NASA-TM-X-3078-1974 Subsonic and supersonic longitudinal stability and control characteristics of an aft tail fighter configuration with cambered and uncambered wings and uncambere.pdf

1、NASA TECHNICALMEMORANDUMNASA TM X-3078COSUBSONIC AND SUPERSONIC LONGITUDINALSTABILITY AND CONTROL CHARACTERISTICSOF AN AFT TAIL FIGHTER CONFIGURATIONWITH CAMBERED AND UNCAMBERED WINGSAND UNCAMBERED FUSELAGEby Samuel M, DollyhighLangley Research CenterHampton, Va. 23665NATIONAL AERONAUTICS AND SPACE

2、ADMINISTRATION WASHINGTON, D. C. AUGUST 1974Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1, Report No. 2. Government Accession No.NASA TMX- 30784. Title and SubtitleSUBSONIC AND SUPERSONIC LONGITUDINAL STABILITY ANDCONTROL CHARACTERISTICS OF AN AF

3、T TAIL FIGHTERCONFIGURATION WITH CAMBERED AND UNCAMBEREDWINGS AND UNCAMBERED FUSELAGE7. Author(s)Samuel M. Dollyhigh9. Performing Organization Name and AddressNASA Langley Research CenterHampton, Va. 2366512. Sponsoring Agency Name and AddressNational Aeronautics and Space AdministrationWashington,

4、D.C. 205463. Recipients Catalog No.5. Report DateAugust 19746. Performing Organization Code8. Performing Organization Report No.L-946310. Work Unit No.760-67-01-0411. Contract or Grant No.13. Type of Report and Period CoveredTechnical Memorandum14. Sponsoring Agency Code15. Supplementary Notes16. Ab

5、stractAn investigation has been made in the Mach number range from 0.20 to 2.16 to determinethe longitudinal aerodynamic characteristics of a fighter airplane concept. The configurationconcept employs a single fixed geometry inlet, a 50 leading-edge-angle clipped-arrow wing, asingle large vertical t

6、ail, and low horizontal tails. The wing camber surface was optimized indrag due to lift and was designed to be self-trimming at Mach 1.40 and at a lift coefficientof 0.20. An uncambered or flat wing of the same planform and thickness ratio was also tested.However, for the present investigation, the

7、fuselage was not cambered. Further tests shouldbe made on a cambered fuselage version, which attempts to preserve the optimum wing loadingon that part of the theoretical wing enclosed by the fuselage.17. Key Words (Suggested by Author(s)Wind-tunnel testsFighter configurationsAerodynamic characterist

8、ics19. Security Qassif. (of this report)Unclassified18. Distribution StatementUnclassified - UnlimitedSTAR Category 0120. Security Classif. (of this page) 21. No. of PagesUnclassified 9622. Price*$4.00For sale by the National Technical Information Service, Springfield, Virginia 22151Provided by IHSN

9、ot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SUBSONIC AND SUPERSONIC LONGITUDINAL STABILITY AND CONTROLCHARACTERISTICS OF AN AFT TAIL FIGHTER CONFIGURATIONWITH CAMBERED AND UNCAMBERED WINGS .AND UNCAMBERED FUSELAGEBy Samuel M. DollyhighLangley Research CenterSUMM

10、ARYAn investigation has been made in the Mach number range from 0.20 to 2.16 todetermine the longitudinal aerodynamic characteristics of a fighter airplane concept.The configuration concept employs a single engine fed by a single fixed geometry inlet, a50 leading-edge-angle clipped-arrow wing, a sin

11、gle large vertical tail, and low horizontaltails. The wing camber surface was optimized in drag due to lift and was designed to be .self-trimming at Mach 1.40 and at a lift coefficient of 0.20. An :uncambered or flat wingof the same planform and thickness ratio was also tested. However, for the pres

12、entinvestigation, the fuselage was not cambered. Further tests should be made on a cam-bered fuselage version, which attempts to preserve the optimum wing loading on that partof the theoretical wing enclosed by the fuselage.The results indicate that the configuration possessed reasonably linear pitc

13、hing-moment characteristics over the test Mach and angle-of-attack ranges, except at Mach0.50 where the configuration pitched down when the wing airflow separated at angles ofattack above 20. The horizontal-tail control effectiveness was found to be adequate overthe test Mach range. The configuratio

14、n with the supersonic cambered wing had drag polarcharacteristics at the higher angles of attack superior to those for the configuration withthe flat wing at all Mach numbers of the test. However, the positive zero-lift pitchingmoment was absent; this would have enabled the cambered wing configurati

15、on to trim athigh lift coefficients with relatively small or no horizontal-tail loads, and thus lower thetrim drag. It was speculated that the absence of significant positive, zero-lift pitchingmoment in the cambered wing configuration was due to the fuselages lack of being cam-bered in such a way t

16、hat the theoretical wing loading was preserved. Trimmed drag dif-ferences between the configuration with the two wings at Mach 1.47, 1.80, and 2.16 werefairly accurately predicted by current supersonic theoretical methods.Provided by IHSNot for ResaleNo reproduction or networking permitted without l

17、icense from IHS-,-,-INTRODUCTIONAs part of a research program on advanced fighter aircraft technology, the NationalAeronautics and Space Administration has undertaken research related to highly maneu-verable fighter aircraft. This report presents the results of wind-tunnel tests of the firstin a ser

18、ies of generalized fighter configurations of research models of an aft tail fighterconcept.The configuration concept is as tightly packaged as possible to keep cross-sectionalarea low. It employs a single engine fed by a single fixed geometry inlet, and the cock-pit features an inclined pilot seat.

19、As a result, the cross-sectional area at the pilot sta-tion is greatly reduced, and the pilot is able to withstand higher sustained g loads. Thewing planform is a clipped arrow with a 50 leading-edge sweep. The wing camber sur-face is designed for minimum drag due to lift and to be self-trimming at

20、Mach number1.40 and at a lift coefficient CL of 0.20 by the method discussed in reference 1. Ideally,designing the wing this way should result in a low drag penalty associated with trimmingthe aircraft by keeping the necessary horizontal-tail deflections or horizontal-tail loadssmall. No attempt was

21、 made to camber the fuselage in order to preserve the wing load-ing on the part of the theoretical wing that was enclosed by the fuselage. A second wingof the same planform and thickness distribution, but with a flat camber surface, wasincluded in the investigation as a reference.Wind-tunnel tests o

22、n a 0.056-scale model were conducted in the Langley 8-foottransonic pressure and Unitary Plan wind tunnels at Mach numbers from 0.2 to 2.16. Theresults of the wind-tunnel investigation along with some supersonic analytical results arereported herein.SYMBOLSThe force and moment coefficients are refer

23、enced to the stability axis system. Themoment reference point was located at fuselage station 39.40 cm (0.40 c) for the wingapex located at 20.353 cm and at fuselage station 40.61 cm (0.30 c) for the wing apexlocated at 23.52 cm.A aspect ratiob wing span, cmCQ drag coefficient, Drag/qSCD c chamber-d

24、rag coefficient, Chamber drag/qSProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-CD i internal-drag coefficient, Internal drag/qSCD 0 drag coefficient at zero liftCL lift coefficient, Lift/qSCL lift curve slope at CL = 0Cm pitching-moment coefficient,

25、 Pitching moment/qScCm o pitching-moment coefficient at zero lift9Cm/9Cr longitudinal stability parameter at CL = 0AC/CL drag-due-to-lift parameter (determined at CL = 0.5)6h tail control effectiveness at zero moment, per degree9Cm/96n pitching effectiveness of horizontal tail at CL = 0c streamwise

26、chord, cmc wing mean geometric chord, cmL/D lift-drag ratioM free-stream Mach numberq free-stream dynamic pressure, N/m2S reference area of wing including fuselage intercept, cm2x longitudinal distance from leading edge of wingy lateral distance from center line of airplanez vertical ordinate of cam

27、ber surface, positive upor angle of attack, degreesProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-r dihedral angle, degrees5, horizontal-tail deflection angle, positive when trailing edge is down, degreesA leading-edge sweep angle, degreesSubscripts

28、:max maximumtrim trimmedDESCRIPTION OF MODELA three-view drawing of the complete model is shown in figure l(a), and drawingsof the wing, vertical tail, and horizontal tail are shown in figures l(b) to l(d). Some geo-metric characteristics are given in table I, and a photograph of the model is presen

29、ted infigure 2. The configuration incorporates an uncambered fuselage with a single externalcompression horizontal-ramp inlet, a clipped-arrow wing, twin horizontal tails, and asingle vertical tail.The wing planform was a clipped arrow with a 50 leading-edge sweep. The taperratio of the theoretical

30、planform was 0.20, and the notch ratio was 0.157. The stream-wise airfoil thickness distribution was a NACA 65A004.5. Two wings were tested, eachhaving the same planform and airfoil thickness distribution but differing in camber sur-face. The first wing had a camber surface that was designed for min

31、imum drag due tolift at Mach number 1.4 and CL - 0.2. The camber surface was also designed so thatthe wing would be self-trimming about the center of gravity of the configuration at thedesign point (M = 1.4; CL = 0.2) with the wing apex located at model station 20.353. Thecamber surface ordinates of

32、 this wing with respect to the leading edge are given in table II.The wing is hereafter referred to as the cambered wing. The second wing was uncam-bered and untwisted (flat) and is hereafter referred to as the uncambered wing. Bothwings could be moved rearward 3.167 cm.The configuration employed lo

33、w twin horizontal tails with a 4 percent biconvex sec-tion. The horizontal tail could be deflected over a range of from -13;33 to 10 and couldbe removed from the model. The relatively large single vertical tail also had a 4 percentbiconvex airfoil section.Provided by IHSNot for ResaleNo reproduction

34、 or networking permitted without license from IHS-,-,-TESTS AND CORRECTIONSThe tests were conducted in the Langley 8-foot transonic pressure and Unitary Planwind tunnels. The conditions under which the tests were conducted were as follows:Mach number0.2.5.8.85.90.951.031.21.471.802.16Stagnation pres

35、sure,kN/m257.4657.4657.4657.4657.4657.4657.4657.4666.03 and 39.6073.07 and 43.8685.61 and 52.38Stagnation temperature,K316320321322323323323323339339339Reynolds numberper meter2.305.187.057.227.387.48 7.687.818.20 and 4.928.20 and 4.928.20 and 4.92At Mach 1.47, 1.80, and 2.16, the Reynolds number pe

36、r meter, as indicated by the lowervalue in the table, was reduced at angles of attack above 10 in order to stay within thebalance load limits. The dewpoint was maintained sufficiently low to prevent measurablecondensation effects in the test section. .The angle-of-attack range was from approxi-matel

37、y -6 to 20. In order to insure boundary-layer transition to turbulent flow atMach 0.2 to 1.2, 0.16-cm-wide transition strips of No. 60 grit were placed on the body3.05 cm aft of the nose of the model, and strips of No. 80 grit were placed 1.02 cm stream-wise on the wings, tails, inlet ramps, and ext

38、ernal inlet surface. At Mach 1.47 to 2.16,strips of No. 50 grit were used to replace the strips of smaller grit used at the lower Machnumbers. These transition strips are shown to be adequate in reference 2.Aerodynamic forces and moments on the model were measured by means of a six-component strain-

39、gage balance which was housed within the model. The balance wasattached to a sting which, in turn, was rigidly fastened to the tunnel support system.Balance-chamber static pressures were measured with pressure tubes located in thevicinity of the balance The model internal-flow total and static press

40、ures were mea-sured with a rake consisting of 29 total-pressure tubes and 5 static-pressure tubes. Therake was placed flush with the base of the model and was removed during the force-measurement tests. The drag data presented herein have been corrected for internaldrag and have also been corrected

41、to the condition of free-stream static pressure in thebalance chamber. Figures 3 and 4 show values of the balance chamber and internal dragProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-coefficients which were used to correct the drag-data. Correcti

42、ons to the angles of attaiof the model have been made for both tunnel-airflow misalinement and deflection of thebalance and sting under load.PRESENTATION OF RESULTSFigur-l-tuHaaoortTSI0)0)JHCO!H0)*-0)aiCDOcl-lCDCQcoiHCO, W V*JO a,3 a0)gaCDOCOa*piti0)a13Provided by IHSNot for ResaleNo reproduction or

43、 networking permitted without license from IHS-,-,-oIMOOasCOQOO sa, fa14Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- c ouat-,VT3oCOIc05SpIa0)OuoUrt (UI I0)Oen16Provided by IHSNot for ResaleNo reproduction or networking permitted without license f

44、rom IHS-,-,-17Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-CD,(0-.001-.002.0010-.001iFMteafi11 rll “*i ,t itlit 1n it_;iiMlsis5 AtJSE“ u:ay -.;j8i i_iit,_4?$tdrJ“i-trti!11-,5$jfr!8:;Ejiji ji,1 iirrH-ImTlHI*:i!iir1ij1-mw*s*rr!1 “l!-r1Sitt=rSBc-,*io

45、-?*44=Hit*i. t.3t!,o-“-?tTI jsSiii*“|nttiwIff:mt_i:-u“f/-*“t.:rf-rr1in:niti*-J“ ir1aoEni-i/fJ:l1/Jl1Usi1“it-tiln1-r1K!tintuti-tii.te-Jf.ite-.,1tr-4-.t- 0-.001 cD,c-.002.002.001CD.C0-.0018-4 0 4 8 12 16 20 24 28 32a.degFigure 3.- Chamber drag correction.18Provided by IHSNot for ResaleNo reproduction

46、or networking permitted without license from IHS-,-,-.008.007CD, c .006. .005.0040-.001cD,c -.002_ nr(4 “*riv1y“*-t/ JVXJyu1XLfA“ 5Ssf:H:-loPir-_-J.Y*4-_y-1J:Jypr-C-c.1_j.r“-J.f“tM-l.20T , k (.M = l.03fJ. 1;M=0.95o f,V OM=0.90-bs.kT_-r-h._rx1Xll-iE2=,/|/_-_Jrif11f*r_r“,/-_JP-_-J-t“r:.*5EMl_ =-J-l-“4

47、HT-1-tTTr5STt11-_-*-5.,SSpr“-_t._“!:,“-1wSTITtCC1,1_“_-tTTTT5S-=HsT:i-_“1_-_-1j.iiE-=M?mrslHE=“;-,-“_=iM-2:1*=3tptI3rTiH-;1-_-1AAJO.007.006.0050-.001-.002-.003CD,-8 -4 . 0“D,C48 12 16 20 24 28 32a, degFigure 3.- Continued.19Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.003.002CD;C.0010.004.003cD,c .002.001n33ltd3HBV .“Hr2=_-_.-,_f.-3=.-f:-“tl-_aag.E?Ja“_=-r-*BE_r-_,r-“-*kfaa,-a“1j-f*lM“atfirTt-1-1-_HFn.-rii-i.tftt?aa1 1.=:-1o-rElt.“-*,i-a-“at-“W“r“K.“ t=MH3M =2.16-( v1M = 1.80L. “S_ “M = l.47itf“-(.STf“t=j-,-S,i- ,!p|fTaaM