Hypervelocity Boundary Layers Research Articles

Analysis of focused laser differential interferometry.

A computational method for predicting the output of a focused laser differential interferometer (FLDI) given an arbitrary density field is presented. The method is verified against analytical predictions and experimental data. The FLDI simulation software is applied to the problem of measuring Mack-mode wave packets in a hypervelocity boundary layer on a 5° half-angle cone. The software is shown to complement experiments by providing the necessary information to allow quantitative density fluctuation magnitudes to be extracted from experimental measurements.

Observations of hypervelocity boundary-layer instability

A novel optical method is used to measure the high-frequency (up to 3 MHz) density fluctuations that precede transition to turbulence within a laminar boundary layer in a hypervelocity flow. This optical method, focused laser differential interferometry, enables measurements of short-wavelength, high-frequency disturbances that are impossible with conventional instrumentation such as pressure transducers or hot wires. In this work, the T5 reflected-shock tunnel is used to generate flows in air, nitrogen and carbon dioxide with speeds between 3.5 and $5~\text{km}~\text{s}^{-1}$ (Mach numbers between 4 and 6) over a 5° half-angle cone at zero angle of attack. Simultaneous measurements are made at two locations approximately midway along a generator of the 1-m-long cone. With increasing Reynolds number (unit values were between 2 and $5\times 10^{6}~\text{m}^{-1}$), density fluctuations are observed to grow in amplitude and transition from a single narrow band of frequencies consistent with the Mack or second mode of boundary-layer instability to bursts of large-amplitude and spectrally broad disturbances that appear to be precursors of turbulent spots. Disturbances that are sufficiently small in initial amplitude have a wavepacket-like signature and are observed to grow in amplitude between the upstream and downstream measurement locations. A cross-correlation analysis indicates propagation of wavepackets at speeds close to the edge velocity. The free stream flow created by the shock tunnel and the resulting boundary layer on the cone are computed, accounting for chemical and vibrational non-equilibrium processes. Using this base flow, local linear and parabolized stability (PSE) analyses are carried out and compared with the experimental results. Reasonable agreement is found between measured and predicted most unstable frequencies, with the greatest differences being approximately 15 %. The scaling of the observed frequency with the inverse of boundary-layer thickness and directly with the flow velocity are consistent with the characteristics of Mack’s second mode, as well as results of previous researchers on hypersonic boundary layers.

Stability of highly cooled hypervelocity boundary layers

The influence of high levels of wall cooling on the stability of hypervelocity boundary layers is investigated. Such conditions are relevant to experiments in high-enthalpy impulse facilities, where the wall temperature is much smaller than the free-stream temperature, as well as to some real flight scenarios. Some effects of wall cooling are well known, for instance, the stabilization of the first mode and destabilization of the second mode. In this paper, several new instability phenomena are investigated that arise only for high Mach numbers and high levels of wall cooling. In particular, certain unstable modes can travel supersonically with respect to the free stream, which changes the nature of the dispersion curve and leads to instability over a much wider band of frequencies. The cause of this phenomenon, the range of parameters for which it occurs and its implications for boundary layer stability are examined. Additionally, growth rates are systematically reported for a wide range of conditions relevant to high-enthalpy impulse facilities, and the stability trends in terms of Mach number and wall temperature are mapped out. Thermal non-equilibrium is included in the analysis and its influence on the stability characteristics of flows in impulse facilities is assessed.

Roughness-Induced Transition of Hypervelocity Boundary Layers

Experiments were performed in the T4 free-piston reflected shock tunnel to explore the transition process of hypervelocity boundary layers. The experiments took place on a 1000-m-long by 300-m-wide flat plate with a sharp leading edge at representative conditions pertaining to those of an unpublished test of a Mach 12 scramjet. A streamwise row of thin-film heat transfer gauges along the centerline of the plate was used to determine the state of the boundary layer. It was found that, without any boundary-layer trips, the boundary layer remained laminar for the entire length of the plate. Multiple combinations of trip geometry and size were then used in an attempt to force the boundary layer into turbulence. The most successful geometry was the diamond trip configuration, although transition did not occur directly behind the trip. Transition occurred after a period of relaminarization, with the length of the delay region agreeing well with the literature.

Free-stream density perturbations in a reflected-shock tunnel

Focused laser differential interferometry is used to quantify the free-stream density perturbations in the T5 reflected-shock tunnel. The investigation of reflected-shock tunnel disturbances is motivated by the study of hypervelocity boundary-layer instability and transition. Past work on hypersonic wind-tunnel noise is briefly reviewed. New results are reported for hypervelocity air flows at reservoir enthalpies between 5 and 18 MJ/kg at Mach ≈ 5.5. Statistical analysis finds no correlation of RMS density perturbations with tunnel run parameters (reservoir pressure, reservoir mass-specific enthalpy, free-stream unit Reynolds number, free-stream Mach number, and shot number). Spectrograms show that the free-stream disturbance level is constant throughout the test time. Power spectral density estimates of each of the experiments are found to collapse upon each other when the streamwise disturbance convection velocity is used to eliminate the time scale. Furthermore, the disturbance level depends strongly on wavelength. If the disturbance wavelength range of interest is between 700 μm and 10 mm, the tunnel noise is measured to be less than 0.5 % with the focused laser differential interferometer.

Scaling of heat transfer augmentation due to mechanical distortions in hypervelocity boundary layers

We examine the response of hypervelocity boundary layers to global mechanical distortions due to concave surface curvature. Surface heat transfer and visual boundary layer thickness data are obtained for a suite of models with different concave surface geometries. Results are compared to predictions using existing approximate methods. Near the leading edge, good agreement is observed, but at larger pressure gradients, predictions diverge significantly from the experimental data. Up to a factor of five underprediction is reported in regions with greatest distortion. Curve fits to the experimental data are compared with surface equations. We demonstrate that reasonable estimates of the laminar heat flux augmentation may be obtained as a function of the local turning angle for all model geometries, even at the conditions of greatest distortion. This scaling may be explained by the application of Lees similarity. As a means of introducing additional local distortions, vortex generators are used to impose streamwise structures into the boundary layer. The response of the large scale vortices to an adverse pressure gradient is investigated. Surface streak evolution is visualized over the different surface geometries using fast response pressure sensitive paint. For a flat plate baseline case, heat transfer augmentation at similar levels to turbulent flow is measured. For the concave geometries, increases in heat transfer by factors up to 2.6 are measured over the laminar values. The scaling of heat transfer with turning angle that is identified for the laminar boundary layer response is found to be robust even in the presence of the imposed vortex structures.

Boundary-Layer Transition Measurements in Hypervelocity Flows in a Shock Tunnel

Experiments to investigate the transition process in hypervelocity boundary layers were performed in the T4 free-piston shock tunnel. An array of thin-film heat-transfer gauges was used to detect the location and extent of the transitional region on a 1500 mm long x 120 turn wide flat plate, which formed one of the walls of a duct. The experiments were performed in a Mach 6 flow of air with 6- and 12-MJ/kg nozzle-supply enthalpies at unit Reynolds numbers ranging from 1.6 x 10(6) to 4.9 x 10(6) m(-1). The results show that the characteristics typical of transition taking place through the initiation, growth, and merger of turbulent spots are evident in the heat-transfer signals. A 2-mm-high excrescence located 440 turn from the leading edge was found to be capable of generating a turbulent wedge within an otherwise laminar boundary layer at a unit Reynolds number of 2.6 x 10(6) m(-1) at the 6-MJ/kg condition. A tripping strip, located 100 mm from the leading edge and consisting of a line 37 teeth of 2 rum height equally spaced and spanning the test surface, was also found to be capable of advancing the transition location at the same condition and at the higher enthalpy condition.

Numerical solutions for radiating hypervelocity boundary-layer flow on a flat plate.

Abstract : The effects of thermal radiation on laminar, flat plate boundary layer flow have been investigated. Non-similar velocity and temperature profiles were obtained as numerical solutions of an appropriate set of finite difference equations. Simplified expressions for the thermodynamic and transport properties were used together with a mean radiation absorption coefficient approximating that of air. On the basis of the results of a previous approximate analysis, it was assumed that the boundary layer would be optically thin; for the velocity range considered (10 to 40 km/s) the numerical results confirmed this hypothesis. It was found that the maximum temperatures in the radiating boundary layers were much lower than in the corresponding non-radiating cases and that the radiating boundary layers were much thinner. Changing the exponent in the temperature-viscosity relationship mu proportional to T to the omega power from 1.0 to .75 had a negligible effect on the maximum temperature in the boundary layer but caused a significant reduction in its thickness and thus in the radiative heat flux to the wall. The predictions of the approximate analysis were compared with the numerical results and found to give useful engineering estimates of the properties of radiating boundary layers. (Author)