NASA-TN-D-8136-1976 Subsonic stability and control derivatives for an unpowered remotely piloted 3 8-scale F-15 airplane model obtained from flight test《飞行试验获得的无动力远程先导控制的3 8比例F-15飞.pdf

《NASA-TN-D-8136-1976 Subsonic stability and control derivatives for an unpowered remotely piloted 3 8-scale F-15 airplane model obtained from flight test《飞行试验获得的无动力远程先导控制的3 8比例F-15飞.pdf》由会员分享，可在线阅读，更多相关《NASA-TN-D-8136-1976 Subsonic stability and control derivatives for an unpowered remotely piloted 3 8-scale F-15 airplane model obtained from flight test《飞行试验获得的无动力远程先导控制的3 8比例F-15飞.pdf（34页珍藏版）》请在麦多课文档分享上搜索。

1、SUBSONIC STABILITY AND CONTROL DERIVATIVES FOR AN UNPOWERED, REMOTELY PILOTED 3/8-SCALE F-15 AIRPLANE MODEL OBTAINED FROM FLIGHT TEST -LOP.N COPY: RETURN TO Kenneth W. Ilz$ Richard E. Maine, AFM-L. TECHNICAL LIBRARY and Mary F. Shufer KIRTLAND AFB, M* y* Flight Research Center Edwards, Cali$ 93523 N

2、ATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. JANUARY 1976 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB,NM IIllill11111Ill11llllllllllIIIIIlllllIll1Ill1 1. Report No. 2. Government Accession NO. ._ 4. Title and S

3、ubtitle SUBSONIC STABILITY AND CONTROL DERIVATIVES FOR ANI UNPOWERED, REMOTELY PILOTED 3/8-SCALE F-15 AIRPLANE MODELI OBTAINED FROM FLIGHT TEST Kenneth W. Iliff, Richard E. Maine, and Mary F. Shafer 9. PerformingOrganization Name and AddressI NASA Flight Research Center P.O. Box-273I Edwards, Califo

4、rnia 93523 112. Sponsoring Agency Name and Address National Aeronautics and Space Administration Washington, D .C . 20546 15. Supplementary Notes 16. Abstract 3. Recipients Catalog No. 5. Report Date January 1976 6. Performing Organization Code 8. Performing Organization Report No. H-905 10. Work Un

5、it No 512-53-03 11. Contract or Grant No. 13. Type of Rewrt and Period Covered Technical Note 14. Sponsoring Agency Code -In response to the interest in airplane configuration characteristics at high angles of attack, an unpowered remotely piloted 3/8-scale F-15 airplane model was flight tested. Thi

6、s report documents the subsonic stability and control characteristics of this airplane model over an angle of attack range of -2OO to 53O. The remotely piloted technique for obtaining flight test data was found to provide adequate stability and control derivatives. The remotely piloted technique pro

7、vided an opportunity to test the aircraft mathematical model in an angle of attack regime not previously examined in flight test. The variation of most of the derivative estimates with angle of attack was found to be consistent, particularly when the data were supplemented by uncertainty levels. 17.

8、 Key Words (Suggested by Authork) 1 18. Distribution Statement F-15 airplane model Stability and control derivatives Unclassified - Unlimited Remotely piloted research vehicle I Category: 08 19. Security Classif. (of this report) 20. Security Classif. (of this page) Unclassified Unclassified *For sa

9、le by the National Technical Information Service, Springfield, Virginia 22161 I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SUBSONIC STABILITY AND CONTROL DERIVATIVES FOR AN UNPOWERED , REMOTELY PILOTED 3/8-SCALE F-15 AIRPLANE MODEL OBTAINED FROM

10、 FLIGHT TEST Kenneth W . Iliff, Richard E. Maine, and Mary F . Shafer Flight Research Center INTRODUCTION The increased concern with airplane characteristics at high angles of attack during stall, departure, and spin has motivated research in this angle of attack regime. There is a lack of complete

11、confidence in the ability of current design methods to predict airplane handling qualities at high angles of attack, so experi mental as well as analytical data are needed. The prediction of the handlingqualities of an airplane relies to a large extent on the prediction of its stability and control

12、characteristics. The proof of a new design must await flight tests, when the measured airplane stability and control characteristics can be compared with those estimated before flight test. The design cycle is reasonably well understood for low speeds and angles of attack for normal maneuvering, but

13、 the desire to utilize high angles of attack has expanded design envelopes beyond previously accepted designlimits. In response to the interest in stall, departure, and spin controllability, the NASA Flight Research Center is flight testing an unpowered remotely piloted 3/8-scale model of the F-15 a

14、irplane to high angles of attack. The remotely pilotedflight test technique (ref. 1) was chosen because of the risks involved in aircraft spin testing. The technique is versatile in that the pilot interacts with the vehicle as he does during normal flight, it is potentially more economical than full

15、-scale flight testing, and it allows the flight envelope to be expanded more rapidly than do conventional flight test methods. The derivative characteristics determined during the flight program were used both to verify the predicted airplane model aerodynam ics and to update a flight support simula

16、tor. This report documents the stability and control derivatives of the F-15 airplane model determined at subsonic speeds over an angle of attack range from -2OO to 53O. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SYMBOLS Cln ClY m m 0 N N trim n

17、 IX IXZ P 4 r t W a normal acceleration, g lateral acceleration, g rolling-moment coefficient pitching-moment coefficient pitching-moment coefficient for zero a and zero 6 e normal-force coefficient normal-force coefficient at trim for the center of gravity at 26-percent mean aerodynamic chord norma

18、l-force coefficient for zero a and zero 6 e yawing-moment coefficient side-force coefficient 2moment of inertia about the longitudinal axis, kg-m cross product of inertia, kg-m 2 2moment of inertia about the lateral axis, kg-m moment of inertia about the normal axis, kg-m 2 roll rate, deg/sec or rad

19、/sec pitch rate, deg/sec or rad/sec yaw rate, deg/sec or rad/sec time, sec weight, N angle of attack of the body axis, deg 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-aO P P a d e etrim r e angle of attack of the principal axis, deg angle of si

20、deslip, deg time derivative of angle of sideslip, radjsec aileron deflection , deg differential tail deflection, deg elevator deflection, deg elevator deflection at trim for the center of gravity at 26-percent mean aerodynamic chord rudder deflection , deg pitch angle, deg roll angle, deg partial de

21、rivative with respect to the subscripted variable DESCRIPTION OF THE VEHICLE AND INSTRUMENTATION The F-15 airplane is a conventional single-placed two-engined fighter with the wing leading edge swept back 45O and twin vertical tails. The model (figs. 1 and 2) and the full-scale F-15 aircraft have si

22、milar elevator , aileron, and rudder control surfaces for the stability augmentation system and for pilot control. The elevator is used for longitudinal control, and the rudder, wing aileron , and differential elevator are used for lateral-directional control. The model is unpowered, and the inlets

23、are blocked. Pertinent airplane model physical characteristics are presented in table 1. Details concerning the airplane model are given in reference 1. The implementation of the remote piloting and stability augmentation aspects of the vehicle are given in reference 2. Two model configurations were

24、 tested: the basic configuration and the so-called production configuration. The production configuration updated the basic config uration to that of the full-scale production airplane. This configuration was used during the last seven flights. The differences between the two versions of the model a

25、re shown in figure 3. The changes (shown dashed) involved rounding the wingtip trailing edges and removing the section of the horizontal stabilizer shown in the figure. The airplane model was flown with three center of gravity locations 3 6 Provided by IHSNot for ResaleNo reproduction or networking

26、permitted without license from IHS-,-,-TABLE 1. - THREE-EIGHTHS-SCALE MODEL CHARACTERISTICS Model -Length, m 7.15 Weight, N 10,964 Wing -Area, m 2 Span,m Aspect ratio Mean aerodynamic chord, m Leading-edge sweep, deg Taper ratio Dihedral, deg . Incidence, deg . Ailerons: Span, m . . . . Deflection,

27、deg . . Horizontal tail -2Planform (exposed), 7.94 4.89 3.0 1.82 45.0 0.25 -1.0 0 1.24 220 m . 1.57 Span, m 3.24 Aspect ratio 2.05 Taper ratio 0.34 Leading-edge sweep, deg 50.0 Mean aerodynamic chord (exposed), m . 0.94 Dihedral, deg Tail length, m Deflection, deg: Symmetrical Differential Vertical

28、tails . 0 . 2.30 . 15, -26 . +11 2Area (both sides), m . 1.64 Span,m 1.18 Leading-edge sweep, deg 36.6 Mean aerodynamic chord, m 0.77 Tail length, m . 2.02 Rudders -2Area (total), m . 0.26 Span, m 0.54 Mean aerodynamic chord, m 0.24 Maximum deflection, deg . ?30 26-percent, 30 .3-percent, and 38.5-p

29、ercent mean aerodynamic chord. The location of the center of gravity is indicated by the configurations designation, as shown in table 2. Table 2 also shows the configuration inertias. The airplane models instrumentation consisted of the standard package used for the measurement of stability and con

30、trol parameters, including three-axis angular rate gyros, attitude gyros, and linear accelerometers, along with control position sensors and boom-mounted angle of attack and angle of sideslip vanes. The data were filtered with 40-hertz passive analog filters. The data were then sampled with a 9-bit

31、pulse code modulation system and telemetered to a ground station in real time to be recorded. Before the flight data were analyzed, corrections for upwash and sidewash were made to the angle of attack and angle of sideslip measurements. The corrections to angle of attack and angle of sideslip for an

32、gular rates, as well as those for accelerometer position, were included in the digital program used for the derivative extraction . A complete description of the instrumentation system, includ ing the accuracy and resolution of each quantity measured, is given in reference 1. 4 Provided by IHSNot fo

33、r ResaleNo reproduction or networking permitted without license from IHS-,-,-TABLE 2. - INERTIA AND WEIGHT CHARACTERISTICS Configuration Center of gravity, percent mean IX IY, IZ IXZ. aerodynamic chord kg-m2 kg-m2 kg-m kg-m 2 Basic 1 26 373 2,579 3,021 16 Basic 2 30.3 373 2,451 2,893 3 Production 1

34、30.3 369 2,451 2,889 3 Production 2 38.5 369 2,402 2,839 0 I I I FLIGHT TEST PROCEDURE W, N 10,960 10,960 10,950 9,119 The 3/8-scale F-15 airplane model was air launched from a modified B-52 air plane at an altitude of approximately 15,000 meters at a Mach number of 0.65. After the launch, the pilot

35、 flew the aircraft remotely through a planned flight profile. In addition to maneuvers like stalls and spins, the pilot performed maneuvers for obtaining stability and control derivatives. These maneuvers were performed either by pilot commands through conventional cockpit controls or through an inp

36、ut pulse panel. The input pulse panel switch initiated programed control inputs once a desired flight condition was attained. These programed inputs allowed more maneuvers to be performed and permitted the pilot to concentrate on keeping flightconditions more nearly constant. MANEUVERS AND FLIGHT CO

37、NDITIONS The results presented in this report were obtained from 168 maneuvers that were performed during 12 of the first 16 flights of the airplane model. The remotely piloted research vehicle technique made maneuvers from which stability and control derivatives could be extracted possible over an

38、angle of attack range of -2OO to 53O, a range never before investigated in flight tests. The maneuvers were made at Mach numbers below 0.60. The maneuvers were performed for small perturbation analysis about the desired steady state flight condition, where linearity of the air plane model could be a

39、ssumed. The maneuvers were initiated with inputs in the longitudinal or lateral-directional mode and analyzed for that mode. Most of the data were obtained without the stability augmentation system engaged in the mode to be analyzed; however, the stability augmentation system was engaged in the othe

40、r mode when the vehicle was difficult to stabilize. Stabilized flight conditions were difficult to maintain at extreme angles of attack. A consequence of the models lack of power was that the high angle of attack maneuvers were of short duration. The shortness of the maneuvers sometimes reduced the

41、accuracy of the estimates of the stability and control derivatives, which were defined by the mathematical model described in reference 3. Coupling motions, present in some high angle of attack maneuvers, were accounted for as described in 5 Provided by IHSNot for ResaleNo reproduction or networking

42、 permitted without license from IHS-,-,-ll1l11ll I Ill1 I I reference 4 to make it possible to determine stability and control derivatives adequately. Most of the maneuvers that did not result in satisfactory matches were made at high angles of attack. There was some aerodynamic flow separation abov

43、e an angle of attack of 15O, and the separation was quite extensive above an angle of attack of 25O. The results presented for angles of attack greater than 30 are the best avail able but may be affected by nonlinearities . METHOD OF ANALYSIS A maximum likelihood estimation method of analysis was us

44、ed to determine the stability and control derivatives from the maneuvers made in flight. The method used (sometimes called the Newton-Raphson method) is an iterative technique that minimizes the difference between the aircrafts measured and computed response by adjusting the stability and control de

45、rivative values used in calculating the computed response. The Newton-Raphson method is used to attain the minimizations. The maximum likelihood estimation method can be modified to include a priori information from previous calculations , flight tests , or wind tunnel tests. The modification is mad

46、e by including a penalty for adjusting the unknown stability and control derivatives away from the a priori values. Therefore , if new information is contained in a flight maneuver , the estimate of the derivative is affected only slightlyby the a priori information. If no new information is contain

47、ed in a flight maneuver , the a priori value results. This method is called modified maximum likelihood estimation and is fully described in reference 3. In this flight test program, a priori information was included in the analysis. A complete description and FORTRAN listings of the digital program

48、 used for the derivative extraction are given in reference 5. In addition to providing estimates of the derivatives, this method of analysis provides uncertainty levels associated with each derivative. Uncertainty levels are proportional to the approximation of the Cram and Deets , Dwain A. : Develo

49、pment of a Remote Digital Augmentation System and Application to a Remotely Piloted Research Vehicle. NASA TN D-7941, 1975. 3. Iliff , Kenneth W .; and Taylor , Lawrence W ., Jr .: Determination of Stability Derivatives From Flight Data Using a Newton-Raphson Minimization Technique. NASA TN D-6579, 1972. 4. Iliff , Kenneth W .; and Maine, Richard E. : Practical Aspects of