REG NASA-CR-176615-1985 Exploratory low-speed wind-tunnel study of concepts designed to improve aircraft stability and control at high angles of attack.pdf

1、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EXPLORATORY LOW-SPEED WIND-TUNNEL STUDY OF CONCEPTS DESIGNED TO IMPROVE AIRCRAFT STABILITY AND COEiTROL AT HIGH ANGLES OF ATTACK BY David Edgar Hahne B.S June 1981, Vi rginia Polytechnic Institute and S

2、tate University A Thesis submitted to the Faculty of The Gradueaie School of Engineering and Applied Sciowe of The George Washington University in partial satisfaction of the requirements for the degree of Master of Science December 1985 Provided by IHSNot for ResaleNo reproduction or networking per

3、mitted without license from IHS-,-,-ABSTRACT A wind-tunnel investigation of concepts to improve the high angle-of-attack stability and control characteristics of a high performance ai rcrsft has been conducted. The effect of vertical tail geometry on stability and the effcc- tiveness of several conv

4、entional and unusual control concepts has been deter- mined. found that vertical tail location, cdnt angle and leading edge sweep could i nfl uence both 1 ongi tudi nal and 1 ateral -di rectional stabi 1 i ty . cepts tested were found to be effective and to provide control into the post- stall angle

5、-of-attack region. These results were obtained over a large angle-of-attack range. It was The control con- ii Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-ACKNOWLEDGEMENTS The author wishes to thank the National Aeronautics and Space Administratio

6、n for providing the opportunity to perform this research, prepare this thesis, and complete all the other requirements for this degree. The author is grateful to Professor S rolling-moment coefficient, Rolling moment/qSb pi tching-moment coefficient, Pitching moment/qSc yawing-moment coefficient, Ya

7、wing mornent/qSb side- force coefficient, Side force/qS mean aerodynamic chord, ft mass moments of inertia about xb, Yb, zb body axes, sIugs-ft2 mach number ai rcraf t mass, slugs angular veloctiy about the Xg, Yb, and z;, body axes, deglsec roll rate about the velocity vector, ! linear velocities a

8、long the xb, Yb, zb body axes, ft/sec freestream velocity, ft/sec body axes span ordinate, ft nondirnensiortal semi span lwation angle cf attack, deg angle of sideslip, deg viii Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Q P incremental roll ing

9、 lnornent coefficient incremental yawing momcnt coefficient Incremental side force coefficient aileron deflection, deg centerline vertical tail deflection, deg canard def 1 ect j on, deg trailing edg? extension flap deflection, deg leading edge flap deflection, deg horizontal tail deflection, deg ru

10、dder deflection, deg aft swept tail rudder deflection, deg forward swept ta5l rudder deflection, deg tip rudder deflection, deg wing tip deflecton, deg root of cha-acterfstic equation afrcraft roll attitude, deg a5r density, slugs/ft3 Stability derivatives 2v 2V 2v 2v ix Provided by IHSNot for Resal

11、eNo reproduction or networking permitted without license from IHS-,-,-Subscripts IS inboard M6 mid board 06 outboard X Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-CHAPTER 1 I NTROBUCTION The emphasis an sustained nonafterburning supersonic cruise

12、 makes aircraft with high-f ineness-ratio fuselages, OH-aspect-ratio, highly-swept wings and highly-integrated control surfaces attractive because of low cruise drag (Reference 1). These sane features, however, make the aircraft more susceptible to the development and sudden breakdovrn of strong vor

13、tical flowfields at angles of attack and speeds encountered during maneuvering. above 15“ angle of attack where the vortical flowfields can produce ronlineari- tier and regions of unstable behavior in the longitudinal, lateral ana direc- tional aerodynamics. In addition strong spanwise flowfields ar

14、ise which reduce This is especially true the effectiveness of conventional control devices. Such aerodynamic charac- teristics complicate control system design especially for unstable aircraft where powerful control surfaces are the key to a sirccessfcrl active flight control system. As illustratzd

15、conceptually in figure 1, maneuverability aecreases with increasing angle of attack such that very little capability is available at and above the stall. tibility to loss of control and spins. On some configurations, limiters are imposed to avoid suscep Providing controls that maintain a high level

16、of effectiveness will allow future high performance aircraft to exploit a much larger angle-of-attack envelope, including brief excursions into post stall flight conditions. A major potential payoff is configuration optimization flexjbility which results frov the fact that the hish levels of conzrol

17、 effec- tiveness combined wi th appropriate control 1 dws and fly-by-w i re technology WS 11 allow increased reliance on st and (2) to stiidy the impact that these advances will have on aircrdft maneuverability and the assc- (11 to develop aerodynamic control devices that ci ated fl i ght dynamics.

18、The generic configuration of thi s study has emerged front a nUmb%!i- of studies aimed at developing design yuidelincs far providing efficient supersonic cruise, transonjc rcancvzjering capability cowparable to current high performance aircraft ad gmd iw-speedhigh angle-ot-attack stabi- 1 ity and co

19、ntrol (Reference 3). This thesis presents the results of the latest phase of tnis group of studies to develop and evaluate conventional an4 unusual control concepts for advanced aircraft. This work. includes an eva;uatian of the effects of tail geometry on stability as well as an evalrratior: of con

20、trol effec- tiveness. Compari sons of manebgerabi 1 i ty Hi th a hi ghly-maneuverabl e ai rcraf t are also presented. The test program was conducted in the Langley Research Centers IP-Foot Low-Speed Wind Tunnet using a strut supporreci scale mode? of tie generic COR- figuration. Data viers taken cwr

21、 a large angle-of-attack range at several si des1 5 p angi es . -2- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-CHAPTER 2 MODEL AND APPARATUS The basic model and its support system are shown in Figure 3. Details of the model forebody, vertical ta

22、ils, control surfaces, and horizontal tails and canards are shown in figures 4, 5, 6, and 7, respectively. 65“ swept leading edge with an aspect ratio of 1.95. canards all had thin flat-plate cross-sections (for simplicity) with beveled leading and trailing edgzs. The wing leading edge was divided i

23、nto three flap segments and the trailing edge had two different flap/aileron configurations along with flaps on the inboard trailing edge extension. The arro: wing had a The wings, tails and 1 It was also possible to deflect the wing tips as ailerons. and could be replaced by horizontal tcsils. to u

24、se canards. The trailing edge extension was removable The fuselage could also be adapted The rnadel could be configured with three di Fferent vertlcal tail arrangements: edge extensions) and a single centerline tail. The vertical tails also had several fore and aft positions as well. as the abi!ity

25、to cant either inboard or outboard. The fuselage, with flow through ducts, was strut mounted to minimize unfaverable effects on vortex bursting as described in Reference 4. forward and aft svrept Win tails (mounted on the traiiing F3rce and moment data were obtained using a six component strain-gage

26、 balance mwnted inside the model fuselage. The voltage outputs of the strain-gage balance were converted to digital signals by a MEFF 620 amp1 i fier/mul tiplexor. The NEFF 620 providea the required signal processing for use Mith the HP-98458 microco.+pttter tcsed fer data reductiors and storage. Th

27、e data acquisition system r,amp! 351. Even though Cn5 is neW- tive or very small for this configuration between a = 20“ and 46O, the stable values Of Ct3 from a = 0“ to GO“ shculd be cnotigh to preveot a n05e slice frm occurring. - 16 - Provided by IHSNot for ResaleNo reproduction or networking perm

28、itted without license from IHS-,-,-Control Reversal .- When adverse yawing moment due to aileron deflection becomes too high, too much sideslip is generated and control response to a pilot roll input will be reversed. This control reversal is predicted when LCDP (as defined in Chapter 3) becomes neg

29、atide. FGr. ailerons only, this con- figuration will experience a reversed response to roll inputs past a = 20“. In order to prevent chis, it is necessary to use rudder deflection to offset the adverse yaw dlre to aileron deflection. pilot or by using an aileron-rudder interconnect (ARI) system. The

30、 ARI system This can be accomplished manually by the provides a prJportiona1 amount of yaw control for a given roll inplrt in order to ensure proper control response. With the proper ARI gain (K = 0.351, correct response to pilot roll commands can be maintained until a = 45“ (figure 33). Maneuverabi

31、l i 9. The maneuverability of an aircraft is generally measured by how fast it can turn about its three axes. This requires a complete data base, such as that used in sirnulation str;dies, and knokledgi? of restrictions that could limft the maximum allowable rates and accelerations. A less exact met

32、hod is to use the equations presented in Chapter 4 to calculate the maximum rate accelerations that can be commanded using a constant velocity and altitude. This gives an indication of how fast an aircraft can initiate a turn. Figdre 34 presents the rate accelerations that can be commanded for a cti

33、rrent highly maneuverable aircraft (Aircraft A) and the configuration of this study (Aircraft B) . The pitch rate acceleration data of figure 34(a) show that belovr the stall (both aircraft stall at a = 35“) Aircraft A can initiate 2 pftch attitude change faster than Aircraft B. Aircraft B, hcwever,

34、 is capable of higher pitch rate acceleration beyond the stall. Also, Aircraft B shows little change in pitch rate acceleration with changing angle of attack. This will provide the pilot P - 17 - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-with v

35、irtually the same longitudinal respsnse to his pitch control inputs throughout this angle-of- attack range. The results of the yaw rate acceleration calculations are presented in figure 34(b). The data indicate that Aircraft A is capable of generating higher yaH rate accelerations than Aircraft 5 th

36、rcughout the angle-of-attack range tested. A consequence of lower attainable yaw rate acceleration for Aircraft 8 will be lower effectiveness of a stability augmentation system &AS) in the yaw axis. Exactly how important this is depends upon the basic airframe static and dynamic stability characteri

37、stics. Figure 34(c) shows the roll rate acceleration capabilities of the two aircraft. Aircraft A is capable of initiating a roll rate quicker than Aircraft B below the stnll whereas the opposite is true above the stail. the region approaching the stall and above for aircraft vlith highly swept wing

38、s to exhibit plring rock. tendencies. It is comrmn in The higher 1-37 rats acceleration capabii ity of Aircraft 5 in this angle-of-attack range will allow more effective use of artificial stabilization to prevent this ding rock. High performance aircraft, vthich are typically fuselage heavy, tend to

39、 pitch p when rolling about the aircraft velocity vector (stability axes) because of inertia coupling. This places a limitation on the maximum stability axes roll rate that can be maintained. This maximum roll rate, as shown in equation 15, is related to the amount of nosedown longitudinal control p

40、ower avail able to overcome the pi tchu p caused by inertia coup1 ing. description of itiertia coupling can be found in Reference 8. the results of this evaluation for the test configuration and current aircraft. aelow the stall, Aircraft A is limiteO less by inertia coupling than Aircraft 3, allowi

41、ng Aircraft 4 to make quicker rolls about the velocity vector. A detailed Figure 35 shows In the - 18 - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-post-stall region, however, Aircraft B is able to sustain a higher roll rate that is almost constant throughout the remainder of the angle-of- attack range. - 19 - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-