1、NASA TECHNICAL NOTE NASA TN 0-7149 oc w= z c WIND-TUNNEL INVESTIGATION OF STATIC LONGITUDINAL AND LATERAL CHARACTERISTICS OF A FULL-SCALE MOCKUP OF A LIGHT SINGLE-ENGINE HIGH-WING AIRPLANE bY H. Dozlglds Greer, James P, Shivers, Muruin P. Fink Ldngley Reseurch Center and C. Robert Curter Lungley Dir
2、ectorute, US. Army Air Mobility R: of zero which represents either a low-power or a high-speed condition (where the thrust coefficient approaches zero), TI: = 0.14 which corresponds to a climb condition, and TI: = 0.30 which cor- responds to a take-off condition. The investigation showed that the mo
3、del has both stick-fixed and stick-free longi- tudinal stability up to the stall for all configurations tested with the center of gravity located at 13.7 percent of the mean geometric chord. Power generally has a small destabilizing effect, but at worst the model could become no more than neutrally
4、stable with the center of gravity as far aft as 40 percent of the mean geometric chord. The model has positive effective dihedral and directional stability for all test conditions. The aileron and rudder effectiveness was maintained up to the stall and was powerful enough to trim out all model momen
5、ts up to the stall. INTRODUCTION The Langley Research Center has been conducting a program to document the aero- dynamic characteristics of a number of general-aviation type aircraft. Full-scale mock- ups of two single-engine low-wing airplanes, two twin-engine low-wing airplanes, and one low-wing g
6、eneral research model which had several nacelle configurations and modes of Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-propeller rotation have been tested in the Langley full-scale tunnel and the results have been reported in references 1 to 5.
7、The present investigation was conducted to deter- mine the static longitudinal and lateral stability and control characteristics of a mockup of a light single-engine high-wing airplane. The present tests were made with various power and flap settings over a range of angles of attack from -4O to 24 a
8、nd over a range of sideslip angles of *8O. The tests at thrust coefficients of 0 and 0.14 were made at a tunnel speed of about 27.4 m/sec (90 ft/sec) giving a Reynolds number of about 2.8 X lo6. The tests at a thrust coefficient of 0.30 were made at 22.8 m/sec (75 ft/sec) giving a Reynolds number of
9、 about 2.3 X 106, The Reynolds number is based on a mean geomet- ric chord of 1.50 m (4.91 ft). SYMBOLS Figure 1 shows the stability-axis system used in the presentation of the data and the positive direction of forces, moments, and angles. The data are computed about the moment center shown in figu
10、re 2 which is at 13.7 percent of the mean geometric chord. Values are given in both SI and U.S. Customary Units. Measurements and calcula- tions were made in U.S. Customary Units. b wing span, 11.20 m (36.75 ft) cD h,a h,e Drag drag coefficient, - qs Hinge moment aileron hinge-moment Coefficient, SS
11、aEa Hinge moment e levator hinge - moment coefficient , qSeEe Hinge moment rudder hinge - mom ent coefficient, h,r qSrEr Lift lift coefficient, - cL qs lift - curve slope La C cZ rolling-moment coefficient, Rolling moment 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without
12、 license from IHS-,-,-C C ba Cm C be Cn cnP CY - C qtp S effective-dihedral parameter, aCz/8p, per deg aileron effectiveness parameter, 8C a6, per deg J Pitching moment q= pitching-moment coefficient, elevator effectiveness parameter, aCm/abe, per deg Yawing moment yawing-moment coefficient, directi
13、onal stability parameter, 8Cn/ap, per deg I rudder effectiveness parameter, aCn/ab, per deg Side force qs side-force coefficient, mean geometric chord, 1.50 m (4.91 ft) aileron mean chord, 0.37 m (1.20 ft) elevator mean chord, 0.46 m (1.50 ft) rudder mean chord, 0.40 m (1.30 ft) f r e e - stream dyn
14、amic pres sur e, N/m2 (lbf /f t2) ratio of dynamic pressure at tail to free-stream dynamic pressure wing area, 16.25 m2 (175 ft2) area of one aileron, 0.65 m2 (6.95 ft2) area of elevator, 1.63 m2 (17.50 ft2) area of rudder, 0.59 m2 (6.40 ft2) 3 Provided by IHSNot for ResaleNo reproduction or network
15、ing permitted without license from IHS-,-,-T T;: x v a! P 6a 6e 6f 6r E 6 effective thrust (at a! = Oo and gear zp), and for flap deflections of loo and 30 with gear down. A range of elevator deflections of loo to -15 was investigated at zero sideslip, and the aileron and rudder effectiveness was me
16、asured over the sideslip range. The tests were made at thrust coefficients of 0, 0.14, and 0.30 qhich represent a flight con- dition of low power or high speed (where the thrust coefficient approaches zero), best climb, and full power as in take-off or wave-off, respectively. The thrust coefficients
17、 are equated to an installed 224 kW (300 hp). A propeller blade angle of 20 was used for all tests. Tail downwash surveys were made with a calibrated pitch-yaw head along the elevator hinge axis with the horizontal tail removed at zero sideslip for flap deflec- tions of Oo, loo, and 30 for T;: = 0,
18、0.14, and 0.30. The tests at TL = 0 and 0.14 were conducted at a tunnel speed of 27.4 m/sec (90 ft/sec) which gave a Reynolds number of approximately 2.8 X 106. The tests at TL = 0.30 were made at 22.8 m/sec (75 ft/sec) giving a Reynolds number of about 2.3 X lo6. The Reynolds number is based on a m
19、ean geometric chord of 1.50 m (4.91 ft). PRESENTATION OF DATA The longitudinal data from these tests have been corrected for blockage, airstream misalinement, buoyancy effects, mounting-strut tares, and wind-tunnel jet-boundary effects. The lift and drag have been corrected for the integrated averag
20、e airstream L-8682 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-misalinement. The lateral data have not been corrected for the lateral variation of the stream angle. It should be pointed out, however, that at least a part of the positive roll- i
21、ng moment noted at the lower angles of attack for most model conditions can be attrib- uted to lateral variation of the tunnel airstream angle as shown in figure 4. Calculations of section rolling moments using the spanwise variations of stream angle of figure 4 indicated that the total measured out
22、-of -trim rolling moment could be approximately accounted for by the airstream angularity. Similar out-of-trim rolling moments were also noted in references 1 to 5. The data are presented as follows: Figure With propeller removed . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 With propell
23、er removed and zero thrust coefficient . . . . . . . . . . . . . 6 With power and flap deflection . . . . . . . . . . . . . . . . . . . . . . . . 7 to 9 With horizontal tail removed . . . . . . . . . . . . . . . . , . . . . . . . . 10 Variation of pitching moment with elevator deflection . . . . . .
24、 . . . . . . 11 Longitudinal characteristics: Lateral characteristics: With propeller removed . . . . . . . . . . . . . . . . . . . . . . . . . . . . With power and flap deflection . . . . . . . . . . . . . . . . . . . . . . . . With vertical tail removed . . . . . . . . . , . . . . . . . . . . . .
25、. , . . With aileron deflection, 6f = 0 . . . . . . . . . . . . . , . . . . . . . . . With aileron deflection, sf = 30 . . . . . . . . . . . . . . . . . . . . . . . With rudder deflection, Sf = 0 . . . . . . . . . . . . . . . . . . . . . . . With rudder deflection, Sf = 30 . . . . . . . . . . . . .
26、. . . . . . . . . . . . . . . . . . . . . Lateral stability characteristics with propeller removed Lateral stability characteristics with propeller removed and at zero thrust. . . . . . . . . . . . . . . . . . . . . . , . . . . . . . . . . . . Lateral and directional stability characteristics with v
27、ertical tail removed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Downwash at tail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Dynamic pressure at tail . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator hinge-moment coefficients . . . . . . . .
28、. . . . . . . . . . , . . . Aileron hinge-moment coefficients . . . . . . . . . . . . . . . . . . . . . . Rudder hinge-mom ent coefficients . . . . . , . . . . . . . . . . . . . . . . . Effect of power on longitudinal characteristics . . . . . . . . . . . . . . . . . 12 13 to 15 16 and 17 18 to 20 2
29、1 to 23 24 to 26 27 to 29 30 31 32 33 to 35 36 to 38 39 40 and 41 42 and 43 44 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Figure 45 Stick-fixed longitudinal characteristics . 46 and 47 Stick-free longitudinal characteristics . 48 Effect of pow
30、er on Lift-curve slope and maximum lift coefficients . Effect of power on stick-free stability . 49 Variation of elevator effectiveness with angle of attack 50 Effect of power on elevator effectiveness 51 Flow conditions at tail 52 Effective dihedral characteristics 53 Directional stability characte
31、ristics 54 Effect of power on directional stability 55 Ai le ron effectiveness . 56 Rudder effectiveness . 57 and 58 Rolling- and yawing-moment coefficients for various power settings and flap deflections . 59 Control capability 60 RESULTS AND DISCUSSION The basic data obtained during the wind-tunne
32、l investigation are presented in fig- ures 5 to 43 without analysis. Summary plots have been prepared from some of these data to illustrate the general static stability and control characteristics of the model. Only the summary plots are discussed. Longitudinal Characteristics The longitudinal chara
33、cteristics of the model with various power conditions are presented in figure 44 for flap deflections of Oo, loo, and 30. As might be expected, increasing power results in an increase in lift-curve slope and maximum lift coefficient because of the increased slipstream velocity over the wing. This ef
34、fect of power on the lift characteristics is summarized in figure 45 where lift-curve slope and maximum lift coefficient are shown as a function of thrust coefficient. The pitching-moment curves shown in figure 44 are fairly linear up to angles of attack near the stall and do not generally exhibit t
35、he nose-down pitching moment at the stall usually associated with a straight-wing airplane. The curves of the pitching moment 7 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-against angle of attack indicate that increasing thrust has little effect
36、on the longitudinal characteristics except for a nose-up trim change for the full-flap configuration. These effects of power on the static longitudinal stability are summarized in figures 46 and 47 where variations of the static stability parameter aCm/8CL are plotted against CL and T;, respectively
37、. These figures show that, in general, the level of stability increased with an increase in lift coefficient and that increased thrust coefficient caused a decrease in stability particularly for the flap-deflected configurations. It might be pointed out here that the model shows a high level of stat
38、ic longitudinal stability for the forward center-of- gravity location. The data indicate that the model would become no worse than neutrally stable with the center of gravity as far aft as 0.40E. The stick-free static stability characteristics , determined from the pitching- and hinge-moment curves,
39、 are shown in figure 48. The data show that the model had stick- free stability for the flap and power conditions tested. The effect of thruqt coefficient on the stick-free longitudinal stability at a lift coefficient of 1.0 is presented in figure 49. These data show power has a more destabilizing e
40、ffect with a full-flap deflection (fjf = 30) than with the lower deflections (6f = Oo, loo), and the model would have stick-free stability with the center of gravity as far aft as 0.4OF. The variation of elevator effectiveness with angle of attack is presented in figure 50 for flap deflections of Oo
41、, loo, and 30. These data show that the effectiveness is only slightly reduced at the higher angles of attack, and effectiveness is still maintained up to the stall. Presented in figure 51 is the variation of elevator effectiveness with thrust coefficient. These data show that power increased the el
42、evator effectiveness. Presented in figure 52 is the variation of the average downwash angle and the dynamic pressure ratio at the tail with angle of attack for the flap and power conditions investigated. These data show a large increase in downwash angle with flap deflection. Also, the effects of po
43、wer on the downwash angle are fairly large, particularly at high angles of attack and full flap deflection. Lateral Characteristics The variation of the effective-dihedral parameter C with angle of attack is lP shown in figure 53 for the various flap and power conditions tested. These data show that
44、 the model has positive effective dihedral (-Cd for all conditions to CLYmax. The values of C vary widely depending upon angle of attack, flap, or power condition which means the response of the airplane to gusts or to rudder inputs to raise a wing could vary with the airplane configuration and flig
45、ht conditions. IP The variation of the directional stability parameter C with angle of attack is “P shown in figure 54 for the flap and power conditions investigated. These data show that 8 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-although the
46、 directional stability generally decreases with increasing angle of attack up to the stall, the model is directionally stable for all conditions studied. The effect of power on the directional stability characteristics is presented in figure 55 and the data show that, in general, power did not have
47、a major effect on the directional stability. The variation of the aileron effectiveness with angle of attack for sideslip angles of Oo and = 0, 0.14, and 0.30. These data show that, in general, the aileron effectiveness remains at a fairly constant level throughout the angle-of-attack range and is r
48、elatively unaffected by flap deflection, power setting, or angle of sideslip. The variation of rudder effectiveness with angle of attack for sideslip angles from 8O to -8O is presented in figure 57. These data show that rudder effectiveness is main- tained throughout the angle-of-attack range for al
49、l test conditions. The effect of power on the rudder effectiveness is shown in figure 58. The data show that power caused an increase in effectiveness with flaps up but had little effect with flaps fully deflected. The basic lateral characteristics of the model, as shown by the variation of the lateral coefficients Cl and Cn with angle of attack for Oo sideslip, are presented
