NASA-TN-D-6440-1971 Dynamic stability derivatives of a jet transport configuration with high thrust-weight ratio and an externally blown jet flap《带有高推重比和外流喷气襟翼喷气式运输飞机结构的动态稳定性导数》.pdf

《NASA-TN-D-6440-1971 Dynamic stability derivatives of a jet transport configuration with high thrust-weight ratio and an externally blown jet flap《带有高推重比和外流喷气襟翼喷气式运输飞机结构的动态稳定性导数》.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-6440-1971 Dynamic stability derivatives of a jet transport configuration with high thrust-weight ratio and an externally blown jet flap《带有高推重比和外流喷气襟翼喷气式运输飞机结构的动态稳定性导数》.pdf（84页珍藏版）》请在麦多课文档分享上搜索。

1、NASA TECHNICAL NOTE d z DYNAMIC STABILITY DERIVATIVES OF A JET TRANSPORT CONFIGURATION WITH HIGH THRUST-WEIGHT RATIO AND AN EXTERNALLY BLOWN JET FLAP by Szie B. Grufton, Lysle P. Purlett, und Churles C. Smith, Jr. Lungley Reseurch Center Humpton, Vu. 23365 NATIONAL AERONAUTICS AND SPACE ADMINISTRATI

2、ON WASHINGTON, D. C. SEPTEMBER 1971 c Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB,NM I 111111 lllll11lIl lllllllllllllll11111Ill 111 0332940 . 1. Report No. 2. Government Accession No. NASA TN D-6440 i-4. Title and Subtitle DYNA

3、MIC STABILITY DERIVATIVES OF A JET TRANSPORT CONFIGURATION WITH HIGH THRUST-WEIGHT RATIO AND AN EXTERNALLY BLOWN JET FLAP 7. Author(s) Sue B. Grafton, Lysle P. Parlett, and Charles C. Smith, Jr. 9. Performing Organization Name and Address NASA Langley Research Center Hampton, Va. 23365 12. Sponsorin

4、g Agency Name and Address National Aeronautics and Space Administration Washington, D.C. 20546 15. Supplementary Notes _. 3. Recipients Catalog No. 5. Report Date September 1971 6. Performing Organization Code 8. Performing Organization Report No. L-7870 10. Work Unit No. 760-72-01-02 . . 11. Contra

5、ct or Grant No. 13. Type of Report and Period Covered Technical Note 14. Sponsoring Agency Code . 16. Abstract The investigation was conducted to determine the dynamic stability derivatives of an externally blown jet-flap transport configuration having clustered inboard pod-mounted engines and full-

6、span triple-slotted flaps. The results showed that the model had positive damping in pitch, roll, and yaw up to the stall angle of attack. The application of power resulted in an increase in pitch damping at high angles of attack and a moderate increase in yaw damping for the higher flap deflections

7、 but had no consistent effects on roll damping. For a given level of total engine thrust, the damping derivatives were generally not affected by frequency or by having one engine inoperative. -17. Ke; Words (Suggested by Authorb) ) 18. Distribution Statement Dynamic stability I Unclassified - Unlimi

8、ted Jet flap High thrust.-weight ratio 19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. NO. of Pages 22. Price* Unclassified Unclassified 81 $3.00 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-DYNAMIC STABILITY DERIVA

9、TIVES OF A JET TRANSPORT CONFIGURATION WITH HIGH THRUST-WEIGHT RATIO ANDANEXTERNALLYBLOWNJETFLAP By Sue B. Grafton, Lysle P. Parlett, and Charles C. Smith, Jr. Langley Research Center SUMMARY A wind-tunnel investigation was conducted in the Langley full-scale tunnel to deter mine the dynamic stabili

10、ty derivatives of an externally blown jet-flap transport configura tion having clustered inboard pod-mounted engines and full-span triple-slotted flaps. The investigation was made at a Reynolds number of 0.35 X lo6 based on the mean aerodynamic chord of the model. The results showed that the model h

11、ad positive damping in pitch, roll, and yaw up to the stall angle of attack. The application of power resulted in an increase in pitch damping at high angles of attack mainly because the tail damping was higher. Power also caused moderate increases in yaw damping for the higher flap deflec tions, bu

12、t the effects on roll damping were inconsistent. The effects of frequency on the damping derivatives were generally relatively small. For a given level of total engine thrust, the damping derivatives were not appreciably affected by having one engine inoperative. INTRODUCTION The present investigati

13、on was conducted to provide some fundamental information on the dynamic stability derivatives of a jet transport configuration with high thrust-weight ratio and an externally blown jet flap. Previous static stability and performance studies (refs. 1 and 2) have shown that the application of this con

14、cept to high-thrust weight-ratio turbofan aircraft was effective for producing the high lift required for short take-off and landing operation. Research was continued with full-span triple-slotted flaps, leading-edge boundary-layer control, and clustered engine arrangement as a means of improving th

15、e aerodynamic efficiency of such a configuration. (See ref. 3.) Because of the promising results achieved in the static stability and performance studies, a pro gram has been initiated to evaluate the dynamic stability, flight characteristics, handling qualities, and general piloting techniques of t

16、his configuration. The research is to be conducted with a fixed-base simulator requiring aerodynamic inputs in the form of static Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-and dynamic stability derivatives. As part of the overall program, the p

17、resent inves tigation was undertaken to measure the dynamic stability derivatives of the jet transport configuration with a high thrust-weight ratio and an externally blown jet flap. Results of a similar investigation for a lower range of thrust-weight ratios are presented in reference 4. The model

18、used in the investigation was powered by four simulated high-bypass ratio turbofan engines mounted in a clustered arrangement relatively close inboard and was equipped with full-span triple -slotted trailing-edge flaps and a fixed leading-edge flap. The model was also equipped with a leading-edge bl

19、owing system for use in some tests. The dynamic stability derivatives were determined in pitching, rolling, and yawing forced-oscillation tests at different frequencies, thrust conditions, and flap deflection angles for an angle-of-attack range from -5 to 35. Additional tests were made to determine

20、the dynamic stability for the model with the vertical tail off and for the model with various engine-out conditions. In order to aid in the interpretation of the dynamic force test data, the static longitudinal and lateral stability characteristics of the model were also determined and are presented

21、. SYMBOLS The dynamic longitudinal and lateral-directional data and the static lateral data are referred to the body-axis system; the static longitudinal data are referred to the stability-axis system. (See fig. 1.) The origin of the axes was located to correspond to the center-of -gravity position

22、(0.40 mean aerodynamic chord) shown in figure 2(a). Measurements and calculations were made in the U.S. Customary Units. They are presented herein in the International System of Units (SI) with the equivalent values in the U.S. Customary Units given parenthetically. Factors relating the two systems

23、are given in reference 5. b wing span, meters (feet) *DCD drag coefficient, qC2 FL CL lift coefficient, qms CLP= aC1 per degree or per radian MY Cm pitching-moment coefficient, q msc 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Cnp -per degree o

24、r per radian ac Cyp = 2,per degree or per radian engine gross-thrust coeffi i wing-semispan leading-edge blowing jet momentum coefficient, R/q,S local chord, meters (feet) mean aerodynamic chord, meters (feet) axial force, newtons (pounds) drag force, newtons (pounds) lift force, newtons (pounds) no

25、rmal force, newtons (pounds) force along X-axis, newtons (pounds) force along Y-axis, newtons (pounds) frequency of oscillation, hertz (cycles per second) horizontal-tail incidence angle, degrees reduced-frequency parameter, wb/2V or wE/2V rolling moment, meter -newtons (foot-pounds) pitching moment

26、, meter -newtons (foot -pounds) yawing moment, meter -newtons (foot -pounds) rolling angular velocity, radians/second pitching angular velocity, radians/second Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-free-stream dynamic pressure, pV2/2, newto

27、ns/meter2 (pounds/foot2) resultant force, newtons (pounds) yawing angular velocity, radians/second wing surface area, meters2 (feet2) total installed engine thrust, newtons (pounds) free-stream velocity, meters/second (feet/second) body reference axes stability reference axes flap coordinates, meter

28、s (feet) angle of attack, degrees or radians rate of change of angle of attack, radians/second angle of sideslip, degrees or radians rate of change of angle of sideslip, radians/second elevator deflection angle, positive when trailing edge is down, degrees deflection of rear element of trailing-edge

29、 flap (same as 6f3 in fig. 2(b), positive when trailing edge is down, degrees leading-edge flap deflection, positive when leading edge is down, degrees air density, kilograms/meter3 (slugs/foot3) angle of roll, degrees or radians angular velocity, 2nf, radians/second Provided by IHSNot for ResaleNo

30、reproduction or networking permitted without license from IHS-,-,-“ 2v Subscripts: left wing R right wing APPARATUS AND MODEL The tests were made in the 9.1- by 18.3-meter (30- by 60-ft) open-throat test sec tion of the Langley full-scale tunnel with the model mounted about 3.05 meters (10 ft) above

31、 the ground board. The model was so small in proportion to the tunnel test section that no wind-tunnel corrections were needed. The investigation was conducted on the four -engine, high-wing jet-transport model illustrated by the three-view drawing of figure 2(a). Dimensional characteristics of the

32、model are given in table I. The wing had a leading-edge sweep angle of 28.3 and incor porated the leading-edge flaps and triple-slotted trailing-edge flaps shown in figures 2(b) and 2(c). Coordinates for the three elements of the trailing-edge flaps are given in table II. The full-span trailing-edge

33、 flaps were divided into three spanwise segments on each wing semispan as indicated in figure 2(a). All three trailing-edge segments were deflected as a unit (a full-semispan flap), except where otherwise specified. 5 L Provided by IHSNot for ResaleNo reproduction or networking permitted without lic

34、ense from IHS-,-,-The leading edge of the wing was also equipped with a blowing system for boundary-layer control (fig. 2(d). Compressed air forced through a tube inserted in the wing leading edge exhausted through many small holes into the leading-edge plenum chamber and from there through the lead

35、ing-edge slot. To facilitate model configuration changes and to insure accurate flap deflection angles, the wing of the model was designed with removable trailing edges. Such remov able trailing edges were provided for the clean (flaps up) configuration and for several different flap-deflected confi

36、gurations with fixed gaps, overlaps, and deflection angles. Figure 2(b) shows the principal flap systems used: one, designated the “take-off flap,“ had deflections of 17/0.50/350 and the others, designated the “landing flaps, had deflec tions of 25/100/500 or 25/100/700. In addition, the landing fla

37、ps were constructed so that the rear flap element could be deflected differentially and locked into position for lat eral trim in the engine-out tests. In the remainder of the text and in all the data figures, only the deflection of the rear flap element is used for identification purposes. The mode

38、l engines (pod mounted) represented high-bypass-ratio turbofans and were installed at -3O incidence (with reference to X-axis) so that the jet exhaust impinged directly on the trailing-edge flap system. The engine turbines were driven by compressed air and turned fans which produced the desired thru

39、st. This model is described in detail in reference 3. A photograph of the model mounted for dynamic forced-oscillation tests in the full-scale tunnel is presented in figure 3. All the static and dynamic force tests were made with a single strut or sting-support system and with an internal strain-gag

40、e balance. Sketches of the forced-oscillation test equipment are presented in figure 4, and the equip ment is described in reference 6. TESTS AND PROCEDURES In preparation for the tests, the gross thrust of each engine was measured as a func tion of engine rotational speed in the static condition. T

41、he tests were then performed by setting the engine rotational speed to give the desired thrust and holding this speed constant through the range of angle of attack. The jet momentum of the leading-edge boundary-layer control system was evaluated by measuring the force produced by the jets in the win

42、d-off condition. Dynamic force tests were made to determine the longitudinal and lateral-directional oscillatory stability derivatives of all model configurations with power off and with power on for values of Cp of 1.74 and 3.48. These force tests were made over an angle-of attack range from -5 to

43、35 for 6f = 35“, 50, and 70. A few tests were also made with the clean configuration. The longitudinal stability derivatives were measured for an 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-amplitude of effective dihedral, however, increased wi

44、th an increase in angle of attack up to the stall. For all tail-on configurations, the application of thrust resulted in notable increases in directional stability throughout the test angle-of -attack range. At angles of attack near the power-off stall angle, thrust also produced increments of effec

45、tive dihe dral (fig. 8). Additional static lateral-directional stability data for the subject configura tion are presented in reference 3. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Dynamic Stability Derivatives Pitching. - The variations of the

46、 oscillatory pitching derivatives with angle of attack are presented in figures 9 to 15. The data for the clean configuration (fig. 9) show that the model had very small values of pitch damping with the tail off and that the addition of the tail resulted in relatively large values of pitch damping t

47、hat remained essentially con stant with angle of attack. The data of figures 10 to 14 show that the pitch damping for the flap-down configuration was much higher than that for the clean configuration, mainly because the tail damping was much higher. As the angle of attack was increased, how ever, th

48、e tail damping contribution decreased and, in the power-off condition, resulted in the model becoming undamped in some configurations. The addition of power is seen to minimize or eliminate the loss in tail damping as angle of attack increases; thus, the power-on model had relatively high values of

49、pitch damping over the test angle-of-attack range. A comparison of the data of figures ll(a) and ll(b), 12(a) and 12(b), and 14(a) and 14(b) shows that changes in oscillation frequency between 0.5 and 1.0 hertz (cps) had no appreciable effect on the pitch damping derivatives in the normal operating angle-of -atta