1、RESEARCH MEMORANDUM EXPERLMENTAL INVESTIGATION AT HIGH SUBSONIC SPEEDS OF THE ROLLING STABILITY DERIVATIVES OF A COMPLETE lVlODEL WITH AN ASPECT-RATIO-2.52 WING HAVING AN UNSWEPT 72-PERCENT-CHOEI;D LSNE AND A HIGII HORIZONTAL TAIL By William C. Sleemm, Jr., and James W. Wiggins Langley Aeronautical
2、Labomtory Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-By William C. Sleem, Jr., and James W. Wiggins Rolling stabiltty derivatives me presentea for a complete model e having e lov-aspect-ratio wwg and tail surfaces for a Mach nunber range of 0.70
3、 to 0.94 and for mgle-of -attack rmge from 0 to 13 for the lower Mach numbers. Tne model had a wing of espect ratio 2.52, a %=per was E modified biconvex section of 3.4-percert-chord thickness hming 2;n elliptical Eose proile. 1 ratio of 0.385, md 19.1 sweep of the querter chord. The wing abfoil The
4、 model test results iadicated regions of neutral or unstable dmping in roll at Mach numbers of 0.85 and 0.90 in t:ne higher angle-of- etteck renge for the basic model. Mdition of bTng-tip tanks approxi- mately doubled the rtnnping in roll at low zsgles of attack and, although 1zge decreeses in dazug
5、ing occurred in going to high angles of ettack, positive daxping -as indicated over the range of test conditions for the cmplete mdel w5th tanka. At Oo angle of attack, ddition of wing-tip tmks increased the aileron effectiver-ess of the basic del; however, the rolling angular velocity which could b
6、e obtained wLth a given aileron deflection xss decreased ebout 30 percent by addition of the whg tanks. Deflection of leadirg-edge flags, in generzl, qpeared to increase the angle of at-ieck at which lazge losses in damping in roll occurred. In addition to the aforementioned dazing results, the othe
7、r rolling deriva-liveo (yawing noment am3 lateral force due to rolling) were obtabed. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA RM L5kI2O A series of tests vere =de in the kngley high-speed 7- by 10-foot tznnel to determine the rolling s
8、tEbility derivEtives of 8 com-glete model having a low-aspect-ratio wing md tail. The mdel ha6 e. wing of aspect ratio 2.52, E taper ratio of 0.383, and zero sweep of the 72-percent- chord lice (19.1O sh-eepback of the quarter-crord line). The wing eirfoil vas a nodified biconvex section of 3.4-perc
9、ent-chord thickness having an elliptical nose profile. Results are presented for %he basic configuration over a Yach num- ber range from 0.70 to 0.94 md for a mximum angle-of-attack range of approxirately 0 to l3O. A number of breakdown tests xere made to deter- mine the contribution of the tail sur
10、faces to the rolling Cerivatives or“ the model wi%h ans are giver? with respect to the center-of-gravity location shm in figure 2 (25-percent mean aero- dynamic chord on the iusehge cecter line). Cl rolling-sonent coefficier?t, Rolling moment q= Cri. CY yawing-moment coefficient , lateral-force coef
11、ficient , Yzwing moment kteral force ss 9 Q“mic pressure, 3V , lb/sq ft 12 P air density, slugs/cu ft v free-stream velocity, fi/sec . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-M S b G-l a a% cnp = - . a- P-0 2v W r; WT . 3 free-stream Mach nmb
12、er wing are leading-edge Tlp deflection, ?aracteristics is given in figure 2, and photogapns of the model momted on the forced- roll sting at Oo and IOo zngle of ettack ere given in figure 3. me nodel Tas constructed of steel. The wing which had loo of negative dihedrel and the tail surfaces could b
13、e removed from the fuselage for break-Zown tests. For these tests, tke comporerxt parts were replaced by snoot3 fairing blocks which continued the fuseiage contour. The air inlets were faired over as shown in fig- ures 2 and 3 an however, these corrections are believed to be snail. The figires prese
14、nting the results are as follows: Figure Rolling Stability Derivatives: wing off 6 Besic model, effect of tail surfaces 7 Basic model with tanks on . 8 Effect of leading-edge fl however, the damping in roll of the complete nzodel was still neutrel or sligntly unsteble . Comparison of figures 7 and 8
15、 indicates that aiiditior- of the wing- tip tanks approxilnately doubIed t:?e danping in roll of the basic model at low angles or“ attack. Although significant losses in damping occurred with increasing angle of attack, positive damping was indicated through- out the test angle-of-attack an on is ap
16、preciably ciifferent fron thst for the wing-off configuretion and this difference is in accord with the siae- wash dEe to roll effect discussed in reference 4. P Effects of ;I?odifica%ions to the basic model such as deflection of leading-edge flags snd addition of wing-tip tadss were relatively smal
17、l with regard to lateral force and y=wIng moment dm to rolling. Aileron Characteristics Aileron control characterfstics obtained from forced steady roll tests of the complete model hovever, the value of pb/2V which could be attained with a given gileron deflection with tanks on was decreased about 3
18、0 percent. This loss in rolling effectiveness with the tanks on is of course due to the increased damsing in roll obtained for this configmation. The aileron effectiveness at all angle of atteck of approximately 6.7O with tanks on was not appreciably dif- ferent from that for tb-e basic model, and t
19、he damping was generglly some- what Less - wnich resulted in en increase in (pb/2V- with tanks on. Results at the highest angle ol“ attack are presented for completeness; however, nonlinearities in the rolling-moment vle-of-attack rmge for the basic nodel. Addi- tion of wing-tip tanks approximately
20、doubled the dmping in roll at low angles of atteck and, elthough large decreases in damping occwred in going-to high angles of attack, positive damping vas insicate6 over tine rmge of test conditions for the cmplete nodel with tanks. At Oo angle of attack, additioo of the wing-tip tanks increesed th
21、e aileron effec- tiveness of the basic nQdel; however, the rolling angular velocity which could be obtahed with a given aileron deflection was decreased 14 06 Pb 2v a“ =O“ “ -06 -04 -02 0 02 04 .06 zv a, =6O - Pb 46 -M -02 0 .OZ .W 06 Pb 2v a, = 12“ - Figure 12.- Variation of labera1 characLeri6tics
22、 of the model with wing-tip helix angle. Configuration FV. I . . - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I I L I 4 I .t . I c p L7,=0“ a, =6“ Figurc 13.- Variation of lateral characteristics of the model with wing- tip helix angle. Configuration FVR. a, =/2“ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
