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The longitudinal stability characteristics of elastic swept wings of high aspect ratio experiencing bending and torsional deformations are calculated for supersonic speed by the application of linearized lifting-surface theory. A parabolic wing deflection curve is assumed and the analysis is simplified by a number of structural approximations. The method is thereby limited in application to wings of high aspect ratio for which the root effects are small. Expressions for the lift, pitching-moment, and span load distribution characteristics are derived in terms of the elastic properties of the wing; namely, the design stress, the modulus of elasticity, the shearing modulus, and the maximum design load factor. The analysis applies to wings with leading edges swept behind the Mach lines. In all cases, however, the trailing edge is sonic or supersonic. Application of the method of analysis to wings with leading edges swept ahead of the Mach lines is discussed.

Estimates for pressures on the surface of a given delta wing at zero incidence in a steady uniform stream of air are obtained by numerically integrating two semi-characteristic forms of equations which govern the inviscid supersonic flow of an ideal gas with constant specific heats. In one form of the equations coordinate surfaces are fixed in space so that the surface of the wing, which has round sonic leading edges, is a coordinate surface. In the other, two families of coordinates are chosen to be stream-surfaces. For each form of the equations, a finite difference method has been used to compute the supersonic flow around the wing. Convergence of the numerical results, as the mesh is refined, is slow near the leading edge of the wing and an extrapolation procedure is used to predict limiting values for the pressures on the surface of the wing at two stations where theoretical and experimental results have been given earlier by another worker. At one station differences between the results given here and the results given earlier are significant. The two methods used here produce consistent values for the pressures on the surface of the wing and, on the basis of this numerical evidence together with other cited numerical results, it is concluded that the pressures given here are close to the true theoretical values.

A NUMERICAL study is made of the supersonic flow around the wedge-shaped leading edges of an axisymmetric body with a channel, in the case of a toroidal separated shock wave. The solution is obtained by the method of integral relations in a one-strip scheme in special variables taking into account the nature of the flow at the top of the edge. The effect of geometrical parameters of the body and of the Mach number of the incident flow is studied.

The properties of a riblet surface at a Mach number of 1.25 are described. Direct measurement has been made of the surface shear stress on a riblet surface attached to a skin-friction balance mounted beneath the turbulent boundary layer growing along a wall of the RAE 2 ft×1 1/2 ft variable density transonic wind tunnel. Comparison of the data with that for a smooth wall indicates a reduction of about 7% in skin friction at optimum flow conditions. Misalignment of the riblet surface with the flow direction reduces this benefit and beyond 30° misalignment a drag penalty occurs. Analysis of the velocity profiles indicates that compressibility effects may be accounted for with the use of an appropriate transformation.

The transient supersonic flow of a slightly conducting gas in an MHD channel with axial symmetry is investigated assuming that the magnetic Reynolds number and the interaction parameter are small. The solutions to the linearized equations for supersonic flow show that in the asymptotic limit as t →∞, these solutions all tend to coincide with the solutions for the corresponding steady flow. This point is remarkably different as compared with the result for the transient subsonic flow in a channel. The reason for this discrepancy is made clear. The time for the transient flow to achieve the steady state is also evaluated.

The subject of present investigation is the diffraction in such ranges of angles between the fronts of weak incident pressure jump and of an oblique compression shock attached to a wedge of finite opening angle in which the intersection of these occurs at the distorted section of the jump. On the side of smaller angles these ranges border on regions of possible regular interaction between two shocks coming from opposite directions [1], while on the side of greater angles these border of regions of shock waves moving in the same direction, which admit uniform streams in the neighborhood of intersection of fronts [2, 3]. The resulting boundary value problem has much in common with the similar problem of regular counter-interaction which was considered in [4]. In that paper, as in the present one, the method of analysis is related to the problem of perturbations of a uniform stream behind a plane shock front considered by Lighthill [5, 6]. The effect of triple shock configuration is represented by the combined dynamic singularity at the triple point. Some of its properties were predicted by Landau [7]. The particular case of motion of a slender wedge at hypersonic speed was considered by Inger [8, 9]. However his analysis contains statements which contradict existing concepts.

A closed form solution is given for the supersonic rolling of an elastic cylinder on a plane using the method of the preceding paper [1]. Tangential and normal forces are transmitted and Coulomb friction is assumed.

A visual study is performed in a supersonic, two-dimensional wake; the high value of the Reynolds number ensures that the wake is turbulent from the trailing edge. The flow is seeded by fluid vaporization in one boundary layer upstream of the trailing edge; a light sheet is generated by a Q-switched, high energy ruby laser. The set of photographs taken from the trailing edge up to the far wake is then processed after digitization of the pictures. A progressive contamination of the lower part of the wake by the fluid initially present in the upper part can be observed. In the far wake region, well organized large scale structures can be visualized. Statistics are performed and the results are compared with previous hot-wire measurements and discussed in terms of downstream wake behaviour.

For wings with supersonic edges and with arbitrary dihedral, twist, camber and thickness distribution, the pressure distribution on the wing exterior to and along the two Mach lines emanating from the vertex of the wing is equal to the corresponding pressure distribution for a planar wing. The problem is to find the pressure distribution inside the two Mach lines. In the present paper, the unknown pressure distribution is approximated by an elementary function of the two surface variables. The (as yet undetermined) constants in the function are then found by the conditions: (i) that the function takes on the corresponding planar values along the two Mach lines, (ii) that it fulfils certain generalized integral relationships (Ting 1959), and (iii) that it satisfies the averaging property of solutions of the wave equation to be developed in this paper. The generalized integral relationship relates the integral of the pressure distribution along the line of intersection of a Mach plane with the wing to the integral along the same line of the prescribed normal velocity. The averaging property relates the pressure distribution along the line of intersection of the surface of the dihedral wing to that on a planar wing.

Pressure distribution and the effect of air injection in the separated flow over a spiked-hemisphere were investigated at a Mach number of 3·3, and Reynolds number around the transitional value. Pressure distribution along the spike as well as over the body was measured in the absence of injection. Air was injected into the separated flow at the spike tip and base and reattachment region through one or more orifices drilled normal to the surface, and the resulting flow patterns were observed using the schlieren technique. The results show that (i) the pressure variation along the spike is similar to a two-dimensional separated flow in the transition régime; and (ii) the mass injection at the spike tip has a strong destabilizing effect regardless of injection rate, while the injection from spike base and reattachment region can be either slightly stabilizing or destabilizing depending on the flow condition.

The Black Brant IV shown in Fig. 1, is a two-stage sounding rocket with a 15KS-25 000 booster motor stabilised with three fins and a 9KS-11 000 motor, a 5·3 fineness ratio conical noise and an 11·5° conical flare forming the second stage. Basically, the second stage is the single stage Black Brant III rocket, but uses a conical flare instead of three fins at the aft end for aerodynamic stability. This paper reviews the static stability of a Black Brant IV second stage with a flare and two fin-flare combinations.