1、NASA_Ieehn_cal1990m=.Experimental andTheoretical AerodynamicCharacteristics of aHigh-Lift SemispanWing ModelZachary T. Applinand Garl L. Gentry, Jr.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-rL.Provided by IHSNot for ResaleNo reproduction or net
2、working permitted without license from IHS-,-,-NASATechnicalPaper2990199ONationalAeronautics andSpace AdministrationOffice of ManagementScientific and TechnicalInformation DivisionExperimental andTheoretical AerodynamicCharacteristics of aHigh-Lift SemispanWing ModelZachary T. Applinand Garl L. Gent
3、ry, Jr.Langley Research CenterHampton, VirginiaProvided 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-,-,-SummaryA study was conducted to compare experimen-tal and
4、 theoretical aerodynamic characteristics of ahigh-lift semispan wing configuration. Experimen-tal data were obtained from a large semispan wingmodel that incorporated a slightly modified versionof the NASA Advanced Laminar Flow Control (LFC)airfoil section. The experimental investigation wasconducte
5、d in the Langley 14- by 22-Foot SubsonicTunnel at test-section dynamic pressures of 15 and30 psf. This provided reference chord Reynolds num-bers of 2.36 106 and 3.33 106, respectively. Atwo-dimensional airfoil code and a three-dimensionalpanel code were used to obtain aerodynamic pre-dictions. Two-
6、dimensional data were corrected forthree-dimensional effects. Comparisons between pre-dicted and measured values were made for the cruiseconfiguration and for various high-lift configurations.Both codes predicted lift and pitching-moment co-efficients that agreed well with experiment for thecruise c
7、onfiguration. These parameters were over-predicted for all high-lift configurations. Drag coeffi-cient was underpredicted for all cases. Corrected two-dimensional pressure distributions typically agreedwell with experiment, whereas the panel code over-predicted the leading-edge suction peak on the w
8、ing.One important feature missing from both thesecodes was a capability for separated flow analysis.The major cause of disparity between the measureddata and predictions presented herein was attributedto separated flow conditions.IntroductionThe purpose of the present effort was to compareexperiment
9、al and theoretical aerodynamic character-istics of a high-lift semispan wing configuration. Theexperimental data were obtained during an investi-gation in the Langley 14- by 22-Foot Subsonic Tun-nel. Theoretical predictions were obtained with atwo-dimensional airfoil code and a three-dimensionalpane
10、l code.Current analytical techniques provide adequateaerodynamic predictions for basic airplane config-urations which have little or no flow separation.However, these techniques typically lack the capa-bility to determine aerodynamic characteristics forconditions of extensive flow separation. Signif
11、icantflow separation can exist on airplanes for severalcommon operational situations. For example, sep-aration may be present on the upper surface oftrailing-edge flaps during high-lift takeoff and land-ing conditions. In addition, recent geometries devel-oped for highly maneuverable fighter airplan
12、es aredesigned for operation at extreme angles of attackwhere separated flow is certain to occur.The primary interest of the present study is inconfigurations with trailing-edge and leading-edgeflaps deployed, where highly viscous interactions andflow separation cause inaccurate and sometimes mis-le
13、ading predictions of aerodynamic characteristics.However, comparisons are also presented for thecruise and trailing-edge-flap-only configurations.The airfoil code used to calculate two-dimensionalaerodynamic characteristics was the Multi-Component Airfoil (MCARF) program describedin references 1 and
14、 2. This program combinesboundary-layer solutions with potential flow pressuredistributions to obtain viscous aerodynamic charac-teristics of airfoil geometries.The panel code, VSAERO, calculates nonlinearaerodynamic characteristics of partial or completeconfigurations in the subsonic flow regime (r
15、efs. 3and 4). Nonlinear effects of vortex flow interactionwith flow fields and surfaces are treated with wake re-laxation techniques in an iterative procedure. In anapproach that is similar to MCARF, VSAERO canaccount for viscous effects by coupling a potentialflow solution with strip boundary-layer
16、 calculations.Several wing configurations were analyzed to deter-mine the viscous effect as predicted by VSAERO.The difference between viscid and inviscid solutionswas insignificant; therefore, only inviscid solutionsare presented herein.SymbolsAll longitudinal aerodynamic data are referredto the wi
17、nd axis system. Dimensions of the cruiseconfiguration were used to nondimensionalize aero-dynamic force and moment data.b wing semispan, 116.01 in.DragC D drag coefficient, -qccSLiftC L lift coefficient, -q_SCL_ lift-curve slope, per degCm pitching-moment coefficient aboutPitching momentquarter-chor
18、d,q_cScstatic pressure coefficient, ps -pocq_reference wing chord, 39.37 in.Liftsection lift coefficient, -qoccsurface static pressure, lb/ft 2GcclPsl Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Pot)qocSX, y_ ZO_C_MCARF5LE_ST E71Abbreviations:fre
19、e-stream static pressure, lb/ft 2free-stream dynamic pressure, lb/ft 2reference wing area, 31.72 ft 2coordinates of pressure taps, in.angle of attack of WRP, degMCARF angle of attack used forpressure distribution comparisons,degleading-edge flap deflection angle,positive trailing edge down, degtrail
20、ing-edge flap deflection angle,positive trailing edge down, degL.E. leading edgeLFC laminar flow controlT.E. trailing edgeWRP wing reference plane of cruiseconfigurationTest SetupThe unswept semispan wing model was tested inthe Langley 14- by 22-Foot Subsonic Tunnel whichis a closed, single-return,
21、atmospheric wind tunnelwith a test section 14.50 ft high by 21.75 ft wide by50.00 ft long. (See ref. 5.) The test-section dynamicpressure is continuously variable from 0 to 144 psf.The tunnel is equipped with a floor boundary-layerremoval system consisting of a floor-mounted suc-tion grid located 8.
22、2 ft upstream of the wing lead-ing edge. The suction grid spans the floor of thetest; section between the tunnel walls and reducesthe boundary-layer thickness to approximately 1.6 in.at the wing location for the empty tunnel condi-tion. The model was mounted vertically, protrudingthrough the floor,
23、on a six-component strain-gaugebalance which was located below a 15.8-ft-diamcterturntable which could be rotated throughout theangle-of-attack range of the wing. Angle of attackof all configurations was referenced to the wing refer-ence plane of the cruise configuration. The yaw angleof the turntab
24、le was detected by a digital shaft en-coder geared to the turntable mechanism. This pro-vided an angle-of-attack accuracy to within -t-0.02 .,The l l6.01-in, semispan, rectangular, untwistedwing model had a 39.37-in. chord incorporating aslightly modified version of the NASA Advanced2Laminar Flow Co
25、ntrol (LFC) airfoil section pre-sented in references 6 through 8. Maximum thick-ness of the airfoil section was 0.13c. The unmodifiedairfoil section was designed to provide shock-free flowover the upper surface at high subsonic Mach num-bers as described in reference 6. The current studyinvestigates
26、 the low-speed characteristics of the mod-ified airfoil shape, with and without high-lift devices.Modifications to the airfoil shape included a shiftin the lower surface lobe rearward by 2 percent ofthe chord and a slight increase in trailing-edge cam-ber. These modifications allowed sufficient leng
27、th inthe chordwise direction, forward of the lower sur-face lobe, for storage of a Krueger-type flap of upto 12 pcrccnt chord. A Krueger-type flap was chosenbecause possible surface discontinuities when stowed(i.e., steps, gaps) would be in a region of favorablepressure gradients generated by the ai
28、rfoil contour.(See ref. 6.) No analysis has been made of the inter-nal volume required for storage of the Krueger-typeflap or for the necessary deployment mechanism.This model was fabricated to investigate aerody-namic characteristics for the high-lift configuration (acondition for which LFC is not
29、practical). Therefore,no provisions were made for an LFC suction system.The model high-lift components included eithera 0.10c or a 0.12c full-span leading-edge flap and afull-span 0.25c trailing-edge flap. All components ofthis semispan model had rounded tips. A sketch ofthe model planform and photo
30、graphs of the modelinstalled in the tunnel are presented in figure 1. Asingle row of pressure taps located at _ = 0.44was used to obtain surfacc pressure distributions.Coordinates of the wing airfoil section for the cruiseand main element of the high-lift configurations aregiven in terms of the loca
31、tions of surface pressuretaps and are presented in tables I and II, respectively.Coordinates of the trailing-edge flap are presented intable III; coordinates of the two leading-edge flapsare presented in table IV. Section contours of theconfigurations tested during this investigation areshown in fig
32、ure 2.The leading- and trailing-edge flaps were posi-tioned using the definitions for deflection, gap, andoverlap presented in reference 9. Reference lines forthese definitions pass through the leading and trail-ing edge of each component, including the main cle-ment of the high-lift configurations.
33、 For the trailing-edge flap, the gap and overlap were 0.02c and 0.00c,respectively. For both leading-edge flaps, the gap andoverlap were 0.012c and 0.016c, respectively. Thesesettings were used for all deflection angles tested inthis investigation.The wing was fabricated from solid aluminumby a nume
34、rically controlled milling machine. TheProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-resultantcontourwaswithin 0.005in. of thespec-ified airfoil coordinates. Surfacepressuretubeswere routed internally to pressuremeasurementinstrumentationlocatedbel
35、owthetunnelfloor. Forconfigurationswith the trailing-edgeflap installed(fig. l(b), the cruisetrailingedgewasreplacedbya covesectionwhichprovidedsupportbracketsandpressure-tuberoutingrecessesfor the flappressuretubes.Leading-edgeflapsweresupportedby brack-etsmountedonthelowersurfaceof theleadingedgeo
36、f thecruisewing.Pressuretubesfromthehigh-liftcomponentswereroutedexternallyalongthesupportbrackets(7/= 0.377)to thewing. Thesetubesweretheninternallyroutedthroughthewingto thepres-sureinstrumentationlocatedbelowthetunnelfloor.Theexternaltubesweretightly tapedto the flapbracketsandstreamlinedwith the
37、 useof modelingclayto produceasmoothsurface.Modelingclaywasalsousedto streamlinethe remainingflapbracketsnotusedto routepressuretubes.Spanwiselocationsof theflapbracketcenterlinesaregivenin tableV.Therewasa 1.5-in-widegapbetweenthewingup-persurfaceandthetunnelfloorplates(0.25in.thick)wherethe wingpr
38、otrudedthroughthetunnelfloor.Thisgapwasprovidedto preventfoulingwhenaero-dynamicloadingcausedthe balanceand wing todeflect. A 1.0-in-widegap wasprovidedfor thelowersurface.Toreduceairflowthroughthisgap,a2-in-thickpadofclosed-cellfoamrubber(whichover-lappedthetunnelfloor)wasattachedtothewingjustbelow
39、the tunnelfloor. An electricalfoulingcircuitalertedthe tunneloperatorif anycontactoccurredbetweenthewingandtunnelfloor.Boundary-layertransitionstrips1/8in.widewereappliedusingNo.60Carborundum grit. The transi-tion roughness was sized according to the procedureoutlined in reference 10. These transiti
40、on strips werelocated on both the upper and lower surfaces at the5-percent-chord station for the cruise configurationand extended across the entire span. For the high-lift configuration, the same grit was located 2 in. fromthe leading edge on the main component and 1 in.from the leading edge on all
41、the flaps.Pressure measurements were obtained with anelectronically scanned pressure (ESP) system. Thissystem consisted of modules which contained a720-psf-range silicon pressure transducer for everyport. These transducers were operated as 144-psf-range transducers by the addition of sensitizing ele
42、c-tronics. The manufacturers quoted accuracy for thesystem when operated in this range is =t=0.5 psf. Thepressure transducers were referenced to atmosphericpressure and had an over range capability. Sixteenpressure ports near the leading edge of the wing wereconnected in parallel to a 720-psf and a
43、144-psf trans-ducer to assure accurate measurement of pressure.above 144 psf. On-line calibration was possible witlthis system and was done before every run to main-tain a high degree of accuracy. When a data pointwas measured, each of the pressure transducers wasscanned electronically at up to 20 0
44、00 measurementsper second; thus all pressure data were acquired atessentially the same time.Aerodynamic force and moment measurementswere obtained with an existing six-component,strain-gauge balance, which had previously been usedon a semispan wing similar in size to the LFC wing.Balance load charac
45、teristics, as well as its effect onthe accuracy of aerodynamic coefficients, are pre-sented in table VI. The previous model incorporatedan NACA 0012 airfoil section (ref. 11). The LFCwing used the same mounting hardware as used forthe NACA 0012 wing. It was determined that the ex-isting balance did
46、not have sufficient load capacity toallow operation of the LFC wing at the maximum liftcondition (stall angle of attack). The investigation ofthe aerodynamic characteristics of the LFC wing wastherefore limited to moderate angles of attack.Test ProceduresThe model was tested in four different config
47、ura-tions as shown in the following table:Configuration 6TE, degCruiseTrailing-edge flap only10-percentleading-edge flap12-percentleading-edge flap1515, 3015, 30_LE, degThe angle-of-attack range varied with model configu-ration and was limited by the load capacity and sta-bility of the balance. Test
48、-section dynamic pressuresof 15 and 30 psf (Mach numbers of 0.10 and 0.14)were used throughout the investigation; this providedreference chord Reynolds numbers of 2.36 x 106 and3.33 x 106, respectively. Unfortunately, due to amalfunction in the data acquisition system, no datawere obtained at qoc -
49、30 psf for the 10-percentleading-edge flap configuration with _LE - -500 and_TE = 15-Although all six force and moment componentswere measured with the balance, only the longitu-dinal aerodynamic data are presented. Since themodel was mounted perpendicular to the tunnel3Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-floor,
