NASA-TN-D-5408-1969 Static and dynamic stability derivatives of a model of a jet transport equipped with external-flow jet-augmented flaps《装配有外流喷气扩张襟翼喷气运输机模型的静态和动态稳定性导数》.pdf

《NASA-TN-D-5408-1969 Static and dynamic stability derivatives of a model of a jet transport equipped with external-flow jet-augmented flaps《装配有外流喷气扩张襟翼喷气运输机模型的静态和动态稳定性导数》.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-5408-1969 Static and dynamic stability derivatives of a model of a jet transport equipped with external-flow jet-augmented flaps《装配有外流喷气扩张襟翼喷气运输机模型的静态和动态稳定性导数》.pdf（53页珍藏版）》请在麦多课文档分享上搜索。

1、STATIC AND DYNAMIC STABILITY DERIVATIVES OF A MODEL OF A JET TRANSPORT EQUIPPED WITH EXTERNAL-FLOW JET-AUGMENTED FLAPS by Delmu C. Freeman, Jr., Sive B. Grafton, und Richurd DAmato Ldngley Reseurcrb Center Lungley Stution, Hampton, Vu, NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C.

2、SEPTEMBER 1969 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM 0332205 1. Report No. 2. Government Accession No. NASA TN D-5408 I 4. Title and Subtitle STATIC AND DYNAMIC STABILITY DERIVATIVES OF A MODEL OF A JET TRANSPORT EQUIP

3、PED WITH EXTERNAL-FLOW JET-AUGMENTED FLAPS 7. Authods) Delma C. Freeman, Jr., Sue B. Grafton, and Richard DAmato 9. Performing Organization Name and Address NASA Langley Research Center Hampton, Va. 23365 2. Sponsoring Agency Name and Address National Aeronautics and Space Administration Washington,

4、 D.C. 20546 5. Supplementary Notes 6. Abstract 3. Recipients Catalog No. 5. Report Data September 1969 6. Performing Orgonizotion Cod 8. Performing Organization Rep L-6521 IO. Work Unit No. 121-01- 11-03-23 11. Controct or Grant No. 13. Type of Report and Period C Technical Note 14. Sponsoring Agenc

5、y Code A wind-tunnel investigation has been made to determine the static and dynamic stability derivatives of a model of a large jet transport equipped with external-flow jet-augmented flaps. The tests were conducted in thi Langley full-scale tunnel, and a model powered by scale-model, compressed-ai

6、r-driven turbofan engines was u The results of the investigation showed that blowing on the flap system increased the lift-curve slope, delayed the stall, and increased the maximum lift coefficient. The data also showed that all model configuratio generally had static longitudinal stability over the

7、 test angle-of-attack range except those at the higher flap-deflection angles where the effects of power were destabilizing. The results also showed that the model had pc tive damping in pitch, roll, and yaw throughout the test angle-of-attack range up to and slightly beyond the stall for all test c

8、onditions. The application of power in the flap system resulted in appreciable increases in r damping but produced essentially no effects in pitch and yaw damping. There were essentially no effects due to oscillation frequency on the damping derivatives. 17. Key Wards Suggested by Author(s) 18. Dist

9、ribution Statement let-augmented flaps Unclassified - UnlimitedDynamic sta bi I ity 19. Security Classif. (of +his report) M. Security Classif. (of this page) Unclassified Unclassified Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-STATIC AND DYNAMI

10、C STABILITY DERIVATIVES OF A MODEL OF A JET TRANSPORT EQUIPPED WITH EXTERNAL-FLOW JE T-AUGMENTE D FLAPS By Delma C. Freeman, Jr., Sue B. Grafton, and Richard DAmato Langley Research Center SUMNIARY A wind-tunnel investigation has been made to determine the static and dynamic stability derivatives of

11、 a model of a large jet transport equipped with external-flow jet-augmented flaps. The tests were conducted in the Langley full-scale tunnel, and a model powered by scale-model, compressed-air-driven turbofan engines was used. The results of the investigation showed that blowing on the flap system i

12、ncreased the lift-curve slope, delayed the stall, and increased the maximum lift coefficient. The data also showed that all model configurations generally had static longitudinal stability over the test angle-of-attack range except those at the higher flap-deflection angles where the effects of powe

13、r were destabilizing. The results also showed that the model had positive damping in pitch, roll, and yaw throughout the test angle-of-attack range up to and slightly beyond the stall for all test conditions. The application of power in the flap system resulted in appreciable increases in roll dampi

14、ng but produced essentially no effects in pitch and yaw damping. There were essentially no effects due to oscillation frequency on the damping derivatives. INTRODUCTION A great deal of interest has been shown in the external-flow jet-augmented-flap concept as a means of achieving high lift coefficie

15、nts. In this concept, the jet efflux from pod-mounted engines is deflected upward to blow over the flaps and through slots of the flaps and, thus, induces very high lift on the wing. Early experimental investigations of external-flow jet-augmented flaps on general research models (for example, see r

16、efs. 1to 5) have demonstrated that desirably high lift coefficients can be generated with this system. Recently, a program has been started at the Langley Research Center to investigate the application of the external-flow jet-augmented flap to a large jet transport with high bypass-ratio turbofan e

17、ngines. The results of conventional static wind-tunnel tests of this configuration are reported in reference 6 and show that the external-flow jet-augmented flap offered a promising means of achieving improved take-off and landing Provided by IHSNot for ResaleNo reproduction or networking permitted

18、without license from IHS-,-,-LllllII I I 1111111111111 11111111.11111.1.1111.11.11.1 .II-I .111-1.1 I performance for large jet transports. Because of these promising results, a program has been initiated to evaluate the dynamic stability, flight characteristics, and general piloting techniques of s

19、uch a configuration. This work is to be conducted with a fixed-base simulator requiring aerodynamic inputs in the form of static and dynamic stability derivatives of the particular configuration under study. As part of the overall program, the present investigation was undertaken to measure the stat

20、ic and dynamic stability derivatives of the jet-transport model with an external-flow jet-augmentedflap. The model used in the investigation was powered by four high-bypass-ratio turbo fan engines and could be equipped with double-slotted trailing-edge flaps for use in an external-flow jet-augmented

21、-flap system. The flap configurations tested represented possible landing-approach and take -off configurations. The dynamic stability derivatives were determined in pitching, rolling, and yawing forced-oscillation tests at two different frequencies over an angle-of-attack range. In order to help in

22、terpret the dynamic data, the static longitudinal and lateral stability characteristics of the model were also deter mined and are presented. SYMBOLS The longitudinal data are referred to the stability-axis system and the lateral data are referred to the body-axis system. (See fig. 1.) The origin of

23、 the axes was located to correspond to the center-of -gravity position (0.25 mean aerodynamic chord) shown in figure 2. In order to facilitate international usage of data presented, dimensional quantities are presented both in the U.S. Customary Units and in the International System of Units (SI). E

24、quivalent dimensions were determined by using the conversion factors given in reference 7. b wing span, feet (meters) aCL cLa lift-curve slope, aa, C local wing chord, inches (centimeters) -C mean aerodynamic chord, feet (meters) FD drag force, pounds (newtons) FL lift force, pounds (newtons) FX for

25、ce along X-axis, pounds (newtons) 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-FY FZ f it k MX MY MZ P q qca r S T lateral force, pounds (newtons) force along Z-axis, pounds (newtons) frequency of oscillation, cycle/second (1 cycle/second = 1her

26、tz) horizontal-tail incidence angle, degrees reduced-frequency parameter, wb/2V or d/2V rolling moment, foot-pounds (meter -newtons) pitching moment, foot-pounds (meter-newtons) yawing moment, foot -pounds (meter-newtons) rolling velocity, radians/second pitching velocity, radians/second free-stream

27、 dynamic pressure, pV2/2, pounds/foot2 (newtons/meterZ) yawing velocity, radians/second wing area, feet (meted) thrust, pounds (newtons) free-stream velocity, feet/second (meters/second) x,y,z body reference axes Xs,Ys, Zs stability reference axes a! angle of attack, degrees or radians a! rate of ch

28、ange of angle of attack, radians/second P angle of sideslip, degrees or radians 3 V Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- rate of change of angle of sideslip, radians/second elevator deflection, positive when trailing edge down, degrees 6f

29、 1 forward trailing-edge flap-segment deflection, degrees 6f 2 aft trailing-edge flap-segment deflection, degrees 6sl leading-edge slat deflection, degrees P air density, slugs/foot3 (kilograms/meterS) angle of roll, degrees or radians w angular velocity, 2d, radians/second cx=-FX cz =-FZ qms qms m

30、cy=-FY qcos rblr - rb cy, =-aCY a-2v a-2v acz aCY czp = ap CYp = ap 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I TESTS, EQUIPMENT, AND TECHNIQUE Wind Tunnel The tests were made in the 30- by 6O-foot (9.1- by 18.3-meter) open-throat test sectio

31、n of the Langley full-scale tunnel with the model mounted about 10 feet (3.05 meters) above the ground board. The model was so small in proportion to the tunnel test section that no wind-tunnel wall corrections were needed or applied. Normal corrections for flow angularity were applied. Apparatus an

32、d Model The investigation was conducted on the four -engine, high-wing, jet-transport rmM illustrated by the three-view drawing of figure 2(a). The dimensional characteristics of the model are given in table I. The wing had an average leading-edge sweep angle of 28 and incorporated leading-edge slat

33、s and double-slotted trailing-edge flaps. A detailed sketch of the flap assembly and engine-pylon arrangement is shown in figure 2(b). The wing airfoil section was the same as that used in reference 6. The forward and aft flap-deflection angles tested represent approximately a landing-approach condi

34、tion (6f1/6f2 = 3Oo/6O0) and take-off conditions (6f1/6f2 = 2Oo/4O0 and 6f1/6f2 = 10/200). The slat-deflection angles represent these same conditions. Photographs of the model and flap system are shown in figures 3(a) and 3(b), respectively. To facilitate model configuration changes and to insure ac

35、curate flap-deflection angles, the wing of the model was designed with removable trailing edges. To convert the model from the clean configuration to each of the flap-deflected configurations, the clean trailing edges were replaced with trailing-edge flaps constructed with fixed gaps, overlaps, and

36、deflection angles. The leading-edge slats were designed so that they could be fastened to the wing leading edge at fixed positions when desired. The model engines represented high-bypass-rativurbofans and were installed at -3 incidence so that the jet exhaust impinged directly on the trailing-edge f

37、lap system. The engine turbines were driven by compressed air and turned fans which produced the desired thrust. All dynamic force tests were made with a single strut and sting support system and a strain-gage balance. Sketches of the forced-oscillation test equipment are presented in figure 4, and

38、the equipment is described in reference 8. The static force tests were made on a conventional sting which entered the rear of the fuselage. Tests and Procedures Calibrations were made to determine the engine-installed thrust as a function of engine speed in revolutions per minute with the model at a

39、n angle of attack of 0 and with 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Illll Ill I II I I I the trailing-edge flaps and leading-edge slats undeflected. The aerodynamic tests were made by setting the engine speed to give the desired thrust

40、at an angle of attack of 0 and then maintaining this engine speed as the model was tested through a range of angles of attack. The thrust calibrations were made at the free-stream dynamic pressure used in the present tests, 7.1 .lb/ft2 (340 N/m2). The value for the thrust used in computing the thrus

41、t coefficients for the forward flight tests is the difference between the longitudinal force with power on and the longitudinal force with power off and flow-through nacelles, both for the same free-stream dynamic pressure. The longitudinal force with power off and flow-through nacelles was not actu

42、ally measured on the present model but was com puted by subtracting the increment of drag coefficient due to the windmilling of rotating parts from the drag coefficient of the model in the windmilling condition. The drag of the windmilling parts was assumed to be the same as that measured for simila

43、r but 60-percent larger scale-model engines in an unpublished investigation. There is the possibility of a small error in this procedure if the drag coefficient of the windmilling parts is not exactly the same on the larger and smaller engines. Even if the assumed drag coefficient of the windmilling

44、 parts of the present engines were 100-percent different, however, the error in thrust coefficient would only be about 5 percent for the lowest thrust coefficient of the tests and would be even smaller for the higher thrust coefficients. The thrust calibrations were made through a range of engine sp

45、eeds up to 60 000 revolutions per minute, at which speed the fans developed, in the static condition, their rated thrusts of approximately 30 pounds (133.44 newtons) each. Dynamic-force tests were made to determine the longitudinal and lateral oscillatory stability derivatives of all model configura

46、tions with power off and for thrust coefficients CT of 0.38, 0.78, and 1.70. These force tests were made over an angle-of-attack range from -4 to 24 for the three flap-deflection combinations previously discussed. The dynamic stability derivatives werg measured for an amplitude of 5.5 and for freque

47、ncies of 0.5 and 1.0 cycle per second corresponding to values of the reduced-frequency param eter k of 0.023 and 0.045, respectively, for the pitching tests and 0.160 and 0.321, respectively, for both the rolling and yawing tests. In order to help interpret the dynamic data, static-force tests were

48、also conducted to obtain the static longitudinal and lateral stability characteristics of the model. All tests were conducted with the rudder unde flected and with the tail incidence and elevator-deflection angle set to give longitudinal trim in the operational angle-of-attack range of the airplane.

49、 The force tests were conducted at a dynamic pressure of 7.1 lb/ft2 (340 N/m2) which corresponds to a Reynolds number of 0.543 X lo6 based on the mean aerodynamic chord of the model. 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-RESULTS AND DISCUSSION Static Stability Derivatives Lon