Hypersonic laminar/turbulent transition: computations and experiments

Journal of Thermal Science, 2005

The results of experimental investigation of laminar-turbulent transition in three-dimensional flow under the high continuous pressure gradient including the flow with local boundary layer separation are presented. The experimental studies were performed within the Mach number range from 4 to 6 and Reynolds number 10~60 X 10 6 l/m, the angles of attack were 0 ° and 5 °. The experiments were carded out on the three-dimensional convergent inlet model with and without sidewalls. The influence of artificial turbulator of boundary layer on transition and flow structure was studied. The conducted researches have shown that adverse pressure gradient increase hastens transition and leads to decrease of transition area length. If pressure gradient rises velocity profile fullness increases and profile transformation from laminar to turbulent occurs. As a result of it the decrease of separation area length occurs. The same effect was reached with Reynolds number increase. These results are compared with the data on two-dimensional model with longitudinal curvature.

2021

Linear stability analysis is performed using a combination of two-dimensional Direct Simulation Monte Carlo (DSMC) method for the computation of the basic state and solution of the pertinent eigenvalue problem, as applied to the canonical boundary layer on a semi-infinite flat plate. Three different gases are monitored, namely nitrogen, argon and air, the latter as a mixture of 79\% Nitrogen and 21\% Oxygen at a range of free-stream Mach numbers corresponding to flight at an altitude of 55km. A neural network has been utilised to predict and smooth the raw DSMC data; the steady laminar profiles obtained are in very good agreement with those computed by (self-similar) boundary layer theory, under isothermal or adiabatic wall conditions, subject to the appropriate slip corrections computed in the DSMC method. The leading eigenmode results pertaining to the unsmoothed DSMC profiles are compared against those of the classic boundary layer theory. Small quantitative, but no significant q...

2013

A probabilistic approach, based on Uncertainty Quantification, is combined to deterministic numerical simulations relying on Linear Stability Theory and Reynolds Averaged Navier-Stokes. The objective is to improve the prediction of natural transition in hypersonic flows. Results are compared to experimental data obtained in the VKI-H3 facility focusing on the transition onset location and on the extension of the transitional zone. Comparisons show that the computed intermittency distribution is able to capture the experimental onset of transition well following the experimental data within the transitional zone. Finally, a preliminary investigation on the distribution of the perturbation spectrum upstream of the transition onset is also performed with satisfying results.

Journal of Spacecraft and Rockets, 2018

While low disturbance ("quiet") hypersonic wind tunnels are believed to provide more reliable extrapolation of boundary layer transition behavior from ground to flight, the presently available quiet facilities are limited to Mach 6, moderate Reynolds numbers, low freestream enthalpy, and subscale models. As a result, only conventional ("noisy") wind tunnels can reproduce both Reynolds numbers and enthalpies of hypersonic flight configurations, and must therefore be used for flight vehicle test and evaluation involving high Mach number, high enthalpy, and larger models. This article outlines the recent progress and achievements in the characterization of tunnel noise that have resulted from the coordinated effort within the AVT-240 specialists group on hypersonic boundary layer transition prediction. New Direct Numerical Simulation (DNS) datasets elucidate the physics of noise generation inside the turbulent nozzle wall boundary layer, characterize the

Progress in Aerospace Sciences, 2006

Turbulence modeling remains a major source of uncertainty in the computational prediction of aerodynamic forces and heating for hypersonic vehicles. Dodson developed a methodology for assessing experiments appropriate for turbulence model validation and critically surveyed the existing hypersonic experimental database. We limit the scope of our current effort by considering only twodimensional/axisymmetric flows in the hypersonic speed regime where calorically perfect gas models are appropriate. We extend the prior database of recommended hypersonic experiments by adding three new cases. The first two cases, the flat plate/cylinder and the sharp cone, are canonical test cases which are amenable to theory-based correlations, and these correlations are discussed in detail. The third case added is the two-dimensional shock impinging on a flat plate boundary layer. The second goal is to review and assess the validation usage of various turbulence models on the existing experimental database. Here we limit the scope to one-and two-equation turbulence models where integration to the wall is used (i.e., we omit studies involving wall functions). In order to preserve a models prior validation history, we omitted corrections to the standard turbulence models in cases where the impact of such corrections on low-speed flows had not been adequately addressed (either through a re-validation of the models on a wide range of low-speed test cases or theoretical arguments). A methodology for validating turbulence models is given, and turbulence model comparisons from various authors are compiled and presented in graphical form. Conclusions are drawn for those models which have been applied to a sufficiently wide range of two-dimensional/axisymmetric hypersonic flows, and recommendations for future experimental and modeling efforts are given. ongoing research efforts in advanced turbulence models such as Reynolds stress models and large eddy simulation, the most complex models currently employed in design studies (where a large number of parametric cases must be considered) are one-and two-equation turbulence models. We therefore limit the current study to these models. We also limit this study to models where integration of the governing equations to the wall is performed, thereby eliminating the use of wall functions. This choice was primarily driven by the fact that a majority of the cases of interest for hypersonic flows include shock-boundary layer interactions, where the assumptions inherent in the use of wall functions are difficult to justify. We further limit our scope to cases where the transition from laminar to turbulent flow occurs naturally, and where this transition location is specified in the experimental description. The focus here is not on the prediction of transition, which itself is a difficult challenge for hypersonic flows. Finally, the effects of surface roughness, ablation, chemical reactions, real gases, and body rotation are all neglected as the existing experimental database does not yet adequately address these phenomena. Typically turbulence models have been developed for incompressible flows and then extended without much change to compressible flows. This approach in many cases is not adequate. For complex turbulent flows, Coakley et al. [86] have recommended corrections to apply to the twoequation and turbulent eddy viscosity models. In addition, Aupoix and Viala [87] have proposed corrections to the model for compressible flows. The authors have used flat plate flows and mixing layers to assess the compressible corrections introduced. Significant efforts to assess turbulence models for compressible flows have occurred at NASA Ames Research Center. The results of these investigations have been published by Horstman [88], Horstman [89], Coakley and Huang [84], Huang and Coakley [90], Coakley et al.[86], Bardina et al. [91], and Bardina et al. [92]. See Appendix A for additional discussion of the compressibility effects for hypersonic flows. 2. Turbulence models 2.1. One-equation models (eddy-viscosity transport models) 2.1.1. Spalart-Allmaras (SA) A transport equation for determining the eddy viscosity with near-wall effects included has been developed by Spalart and Allmaras [7], [8]. The accuracy of the predictions with the Spalart-Allmaras model is fairly insensitive to the y + spacing at the wall relative to the two-equation models, at least for high-speed flows [9]. Our experience with this model suggests that it has a good combination of accuracy and robustness for attached flows. While stable for large y + values, the maximum for accurate solutions should be roughly y + ≤ 1. 2.1.2. Goldberg (UG) Goldberg has developed a one-equation turbulence model [158], [97]. 2.1.3. Menter one-equation model (MTR) Menter has developed a one-equation turbulence model [6]. 2.2. Two-equation models 2.2.1. Jones and Launder high Reynolds number k-ε (kεJL) The basic k-ε model was developed by Jones and Launder [10] in 1972, and is valid for high Reynolds number flows only. Damping functions or wall functions must be used in order to handle wall-bounded flows.

33rd Aerospace Sciences Meeting and Exhibit, 1995

their strength at yet higher temperatures. Failing this, the designers must add heat shielding, further weighing The nonparallel effects on the stability of the flow over the vehicle down, a Mach 8, 7 ' half-angle cone at 2' angle of attack are The prediction of transition to turbulence of hyperinvestigated with the Parabolized Stability Equations. sonic flows is a daunting task. While much progress The results are compared to those from local stability has been made in the of transition in analyses. Both analyses are performed with two difsubsonic floWs12, the methods used are difficult to exferent sets of flow variables. It is found that the local tend to hypersonic flows. analyses produce different results for different Sets of M~~~ of these difficulties are experimental in navariables when nonparallel terms are evaluated locally. ture, Controlled disturbances in subsonic boundary This is an artifact of the local approximation. The PSE layers can be excited with vibrating ribbons, The die analyses converge to the same solution downstream inturbances can then be carefully measured as they grow dependent of the variables chosen. The nonparallel efand evolve downstream. This is much difficult to fects for this flow are thus very significant but can be do in hypersonic flows. Rypersonic test facilities instead modeled accurately with the PSE. are often plagued with noise generated by the turbulent boundary layers on the wind tunnel walls. This excites a broad and unknown spectrum of disturbances in the

Wall temperature variations significantly influence the stability and transition dynamics of hypersonic boundary layers. This study numerically investigates the stability response to thermal gradients, with wall temperatures ranging from 262 K to 362 K, for double-cone and cone-cylinder-flare (CCF) geometries. The CCF geometry, modeled using data from the Research Task Group NATO-AVT-346, consists of a 5°half-angle cone followed by a 147.3 mm cylindrical section and a 12°flare, with the sharp-nosed variant (0.1 mm nose radius) analyzed to emphasize instability mechanisms such as the first and second Mack modes. For the double cone, modeled after the ROTEX-T flight-test vehicle (457.31 mm in length, 1 mm nose radius, and 7°half-angle) at Mach 5.7, flow conditions were representative of the AFOSR-Notre Dame Large Mach 6 Quiet Wind Tunnel. Heating the entire wall stabilized second-mode instabilities while amplifying the first mode, particularly within the separation bubble. In contrast, for the CCF geometry, wall heating upstream of the separation bubble consistently stabilized both first-and second-mode instabilities throughout the flowfield. Numerical stability analyses were conducted using the JoKHeR Parabolized Stability Equations (PSE) code. These results underscore the geometry-dependent nature of thermal stabilization strategies and suggest that targeted upstream heating can effectively suppress boundary layer instabilities and delay transition in hypersonic flows.

2014

This paper details five different projects in the Boeing/AFOSR Mach 6 Quiet Tunnel (BAM6QT) at Purdue University. In the first project, a highly swept fin on a 7 • half-angle cone was examined. Temperature sensitive paint and pressure fluctuation measurements showed the presence of an instability near 200 kHz. In the second, the effect of roughness on the stationary and travelling crossflow instabilities was investigated on a cone at an angle of attack. It was found that when the roughness generated larger stationary waves, the travelling waves were damped. In the third, discrete roughness elements made of epoxy were applied to a flared cone. The roughness elements were sized to interact with the second-mode waves without becoming a trip. Using this technique, the spacing of streaks of increased heating was controlled. The fourth project involves improving the current pulsed jet perturber used to study nozzle-wall boundary-layer perturbations. Two improvements have been designed and have undergone preliminary testing. Finally, a new model was designed to measure what appears to be an entropy-layer instability at Mach 6. Surface pressure measurements and hot-wire data are used to investigate the entropy layer.

Journal of Spacecraft and Rockets, 2001

Hypersonic transitional flows over a flat plate and a sharp cone are studied using four turbulence models: the one-equation eddy viscosity transport model of Spalart-Allmaras, a low Reynolds number k-ε model, the Menter k-ω model, and the Wilcox k-ω model. A framework is presented for the assessment of turbulence models that includes documentation procedures, solution accuracy, model sensitivity, and model validation. The accuracy of the simulations is addressed, and the sensitivities of the models to grid refinement, freestream turbulence levels, and wall y + spacing are presented. The flat plate skin friction results are compared to the well-established laminar and turbulent correlations of Van Driest. Correlations for the sharp cone are discussed in detail. These correlations, along with recent experimental data, are used to judge the validity of the simulation results for skin friction and surface heating on the sharp cone. The Spalart-Allmaras performs the best with regards to model sensitivity and model accuracy, while the Menter k-ω model also performs well for these zero pressure gradient boundary layer flows.

Procedia IUTAM, 2015

The laminar-turbulent transition of a Mach 4.6 flat-plate boundary layer forced by wall injection is investigated using Implicit Large Eddy Simulation. The boundary layer edge conditions match those of a 1:3 scale hypersonic vehicle forebody at Mach number 6 in the T-313 blow-down wind tunnel of ITAM-Russian Academy of Science. Two different injection pressures are considered. The breakdown to turbulence is observed through instantaneous Q-criterion visualizations and wall pressure fluctuations. Global instability analysis is performed by the computation of Koopman modes from 2D pressure and density snapshots of the ILES flow field. Some coherence is found in the spatial structures and temporal frequencies of global modes computed on two different grids, indicating a weak dependency on grid resolution. Physical insight is gained, confirming different conjectured transition mechanisms for the different injection pressures.

53rd AIAA Aerospace Sciences Meeting, 2015

A summary of the activities performed over the last years at the von Karman Institute for Fluid Dynamics in the frame of hypersonic boundary layer transition studies is presented. Free-stream noise levels have been determined in the H3 Mach 6 conventional wind tunnel using double hot-wires and modal analysis. In the Longshot wind tunnel at Mach 10, an improved free-stream characterization method, based on the use of free-stream static pressure probes, has been applied, alleviating the needs for the limiting adiabatic and isentropic nozzle flow assumptions. Based on these improved flow characterization, natural transition experiments have been performed in both wind tunnels on 7 • half-angle conical geometries at 0 • angle of attack and with different nosetip radii. Measurements techniques include either infrared thermography or flush-mounted fast response thermocouples in order to determine the transition onset location. Boundary layer instabilities are visualized using a LIF-based Schlieren technique at Mach 10, revealing rope-shape structures typical of the second mode disturbances. Wall measurements using fast-response pressure sensors complete the investigations. Dominant boundary layer disturbances at various locations along the cone are determined and compared with theoretical predictions. The corresponding N-factor is inferred for each wind tunnel. A comparison of the different measurement techniques is finally reported.

2012 IEEE International Conference on Computational Intelligence and Computing Research, 2012

This paper presents the effect of temperature inside the settling chamber of a hypersonic wind tunnel which is a test facility used to study the effect of aerodynamic forces on the specimen under test. Regulation of pressure inside the settling chamber is a very important task for the efficient operation of the tunnel system. The effects of aerodynamic forces and temperature are to be considered for the flight vehicle design and to be tested in the test section of the tunnel system. Here the stability analysis and the open loop response of the nonlinear model of the hypersonic wind tunnel is determined. The response of the settling chamber pressure incorporating the variation in temperature in all three vessels, high pressure vessel, heater and settling chamber is also obtained.

52nd Aerospace Sciences Meeting, 2014

During the descent phase of the transition flight experiment HIFiRE-1 the angle of attack was higher than expected, since an anomaly occurred in the exoatmospheric pointing maneuver. All pre-flight ground tests were carried out at angles of attack below 6 •. Therefore several post-flight experiments at high angles of attack were performed in the hypersonic wind tunnel (H2K) of the German Aerospace Center in Cologne. The selected Mach number of 7 and the Reynolds number range cover the flow conditions of the flight phase which are relevant for the transition experiment. The test campaign included highfrequency surface pressure measurements with PCB R sensors and heat flux measurements by means of quantitative infrared thermography. The results show significant effects of the emerging crossflow vortices on the transition location and the pressure fluctuations in the boundary layer.

48th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, 2007

A review of the state-of-the-art in hypersonic aeroelasticity and aerothermoelasticity is provided. Recently, the main focus in this area has been on the development of computational aeroelastic and aerothermoelastic methods capable of studying complete hypersonic vehicles. Thus, in addition to a survey of studies conducted in this area over the past six decades, two important issues are a focus of this review, namely: 1) modeling unsteady hypersonic aerodynamics; and 2) incorporation of the heat transfer between the fluid and the structure into the aeroelastic solution process. Finally, future directions of hypersonic aeroelasticity and aerothermoelasticity are outlined. Since air-breathing hypersonic vehicles exhibit strong coupling between the airframe, propulsion, and control systems, future directions point to the incorporation of advanced computational aerothermoelastic methods into a comprehensive vehicle analysis.

2007

Prediction of heat load on the surface of vehicles (re-) entering a planetary atmosphere is important for heat-shield design. Turbulent flow induces a much higher heating in comparison to laminar flow. Therefore, prediction of the location of laminar-turbulent transition is a key factor in defining the dimensions and materials used for the thermal protection system (TPS). Yet, the fundamental physical processes related to the laminarturbulent transition in high-speed boundary layers are not well understood.