High resolution schemes for hyperbolic conservation laws

This article reviews stability and laminar-turbulent transition in high-speed boundary-layer flows, emphasizing qualitative features of the disturbance spectrum leading to new mechanisms of receptivity and instability. It is shown that the extension of subsonic and low-supersonic stability concepts and transition prediction methods to hypersonic speeds is not straightforward. The discussion focuses on theoretical models providing insights into the physics of instability and helping make proper decisions on transition control strategies.

Transition and Stability of High-Speed Boundary Layers

https://engineering.purdue.edu/~aae519/BAM6QT-Mach-6-tunnel/reviewpapers/cone-scram-blt-review-pas-feb04.pdf

Hypersonic laminar–turbulent transition on circular cones and scramjet forebodies

It is well known that the high levels of noise present in conventional hypersonic ground-test facilities cause transition to occur earlier than in e ight. Flight measurements of incoming noise are reviewed and compared with measurements in ground-test facilities, of both conventional and quiet design, at hypersonic and high supersonic speeds. The low noise present in e ight is apparently the reason for the very large transition Reynolds numbers sometimes measured in e ight, when roughness, crosse ow, and other factors are controlled. Design will usually involve consideration of the trend in transition when a parameter is varied. The effect of facility noise on these trends is reviewed. In some cases, the trend of conventional-tunnel data is opposite to the trend in quiet-tunnel data. Thus, transition measurements in conventional ground-test facilities are not reliable predictors of e ight performance, except perhaps in special cases.

Effects of High-Speed Tunnel Noise on Laminar-Turbulent Transition

This paper reviews the state of phosphor thermometry, focusing on developments in the past 15 years. The fundamental principles and theory are presented, and the various spectral and temporal modes, including the lifetime decay, rise time and intensity ratio, are discussed. The entire phosphor measurement system, including relative advantages to conventional methods, choice of phosphors, bonding techniques, excitation sources and emission detection, is reviewed. Special attention is given to issues that may arise at high temperatures. A number of recent developments and applications are surveyed, with examples including: measurements in engines, hypersonic wind tunnel experiments, pyrolysis studies and droplet/spray/gas temperature determination. They show the technique is flexible and successful in measuring temperatures where conventional methods may prove to be unsuitable.

https://www.mdpi.com/1424-8220/8/9/5673/pdf

Thermographic Phosphors for High Temperature Measurements: Principles, Current State of the Art and Recent Applications

An experimental investigation was conducted on a 5-degree half-angle cone and a 5-degree half-angle flared cone in a conventional Mach 6 wind tunnel to examine the effects of facility noise on boundary layer transition. The influence of tunnel noise was inferred by comparing transition onset locations determined from the present test to that previously obtained in a Mach 6 low disturbance quiet tunnel. Together, the two sets of experiments are believed to represent the first direct comparison of transition onset between a conventional and a low disturbance wind tunnel using a common test model and transition detection technique. In the present conventional hypersonic tunnel experiment, separate measurements of heat transfer and adiabatic wall temperatures were obtained on the conical models at small angles of attack over a range of Reynolds numbers, which resulted in laminar, transitional, and turbulent flow. Smooth model turbulent heating distributions are compared to that obtained with transition forced via discrete surface roughness. The model nosetip radius was varied to examine the effects of bluntness on transition onset. Despite wall to total temperature differences between the transient heating measurements and the adiabatic wall temperature measurement, the two methods for determining sharp cone transition onset generally yielded equivalent locations. In the 'noisy' mode of the hypersonic low disturbance tunnel, transition onset occurred earlier than that measured in the conventional hypersonic tunnel, suggesting higher levels of freestream acoustic radiation relative to the conventional tunnel. At comparable freestream conditions, the transition onset Reynolds number under low disturbance conditions was a factor of 1.3 greater than that measured on flared cone in the LaRC conventional hypersonic tunnel and a factor of 1.6 greater that the flared cone run in the low disturbance tunnel run 'noisy'. Navier-Stokes mean flow computations and linear stability analysis were conducted to assess the experimental results and have indicated N factors associated with sharp flared cone transition onset to be approximately a factor of 2 lower than that inferred from the corresponding low disturbance tunnel measurements.

/pdf/boundary-layer-transition-on-slender-cones-in-conventional-138983wspr.pdf

Boundary Layer Transition on Slender Cones in Conventional and Low Disturbance Mach 6 Wind Tunnels

Boundary layer and aeroheating characteristics of several X-33 configurations have been experimentally examined in the Langley 20-Inch Mach 6 Air Tunnel. Global surface heat transfer distributions, surface streamline patterns, and shock shapes were measured on 0.013-scale models at Mach 6 in air. Parametric variations include angles-of-attack of 20-deg, 30-deg, and 40-deg; Reynolds numbers based on model length of 0.9 to 6.6 million; and body-flap deflections of 0, 10 and 20-deg. The effects of discrete and distributed roughness elements on boundary layer transition, which included trip height, size, location, and distribution, both on and off the windward centerline, were investigated. The discrete roughness results on centerline were used to provide a transition correlation for the X-33 flight vehicle that was applicable across the range of reentry angles of attack. The attachment line discrete roughness results were shown to be consistent with the centerline results, as no increased sensitivity to roughness along the attachment line was identified. The effect of bowed panels was qualitatively shown to be less effective than the discrete trips; however, the distributed nature of the bowed panels affected a larger percent of the aft-body windward surface than a single discrete trip.

/pdf/x-33-hypersonic-boundary-layer-transition-3wj95wtf1s.pdf

X-33 Hypersonic Boundary-Layer Transition

The importance of discrete roughness and the correlations developed to predict the onset of boundary layer transition on hypersonic flight vehicles are discussed. The paper is organized by hypersonic vehicle applications characterized in a general sense by the boundary layer: slender with hypersonic conditions at the edge of the boundary layer, moderately blunt with supersonic, and blunt with subsonic. This paper is intended to be a review of recent discrete roughness transition work completed at NASA Langley Research Center in support of agency flight test programs. First, a review is provided of discrete roughness wind tunnel data and the resulting correlations that were developed. Then, results obtained from flight vehicles, in particular the recently flown Hyper-X and Shuttle missions, are discussed and compared to the ground-based correlations.

/pdf/discrete-roughness-transition-for-hypersonic-flight-vehicles-32ts51cmvp.pdf

Discrete Roughness Transition for Hypersonic Flight Vehicles

/pdf/discrete-roughness-transition-for-hypersonic-flight-vehicles-4k2m4sk6yr.pdf

Discrete-Roughness Transition for Hypersonic Flight Vehicles

This report describes a NASA Langley led endeavor sponsored by the NASA Engineering Safety Center, the Space Shuttle Program Office and the NASA Aeronautics Research Mission Directorate to demonstrate a quantitative thermal imaging capability. A background and an overview of several multidisciplinary efforts that culminated in the acquisition of high resolution calibrated infrared imagery of the Space Shuttle during hypervelocity atmospheric entry is presented. The successful collection of thermal data has demonstrated the feasibility of obtaining remote high-resolution infrared imagery during hypersonic flight for the accurate measurement of surface temperature. To maximize science and engineering return, the acquisition of quantitative thermal imagery and capability demonstration was targeted towards three recent Shuttle flights - two of which involved flight experiments flown on Discovery. In coordination with these two Shuttle flight experiments, a US Navy NP-3D aircraft was flown between 26-41 nautical miles below Discovery and remotely monitored surface temperature of the Orbiter at Mach 8.4 (STS-119) and Mach 14.7 (STS-128) using a long-range infrared optical package referred to as Cast Glance. This same Navy aircraft successfully monitored the Orbiter Atlantis traveling at approximately Mach 14.3 during its return from the successful Hubble repair mission (STS-125). The purpose of this paper is to describe the systematic approach used by the Hypersonic Thermodynamic Infrared Measurements team to develop and implement a set of mission planning tools designed to establish confidence in the ability of an imaging platform to reliably acquire, track and return global quantitative surface temperatures of the Shuttle during entry. The mission planning tools included a pre-flight capability to predict the infrared signature of the Shuttle. Such tools permitted optimization of the hardware configuration to increase signal-to-noise and to maximize the available dynamic range while mitigating the potential for saturation. Post flight, analysis tools were used to assess atmospheric effects and to convert the 2-D intensity images to 3-D temperature maps of the windward surface. Comparison of the spatially resolved global thermal measurements to surface thermocouples and CFD prediction is made. Successful demonstration of a quantitative, spatially resolved, global temperature measurement on the Shuttle suggests future applications towards hypersonic flight test programs within NASA, DoD and DARPA along with flight test opportunities supporting NASA's project Constellation.

/pdf/the-hythirm-project-flight-thermography-of-the-space-shuttle-5hmdoolkzv.pdf

Thomas J. Horvath

Papers

Boundary Layer Transition on Slender Cones in Conventional and Low Disturbance Mach 6 Wind Tunnels

X-33 Hypersonic Boundary-Layer Transition

Discrete Roughness Transition for Hypersonic Flight Vehicles

Discrete-Roughness Transition for Hypersonic Flight Vehicles

The HYTHIRM Project: Flight Thermography of the Space Shuttle During the Hypersonic Re-entry