1、I 0 c I , NASA TN D-7428 N74-11821 1 4 .) 7 LOW-SPEED AERODYNAMIC CHARACTERISTICS OF A 17-PERCENT-THICK AIRFOIL SECTION DESIGNED FOR GENERAL AVIATION APPLICATIONS Robert J. McGhee, et a1 Langley Research Center Hampton, Virginia LOAN COPY: RETUR AFWL TECHNICAL LIE “KIRTLAND AFB. N. December 1973 DIS
2、TRIBUTED BY: National Technical information Service U. S. DEPARTMENT OF COMMERCE TO !ARY n. Provided 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-,-,-w 4. Title u
3、xl Subtitle D-7428 LOW-SPEED AERODYNAMIC CHARACTERISTICS OF A 17-PERCENT-THICK AIRFOIL SECTION DESIGNED FOR GENE Ff Research INTRODUCTION on advanced aerodynamic technology airfoils has been conducted over the last several years at the Langley Research Center. Results of this research have been appl
4、ied to the design of a 17-percent-thick airfoil suitable for a propeller driven light airplane. 3 I- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The subcritical characteristics of thick supercritical airfoil section research of Some reference 1 i
5、ndicated performance increases over conventional airfoil sections. of the features that produce these favorable aerodynamic characteristics have been applied in the design of a new low-speed airfoil section. This new airfoil is one of sev- eral being developed by NASA for light airplanes and has bee
6、n designated as General Aviation (Whitcomb) -number one airfoil (GA(W)- 1). dimensional aerodynamic characteristics of the NASA GA(W)- 1 airfoil section. In addi- tion, the results are compared to a comparable NACA 65 series airfoil section. Such sections are presently used on some light airplanes.
7、Also, the experimental results are compared with results obtained from an analytical aerodynamic performance prediction method. The present investigation was conducted to determine the basic low-speed two- The investigation was performed in the Langley low-turbulence pressure tunnel over a Mach numb
8、er range from 0.10 to 0.28. The Reynolds number, based on airfoil chord, varied from about 2.0 X lo6 to 20.0 X lo6. The geometrical angle of attack varied from about -loo to 24O. SYMBOLS Values are given both in SI and the U.S. Customary Units. The measurements and calculations were made in the U.S.
9、 Customary Units. a mean- line designation cP PL - pco qco pressure coefficient, C chord of airfoil, cm (in.) section chord-force coefficient, CC max s forward (t/c), Cd section profile-drag coefficient determined from wake measurements, 2 Provided by IHSNot for ResaleNo reproduction or networking p
10、ermitted without license from IHS-,-,-d Cl c2, i cn h l/d M P R t X z CY P point drag coefficient, section lift coefficient, design section lift coefficient Cn cos CY - cc sin CY section pitching-moment coefficient about quarter chord, Cp d($) - Cp d(5) L.S. U.S. section normal-force coefficient, ve
11、rtical distance in wake profile, cm (in.) section lift-drag ratio, Cl/Cd free-stream Mach number s tatfc pr e s sur e, N/ m2 dynamic pressure, N/m2 (lb/ft2) (lb/f t2) Reynolds number based on free-stream conditions and airfoil chord airfoil thickness, cm (in.) airfoil abscissa, cm (in.) airfoil ordi
12、nate, cm (in.) mean line ordinate, cm (in.) angle of attack of airfoil, angle between chord line and airstream axis, deg density, kg/m3 (slugs/ft3) 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Ill1 II Illll Subscripts: L local point on airfoil m
13、ax maximum t thickness 1 tunnel station 1 chord length downstream of model 2 tunnel station downstream of model where static pressure is equal to free-stream static pressure m undisturbed stream conditions Abbreviations: GA(W)- 1 General Aviation (Whitcomb)-number one 1. s. lower surface U.B. upper
14、surface AIRFOIL DESIGN The airfoil section (fig. 1) was developed by employing some of the favorable char- acteristics of the thick supercritical airfoil of reference 1, which indicated performance increases over conventional airfoils at subcritical conditions. In order to expedite the airfoil devel
15、opment, the computer program of reference 2 was used to predict the resuits of various design modifications. The final airfoil shape was defined after 17 iterations on the computer. The airfoil is 17 percent thick with a blunt nose and a cusped lower sur- face near the trailing edge. The design crui
16、se lift coefficient was about 0.40 at a Reynolds number of about 6 X 10 6 . In defining the airfoil emphasis was placed on providing good lift-drag ratios at cl = 1.0 for improved climb performance, and on providing a maximum lift coefficient of about 2.0. Several key design features of the airfoil
17、are: 1. A large upper surface leading-edge radius (about 0.06) was used to attenuate the peak negative pressure coefficients and therefore delay airfoil stall to high angles of attack. 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2. The airfoil
18、was contoured to provide an approximate uniform chordwise load distribution near the design lift coefficient of 0.40. To account for viscous effects this airfoil incorporated more camber in the rear of the airfoil than the NACA mean camber line (fig. 2). 3. A blunt trailing edge was provided with th
19、e upper and lower surface slopes approximately equal to moderate the upper surface pressure recovery and thus postpune the stall. The airfoil thickness distribution and camber line are presented in figure 2. Table I presents the measured airfoil coordinates. 1 APPARATUS AND PROCEDURE Model Descripti
20、on The airfoil model was machined from an aluminum billet and had a chord of 58.42 cm (23 in.) and a span of 91.44 cm (36 in.). The airfoil surface was fair and smooth. upper and lower surface orifices located at the chord stations indicated in table II. A base pressure orifice was included in the b
21、lunt trailing edge of the airfoil (x/c = 1.0). In order to provide data for a simple flap deflection, an aluminum wedge was installed on the model to simulate a split flap deflected 60. Orifices were installed on this simulated flap as indicated in table II. Figure 3 shows a photograph of the model.
22、 The model was equipped with both I Wind Tunnel The Langley low-turbulence pressure tunnel (ref. 3) is a closed-throat single- 2 return tunnel which can be operated at stagnation pressures from 101.3 to 1013 kN/m (1 to 10 atm) with tunnel-empty test-section Mach numbers up to 0.46 and 0.23, respec-
23、tively. The maximum unit Reynolds number is about 49 X 106 per meter (15 X 106 per foot) at a Mach number of 0.23. (7.5 ft) high. The test section is 91.44 cm (3 ft) wide by 228.6 cm Circular end plates provided attachment for the two-dimensional model. The end plates are 101.6 cm (40 in.) in diamet
24、er and are flush with the tunnel wall. They are hydraulically rotated to provide for model angle-of-attack changes. The airfoil was mounted so that the center of rotation of the circular plates was at 0.25 on the model chord line. The air gaps at the tunnel walls were sealed with flexible-sliding me
25、tal seals (fig. 4). 5 Y Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Wake Survey Rake A fixed wake survey rake (fig. 5) at the model midspan was mounted from the tun- nel sidewall and located 1 chord length rearward of the trailing edge of the air
26、foil. The wake rake utilized 91 total-pressure tubes and five static-pressure tubes 0.1524 cm (0.060 in.) in diameter. The total-pressure tubes were flattened to 0.1016 cm (0.040 in.) for 0.6096 cm (0.24 in.) from the tip of the tubes. The static pressure tubes had four flush orifices drilled 90 apa
27、rt and located 8 tube diameters from the tip of the tube and in the measurement plane of the total-pressure tubes. Three tunnel sidewall static total-pressure tubes. One static orifice was located on the center line of the tunnel and the other two orifices were about 0.35 above and below the center
28、line of the tunnel. pressures were also measured from orifices located in the measurement plane of the r Inst rumentation Measurements of the static pressures on the airfoil surfaces and the wake rake pressures were made by an automatic pressure-scanning system utilizing variable capacitance type pr
29、ecision transducers. Basic tunnel pressures were measured with precision quartz manometers. Angle of attack was measured with a calibrated potenti- ometer operated by a pinion gear and rack attached to the circular plates. Data were obtained by a high-speed data-acquisition system and recorded on ma
30、gnetic tape. TESTS AND METHODS The airfoil was investigated at Mach numbers from 0.10 to 0.28 over an angle-of- attack range from about -loo to 24O. Reynolds number based on the airfoil chord was varied from about 2.0 X lo6 to 20.0 X 10 6 , primarily by varying the tunnel stagnation pressure. The mo
31、del was tested both with the wake rake installed and removed to deter- mine its influence on the flow over the airfoil. Figure 6 shows typical lift coefficient and pitching-moment-coefficient data and no effects were indicated. The pressure distribu- tion data also indicated no effect of the wake ra
32、ke on the flow over the airfoil. The air- foil was tested both smooth (natural boundary-layer transition) and with roughness located on both upper and lower surfaces at 0.08. The roughness was sized according to refer- ence 4 which indicated a nominal roughness particle height of 0.0107 cm (0.0042 i
33、n.) at a Reynolds number of 6 x 106 and 0.0257 cm (0.0101 in.) at a Reynolds number of 2 X 106. The corresponding commercial grit numbers required are number 120 and number 60. The transition strips were 0.25 cm (0.10 in.) wide. The roughness was sparsely spaced and attached to the airfoil surface w
34、ith lacquer. Several different roughness sizes were used for the same test conditions and these results are shown in figure 7. For several runs the standard NACA method of applying roughness (number 60 grit wrapped around leading edge on both surfaces back to 0.08) was employed (ref. 5). For several
35、 test-runs 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-oil was spread over the airfoil upper surface to determine if any local flow separation was present. Tufts w2re attached to the airfoil and tunnel sidewalls with plastic tape to determine s
36、tall patterns on both the airfoil and adjacent sidewalls. The static-pressure measurements at the airfoil surface were reduced to standard pressure coefficients and then machine integrated to obtain section normal-force and chord-force coefficients and section pitching-moment coefficients about the
37、quarter chord. Section profile-drag coefficient was computed from the wake rake measurements by the method of reference 6. influence on the static pressures due to the presence of the rake body; therefore, the tunnel sidewall static pressures were used in computing the section profile-drag coefficie
38、nts. The wake rake static-pressure measurements indicated some An estimate of the standard low-speed wind-tunnel boundarp corrections as cal- culated by the method of reference 7 is shown in figure 8. These corrections amount to about 2 percent of the measured coefficients and have not been applied
39、to the data. An estimate of the total head tube displacements effects on the values of Cd showed these effects to be negligible. RESULTS The results of this investigation have been reduced to coefficient form and are presented in the following figures: Figure Tuft photographs of NASA GA(W)- 1 airfoi
40、l . 9 10 Effect of Reynolds number on section characteristics, model smooth . Effect of Mach number on section characteristics, model smooth, R=6X106 11 Effect of Reynolds number on section characteristics, roughness located at0.08 12 Effect of Mach number on section characteristics, roughness locat
41、ed at O.O8c, R = 6 x lo6. 13 Effect of roughness on section characteristics . Comparison of section characteristics between NASA GA(W)- 1 and 14 NACA 652-415 and 653-418 airfoils . 15 for various airfoils without flaps 16 Variation of maximum section lift coefficient with Reynolds number 7 Provided
42、by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I I lllIlIlllIllllllllIllllll llllll llllll Figure Section characteristics for 0.20 simulated split flap deflected 60 . . . . . . . . 17 Effect of angle of attack on chordwise pressure distributions . . . . . .
43、. . . . . 18 Comparison of experimental and theoretical aerodynamic characteristics . . . . 19 Comparison of experimental and theoretical chordwise pressure distributions . . 20 DISCUSSION OF RESULTS Experiment a1 Results Lift.- Figure 10 shows that with the airfoil smooth (natural boundary-layer tr
44、ansi- tion) a lift-curve slope of about 0.12 per degree and a lift coefficient of about 0.52 at a = 0 was obtained for all Mach numbers and Reynolds numbers investigated, Maxi- mum lift coefficients increased from about 1.64 to about 2.12 as the Reynolds number was increased from about 2 X lo6 to 12
45、 X lo6 at M = 0.15 (fig. 16), with the most rapid increase occurring between Reynolds numbers of 2 x 106 and 6 x 106. Increas- ing the Reynolds number above 12 X lo6 had no additional effect on maximum lift coef- ficient as shown by figure 1O(b) (M = 0.20). The GA(W)- 1 airfoil section encounters a
46、gradual type stall (fig. lo), particularly in the lower Reynolds number ranges. Tuft pictures (fig. 9) indicated the stall is of the turbulent or trailing-edge type. (See also pressure data of fig. 18.) At a Reynolds number of 6.0 X lo6, increasing the Mach number from 0.10 to 0.28 had only a minor
47、effect on the lift characteristics as shown by the results presented in figure ll(a). The stall angle of attack was decreased about 2O and maximum lift coeffi- cient about 5 percent. The addiiion of roughness at 0.08 (figs. 12 and 14) did alter the effective airfoil shape because of changes in bound
48、ary-layer thickness, particularly for R = 2.0 X 106 as shown in figure 14(a). For example, the angle of attack for zero lift coefficient changed from about -4O to -3.6. No measurable change in lift-curve slope was indicated; there- fore, the lift coefficient at (Y = Oo decreased from about 0.52 to a
49、bout 0.43. These effects on the lift characteristics decreased as the Reynolds number was increased above 2.0 x 106 as might be expected because of the related decrease in boundary-layer thick- ness. Figure 13(a) indicates that the effects of Mach number with roughness applied to the airfoil were similar to those with
