.Am-^io^oi RM E50J09 RESEARCH MEMORANDUM EXPERIMENTAL PRESSURE DISTRIBUTIONS OVER TWO WING-BODY COMBINATIONS AT MACH NUMBER 1.9 By Barry Moskowitz and Stephen H. Maslen Lewis Flight Propulsion Laboratory Cleveland, Ohio UN^yeRSlTY OF FtOBtDA txX^UMENTS DEPAPrrMENT 120MARST0N SCIENCE UBRARY ^0^^^3261.7011 USA NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON February 5, 1951 ^-^.f-^JVl^U? ry^ l^^^f NACA RM E50J09 NATIONAL ADVISORY COMMITrEE FOE AERONAUTICS RESEARCH MEMORANDUM EXPERIMENTAL PRESSURE DISTRIBUTIONS OVER TWO WING- BODY COMBINATIONS AT MACS NUMBER 1.9 By Barry Moskowitz and Stephen H. Maslen SUMMARY Pressure distri'butions on two ving-TDody combinations have teen ohtained at a Mach numher of 1.9 to investigate the wing -"body inter- ference. A rectangular wing, a triangular wing, and a cylindrical body with an ogive nose were studied alone and in comhination. Both wings had a span of 14 inches, an aspect ratio of 2=-, and a 5 -percent- thick douhle-wedge cross section. The wings were mounted on the cylin- drical portion of the hody. The investigation was conducted over a range of angles of attack varying in 2° increments from -4° to 4°. The pressiure distributions over the wing-hody combination com- pared favorably with theoretical calculations based primarily on a generalization of the method of Nielsen and Matteson, except at the root section of the wings where the boundary layer of the body modi- fied the flow. INTRODUCTION Although the characteristics of the flow about thin wings and about slender bodies of revolution in a supersonic stream have been extensively studied in the past few years, information concerning the effects of interference between wing and body is still needed. Theoretical studies of the problem using linearized theory are pre- sented in references 1 to 5. In reference 6, the method of charac- teristics was used. Little experimental data are available, however, particularly on pressure distributions. Van Dyke (reference 7) has reported force measurements and Cramer (reference 8) has given some results that are compared with Ferrari's studies (reference 2). NACA RM E50J09 In the investig3.tion reported herein, which was conducted at the NACA Lewis laboratory, static -pressure surveys were made at a Mach numher of 1.9 on two wing-hody comhinations; one wing was rectangular and the other triangular. The wings and the body were investigated separately for comparison. The results are compared with theoretical calculations based primarily on the method of reference 4, which was generalized to include sotae angle -of -attack effects. APPARATUS AND PROCEDUEE The wing-body combinations were investigated in the NACA Lewis 18- by 18 -inch supersonic wind tunnel. The Mach number, obtained from a tunnel calibration survey, was 1.90 ±.01 in the vicinity of the models. The total -pressure variation along the models was ±0.5 percent and was therefore neglected. The test Reynolds number, based on the mean wing chord, was 1.64 x 10 . A photograph of the models Investigated is shown in figure 1 and a sketch of the models showing principle dimensions and the loca- tion of static -pressure orifices is presented in figure 2. Pressure surveys were taken at 2° increments in angle of attack from -4° to 4°. The pressures were photographically recorded from multiple-tube mano- meter boards using tetrabromoethane as the working fluid. The angle of attack of the model was measured with a catheto- meter during each test. Angles were accurate to t.oi°. Because of the low aspect ratio of the wings, aeroelastic deformation of the wing sections was considered negligible. THEORY The differential equation for linearized potential flow is assumed to apply. Solutions may be linearly superposed to obtain the flow over a particular configuration. The boundary condition for the flow is that the normal velocity is zero at the surface. The wing-body problem may be solved by starting with the solu- tion for the wing alone (reference 9 or 10) and the body alone (refer- ence 11 or 12). When these solutions are combined, the boundary con- ditions are no longer satisfied because the body (at angle of attack) generates an upwash in the plane of the wing and the wing causes a flow through the body. A wing -interference potential and a body- interference potential are introduced, whose respective functions are to cancel the upwash in the plane of the wing and to cancel the flow through the body. NACA RM E50J09 Consider the ■'ri.nc in the presence of the known ■body upwash field (reference 12 or 13). The required wing interference potential is readily found hy the method given in reference 14, which amounts to considering a wing with a local angle of attack equal to the local upwash induced by the body divided by the free-stream velocity. In order to avoid infinite sidewash at the wing-body Juncture, the upwash is taken to be continuous through the blanketed portion of the wing (region of the wing covered by the body). The wing-interference potential plus the two basic solutions satisfy the bounds.ry conditions on the whole wing and on the body in the region of no interaction. The renaining problem consists in can- celing the flow generated tlirough the body by the two wing potentials, without disturbing the boundary conditions on the wing. The latter condition id-ll automatically be satisfied if the problem can be treated as symmetric with respect to the plane of the wing. In reference 4, Nielsen and Ifetteson describe an approximate method to determine the body interference solution for symmetric - flow problems. The body is sjinmetrically divided into plane control areas with finite line pressure sources placed on each control surface normal to the free stream. The normal velocity is averaged over each area and the strengths of successive sources are determined taking into account the effect of one line so^orce on another. Because the procedure of reference 4 gives a symmetric solution, no flow is induced across the plane of the wing. For a symmetrical wing mounted on the body at the center line with both wing and body at zero angle of attack, the flow is symmetric. Also, those portions of an unsymmetric flow may be treated as symmetric that are not influenced by any portion of the body (in the region of intera.ctlon) lying on the opposite side of the plane of the wing. For example, in the configurations shown in the following sketch, no body-interference potential is required in region I; in region II, the body interference potential may be found by considering a sjTnmetri- cal problem; and in region III the problem must be considered as unsymmetric. NACA KM E50J09 V' II in (c) With the procedures previously discussed, the pressure distri- ■bution for a rectangular wing-body comhination can he calculated in regions uninfluenced by the wing trailing edge (similar to the con- figuration in sketch (a)). The triangular wing investigated actually had a slightly subsonic leading edge but was assumed, for ease in com- putations, to have a sonic edge and is conseq_uently similar to sketch (b). The triangular -wing solution yields infinite sidewash at the intersection of the leading edge with the body. This infinite sidewash was reduced to finite values, before calculating the body interference potential, by distribution line sources and sinks in the manner discussed in reference 4. Superposition of the rectangular -wing solution, the wing- interference solution, and the body solution results in discontin- uities in the pressure distribution on the body. The body -interference solution should therefore have corresponding discontinuities of the same magnitude but of opposite sign, inasmuch as the pressixre distri- bution on the body should be continuous. This continuity of the pres- sures on the body is known from the fact that all disturbances on the body due to the wing are propagated along Mach cones, which generate no pressure discontinuities in linearized theory. The normal velo- cities, however, are averaged over given regions of the body; con- sequently, the body-interference solution fails to give these dis- continuities. For this reason, the body-interference pressures were so adjusted on the body that the pressure distributions were contin- uous on the body. NACA EM E50J09 Another effect of averaging the normal velocities is that the pressures due to the presence of the wing acting in the region on the 'body between the Mach wedge from the leading edge of the rectan- gular wing and the Intersection of the hody with the Mach cones from the foremost part of the wing-hody juncture are not zero. An adjust- ment was therefore made so that the "body -interference pressures exactly canceled the pressure on the hody due to the wing solution in this region. Nielsen of the KACA Ames lahoratory pointed out that this region should actually extend slightly farther downstream because the disturbances generated by the junction of the body and the wing leading edge will be propagated along the body surface at the Mach angle rather than along Mach cones. Thus, for example, the disturb- ance will first reach the top of the body a distance ^'Rii/Z, where g is the cotangent of the Mach angle downstream of the wing leading edge and R is the body radius, rather than PR/n/2 as assumed in the computation. The last value was used because it is the result indicated by ordinary linear theory. RESULTS Experimental data are presented and compared with theoretical results in figures 3 to 13. The pressure coefficient Cp and the change in pressure with angle of attack dCp/da, both evaluated at a = 0, are plotted for each orifice location. Within the limited range of angles of attack of the investigation, the faired variation of Cp with angle of attack was approximately linear. The experi- mental results for the body alone, rectangular wing alone, and tri- angular wing alone are given in figures 3, 4, ajid 5, respectively. In dCT,/da figure 3 the experimental value of — ^ , where 6 is the angle sin measured around the body from the plane of the wing, is obtained by dC^/da averaging the various values of — ^ The agreement between theory sin 6 and experiment in figures 3, 4, and 5 is the basis for determining the accuracy of the interference calculation. For example, if the difference between experiment and theory for a component in combi- nation is within the order of agreement existing between theory and experiment for the component alone, the agreement is considered good. Rectangular Wing and Body Results for the rectangular wing in the presence of the body eire presented in figure 6. Because of the large nose angle of the body NACA RM E50J0S (30^^ half angle), a fairly strong shock was generated (fig- 14). In the neighborhood of the wing tip, the shock wave occurred approxi- mately 2 inches upstream of the Mach wave assiimed hy linear theory. As a result the pressures predicted hy linear theory, in the forward part of the tip region, are higher than the experimental values. As shown in the schlieren photograph, the nose shock reflects off the tunnel walls and intersects the wing tip. Because of the increase in pressure across the reflected shock, the experimental values are higher than those ohtained "by linear theory in the rearward portion of the tip region. In the plane of wing, however, the "body does not directly cause any change in pressure due to angle of attack. More- over, the shock position does not vary appreciably over the range of angles of attack of the investigation. The effects of the shock and its reflection therefore cause no discrepancy "between the theoretical and experimental values of dCp/da on the wing. At the root section, the "boundary layer of the "body modified the flow over the wing. In particular, the discontinuity in the slope of the wing was softened, there"by decreasing the pressure change pre- dicted "by linear theory. At the wing mldspan position, none of these difficulties occurred and close agreement "between linear theory and experiment was o"btained. In order to illustrate the effect of the presence of the "body on the pressures acting on the rectangular wing, the increments ACp and A — E. due to the presence of the "body are presented in figure 7 . The da theoretical curves and the experimental points were o'btained by taking the difference between corresponding values in figures 6 and 4. For zero angle of attack, the section wave drag of the wing in combination is less than the corresponding section wave drag of the wing alone, if the increment in pressure is ne^tive over the positively sloped portion of the wing and positive over the negatively sloped portion. In the root section the experimental points, but not the theoretical curve, indicate less section drag. In the midspan region the experi- mental points indicate slightly less section drag, whereas the section drag indicated by the theoretical curve remains about the same. In the tip region, both theoretical and experimental values indicate an increase in the section drag, which is to be expected inasmuch as the tip region of the rectangular wing is influenced by the pressure gra- dient associated with the nose of the body. Because of a large reduc- tion in the drag of the wing due to the blanketing of the center sec- tion, the over-all wave drag of the combination is probably less than the sum of the wave drags of the components . NACA RM E50J09 Iimsmuch as the Increment in dCp/da is negative on the top sur- face of the wing, the sectional lift is greater than the corresponding sectional lift of the wing alone. This resxilt is to he expected "because the presence of the hody at angle of attack causes an upwash field, which increases the effective wing angle of attack. (See the section THEORY.) The experimental variation of Cp and dCp/da on the hody in the presence of the rectangular wing is presented in figures 8(a) ajid 8(h), respectively. Close agreement was ohtained for the slope of the pressure coefficient curve. For zero angle of attack, the trend of the experimental points and the theoretical curves appear to he similar. The q^uantitative agreement is of the order ohtained for the wings and the hody alone. The increments in Cp and dCp/da on the hody due to the presence of the wing are presented in figures 3 (a) and 9(h), respectively. In figure 9(a) the direct influence of the wing is noted in the increases and decreases in hody pressure coefficient in the regions influenced hy the positive and negative wing slopes, dCr) respectively. Because the A— r^ curve in figure 9(h) is negative, an increase in lift results on the hody due to the presence of the wing. Triangular Wing ajid Body The experimental results for the triangular wing in the presence of the hody are shown in figure 10. Reasonably good agreement with linear theory was ohtained at all stations except for the root sec- tion at a = 0, where, as was noted in the case of the rectangular wing, the theoretical discontinuity in pressure is modified hy the boundary layer of the hody. Inasmuch as the hody nose shock was not near the triangular wing at any point, the poor correlation between theory and experiment noted in the outboard region of the rectangular wing does not occur for this case. dC The increments AC and ^-^ on the triangular wing due to the presence of the body are presented in figure 11. At the outboard section A, the experimental points and the theoretical curve for a = indicate a slight decrease in section drag relative to the drag of the wing alone. At the midspan section B, the experimental points indicate a slight decrease in section drag, whereas the theoretical MCA EM E50J09 curve Indicates that the section drag remains about the same for a = Both the experimental points and the theory at the inhoard section C show a sizeable decrease in section drag over that for the triangular wing alone at a = 0. Thus at zero angle of attack, the over -all wave drag of the triangular wing-hody combination is less than the sum of the wave drags of the com.ponents- At sections A and B, the fact that the experimental points of AdCp/da are ne^tive indicates a slight increase in section lift, although the theoretical curves at these stations indicate that the section lift remains unchanged. At the root section C, both linear theory and experiment show a decrease in section lift. Experimental results for Cp and dCp/da on the body in the presence of the triangular wing are presented in figures 12(a) and 12(b), respectively. The difference between experiment and linear theory in figure 12(a) is very nearly that displayed by the curves for the body alone (fig. 3), v^hich indicated that the interference was accurately predicted by the theory. This agreement is also shown in figures 13(a) and 13(b), where the differences in Cp and dCp/da for the body in combination with the triangular wing and the body alone are presented. CONCLUDING REMARKS Experimental pressure distributions at a Mach number of 1.9 were obtained for a rectangular wing ajid a triangiilar wing in combination with an ogive -nose body. The procedure of Nielsen and Matteson was applied at zero angle of attack and a generalization of this method was used to calculate interference effects at angle of attack. The experimental results compared favorably with these theoretical calculations, except at the root section of the wings, where the boimdary layer of the body modi- fied the flow over the slope discontinuity of the wing and thereby decreased the pressure change predicted by linear theory. Lewis Flight Propulsion laboratory. National Advisory Committee for Aeronautics, Cleveland, Ohio. NACA EM E50J09 REFERENCES 1. Lagerstrom, P. A., and Van Dyke, M. D. : General Considerations about Planar and Non -Planar Lifting Systems. Rep. No. SM-13452, Douglas Aircraft Co., Inc., June 1949. 2. Ferrari, Carlo: Interference between Wing and Body at Supersonic Speeds - Theory and Numerical Application. Jour. Aero. Sci., vol. 15, no. 6, June 1948, pp. 517-536. 3. Spreiter, John R.: The Aerodynamic Forces on Slender Plane- and Cruciform-Wing and Body Combinations. MCA Rep. 962, 1950. (Formerly NACA TWs 1897 and l662.) 4. Nielsen, Jack N., and >fe.tteson, Frederick H. : Calculative Method for Estimating the Interference Pressure Field at Zero Lift on a Symmetrical Swept-Back Wing Mounted on a Circular Cylindrical Body. NACA RM A9E19, 1949. 5. Browne, S. H., Friedman, L., and Hodes, I. : A Wing-Body Problem in a Supersonic Conical Flow. Jour. Aero. Sci., vol. 15, no. 8, Aug. 1948, pp. 443-452. 6. Ferrari, Carlo: Interference between Wing and Body at Supersonic Speeds - Analysis by the Method of Characteristics. Jour. Aero. Sci., vol. 16, no. 7, Jvily 1949, pp. 411-437. 7. Van Dyke, Milton D. : Aerodynamic Charcteristics Including Scale Effect of Several Wings and Bodies Alone and in Combination at a Mach Number of 1.53. NACA RM A6E22, 1946. 8. Cramer, R. H- : Some Theoretical and Experimental Results for Wing -Body Interference at Supersonic Velocities. Bull. Bumblebee Aero. Symposum (Johns Hopkins Univ.), Nov. 4-5, 1948, pp. 267-276. 9. Eward, John C.: Use of Source Distributions for Evaluating Theoretical Aerodynamics of Thin Finite Wings at Supersonic Speeds. NACA Rep. 951, 1950. 10. lagerstrom, P. A.: Linearized Supersonic Theory of Conical Wings - (Corrected Copy). NACA TW 1685, 1950. 11. von Earman, Theodor, and Moore, Norton B. : Resistance of Slender Bodies Moving with Supersonic Velocities, with Special Reference to Projectiles. Trans. A.S.M.E., vol. 54, no. 23, Dec 15, 1932, pp. 303-310. 10 MCA EM E50J09 12. Tsien, Hsue-Shen: Supersonic Flow over an Inclined Body of Eevolution. Jour. Aero. Sci-, vol. 5, no- 12, Oct. 1938, pp. 480-483. 13. Beskin, Leon: Determination of Upwash aroimd a Body of Revo- lution at Supersonic Velocities. DEVF Memo. BB-6, Consoli- dated Vu_Ltee Aircraft Corp., May 27, 1946. 14. Mlrels, Harold: Theoretical Mfethod for Solution of Aerodynamic Forces on Thin Wings in Nonuniform Supersonic Stream with an Application to Tail Surfaces. NACA TW 1736, 1948. NACA EM E50J09 11 o g ■H o •P •H > ri - §>■ aJ •H ■P CO -p o (D M (D to •H NACA EM E50J09 13 200 Typical croes section J—^ 14" 3" -^^r jL 1" T~"2 4 jL 0.3" Rectangular wing 9 Position of wings relative to body O Orifice on top surface D Orifice on 'bottom, surface Triangular wing Figure 2. - Dimensioned sketch of models showing pressure orifices. 14 KACA EM E50J09 Pi O CD D •O U P< JO p a .4 r*i \ .3 \ c \ .2 \ \ - Theory Experiment t o .1 \ \ \ o\ .^ O 0^ ) o o c ) -.1 o o ^ ^ > ( } Dists LO moe fi •om not L4 ie , in 1 8 22 -.020 -.016 -.012 -.006 -.004 Ck ^ N, o^ \ °\ \, \ o V >> O { •> O r\ f ^ .004 1 Figure 3. - Values and slopes of pressure coefficient for body alone at zero angle of attack. NACA EM E50J09 15 .08 .04 o J a. o -.04 -.08 ,12 -.04 to Tip / O ' ) » o Two -d Imens 1 ona 1 region / O (, , " o C ) 3 3 _j_i" >^-- -- ^ Flow X \ \ *s \ o o < ) O < > o ( ) V o < Tip ) n Two-dimensional region < < ) O "^ ( > r^ )-£> 1 "^.t^ 4 6 2 Distance from leading edge, in. Figure 4. - Values and slopes of pressure coefficient for rectangular wing alone at zero angle of attack. 16 MCA RM E50J09 o Pi 0) u i) o d U8 \, Theory Experiment o o o ^ 04 o o o ! 04 o o 08 o / ^ V \ b 1 °A / / \ ^ 1 12 04 Sect' Lon C _/ y 1 1 1 / V < *- 02 > d^> O o o o - ^ -''-^ O o o 1 4 2 Distance from leeidlng edge. In. (to) Oufboard stations. Figure 5. - Concluded. Values and slopes of pressure coefficient for triangular wing alone at zero angle of attack. 18 NACA EM E50J09 -p a >i o o o a (D to ® P. ai rH C! (d -p ■H O o o tn a! -P ■P +J a a B •H Hh o o •rH <*-! " ^^ E-i O Ci o -©- c o m 0) -P a % u o fe 20 NACA EM E50J09 12 14 16 Distance from body nose, In. 18 20 (a) Pressure coefficient Figure 8. - Values and slopes of pressure coefficient on body in presence of rectangular wing at zero angle of attack. NACA EM E50J09 21 I .03 -.02 .01 .03 .02 .01 c •a u o Pi o 3 -.02 .04 -.01 -1' T3 h -.03 .02 -.01 -.04 .03 .02 .01 1 . . 1 1 JvjACA -^ Theory O Experiment n >' V e (deg] < > -> o ^k .^_o^ 90 o y / < x,,^ ■"'^"^ 'X < ) 65 ^ ' - / ^"^ O o \ __0__ 40 rv{ / < ) r ■^ ^ \ ) o\ - 20 \, o ^ ^. >—- - 'v o 10 0-' 1 » 10 650 — 40° — 20° _ 10° — 12 14 16 Distance from body nose, In. ^ 18 ^^ 20 (b) Slope of pressure coefficient. Figure 8. - Concluded. Values and slopes of pressure coefficient on body in presence of rectangular wing at zero angle of attack. 22 NACA EM E50J09 .08 -.04 .04 -.06 .04 .08 - . 08 .04 O < .12 .08 - . 04 .04 .04 .08 .08 10 12 14 16 Distance from tody nose, in. (a) Incremental pressure coefficient. Figure 9. - Incremental values and slopes of pressure coefficient on body due to presence of rectangular wing at zero angle of attack. MCA EM E50J09 23 -.01 -.02 .01 to a u o & -. 04 -.01 .03 -.02 .01 -.04 10 12 14 16 Distance from tody nose, in. ("b) Incremental slope of pressure coefficient. Figure 9. - Concluded. Incremental values and slopes of pressure coefficient on body due to presence of rectangular wing at zero angle of attack. 24 NACA EM E50J09 o (D +J as S>> ■d O ,Q V( O (O O G 0) CO o i" 3 60 •H -C! u 4j +J 60 ^H ^ o o tM a! ■^ ■p CIS 4J -P r-{ tH tH H O O o •H o 3 (D CM O m © o ^ o (D 0=D ro) O'T) / 8ep jsd /-2£ l'3P. NACA EM E50J09 25 \ § = - y -((CD X 1 z/ { u ^ t mY"~\- \- o / PQ / o / o * "\ \ o v.^ o ^ I o o \ o o o \o o > o ■p i o o )°0 m O 1 o O c c ■p a < \ ^ >.5 °\ c \ 1 \ o a> m 1 4) Q °/ a o B W /° o •H -p o 4) o / o o o / r CO c 6 °/ 3 °/ < /^ / O j s o / o/ •H -P o 01 o°l »° 1 o\ \ ra o o CO o o I 0^,d, (^OV) Sep j:9d o 0=T0 !« o (D o a (D m ^ (3 rH :3 M § •H ^ -P fl u o tH ■\ 111 -P ^i! rio a o Ti O -P a) ■H cd a =H •H 'm 'm TIJ 5 H (D t3 ft o o ^ M m ■d C (d to Q) :i r-t m > rH oj -P a o a (D ^^ o •pp , 5 CD •H 26 NACA EM E50J09 .04 .04 ,04 -.08 -.04 -.08 .04 o Pi .04 .04 -.08 -.04 • . 08 .04 -.04 -.08 10 Theory Experiment e (deg) ^ N ______ __90__ — ' o ( ) o o X o o < ) 65 o ^ o <\ > -^ V ^ "-WL __ o o 12 14 16 Distance from body nose, In. 18 20 (a) Pressure coefficient. Figure 12. - Values and slopes of pressure coefficient on body in presence of trianmilar wing at zero angle of attack. NACA EM E50J09 27 o & •a l-o . -.02 -.01 -.02 .01 .02 .01 .02 .01 .02 .01 10 Theory- Experiment o e (deg) < ) n n ^.^O 90 o o^^ ) ^ -i J o o 65 O ' > O ^^ o O o 40/^ ) 20 '/ ^-^.^^O ^ o^ < ) o ^ J ^■. ^ 10 \ o^- 1 o — '6' 12 14 16 Distance from "body nose, in. 18 20 (b) Slope of pressure coefficient. Figure 12. - Concluded. Values and slopes of pressure coefficient on body in presence of triangular wing at zero angle of attack. 28 WACA RM E50J09 .04 -.04 .08 o o .08 .04 .04 -.08 Theory Experiment o .04 e (deg) 90 x^ ^^ ' o~^ .04 o o O \C ) 65 ^ c > o \ o c > .04 >- -^ r"\^ ( .04 V .,_^ < ) 20 / o" o \ \: ) o n < > .04 -■ 10 ' o" rt N, -"-■-^ .04 < ?^^ "^O < > 10 12 14 16 Distance from body nose, in. 18 20 (a) Incremental pressure coefficient. Figure 13. - Incremental values and slopes of pressure coefficient on body due to presence of triangular wing at zero angle of attack. NACA EIvl E50J09 29 to o d -o I'd .02 -.01 -.02 -.01 -.02 .01 .02 .01 .02 .01 - Theory Experiment o c » — o — o e (deg) 90 O o^-- "^ i < ) . r\ ~o 65 o o I ) o n - ( o o *-» 40/^ ( ) 20 J L^^ ^^ o^ ( ) 1 fl o o J ^^ 10 \ o,-- o o" 1 > -*< 10 12 14 16 Distance from "body nose, in. 18 20 (t) Incremental slope of pressure coefficient. Figure 13. - Concluded. Incremental values and slopes of pressure coefficient on body due to presence of triangular wing at zero angle of attack. NACA EM E50J09 31 e o o 3 O ft O o o CO •H NACA - LanRU'y Hula. vj. to -* I u 0) m OJ ft 0) -p CQ Ti oj N •p •H CQ O t: u a pq O H Ha u^ o o\ in rH is; to CO * H H -H S o s to CO o H (5q •H ? 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