1、RM No. L8K12 RESEARCH MEMORANDUM LATERAL-CONTROL INVESTIGATION ON A 37 SWIWTBACK WING OF T =TI0 6 AT A REYNOLDS NUMBER OF BY s QJ 0 a? Robert R. Graham and William Koven c3 rn ,114 ,111 it : Langley Aeronautical Laboratory Langley Field, Va. 34; : 6,800,000 1 NATIONAL ADVISORY COMMITTEE FOR AERONAUT
2、ICS WASHINGTON January 27,1949 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-V c . LATERAL-CONTROL INITSTIGATION ON A 37O EWEPIBACK WING OF ASPECT MTIO 6 AT A FWNO OF 6,800,000 3y Robert R. Graham and William Koven The low-speed lateral-control cha
3、racterietics of a 37 sweptback semispan Wfng of aspect ratio 6 and NACA :Llj Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA RM No. BKl2 “ on the same line but turned perpendicular to the air stream had practically no effect on the roll- effec
4、tiveness, but moving the segments inboard caused an increase in effectiveness. c Vwyfng the span of the spoiler showed that the inboard portions of the spoiler were considerably more effective than the outboard portions in producing rolling moments. At high lift coefficients on the WFng with slat an
5、d double Blotted flap,. the half-span plain outboard spoiler with a 10-percent-chord pro- jection produced about the 6- rolling mment as a total aileron deflec- tion of 30, but at low lift coefficients on the plain wing the spoiler produced only about one-third the rolling moment of the ailercms. Th
6、e yawing maments due to oppositely deflected aileron6 were generally unfavorablb and became more unfavorable as the angle of attack was increased. Those due to spoiler projection were favorable but became less favorable as the angle of attack waa increased or as the spoiler w80 moved inboard. The st
7、all-contzol and high-lift devices had a negligible effect on the yaKing moments. OMJCTION The me of sweptback wings on high-speed airplanes introduces several stability and control problem in the low-speed range. Two of these pro- Pressure difference across seal Pressure difference across vents 1 E
8、L Reynolds nmiber (q) R angle of attack of root chord line, degrees L lift D bag pitch- moment about 0.23 M Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-local wing chard pargllel to plane of symmetry local wfng chord perpendicular to 0.27 Une wing
9、 span perpendicular to plane of symne mcBpBnt of area of aileron rearward of hinge line about hinge aXi8 aileron span measured along hinge llne spoiler span msaElured perpendicular to plane of syrmetry aileron chord rearward of hinge llne measured perpendicular to 0.27 line aileran nom-balance chord
10、 forward of hinge lhe measured perpendicular to 0.27 line a;snamic pressure coefficient of rate of change deflection rate of change viscosity of rolling-mcrment coefficient with ailerm of lift coefficient with aileron deflection Provided by IHSNot for ResaleNo reproduction or networking permitted wi
11、thout license from IHS-,-,-chs rate of change of aileron hinge-moment coefficient ufth aileron deflection PRS rate of change of resultant-pressure coefficient with aileron deflectim % rate of change of aileran hinge-mment coefficient with angle of attack R, rate of change of resultant-pressure coeff
12、icient with angle of attack All coefficients and dimensicm sylribols refer to the model as a cmplete wfng. The effects of the spoiler controls on lift, drag, and pitching-moment coefficients are presented as the effects of spoiler projection on one side of a cmplete wing. The model wed in the invest
13、igation was a semispan wing mournted in the presence of a reflection plane as shown in fi e8 1 and 2. It was of steel construction and had an aspect ratio of ra taper ratio of 0.5, and 37O sweepback of the leading edge. The afrfoil section perpendicular to tihe 2j“percent-chord lFne (25-percmt-chord
14、 Une of wlng when In the unswept condition) was an WCA 641-212 pof ile The model was fenished . with lacquer and w88 .maintained in 871 aerodynamically smooth condition throughout the tests. The general plan form and some of the more perti- nent dimensions of the model me shown in figure 3. Details
15、of the lateral-control devices are shown in figure 4. The aileron wa8 of the constantpercentage-chord type (0.20 or 0 -183) and had the sms contour as the corresponding portion of the airfoil section. It was arranged to simulate a sealed internally balanced type of aileron with zero balance. The sea
16、l was simulated by a steel plate beveled to a knife edge with the edge a8 close as possible to the nose of the aileron at the hhge -e. Although thfs method did not completely seal the afleron, the reeulting gap was only a Ermall fraction of the balance-cmprtnent vente at the upper and lower surfaces
17、 of the wing. The balance capm-tmsnt was prodded with orifices for measuring pressures above and below the seal. The aileron was attached to the wing by strain gages which indicated electrically the aileron hinge moments. Two configurations of spoiler lateral controls were investigated. One extended
18、 along a comtant-percentage-chord line and the other con- sisted of a series of spoilers, each 10 percent of the w3ng semispan in length and placed perpendicular to We plane of symmetry. The first is referred to herein as the plain spoiler and the second as the step spoiler. Provided by IHSNot for R
19、esaleNo reproduction or networking permitted without license from IHS-,-,-6 RAM m No. L8m Both configuratiom simulated retractable circular-arc spoilers. Varioue projections of the plain spoiler were tested at the 0.65 and 0.73 positions, but only one projection of the step spoiler was testedwlth th
20、e midpoint of each step at 0.65 . The projection of the plab spoiler was a constant percent chord along the span, but the step spoiler prodectfon varied in steps along the span. The height at the center of each step was a constant percent chord but the individual steps were a constant height along t
21、he span of each- step- The aileron and plain spoiler extended from 0.50% to 0 -97%. The b step spoiler extended from 0 -27% to 0-97%, but the span and spanwise location could be varied by varying the ng correctim, obtained by coaibining the methods of references 3 and 4 were made to the angle of att
22、ack and to the drag, pitching-moment, yawing-mcanent, and rolling-moment coefficients The corrections were applied a6 follows: J Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-7 a = ageom3tric + 1.12 CL + 0.0164 Lq c, = sass + 0.0101 CL cD = GDpOSS
23、cn = %gross - %lm - where the adscript “gross“ refers to the uncorrected coe fficients, the subscript “tare“ refers to the uncorrected coefficients obtained with aileron or spoiler neutral, and K2 and are We rolling-moment and yawing-moment jet-boundary-correction canstanta Values for K2 are present
24、ed in figure 6 for various spans and spanwise locations of the lateral control. The values of momsnt data ed are given: namely, 0,0481 for a -span lateral control with the outboard end at 0 .g“$ and 0.0578 for a half-span lateral control with the outboard end at 0.779. No jet-boundary corrections ha
25、ve been applied to the 2 aileron hinge-manent data. A calibration of the aileron seal indicated a leakage factor E of 0.14. The balance-compar-tment pressures have been corrected for this leakage so that they represent pressures with a complete seal. The eff ecta of the lealrage on the rolling-mcane
26、nt and hinge-mament coefficients, however, are believed to be mall and have been neglected- The tare and interference effects of the model supports were not determined but are believed to have only a -1 effect on the character- istics of the Wlhg. The results of the aileron investigation are present
27、ed in figure 7 for the basic wing and Fn figures 8 to 12 for the wing with the varfous stall- control and high-lift devlces. Summary figures showing the effects of the various devices on the aileron hinge-molllsnt and rolllnn-effectiveness parameters are shown in figures 13 to 16. The results of the
28、 epoiler investigatlan are presented in figures 17 to 20 for the basic wing and fiEures 21 to 27 lor the wing with the various stall-control and high-lift devices. A canparison of the ailerm and spailers is presented in f igwe 28 Provided by IHSNot for ResaleNo reproduction or networking permitted w
29、ithout license from IHS-,-,-8 Ailercm Characteristics EXACA RM No. L8m2 Aileron characteristics on basic wlnq.- The data for the aileron tests on the plain wing are presented in figure 7. The cantzol-effective- ness parameter Cz8 was obtained from cross plots of the data of figure 7 and is presented
30、 aa a function of angle of attack in figure 13- The loss Fn aileron effectiveness that is usually found on sweptback wings at high #es of attack is clearly shown in figure 13- The reduction in Cz8 at angles of attack below the stall is probably cawed by the thickened bomdary layer due to the cross f
31、low along the trailing edge near the wing tips and the large reduction at the stall is.attributed to tip stalling. The value of Czg for low angles of attack has been calculated by the method given in reference 5 The cclmputed Cz8 when reduced by cos2A to account for sweep (reference 6) and corrected
32、 for section- lif t-curve slope (0 -109 for he high angles of attack. Split flape in cmibinatim with the slat generally caused a reduction in Cz8 from that for the wlng alone throughout the angle-of -attack range Double slotted flaps in cmbinatian with the slat effected an lncrease in Czs at all ang
33、les of attack. - The effects of the various stall-control devices on -the mUing- mcsnent coefficiant for a total aileron deflection of 30 are he stall of the basic wing. The leadlng-edge flap caused the Mgeet reduction in C at a = Oo and all the devices produced about the 881118 change at high angle
34、s of attack. The split flap in cabination with the slat caused a slight reduction in C2 from that with the slat alone In the high angle-of-attack range. The C2 Mth double Slotted flap in combination with -the slat was larger than that for any other configuration in the low and -moderate angle-of - a
35、ttack range The stall-control and high-lift devices had a negligible effect on the yawtng-moment coefficients due to oppositely deflected ailerons. (see figs. 7 to 12.1 The effects of the stdJ“cmto1 devices on the aileran hinge-moment parameters me he trailing edge from o -6% to o .75c reduced the y
36、awing-mcanen-b coeff icients . The characteristics of the step spoiler are Shown in figure 19. Comparison of figure 19 with figure 17 shows tbat a step spoiler of the stme span, spanwise location, and projection aa the plain spoiler pro- duced about the same rolling moment a8 the plain spoiler excep
37、t at angles of at-bck just below the stall where the step spoiler showed a slight improvement over th plain spoiler. These data are in disagreement with of reference u where a step spoiler on a wing of lower aspect ratio Provided by IHSNot for ResaleNo reproduction or networking permitted without li
38、cense from IHS-,-,-12 NACA RM No. L8m and slight9 greater sweep showed greater rolling effectiveness than a comparable plain spoiler. No explanation ha been found for the disagree- ment but it may be due to the difference in the geometric oharacteristics of the wings. The step spoilers caused a slig
39、htly BmaLler yawing mament than the plain spoiler. U Moving the spoiler Fnboard but maintaining the same span caused an appreciable increme in fte effectivenese, which is in general agree- ment with the resul-h of reference ll. The inboasd movement of the spoiler also cawed a reduction in yawing-mom
40、ent coefficient. The effects on the roUFng effectiveness of vqhg the span and the spanwise location of the step spoiler are shown in figure 20. It can be seen that the addition of the outboard portions of the step spoiler adds little to the effectiveness of the inbowd spoiler. Inboard addltiona to t
41、he outboard spoiler, on the other hand, produced caneiderable increases Fn ef f ec tiveness . It is believed that the plain spoiler would exhibit similar charac- teristics if its span and spamwise location were wrled. It fe aUo probable that changes Fn the gemetrg of the aileron would produce simila
42、r changes Fn effectiveness. 4 Effects of stall-cantrol and hingles of attack. Deflecting the split or double slotted flaps in conjunction with the slat caused a cmiderable increase in the spoiler yawing maments through the angle-of-attack range. I The effects of changing the span of the step spoiler
43、 in the presence of the leading-edge flap are shown in figure 25. It can be seen that in the low and moderate angle-of-attack ranges, changing the span of the Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-c spoiler had about the s effectiveme of th
44、e plain spoiler and the, ail.ercm is shown in figure 28. It shows that near zero lift the rolling-mament coefficfent produced by the spoiler wa8 about the same as that produced by a total aileron deflectian of only go. At a lift coeffi- cient of 1.6, however, (slat and dodle slotted flap extended) t
45、he spoiler produced about We same rolling Mrment a8 a total aileron deflection of 309- The 0.10 projection of the spoiler WRB chosen as the lmximlrm that could be obtained with a retractable-arc-type Bpoiler. The ma3dmum deflection that odd be obtained with a sealed intew balanced aileron on a wing
46、swaz to the model tested woula be about 215O. It can be seen, therefore, that at high lift coefficients the epoiler has about the same effectiveness as the aileron. At low lift coefficients the Spoiler apears to be Considerably less effective than tihe aileron, but spoilers have been aham to produce
47、 smaller wlng twietj_ng moments than ailerons (reference 13) and under high-speed-fley Aeronautical Laboratoq National Advisory CoIlPnittee for Aeronautics Langley Field, Va. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-16 NACA RM NO* L8lU2 1. Kov
48、en, William, and Graham, Robert R. : Wind-Tunnel Investigation of the High-Lift and Stall-Cmtrol Devices on a 37 Sweptback Wing of Aspect Ratio 6 at Hi&- Reynolds Numbera. NACA RM No- L8D29, 1948. 2. Graham, Robert R., and Canner, D. William: lhvestig$don of High-Lift and Stall-Conbol Devices on an NACA &-Series 42 Sweptback Wing with and without Fueelage- NACA RM No. L7GO9, 1947. 3. SiveUB, James C ., and Deters, Owen J. : Jet-Boundary and Plan-Form Corrections for Partial-Span Models with Reflection Plane, End Plate, or No End Plate In a Closed Circular Wind
