1、ImNATIONAL ADVISORY COMMITTEEFOR AERONAUTICSEFFECT OF HIGH-Lll?T DEVICES ON THE LOW-SPEEDTECHNICAL NOTE 2689STATICLATERAL AND YAWING STABILITY CHARACTERISTICS OFAN UNTAPERED 45 SWEPTBACK WINGBy Jacob H. LiechtensteinLangley Aeronautical LaboratoryLangley Field, Va.WashingtonMay 1952=- . - - . . . -
2、. . . . . + -_m _-Provided by IHSNot for Resale-,-,-K.NATIONALADVISORYTEGHLIBRARY-M: MIlllllllll!nllflllllululCCMMITTEE FOR AERONAUTICS IIOL5L5LTECHNICAL NOTE 2689EFFECT OF HIGH-LIIFPDEVICES ON THE LOW-SPEED STATICLATERAL AND YAWING STABILITY CHARACTERISTICSOFAN UNMHW2D 45 SWEPI!WCKWING1 By JacobH.L
3、iechtensteinsuMMARYA wind-tunnel inve”stigatibnwas made)in the Langley stabilitytunnel to determine the effect of lift flaps (leading edge and splittrailing edge) on the static lateral stability“derivati=s md theyawing derivatives of an untapered 45sweptbackwing at low speeds(Machnumber O.13).1 The
4、results of the tests indicated that, in general, the additionof inboard trailing-edge split flaps tened to displace the curves forboth the rolling moment due to yaw and the rolltng moment due to yawingvelocity“ina negative direction,whereas addition of O.9-SPaR outard-splitflaps tended to displace t
5、he curves for both rolling moments in apsitive direction. The addition of trailing-edge flaps tended, ingeneral, to tncreasethe -directionalstibil-ityand the dampg fi PW*Leading-edge flaps, however, generally caused the trends obsed at lowlift coefficientsto extend to higher lift coefficients for th
6、e staticlateral and yawtig stability derivatives. The effect of flaps on eitherthe lateral force due to yaw or the lateral force due to yawtig velocity 0apared to be unimportant. Because of the stiilar effect of the flapson the derivatives due to yaw and yawing velocity,the effect of theflaps on the
7、 derivatives in wing velocity appeared to be indicatedbythe manner in which the flaps affect the derivative in yaw.INTRODUCTIONEst=tion of the dynsmic flight characteristicsof airplanesrequiresa knowledge of the ccmqmnent forces and moments resulting from%upersedes the recentlyLift Devices on the Lo
8、w-S,Redteristics of an Untapered 451948.declassified NACA RM L8G20, “Effect of High-Static Lateral and Yawing Stability Charac-SweptbackWing” by Jacob H. Liechtenstein,. .- _ ._, _ ._. _- . - -. - .-.- - -Provided by IHSNot for Resale-,-,-2 NACA TN 2689the orientation of “theairplane with respect to
9、 the air stream andfrom the rate of angular motion of the airplane about each of its threesxes. The forces and moments resulting from the orientation of the air-plane usua13y =e expressed as the static stability derivatives,whichare readily determined in conventionalwind-tunnel tests. The forcesand
10、moments related to the angular motions (rotaryderivatives) havegenemlly been estimated from theory because of the lack of a conven-ient expertiental technique.The recent application of the rolling-flowand curved-flow princi-ple of the Langley stability tunnel has made equally possible the deter-mina
11、tion of both rotary and static stability derivatives, and this prin-ciple is now being utilized in a comprehensiveprogram of research todetemnine the effects of various geometricvariables on both-rotary andstatic stability characteristics.The results of an fivestigationof the static and yawing stabi
12、litycharacteristicsof a number of untapered swept wings, without high-liftdevices, have been presented in reference 1. An investigationof theinfluence of fuselage and tail surfaces is rerted in reference 2. Thepresent investigationis concernedwith the determinationof the influ-ence of various high-l
13、ift devices on the low-speed static lateral and yawing stability characteristicsof one of the sweptbackwings consid-ered in reference 1. lhasmuch as the experimentalresults for the wing-alone tests were compsredwith theoretical results in reference 1 and no.adequate theory for predicting the effect
14、of flaps on sweptbackwings isavailable,no comparisonsbetween experiment and theory are made in thispaper.SYMBOLS ,The results of the tests are presented as standard NACA coeffi- 1cients of forces and moments, which are referred to the stability axesfor which the origin is assumed at the projection o
15、n the plane of sym-metry of the quarter-chordPint of the mean aerodynamic chprd of thewing of the model tested. The stability-axessystem is shown in fig-ure 1 with sitive forces and momentsand symbols used herein are defined asindicated. The coefficientsfollows:CL lift coefficient (L/qS)drag coeffic
16、ient (D/qS).Cy lateral-force coefficientcl rolling-moment coefficient(Y/qs)(L/qsb).-. -Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS.,*NACATN%LDYLI?qPvsbcxFAA. 4-a,! rrbE2689yawing-momentlift, pmndsdrag, prods(N/ however,it may be mentioned here that t
17、he change in the slope of the curveat CL of about 0.5 is probably due to the early tip-stall characteristicof sweptbackwings. In view of the fact that the forces at the tip,because of the longer arm, exert conside=bly more influence on themoment derivatives than forces near the center, it is easily
18、understood, why the tip stall should result in such a chmwe h c. The fact thtthe slope of the Cr curve chau therefore, at this radial linewhereas, rearward of theaerodynamic center, the streamlinesapproach at effective pasitive yaw.Inasmuch as the tips are rearward of the aerodynsa(ccenter, it could
19、 besaid that the wtig is effectively at sitive yaw. Positive yaw tendsto reduce the effective sweepback of the left wing semispan and toincrease it on the right wing semispan. Because increased sweepback ,tends to delay the stall, the left semispanwouldbe expected to stall ,first and cause the slope
20、 of the rolling-moment curves-to change in anetive dtiection. The wing plus semispan trailing+dge-flap curvedoes not exhibit this decrease until the stall is more closely approached, ,and the curve, in general, is displaced negatively from the wing-aloneresults. The delay of the change b the slope o
21、f the curve is probably1 due to the fact that the semis- flaps increase the loading or thecenter part of the wing much more than at the wing tips and, consequently,the wing tends to exhibit somewhatmore uniform stalling characteristics. Because of the high center-sectionloading, in order to obtati z
22、erototal lift, the.angle of qttack must be decreased until the negative liftobtained at the tips is equal to the positive lift at the center. Thiseffect, h combinationwith the spanwise velocity gradient under yawingconditions, should cause a negative displacement of Czr with respectto the wing alone
23、. For the wing with 0.9-span outboard split-flaps thiscondition is reversed; in this case, the tips tend to load up more thanthe centerwith the consequencethat the value of c2r at zero lift ispsitive with respect to the wing alone. Addition of leading+dge flapsto either the wing alone or the wing wi
24、th 0.9-sp9n split flaps had onlyslight effect in the lift-coefficientrange between zero and about 0.7.At the high lift coefficients,although the leading-edge flaps wereunable to prevent the negative change in the slope of the C2r curve,they did prevent an appreciable decrease in Czr until maximum li
25、ftwas almost attained. ,The dsmping-in-yaw characteristics Cnr for the test configurationsare presented in figure 9. The results show that addition of eitherleading-edge flaps or semispan trailing-edge flaps to the wing alone didnot affect Cnr over the r“mge for which they are comparable. Additionof
26、 0.9-spantrailing-edge flaps or both leading-edgeand 0.9-s ant)trailing-edge flaps considerably increased the damping yaw -Cnr At high lift coefficients,the damping for the configurationwithleading-edge and 0.9-span trailing-edgeflaps was almost as much as thatfor a conventionalmodel with a vertical
27、 tail. Although Cnr iS IMhlya function of drag, for trailing-edge-flapconfigurationswhere thecenter of pressure is considerablyrearward of normal, the side forcealso can influence Cnr. This effect can be observed for both the wingProvided by IHSNot for Resale-,-,-I; jx NACA 2689 9with 0.9-spsn split
28、 flaps and the wing with leadtig-edge and 0.9-spansplit flaps by noting that where the CYr curve (fig. 10) was somewhatpositive with reswct to the wing alone, the r curve for the config-uration with flaps was considerablymore negative than the configurationwithout flaps.The magnitude and variation o
29、f cYr with lift coefficient for thewing alone was so small that it is believed to be of slight significance(fig. 10) and the addition of flaps did not appreciably alter thesecharacteristics.An interesting general observation that can be made is the verysimilarmarine in which the flaps affect the sta
30、tic lateral stabilityderivatives Cb Cnv and C) and the correspondingyawing stabilityderivatives (%r cnr, and cyT). This similarity seems to indicatethat the manner in which flaps are.likely to affect the yawing stabilityderivatives of a wing configuration can be predicted by observing theeffect the
31、flaps have on the static stability derivatives of the wing.CONCLUSIONSA wind-tunnel investigation of a 45 swepthackuntared wing withlift flaps in straight and yawing flow indicated the following generalconclusions: .1. At a given lift coefficient,the curves of rolling moment dueto yaw and rolling mo
32、ment due to yawing velocity were, in most instances,displaced in a negative direction by the addition of inboard trailing-edge split flaps and displaced b a positive direction by the additionof 0.9-span outboard trailing-edge split flaps.2. Trailing=dge split flaps were generally found to increase t
33、hedirectional stability and the damping in yaw.3. Leading-edge flaps generally caused an extension to higher liftcoefficientsof the trends usually noted at low lift coefficients forthe static lateral and yawing stability derivatives.,4. looking upstremn.NmQ?1 Provided by IHSNot for Resale-,-,-14 NAC
34、ATN2689.98.7.6.54.32./o3632282420/6840-4-8“-/2-6-4-2 0 2 4 .6 B 0/!2 /.4. -Provided by IHSNot for Resale-,-,-NACATN 2689 15.006.W4.008 . . JIIIII 1. . XI I.002.0?0027004.Cw.ax., .002-m2; 7(W4:006-8 -. -U -Z O .2 U .6 .8 10 /2 4CL.O 78 -.6 -4 -2 0 .2 .4 =6 .8 10 12 1.4 % =%=, FWIP 5. Vanotlon of Ctti
35、 with hff cwffkxwf hrk wmas Ad codguruiiom. . .-. - , .- .-. - - . .- - .- .-Provided by IHSNot for Resale-,-,-NACA l!N2689.Cnti.00207002-004I l-l II 1“I-mw-i$-l.:8 76 -.4 72 0 2 4 .6 B LO /2 /.4CL.004.00207002:004:8 :6 -4 :2 0 .2 4 .6 .8 10 G? 1.4CL“ Fgure 6- Vurttion of Cnti wlthl(ff cm%aent for t
36、he vunms teticonfigumtlons.,. . . . . . - . . _ - . - -Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS3X NACATN 2689 17.0/2.00807004:8 & :2$ .0 .2 4 .6 .8 LO / f.4%.012CyP“.004%-t-t-t/-P /J I I 1,1. I-78 76 -4 -2Cihff coeftick?ti for th vamws hs?conflgum
37、fiom. - - -z _._ _ . . . . . _ - .- . . .Provided by IHSNot for Resale-,-,-i8 NACAm 2&9.2Czr o-2-4I I I I I I I 1 1y-+. 1 I 1 1 IIu.2.CL:8 :6 -4 G O .2 4 E .8 LO /Z /.4c-figum8.-Vorm+onof Cjr wifh M coefilcfent h the vormustest conf.ons.,. . . - . -. - -Provided by IHSNot for ResaleNo reproduction o
38、r networking permitted without license from IHSNACAm 2689 19,.2./-278 -.6 -4 72 0 .2 .4 .6 .8 LO I!2 /#.,Fgum9.- Vamtlon Of Cnr h M COdt7CI& h & WriYZS/W co/u/Wo. . . . . .- - - . - - . .- .-. ,.- . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS20 NACA TN 29.Cyr420-2-4.164.20-2-478 -.6 -4 72 0 .2 4 .6 .8 LO 2 1.4I“NAC!A-Lan6y-6.7-62.1000.- . . . - - . .Provided by IHSNot for Resale-,-,-
