1、NASA TECHNICAL NOTE NASA TJN D-5102q_ _ -LOAN COPY: RETURN TO AFWL (WLIL-2) KIRTLAND AFB, N MEX LARGE-SCALE TESTS OF AN AIRPLANE MODEL WITH A DOUBLE-DELTA WING, INCLUDING LONGITUDINAL AND LATERAL CHARACTERISTICS AND GROUND EFFECTS by Victor R. CorsigZia, David G. Koenig, and Joseph P. MoreZZi Ames R
2、esearch Center Moffett Field, CaZ.3 NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. MARCH 1969 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY M,NM 0333843 NASA TN D-5102 LARGE-SCALE TESTS OF AN AIRPLANE MODEL WITH A DOUB
3、LE-DELTA WING, INCLUDING LONGITUDINAL AND LATERAL CHARACTERISTICS AND GROUND EFFECTS By Victor R. Corsiglia, David G. Koenig, and Joseph P. Morelli Ames Research Center Moffett Field, Calif. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION - _ . For sale by the Clearinghouse for Federal Scientific and
4、Technical Information Springfield, Virginia 22151 - CFSTI price $3.00 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-LARGE-SCALE TESTS OF AN AIRPLANE MODEL WITH A DOUBLE-DELTA WING, INCLUDING LONGITUDINAL AND LATERAL CHARACTERISTICS AND GROUND EFFEC
5、TS By Victor R. Corsiglia, David G. Koenig, and Joseph P. Morelli Ames Research Center SUMMARY A wind-tunnel investigation has been undertaken to determine the lateral directional, and iongitudinal aerodynamic characteristics of a supersonic transport configuration with a double-delta wing of aspect
6、 ratio 1.66. Ground effects were also investigated. There was no large reduction in static longitudinal stability at high angles of attack. The trimmed lift coefficient in ground effect, at gear height, was 1.25 times that out of ground effect at the angle of attack for takeoff or landing. Calculati
7、ons of takeoff performance showed that the take off distance can be reduced by increasing the speed beyond that corresponding to the ground limit value of angle of attack. To improve the lift-drag ratio in the takeoff climb, it is desirable to accelerate after lift-off, because the speed for maximum
8、 L/D is about 250 knots relative to a takeoff speed of about 190 knots. The improved L/D results in reduced noise at distances greater than 3 miles from brake release and results in more efficient flight. INTRODUCTION One aircraft configuration that has been considered in recent supersonic transport
9、 design studies has a fixed double-delta wing and no horizontal tail or canard. Elevons are used for longitudinal and lateral control. The double-delta wing has a sharp leading edge that produces flow separation at angles of attack which would be used for landing and takeoff. This separated flow for
10、ms two vortex cores that extend above the wing. These vortex cores affect the aerodynamic forces and moments by interacting with the wing and vertical tail. In reference 1, it is pointed out that the use of strakes on a delta wing to form a double-delta wing increases the lift at a given angle of at
11、tack but reduces longitudinal stability. The double-delta wing, then, must be designed to obtain adequate lift at a given angle of attack and to avoid excessive reductions in stability with increasing angle of attack. A large-scale wind-tunnel investigation has been undertaken to determine the later
12、al, directional, and longitudinal aerodynamic characteristics of a supersonic transport configuration with a double-delta wing of aspect ratio 1.66. Longitudinal aerodynamic characteristics in the proximity of the ground Provided by IHSNot for ResaleNo reproduction or networking permitted without li
13、cense from IHS-,-,-were also obtained. The model was equipped with leading-edge flaps, a rudder, and elevons. The effects of Krueger flap and of increasing fuselage length were also investigated. NOTAT ION wing span, 29.0 ft wing local chord, in. reference chord, 2/S c2 dy, strakes off, 21.68 ft dra
14、g coefficient, drag/qS lift coefficient, lift/qS pitching-moment coefficient, pitching moment/qSE, positive nose up yawing-moment coefficient, yawing moment/qSb, positive nose right side-force coefficient, side force/qS, positive right rolling-moment coefficient, rolling moment/qSb, positive right w
15、ing down fuselage configuration, see figure 2(a) height of moment center above ground plane or above the runway, ft lift-drag ratio engine nacelles (no engines) load factor 2free-stream dynamic pressure, lb/ft 2 reference area, strake off, 505.9 ft thrust-weight ratio wing- sect ion thickness vertic
16、al tail or speed wing including strake or gross weight ordinate of wing section mean line measured from wing reference line Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a B however, at climbout values of CL (0.3 to 0.4) the L/D improvement increas
17、es to about 1.0. The angle-of-attack penalty due to deflection of the leading-edge flaps is about 1.0“. As an alternative to using leading-edge flaps, the use of Krueger flaps to improve L/D at high lift coefficients was considered. A sketch of these flaps is seen in figure 2(a). The data with the K
18、rueger flaps installed are presented in figure 6. A sununary showing the L/D and angle-of-attack varia tion with Krueger flap deflection appears in figure 7. At CL = 0.5 the maxi mum L/D increase is about the same as obtained with the use of leading-edge flaps; at lower CL (0.3), however, the L/D in
19、crease is less. The angle-of attack penalty remained about the same as for leading-edge flaps at all CL. As shown in figure 2(a), the leading-edge flaps had three panels on each side, each with approximately the same spanwise extent. Data with leading-edge flap deflection varying spanwise are presen
20、ted in figure 8. The spanwise variation of leading-edge flap deflection produced modest effects on pitching-moment characteristics, and only minor effects on lift and drag characteris tics. These results suggest that moderate improvements in stability can be achieved by spanwise variation of leading
21、-edge flap deflection. In the data shown above, the leading-edge flaps were deflected down to increase L/D at a given CL. However, this was accompanied by an increase in angle of attack required for a given CL. Therefore, the use of upward deflection of the leading-edge flaps to reduce the angle of
22、attack required for a given CL was considered. The results of this investigation are pre sented in figure 9. It is shown that an upward 20“ deflection of the leading-edge flaps reduced the angle of attack by only 0.3“ at CL = 0.5, compared to the configuration with 0“ flap deflection. In addition, t
23、here was an increase in CD for a given CL. -. -Effects of ground proximity.- The characteristics of the model in the presence of the ground- are presented in figures 10 through 12. Figure 13 is a summary of the ground effects. The effect on elevon effectiveness was slight. Lift-curve slope and stati
24、c margin changed 0.030 and 0.08, respec tively, for a change in ground height from out of ground effect to h/c = 0.19. This corresponds to about a SO-percent increase in lift-curve slope. In the same range of ground height, the drag was reduced 20 percent 6 L Provided by IHSNot for ResaleNo reproduc
25、tion or networking permitted without license from IHS-,-,- of the out of ground effect value. With the elevons fixed (6, = -5“), the lift coefficient in ground effect was about 1.6 times the value out of ground effect. However, with the pitching moment trimmed by use of the elevons, this multiple wa
26、s reduced to 1.25. Lateral-Directional Characteristics Figure 14 presents data with various angles of sideslip for the short fuselage configuration. Data with the vertical fin and the nacelles removed are presented in figures 15 and 16, respectively. Lateral stability.- A summary of the data showing
27、 the variation of C with sideslip and lift coefficient is presented in figure 17. The value of 2B varies between -0.0018 and -0.0030 for CL from 0.4 to 1.0. CLB was ZB obtained by crossplotting C2 versus B and then measuring the slope. Directional stability.- The data for the configuration with the
28、extended fuselage are presented in figure 18. Figure 19 is a summary of the yawing moment due to sideslip characteristics of the model as shown in figures 14(b) and 18(b). The variation of yawing moment with angle of sideslip is stable at lift coefficients up to 0.8, but the level of stability is re
29、duced consid erably at a lift coefficient of 1.0. This is due to the effect of the inter action of the windward wing leading-edge vortex and the forebody vortices with the vertical tail. The only significant effect of extending the fuselage was to reduce CnB for low values of sideslip and to cause t
30、he destabilizing tendency at high lift to become more severe. The value of Cn measured at low values of lift coefficient and sideslip is 0.0020 for the gxtended fuselage configuration. Lateral control.- Data with various aileron deflections at zero elevon are presented in figure 20. of CL from 0 to
31、1.0. The value of C was 0.00088 at zero lift. The The ailerons are seen to be effective for values 6, yawing-moment coefficient due to aileron deflection is adverse at high lift coefficients but not at low lift coefficients (CL below 0.23). The values of Cn6, are -0.0001 and -0.0002 for CL = 0.4 and
32、 0.8, respectively. Figure 21 presents data for various angles of sideslip with the ailerons deflected. An aileron deflection of 6, = 40“ was adequate to trim up to 12“ CL = 0.8.of sideslip at Directional control.- Data with various rudder deflections are presented in figure 22: The rudder effective
33、ness is linear and can be seen to be inde pendent of CL for values of CL from 0 to 1.0. The value of Cn was 6r measured to be -0.0014. Figure 23 presents data for various angles of side slip with the rudder deflected. The maximum rudder deflection tested CL(6, = +27“) was not adequate to trim 12“ of
34、 sideslip (B = 12“) for from 0.2 to 0.8. sideslip. At CL = 0.6 this rudder deflection would trim about 10“ of 7 C Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-COMPARISON WITH THEORY Out of Ground Effect A theoretical method to predict the normal f
35、orces and pitching moments on low-aspect-ratio wings with leading-edge vortex flow is presented in reference 2. This method has been applied to the wing fuselage configuration used in the present investigation, and the results are shown in figure 24. As shown, the lift curve is slightly under that p
36、redicted with a maximum error in CL of 0.07. The aerodynamic center location at low lift is predicted well. However, the theory predicted an increase in longitudinal stability with increasing CL instead of the slight reduction in stability obtained experimentally. At CL = 0.5, the error in longitudi
37、nal stability is 6 -percent c. In Ground Effect For the prediction of ground effect, the theory of Gersten (refs. 3 and 4) was used. A summary of the ground effect on lift results, as measured in the present investigation, is compared with Gerstens theory in figure 25. It is seen that for the range
38、of ground heights tested the agreement between theory and experiment is good. TAKEOFF PERFORMANCE A double-delta wing transport has aerodynamic characteristics consider ably different from those of conventional subsonic jet transports. Three characteristics are especially noteworthy. First, the lift
39、 coefficient available for takeoff and landing is limited by the low lift-curve slope and the ground clearance limit on angle of attack (about a = 12). Second, the lift-drag ratio is low (about 5) at values of lift coefficient used for take-off and landing. Finally, the thrust-weight ratio is quite
40、high (about 0.38). Figure 26 presents trimmed values of CL as a function of L/D, and a. for this double-delta configuration both in and out of ground effect. A representative value of angle of attack for lift-off is about a. = 12“, so that the maximum available lift coefficient for lift-off with lea
41、ding-edge flaps deflected 30 is CL = 0.64. The value of L/D out of ground effect at this lift coefficient is approximately half the maximum value. Takeoff Velocity and Distance The takeoff characteristics have been predicted for an airplane having the aerodynamic characteristics presented here and t
42、he thrust and weight characteristics representative of a supersonic transport airplane. Some of 8 . . .- ._ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-the parameters used in this computation are listed in table 111, and a brief description of th
43、e computational technique is given in appendix A. Figure 27 presents the takeoff distances to flight speeds corresponding to lift-off and 50-foot altitude for various values of gross weight. For comparison, the flight-test results of a KC-135A jet tanker are also presented (ref. 5). Per tinent param
44、eters of the KC-135A are also listed in table 111. Lift-off velocities are about 20 knots higher for the SST (due to the low value of CL available). However, because of its superior acceleration ability (due to the higher thrust-weight ratio), the takeoff distance to an altitude of 50 feet for the S
45、ST (for W = 590,000 lb) is about 1500 feet less than that of the KC-135A (for W = 240,000 lb). Optimum Lift-off Speed Minimum lift-off speed is not the speed for optimum takeoff for this aircraft. For the takeoff computation shown in the previous figure, lift-off occurred near the limit value of ang
46、le of attack. This was done to limit the speed at lift-off to a near minimum value. An improved L/D at takeoff can be achieved by increasing the lift-off speed. Increased lift-off speed will reduce the CL below that available at a = 12 and thereby increase the L/D (see fig. 26(b). As shown in figure
47、 28, increasing the lift-off speed from 170 to 189 knots improved the second segment rate of climb by 9 percent without increasing the distance required to attain an altitude of 35 feet. Takeoff Climb Profile To operate at an improved L/D after lift-off, it is necessary to accelerate to higher speed
48、 (i.e., lower CL) rather than to climb at the max imum angle available at takeoff speed. This will result in reduced airplane altitude in the vicinity of the airport. However, to reduce the noise heard on the ground, it is desirable to attain high altitude and high L/D (low thrust). These climb tech
49、nique considerations are shown in figure 29, where two climb profiles are presented. One represents a high climb gradient tech nique, the other an acceleration type. For the high climb gradient takeoff (i.e. , aircraft A, fig. 29) , the angle of attack was increased from 12 at lift-off to 14, resulting in a load factor of 1.2 (standard day). This load factor was maintained until
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