1、_:j),- 0.30-AIEUK)IL-CHORDBEVEIED-TRAILIIVGEZ)GEFLAPSBy H. PageHoggard,Jr.,andMarorieE. BullmhIangleyMemorialAeronauticalLaboratoryeY Field,Va. . FORREFERENCE.,.?+ !$.,.liO TO BE ZAKENFROIfTF$13ROOMTNqK-2i-N A CA LIBRARYWMHINGTON LANGLEYMEMORIALAERONAUIIwME4XLKTORY9 ey Fieldj Va.NACA WARTIME REPORTS
2、 are reprints f papers orlgimdly issued to provi rapid distribution ofadvance research results to an authorized group rq the results of these tests arepresented in reference 3 (also summarized in reference 1).From the results of these force tests of trailing-edgeshapes having various included ,trail
3、ing+dge angles andother airfoil tests, a method based on the included angleat the trailing edge has been found for predicting thevalues of hinge-mom?nt parameters to be expected from abevel. This correlation can be found in figwe 150 ofreference 1.The two-dimensional-flew tests presented herein were
4、made to investigate the pressure acting on a control sur-face with a beveled trailing edge. Such data should bevaluable for structural design of the control surfaces,for explanation of the balancing action of the bevel,and for study of boundary-layer conditions. The investi-gation was made at all an
5、gles of attack and flap deflec-tions considered necessary for the structural design ofailerons, elevators, and rudders.SYMBOLSCf flap chord rearward of flap hinge axis, percentairfoil chordc chord of basic airoil with flap neutral2 dynamic prssure of free air streamP pressure coefficient.Provided by
6、 IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA ARIR?0 L!03 3.PR.APRPPoa.6fMCn%CntchfnmnfhfCna.resultant pressure coefficientincrement of resultant pressure coefficientstatic pressure at a point on airfoilstatic pressure in free air stream ,angle of attack
7、 for infinite aspect ratioflap deflectionMach number, ratio of local velocity to speed ofSoundairfoil sectionairfoil sectionnormal-force coefficient (n/qc)pitching-moment coefficient .about quarter-chord point of airfoil (m/qc.? ;,. :.: :4. - -.; ,. , :., -: ; ,.: ,“”. : .: : .“ . . ,“Provided by IH
8、SNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-,4()bChfChfa = 0 6fThe subscripts outside the parentheses indicate thefactors held constant during the measurement of theparameter.Subscripts:u point on upper surfaceL , point on lower surfaceR resultantAPPARATUS AND
9、 KODELSThe tests were made in the NACA 4-”b 6-foot verticaltunne1. The test section of this tunnel has been con-verted from the original open, circular, 5-foot-diameterjet (reference .4.) to a closed rectangular 4.-by 6-foot .test section, .as.shown in figure 1. The model completelyspanned the test
10、section; therefore, two-dimensional flowwas approximated-. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA ARR 0.4D03 5.,.The znodelused for the pressure-distribution testsof this investigation was designed to be an exact copy ofthe model used f
11、or the force tests in reference 3 butwith only the 0.15cf and 0.20cf beveled-trailing-edgeshapes. The 0.15c bevel was tested with the bevelcorner faired with both large and small radii. The2-foot-chord model was made of linated mahogany to themodified NACA 0009 profile (table I). The airfoil wasequi
12、pped with a 0.30c plain flp, as shown in figure 2(a).A gap of 0.005c was pr,ovidedat the flap nose. The flapwas constructed with interchangeable blocks that formed a -beveled trailing edge and a thickened profile, as shownin figure 3 of reference 3.A single chordwiserow of pressure orifices wasbuilt
13、 into the upper and lower surfaces of the airfoiland flap at the midspan location. The orifice loca-tions are presented in figure 2(b) in percent of airfoilchord from the leading.edge. The copper tubes from thepressure orifices were brought out of the model at oneend through the torque tube and the
14、tunnel wall to amultiple-tube, open-faced manometer. Readings wererecorded by a camera.TESTSAll of the tests, except those with large flap de-flection and high positive angle of attack (flap defect-ion, 30 and 45; angle of attackt 143and 19.3) wererun at an average damic presmre of 15 pounds persqua
15、re foot. The large flap deflections at high positiveangles of attack required more power than was available tomaintain a dynamic pressure of 15 pounds per square foot;therefore, these tests were run at an average dynamicpressure of 12 pounds per square foot. The airspeed inthe test section at dynami
16、c pressures of 15 and 12 poundsper square foot is about 76 and 69 mi,lesper hour, respec-tively; at standard sea-level conditions. The corre-sponding values of ffective Reynolds number are2,760,000and 2,208,000. (Effective Reynolds num-ber = Test Reynolds number X .Turbence factor; the tbu-lence fac
17、tor of the NACA 4- by 6-foot vertical tunnel is1.93.) “The tests were made at aiiglesof attack ranging from-20 to 20 at intervals of “ sad at angles giving maxi-mum positive and negative lift. It may be noted that alliIIIIProvided by IHSNot for ResaleNo reproduction or networking permitted without l
18、icense from IHS-,-,-II3iI,.,III6 NACA Ml? Nc),I203angles of attack are offset from the exact values of00, 50; 10o, 150, and 20 by-0.7 oioeorin setting the zero angle ofattack.found to be consistent throughout thg tests and the datawere corrected accordingly. The model was tested withthe 0.30c plain
19、flap deflected 0,10, 2, 5, 10,150, 200, 250, 300, and 450. The tests were run withthe flap gap both open (0.005c”gap) and sealed withplasticize. During the tests with 30 and 45 flapdeflection, pressure orifice 15 for the lower surface(fig. 2(b) was sealed because its position at bothlarge flap defle
20、ctions was inside the gap.Check tests were made for.each flap deflection asan indication of the accuracy of the test results.When the 0.005c gap was used, the check tests weemadeafter both angle of attack and flap deflection had beenreset. The sealed-gap check tests had only the angleof attack reset
21、$ because the plasticize seal would haveto be refaired if the flap deflection were changed.I!,.-.a7, ,: - $. -,:J:. e L- . . . , .,., , ; . “-:.: .). - .,.; ; -+. ., .:-. 1.,. :,+.,-,.: ,. ;: ,-., - .- J,-.-: . . . ,-: .*:,. *. :. ,. . ,: .; :-, ,. f. :.,- .-, ,Provided by IHSNot for ResaleNo reprod
22、uction or networking permitted without license from IHS-,-,-.- -. 4.8 NACA ARR Ifoo 4D03the various bevel and gap combinations in figures 15to 18 for convenience in calculating distributions atsmall values of =0 and df. l!heflap section normal-force coefficient as a function of flap deflection ispre
23、sented tor all combinations of bevel and gap infigures 19 and 20 at several angles of attack. conl-plete chordwise pressure distributions for variouscombinations of a. and df that might occur on thehorizontal tail of a dive bomber in highly acceleratedmaneuvers at various speeds are presented in fig
24、ure 21for the 0.15cf-bevel flap with sealed gap.The section aerodpmmic coefficients of the airfoiland flap are presented as functions of angle of attackfor all bevel and gap combinations in figures 22 to .The coefficients were obtained in each case by mechanicalintegration of the original pressure d
25、iagrams.The parmeter values for beveled flaps are pre-sented in table II along with values for the plain-airj?oil-contourflap for convenient comparison. Theplain-flap parameter values were obtained from refer-ences 1 and 6,PrectsionThe angles of attack are believed accurate within*O.1O.Flap deflecti
26、ons are believed accurate within*0.20. Plotted values ofpressure coefficient P arecorrect within :2 percent except for peaks at theleading edge and flap hiage axis or for stalled con-ditions. .,.Coefficient values talc-dated from check test pointshave been plotted in figures 19 and 22 and are design
27、atedby flagged symbols. Many of the points come within theaccuracy of the plot; others vary a negligible amount.The accuracy of the corrected zero angle of attack isindicated by the deviation from zeio of lift and momentcoefficients at zero angle of attack. From figures 19and 22, it appears that the
28、 maximum error in setting theangle of attack at zero lift is 0.2. This discrepancymay be caused by flow misalinement in the tunnel or by anasymmetrical model.Two-dimensional flow having been approximated, theresults may be considered as section characteristics.,.Provided by IHSNot for ResaleNo repro
29、duction or networking permitted without license from IHS-,-,-1-.NACA ARR No. 14.D03 9Experimental tunnel corrections were applied onlj to theairfoil section normal-force coefficient Cn. Althoughno corrections were made for the other coefficients, thetunnel values are believed to be higher than the f
30、ree-airvalues “andhence are on the conservative side for struc-tural purposes. The magnitude of the airfoil resultantpressure coefficients as represented in the resultant-pressure diagrams (Tigs. 3 to 10) is own to be too largeby about 7 percent because these curves were plotteddirectly from manomet
31、er records without the applicationof the experimental tunnel correction, which allows forthe increase in lift produced by tunnel-wall interference.DISCUSSIONResultant-Pressure DistributionThe resultant-pressure diagrams should prove usefulin determining loading conditions for the structuraldesign of
32、 ailerons and horizontal and vertical controlsurfaces. Tests have indicated that the increments ofpressure and the increments of section aerodynamiccoefficients caused by flap deflection are approximatelyindependent of the airfoil section for airfoils ofapproxtiately the same maximum thickness and t
33、hicknessdistribution (references 7 and 8). It is thereforebelieved that, for structural design, the incrementaldata presented herein may be applied to other basicsections of approximately the same thickness and thick-ness distribution. The increments of the section aero-dynamic coefficients may be t
34、aken from figures 22 to 2.$.by using the flap-neutral curve as a reference line.From a study of the incremental-resultant-pressurecurves for the stalled conditions (a. = 19.30 d -20.70)for both bevel chords and gap conditions (figs. 4, 6, 8,and 10), it appears that the bevel continues to reducethe f
35、lap hinge moment in the stalled condition from thehinge moment for a plain flap under the same conditions.The tests of beveled elevators on the fuselage of atypical pursuit airplane also indicated that the bevelwas effective in the stalled attitude and reduced thefloating angle of the elevators by a
36、bout 10 (reference 9)from the angle at which airfoil-contour elevators wouldfloat. The resultant-pressure curves (figs. 3 to 10),especially for the 0.005c gap, show a tendency toward adecrease of resultant pressure over the main airfoil justahead of the flap.Provided by IHSNot for ResaleNo reproduct
37、ion or networking permitted without license from IHS-,-,- .- -_+10 NACA”ARR No. L4D03The results indicate that the size of the radius atthe bevel juncture is relatively unimportant in itseffect on the loads over a Develed-trailing-edge flap(fig. 11).Pressure Distribution over Upper snd LowerSurfaces
38、 of Beveled Flap.The distributions presented at various angles ofattack and flap deflections”in figure 12 indicate thatonly on the surtace of the fiap which is deflectedagainst the relative wind does the bevel affect thepressure distribution to any great extent. The onlyexceptions occur .atlow angle
39、s of attack and small flapdeflections, for which the upper- and lower-surfacedistributions show nearly equal effeetof bevel. Thepressure distribution on the side away from the rela-tive wind, when at large angles of attack or flap deflec-tion, Vesembles that.of a flap and tab in a stalledcondition.I
40、t will be noticed in figure 13 that the resultant-pressure peak at the flap hinge sxis is higher for the .beveled flap with the 0.005c gap than for the beveled.flap with the sealed gap. Inasmuch as the resultantpressure is the algebraic difference of the upper- and lower-surface pressures at any poi
41、nt; the positive peakon the lower surface makes the resultant-pressure peakhigher.(See fig. 14. )The pressure distribution produced over the upperand lower surfaces of a flap by a beveled trailing edgeis compared with the pressures over a plain flap infigure 4. The effect on the pressure distributio
42、n of “the bevel on the surface deflected against the relativewind is more pronounced when the gap is open. The maineffect of the open gap on the flap pressure distributionappears to be the decrease in magnitude of the negativepressures over the upper surface of the flap, whichresults in a tendency t
43、oward lower or even overbalancedhinge moments. .Curves of Pa and PFor convenience in calculat-ingthe pressure distri-butions over both surfaces for small values of aProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACilARE?NO. L4D03 11.and 6f the curv
44、es of Pa and P were calculated andare presented in figures 15 to18. From the experi-mental data, it was found impossible to predict with anydegree of accuraqy the variation of pressure with angleof attack over the nose of the airfoil because thestagnation point moves considerably and the pressuresch
45、ange rapidly aridare not linear with angle of attack.The variation of pressure with mgle of attack over therest of the airfoil appeare,dfrom these tests to remaina linear variation only from O0 to 50; therefore, thepa-curves should not be used for calculating pressuresbeyond a value of a. of 50.The
46、variation of pressure with flap deflection forany point on the airfoil contour appeared from thesetests to be linear to 5. The PG-curves thereforeshould not be used.for flap deflection greater than 5.The final pressure distribution required is found bymultiplying the values of Pa and P by the values
47、of a. and af for which the distribution is desiredand adding algebraically to the basic distribution(P at a. = Gf = Oo) given in the lower part of fig-ures 15 to 18.Flap Section lTornal-ForceCoefficientFor all combinations of bevel snd gap tested, thevalues of cnf were smaller than for the plain fla
48、p withsealed gap at the same angles of attack. The valuesof cnf d cnf for beveled and plain flap may beu.convenintly compared in table II. The variationOf Cnf as a function of angle of attack is clearlyshowm in figures 19 and 20. The effect of a is smallat Gf = 28o with the gap open and at Gf = 200 withthe Gap sealed.Pressure Distribution on Horizontal Tail ForHighly Accelerated ManeuversThe flight condition during which high structuralloads and the formation of a compression shock on thehorizontal tail are most likely to occur
