1、NASA Contractor Report 3081 Wind Tunnel Force and GRANT NSG-1165 JUNE 1979 nJAsA Tests of a 21% Thick General Aviation Airfoil With 20% Aileron, 25% Slotted Flap and 10% Slot-Lip Spoiler W. H. Wentz, Jr. and K. A. Fiscko FOR EARLY DOMESTIC DISSEMINATION Because of its significant early commercial po
2、tential, this information, which has been developed under a U.S. G6v- ernment program, is being disseminated within the United States in advance of general publication. This information may be duplicated and used by the recipient with the ex- press limitation that it not be published. Release of thi
3、s information to other domestic parties by the recipient shall be made subject to these limitations. Foreign release may be made only with prior NASA ap- proval and appropriate export licenses. This legerid Mall . . . . be marked on any reproduction of this infoTmation in whdle 1, or in part. . _I D
4、ate for general release -November:197915 1. ;I: :, r, i: :; Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I TECH LIBRARY KAFB, NM ,_. _ .- OOb3883 NASA Contractor Report 3081 Wind Tunnel Force and Pressure Tests of a 21% Thick General Aviation Airf
5、oil With 20% Aileron, 25% Slotted Flap and 10% Slot-Lip Spoiler W. H. Wentz, Jr. and K. A. Fiscko Wichita Stccte University Wichita, Karzsas Prepared for Langley Research Center under Grant NSG- 116 5 National Aeronautics and Space Administration Scientific and Technical Information Office 1979 Prov
6、ided 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-,-,-I SUMMARY Force and surface pressure distributions have been measured for the 21% LS(l)-0421 modified airfoi
7、l fitted with 20% aileron,25% slotted flap and 10% slot-lip spoiler. All tests were conducted in the Walter Beech Memorial Wind Tunnel at Wichita State Uni- versity at a Reynolds number of 2.2 x lo6 and a Mach number of 0.13. Results include lift, drag, pitching moments, control sur- face normal for
8、ce and hinge moments, and surface pressure distri- butions. The basic airfoil has a ckmax of 1.31 with nearly con- stant cR beyond the stall at 2.2 x 106 Reynolds number. Incre- mental performance of flap and aileron are similar to that ob- tained on the GA(W)-2 airfoil. Spoiler control shows a slig
9、ht re- versal tendency at high 1, low spoiler deflection angle conditions with flap nested. Flap extended spoiler control is non-linear but positive. iii Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INTRODUCTION AS part of NASAs program for develo
10、ping new airfoil sec- tions for general aviation applications (ref. l), Wichita State University is conducting flap and control surface research for the new airfoils. This report documents two-dimensional wind tunnel tests of the 21% thick LS(l)-0421 modified airfoil section with: (a) 20% chord aile
11、ron, (b) 25% chord slotted flap; and (c) 10% chord slot lip spoiler. High Reynolds number tests of the LS(l)-0421 modified air- foil have been reported in reference 2. All experimental tests reported herein were conducted in the Walter Beech Memorial Wind Tunnel at Wichita State University. SYMBOLS
12、The force and moment data have been referred to the .25c location on the flap-nested airfoil. Dimensional quantities are given in International (SI) Units. Measurements were made in U.S. Customary Units. Conversion factors between the various units may be found in reference 3. The symbols used in th
13、e pre- sent report are defined as follows: C Airfoil reference chord (flap-nested) Cd Airfoil section drag coefficient, section drag/ (dynamic pressure x c) cf Flap chord Ch Control surface hinge moment coefficient, section moment about hingeline/(dynamic pressure x control surface reference chord2)
14、 CR Airfoil section lift coefficient, section lift/ (dynamic pressure x c) cm Airfoil section pitching moment coefficient with respect to the .25c location, section moment/(dynamic pressure x c2) cma Airfoil forward section moment coefficient, moment about leading edge/(c2xdynamic pressure) -_ _ - -
15、 -. _-. -ll- _._ -_.- -. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-L-A - .- - _- -._ -_ L. - _-. _-_ - Cmf cn cna Cnai (+f cP Ah P X Z a A a 6f %3 Subscripts: a f P S OJ Flap moment coefficient, moment about leading edge/ (c2xdynamic pressure)
16、Airfoil or flap normal force coefficient, section normal force/(dynamic pressurexc) Airfoil forward section normal force coefficient, normal force/(cxdynamic pressure) Aileron normal force coefficient, normal force/ (cxdynamic pressure) Flap normal force coefficient, normal force/(cx dynamic pressur
17、e) Coefficient of pressure, (p-pm)/dynamic pressure Spoiler projection height normal to local airfoil surface Static pressure Coordinate parallel to airfoil chord Coordinate normal to airfoil chord Angle of attack, degrees Increment Rotation of aileron from nested position, degrees Rotation of flap
18、from nested position, degrees Rotation of spoiler from nested position, degrees Aileron Flap Pivot Spoiler Remote free-stream value 2 . 7-m1 .-._ -;- - -._- - _-._. -. -_-_ .-_ _ ._ _,- . _ . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-APPARA
19、TUS AND TEST METHODS Model Description The LS(l)-0421 modified airfoil section is a 21% maximum thickness airfoil with a design lift coefficient of 0.4, derived from the 17% thick LS(l)-0417 (formerly designated GA(W)-1) air- foil. The LS(l).-0421 modified section is the result of several iterations
20、 of testing and theoretical analysis by the NASA Langley Airfoil Research Group to develop a highly efficient 21% thick sec- tion (ref. 2). For tests in the WSU two-dimensional facility, models were sized with 91.4 cm span and 61.0 cm chord. The forward 70% of the airfoil was fabricated from laminat
21、ed mahogany bonded to a 2.5 cm x 34.8 cm aluminum spar. Trailing edge sections were fabricated from solid aluminum for the aileron, flap and spoiler configurations. Geometric details are given in figure 1. The 20% chord aileron was designed with a 0.5% leading edge clearance gap. The 25% slotted fla
22、p and 10% spoiler were designed with an airfoil forward section which terminates at 87.5% chord. The 10% spoiler was arranged in a slot-lip configuration with the 25% slotted flap. The spoiler was fitted with ball bearing hinges at three spanwise locations, and strain-gaged cantilever beam flexures
23、at each end for hinge moment measurement. All components were equipped with 1.07 mm inside diameter pressure taps for pressure distribution surveys. Flap and aileron positioning was provided through a set of guide rails mounted on the end plate disks, external to the test section. The model and end
24、plates were mounted on the wind tunnel main balance system by means of pivot pins located at the airfoil 50% chord station. Foam seals around the circumference of the 1.07 m diameter end plates protected against flow leakage. These seals were care- fully adjusted during static calibration to avoid i
25、nterference friction forces. The model was fitted with 2.5 mm wide transition strips of #80 Carborundum grit located at 5% chord on the upper surface, and 10% chord on the lower surface. 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Instrumentati
26、on Three-component force measurements were obtained from the tunnel main balance. Spoiler hinge moment measurements were ob- tained directly from strain-gage flexures, and aileron hinge moments were obtained from integration of surface pressures. Pressure measurements were made with 96 pressure tube
27、s multiplexed to 4 unbonded pressure transducers through a system of pressure switches (see fig. 2). Resolution of the various instrumentation systems are given in Table 1: Table 1 - Instrumentation Resolution Item Resolution lift f0.9N (f0.2 lb) drag (wake survey) f0.06N (f0.014 lb) (force balance)
28、 +0.2N (to.05 lb) pitching moment +O.lN-m (+l in-lb) hinge moment pressure transducers dynamic pressure angle of attack flap and aileron angles +O.O2N-m (to.2 in-lb) f4.8N/m2 (kO.1 psf) f4.8N/m2 (fO.l psf) +o.o5O +0.5O spoiler angle f0.25O flap longitudinal and vertical settings 2.001 c Experimental
29、 data were obtained, stored and processed into final corrected form using the WSU wind tunnel on-line mini-computer system. This system had a 32 kilo-byte random access memory, two 110 kilo-byte cassette tape drives for program and raw data storage, 4 Provided by IHSNot for ResaleNo reproduction or
30、networking permitted without license from IHS-,-,-a 120 character/set printer, and 28 cm plotter with a 0.4 mm reso- lution. With this system, final data which included one-component plots were available 6 seconds after data acquisition. Final three- component plots were available 3 minutes after en
31、d of run. Incre- mental control effectiveness and pressure integrations were ob- tained by off-line computer runs on the same computing system. Test Procedure Three-component force measurements were made using the wind tunnel main balance system. Flap-nested drag measurements were made using the wak
32、e survey method. A scanning five tube pressure probe was used for this purpose. Surveys were conducted at one chord-length downstream from the model trailing edge. The difference between force balance drag and wake survey drag is end plate tare drag, which depends upon lift coefficient as well as ai
33、rfoil sec- tion. The wake survey method cannot be utilized when separation is present. For this reason it was not applied to flap extended tests. However under high drag conditions the end plate tare is a relatively small portion of total drag. This reasoning has led to the following procedure: (a)
34、for flap-nested cases the wake survey drag is used directly, (b) for flap, aileron or spoiler extended cases the drag as measured by the force balance is cor- rected by subtracting the end plate tare. The end plate tare curve is extrapolated for high lift-coefficient conditions. De- tails of this ex
35、trapolation are given in appendix A. Wind Tunnel The WSU Walter Beech Tunnel is a closed return tunnel with atmospheric test section static pressure. The test section with two-dimensional inserts is 0.91 m x 2.13 m. Complete description of the insert and calibration details are given in reference 4.
36、 Special corrections for circulation effects on the test secton static pressure system have been applied as described in Appendix B. RESULTS AND DISCUSSION Presentation of Results Test results and comparison with theory and other experi- mental results are shown in the figures as listed in Table 2.
37、5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Table 2 - List of Figures - Configuration airfoil, aileron, flap and spoiler pressure system schematic basic section basic section basic section 20% aileron 20% aileron 20% aileron 25% flap 25% flap 2
38、5% flap 25% flap 25% flap 25% flap 10% spoiler 10% spoiler Type Data model geometry CllfCdrCm pressures tufts CRICdrcm AC%, ACdrAc+,r Ch pressures optimum flap settings OR max contours CRICdrcm flap effectiveness experimental pressures pressures effect of spoilers on lift for various flap settings i
39、ncremental spoiler effectiveness and hinge moments Comparisons Figure m- 1 - 2 data of ref.2 3 theory 4 - 5 - 6 - 7 - 8 - 9 - 10 theory 11 GA(W)-2 12 - 13 theory 14 -17 - 18 - 19 Discussion Flap Nested: (figures 3 through 5). The force data show that the basic section has a very unusual stalling cha
40、racteristic. Initial stall occurs at a cRmax of 1.31 and an angle of attack of 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-11.3O. This is substantially lower than the 1.54 cRmax of the 17% thick GA(W)-1 section (ref. 5). The post-stall cLmax cu
41、rve for the 21% section is quite flat, dipping to about 1.26 at 18O and subsequently recovering to a higher level above 20. The drag and pitching moment results are similar to the lift, showing progressive changes through cRmax with no indication of abrupt separation. The NASA tests of ref. 2 show s
42、imilar results for lift and moment at 2.0x10 6 Reynolds number, but abrupt stalling character- istics at higher Reynolds numbers. The drag measurements from the present tests show the same minimum drag level as the NASA tests, but somewhat higher drag levels for lift coefficients above 0.4. The pres
43、sure distributions and tuft studies for the flap nested configuration confirm the implications of the force measurements. The separation progression is quite slow as angle of attack is in- creased. In fact both tuft pattern and pressure distributions in- dicate that even at 30 angle of attack, separ
44、ation has not reached the leading edge. Pressure distributions are characterized by very modest nose suction peaks and mild gradients. Theoretical results using the method of reference 6 show relatively poor agreement with experiment for all positive angles of attack. The discrepancies become quite
45、large for high angles of attack when massive separa- tion is present. 20% Aileron: (figures 6 through 8). Lift characteristics with aileron show that as aileron downward deflection is increased, the stalling characteristic becomes progressively more abrupt. Aileron drag, pitching moment and incremen
46、tal control effectiveness are similar to the 17% thick GA(W)-1 airfoil (ref. 8). Aileron hinge moments are similar to the GA(W)-1, but show considerable non-linearity at high angles. Pressure distributions show mild peaks and relatively slow progression of separation with angle of attack. 25% Flap:
47、(figures 9 through 17). cRmax contours for flap de- flections from loo to 35O show that the optimum flap settings are quite similar to other airfoils (for example, ref. 9). cRmax values for all flap settings are lower than comparable data for the 13% 7 Provided by IHSNot for ResaleNo reproduction or
48、 networking permitted without license from IHS-,-,-thick GA(W)-2 section (ref. 9). Theoretical results over-predict lift at 30 and 35“ flap deflection at all angles of attack. At 10“ and 20 flap settings the theory under-predicts the lift, even at low angles of attack. While the under-prediction discrepancies are not large, they are consistent with trends observed on other air- foil-flap combinations (see ref. 9). Over-prediction of lift has been attributed to boundary layer thickness exceeding theoreti
