超音速翼型の造波抵抗計算法について

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Three methods of calculating wave drag of supersonic wing sections are presented here. These are : (a) Excess pressures across the leading edge shock and along the wing surface are calculated numerically by dividing the wing section into 16 parts. The effect of reflected waves due to the interaction between the shock and the expansion waves is ignored. (b) The drag coefficient is calculated by applying the Buse-mann's series expansion to the flows across the shock and along the wing surface. The third term of the flow deflection angle, θ, is taken into accounts to retain the effect of the entropy change. (c) The third method is all the same as (b) except that the pressure along the surface is not expanded around the point just behind the leading edge shock but around the point where θ=0. Using these methods, the wave drag coefficient of a symmetrical circular arc wing section having 0.081 thickness ratio and zero angle of attack is calculated at various Mach Numbers, and the results are compared with those by Ackeret's method and experimental one at Goettingen. It is concluded in this study that the correction due to the Busemann's D-term is not so large between M=2 and M=5, and that, in the region of the Mach Number near to unity it seems to diverge. The value obtained by the method (a), which is about 20% larger than that obtained by the Ackeret's method, nearly coincides with the experimental one. For the purpose of profile design the method (a) seems to be most adequate and at that time the Ackeret's theory will be useful as a pointer. (Received October 2. 1952)
宇宙航空研究開発機構の論文
1953-01-10

超音速翼型の造波抵抗計算法について

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