As an example of lifting wings in nonuniform flight speeds, slender wings are studied. The flight speeds may be subsonic, transonic, or supersonic. It is found that the lift force and lift distribution are affected explicitly by the instant acceleration, and not by the whole history of the past flight. It makes a clear contrast to usual non-slender wings. An acceleration (a deceleration) increases (decreases) the lift force, from that for a constant flight speed. The reason of this result is inferred. An important dimensionless quantity, "reduced acceleration, " is defined by UC0/UU2, where U is acceleration, C0 is maximum wing chord, and U is the instant flight speed.
Monocoque structures, which are considered as an assembly of a number of rods and pannels, are generally analyzed by the usual method. But the stiffness of the thin plate element is largely changed due to effects of buckling, large deflection and others. To account for these features, some particular devices may be introduced, otherwise the use of the more refined mesh or adequate discretion becomes essential to obtain suitable results. The present paper gives a convenient representation for the stiffness matrix of the rectangular and curved panel surrounded by stiffeners. The expression is an incremental relation between some modes of membrane stresses distribution, surface pressure and deflections or deformation of the plate, where the matrix elements also depend on the total membrane stresses and deflections. Using this relation, analysis of the total structure may be easily performed by the well-known incremental load method. Several numerical examples are shown, where number of modes used is 27 in stress distribution, and 15 in deflection. Comparing these results with the exact solution, it is proved that the expression is sufficient for practical use.
The Semi-empirical formula of estimating the induced rolling moment has been developed. This formula is based on the concept that the magnitude of induced rolling moment is ruled by the dimension and structure of vortex wake behind missile, and is represented as a function of the parameter describing the relative dimension of vortex wake to the fin span. Since the formula includes a few assumptions and experimental parameters, agreement between the theory and experiment is not always good. However, the formula is useful to get qualitative understanding of the effect of induced rolling moment on missile, and moreover will give correct quantitative results if employed under accurate underStanding of the flow field around missile. Presented in this paper are application of the formula to cruciform missile, comparisons between the flight data of sounding rockets and the simulation results obtained by this formula, and finally the effect of fin design on induced rolling moment.