1、NASA TECHNICAL NOTEIZI 1V_ZNASA TN D-8524AERODYNAMIC CHARACTERISTICSOF WING-BODY CONFIGURATIONWITH TWO ADVANCED GENERALAVIATION AIRFOIL SECTIONSAND SIMPLE FLAP SYSTEMSHarry L. Morgan, Jr., and John IV. Paulson, Jr.Langley Research CenterHampton, Ira. 23665NATIONAL AERONAUTICSAND SPACE ADMINISTRATION
2、 WASHINGTOND. C. AUGUST1977Provided 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. Report No. 2. GovernmentAccessionNo. 3. R_ipients _I_ No.NASA TN D-85244.
3、Title and Subtitle 5. Repo_ DateAugust 1977AERODYNAMIC CHARACTERISTICS OF WING-BODY CONFIGURA-TION WITH TWO ADVANCED GENERAL AVIATION AIRFOILSECTIONS AND SIMPLE FLAP SYSTE24S7. Aurar(s)Harry L. Morgan, Jr., and John W. Paulson, Jr.9. Perfuming Or_nization Nameand Addre_NASA Langley Research CenterHa
4、mpton, VA 2366512. SponsoringAgency Name and Addr_sNational Aeronautics and Space AdministrationWashington, DC 20546,15. SupplementaryNote=6. PerformingOrganization Code8. PerformingOrganization Rel:)xtNo.L-1130510. Work Unit No,505-10-1 1-1011. Contract or Grant No.13. Type of Report and Period Cov
5、eredTechnical Note14. Sponsoring Agency Code16. Abstractinvestigation was conducted in the Langley V/STOL tunnel to determine theaerodynamic characteristics of a general aviation wing equipped with NACA 652-415,NASA GA(W)-I, and NASA GA(PC)-I airfoil sections. The NASA GA(W)-I wing wasequipped with
6、plain, split, and slotted partial- and full-span flaps and ailerons.The NASA GA(PC)-I wing was equipped with plain, partial- and full-span flaps.Experimental chordwise static-pressure distribution and wake drag measurements wereobtained for the NASA GA(PC)-I wing at the 22.5-percent spanwise station
7、. Compari-sons were made between the three wing configurations to evaluate the wing perfor-mance, stall, and maximum lift capabilities. The tests were conducted over anangle-of-attack range of -4 to 22 and a Reynolds number range of 1.21 x 106 to1.92 x 106 based on wing chord.The results of this inv
8、estigation indicated that the NASA GA(W)-I wing had ahigher maximum lift capability and almost equivalent drag values compared withboth the NACA 652-415 and NASA GA(PC)-I wings. The NASA GA(W)-I had a maximumlift coefficient of 1.32 with 0 flap deflection, and 1.78 with 41.6 deflectionof the partial
9、-span slotted flap. The effectiveness of the NASA GA(W)-I plainand slotted ailerons with differential deflections were equivalent. The NASAGA(PC)-I wing with full-span flaps deflected 0 for the design climb configura-tion showed improved lift and drag performance over the cruise flap settingof -10 .
10、17, Key Words (Suggestedby Author(s)General aviationAirfoilsWings18. Distribution StatementUnclassified - UnlimitedSubject Category 0219. Security Clar4if.(of thisreport) 20. SecurityClassi.f.(of this page) 21. No. of Pages 22. Price“Unclassified Unclassified 69 $4.50* For sale by the National Techn
11、ical Information Service, Springfield, VJfg_nia 22161Provided 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-,-,-AERODYNAMIC CHARACTERISTICS OF WING-BODY CONFIGURAT
12、ION WITH TWO ADVANCEDGENERAL AVIATION AIRFOIL SECTIONS AND SIMPLE FLAP SYSTEMSHarry L. Morgan, Jr., and John W. Paulson, Jr.Langley Research CenterSUMMARYAn investigation was conducted in the Langley V/STOL tunnel to determinethe aerodynamic characteristics of a general aviation wing equipped with N
13、ACA652-415 , NASA GA(W)-I, and NASA GA(PC)-I airfoil sections. The NASA GA(W)-Iwing was equipped with plain, split, and slotted partial- and full-span flapsand ailerons. The NASA GA(PC)-I wing was equipped with plain, partial- andfull-span flaps. Experimental chordwise static-pressure distribution a
14、nd wakedrag measurements were obtained for the NASA GA(PC)-I wing at the 22.5-percentspanwise station. Comparisons were made between the three wing configurationsto evaluate the wing performance, stall, and maximum lift capabilities. Thetests were conducted over an angle-of-attack range of -4 to 22
15、and a Reynoldsnumber range of 1.21 x 106 to 1.92 x 106 based on wing chord.The results of this investigation indicated that the NASA GA(W)-I wing hada higher maximum lift capability and almost equivalent drag values compared withboth the NACA 652-415 and NASA GA(PC)-I wings. The NASA GA(W)-I wing ha
16、d a max-imum lift coefficient of 1.32 with 0 flap deflection, and 1.78 with 41.6 deflection of the partial-span slotted flap. The effectiveness of the NASAGA(W)-I plain and slotted ailerons with differential deflections were equivalent.The NASA GA(PC)-I wing with full-span flaps deflected 0 for the
17、design climbconfiguration showed improved lift and drag performance over the cruise flapsetting of -10 .INTRODUCTIONResearch on advanced aerodynamic technology airfoils for general aviationapplications has been conducted over the last several years at the LangleyResearch Center and reported in refer
18、ences I to 4. The first of these airfoilswas developed from a 17-percent-thick supercritical airfoil to provide an air-foil with improved low-speed characteristics. This airfoil designated NASAGA(W)-I in reference I showed a 30-percent increase in maximum lift coefficientand more gradual stall chara
19、cteristics than a typical older NACA 65 series air-foil used for comparison.Wings using this improved low-speed section would be suitable for applica-tion to light general aviation aircraft. These aircraft usually have limitedpayload weights because of low-powered engines and generally have poor rid
20、equality because of large wing areas. Application of the improved airfoil sec-tion should increase payload capability because of the lighter wing weightProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-attainable with thicker sections and should improv
21、e ride quality becauseof thesmaller wing areas possible with an increase in lift capability.This investigation was conducted to determine the longitudinal aerodynamiccharacteristics of an aspect-ratio-9 wing with the NASAGA(W)-I airfoil sectionequipped with typical simple flaps and ailerons. This wi
22、ng was attached to arepresentative fuselage shape with a fineness ratio of approximately 8. Anadditional airfoil section, designated NASAGA(PC)-I, which was designed foroptimum drag coefficient at a climb lift coefficient of 0.9 was also tested dur-ing this investigation. This additional wing was eq
23、uipped with a plain flap andwas intended for particular application to single-engine aircraft which, in gen-eral, hav_ poor lift-drag ratios in climb. A wing with a NACA65-415 airfoilsection was also tested to provide baseline comparison data for the other wings.The tests were conducted in the Langl
24、ey V/STOLwind tunnel through an angle-of-attack range of -4 to 22 and a sideslip range of -5 to 5 . Reynolds numberbased on wing chord was also varied from 1.21 106 to 1.92 106. The chord-wise pressure distribution and corresponding wake velocity profile were measuredYat the _ = 0.225 spanwise stati
25、on on the NASAGA(PC)-I wing.b/2SYMBOLSValues are given in both SI and U.S. CustomaryUnits. The measurementsand calculations were madein the U.S. CustomaryUnits. The model force andmomentdata are referred to the stability axis system shownin figure I. Themodel momentreference center was located at th
26、e quarter-chord location of thewing root chord as shownin figure 2.bCDChCLCLeCICmCnCpCycwing span, 4.013 m (13.17 ft)drag coefficient, Drag/qSaileron hinge-moment coefficient,lift coefficient, Lift/q Slift-curve slope per degreerolling-moment coefficient,pitching-moment coefficient,yawing-momentcoef
27、ficient,pressure coefficient, (Ps - P_)/q_side-force coefficient, Side force/q Swing chord, 44.7 c_ (17.6 in.)Hinge moment/qCa2baRolling moment/qSbPitching moment/q_ScYawingmoment/qSbProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-CdCnL/DPsPtp_%RSVXY
28、Zsection profile drag coefficient determined from wake measurements,_01 _Pt- Ps I IPt - P_I (_)2 q_ q_ d (see eq. 24.16, ref. 9)section normal-force coefficient,lift-drag ratiowake rake height, 16.76 cm (6.60 in.)local static pressure, Pa (lbf/ft 2)total pressure, Pa (lbf/ft 2)free-stream static pre
29、ssure, Pa (ibf/ft 2)free-stream dynamic pressure, kPa (lbf/ft 2)Reynolds number based on free-stream conditions and airfoil chordwing area, 1.795 m2 (19.307 ft2)free-stream velocityairfoil abscissa, cm (in.)vertical distance in wake profile, cm (in.)airfoil ordinate, cm (in.)angle of attack, measure
30、d vertically between free stream and fuselagecenter line (positive direction, nose up), degB angle of sideslip, measured laterally between free stream and fuselagecenter line (positive direction, nose left), deg6 control surface deflection, measured vertically between wing chordlineand control surfa
31、ce chordline (positive direction, control surfacedown), degSubscripts:a aileronf flapmax maximum3Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-s statict totalfree-stream conditionsNotation:I lower surfaceu upper surfaceGA(I)-I airfoil designation,
32、General Aviation (Initial of designers name) -Identification number of particular airfoil designMODELSThe configurations tested during this investigation consisted of threeaspect-ratio-9 rectangular wings mounted on a fineness-ratio-8 tailless fuselage.The planform details of the wing and fuselage a
33、re presented in figure 2 and pho-tographs of the model installed in the Langley V/STOL tunnel, in figure 3. Allthe wings had a span of 4.013 m (13.17 ft), a wing chord of 44.7 cm (17.6 in.),and a wing area of 1.795 m2 (19.307 ft2). The first wing had a NACA 652-415 air-foil section; the second, a NA
34、SA GA(W)-I General Aviation (Whitcomb) - NumberOne; and the third, a NASA GA(PC)-I General Aviation (Peterson and Chen) -Number One airfoil section. Plots of these airfoil shapes are presented in fig-ure 4 and their tabulated coordinates, in tables I, II, and III. The NACA652-415 and NASA GA(W)-I wi
35、ngs had a positive 2 incidence at the root with 2washout at the wing tips. The NASA GA(PC)-I wing had 0 incidence of the rootwith 2 washout at the wing tip.The NACA 652-415 airfoil section is a member of the family of low-drag air-foils developed by the NACA and are often referred to as the “laminar
36、 flow“ air-foils. (See ref. 5.) These airfoils have been used successfully on sailplanes;however, on general aviation wings laminar boundary-layer conditions are diffi-cult to maintain because of surface roughness near the leading edge caused eitherby wing fabrication techniques or by insect remains
37、 gathered during flight. TheNACA 652-415 airfoil section has leading-edge flow separation characteristics athigh angles of attack in two dimensions which results in unfavorable wing stallcharacteristics. This airfoil, nevertheless, is used on many current generalaviation aircraft and was tested duri
38、ng this investigation to obtain baselinecomparison data. This wing was not equipped with flaps or ailerons.The NASA GA(W)-I was designed by Richard T. Whitcomb specifically for low-speed application. (See ref. I.) This airfoil section was designed for acruise lift coefficient of 0.4, for a good lift
39、-drag ratio at a climb lift coef-ficient of 1.0, and for a maximum lift coefficient of 2.0. The key design fea-tures of this airfoil are (I) a large upper surface leading-edge radius; (2) anapproximate uniform loading at the cruise lift coefficient; and (3) a bluntProvided by IHSNot for ResaleNo rep
40、roduction or networking permitted without license from IHS-,-,-trailing edge. The large upper surface leading-edge radius was used to attenu-ate the peak negative pressure coefficients and thereby to delay airfoil stallto a high angle of attack. A blunt trailing edge provided the airfoil withapproxi
41、mately equal upper and lower surface slopes to moderate the upper surface-pressure recovery and thus further delay stall. A 17-percent-thick NASAsuper-critical airfoil was used as a starting geometry for the low-speed airfoildesign because the highly aft-cambered supercritical airfoils had indicated
42、 goodlow-speed characteristics. The final low-speed airfoil geometry was obtainedby tailoring the supercritical airfoil geometry until the desired cruise, climb,and maximumlift conditions were satisfied. The computer program of reference 6was used to predict the design and off-design characteristics
43、 of the airfoilduring the tailoring process.The NASAGA(W)-I wing was equipped with full-span plain, slotted, and splitflap systems as shownin figure 5. The chord of both the plain and slotted flapwas 18 percent of the wing chord, and the chord of the split flap was 24.6 per-cent of the wing chord. E
44、ach flap system was divided at the mid-semispan loca-tion to allow for independent movementof the inboard and outboard sections.The inboard section had a range of deflection from 0 to 40 down, and the out-board, a range of deflection from 0 to 10 down. The outboard section of theleft wing panel of t
45、he plain and slotted flap systems could be deflected from30 up to 20 downand was used as a representative aileron. These aileron sec-tions were equipped with a push-rod type hinge-momentgage as shownin figure 6to determine aileron control forces.The NASAGA(PC)-I was designed by John B. Peterson, Jr.
46、, of Langley ResearchCenter and Allen W. Chen, NRC-NASAResident Research Associate, for optimum dragat a climb lift coefficient of 0.9. Details of the design procedure used forthis airfoil are given in the appendix. A suitable airfoil shape for cruiseflight was obtained by deflecting a 19-percent-ch
47、ord simple flap 10 upward(_f = -10 ) with the center of rotation on the lower surface at the 80.8-percent-chord location. A representative landing shape was obtained by deflecting thesimple flap down10 (6f = 10 ) as shownin figure 7. This flap system, likethose on the NASAGA(W)-I wing, was divided a
48、t the mid-semispan location to allowfor independent movementof the inboard and outboard sections. Partial- andfull-span flap combinations with flap deflections from -10 to 10 were tested.The left wing panel was equipped with a chordwise row of surface-pressureorifices at a spanwise location equal to
49、 22.5 percent of the span to determinethe sectional characteristics of the NASAGA(PC)-I airfoil. The pressure ori-fice locations are given in table IV. The pressure data were integrated toobtain the section normal-force coefficients. A wake rake was positioned5.08 cm (2.0 in.) downstreamof the wing trailing edge at t
