Effects of Cone Tip Changes On Wall-Cooled Hypersonic Boundary Layer Transition and Turbulence

Boundary layer measurements took place at Air Force Research Lab’s (AFRL) Mach-6 Ludwieg tube hypersonic wind tunnel on a 7-degree half angle spherically blunted nose cone with a 1.5 mm radius to study the effects of a cooled surface on the transition process. Experiments compared uncooled and cooled flow conditions on the blunt-nosed model, and then compared to past data for a sharp-tipped cone of similar geometry. Cooling the surface delayed the onset of transition on the blunt nose model. Combining the effects of nose blunting and cooling further increased transition delay than each acting individually compared to the sharp-tipped case. This outcome provides much needed data in an area of study with inconclusive results

Shock-Induced Separation of Transitional Hypersonic Boundary Layers

This thesis presents a joint experimental/CFD investigation of shock-induced boundary layer separations in hypersonic transitional boundary layers with an emphasis on collapse and re-establishment times of the separation bubble. This study also provides high fidelity measurements and excellent characterisation of the flow field in order to provide benchmark data of a challenging flow configuration with which to benchmark next generation CFD solvers. The experiments were conducted in the Imperial College Aeronautics Department Number Two Gun Tunnel, a Mach 8.9 axisymmetric facility with a freestream unit Reynolds number of 47 million An axisymmetric blunt-nosed cylinder fitted with an 8 degree flare forms the primary vehicle for this study, although a 1.3 degree cowl geometry was also used to impinge a shock onto the blunt-nosed cylinder.. The shock boundary layer interaction was designed such that it was separated for a laminar boundary layer and collapsed for a turbulent one. Carefully controlled turbulent spots were generated upstream of the interaction region which passed through the separation causing its collapse and subsequent re-establishment. Two intermittency cases are considered, one where turbulent spot spacing is large and collapse/re-establishment pairs can be considered independent of each other and one where they can not. Experimental surface quantities through the interaction region are measured using either heat-transfer or pressure measurements and schlieren video is used to diagnose the larger shock structure. Further a non-intrusive toluene PLIF method is assessed for use in this facility and shows promise. CFD simulations are done using an in-house operator split Godunov solver with a Baldwin-Lomax turbulence model. CFD simulations show good agreement with experiment and provides information on flow quantities that would be extremely difficult to measure otherwise. Collapse times of the separation bubble were found to be fast in relation to characteristic spot passage times. The collapse process is also fast in relation to the surrounding flows ability to adjust, with collapse associated with significant shock curvature of the immediate outboard shock structures. This leads to unsteadiness, with surface pressure measurements exceeding the range bounded by the laminar separated and turbulent collapsed cases. The severity of the unsteadiness appears to be driven by turbulent spot spacing. Re-establishment is considerably slower, showing asymptotic recovery that is likely driven by viscous diffusion rates, taking many characteristic spot passage times to recover.Open Acces

Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels by means of a slender wedge probe and direct numerical simulation

Combined free-stream disturbance measurements and receptivity studies in hypersonic wind tunnels were conducted by means of a slender wedge probe and direct numerical simulation. The study comprises comparative tunnel noise measurements at Mach 3, 6 and 7.4 in two Ludwieg tube facilities and a shock tunnel. Surface pressure fluctuations were measured over a wide range of frequencies and test conditions including harsh test environments not accessible to measurement techniques such as Pitot probes and hot-wire anemometry. A good agreement was found between normalized Pitot pressure fluctuations converted into normalized static pressure fluctuations and the wedge probe readings. Quantitative results of the tunnel noise are provided in frequency ranges relevant for hypersonic boundary-layer transition. Complementary numerical simulations of the leading-edge receptivity to fast and slow acoustic waves were performed for the applied wedge probe at conditions corresponding to the experimental free-stream conditions. The receptivity to fast acoustic waves was found to be characterized by an early amplification of the induced fast mode. For slow acoustic waves an initial decay was found close to the leading edge. At all Mach numbers, and for all considered frequencies, the leading-edge receptivity to fast acoustic waves was found to be higher than the receptivity to slow acoustic waves. Further, the effect of inclination angles of the acoustic wave with respect to the flow direction was investigated. An inclination angle was found to increase the response on the wave-facing surface of the probe and decrease the response on the opposite surface for fast acoustic waves. A frequency-dependent response was found for slow acoustic waves. The combined numerical and experimental approach in the present study confirmed the previous suggestion that the slow acoustic wave is the dominant acoustic mode in noisy hypersonic wind tunnels.The present study was supported by an ESA funded Technology Research Project (ESA-Contract number $4200022793/09/NL/CP$).Published versio

Stability of High-Speed, Three-Dimensional Boundary Layers

Boundary-layer experiments are performed in the low-disturbance, Mach 6 Quiet Tunnel (M6QT) at Texas A&M University. The experiments are focused specifically on investigating the physics of two three-dimensional phenomena in hypersonic boundary-layer stability and transition: the breakdown of second-mode waves and the growth and breakdown of crossflow waves. In order to enable these experiments, a new, three-dimensional probe traversing mechanism was designed and constructed. In order to investigate the breakdown of second-mode waves, experiments are conducted on a flared cone with a 5° half angle at the tip at zero angle of attack. Experiments were typically performed at unit Reynolds number Re’ ≈ 10 × 10^6/m with a slightly hot wall, T/Taw ≈ 1.05. A new, durable method of roughness element application is discussed for the purpose of exciting the unstable waves. Hot-wire measurements were made of the boundary layer and it is shown that even with roughness elements, transition to turbulence does not occur on the model. Therefore, the expected Λ vortices are not observed. The crossflow instability in a hypersonic boundary layer is studied on a 7° right circular cone at 5.6° angle of incidence. Experiments were performed at Re’ ≈ 10 × 10^6/m with an adiabatic wall. Hot wire measurements are made at a series of axial locations to generate contours of streamwise mass flux. The stationary vortex structure is shown through its saturation. Traveling waves are observed in the expected frequency range, 10 kHz to 60 kHz, predicted by computations and are located generally in the high-speed troughs in the vortex structure. Secondary instability is observed between 80 kHz and 130 kHz. Frequency scaling and location is shown to agree with low-speed experiments and good preliminary agreement with hypersonic computations is obtained. Transition does not naturally occur on the model. Distributed roughness is applied to the tip in order to excite crossflow and cause transition. Transition is shown to occur with the rough tip, but is not likely a result of crossflow

Molecular-Based Optical Measurement Techniques for Transition and Turbulence in High-Speed Flow

High-speed laminar-to-turbulent transition and turbulence affect the control of flight vehicles, the heat transfer rate to a flight vehicle's surface, the material selected to protect such vehicles from high heating loads, the ultimate weight of a flight vehicle due to the presence of thermal protection systems, the efficiency of fuel-air mixing processes in high-speed combustion applications, etc. Gaining a fundamental understanding of the physical mechanisms involved in the transition process will lead to the development of predictive capabilities that can identify transition location and its impact on parameters like surface heating. Currently, there is no general theory that can completely describe the transition-to-turbulence process. However, transition research has led to the identification of the predominant pathways by which this process occurs. For a truly physics-based model of transition to be developed, the individual stages in the paths leading to the onset of fully turbulent flow must be well understood. This requires that each pathway be computationally modeled and experimentally characterized and validated. This may also lead to the discovery of new physical pathways. This document is intended to describe molecular based measurement techniques that have been developed, addressing the needs of the high-speed transition-to-turbulence and high-speed turbulence research fields. In particular, we focus on techniques that have either been used to study high speed transition and turbulence or techniques that show promise for studying these flows. This review is not exhaustive. In addition to the probe-based techniques described in the previous paragraph, several other classes of measurement techniques that are, or could be, used to study high speed transition and turbulence are excluded from this manuscript. For example, surface measurement techniques such as pressure and temperature paint, phosphor thermography, skin friction measurements and photogrammetry (for model attitude and deformation measurement) are excluded to limit the scope of this report. Other physical probes such as heat flux gauges, total temperature probes are also excluded. We further exclude measurement techniques that require particle seeding though particle based methods may still be useful in many high speed flow applications. This manuscript details some of the more widely used molecular-based measurement techniques for studying transition and turbulence: laser-induced fluorescence (LIF), Rayleigh and Raman Scattering and coherent anti-Stokes Raman scattering (CARS). These techniques are emphasized, in part, because of the prior experience of the authors. Additional molecular based techniques are described, albeit in less detail. Where possible, an effort is made to compare the relative advantages and disadvantages of the various measurement techniques, although these comparisons can be subjective views of the authors. Finally, the manuscript concludes by evaluating the different measurement techniques in view of the precision requirements described in this chapter. Additional requirements and considerations are discussed to assist with choosing an optical measurement technique for a given application

Aeronautical engineering: A continuing bibliography with indexes (supplement 269)

This bibliography lists 539 reports, articles, and other documents introduced into the NASA scientific and technical information system in August, 1991. Subject coverage includes: design, construction and testing of aircraft and aircraft engines; aircraft components, equipment and systems; ground support systems; and theoretical and applied aspects of aerodynamics and general fluid dynamics

Aeronautical engineering: A continuing bibliography with indexes (supplement 279)

This bibliography lists 759 reports, articles, and other documents introduced into the NASA scientific and technical information system in May 1992. Subject coverage includes: design, construction, and testing of aircraft and aircraft engines; aircraft components, equipment, and systems; ground support systems; and theoretical and applied aspects of aerodynamics and general fluid dynamics

The Fifth Symposium on Numerical and Physical Aspects of Aerodynamic Flows

This volume contains the papers presented at the Fifth Symposium on Numerical and Physical Aspects of Aerodynamic Flows, held at the California State University, Long Beach, from 13 to 15 January 1992. The symposium, like its immediate predecessors, considers the calculation of flows of relevance to aircraft, ships, and missiles with emphasis on the solution of two-dimensional unsteady and three-dimensional equations

Characterisation of the USQ hypersonic facility freestream

Hypersonic ground testing can make significant contributions to the design process for hypersonic flight vehicles. However, experimentation in conventional hypersonic ground testing facilities is complicated by the high levels of freestream fluctuations which are typically one-to-two orders of magnitude greater than in flight. This noisy test environment can have a significant impact on flow phenomena, such as boundary layer transition, and this leads directly to uncertainties in the prediction of essential hypersonic vehicle design parameters. To assess the noise level in ‘TUSQ’, the hypersonic wind tunnel at the University of Southern Queensland, the Mach 6 nozzle exit flow was characterised by measurements which provided: (1) the time-averaged and fluctuating components of Pitot pressure; (2) the time-averaged and fluctuating components of stagnation temperature; and (3) the fluctuating component of density. The Pitot pressure measurements were made using Kulite XTL-190M B-screen pressure transducers which were exposed directly to the flow. The stagnation temperature was determined from the experimental measurement of heat flux using microsecond response time coaxial surface junction thermocouples mounted in a stagnation point heat transfer gauge. A focused laser differential interferometer was designed for TUSQ, and this instrument was used to measure the freestream density fluctuations. Using the Pitot pressure measurements and the measurements of the stagnation pressure in the nozzle reservoir (the barrel), the Mach number was found to decrease over the flow duration from 5.95 to 5.85. Through the measurement of stagnation temperature, the piston compression and the nozzle expansion of the test gas were found to be approximately isentropic for the first 65ms of hypersonic flow. Thereafter, the stagnation temperature reduces due to the heat lost to the cold barrel. Thermodynamic modelling based on the measured pressure history in the barrel combined with empirical heat transfer correlations can be used to simulate the stagnation temperature in TUSQ to within 2% of the actual value for t=0_150ms, increasing to within 5% at t=170ms. The heat transfer process in the barrel was found to significantly affect the fluctuations in the hypersonic freestream. For t less than 65ms, the freestream fluctuations of Pitot pressure, stagnation temperature and density were found to be broadband in nature, consistent with a disturbance environment dominated by the radiation of acoustic noise from the turbulent boundary layer on the nozzle walls. At t approximately equal to 65ms, a 3_4 kHz narrowband disturbance was detected in the barrel and in the freestream flow, and this disturbance remains superimposed on the broadband disturbance environment for the remainder of the test flow. Because the characteristics of the flow changed during the run, it is appropriate to specify two RMS Pitot pressure fluctuation magnitudes in the 300 Hz to 25 kHz bandwidth: 2.52% for t=5_65ms; and 2.86% for t=65_200ms for Reu = 6.94 x 106 m-1. The RMS Pitot pressure fluctuations in the TUSQ freestream are similar to comparable Ludwieg and blowdown facilities. RMS stagnation temperature fluctuations were resolved for f=4 Hz_5kHz and were found to increase throughout the flow period from approximately 1.5% at the start of the run to 2.4% at the termination of the nozzle flow. RMS freestream density fluctuations were determined for f=1_250kHz, increasing from 0.4% to 0.6% over he flow period. The bandwidth of the density fluctuation measurement was sufficient to resolve the classic Kolmogorov _5/3 rolloff in the inertial subrange. Preliminary measurements of the boundary layer on a conical nose cylinder were made sing the focused laser differential interferometer. These measurements identified the second mode instabilities in the transitional boundary layer, and identified the amplification f the narrowband 3_4 kHz freestream fluctuations within the boundary layer. Further opportunities to explore boundary layer transition in the TUSQ facility are expected o arise in the near future, at which time the FLDI instrument can be deployed with improved focusing ability

An Experimental Characterization of 3-D Transitional Shock Wave Boundary Layer Interactions at Mach 6

Hypersonics is of current national interest, but improved understanding of fundamental flow physics is required for safe and efficient vehicle design. The objective of the present study was to expand the knowledgebase of shock wave boundary layer interaction flows during transition at a high Mach number since these interactions are likely to occur on hypersonic aircraft. The approach was to experimentally determine how the dynamics of the shock structure, the fluctuations within it, and the resulting thermal and acoustic loads, change as the flow evolves through its transitional regime. Tests were conducted on a canonical cylinder-induced 3-D shock wave boundary layer interaction geometry at Mach 5.8 in the Actively Controlled Expansion hypersonic wind tunnel. The model was tested in different configurations to isolate the effects of the boundary layer trips and the shock generator. The interaction excited a 40 kHz (possibly second mode) in-stability, causing transition just downstream of the separation shock. A transitional boundary layer was only achieved on the baseline model with trips at Re=7M/m, which demonstrated that a transitional incoming boundary layer is not required to produce a transition interaction. Time-resolved schlieren imaging revealed disturbances emerging from the supersonic jet and ascending the cylinder with a characteristic frequency near 20 kHz. The separation shock motion frequency was O(1 kHz) and was fed by disturbances originating near the base of the cylinder. The film coefficient was found to be the heat transfer parameter of interest since it scaled roughly linearly with Reynolds number. It revealed fundamental differences in heating at the reattachment arc for configurations with and without trips and indicated higher heating in that region for a laminar SBLI. Cylinder sweep had an impact on fluctuation levels and thermal loads. Sweeping the cylinder back 15 degrees significantly reduced the extent of the interaction and dropped the RMS pressure fluctuations and heating loads at the base of the cylinder by roughly 50%. Alternatively, sweeping the cylinder forward 15 degrees led to fluctuations on the order of the freestream static pressure and the highest heating levels observed for this campaign