The flow past families of two-dimensional cylinders and bodies of revolution with elliptic noses at zero incidence has been examined at free stream Mach numbers between 1·42 and 1·82. The observations include schlieren photographs and measurements of the pressure distributions at the surface.

The measured pressure distributions and positions of the detached bow waves are found to be in fair agreement with values calculated by a number of methods. The drag coefficients of the slender two-dimensional elliptic noses are considerably higher than those calculated for wedges of the same axis ratio with attached shock waves, but the drags become almost equal when the axis ratio approaches the value for subsonic flow behind the bow wave of the wedge. For most axis ratios the drag coefficients of the elliptic-nosed bodies of revolution are lower than those for cones of the same axis ratio.

The noise levels of a jet issuing from a long pipe are compared with those of a jet having a square velocity profile at the exit. A subsonic noise reduction of between 2 and 5 decibels for various conditions is found for the case of the flow emerging with an approximately “turbulent pipe-flow” velocity profile for the same maximum jet velocity, but this is at the expense of a loss in thrust of a quarter. On comparison with a jet of smaller diameter which has an equal thrust for the same maximum jet velocity, it is found that the changes in noise level are rather smaller. For jets of equal diameters, the effects on the subsonic aerodynamic noise generated of a reduction of velocity gradient near the boundary are more than offset by the increased velocities necessary near the centre of the jet to obtain equal thrust. It is concluded that if the effect of differences in initial turbulence can be neglected the use of an auxiliary flow forming a comparatively thin sheath of slower moving fluid at the exit is not likely to result in large decreases in the subsonic noise level, and that a general reduction in jet velocity is more effective.

Above the critical pressure larger reductions of up to 10 decibels are found. These are consistent with a delay of the onset of the self-maintained shock-produced noise.

This paper gives a new calculation of the solid and wake blockage for compressible subsonic flow about a symmetrical two-dimensional aerofoil, midway between symmetrically disposed tunnel walls, which need not be straight. Previous calculations have been based on the theory of sources, and the results obtained have usually involved only first order terms. At high subsonic Mach numbers the second order terms become important; they are given in this paper. The theory is based on an integral equation, which is exact for incompressible flow, and which is more accurate than linear pertubation theory in compressible flow. The effect on blockage of a possible increase in the boundary layer displacement thickness on the tunnel wall, due to the presence of the aerofoil, is investigated, and finally a method of calculating the total blockage from wall pressure measurements is given.

The paper contains an analysis of the flow of a perfect gas with constant specific heats through a rotating channel of constant cross-sectional area, as used in certain helicopter propulsion systems and wind-driven gas turbines. The analysis is restricted to the adiabatic one-dimensional treatment, the Coriolis accelerations acting across a section being disregarded.

The equations of motion and energy are deduced and, together with the equation of continuity, reduced to an ordinary non-linear differential equation of the first order, involving a dimensionless form of velocity and distance.

The patterns of the integral curves of the differential equation are discussed and sketched by examining their asymptotic behaviour as well as that in the neighbourhood of singular points. It is shown that there exists a critical value of the angular velocity below which the flow remains subsonic in the pipe, if the entrance velocity is subsonic; it may, however, become sonic at the exit of the pipe. For supercritical angular velocities the flow may become sonic or supersonic in the pipe if the entrance velocity attains a given “ correct” but subsonic value. A method of examining for the possibility of shock formation is indicated.

The initial conditions for the differential equation are deduced for the design and for the performance problem, two new flow functions being introduced and tabulated to facilitate practical calculations. Formulae are also deduced for the calculation of the pressure and Mach number variation from the previously calculated velocity variation along the pipe.

Finally an approximate solution in closed terms is given for the case of small entrance velocities.

An instrument is described which detects total pressure, total temperature, and direction of flow simultaneously in a stream tube of one-sixteenth inch diameter. The observed directions and pressures are transmitted electrically to a remote station where they appear as pencil traces on a paper record which moves at a rate proportional to the traverse rate of the instrument. The temperature is noted at intervals and added manually to the trace. Integrators are included which permit rapid assessment of mean pressure and direction at the conclusion of a traverse, the control and maintenance of setting of the instrument being semi-automatic. A further development is also noted whereby total pressure and direction in three dimensions can be observed either manually, or remotely, by automatic means.