1、NASA Technical Memorandum 85674 NASA-TM-8567419850007386 1Aileron Effectiveness fora Subsonic Transport ModelWith a High-Aspect-RatioSupercritical WingPeter F. JacobsDECEMBER 1983 %,25th Anniversary1958-1983Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS
2、-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NASA Technical Memorandum 85674Aileron Effectiveness fora Subsonic Transport ModelWith a High-Aspect-RatioSupercritical WingPeter F. JacobsLangley Research CenterHampton, VirginiaNIANational Aerona
3、uticsand Space AdministrationScientific and TechnicalInformation Branch1983Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SUMMARYThe purpose of this in
4、vestigation was to determine aileron effectiveness for asubsonic energy-efficient transport (EET) model with a high-aspect-ratio supercriti-cal wing. This investigation was conducted in the Langley 8-Foot Transonic PressureTunnel. Data were taken over a Mach number (M) range of 0.30 to 0.86. The Rey
5、noldsnumber was 3.0 106 per foot for M_ = 0.30 and 5.0 106 per foot for the otherMach numbers. Data are presented for ailerons located at three positions along thewing span. The ailerons were designed as a preliminary active-control concept withgust-load alleviation, maneuver-load alleviation, and f
6、lutter-suppression systems.The data indicate a linear variation of rolling-moment coefficient with angle ofattack for individual and multiple aileron deflections at Mach numbers up to 0.81.For Mach numbers greater than 0.81, the rolling-moment-coefficient data become non-linear with increasing angle
7、 of attack. At Mach numbers near the design value(M = 0.81), increased aileron effectiveness resulted from aft transition locations,which produced relatively thin boundary layers (higher effective Reynolds number) andgreater effective aileron deflections. Individual aileron deflections on the rightw
8、ing panel produced only small effects on yawing-moment and side-force coefficients.INTRODUCTIONSince the development of advanced-technology supercritical airfoils by theNational Aeronautics and Space Administration, great strides have been made towardimproving the cruise performance of future jet tr
9、ansport aircraft. Extensive theo-retical studies and experimental wind-tunnel investigations have produced aerodynami-cally efficient transport wings which have higher lift-drag ratios, thicker airfoilsections, less sweep, and higher aspect ratios than the wings on current wide-bodyaircraft. The per
10、formance characteristics of these configurations have been docu-mented in references I and 2; however, data on the effectiveness of lateral-controlsurfaces for these supercritical wings have not generally been available.The purpose of this investigation was to determine aileron effectiveness for ahi
11、gh-aspect-ratio supercritical wing configuration. The control surfaces investi-gated were representative of a preliminary active-control technology concept withgust-load alleviation, maneuver-load alleviation, and flutter-suppression systems(ref. 3). These controls did not correspond directly to con
12、ventional ailerondesigns, either in size or location. It is anticipated, however, that this investi-gation will provide insight into the sizing of a more conventional set of aileronsfor a high-aspect-ratio supercritical wing configuration.SYMBOLSForce and moment data presented in this paper have bee
13、n reduced to conventionalcoefficient form based on the wing trapezoidal planform area (extended to the fuse-lage centerline). Longitudinal aerodynamic characteristics are referred to thestability-axis system, and lateral-directional aerodynamic characteristics arereferred to the body-axis system. Mo
14、ments are referenced to the quarter chord of theProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-mean geometric chord. All dimensional values are given in U.S. Customary Units.Symbols are defined as follows:al,a2,a 3 ailerons I, 2, and 3, respectively
15、 (fig. 2)b wing span, 52.97 in.CD drag coefficient Dragq SLiftCL lift coefficient, q SC rolling-moment coefficient Rolling moment1 q Sb,C I ,C control-effectiveness parameter for ailerons I, 2, and 3,C16ai 6a2 16a3 AC1respectively, _, per degreeC pitching-moment coefficient, Pitching momentmq S_Cn y
16、awing-moment coefficient Yawing moment q SbSide forceCy side-force coefficient,q Sc local streamwise chord of wing, in.c mean geometric chord of reference wing panel, 5.74 in.M free-stream Mach numbercoq_ free-stream dynamic pressure, ib/ft2R Reynolds number per footS wing planform reference (trapez
17、oidal) area, 1.988 ft 2t/c local wing maximum thickness-to-chord ratiox chordwise distance from wing leading edge, positive aft, in.y spanwise distance from model centerline, in.z vertical coordinate of airfoil, positive upward, in.angle of attack, degA incremental valueProvided by IHSNot for Resale
18、No reproduction or networking permitted without license from IHS-,-,-6a deflection angle of aileron, positive for trailing-edge down, deg local wing incidence angle measured from fuselage waterline, positive forleading edge up, deg2ysemispan station, _-Subscripts:1,2,3 ailerons I, 2, and 3, respecti
19、velyEXPERIMENTAL APPARATUS AND PROCEDURESTest FacilityThis investigation was conducted in the Langley 8-Foot Transonic Pressure Tunnel(ref. 4). This facility is a continuous-flow, single-return tunnel with a rectangu-lar, slotted test section. Tunnel controls allow independent variation of Mach num-
20、ber, density, stagnation temperature, and dew-point temperature. The test section isapproximately 7.1 ft square (same cross-sectional area as that of a circle with an8.0-ft diameter). The ceiling and floor are slotted axially and have an averageopenness ratio of 0.06. These features permit the test-
21、section Mach number to bechanged continuously throughout the transonic speed range. The stagnation pressurein the tunnel can be varied from a minimum of 0.25 atm (I atm = 2116 ib/ft2) at allMach numbers to a maximum of approximately 2.00 atm at Mach numbers less than 0.40.At transonic Mach numbers,
22、the maximum stagnation pressure that can be obtained isapproximately 1.5 atm.Model DescriptionDrawings of the model are shown in figures 1 and 2. A photograph of the modelin the Langley 8-Foot Transonic Pressure Tunnel is shown in figure 3.Fuselage.- The fuselage used in this investigation had a max
23、imum diameter of5.74 in. and was 49.56 in. long. The fuselage wetted area was approximately5.63 ft2. The fineness ratio of the fuselage (8.6) was typical of second-generationor wide-body jet transports. The lower surface of the wing was faired into the fuse-lage to produce a relatively flat bottom t
24、hat extended from near the wing leadingedge to approximately 6.0 in. aft of the trailing edge.Win_- The reduced-camber wing of reference 2 was used in this investigation.The wing had 5 of dihedral and 30 of sweep at the quarter chord. Based on thetrapezoidal planform (extendedto the fuselage centerl
25、ine), the wing had a referencearea of 1.988 ft2, an aspect ratio of 9.80, and a taper ratio of 0.397. Twist andthickness distributions are shown in figures 4 and 5, respectively. Airfoil sectionsat several spanwise locations are shown in figure 6.Ailerons.- An existing set of wing panels was modifie
26、d for trailing-edge con-trols. The three ailerons represented a preliminary active-control design and werelocated on the right wing panel only. Details of the ailerons are shown in figure 2.Each aileron had angle brackets which produced deflections of 0, 5, and 10. Aplastic filler material prevented
27、 airflow through the gap between the wing and theProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-leading edges of the ailerons to simulate a sealed aileron configuration. The fillermaterial was shaped to provide a smooth contour between the wing and
28、ailerons.Transition StripsBoundary-layer transition strips were applied to the fuselage and the wing.These strips were comprised of a 0.10-in-wide band of carborundum grit set in a plas-tic adhesive. The grit was sized on the basis of reference 5.A transition strip of No. 120 grit was applied to the
29、 fuselage I in. aft of thenose. The transition strip patterns on the wing are shown in figure 7. The transi-tion strips on the wing were located rearward in an attempt to simulate a highereffective Reynolds number (ref. 6).MeasurementsForce and moment data were obtained by use of a six-component ele
30、ctrical strain-gauge balance housed within the fuselage cavity. Angle of attack was measured by anaccelerometer that was also housed within the fuselage. Static pressures were mea-sured in the model sting cavity by using differential-pressure transducers referencedto free-stream static pressures.Cor
31、rectionsThe angle of attack of the model was corrected for flow angularity in the tunneltest section. This correction (approximately 0.1) was obtained from upright andinverted tests of the basic wing configuration. The drag data have been adjusted tocorrespond to the condition of free-stream static
32、pressure in the sting cavity. NoMach number correction was made for blockage effects, which were estimated to benegligible. Control-effectiveness-parameter values were computed using nominalcontrol-deflection angles and were not corrected for control deflections under load.Test ConditionsThroughout
33、the entire test, stagnation temperature was maintained at 120F, andthe air was dried until the dew point was sufficiently low to prevent condensationeffects. The test conditions for which data were taken are presented in the follow-ing table:M_ _, deg R, per foot q_, lb/ft20.30 -4 to 14 3.0 x 106 21
34、0.60 -4 to 16 5.0 660.70 -4 to 10 5.0 747.77 -4 to 6 5.0 802.81 -4 to 8 5.0 83384 -4 to 8 5.0 853.86 -4 to 8 5.0 867Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-PRESENTATION OF RESULTSThe results of this investigation are presented in the figures,
35、 as indicated inthe following table:Aileron configuration Figure numbers6ai 6a2 6a3 CD vs CL CL vs _ Cm vs CL C , C , and C vsdeg deg deg 1 n Y0 0 0 8 18 28 38-10 to 10 0 0 9 19 29 390 -10 to 10 0 10 20 30 400 0 -10 to 10 11 21 31 41-5 0 -5 12 22 32 425 0 -5 13 23 33 435 0 5 14 24 34 445 5 0 15 25 3
36、5 450 5 5 16 26 36 465 5 5 17 27 37 47The variation of aileron-effectiveness parameter with Mach number is shown infigure 48.DISCUSSION OF RESULTSThe purpose of this investigation was to determine aileron effectiveness for asubsonic EET model with a high-aspect-ratio supercritical wing. Since the ai
37、leronswere located on the right wing panel only, data for positive and negative deflectionsof each surface were summed in the calculation of the control-effectiveness values(i.e., on the basis of asymmetric control deflections A6a = 6a - 6a ).down upLongitudinal Aerodynamic CharacteristicsThe static
38、 longitudinal aerodynamic characteristics of the model (figs. 8 to 37)are not representative of an actual aircraft because ailerons were located on theright wing panel only. The data are included and may be used judiciously, but theyare considered of secondary importance to the lateral aerodynamic c
39、haracteristics andwill not be discussed.Lateral-Control CharacteristicsThe static lateral aerodynamic data for the baseline configuration (no aileronsdeflected) indicate a slight asymmetry of the model (fig. 38). The model asymmetryaffects the absolute value of the baseline rolling-moment coefficien
40、t but has noeffect on the calculation of the control-effectiveness parameter, which depends onincrements in rolling-moment coefficient. Nolling-moment-coefficient values for thebaseline configuration are positive for most angles up to stall and increase slightly5Provided by IHSNot for ResaleNo repro
41、duction or networking permitted without license from IHS-,-,-with Mach number. Poststall rolling-moment-coefficient values become very nonlinearfor the baseline configuration at M = 0.60 (fig. 38(b). This effect is probablycaused by varying amounts of flow separation and aeroelastic deformation at t
42、he wingtip. A similar trend is shown for 5 and 10 deflections of the outboard aileron a3(fig. 41(b). Positive deflections of a3 increase the loading at the wing tip;however, the increased loading is partially offset by aeroelastic deformation of thewing (washout).Data for individual aileron deflecti
43、ons (figs. 39 to 41) and multiple ailerondeflections (figs. 42 to 47) indicate a linear variation of rolling-moment coeffi-cient with angle of attack up to stall for Mach numbers up to and including thedesign Mach number (M = 0.81). For Mach n_nbers greater than 0.81, rolling-moment-coefficient valu
44、es become nonlinear with increasing angle of attack, probably as aresult of increased trailing-edge boundary-layer separation near the deflected aile-rons. Aileron aI shows a control reversal anomaly in the rolling-moment data forMach numbers of 0.84 and 0.86 (figs. 39(f) and 39(g). This anomaly is
45、probablycaused by shock wave interaction between the wing and fuselage, since aileron aI islocated approximately 0.2 in. from the fuselage. _olling-moment data for multipleaileron deflections can be estimated by summing the data for individual deflections.The effect of individual aileron deflections
46、 (rightwing panel) on yawing-moment andside-force coefficients is very small.Aileron EffectivenessFor the Reynolds numbers of this test, aileron effectiveness is highly dependenton the transition location and growth rate of the boundary layer. Data in refer-ence 7 indicate that thicker boundary laye
47、rs (resulting from a forward movement ofthe transition strips) caused reduced aileron effectiveness. The thickening of theboundary layer on the ailerons may be considered as an effective reduction of theaileron deflection.As previously stated, the transition strips for this investigation were locate
48、drearward to simulate a higher effective Reynolds number at near-cruise Mach numbers.However, the actual point of transition varies as a result of changes in chordwisepressure distribution with Mach number (ref. 8). For off-design conditions(M 0.70), the supercritical wing of this investigation exhi
49、bits “peaky“-typechordwise surface pressure distributions. The adverse pressure gradient of thisupper-surface pressure peak causes transition near the leading edge instead of at thetransition strip location and results in a thicker boundary layer over the aft por-tion of the wing. As the Mach number approaches
