1、4,Copy No. _,P.M-No. L8DZ9_WRTLTEDProvided 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-,-,-NACA RM No. LSD29NATIONAL ADVISORY C0_94ITTEE FOR AERONAUTICSRESEARCH
2、MEMORANDUMWIND-T_ INVESTIGAT_ION OF HIGH-LIFt AND STALL-CONTROLDEVICES ON A 37 SWEPYBACK WING OF ASPECT RATIO 6AT HIGH REYNOLDS NUMBERSBy William Koven and Robert R. GrahamSUMMARYResults are presented of an investigation in the Langley 19-footpressure tunnel of the longitudinal characteristics of a
3、semlspan modelwing having 37 sweepback of the leading edge, an aspect ratio of 6,and NACA 641-e12 airfoil section perpendicular to the 27-percent-chordline. Several types of stall-control devices including extenslbleround-nose leading-edge flaps, a leading-edge slat, and a droopedleading edge were i
4、nvestigated_ partial- and full-span trailing-edgesplit and double slotted flaps were also tested. In addition, variouscombinations of the aforementione a critical span of the leading-edge device wasfound, however, below which reductions in maximum lift resulted.Provided by IHSNot for ResaleNo reprod
5、uction or networking permitted without license from IHS-,-,-2 NACARMNo. LSD29The maximumlift coefficient of the plain wing was about 1.27.Maximumlift coefficients of about 1.9 and 2.0 were obtained forcombinations of an outboard half-span leading-edge device with. inboardhalf-span split and double s
6、lotted flaps, respectively. The highestmaximumlift coefficients were obtained with drooped leading edgeplus fence combinations with trailing-edge flaps. An increase intrailing-edge flap span from half to full span did not produce appreciableincreases in maximumlift when the accompanying changes in t
7、rim weretaken into account.INTRODUCTIONNumerous investigations have been devoted to a study of the low-speed longitudinal characteristics of swept wings. (For example, seereferences i to 3-) As indicated by these studies, two of the majordifficulties associated with sweptback wings are low values of
8、 maximumlift coefficient comparedwith unswept wings and instability at thestall due to tip stalling.As far as maximumlift is concerned, the available data are confinedmainly to investigations of plain wings and wings with split flaps.Even with split flaps, the maximumlift coefficients have been rela
9、tivelylow and it is indicated that investigation of additional high-lift devicessuch as a double slotted flap would be desirable.One method of eliminating tip stalling which has been used suc-cessfully (reference 4) involves the use of a leading-edge device locatedon the outboard sections of the win
10、g span. Several types of leading-ed_devices have been tried, that is, extensible round-nose leading-edgeflap, leading-edge slat, and so forth; but no direct comparison to assistin the selection of the most satisfactory device has been made.With the above considerations in mind, an investigation has
11、beenconducted in the Langley 19-foot pressure tunnel on a wing having 37sweepback of the leading edge and an aspect ratio 6. It should bepointed out that the wing plan-formvariables were such that, accordingto the stability boundary presented in reference l, tip stalling andinstability at the stall
12、would be expected. In addition to the basicwing characteristics at high Reynolds number, the investigation wasconcerned mainly with (a) the effectiveness of double slotted flapsand split f.laps, (b) whether a leading-edge device would eliminate tipstall on the particular plan form used, (c) the dete
13、rmination of therelative merits of several types of leading-edge devices, and (d) themagnitude of maximumlift coefficients and the type of stall associatedwith various combinations of leading- and trailing-edge devices.Provided by IHSNot for ResaleNo reproduction or networking permitted without lice
14、nse from IHS-,-,-NACARMNo. L8D29 3The semispan reflection-plane model was equipped with three typesof leadlng-edge or stall-control devices, namely, a round-nose extensibleleadlng-edge flap, a leading-edge slat, and a drooped leading edge. Inaddition, the wing was provided with partial- and full-spa
15、n split anddouble slotted flaps. Additional devices, such as a fence and outboardpitch flaps, were also investigated. The model configurations weretested alone and in combination through a large angle-of-attack rangeat Reynolds numbers varying from 2.00 106 to 9-B_ lO6. Lift,drag, and pitchlng-momen
16、t data and stall studies are given for someofthe more important configurations.COEFFICIENTS AND SYMBOLSThe data are referred to the wind axes with the origin at thequarter chord of the mean aerod_vnamlcchord. The data have beenreduced to standard NACAnondimensional coefficients which are definedas f
17、ollows:CLC mR_maxZlift coefficient _qL-s_maximnmlift coefficientdrag coefficient D_-S_pitching-moment coefficient _q“-_Reynolds number _stream Mach numberangle of attack of root chord line, degreesangle of attack at CLmaxlift-curve slope _dCL_downwash angle, degreesvertical distance above chord plan
18、e extendedL liftProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-h NACARMNo. LSD29D dragM pitching momentabout 0.256S wing areab wing spanbf flap spanmean aerodynamic chord c2c local wing chord parallel to plane of symmetryy lateral coordinatelateral
19、coordinate of centroid of liftq dynamic pressure _qt dynamic pressure at tailV free-stream velocitycoefficient of viscosityp density of air5 flap deflectionSubscripts :n nosea aileronMODEL AND TESTSMODELThe model used in the investigation was a semlspan wing mounted ona reflection plane and single s
20、trut as shown in figure 1. It was ofsteel construction and had an aspect ratio of 6, a taper ratio of 0.50,and 37 sweepback of the leading edge. The airfoil section perpendicularProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACARMNo. LSD29 5to the
21、27-percent-chord line was the NACA641-212 profile. The generalplan form and some of the principal dimensions of the model are given infigure 2(a).Details of the geometry of the various stall-control devices areshown in figures 2(b) to 2(d). The drooped leading edge (which could bedeflected to three
22、positions) and the leading-edge slat covered halfthe wing semispan extending from 0.45 b to 0.95 b. The round-nose extensibleleading-edge flap, on the other hand, was constructed so that severalg- E_.flap spans could be investigated at one deflection. The leading-edgeflap was of constant chord, wher
23、eas both the slat and the drooped leadingedge were of constant percent chord.The model was so constructed that when the leading edge was drooped,the slat was in the retracted position. Thus, slight discontinuities incontour existed at O.14c and 0.02c of the upper and lower surfaces ofthe wing, respe
24、ctively, for the drooped leading-edge configurations.No such discontinuities were present, however, on configurations withoutstall-control devices or configurations with leading-edge flap where adifferent leading edge was used.The stall-control fence is shown in figures 2(e) and 2(f). Thefence was l
25、ocated at 0.50_, had a constant height of 0.60 th9 maximumthickness of the wing at that spanwise location, and extended over thechord as indicated on the figure.The model was equipped with two types of trailing-edge flaps,namely, split and double slotted, both of which could be tested half andfull s
26、pan. The design parameters for the double slotted flap werechosen on the basis of two-dimensional wlnd-tunnel data given inreference 5. A schematic drawing showing the design details of theseflaps is presented in figures 2(g) and 2(h).Photographs of the model and reflection plane mounted in the tunn
27、eland of the various stall-control devices installed on the model arepresented as figure 3.For model configurations with leading-edge roughness, No. 60(O.Oll-inch diameter) carborundum particles were applied by means ofa thin coat of shellac to the forward 8- and 2-percent of the wing_pper and lower
28、 surfaces, respectively. Roughness for the slat-extendedconfiguration was applied in the same manner to the leading edge of theslat and to the leading edge of the inboard sections of the wing.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 NACARMNo
29、. LSD29TESTSTests were made in the Langley 19-foot pressure tunnel with theair compressed to approximately 33 pounds per square inch. In orderto cover as wide a range of Reynolds numbers as possible, severaltests were made at atmospheric pressure. The Reynolds numbers andtheir corresponding Mach num
30、bers obtained in this investi6ation areas follows :R %2.00 x 106 0.083.00 .124.36 .085.30 .106.8o .138.1o .159.35 .18Lift, drag, and pitching moment were measured through an angle-of-attack range extending well beyond maximum llft. In addition, stallstudies of some of the more interesting configurat
31、ions were made byvisual observation and from motlon-picture records of the behavior ofwool tufts attached to the upper surface of the wing. The majorityof the tests and the stall studies were conducted at a Reynolds numberof about 6,800,000. Downwash and dynamlc-pressure surveys were madebehind the
32、wing for the slat and half-span double-slotted-flapconfiguration.CORRECTIONS TO DATAThe lift, drag, and pitching-moment data presented herein havebeen corrected for air-streammisalinement but have not been correctedfor support tare and interference effects. Previous experience oncomplete models indi
33、cates that corrections for the effects of the tareand interference caused by the model supports consist of (a) a constantshift in the pitching-moment curve (about -0.008), (b) a slight increasein lift-curve slope (about 0.0008), and (c) a decrease in drag in thelow lift range.Jet-boundary correction
34、s obtained by combining the methods ofreferences 6 and 7 were made to the angle of attack and to the dragcoefficient and are as follows:= 1.12C LACD = 0.0164CL2Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACARMNo. LSD29 7The correction to the pit
35、ching-moment coefficient caused by th_ tunnel-induced distortion of the loading isACm = O.0101CLAll corrections were added to the data.RESULTS AND DISCUSSIONThe results of the investigation of the plain wing and wing withtrailing-edge flaps are presented in figures 4 to 7. Figures 8 to 12show the ef
36、fect of leading-edge devices, and figures 13 to 22 show theeffect of various combinations of leading-edge and trailing-edge devices.Several additional tests were made to determine the effect of varyingthe leadingTedge flap span; only the maximumliftand pitching-momentcharacteristics of these configu
37、rations are presented (fig. 16). Thespanwise location of the centroid of lift is presented for severalconfigurations in figure 23. A summaryof the more important resultsof the investigation is presented as table I.PLAIN WINGANDHIGH-LII_f DEVICESLift CharacteristicsThe data for the plain wing and win
38、g with split and double slottedflaps are presented in figures h to 7. The lift curves for all conditionswere relatively linear up to maximumlift except for a slight roundingat high angles of attack. In all cases the maximumlift coefficientand angle of attack at maximumlift were very well defined ind
39、icatinga rather sudden breakdown of the flow at the critical angle.Lift-curve slope.- The lift-curve slope was calculated from two-dimensional data using the method suggested in reference 8 where theaspect ratio is based on the true length of the quarter-chord line.The lift-curve slope was also obta
40、ined from the charts of reference 9which assume a section lift-curve slope of 2_. The two methodspredicted values of lift-curve slope of 0.071 and 0.066, respectively,as compared with the value of 0.070 obtained experimentally.Effect of flap deflection.- Increments in lift at zero angle ofattach and
41、 at maximum lift are presented in figure 5 as a function offlap span. The data for the half- and full-span flaps were taken fromfigure 4; in order to obtain more complete data on the effects of flapProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 NA
42、CARMNo. LSD29spar_someadditional tests on intermediate split-flap spans were made.Only the lift increments for the supplementary tests are presented.An attempt-was made to estimate the increments in lift at zeroangle of attack from two-dimensional data utilizing amethod for unsweptwings outlined in
43、reference lO. The equation was modified and sweeptaken into account as follows:ACL = J this abnormalloss of outboard flap effectiveness maybe typical of split flaps onsweptback wings.The data for the double-slotted-flap configurations are considerablydifferent from those for the split flap. The doub
44、le slotted flapproduced larger increments in lift throughout the flap-span range thanthe theory predicted, and the outboard span did not lose its effectivenessbeyond what might be expected from the simplified theory.The reason for this difference between the split and double slottedflap is not appar
45、ent. The effects of sweepback on the variation withflap span of the increment in lift due to flap deflection appear to bedependent on the type of flap under consideration.Figure 5 also shows that the increments in lift at maximumlift areconsiderably less than at zero angle of attack. The magnitude o
46、f thiseffect, however, appears to be of the sameorder as on unswept wings ofsimilar airfoil section. (See reference ll.)Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACARMNo. L8D99 9Maximum lift.- As far as the maximum lifts are concerned, they ca
47、n bestbe summarized in the following table. The values of CLmax listed beloware untrimmed values:FlapNoDe0.5b split0.5b double1.0_ split1.0 _- double2i.97i.551.g21.652.32Pitching-Moment CharacteristicsExcept for the full-span double-slotted-flap condition, the pitching-moment curves were fairly line
48、ar, and for the most part, parallel to oneanother (fig. 4). In all cases the moment at the stall broke in anunstable direction, that is, in a nose-up direction.The trim changes brought about as a result of flap deflection are ofspecial interest. A comparison of the data from figure 4 with similardata from reference ll shows that the full-span split and double slottedflaps produced changes in trim which were of the same magnitude as on anunswept wing with approximately the same airfoil section. The semispanflaps, however, produced considerably smaller trim changes than was notedon the unswept
copyright@ 2008-2019 麦多课文库（www.mydoc123.com）网站版权所有
备案/许可证编号：苏ICP备17064731号-1