Roger L. Kimmel

On hypersonic boundary-layer stability

This paper reviews experimental hypersonic boundarylayer stability results obtained using hot-wire anemometry techniques. Data are obtained at a freestream Mach number of 8 on water-cooled and uncooled 7-degree half angle cones and on a water-cooled cylinder. A limited amount of cone data were obtained at M, = 6 . It is shown that one can not just extend subsonic and supersonic stability concepts and transition data to hypersonic Mach numbers. Hypersonic boundary-layer transition phenomena have several unique features and the topics must be treated independently. In low speed boundary layers one is accustomed to thinking of the vorticity instability mode which produces low frequency, low amplitude velocity fluctuations. A unique feature of a hypersonic boundary layer is the presence of the higher instability modes, the Mack modes. These instabilities v produce high frequency, large amplitude density fluctuations which can dominate the transition process. Some hypersonic trends are different from lower Mach number trends. Surface temperature effect is a good example. Cooling the surface stabilizes low Mach number boundary layers, but can destabilize a hypersonic boundary layer. Many of the parametric effects are very sensitive to Mach number. For example, it is shown that a small nosetip bluntness can completely dominate the stability of a hypersonic boundary layer, resulting in very large critical Reynolds numbers. This paper reviews general hypersonic stability characteristics, comparisons with theory, several parametric effects, and cone versus planar boundary-layer stability.

Transition Analysis for the HIFiRE-5 Vehicle

The Hypersonic International Flight Research and Experimentation (HIFiRE) 5 flight experiment by Air Force Research Laboratories and Australian Defense Science and Technology Organization is designed to provide in-flight boundary-layer transition data for a canonical 3D configuration at hypersonic Mach numbers. This paper outlines the progress, to date, on boundary layer stability analysis for the HIFiRE-5 flight configuration, as well as for selected test conditions from the wind tunnel experiments supporting the flight test. At flow conditions corresponding to the end of the test window, rather large values of linear amplification factor are predicted for both second mode (N>40) and crossflow (N>20) instabilities, strongly supporting the feasibility of first in-flight measurements of natural transition on a fully three-dimensional hypersonic configuration. Additional results highlight the rich mixture of instability mechanisms relevant to a large segment of the flight trajectory, as well as the effects of angle of attack and yaw angle on the predicted transition fronts for ground facility experiments at Mach 6. 1. Background

Space-time correlation measurements in a hypersonic transitional boundary layer

An experimental investigation of the spatial structure of second-mode instability waves was carried out in the boundary layer of a 7-deg half-angle, sharp-nosed cone at an edge Mach number of 6.8. Measurements were made at Reynolds numbers of 2.3 x 10 6 to 9.1 x 10 6 based on boundary edge conditions, spanning the range from unstable laminar flow to nearly turbulent flow. Simultaneous measurements with two hot-film probes in the boundary layer comprise the primary data set. The mean boundary-layer state was measured with pitot and total temperature surveys. Correlations were taken with circumferential, streamwise, and vertical probe separations. The second-mode waves preceding transition are relatively limited in their circumferential dimension, typically less than four boundary-layer thicknesses, based on a coherence level of 20%.

Traveling Instability Waves in a Mach 8 Flow over an Elliptic Cone

Simultaneous measurements were carried out with three hot-film probes in the Mach 8 flow over an elliptic cone of 2:1 aspect ratio, and the data obtained were compared to the results of computations using the parabolized Navier-Stokes equations and linear stability theory. The elliptic-cone flow was found to be significantly different from the flows studied in previous hypersonic-flow stability experiments, which have focused exclusively on wind-tunnel models with two-dimensional, planar or axial symmetry. At least two instability mechanisms appear to be active in the present flow: one associated with the region of maximum crossflow in the vicinity of the shoulder of the cone and the other associated with the inflectional velocity profiles on the top centerline. Between the shoulder and leading edge of the cone, the dominant flow instability occurred at relatively low frequency, and the direction of the phase velocity was significantly skewed from that of the boundary-layer-edge streamlines. The results were found to be in rough agreement with linear stability calculations and are suggestive of a traveling crossflow instability mode, which apparently has not heen observed before in hypersonic flow

Hypersonic Flow Control Using Surface Plasma Actuator

Plasma-fluid-dynamic interaction has been shown to be a viable mechanism for hypersonic flow control. An effective and verified flow control process using direct current surface discharge is summarized. The operating principle is based on a small electromagnetic perturbation to the growth rate of the displacement thickness of a shear layer that is strongly amplified by a subsequent pressure interaction. The aerodynamic control is delivered in less than a millisecond time frame and produces no parasitic effect when deactivated. The magnitude of the resultant aerodynamic force and moment can be significant and does not require a large amount of power for plasma generation to overcome the inefficient ionizing process, thus reducing the weight of a high-speed vehicle. The electromagnetic perturbation is derived from a surface gas discharge with or without an externally applied magnetic field. An embedded plasma actuator near the leading edge of a flat plate has produced high surface pressure equivalent to more than a 5 deg flow deflection at Mach 5, and the flow control effectiveness will increase with an increasing oncoming Mach number. The detailed flow structure of weakly ionized airstreams has been investigated by a combination of experimental effort and computational simulation solving the magneto-fluid-dynamic equations in the low magnetic Reynolds number limit with a drift-diffusion plasma model. The identical plasma actuator is investigated as a variable geometry cowl of a hypersonic inlet. All phenomena are replicated by computational results and are fully validated by experimental observations.

Stability Analysis for HIFiRE Experiments

The HIFiRE-1 flight experiment provided a valuable database pertaining to boundary layer transition over a 7-degree half-angle, circular cone model from supersonic to hypersonic Mach numbers, and a range of Reynolds numbers and angles of attack. This paper reports selected findings from the ongoing computational analysis of the measured in-flight transition behavior. Transition during the ascent phase at nearly zero degree angle of attack is dominated by second mode instabilities except in the vicinity of the cone meridian where a roughness element was placed midway along the length of the cone. The growth of first mode instabilities is found to be weak at all trajectory points analyzed from the ascent phase. For times less than approximately 18.5 seconds into the flight, the peak amplification ratio for second mode disturbances is sufficiently small because of the lower Mach numbers at earlier times, so that the transition behavior inferred from the measurements is attributed to an unknown physical mechanism, potentially related to step discontinuities in surface height near the locations of a change in the surface material. Based on the time histories of temperature and/or heat flux at transducer locations within the aft portion of the cone, the onset of transition correlated with a linear N-factor, based on parabolized stability equations, of approximately 13.5. Due to the large angles of attack during the re-entry phase, crossflow instability may play a significant role in transition. Computations also indicate the presence of pronounced crossflow separation over a significant portion of the trajectory segment that is relevant to transition analysis. The transition behavior during this re-entry segment of HIFiRE-1 flight shares some common features with the predicted transition front along the elliptic cone shaped HIFiRE-5 flight article, which was designed to provide hypersonic transition data for a fully 3D geometric configuration. To compare and contrast the crossflow dominated transition over the HIFiRE-1 and HIFiRE-5 configurations, this paper also analyzes boundary layer instabilities over a subscale model of the HIFiRE-5 flight configuration that was tested in the Mach 6 quiet tunnel facility at Purdue University.

Three-Dimensional Hypersonic Laminar Boundary-Layer Computations for Transition Experiment Design

The stability of boundary layers on sharp-nosed cones with elliptical cross sections is assessed using linear stability theory and crosse ow correlations. The objective is to identify a cone guration for wind-tunnel testing that exhibits signie cant crosse ow but also possesses a sufe cient laminar region for boundary-layer stability probing. Parabolized Navier ‐Stokes computer codes were used to calculate the mean e ow about cones with eccentricities of 1.5:1, 2.0:1, and 4.0:1 at a freestream Mach number of 7.95 and freestream unit Reynolds number of 3:3 £ 10 6 m i 1 . Correlations indicated that transition was possible on each cone guration at the above conditions. All three cone gurations showed unstable, ine ectional velocity proe les and boundary-layer thickening along the centerline (minor axis) due to the ine ux of low-momentum e uid. Crosse ow separation was observed on the 2.0:1 cone guration. Linear stability theory was used to calculate stationary crosse ow N factors on all three cone gurations, and to calculate traveling-wave N factors on the 1.5:1 and 2.0:1 cone gurations. All three cone gurations showed crosse ow instability, with the 4.0:1 cone guration attaining the highest N factors. The 1.5:1 and 2.0:1 cone gurationswere unstable to a broad spectrum of traveling waves, with thehighest Nfactors attained on centerline, due to the unstable proe les there.

HIFiRE-5 Flight Vehicle Design

The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program executed by the Air Force Research Laboratories (AFRL) and Australian Defence Science and Technology Organization (DSTO). HIFiRE flight 5 is devoted to measuring transition on a three-dimensional body. This paper summarizes payload configuration, trajectory, vehicle stability limits and roughness tolerances. Results show that the proposed configuration is suitable for testing transition on a three-dimensional body. Transition is predicted to occur within the test window, and a design has been developed that will allow the vehicle to be manufactured within prescribed roughness tolerances

Spectral Characteristics of Separation Shock Unsteadiness

Spectra of wall-pressure fluctuations caused by separation shock unsteadiness were compared for data obtained from wind-tunnel experiments, the Hypersonic International Flight Research Experimentation flight test 1, and large-eddy simulations. The results were found to be in generally good agreement, despite differences in Mach number and two orders of magnitude difference in Reynolds number. Relatively good agreement was obtained between these spectra and the predictions of a theory developed by Plotkin. The predictions of this theory are also qualitatively consistent with the results of experiments in which the shock motion was synchronized to controlled perturbations. The results presented here support the idea that separation unsteadiness has common features across a broad range of compressible flows and that it behaves as a selective amplifier of large-scale disturbances in the incoming flow.

Aerothermodynamic Testing and Boundary-Layer Trip Sizing of the HIFiRE Flight 1 Vehicle

An experimental wind tunnel test was conducted in the NASA Langley Research Center’s 20-Inch Mach 6 Air Tunnel in support of the Hypersonic International Flight Research Experimentation Program. The information in this report is focused on the Flight 1 configuration, the first in a series of flight experiments. This report documents experimental measurements made over a range of Reynolds numbers and angles of attack on several scaled ceramic heat transfer models of the Flight 1 payload. Global heat transfer was measured using phosphor thermography and the resulting images and heat transfer distributions were used to infer the state of the boundary layer on the vehicle windside and leeside surfaces. Boundary layer trips were used to force the boundary layer turbulent, and a brief study was conducted to determine the effectiveness of the trips with various heights. The experimental data highlighted in this test report were used to size and place the boundary layer trip for the flight test vehicle.