at 0.5~ 4-,6092-, ,1519,0622-0,3887- a0160-,8769- $5267- ~1564, $1486-, ,2156,0926-6, ,3427,2569,2392 e a ,4996,2545 a a ,6782 a a6844,3632, el391,3671 e ,8437 a1052-1,4057 ~7440 a a ,4159,1612-

12 lkei.3 N.- JNEGRR'JED swmoii DATA [w-span out- spoiler] a, deg CN,w 'm,w 'N,N ' m. ~ 'N,f 'm.f e, deg 'N,W 'm,w CN,N ' m, ~ ' ~, f 6f = OO; 6N = 50: Spoiler hinged at 0.5~ 6f = 0'; % = 30'; Spoiler hinged at 0.9~ - 'm,f 8-,6978-, ,7520- e a5683-, ,7049- a1969,1205-0, ,6218-, , ,4516-,3095,140W e.i382.o68l-,1838-,2253-, ,0858-,6387 a1244,3776, ,7348, ,7351,4935, , , ,5646 e , , , , ,5344 0,6076-,1553, ,7459-,4118 4,2881-,1083,5305- ~4053-,6578- a3802 8,0100 a , ,3383, , ,6099, ,9173,4070-, , ,1514-, , ,3096,0434- hf = 0'; 6~ = 30"; Spoiler hinged at 0.7~ hf = 15'; b~ = 30'; spoiler hinged at 0.X 8-,8486-, ,7556-,2552-, , ,7138-,0152-,1947 0,4452-,0192,9723-,6058-,2144, ,0055-,5233-,3991- alb70,2955-8,0970,0045-,0187-,1578-,1098, ,3912,0053- ~9069,2480, bb- 16,7183 ~ ,8747,3408, ,9141, ,3929, , ,4125, ,7348-, , , ,5090-, ~6756-, o,2545-,0540-,8128-,5187-,3804.iaoa- 4,0900,1205-,4054-,3307-,5376, ,4631,1487-,1428,1213-,6300 e ,7678 a ,3653,7182 $ , ,9787,8099, , ~9000 a3b42-

13 NACA RM ~58~05

14 NACA RM ~ Moment center Quur ter-chrd Tunnel ceilhg End p lo te ---- Pressure-or i f ices IOCU t ion Figure 1. - General characteristics of model. All dimensions are in inches unless otherwise noted.

15 NACA RM L58~05

16 NACA RM L58BO5. (a) Cm, CD, and ac plotted against CL. Figure 3.- Effect of chordwise location of half-span inboard sp oiler the aerodynahic characteristics of the semispan model. 6~ = 00; 6f = 00.

17 NACA RM ~58~05 (b) Cn md C2 plotted against a,. Figure 3. - Concluded.

18 NACA RM ~58~05 (a) Cm, CD, and ac plotted against Cb. Figure 4.- Effect of chordwise locat2on of half-span outboard spoiler on the aerodynamic characteristics of the semispan model. 6~ = OO; 8f = 0'.

19 NACA RM ~58~05 (b) Cn and Cz plotted against a,. Figure 4.- Concluded.

20 NACA RM ~58~05 (a), C, CD, and ac plotted against CL. Figure 5.- Effect of chordwise location of full-span spoiler on the aerodynanic characteristics of the semispan model. 6~ = 0'; 6f = 0'.

21 NACA RM ~58~05 (b) Cn and C2 plotted against a,. Figure 5.- Concluded.

22 NACA RM ~58~05 Spoiler hjnge line, Sf, deg SN, deg percent wlng chord Ploin wing 0 30 Spoiler off Spoiler off (a) C, CD, and ac plotted against CL. Figure 6. - Effect of chordwise location of half -span outboard spoiler and effect of trailing-edge-flap deflection on the aerodynamic characteristics of the semispan model.

23 NACA RM L58~05 (b) Cn and C2 plotted against ac. Figure 6. - Concluded.

24 NACA RM L58~05 (8) ern, CD, and ac plotted against CL. Figure 7.- Effect of chordwise location of full-span spoiler and effect of nose-flap deflection on the aerodynamic characteristics of the semispan model. 6f = OO.

25 NACA RM ~58~05 SN, deg Spoiler hinge line, percent wing chord A b 30 a (b) Cn and Cz plotted against ac. Figure 7.- Concluded.

26 NAC A

27 NACA RM ~58~05

28 NACA RM ~58~05 (d) a = -6'. Figure 9.- Continued. (e) a = -4O.

29 NACA RM ~38~05

30 NACA RM L58~05

31 NACA RM ~58~05 (k) a = 8O. (2) a = 10'. Figure 9.- Continued.

32 NACA RM ~58~05 o Upper surface Figure 9.- Continued.

33 NACA RM L58BO5 (0) a = 14O. (p) a = l5o. Figure 9.- Continued.

34 NACA RM ~58~05 (q) a = 16'. (r) a = lto. Figure 9.- Continued.

35 NACA RM L58~05 0 Upper surface (s) a = 18O. Figure 9.- Concluded.

36 NACA RM ~58~05

37 ppp--p- ppp--pp NACA RM L58~05 0 Upper surfuce till! ii%q IIB 1.W5, , , @2b5 e ew(t--" ". - -' o Lower surfuce n o A 2.371, , , , , , o , a , , , s 1.354, llol PI D %," m D , m s LL O70BI,ilO0 EBf ~ , n 0 n 0 D 0 0 o n 0 rn ::::I 1:los 8:8? kp8i 8E8 ::2E ) rn L1 0 D D n e a L1 D lab Q m OPbLS m 0 iio ?a rrsad--hi6fioo(te, m 0 n 0 I0 (d) a = 6O. (e) a = 8'. Figure 10.- Continued.

38 NACA RM ~58~05 (f) a = 10'. Figure 10.- Continued.

39 NACA RM L58~05

40 NACA RM ~58~05 a- -P 0 % 0 Oal w rl 0 k -P LO P 0 +I, 0 I1 rl R C4 LO m +' F:.rl * r l o ala 0 2 %I II El 0 6 kld - g 2 o a P d a 0 a,

41 NACA RM ~58~05 Figure Continued. (e) a = 8'.

42 NACA RM L58~05 (f) a = lo0. (g) a = 12O. Figure 11.- Continued.

43 NACA RP/I ~58BO5

44 NACA RM ~58~05

45

46 NACA RM ~58~05

47 NACA RM L58BO5 (i) a = 16O. (j) a = la0. Figure 12.- Concluded.

48 NACA FiM ~58~05 47

49 NACA RM L58B05

50 - NACA RM ~58~05

51 NACA RM ~58~03 (i) a= 16'. (j) a = 18'. Figure 13.- Concluded.

52 NACA RM ~58~05 $ c 0 a, o\ 3 6 rn. rn+' * 0 a,d 3 1 $2 " $ $ d.$ 2 - k d oa, 82 I f rt 0 PI rn

53 NACA RM ~58~05

54 NACA RM L58BO5 (f) a = lo0. (g) a = 120. Figure 14.- Continued. Lb

55 NACA RM ~58~05

56 NACA RM ~58~05 * w4-m 0 -o*" Z FECE 0 w-0., -40 mr-.,, An -0.- n.+_(p** *me """A"

57 NACA RM L58~05

58 8 J NACA B4 L58~05

59 NACA RM ~58~05 (i) a= 16'. (j) a = 18'. Figure Concluded.

60 NACA RM ~58~05 d 8 0?Q -P Pi k ri k cdo 0 d -P O.=, q-, II 0 a, 0 ri II 0 -P n k P cod v aj.-.-i= 0 -P 0 0 rl I1 Pi R m co +'.a!% c; B cd k -P g 3 a E! $ k -P -P z m.=,.rl 0 0' $ 6 f I m-p m cd II a, k d 6 ma, M - n g.5 cd.rl d a s k k a, 0 ri e -$ I $ V3 ri a, 5 M.ri kc

61 CP CP CP (c) a = 4O. (6) a = 60. (e) a = 8'. Figure Continued.

62

63 NACA RM ~58~05 (i) a = 16'. (j) a= 18O. Figure Concluded.

64 NACA RM ~58~05 II +' 0 =I4 0 d a, -I= 0 +' 0 0 rl II F4 Zi m 0 t-' F:.d,4 g, 3 %4 0 0 b 2 * a $ k +' %.- a; a,. 2 O t-' m cd a, k a PC a, a, 2' m a d -4 d 3 a k 8 3 d -4 U 0 I $4 m

65 NACA RM ~ 58~05 0 (d) CL = 6. (e) a = 8'. Figure Continued.

66 NACA RI\I ~ (f) a = 10'. (g) a = 12O. Figure 17.- Continued.

67 NACA RM ~58~05 (h) a = 14O. (i) a = 16'. Figure 17.- Continued.

68 NACA RM ~58~05 o Upper surface Lower surface 'P (j) a= la0. Figure 17.- Concluded.

69 NACA RM ~58~03

70 NACA RM L58~05 (c) a = 4O. Figure Continued.

71 NACA RM ~58~05 (e) a = 8O. Figure 18.- Continued.

72 NACA RM L58~05 (g) a = 12O. (h) a = 14O. Figure 18.- Continued.

73 NACA RM L58BO5 (i) a = 16O. (j) a= 18O.. Figure 18.- Concluded.

74 OJ NACA RM ~58~05

75 NACA RM L58~05

76 NACA RM ~58~05 (g) a = 12'. Figure Continued.

77 NACA RM L58~05

78 NACA RM L58~05 Ti.

79 NACA RM L58~05

80 NACA RM ~58~05

81 NACA RM L58~05 (i) a = 16'. (j) a = 18'. Figure Concluded.

82 J NACA RM ~58~05

83 NACA RM ~58~05

84 NACA RM L58~05

85 NACA RM ~58~05 o Upper surface (i) a = 16O. Figure Concluded.

86 NACA HL/I L58~05

87 NACA RM ~58~05

88 NACA R v l ~58~05

89 Figure 22.- Concluded. NACA RM L58~05

90 NACA RM ~58~05

91 NACA RM L58~05

92 NACA RM ~58~05 Figure 23.- Continued.

93 NACA RM ~58~05 (h) a = 12'. (i) a = 14O. Figure 23.- Continued.

94 NACA RF/I ~58~05 Figure 23.- Concluded.

95 NACA RM L58~05

96 NACA I34 ~58~05 Figure Continued.

97 Figure 24.- Continued. NACA R4 L58~05

98 NACA RM ~58~05 0 (g) a = 0'. (h) a = 2. Figure 24.- Continued.

99 NACA RIvI L58~05 0 Upper surface 0 Lower surface 0 (i) a = 4. (j) a= 6'. Figure 24.- Continued.

100 NACA RM ~58~05

101 Figure 24.- Continued. NACA RM ~58~05

102 NACA RM ~58~05 Figure Concluded.

103 RM L58~05 d 8 0 e 4-' 9. %'%. Irl "op4 cd k I h +' a, n rl 0 w I' 0 k +' 0.., 2 0 4' 0 +' to 0 II rl P4 Z m 0 +' F:.rl 0. A51? ;% It 8 &i 2 8 n d 2 g: -I-='.A cd 4'rl &i +' 4' m -n 2 g 5. m+' 0 m d CO I E d F4 a, I1 40 e, F: d k d ha,.-, O r l e -2 P4 I m In (U e, 2.ri 61

104 NACA RM ~58~05

105 NACA RM L58~05

106 NACA RM ~58~05 (i) a = 24'. Figure 25.- Concluded.

107 NACA RM ~58~05

108 NACA RM ~58~05 (d) a = 4O. (e) a = 80. Figure Continued.

109 NACA RM ~58~05

110 NACA RM ~58~05

111 NACA FtM ~58~05

112 NACA RM ~58~05 CP 0 (d) a = 4. Figure Continued. (e) a = 8O.

113 NACA RM ~58~05

114 NACA RM ~58~05

115 NACA RM ~58~05 a t;l 0 e c, i? 8 i? 4 d 3 rl QI 0 2% t o t ' 117 % II rl 6 11 k +> - i.0 - P f, 0 c, M 0 rl 11 $4 z m (0 c,.5 0 rl $4 8 % 0 0 cd k t' g 3 d F: a, 0 t' 2 4 P P -ri cd k c, -$ *'. -rt 0 a S: 0' g d a3 I m t ' m cd " Ra d Plat M - :.5 cd.ri d w 3 a k 8 3 s;'.ri U 0 I % a5 CU

116 NACA RM ~58B05

117 NACA RM ~58~05

118 NACA RM L58BO5 LP (h) a = 20'. Figure 28.- Concluded.

119 NACA RM ~58~05

120 NACA RM ~58~05

121 NACA RM L58~05

122 NACA RM ~58~05 CP (i) a = 24O. Figure 29.- Concluded.

123 122 NACA RM ~58~05

124

125 Figure 30.- Continued. NACA RM L~~BOS

126

127 NACA RM L58~05

128 NACA RM ~581305

129 NACA RM ~58~05

130 NACA RM ~58~05 o Upper surface Lower surface (i) a= 24O. Figure 31.- Concluded.

131 NAC

132

133 NACA RM ~58~05 '0 v Q 0 0 (U II d i: d w %.,+ 5 Q : $ P., : 2 * U, S 0 0 I:",\.' Q - =2 '4 2 0 \I) i-4 II. d - hi) w rp'

134 NACA PJl: L58~05 (i) a = 24'. Figure 32.- Concluded.