1、NASA-TECHNICAL NOTE c o* NASA TN- 0-3591-STATIC AND DYNAMIC LONGITUDINAL STABILITY DERIVATIVES OF A POWERED 1/9-SCALE MODEL OF A TILT-WING V/STOL TRANSPORT by Joseph R. Chumbers und Szce B. Grupon LungZey Reseurch Center LungZey Stution, Humpton, Vu. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WAS
2、HINGTON, D. C. a SEPTEMBER 1966 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHSI TECH LIBRARY KAFB,NM Illlllllllll111lllllll Ill11llllltillIll 0079994 NASA I”U-SDYJI STATIC AND DYNAMIC LONGITUDINAL STABILITY DERIVATIVES OF A POWERED 1/9-SCALE MODEL OF A
3、TILT-WING V/STOL TRANSPORT By Joseph R. Chambers and Sue B. Grafton Langley Research Center Langley Station, Hampton, Va. NATIONAL AERONAUT ICs AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $2.00 Provided b
4、y IHSNot for ResaleNo reproduction or networking permitted without license from IHSSTATIC AND DYNAMIC LONGITUDINAL STABILITY DEFUVATIVES OF A POWERED 1/9-SCALE MODEL OF A TILT-WING V/STOL TRANSPORT By Joseph R. Chambers and Sue B. Grafton Langley Research Center SUMMARY Static and oscillatory force
5、tests were conducted to determine the power-on longi tudinal stability derivatives of a model of a tilt-wing V/STOL transport aircraft. The model had four propellers and the wing could be tilted from an incidence of 00 (for con ventional forward flight) to 90 (for hovering flight). The investigation
6、 consisted of tests at several wing incidence angles and thrust conditions for an angle-of -attack range of *30. The forced oscillation equipment and dynamic data readout system are also described. The results indicate that the model was statically unstable for wing incidence angles above about 15 b
7、ut was stable for lower angles. The model had positive damping in pitch (negative values of the damping-in-pitch parameter MYq + Myb) throughout the cony,z a! wing incidence angle, degrees reduced frequency parameter, oC/2V balance calibration factor, volts output per volt-foot-pound (volts output p
8、er volt-meter -newton) value of lift for longitudinal acceleration equal to zero at an angle of attack of 00,pounds (newtons) pitching moment, foot-pounds (meter-newtons) static pitching moment at mean angle of attack of oscillation, foot-pounds (meter-newtons) pitching velocity, radians per second
9、pitching velocity increment, radians per second free-stream dynamic pressure, pV2/2, pounds per square foot (newtons per square meter) wing area, square feet (square meters) period of oscillation, seconds time, seconds reference time, seconds free-stream velocity, feet per second (meters per second)
10、 weight, pounds (newtons) wing loading, pounds per square foot (newtons per square meter) body reference axes (see fig. 1) angle of attack, degrees or radians 3 V Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS- angle-of-attack increment, degrees or radi
11、ans amplitude of incremental angle of attack during oscillation, degrees or radians flap deflection, degrees pitch angle, herein defined as angular displacement of X body axis from wind-tunnel center line, radians pitch-angle increment, radians amplitude of incremental pitch angle during oscillation
12、, degrees or radians air density, slugs per cubic foot (kilograms per cubic meter) angular velocity, 27rf, radians per second aFX - aFZ Fx,=Ti7 FZa -aa! - aMYFxd! - a FX Fzd!=-aFZ My and the wing was also fitted with a double-slotted flap whose geometric characteristics are shown in figure 4. The fl
13、ap was programed with a cam and follower to deflect as wing incidence changed. The variation of flap deflection angle with wing incidence angle is shown in figure 5. Also shown in figure 5 is the programed variation of the incidence of the all-movable horizontal tail. Wing-fuselage ramps were used t
14、o improve the airflow in the area of the wing ten ter section as the wing incidence was changed. Additional information relating to the model can be found in reference 1. Apparatus All force tests were made with a single strut or sting support system and strain-gage balances. The static force tests
15、were made with the model mounted on a conventional sting which entered the rear of the fuselage. The forced oscillation tests were made with the equipment sketched in figure 6. During the rigidly forced oscillation tests the model was mounted with its wings in a vertical plane. The strain-gage balan
16、ce to which the model was attached was mounted in a steel C-channel which was allowed a single degree of rotational freedom in a yoke-pivot assembly. The C-channel was forced to oscillate about a vertical axis by a 3-horsepower variable-speed electric motor and flywheel which were mounted directly o
17、n the vertical support column. The rotary motion of the flywheel was transformed into oscillatory motion by the vertical and horizontal connecting rods which were joined by a bellcrank. The amplitude of the oscillatory motion (limited to and second, in the process of subtracting the effects of inert
18、ia from the in-phase forces and moments, a certain amount of aerodynamic forces and moments which correspond to the still-air aero dynamic damping were also subtracted. Still-air damping has not been important with conventional aircraft because the still-air damping is usually an insignificant perce
19、ntage of the wind-on aerodynamic damping, but for V/STOL aircraft at very low airspeeds, or in hovering, this factor may become significant. 7 Provided by IHSNot for Resale-,-,-TESTS The static and oscillatory force tests were made for an angle-of-attack range of -30 to 30 for wing incidence angles
20、of IOo, 25O, 50, and 65O. The tests were made by setting the wing incidence and varying the angle of attack while holding a constant air speed and a constant power input to the model propellers. Additional static tests were conducted with a wing incidence angle of 90 to determine the longitudinal st
21、ability deriva tives of the configuration in hovering flight (variation of forces and moments with veloc ities along the X and Z body axes). A model cross-shafting failure eliminated the possibility of obtaining any power-on data for a wing incidence angle of Oo; however, static and oscillatory test
22、s were conducted for a wing incidence angle of Oo with propellers windmilling for an angle-of-attack range of -loo to 20. The damping data obtained during this phase of the investigation are believed to be applicable to the trimmed level flight condition at a wing incidence of Oo. The forced oscilla
23、tion tests were made with an oscillatory amplitude of *50. The range of oscillation frequencies was from 0.2 to 1.4 cycles per second. The frequency of the oscillation was held constant for a wing incidence of 65O and the reduced frequency parameter k was held constant for the other incidence angles
24、 investigated. (It is believed that the reduced frequency parameter k loses significance for very low speeds; the period of the oscillation was therefore held constant.) The location of the moment reference center for all tests is given as a function of wing incidence angle in figure 7. These locati
25、ons correspond to the center-of -gravity locations of the model during the free-flight tests of reference 1. During all tests the main propeller blades were set at an angle of 120 at the 0.75 radius location. The tail rotor blades were maintained at a blade angle of Oo during the investigation. The
26、primary purpose of the forced oscillation tests was to determine the dynamic stability derivatives for the condition of zero longitudinal acceleration at an angle of attack of 00 for each value of iw. Constant power input to the model was maintained as angle of attack was changed during each test. T
27、he effect of thrust condition (simulating accelerated and decelerated flight conditions) on the dynamic stability derivatives was determined by conducting similar tests at different tunnel airspeeds with the same con stant power input. RESULTS AND DISCUSSION The longitudinal stability characteristic
28、s measured in the investigation are discussed individually as static stability characteristics, in-phase oscillatory derivatives, and out-of -phase oscillatory derivatives. Longitudinal static stability characteristics measured in the investigation for wing incidence angles of 90, 65“, 50, 25O, loo,
29、 and 0 are presented 8 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHSin figures 8 to 13. The in-phase oscillatory derivatives are presented in figures 14 to 18 and are compared with the static data in figures 19 to 22. The out-of-phase oscillatory deriv
30、atives are presented in figures 23 to 27. In addition, scaled-up values of the model stability characteristics are presented in figures 28 and 29 for a hypothetical full-scale airplane for trimmed level flight throughout the forward speed range from hovering to conventional forward flight. Presented
31、 in table II are average values of the repeatabil ity obtained during measurement of the oscillatory derivatives. Static Stability Characteristics The data of figure 8 show that, for the model in the hovering configuration, forward velocity produced positive pitching moments. This speed stability, a
32、lthough statically stabilizing, is primarily responsible for the controls-fixed unstable oscillation which occurred during the free-flight tests of reference 1. (A mathematical treatment of the importance of speed stability on handling qualities during hovering flight is given in ref. 3.) The data a
33、lso indicated no appreciable variation of vertical force FZ with forward speed. This aerodynamic characteristic, as pointed out in reference 3, leads to uncoupling of the vertical degree of freedom from the horizontal and angular degrees of freedom and makes the longitudinal oscillation essentially
34、a two-degree-of -freedom oscillation involving horizontal displacement and pitching motion. For the transition conditions, the variations of forces and moments presented in figures 9 to 12 show a progressive variation of the static longitudinal stability with wing incidence for angles of attack from
35、 about Oo to loo, the range of greatest interest. For wing incidences of 65O and 50 the model was statically unstable with respect to angle of attack positive Mya1 and was statically stable with respect to speed cpositive MyV At a wing incidence angle of 250, the model pitching-moment characteristic
36、s were ) extremely nonlinear with respect to angle of attack, although they generally indicated static instability in the angle-of-attack range from Oo to loo. In addition, the speed sta bility derivative MyV is positive for low speeds and negative for high speeds, although in the angle-of-attack ra
37、nge of Oo to loo it is in all cases negative for small variations in velocity about trimmed level flight. For a wing incidence of 100, the model was statically stable with respect to angle of attack and had negative speed stability (negative values of Myv). The foregoing stability trends are in agre
38、ement with the flight test results of reference 1where the stick-fixed dynamic stability of the model was observed to change from unstable oscillations for high wing incidence angles to stable motions at low angles. Presented in figure 13 are the variations with angle of attack of the conventional f
39、orce and moment coefficients for the model with a wing incidence of Oo and propellers wind-milling. These data are presented primarily as an aid in the interpretation of the damping data to be presented in a later section of this report. 9 Provided by IHSNot for Resale-,-,-In-Phase Oscillatory Deriv
40、atives The variation of the in-phase oscillatory derivatives presented in figures 14 to 18 , indicates that the model, in general, had positive values of the static stability parameter Mya - w2My4 for wing incidence angles of 25O and above. This result is the same as that shown by the static data of
41、 figures 19 to 22. Values of the static stability parameters for unaccelerated level flight at an angle of attack of Oo as obtained from the static data of figures 9 to 13 and the oscillatory tests of figures 14 to 18 are compared in figures 19 to 22. The comparisons are generally good, but noticeab
42、le differences are present for some conditions. For example, in figures 21 and 22 at angles of attack between about 100 and 25O, there are large differences between the static and oscillation data. The wing is generally stalled in these conditions, and the differences are believed to be caused by th
43、e lag in buildup of the stalled flow conditions during the oscillation tests. Out-of -Phase Derivatives The variation of the out-of-phase oscillatory derivatives presented in figures 23 to 27 indicates that thrust conditions had significant effects on the out-of -phase oscillatory derivatives. At fi
44、rst glance, the unsystematic variation of the derivatives with thrust condition implies a great deal of scatter in the measured data but a study of table 11 indicates reasonable repeatability. The large effects of thrust condition on the damping derivatives probably result from extreme changes in th
45、e flow patterns about the model. The model had positive damping in pitch (negative values of Myq + MYiY ) for all condi tions investigated. The magnitude of the damping in pitch increases as the wing incidence angle is decreased, as might be expected. The conventional damping-in-pitch parameter Cmq
46、+ Cmh is approximately independent of angle of attack for a wing incidence of Oo with the propellers windmilling. Stability Characteristics During Transition The results of the investigation for trimmed level flight at an angle of attack of 00 are summarized by the data of figures 28 and 29 which pr
47、esent the longitudinal stability characteristics of a full-scale airplane as functions of forward velocity from hovering to conventional forward flight. These data were obtained by scaling up the model data. The dimensional damping-in-pitch model data were scaled up by multiplying by the fourth powe
48、r of the model scale factor (94) and by the ratio of full-scale speed to model speed. The full-scale airplane was assumed to have a wing loading of 70 pounds per square foot (3352 N/m2) and a moment of inertia in pitch Iy of 125 000 slug-ft2 (169 476 kg-ma) on the basis of data presented in referenc
49、e 1. The unbalanced pitching moment which would have to be trimmed by the tail rotor or horizontal tail (a different tail incidence program being used) is presented as a 10 I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHSI function of velocity in figure 28. The pitching-moment values are positive for velocities below about 28 knots and negative for speeds above this point. The variation of the static stability parameter MyD - dMy4 presented in figure 29 shows the airplane
