1、NASA TECHNICAL NOTE NASA TN D-5700 0 0 FULL-SCALE WIND-TUNNEL INVESTIGATION OF THE STATIC LONGITUDINAL A LATERAL CHA CTERISTICS OF A LIGHT SINGLE-ENGINE AIRPLANE by Marvin P. Fink, Delma C. Freeman, Jr., and H. Dogglus Greer Langley Research Center La% = 0, 0.24 and 0.46. Control effectiveness was t
2、aken for a full range of deflections on the aileron, elevator, rudder, and flap. Downwash measurements at tail were also obtained for the range of thrust coefficient and flap deflection. 17. Key Words Suggested by Author(s1 118. Distribution Statement Light single-engine airplane Stability and contr
3、ol Tail downwash Unclassified - Unlimited I 19. Security Classif. (of this report) 20. Security Classif. (of this page) 21- No. of Pages 22. priceB Unclassified Unclassified 138 $3.00 “For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 Provided
4、 by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-FULL-SCALE WND-TUNNEL INVESTIGATION OF THE STATIC LONGITUDINAL AND LATERAL CHARACTERISTICS OF A LIGHT SINGLE-ENGINE AIRPLANE By Marvin P. Fink, Delma C. Freeman, Jr., and H. Douglas Greer Langley Research Cente
5、r SUMMARY A force test investigation has been conducted in the Langley full-scale tunnel -to determine the static longitudinal and lateral stability and control characteristics of a full-scale, light, single -engine airplane. The investigation was made over an angle -of - attack range of -4O to 20 a
6、t various angles of sideslip between 58 for various power and flap settings. The power conditions were a thrust coefficient T: of zero which represents either a low-power or a high-speed condition (where the thrust coefficient approaches zero), T; = 0.20 which corresponds to a climb condition, and T
7、; = 0.46 which corresponds to a take-off condition. The investigation showed that the airplane has stick-fixed longitudinal stability ior angles of attack up to and through the stall for all configurations tested with the center of gravity located at 0.10 mean aerodynamic chord. Power generally has
8、a small destabi- lizing effect but the airplane is statically stable even with the most rearward center-of- gravity location. The airplane is directionally stable and has positive effective dihedral through the stall for all conditions tested. The aileron and rudder effectiveness viras maintained th
9、rough the stall and was powerful enough to trim out all airplane moments through the stall. INTRODUCTION For the past several years the NASA Flight Research Center has been concluetishg a program to evaluate the flying qualities of a number of general aviation aircraft. The results of these investig
10、ations have been reported in reference 1. As a part of the eon- tinuing investigation, one of the airplanes investigated in reference 1, a light twin-engine configuration, was tested in the Langley full-scale tunnel, and the results have been reported in reference 2. The next phase of the wind-tunne
11、l program was to investigate the characteristics of the single-engine version of the airplane of reference 2. The investigation was made to determine the static longitudinal and lateral stability and Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-co
12、ntrol characteristics with various power and flap settings over a range of angles of attack from -4O to 2Q0 and over a range of sideslip angles of *8. The tests except those at thrust coefficients of 0.46 and 0.55 were made at a tunnel speed of about 93 feet per second which gives a Reynolds number
13、of approximately 2.96 X lo6. Tests at 0.46 and 0.55 thrust coefficient were made at tunnel speeds of 54.8 and 77.0 feet per second, re speetiv ely . 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 da
14、ta are computed about the moment center shown in figure 2 which is at airplane longitudinal station 85, or 10.0 per- cent of the mean aerodynamic chord and 1.0 ft (0.30 m) below the reference line. b wing span, 35.98 ft (10.97 m) lC 1) drag coefficient, Drag/qS 1 “ L, lift coefficient, ift/qS Crr, p
15、itching-moment coefficient, Pitching moment/qc ell yawing-moment coefficient, Yawing moment/qSb C rolling-moment coefficient, Rolling moment/qSb C Y side-force coefficient, Side force/qS ol lift-curve slope at cu = 0, untrimmed Cz lateral stability parameter Cnp directional stability parameter - lon
16、gitudinal stability parameter CL l8,a aileron effectiveness parameter - C%,r rolling effectiveness of rudder 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-aCm horizontal-tail effectiveness parameter, - , Per deg yawing effectiveness of aileron ru
17、dder effectiveness parameter, -, per deg 8%- mean aerodynamic chord, 5 ft (1.53 m) propeller diameter, 6.42 ft (1.96 m) propeller speed, revolutions/sec free-stream dynamic pressure, lbf/ft% (/m2) ratio of dynamic pressure at tail to free-stream dynamic pressure wing area, 178 ft2 (16.50 m2) effecti
18、ve Dragpropellers removed - Dragpropellers operating thrust coefficient, T/qS at a = O0 free-stream velocity, ft/sec (m/sec) propeller advance ratio longitudinal axis angle of attack of fuselage reference line, deg angle of sideslip, positive when nose is to left, deg total aileron deflection, posit
19、ive when right aileron is down, (6a)Left - (6a)Righty deg flap deflection, positive when trailing edge is down, deg rudder deflection, positive when trailing edge is left, deg Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-E Subscript: max horizonta
20、l-tail deflection, positive when trailing edge is down, deg downwash angle at tail, deg maximum AIRPLANE The airplane tested was a light, single-engine, low-wing monoplane having a maxi- mum take-off weight of 3100 lb (13 800 N). Figure 2 gives the principal dimensions and figure 3 shows the airplan
21、e mounted in the tunnel test section. The airplane had a wing span of 35.98 feet (10.97 m), a wing area of 178 ft2 (16.50 m2), an aspect ratio of 7.3, and a mean aerodynamic chord of 5 feet (1.53 m) based on projection of the outboard leading edge of the wing through the fuselage. The wing airfoil s
22、ection was a modified NACA 642A215 airfoil with the trailing-edge cusp faired out. The wing had 5O of geometric dihedral and was at 2O positive incidence with respect to the fuselage reference line. Normally the airplane wing has no twist, but measurements of the test vehicle taken at the wing root
23、and wing tip showed that the left wing tip had lo of positive incidence with respect to the wing root. Power was provided by a variable-frequency electric motor. The thrust axis was canted 3O to right and 2.75O downward to the reference line. The airplane had a standard three-control system. The hor
24、izontal tail was of the all-movable type with a travel of 3.4O to -12.8O. The horizontal tail had a geared trailing-edge tab which moved in the same direction as the tail with a deflection ratio (tab deflectionltail deflection) of 1.5. The travel of each aileron was from 15O to -17.80. The rudder tr
25、avel was *25O. TESTS The tests were made to determine the static longitudinal and lateral stability and control characteristics of the airplane over a wide range of flight conditions. The air- plane was tested over an angle-of-attack range of -4 to 20 and over a sideslip range of i8O for the clean c
26、ondition (6f = 0, gear up) and for 15 and 32 flap deflections with gear down. A range of tail incidence angles from 3.4O to -18.8O was investigated at zero side- slip, and the aileron and rudder effectiveness was measured over the sideslip range. The tests were made at thrust coefficients of T: = 0,
27、 0.20, and 0.46 which represent flight conditions of low power or high speed, a climb at best angle at about 90-percent power, and at full power as in take-off, respectively. The test thrust coefficients were based on Provided by IHSNot for ResaleNo reproduction or networking permitted without licen
28、se from IHS-,-,-installed horsepower of 260 . Several tests were made at a thrust coefficient of 0.55 which would be representative of a 355 horsepower engine . The propeller blade angle. and consequently the advance ratio. for each thrust coefficient was set at a fixed value which was representativ
29、e of that for flight conditions at which the particular value of thrust coefficient could be achieved . The values of V/nD were 0.98, 0.49, 0.33, and 0.33 for values of T; of 0. 0.20, 0.46, and0.55, respectively . Apropeller blade angle of 19.5O was set for T; = 0 and 0.46; and 230 was used for T; =
30、 0.20. Tail downwash surveys were made along the horizontal tail hinge axis with the tail off at zero sideslip for flap deflections of OO. 15O. and 32O for T; = 0. 0.20, and 0.46. PRESENTATION OF DATA The data from these tests have been corrected for airstream misalinement. hori- zontal buoyancy eff
31、ects. mounting strut tares. and wind-tunnel jet -boundary effects . The data are presented in the following figures: Figure Longitudinal characteristics with propeller removed . 4 Longitudinal characteristics with propeller removed and zero thrust 5 Longitudinal characteristics with power and flap d
32、eflection . 6 to 8 Longitudinal characteristics with horizontal tail removed 9 Variation of pitching-moment coefficient with tail deflection . 10 Lateral characteristics with propeller removed . 11 . Lateral characteristics with power and flap deflections 12 to 14 Lateral characteristics with vertic
33、al tail removed 15 and 16 . Lateral characteristics with aileron deflection. 6f = o0 17 and 18 Lateral characteristics with aileron deflection. 6f = 32O . 19 and 20 Lateral characteristics with rudder deflection. 6f = O0 21 to 24 Lateral characteristics with rudder deflection. 6f = 32O . . 25 to 28
34、Lateral stability characteristics with propeller removed 29 Lateral stability characteristics with propeller removed and at zero thrust . 30 Lateral and directional stability characteristics with vertical tail removed . 31 Downwash at tail 32 to 34 Dynamic pressure at tail 35 to 37 Effect of power o
35、n longitudinal characteristics 38 Effect of power on lift-curve slope and maximum lift coefficient 39 Effect of power on longitudinal stability 40 . Effect of power on horizontal-tail control power 41 and 42 Flow conditions of tail . 43 Effective dihedral characteristics . . 44 Provided by IHSNot fo
36、r ResaleNo reproduction or networking permitted without license from IHS-,-,-Figure Directional stability characteristics . 45 and 46 Aileron effectiveness . 4 7 Yawing effectiveness of aileron 4 8 Rudder effectiveness . 4 9 Rolling effectiveness of rudder 50 Effect of power on rudder effectiveness
37、51 Comparison of rolling- and yawing-moment coefficients . fo:r various power and flap deflections 5 2 . Control capability 5 3 RESULTS AND DISCUSSION The basic data obtained during the wind-tunnel investigation are presented in fig- ures 4 to 37 without analysis. Summary plots have been prepared fr
38、om some of these data to illustlrate the general static stability and control characteristics of the airplane. Only the summary plots are discussed. Longitudinal Characteristics The longitudinal characteristics of the airplane with various power conditions are presented in figure 38 for flap deflect
39、ions of 0, 15O, and 32. 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 effect of power on the lift characteristics is summarized in figure 39 where lift-curve slope and ma
40、ximum lift coefficient are shown as functions of thrust coefficient. The pitching-moment curves shown in figure 38 are virtually linear through the stall and do not exhibit the nose-down pitching moment at the stall usually associated with a straight-wing airplane. The variation of the pitching-mome
41、nt curves with angle of attack indicates that increasing thrust has little effect on the longitudinal characteristics except for a trim change. These power effects are further illustrated in figure 40 where .the variation in static margin acm/8cL with thrust coefficient is presented. These data are
42、a measure of the stick-fixed stability and show that power is destabilizing. The effects are generally small, however, and the airplane would have high static stability (even at the aft center -of -gravity location (longitudinal station 92 or 0.22c). The variation of horizontal-tail effectiveness wi
43、th angle of attack is presented in figure 41 for flap deflections of oO, 15O, and 32O. These data show that there is a rela- tively small reduction in tail effectiveness at the higher angles of attack, particularly with flaps down, and that the general level of tail effectiveness is little affected
44、by flap deflec- tion for a given power condition. The tail effectiveness is presented as a function of Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-thrust coefficient in figure 42 for each flap deflection of the tests. These data show that power.i
45、ncreased the tail effectiveness as would be expected, the slipstream acting on part of the tail. Presented in figure 43 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 were obta
46、ined from surveys and show a large increase in down- wash angle with flap deflection. Also, there is little effect of power on the downwash angle except for tif = 32O and T: = 0.46. The dynamic-pressure ratio is relatively unaffected by flap deflection but increases with power, as might be expected.
47、 Lateral Characteristics The variation of the effective-dihedral parameter Cz with angle of attack is P shown in figure 44 for the various flap and power conditions of the test. The data show that the airplane has positive effective dihedral (-CZ) in all conditions except for Of = 32O and T; = 0.55
48、where Cz is about zero in the middle angle-of-attack range. P The value of -CzP varies widely depending upon the angle of attack, flap, or power con- dition; this statement means the response of the airplane to gusts or to rudder inputs to raise a wing could vary with the airplane configuration and
49、flight condition. The variation of the directional stability parameter Cn with angle of attack is P shown in figure 45 for the various flap and power conditions. These data show that the airplane is directionally stable in all conditions over the entire angle-of -attack range. As was the case for the effective dihedral, the value of C varies considerably for the nP differen
copyright@ 2008-2019 麦多课文库（www.mydoc123.com）网站版权所有
备案/许可证编号：苏ICP备17064731号-1