COMYITTEE FOR AERONATJTIGS
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SPIBBIXG VXsA.G.A. 0009,
OP TFINGS MONOPLANE TIXGS ifousc ..-., &abO.ratCrTY
c ... _._
_ .: 7
2sLO1!3. AND 6718 0.
By u. J. Bmabcr ar,d R. Lz?lue of C, decressos as &222z3" kzB - kXacrosses oxcept for the N.A.C.A. 0009 and K.A.C.A. 23018 win-s0 at SO0 , 6o”, The rat-c of and 70° angle of attack. change 6718 of wing 0, (figs. with kZ2 - kya -T--k3 - kX2 14 and 22). is largest for the N.A.C.A,
The results shoa that, generally, the algebraic value of the sideslia will increase for the wings in the folloainq order: 23-018, Clark Y, and N.A.C.A. 0009, 1J;A.C.A; B.B.C,A.6718. It is interesting to not-e. that the amount of camber-of the wing sections IncreasMn this came order (fiz. l), that the sideslip is more It appears, then, rlependant.upon. the camber than upon the thickness-of the T:ring section. The general indications are that the algebraic values N,A.C.A. of -c, roquircd decrease in Gho follorting ordor: 23018: n';A.C.A. 0009,,NbA.C.A. 6718, and Clark Ywhen kZ2 - kYs -- .--..-_I fs greater than 1.0; and N.A.S.A. 6718, N.A.C.AkZa - kXa EL-kY; 23018, is less 0009, and Clark IT.A.C.A. Y when kZs - kX The than leO (figs. 8, 10, 12, 14, 16, 18, 20, and 22). lr. za - kYa 1 is greater reason for the order to change when --kZ2 - kXa or loss than 1.0 is that the inertia yarning moment changes sign at that value. Pred&cti-o.n nlnno I-..from
spinninE characteristics Prodictio_n.. of-,thc
actoristics of an airplane in which any of these monoplane .- wings is used depends largely upon the aerodynamic-yawingmoment characteristics of the particular airplane. The value of C, as given in this report,is numericrequired, ally equal and of opposite sign to the sum of the wing yam? --. - .-.-At any angle of ating-moments and the inertia couples. is -supplied by- the--empennage, tack, mhen this value of' C, fuselage, and interference effects a steady spin will result provided that the equilibrium is stable; for any othor value of c, the airplane mill not spin at that at.titudo. In LT.7 dcr toinsure against a steady spin in any attitude, a ~aluo .opposing the spin mu.st be provided that is larger of %l required for that particuth>n any attainable value of C, lar loading condition. The yawing moment supplied by tho .:. -_ emDonnS.go, fuselage, and intorfcronce effects depends upon the sidoslip, the sizo and shnpa of the fuselage and tail surface 9, the location of the horizontaltail-surfaces nith respect to the fuselage, fin, and rudder, the amount of .fin aroa chead of the center of gravity, the interference offacts betnoon the rring and the rest of the airplane, and Data on some of these---the limits of the control movements. effects are reported in rofcronce 6 and in rcfarenccs 8 to _. 13. The geometry of the syin indicates that the vert.fcal ._ -__ ..tail surfaces shouldbecome more effective in producing & yaning moment opposing the spin as the sideslip becomes more outward. Fin area ahead of the canter of- gravity nil1 ixvo--~ moments aiding the spin if the sideslip is outward. yzvring (See re?oronce 12, fig. 2.) for all parts of the airplane values of C, the prediction of the spin would depend on.tho Twlgebraic-sum of the individual values of Cn for each part foreach angle In any estimate, f-or normal of attack. airglcncs, it will be found that a change in some-factors mill chmgo the yaring moment for some parts in a sense to opposo the skin; whereas, for other parts;the &T-f&c% Yill for be rcvorsed so that the magnitude of the chan.go in cn each part must be considered. more If the known, The airplane least likely hcving large algebraic values nlgebraic...