Maximum values of gas-dynamic flux densities

A general result valid for any compressible fluid is noted. It gives the maximum values of the flux densities of mass, momentum, and kinetic energy in steady and unsteady flows which are expanding isentropically from a reservoir

The effect of flow oscillations on cavity drag

An experimental investigation of flow over an axisymmetric cavity shows that self-sustained, periodic oscillations of the cavity shear layer are associated with low cavity drag. In this low-drag mode the flow regulates itself to fix the mean-shear-layer stagnation point at the downstream corner. Above a critical value of the cavity width-to-depth ratio there is an abrupt and large increase of drag due to the onset of the ‘wake mode’ of instability. It is also shown by measurement of the momentum balance how the drag of the cavity is related to the state of the shear layer, as defined by the mean momentum transport $\rho\overline{u}\overline{v}$ and the Reynolds stress $\rho\overline{u^{\prime}v^{\prime}}$, and how these are related to the amplifying oscillations in the shear layer. The cavity shear layer is found to be different, in several respects, from a free shear layer

Vortical structure in the wake of a transverse jet

Structural features resulting from the interaction of a turbulent jet issuing transversely into a uniform stream are described with the help of flow visualization and hot-wire anemometry. Jet-to-crossflow velocity ratios from 2 to 10 were investigated at crossflow Reynolds numbers from 3800 to 11400. In particular, the origin and formation of the vortices in the wake are described and shown to be fundamentally different from the well-known phenomenon of vortex shedding from solid bluff bodies. The flow around a transverse jet does not separate from the jet and does not shed vorticity into the wake. Instead, the wake vortices have their origins in the laminar boundary layer of the wall from which the jet issues. It is argued that the closed flow around the jet imposes an adverse pressure gradient on the wall, on the downstream lateral sides of the jet, provoking 'separation events’ in the wall boundary layer on each side. These result in eruptions of boundary-layer fluid and formation of wake vortices that are convected downstream. The measured wake Strouhal frequencies, which depend on the jet-crossflow velocity ratio, match the measured frequencies of the separation events. The wake structure is most orderly and the corresponding wake Strouhal number (0.13) is most sharply defined for velocity ratios near the value 4. Measured wake profiles show deficits of both momentum and total pressure

Streamwise vortex structure in plane mixing layers

The development of three-dimensional motions in a plane mixing layer was investigated experimentally. It is shown that superimposed on the primary, spanwise vortex structure there is a secondary, steamwise vortex structure. Three aspects of this secondary structure were studied. First, the spanwise vortex instability that generates the secondary structure was characterized by measurements of the critical Reynolds number and the spanwise wavelength at several flow conditions. While the critical Reynolds number was found to depend on the velocity ratio, density ratio and initial shear-layer-profile shape, the mean normalized wavelength is independent of these parameters. Secondly, flow visualization in water was used to obtain cross-sectional views of the secondary structure associated with the streamwise counter-rotating vortices. A model is proposed in which those vortices are part of a single vortex line winding back and forth between the high-speed side of a primary vortex and the low-speed side of the following one. Finally, the effect of the secondary structure on the spanwise concentration field was measured in a helium-nitrogen mixing layer. The spatial organization of the secondary structure produces a well-defined spanwise entrainment pattern in which fluid from each stream is preferentially entrained at different spanwise locations. These measurements show that the spanwise scale of the secondary structure increases with downstream distance

On Reflection of Shock Waves from Boundary Layers

Measurements of the reflection characteristics of shock waves from a flat surface with a laminar and turbulent boundary layer are presented. The investigations were carried out at Mach numbers from about 1.3 to 1.5 and a Reynolds number of 0.9 x 10^4. THe difference in the shock-wave interaction with laminar and turbulent boundary layers, first found in transonic flow is confirmed and ,investigated in detail for supersonic flow. The relative upstream influence of a shock wave impinging on a given boundary layer has been measured for both laminar and turbulent layers. The upstream influence of a shock wave in the laminar layer is found to be of the order of 50 bounday-layer thicknesses as compared with about 5 in the turbulent case. Separation almost always occurs in the laminar boundary layer. The separation is restricted to a region of finite extent upstream of the the shock wave. In the turbulent case no separation was found. A model of the flow near the point of impingement of the shock wave on the boundary layer is given for both cases. The difference between impulse-type and step-type shock waves is discussed and their interactions with the boundary layer are compared. Some general considerations on the experimental production of shock waves from wedges and cones are presented, as well as a discussion of boundary layer in supersonic flow. A few exampies of reflection of shock waves from supersonic shear layers are also presented

Observations of supersonic free shear layers

Visual spreading rates of turbulent shear layers with at least one stream supersonic were measured using Schlieren photography. The experiments were done at a variety of Mach number-gas combinations. The spreading rates are correlated with a compressibility-effect parameter called the convective Mach number. It is found that for supersonic values of the convective Mach number, the spreading rate is about one quarter that of an incompressible layer at the same velocity and density ratio. The results are compared with other experimental and theoretical results

Incipient Separation of a Turbulent Boundary Layer at High Reynolds Number in Two-Dimensional Supersonic Flow over a Compression Corner

An experimental study was made of the conditions necessary to promote incipient separation of a turbulent boundary layer in two-dimensional supersonic flow over a compression corner. The aim was to extend Kuehn's earlier results to higher Reynolds numbers. Measurements were obtained for Mach numbers in the range 2 to 5 and at Reynolds numbers , based on the boundary-layer thickness, in the range 10^6 to 10^7, nearly two orders of magnitude greater than those reported earlier. The main result was that the trend with Reynolds number established by Kuehn for the pressure rise for incipient separation does not continue to the high Reynolds number values of the present experiments; in fact, it is reversed. Pressure distributions were also obtained for conditions with and without separation. For the latter case, the upstream influence was considerably less than one boundary-layer thickness end the initial part of the pressure rise was practically a jump, suggesting that the oblique shock has its origin deep in the boundary layer

Measurements of test time in the GALCIT 17-inch shock tube

Experimental measurements of test time were obtained in the GALCIT 17-in. shock tube using both air and argon for driven gases. One series of tests was conducted using a constant driver pressure (pure helium) for various initial pressures of the driven gases. Another series was conducted using air for the driven gas at various initial pressures holding the shock Mach number constant. The data are presented and compared to theoretical predictions computed from the theory in two recent papers by Mirels for the case of a laminar and turbulent wall boundary layer

Flare-Induced Interaction Lengths in Supersonic, Turbulent Boundary Layers

Experimental results are presented for the effects of Mach number, Reynolds number, and corner angle on flare-induced separation of a supersonic, turbulent boundary layer. Measurements were obtained for upstream interaction distance ℓ_0 from the flare to the beginning of the interaction for Mach numbers 2≤M≤4.5, boundary-layer thickness Reynolds numbers 10^5 < R_δ < 10^6, and adiabatic wall conditions. Flares of angle α≤40° were attached to a hollow-cylinder model of 12 in. diam at either x_c= 14 or 18 in. downstream from the sharp leading edge. It was found that ℓ_0/δ_0 decreases with increasing Mach number and Reynolds number and increases with flare angle. For constant α, when ℓ_0/δ_0 is plotted vs the local skin-friction coefficient, C_(f0), the Mach number dependence disappears. From this observation, a simple correlation formula was obtained and used to compare results from other investigations, and also to correlate incipient separation data