1、f “. :- . d RESEARCH MEMORANDUM EFFECT OF LEADING-EDGE HIGH-LIFT DEVICES AND SPLIT FLAPS ON THE M“-LIFT AND LATERAL CHARACTERISTICS OF A RECTANGULAR WING OF ASPECT RATIO 3.4 WITK CEULA=ARC -OIL SECTIONS AT REYNOLDS NUMBERS FROM 2.8 X la6 TO 8.4 X lo6 Roy H. Laage and Ralph W. May, Jr. Langley Aerona
2、utical Laboratory NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON UNCMsS!FIEc November 10, 1948 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- n t NATIONAL ADTISORY CO- FOR AERORATICS The results of an investition at high Reynolds numbers an
3、d low Mach nunibere in the Langley full-scale tunnel to determine the effect1 of leatiing-eclge high-3ift devices and spUt flaps cm the mximuzn-lift and lateral characteristic8 of a rectengular wing of aspect ratio 3.4 with circular-arc airfoil sectians are presented in this,report. The Fnvestigatio
4、n included measurements of the aerodynamic characteristicf in pitch and in yaw of the basic wing and of the wFng wfth several leading-edge high-lift devices and. 0.20-chord aplit flaps deflected alone and in cardbination with me another. Scale effects were inveeti- gated at Regnolh nders ranging fra
5、m 2.9 x 106 to 8.4 x 106. In addition to the force mBasuremnts, the stug characteristics of the wing were determined. The nmxinnun lift coefficient of the basic w5ng is 0.58. me addition. of -span =-span split fags aflected 60 increases this value to 1.00 and 1.24, respectively. The ElgMemsnt betwee
6、n the experimsntal values of the muclmum Iff t coefficient and Uft-curve slope of the basic wing and the increments in lift coeff icients due to flap deflectian and those calculated by the beet availabze methods is good. Maximum lift coefficients of 0.89, 1.20, and 1.21 me obtained for the wing with
7、 the drooped-nose flap deflected 20, uith the extensible leadin;-edge flap, and with the conibkation of drooped-nose flap deflected 10 with 0.032-chord rohe wing is high throughout the moderate to high angle-of -attack range. The addition of split flaps causes a kge drag increase; howenr, an appreci
8、able reduction in the drag in this range is obtainea by deflect- either the drooped-nose flag or by the inatallaticm of the extensible leading-edge flap. The pitching-moment characteristice of the baeic wing asd of the win: with the leading-edge high-lift devices givFng highest maxfmum lift indicate
9、 that below the stall the center-of-pressure locatian is slightly forward of the quarter chord. A stable pitching-moment break is shown at the stall for all configurations except those with the extensible Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,
10、-,-2 NACA RM no. -30 leading-edge flap and wlth the cmbination of the drooped-nose flap 1 deflected XIo with the 0.032-chord round leadlug edge, which have nrarginal stability. In general, the addif;ian af split fhp8 to CanfigUratimB cawes a slightly rearward shift of the cfmbr-of-preseum 10Catia;l.
11、 For the basic wlng the dihedral effect increases parabolically with lift coefficient-and the directid stability lncreaaes essentially linearly with lift coefficient and the respective paramsters attain valuee of 0.0023 per degree and -0.00050 per degree near maxh lfft. Values of the side-force para
12、meter are low. All the leading-edge high-lirt device8 investigated an this wFng with circular-arc section produce almost lFnsar dihedral-effect variations with lift coefficfent, which Is consietent with the characteristics of conventional blunt-nose aixfoila and xith theory; the directional stabilit
13、y and lateral-force oharacterietics are not materially affected. The split flaps decrease the dihedral effect of the basic wing at a given lift coefficient, but they generally do not materially affect the lateral characteristics of the wbg when imtalled in combinatian with the lead-edge high-lift de
14、vices. 3 In order to provlde large-scale data on the high-angle-of-attack I CharaCteri8tiCe of wing8 having airfoll sections xith eharp Leading edges, an investigation is being conducted in the Langley full-scale turmel at high Reynolds numbers- and low Mach nbem of several typical transanic and sup
15、ersonic swept and uncrwept wing plan fom havFng 10-percent-thick, circular-arc airfoil sections. One of the win ixW0S- tigated wa a trapezoidal win; of aspect ratio 4, and the maxllmun-lift and staUrr characteristics have been reported in ref ereace 1. The results of reference 1 show that the inhere
16、ntly low raximum lift and hi critical preesure coefficient; pressure coefficient at a local Mach nWer of 1.00 free-stream Mach mber pitching mcnnent Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-aspect ratio (9 dietance along semispan Tram phe of s
17、ymmetry taper ratio; ratio on the lift, the drag, and the pi ching-mamsnt coefficients in the Reynold8 number range fram about 3.27 x 18 to 7.67 x 106. Due to structural limi- tation, the extemible leadlng-edge flap could not be teated at a tunnel airspeed hlgher than that corresponding to a Reynold
18、s nmiber of about 3.90 X ld. d - “he stLUing chmacteristics mre detarmFned by obeerving the action of wool tufts attached to the upper wing Mace. These tuft studies ware made of the basic wing and of the wing with the more effective high-Ut arrangements. The tuft studies were made at a Reynoldfl nWe
19、r of about 4.1 x 10 for the %Trig with the sharp leading edge and at about 4.1 x 1ra of several wing-flap configurations. 121e critical compressibflity speed of the Xing with the 0.032 round le- edge installed, with the drooped-nose flag deflected loo, and with ha-span split flaps installed is given
20、 in figure 24. Force measuremsnts. - The mzfmum Ut caefficient of the basic wing is 0.58 at an =-Eiige High-Lift Devices Drooped-nose flap. - The maxFmum lfft coefficient of the whg with the drooped-nos8 flap deflected 20 is 0.89. (See fig. 6.) This value is 0.31 higher than that obtained for the ba
21、sic wing. Although a maxi- ruum lift coefficient of 0.92 is obtalsed witk -e drooped-nom fhg deflected eo, it is cbtained at a higher angle of attack and Wfth consid- erably more drag than for the case with 6, = 20. (See figs. 6 and 17.) The increases in maxfmum lift coefficient and angle .of attack
22、 for maximum Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 NACA RM No. L8D30 lift with the drooped-nose flap deflected reault prfmarily from the improved flow canditiona at the lead- edge by more nearly dining the wing contour with the air strea
23、m and themby delaying stall to higher angles of attack. This alinedt of the leading edge tea to alleviate the negative pressure peaks .and tbsreby to decmaae the however; the gliding weed of about 145 miles per hour to obtain a einking epeed of 25 feet per second is above the range of present practi
24、ce. It should be realized that the drag coefficients plotted in figure 23 are for the wing done and, there- * fore, the suing speeds of the complete airplane would be somewhat greatsr. Power could b4 uaed for the landhg approach and landing condi- tims to offset the high drags shd fi PiWe-23, but tM
25、s practice c, for emergency Unga with power off. .I . “ The pitching-moment curves show no significant change in the longi- tudinal stability of the wing aa compared with. the basic wing (fig. 6). A smaller change in trim due to drooped-nose-flap deflection is noted than waa rnoasured with the split
26、 flaps .deflected. (See fig. 4.) Extensible leading-edge flap. - The I21he favorable effects of the round leading edge rather than to the effect of leading-edge3-fla-p defbctiun. TW section data (reference 2) show that the increase in maximm lift coefficient mer that for the basic wing du9 to a 0.10
27、 extensible leading-edge flap vlth sharp leading edge8 iR only about one-half the magnitude of that obtained wit21 the droopcd-n3sc, flap deflected. The higher 8lope.s of the lift curves y near the stall may lead to serious rolling instability. The mxirmxm lift coefficient increases wfth increashg R
28、eynold8 number for the range OP egnold nunibera investigated. (see figs. 7(a) and =(a).) Tuft studies of the King (fig. 14(e) show early separation over the nose flap, but as conpared .with the wing xith Sn = 200 (fig. 14(b) the root staa at the wing traiUng e- haa been delayed to higher angles of a
29、ttack and covers a smaUer area of the wing near maximum lift. The extensible leading-edge flap cauees a cunsiderable reduction In drag of the wing at the higher angles of attack as ccrmpared with the basic wing or the wlng with 8, = 20. As shown in figure 23 (a), the gliding Geed required to maintai
30、n a E- speed of 25 feet -per second is reduced from 145 miles per hour to about 117 miles per hour. The variations of khe pitching-momsnt coefficient with lift coef- ficient (fig. 7(b) ) show a center-of-pressure locatim which is slightly ahead of that for both the basfc wing and the wing with 6, =
31、20 through- out the lift-coefficient range. The stability a% the stall is not appre- ciabl;y affected br Fncreaaing Reynolds nmber. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 1- . TTACA RM No. L8D30 The tuft studies of the wing with the En= .
32、loo, 0,032 round- . leading-edge configuration at a Regno% nlllILber of about 4.1 x 10 (fig. 14(d) ) ehow a stall progrwssion which reaenfblee that obtained fur the wing with 6, = 20 (fig. 14(b) ) except that the lnitlal leading- i edge separatim is confined to a region at about 0.5 - and that a sma
33、ller area of the wing is stalled in the region of h. The tuft studiea at a Reynolds mmib0r of about 6.0 x 10 show no chanp in the stall progress.ion and consequently are not pre8ente.d. 6 b 2 6 Inspection of the gUQing-speed a;na ai8peed chart (fig. 23(d shows slightly lower drag coefficients in the
34、 moderate and high lift- coefficient range for the En = loo, 0.032 round-bading-edge conffgura- tion than for the extemible leading-edge flap; however, in the region of Ch the gliaFng speed reluired to maintain a einklng speed of 25 feet per second is about the same (U.5 milee per hour) as for the e
35、xtern ible leading-edge flap. I.“ _- As in the csse of the extsnsfble leadine;-e-moment characterietics of adding the half-ep- split flap (fig. 19(b) ) is to give slightly mre stabili.t;y below ths stall but considerablg leas stability at than for the 6, = loo round-leading-edge ccmfiguration. The d
36、es-kbilizing effect at high lift coefficients becamea mre pronounced with increases in Reynolds numbsr BO 6 that at a Reynolds nmiker of 8.0 x 10 the wLng fa unstable at stall. tuft etudies of f igwe U(c) (taken at a Reynolds number af about 4.1 x %6 10 where the pitchLng-moment break is stable at t
37、he atall) ehar that the Bplit t flap reduces the trailing-edge sepazation that occurred for ths wing with tha split flap removed (fig. 14(d). As for the other cdbinati canfig- uratiom already discussed, the half-span split flap increaseer the drag of the 6, = loo, 0.032 round 1eading-e- caution BO t
38、hat a minimum sinking speed af 31 feet per second is obtained at a gliding sped of 113 miles per hour (fig. 23 (b) ) . k Th3 0.032 round leadlng edge with 6, = 0. - With the 0.032 round- leading-edge configuration with halp-span lit flaps installed,a %= of 1.25 Is obtained (fig. X). This value of is
39、 0.37 higher than that for the round leading edge alone but is slightly lower than that for the drooped-nose, half-span, split-flap configpation. Th acale effect on maximum lift ia similar to that for the other round-leading-edge config- uratiane previously diecussed (see fig. in tha a marked increa
40、se in occur8 between ReynolrLa nmbsrs of 4.2 x 10 k and 6.2 x 106. Gener- ally marginal stability is shown at low and moderate lift coefficlenta and the etabilizing trend at Mgh lift ccefficfents occurs at Increasing values of CL as the Reynolds number is increased. V I Provided by IHSNot for Resale
41、No reproduction or networking permitted without license from IHS-,-,-MACA RM No. L8D3O Leading-Edge Pressure Measurements 15 The maximum observed negative pressure coefficient for the canfig- uration of 6n = loo with the 0.032 round leading edge and the half- span split flap hat-ls nunibsr. %CL Exte
42、nsible Leadhg-Edge Flap The dihedral effect with the extensible leading-edge flap Fnetalled has a fairly normal variation with ltft cosfficient (fig. 34) and an Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-18 NACA RM No. -30 average C slope that a
43、grees well with theory. Actually, however, 2*cL 1 the C slope is sur at moderate lift coefficients and grea-r than the calculated value of 0.0023 at low and high lift coefficients. The directiansl stability is somewhat different than for any other cdlg- uration tested in that it is zero at a Qf 0.25
44、 and increases rapidly for both higher and lower lift coefficien 8. Of course, this difference at low lift coefficients is relatively uninTport;ant bscauee the flap would probably not be deflected at such lift coefficient8 corresponding to high flight speeds. The %$ variation is quite similar to tha
45、t for the drooped-nose configuraticm with the 0.032 round leehding edge installed in that it Fncreaees about 1Fnesrl.y to -0.0025 at scale effects on the lateral-characteristic parameters are sham for the Reptolde number range investigated. 2*CL : 2 Lo c-iatent The results of an investigation at hig
46、h Reynolde nuzlibers and low Mach numbers in the hgley full-scah tunnel of the ma;rimum-lift and lateral characteriertics of a rectangular wing of aspect ratio 3.4 wfth circular-arc airfoil sectIan are sunumrfzed as f ollowe: 1. The maximum lift coefficient of the basic wing is 0.58. The mition of -
47、span and fu-apan split flaps deflected 60 acreasee this value to 1.00 and 1.24, respectively. The agreement between the experimntal valuea of the maxlnarm lWt coefficient and lift-curve slope of the basic wing and the Fncremts in lift coefficient due to flap deflection with thoae calculated by the b
48、est available methode is good. 2. Maximum lift coefficients of 0.89, 1.20, and 1.21 are obtained for the wing with the drooped-nose flap deflected 2oo, for the wing with the extensible leeLb-lng-edge flap, and far the uhq with the couiblnation of drooped-noee flap deflected loo and the 0.032 round l
49、eading edge, respectively. These values are increased to 1.26, 1.58, and 1.47, reapectve, with the adition of -span split fhpa deflected 60. 3. The drag of the wing is high throughout the derate to high angle-of-attack range. The addition of split flaps caums a larga drag increase; however, an appreciable reduction Fn the drag Fn this range i6
