1、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NASATechnicalPaper24781985NASANational Aeronauticsand Space AdministrationScientific and TechnicalInformation BranchEffect of Aileron Deflections onthe Aerodynamic Characteristicsof a Semispan Model of
2、a SubsonicEnergy-Efficient TransportPeter F. JacobsLangley Research CenterHampton, VirginiaProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SUMMARYAn investigation was conducted in the Langley 8-Foot Transonic Pressure Tunnelto determine the effect of
3、 aileron deflections on the aerodynamic characteristics ofa subsonic energy-efficient transport (EET) model. The semispan model had anaspect-ratio-10 supercritical wing and was configured with a conventionally locatedset of ailerons (i.e., a high-speed aileron located inboard and a low-speed aileron
4、located outboard). Data for the model were taken over a Mach number (M) range from0.30 to 0.90 and an angle of attack range from approximately -2 to 10 . TheReynolds number was 2.5 x 106 per foot for M = 0.30 and 4.0 x 106 per foot for theother Mach numbers. Model force and moment d_ta, aileron-effe
5、ctiveness parameters,aileron hinge-moment data, chordwise pressure distributions, and spanwise load dataare presented.The data indicate positive aileron effectiveness for the inboard aileron (basedon averaged, equal-magnitude, positi%_ and negative deflections) for all test condi-tions except for de
6、flections of _2.5 at a Mach number of 0.90 and angles of attackfrom approximately 0 to 2 Control reversal at these conditions may be caused byshock-induced flow separation on the wing and interference effects from the nacelleand pylon. The outboard aileron had positive aileron effectiveness at all t
7、estconditions. The effectiveness of both ailerons increased near M = 0.82, mainlybecause of the influence of negative aileron deflections on shock location and thechordwise extent of the upper surface pressure plateau. The effectiveness of theoutboard aileron did not increase as much as that of the
8、inboard aileron, since thewing loads were lower at the tip. The effectiveness of both ailerons decreased atabout M = 0.84 because of shock-induced boundary layer separation over much of thewing. The extensive aft camber of this supercritical wing produced negative hingemoments for most deflections o
9、f the ailerons.INTRODUCTIONAs part of the National Aeronautics and Space Administrations Aircraft EnergyEfficiency (ACEE) project, extensive theoretical studies and experimental wind tunnelinvestigations have produced a group of aerodynamically efficient jet transportwings. These wings have higher l
10、ift-drag ratios, thicker airfoil sections, lesssweep, and higher aspect ratios than the wings on current wide-body jet transports.The performance characteristics of these configurations have been documented inreferences I to 3, and aileron-effectiveness data for a preliminary active-controlconfigura
11、tion are presented in references 4 and 5.Further tests of lateral controls on an energy-efficient transport configurationwere undertaken for two reasons. First, data for a more conventionally sized andlocated set of ailerons than those used in references 4 and 5 were desired. Second,since aileron ef
12、fectiveness is sensitive to Reynolds number, a model larger than theone used in references 4 and 5 was required to increase the test Reynolds number.Therefore, a semispan model twice the scale of one of the configurations (SCW-2C) inreference I was constructed for the present investigation. The mode
13、l had a conven-tional set of ailerons (i.e., a high-speed aileron located inboard and a low-speedaileron located outboard). Model force and moment data, aileron-effectivenessProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-parameters, aileron hinge-mo
14、ment data, chordwise pressure distributions, and spanwiseload data are presented in this report,SYMBOLSThe longitudinal and lateral-directional aerodynamic characteristics presentedin this report are referred to the stability and body axis systems, respectively.Force and moment data ha%_ been reduce
15、d to conventional coefficient form based on thetrapezoidal planform of the semispan wing panel (extended to the fuselage center-line). All dimensional values are given in the U.S. Customary Units_a I,a 2b/2CDCHMIailerons I and 2 (fig. 3)wing semispan, 53.17 in.drag coefficient, Dragqshinge-moment co
16、efficient for aileron I, Hinge momentq_SalCa ICHM2CLCLuhinge-moment coefficient for aileron 2,Liftlift coefficient, qslift-curve slope, per degreeHinge momentq=Sa2Ca 2CIRolling momentrolling-moment coefficient, q Sb/2C16aC mlateral aileron-effectiveness parameter,C_ 16a - C 10 o, per degree6a - 0pit
17、ching-moment coefficient,Pitching momentq ScC nYawing momentyawing-moment coefficient, q Sb/2C !nsection normal-force coefficient obtained from integration of pressuremeasurementsCn6 aCpCnl6a - Cnl0Odirectional aileron-effectiveness parameter, 6a 0degreep - ppressure coefficient, q., perlocal chord,
18、 in.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-CaICa 2Ppq_RigSSa 1Sa 2txYz5al,6aaverage chord of al, 3.28 in.average chord of a2, 1.87 in.mean geometric chord of reference wing panel, 11.48 in.free-stream Mach numberlocal static pressure, lb/ft
19、2free-stream static pressure, ib/ft 2free-stream dynamic pressure, ib/ft 2Reynolds number, per footwing planform reference (trapezoidal) area, 3.992 ft 2planform area of al, 0.141 ft 2planform area of a2, 0.124 ft 2local maximum thickness of wing, in.chordwise distance from wing leading edge, positi
20、ve aft, in.spanwise distance from model centerline, in.rvertical distance, positive up, in.angle of attack, degdeflection of ailerons I and 2, positive for trailing edge down, deglocal wing incidence measured from fuselage waterline, positive forleading edge up, degsemispan station, y/(b/2)Abbreviat
21、ions:F,S.L.E.L.S.T.E.U.S.fuselage station, in.leading edgelower surfacetrailing edgeupper surfaceProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EXPERIMENTALAPPARATUSANDPROCEDURESTest FacilityThis investigation was conducted in the Langley 8-Foot Tra
22、nsonic Pressure Tunnel(ref. 6). This facility is a continuous-flow, single-return tunnel with a slottedtest section. Tunnel controls allow independent variation of Machnumber, density,stagnation temperature, and dew-point temperature. The test section is approximately7.1 ft square (samecross-section
23、al area as that of a circle with an 8.0-ft diam-eter). The ceiling and floor are slotted axially and have an average openness ratioof 0.06. These features permit the test section Mach number to be changed continu-ously throughout the transonic speed range. The stagnation _ressure in the tunnelcan be
24、 varied from a minimum of 0.25 atm (I atm = 2116 Ib/ft _) at all Mach numbers toa maximum of approximately 2.00 atm at Mach numbers less than 0.40. At transonicMach numbers, the maximum stagnation pressure that can be obtained is approximately1.5 atm.Model DescriptionA photograph of the model in the
25、 L_ngley 8-Foot Transonic Pressure Tunnel isshown in figure I. Drawings of the model are shown in figures 2 and 3.Fuselage.- The fuselage for the semispan model used in this investigation wasmetric, had a maximum radius of 5.75 in., and was 108.61 in. long. Typical fuselagecross sections are shown i
26、n figure 4. The fineness ratio of this fuselage (9.44) isslightly higher than typical second-generation (wide-body) jet transports. The lowersurface of the wing was faired into the fuselage to produce a relatively flat bottomthat extended approximately 7 in. both ahead of the wing leading edge and a
27、ft of thewing trailing edge. The center section of the fuselage consisted of two fiberglasscover plates that surrounded the wing. The parting line for the cover platesextended horizontally fore of the leading edge and aft of the trailing edge of thewing root. A rubber seal between the wing and fusel
28、age prevented airflow through thejuncture. In order to prevent fouling, the fuselage was mounted approximately0.2 in. away from the tunnel wall. A splitter plate was not used.Wing.- A modified version of wing configuration SCW-2C from reference I was usedin this investigation. The wing had 5 of dihe
29、dral and 30 of sweep at the quarter-chord. Based on the trapezoidal planform (extended to the fuselage centerline), thewing had a reference area (semispan) of 3.992 ft 2, an aspect ratio of 9.8, and ataper ratio of 0.397. The wing was designed for a cruise lift coefficient of 0.55 ata Mach number of
30、 0.82.The planform and airfoils of the original configuration were modified (ref. 2)from the wing root outboard to the trailing-edge break. Trailing-edge and leading-edge extensions were added to allow a substantial increase in the thickness of thewing root where the rear spar and landing-gear attac
31、hment would be located (seefig. 5). The wing trailing edge near the original planform break was also modifiedto correct local upper surface flow separation. To reduce the curvature of the wingupper surface at the trailing edge, the chord was shortened slightly. The planformof the trailing edge was a
32、ltered from a simple break to an elliptical fairing of theoutboard geometry into the new inboard trailing-edge extension. Twist and thicknessdistributions for the original and modified wings are shown in figure 6. Note thatthe twist remained the same for both wings, and although t/c for the modified
33、 wingwas reduced inboard, the local wing thickness was greater than that of the original4Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-wing because of the increased chords inboard for the modified wing. Several airfoilsections from the modified win
34、g are shownin figure 7 and measuredcoordinates aregiven in table I.Ailerons.- The wing was equipped with a high-speed aileron inboard and a low-speed aileron outboard. The locations of the ailerons are shown in figure 3. Theaileron chords were 25 percent of the local wing chord. The inboard and outb
35、oardailerons could be deflected 0 , 2.5 , 7.5 , and 12.5 . In addition, the outboardaileron could be deflected 17.5 . The ailerons had a thin plastic wiper strip atthe leading edge to prevent airflow through the wing-aileron gap.Nacelle.- A flow-through, long-duct nacelle representative of an advanc
36、edenergy-efficient engine was used throughout this investigation. The axis of thenacellewas canted inward 2 (leading edge of the nacelle toward the fuselage).Details of the nacelle are shown in figure 8. The centerline of the nacelle waslocated at D = 0.400.Pylon.- A cambered pylon was used to attac
37、h the nacelle to the wing. Previousinvestigations with cambered and symmetrical pylons (ref. 2) have shown lessinstalled drag with a cambered pylon. The pylon had an NACA 4412 airfoil section(upper surface was outboard) that was modified by local area rule shaping to reducethe installed drag at crui
38、se conditions. A typical cross section of the pylon isshown in figure 9. The pylon for this investigation was developed in previouspropulsion-integration tests of a similar configuration that did not have lateral-control surfaces. The pylon extended to the trailing edge of the wing and waslocated (s
39、panwise) near the middle of the inboard aileron. To allow deflections ofthe inboard aileron, a section was cut out of the aft end of the pylon (fig. 8). Foreach deflection, plastic material was used to fill in all but a 0.06-in. gap in thepylon. The gap allowed hinge moments to be measured without f
40、ouling.Transition StripsBoundary layer transition strips were applied to the model wing, nacelle, andpylon. The strips were comprised of a 0.1-in. wide band of Carborundum grit set in aplastic adhesive. The grains were sized on the basis of reference 7.Transition strips of No. 120 grit were located
41、on the inside and outside of thenacelle 0.7 in. aft of the lip. Transition strips of No. 120 grit were also locatedon both sides of the pylon 0.5 in. aft of the leading edge. The transition strippatterns on the wing are shown in figure 10. The positions of the transition stripson the wing were deter
42、mined from analyses of oil-flow photographs (ref. 8) of thebaseline configuration near its drag-divergence Mach number and cruise lift coef-ficient. The aft transition strip locations on the wing were used to simulate ahigher effective Reynolds number by producing a thinner boundary layer over the a
43、ftportion of the wing (ref. 9). The transition strips on the wing in the region of thenacelle were moved forward, since the influence of the nacelle caused the flowtransition to occur naturally near the wing leading edge.MeasurementsForce and moment data on the wing and fuselage were obtained with a
44、 five-component electrical strain-gage balance. Side force was not measured. An acceler-ometer attached to the wing mounting block inside the fuselage was used to measure5Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-angle of attack. Chordwise stat
45、ic pressures were measuredwith pressure-scanningvalves at seven semispanstations (fig. 3) of the wing. Chordwise locations of thepressure tubes are given in table If. Electrical strain gages were used to measurehinge momentsfor the inboard and outboard ailerons. Massflow through the nacellewas not m
46、easured; therefore, no internal drag correctionswere made.Test ConditionsData were taken at the test conditions presented in the following table:, R N , q_,Mw deg per foot ib/ft 20.30.60.80.82.84.86.90-2 to 10-2 to 10-2 to 7-2 to 7-2 to 7-2 to 7-2 to 72.5 x 1064.04.04.04.04.04.0176528660671683693714
47、Throughout the entire test, stagnation temperature was maintained at 120F, and theair was dried until the dew point was sufficiently low to prevent condensation.PRESENTATION OF RESULTSA large quantity of force and pressure data were obtained for the many configu-rations of this investigation. The da
48、ta are presented in the figures, as indicatedin the following list and table:FigureVariation of lift-curve slope with Mach number forthe baseline configuration. CL = 0.55 . 11Variation of drag coefficient with Mach number forthe baseline configuration. C L = 0.55 . 12ConfigurationLBaseline (6_ I = 0
49、, 6s 2 = O)6_ I = 0, _+2.5, _+7.5, +12.56e 2 = 0, _+2.5, +7.5, +12.5, :1:17.5Figure numbers presentingC L C D C m C I , C n CHM I, CHM2vs. _ vs. C L vs. C L vs. _ vs.20 21 22 23 2436 37 38 39 4025412642Ctccp _vs. X/C vs. n13-1927-33 134-3543-58 59-60Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-
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