Showing papers on "Freestream published in 2008"

Investigation of the vortex induced unsteadiness of a separation bubble via time-resolved and scanning PIV measurements

Abstract: A transitional separation bubble on the suction side of an SD7003 airfoil is considered. The transition process that forces the separated shear layer to reattach seems to be governed by Kelvin–Helmholtz instabilities. Large scale vortices are formed due to this mechanism at the downstream end of the bubble. These vortices possess a three-dimensional structure and detach from the recirculation region, while other vortices are formed within the bubble. This separation of the vortex is a highly unsteady process, which leads to a bubble flapping. The structure of these vortices and the flapping of the separation bubble due to these vortices are temporally and spatially analyzed at angles of attack from 4° to 8° and chord-length based Reynolds numbers Rec = 20,000–60,000 using time-resolved PIV measurements in a 2D and a 3D set-up, i.e., stereo-scanning PIV measurements are done in the latter case. These measurements complete former studies at a Reynolds number of Rec = 20,000. The results of the time-resolved PIV measurements in a single light-sheet show the influence of the angle of attack and the Reynolds number. The characteristic parameters of the separation bubble are analyzed focusing on the unsteadiness of the separation bubble, e.g., the varying size of the main recirculation region, which characterizes the bubble flapping, and the corresponding Strouhal number are investigated. Furthermore, the impact of the freestream turbulence is investigated by juxtaposing the current and former results. The stereo-scanning PIV measurements at Reynolds numbers up to 60,000 elucidate the three-dimensional character of the vortical structures, which evolve at the downstream end of the separation bubble. It is shown that the same typical structures are formed, e.g., the c-shape vortex and the screwdriver vortex at each Reynolds number and angle of attack investigated and the occurrence of these patterns in relation to Λ-structures is discussed. To evidence the impact of the freestream turbulence, these results are compared with findings of former measurements.

Turbulence Model Behavior in Low Reynolds Number Regions of Aerodynamic Flowfields

Abstract: The behaviors of the widely-used Spalart-Allmaras (SA) and Menter shear-stress transport (SST) turbulence models at low Reynolds numbers and under conditions conducive to relaminarization are documented. The flows used in the investigation include 2-D zero pressure gradient flow over a flat plate from subsonic to hypersonic Mach numbers, 2-D airfoil flow from subsonic to supersonic Mach numbers, 2-D subsonic sink-flow, and 3-D subsonic flow over an infinite swept wing (particularly its leading-edge region). Both models exhibit a range over which they behave 'transitionally' in the sense that the flow is neither laminar nor fully turbulent, but these behaviors are different: the SST model typically has a well-defined transition location, whereas the SA model does not. Both models are predisposed to delayed activation of turbulence with increasing freestream Mach number. Also, both models can be made to achieve earlier activation of turbulence by increasing their freestream levels, but too high a level can disturb the turbulent solution behavior. The technique of maintaining freestream levels of turbulence without decay in the SST model, introduced elsewhere, is shown here to be useful in reducing grid-dependence of the model's transitional behavior. Both models are demonstrated to be incapable of predicting relaminarization; eddy viscosities remain weakly turbulent in accelerating or laterally-strained boundary layers for which experiment and direct simulations indicate turbulence suppression. The main conclusion is that these models are intended for fully turbulent high Reynolds number computations, and using them for transitional (e.g., low Reynolds number) or relaminarizing flows is not appropriate.

Effect of Freestream Noise on Roughness-Induced Transition for the X-51A Forebody

Abstract: A 20%-scale X-51A forebody model was tested in the Boeing/AFOSR Mach-6 Quiet Tunnel. Repolishing the nozzle throat has restored quiet flow at high Reynolds numbers. The effect of a smooth blank and two different trip strips on windward-forebody transition was measured using temperature-sensitive paint and hot-wire anemometry. Reducing freestream noise from conventional to quiet levels increased the smooth-wall transition Reynolds number by a factor of at least 2.2. In addition, the transition Reynolds number based on the distance from the trips increased by a factor of 2.4 for the smaller trips and by a factor of 1.7 for the larger trips. Thus, tunnel noise had a substantial effect on roughness-induced transition.

Direct Numerical Simulation Of Hypersonic Boundary-Layer Transition Over A Blunt Cone

Abstract: Direct numerical simulation of transition How over a blunt cone with a freestream Mach number of 6, Reynolds number of 10,000 based on the nose radius, and a 1-deg angle of attack is performed by using a seventh-order weighted essentially nonoscillatory scheme for the convection terms of the Navier-Stokes equations, together with an eighth-order central finite difference scheme for the viscous terms. The wall blow-and-suction perturbations, including random perturbation and multifrequency perturbation, are used to trigger the transition. The maximum amplitude of the wall-normal velocity disturbance is set to 1% of the freestream velocity. The obtained transition locations on the cone surface agree well with each other far both cases. Transition onset is located at about 500 times the nose radius in the leeward section and 750 times the nose radius in the windward section. The frequency spectrum of velocity and pressure fluctuations at different streamwise locations are analyzed and compared with the linear stability theory. The second-mode disturbance wave is deemed to be the dominating disturbance because the growth rate of the second mode is much higher than the first mode. The reason why transition in the leeward section occurs earlier than that in the windward section is analyzed. It is not because of higher local growth rate of disturbance waves in the leeward section, but because the growth start location of the dominating second-mode wave in the leeward section is much earlier than that in the windward section.

On the unsteady inviscid force on cylinders and spheres in subcritical compressible flow.

Abstract: The unsteady inviscid force on cylinders and spheres in subcritical compressible flow is investigated. In the limit of incompressible flow, the unsteady inviscid force on a cylinder or sphere is the so-called added-mass force that is proportional to the product of the mass displaced by the body and the instantaneous acceleration. In compressible flow, the finite acoustic propagation speed means that the unsteady inviscid force arising from an instantaneously applied constant acceleration develops gradually and reaches steady values only for non-dimensional times c(infinity)t/R approximately >10, where c(infinity) is the freestream speed of sound and R is the radius of the cylinder or sphere. In this limit, an effective added-mass coefficient may be defined. The main conclusion of our study is that the freestream Mach number has a pronounced effect on both the peak value of the unsteady force and the effective added-mass coefficient. At a freestream Mach number of 0.5, the effective added-mass coefficient is about twice as large as the incompressible value for the sphere. Coupled with an impulsive acceleration, the unsteady inviscid force in compressible flow can be more than four times larger than that predicted from incompressible theory. Furthermore, the effect of the ratio of specific heats on the unsteady force becomes more pronounced as the Mach number increases.

Effect of freestream noise on roughness-induced transition at mach 6

Abstract: The effects of freestream noise on roughness-induced transition were measured for the HIFiRE forecone model. Temperature-sensitive paint measurements were used to visualize the wake of the roughness elements under quiet and noisy flow. The transition Reynolds number increased under quiet flow, for a less-than-effective trip, by a factor of up to 6.4 when referenced to the trip location. The difference in transition location between quiet and noisy flow conditions is significantly reduced when effective trips are used, but quiet flow still delays the transition location for an effective trip. STABL calculations of the flow over the model at 0 and 6 angle of attack were used to quantify the results. The effect of an isolated roughness on the nozzle wall of the BAM6QT is also being studied. The nozzle wall features a thick laminar boundary layer, allowing measurements of the growth of instabilities. Temperature-sensitive paint measurements of the wake behind an isolated roughness element were obtained. In quiet flow, several hot streaks are visible.

Investigation of Supersonic Combustion with Angled Injection in a Cavity-Based Combustor

Abstract: Self-ignited supersonic combustion experiments have been performed using a cavity-based injector in the T3 free-piston shock tunnel, using various combustor inlet and fuel-flow conditions. Planar laser-induced fluorescence on the hydroxyl radical and fast-acting pressure transducers are used to investigate the flow characteristics. Four hydrogen injectors are located upstream of an open cavity. The separated shear layer reattaches, generating an oblique shock at the cavity's trailing edge and establishing the major flow structure. The normalized pressure rise due to combustion increases as the equivalence ratio increases and the freestream stagnation enthalpy decreases, over the range of conditions' tested. Angled injection upstream of the cavity allows the cavity to act as a flame holder. High injection pressure helps to ignite immediately upstream of the injector and forms two flame layers over the cavity. The fluorescence peak signal shows periodic maxima near the cavity, and the interval between peaks decreases as the equivalence ratio is increased. Low-total-enthalpy conditions also exhibit longer ignition-delay distances. Comparison of fluorescence images and static pressure measurements indicates that, at these conditions, the heat release is mostly initiated by the shock wave from the cavity's trailing face and the ignition above the cavity does not have a strong influence on the downstream combustion.

Large-eddy Simulation of Film Cooling Flows with Variable Density Jets

Abstract: The present paper investigates the impact of the velocity and density ratio on the turbulent mixing process in gas turbine blade film cooling. A cooling fluid is injected from an inclined pipe at α=30° into a turbulent boundary layer profile at a freestream Reynolds number of Re ∞ = 400,000. This jet-in-a-crossflow (JICF) problem is investigated using large-eddy simulations (LES). The governing equations comprise the Navier–Stokes equations plus additional transport equations for several species to simulate a non-reacting gas mixture. A variation of the density ratio is simulated by the heat-mass transfer analogy, i.e., gases of different density are effused into an air crossflow at a constant temperature. An efficient large-eddy simulation method for low subsonic flows based on an implicit dual time-stepping scheme combined with low Mach number preconditioning is applied. The numerical results and experimental velocity data measured using two-component particle-image velocimetry (PIV) are in excellent agreement. The results show the dynamics of the flow field in the vicinity of the jet hole, i.e., the recirculation region and the inclination of the shear layers, to be mainly determined by the velocity ratio. However, evaluating the cooling efficiency downstream of the jet hole the mass flux ratio proves to be the dominant similarity parameter, i.e., the density ratio between the fluids and the velocity ratio have to be considered.

Experimental Flow Structure Investigation of Compound Angled Film Cooling

Abstract: The experimental investigation of film-cooling flow structure provides reliable data for calibrating and validating a 3D feature based computational fluid dynamics (CFD) model being developed synchronously at the ETH Zurich. This paper reports on the flow structure of a film-cooling jet emanating from one hole in a row of holes angled 20 deg to the surface of a flat plate having a 45 deg lateral angle to the freestream flow in a steady flow, flat plate wind tunnel. This facility simulates a film-cooling row typically found on a turbine blade, giving engine representative nondimensionals in terms of geometry and operating conditions. The main flow is heated and the injected coolant is cooled strongly to obtain the requisite density ratio. All three velocity components were measured using a nonintrusive stereoscopic particle image velocimetry (PIV) system. The blowing ratio and density ratio are varied for a single compound angled geometry, and the complex three dimensional flow is investigated with special regard to vortical structure.

Modeling of Stardust Reentry Ablation Flows in the Near-Continuum Flight Regime

Abstract: The ablation process of the Stardust thermal protection material is designed to reduce aerodynamic heating during reentry for extreme conditions The coupling of ablation species with the flowfield is investigated in this work using the direct simulation Monte Carlo method for transitional to near-continuum flows To model surface thermal and chemical ablation processes, a variable surface temperature wall is obtained assuming a radiative heat flux balanced by convective heat flux It is found that chemical ablation due to the reaction between thermal protection system carbon materials and gaseous oxygen and nitrogen atoms is dominant compared with thermal ablation As the altitude decreases, the forebody surface temperature increases, the ablation process becomes more intensive, and the influence of ionization reactions on the flowfield becomes more important due to denser freestream conditions

Transverse Injection Through Diamond and Circular Ports into a Mach 5.0 Freestream

Abstract: Sonic transverse gaseous injection into a Mach 5.0 freestream flow was numerically simulated using two-equation and detached-eddy turbulence models. Circular- and diamond-shaped injectors were investigated in this study. The numerical simulations were compared with available experimental results and it was determined that both the Reynolds-averaged Navier-Stokes and detached-eddy simulation models captured the secondary flow structure. A detailed comparison of the secondary structures was performed for both injectors. Two new vortex structures of practical importance were observed in the diamond-injector flowfield. First, a leading-edge mixing mechanism was identified. Second, a trapped lateral counter-rotating vortex pair was produced. These new structures were observed in both Reynolds-averaged Navier-Stokes and detached-eddy simulation simulations. The detached-eddy simulations indicated that the large-scale structures observed in the plume/wake region of the flowfield were more organized in the diamond-injector test case. To better understand the secondary flow advection mechanism, the magnitudes of the terms in the compressible vorticity transport equation were compared. The diamond-injector structured-grid Reynolds-averaged Navier-Stokes solution was used as a baseline for this study. The inviscid compressibility, vortex-stretching, and baroclinic-torque terms were dominant. Downstream of the barrel-shock region, the baroclinic term was found to diminish when compared with the other inviscid terms. Planar-averaged results for the transport quantities confirmed this behavior. Vortex stretching was found to persist the longest.

Hypersonic Boundary-Layer Transition: Application to High-Speed Vehicle Design

Abstract: This paper reports the boundary-layer transition characteristics of airbreathing hypersonic configurations and their impact on design environments. It discusses the evolution of the National Aerospace Plane configuration from an axisymmetric baseline to a wedgelike configuration where the transition mechanism is dominated by “natural” transition. Thenatural transition is characteristic of the “quiet” environment in freeflight of smooth slender vehicles. This type of transition mechanism gradually evolves spatially in different modes that can be computed analytically and verified experimentally. The paper discusses the effect of leading-edge bluntness, surface wall temperature, and adverse pressure gradient in compression ramp on transition. The effect of freestream Mach number, Reynolds number, and angle of attack are also studied over the range of peak aerodynamic heating conditions of the National Aerospace Plane environment. The transition behavior was investigated using e-type calculations based on a linear stability code known as the eMalik code.

Improved Film Cooling Effectiveness by Placing a Vortex Generator Downstream of Each Hole

Abstract: Calculations are presented demonstrating the effect of placing a delta vortex generator downstream of a film cooling hole. The effects of blowing ratio, density ratio, and spanwise pitch are included in the study. Flow over a flat plate with film cooling holes oriented at a 30 degree angle was investigated. The Reynolds numbers based on the freestream velocity and the hole diameter was 11,300. The simulation was performed using the Glenn-HT code, a full three-dimensional Navier-Stokes solver using the Wilcox k-ω turbulence model. A structured multi-block grid was used with approximately one million cells, and average y+ values on the order of unity. Local and span averaged effectiveness are presented. Analysis and visualization of the flow are presented as well as a discussion on the mechanisms which contribute to the dramatic improvement in effectiveness. The results demonstrate that the delta vortex generator was able to annihilate the up-wash vortex pair produced by the film hole and produce a down-wash pair downstream.

Separated Flow Measurements on a Highly Loaded Low-Pressure Turbine Airfoil

Abstract: Boundary layer separation, transition, and reattachment have been studied on a new, very high lift, low-pressure turbine airfoil. Experiments were done under low freestream turbulence conditions on a linear cascade in a low speed wind tunnel. Pressure surveys on the airfoil surface and downstream total pressure loss surveys were documented. Velocity profiles were acquired in the suction side boundary layer at several streamwise locations using hot-wire anemometry. Cases were considered at Reynolds numbers (based on the suction surface length and the nominal exit velocity from the cascade) ranging from 25,000 to 330,000. In all cases, the boundary layer separated, but at high Reynolds number the separation bubble remained very thin and quickly reattached after transition to turbulence. In the low Reynolds number cases, the boundary layer separated and did not reattach, even when transition occurred. This behavior contrasts with previous research on other airfoils, in which transition, if it occurred, always induced reattachment, regardless of Reynolds number.

Velocity Distributions for Low Pressure Turbines

Abstract: A parametric set of velocity distributions has been investigated using a flat plate experiment. Three different diffusion factors and peak velocity locations were tested. These were designed to mimic the suction surfaces of Low Pressure (LP) turbine blades. Unsteady wakes, inherent in real turbomachinery flows, were generated using a moving bar mechanism. A turbulence grid generated a freestream turbulence level that is believed to be typical of LP turbines. Measurements were taken across a Reynolds number range of 50,000–220,000 at three reduced frequencies (0.314, 0.628, 0.942). Boundary layer traverses were performed at the nominal trailing edge using a Laser Doppler Anemometry system and hot-films were used to examine the boundary layer behaviour along the surface. For every velocity distribution tested, the boundary layer separated in the diffusing flow downstream of the peak velocity. The loss production is dominated by the mixing in the reattachment process, mixing in the turbulent boundary layer downstream of reattachment and the effects of the unsteady interaction between the wakes and the boundary layer. A sensitive balance governs the optimal location of peak velocity on the surface. Moving the velocity peak forwards on the blade was found to be increasingly beneficial when bubble-generated losses are high, i.e. at low Reynolds number, at low reduced frequency and at high levels of diffusion.Copyright © 2008 by ASME

Large-eddy simulation of flow separation on an airfoil at a high angle of attack and Re= 10 5 using Cartesian grids

Abstract: Incompressible flow separating from the upper surface of an airfoil at an 18° angle of attack and a Reynolds number of Re = 105, based on the freestream velocity and chord length c, is studied by the means of large-eddy simulation (LES) The numerical method is based on second-order central spatial discretization on a Cartesian grid using an immersed boundary technique The results are compared with an LES using body-fitted nonorthogonal grids and with experimental data

Laminar Flow Control on a Swept Wing with Distributed Roughness

Abstract: The work cumulated in a series of laminar-turbulent transition flight-test experiments on a swept wing with the goal of validating the spanwise-periodic distributed roughness elements (DRE) technology in a Reynolds number range applicable to SensorCraft technology. Phase I of the program measured freestream turbulence levels that were nominally 0.05% to 0.06% of the freestream speed and thus established the suitability of the flight environment for the laminarization flights. Phase II of the program did the baseline transition measurements on the airfoil i.e. with and without DRE technology. The region of laminar flow was extended from 30% to 60% chord at a chord Reynolds number of Rec = 8 x10 6 and sweep angle, Λ = 30°. Establishing the origins of turbulent flow and transition from laminar to turbulent flow remains an important challenge of fluid mechanics. The common thread connecting aerodynamic applications is the fact that they deal with bounded shear flows (boundary layers) in open systems (with different upstream or initial amplitude conditions). It is well known that the stability, transition, and turbulent characteristics of bounded shear layers are fundamentally different from those of free shear layers. Likewise, open systems are fundamentally different from those of closed systems. The distinctions are trenchant and thus form separate areas of study. For the classic open system, no mathematical model exists that can predict the transition Reynolds number on a simple flat plate because the influences of freestream turbulence, sound, and surface roughness are incompletely understood. With the maturation of linear stability methods and the conclusions that breakdown mechanisms are initial-condition dependent, more emphasis is now placed on the understanding of the source of initial disturbances than on the details of the later stages of transition.

Film Cooling Effectiveness Distributions on a Turbine Blade Cascade Platform With Stator-Rotor Purge and Discrete Film Hole Flows

Abstract: An experimental investigation to obtain detailed film cooling effectiveness distributions on a cooled turbine blade platform within a linear cascade has been completed. The Reynolds number of the freestream flow is 3.1 × 10 5 , and the platform has a labyrinthlike seal upstream of the blades to model a realistic stator-rotor seal configuration. An additional coolant is supplied to the downstream half of the platform via discrete film cooling holes. The coolant flow rate through the upstream seal varies from 0.5% to 2.0% of the mainstream flow, while the blowing ratio of the coolant through the discrete holes varies from 0.5 to 2.0 (based on the mainstream velocity at the exit of the cascade). Detailed film cooling effectiveness distributions are obtained using the pressure sensitive paint (PSP) technique under a wide range of coolant flow conditions and various freestream turbulence levels (0.75% or 13.4%). The PSP technique clearly shows how adversely the coolant is affected by the passage induced flow. With only purge flow from the upstream seal, the coolant flow rate must exceed 1.5% of the mainstream flow in order to adequately cover the entire passage. However, if discrete film holes are used on the downstream half of the passage, the platform can be protected while using less coolant (i.e., the seal flow rate can be reduced).

The Effects of Endplates on a Rotating Cylinder in Crossflow

Abstract: The effects of endplates on the aerodynamic behavior of a rotating circular cylinder in crossflow have been investigated experimentally for Reynolds numbers, based on the cylinder diameter, of 1.6x10 4 � Re � 9.5x10 4 and peripheral velocity to freestream velocity ratios of Ω � 8. The primary focus was on the impact of endplate size, though the effects of having one free end, and whether the endplates were spinning or stationary, were also assessed. The findings showed that, although endplates can significantly enhance lift and improve the lift-to-drag ratio, a limiting lift coefficient was always reached, regardless of end conditions. Larger endplates were found to delay this plateau to higher velocity ratios. The influence of endplate size on drag behavior was more complex, exhibiting a dependency on velocity ratio. The power required to spin the cylinder was generally unaffected by either Reynolds number or, except for the largest plate sizes, end conditions. Measurements of the time-averaged total pressure variation in the wake indicated that force behavior was dependent on the formation and subsequent development of large vortices at the cylinder tips.

Effect of freestream isotropic turbulence on heat transfer from a sphere

Abstract: We consider direct numerical simulation (DNS) based on pseudospectral methods to study the heat transfer around a stationary sphere held at a constant temperature and subject to an ambient turbulent velocity and temperature condition. The sphere Reynolds number is in the range of 63–400, and the sphere diameter (d) varies from one to eight times the Kolmogorov scale (η). The ambient turbulent field is isotropic, and the Taylor microscale Reynolds number Rλ varies from 38 to 240. Results from two sets of DNS are presented. In the first set, the ambient velocity field is turbulent, but the ambient temperature is held constant. In the second set of simulations, both the ambient velocity and the temperature fields are turbulent. These two sets of simulations allow us to isolate the role of freestream velocity fluctuations and temperature fluctuations in modifying the mean and time-dependent heat transfer from the sphere. The mean Nusselt number is observed to be independent of Rλ. It is shown that the freestr...

Computational Study of Unsteady Low-Reynolds-Number Airfoil Aerodynamics Using Moving Overlapping Meshes

Abstract: One of the challenging areas of research motivated by insects and birds is to predict instantaneous unsteady forces on a rapidly maneuvering object. The basic underlying mechanisms of force production are well known and have been established by early researchers through extensive experiments and numerical computations supported by classical theories of unsteady aerodynamics. Nevertheless, to date, experiments and numerical computations of this problem have been focused either on fluid mechanics alone or on fluid mechanics coupled with structural mechanics. The rigid-body dynamics of plunging/pitching/flapping bodies has received less attention. In view of this, an attempt is made here to simulate the rigid-body dynamics of plunging airfoils in incompressible low-Reynolds-number flow in the range of 10 < Re < 4.5 x 10 4 by solving the incompressible Navier-Stokes equations on moving overlapping meshes. The need for a sharp trailing edge is demonstrated by comparing the computed thrust and forward speed for airfoils with different trailing-edge topologies when plunged in zero freestream velocity. The effect of freestream and ground proximity was also analyzed for active flight.

Numerical Investigation of Jet Interaction in a Supersonic Freestream

Abstract: The objective of this investigation is to numerically evaluate effects of a reaction control (divert) jet on the aerodynamic performance of a generic interceptor missile operating at supersonic flight conditions. These effects include transient operation, external chemical reactions, burning, and geometric scale (full scale versus subscale). Three-dimensional computations of the highly turbulent flow field produced by a pulsed, supersonic, lateral-jet control thruster interacting with the supersonic freestream and missile boundary layer of a generic interceptor missile are evaluated. A generic missile interceptor configuration consisting of a long, slender body containing fixed dorsal and tail fins is simulated in this study. Parametric computational fluid dynamic solutions are obtained at altitude conditions corresponding to 19.7 km for the following scenarios: 1) steady-state conditions with the lateral control jet turned off; 2) steady-state conditions with the lateral control jet turned on; 3) steady-state conditions with the lateral control jet turned off with 1/10 subscale; 4) steadystate conditions with the lateral control jet turned on with 1/10 subscale; 5) transient jet startup conditions; 6) transient jet shutdown conditions; 7) steady-state, finite-rate chemistry; and 8) steady-state, frozen calculations with the chemical reactions “turned off.” A thermally and calorically perfect gas with a specific heat ratio equal to 1.4 was assumed for both the transient and geometric scale calculations. Vehicle forces and moments are assessed from each solution by integrating the surface pressures and viscous shear stresses computed on the missile surfaces. These results are used to determine the influence of the jet interaction effects on the transient, external burning, and geometric scale aerodynamic performance of the missile. The analysis predicts strong transient influences, small external burning influences, and very small full-scale vs 1/10-subscale effects for the integrated normal force and pitching moment.

Comparison of High-Altitude Hypersonic Wake Flows of Slender and Blunt Bodies

Abstract: The gas dynamic features of the laminar, near-wake flow behind slender and blunt hypersonic vehicles are studied using the direct simulation Monte Carlo method. Near-wake flows are characterized by features of low density, low Reynolds number, high temperature, thermal nonequilibrium, species separation, and recirculation. The impact of freestream number density and velocity on the near-wake flowfield is considered and compared for slender and blunt bodies. The near-wake structure postulated by theory and observed in numerical continuum calculations is also observed in the kinetic simulations, which are more accurate in the high-altitude, rarefied near-wake flow. The paper discusses the validation of the direct simulation Monte Carlo computational tool with experimental data for slender and blunt shapes and a previously published blunt direct simulation Monte Carlo geometry case. Then, the near-wake flows generated by a 10 deg slender cone and a 70 deg blunt body are analyzed. The near-wake flows behind slender and blunt bodies are similar in that the freestream Mach number has little impact on the near-wake flow structure and the recirculation length is not found to be related to the local Reynolds number. For both geometries, the base radius was found to be the characteristic length in the near-wake flow. Significant differences in the near-wake flow for the two geometries were observed in the spatial distribution of gas temperatures, the degree of chemical dissociation, and the sensitivity of recirculation length to freestream number density.

Simulation of CO2-N2 Expansion Tunnel Flows for the Study of Radiating Shock Layers

Abstract: A 25MJ/kg CO2–N2 expansion tunnel condition has been developed for the study of radiating shock layers in the X2 impulse facility at the University of Queensland. A hybrid Lagrangian and Navier–Stokes computational simulation technique is found to give good correlation with experimentally measured shock speeds and pressure traces. The use of a decaying inertial diaphragm model for describing secondary diaphragm rupture is found to predict between 4% and 25% more CO2 recombination over the test time than the widely accepted holding-time model. Inviscid simulations of the hypersonic nozzle expansion process with a two-temperature model indicate the final test gas is in both chemical and thermal nonequilibrium. The obtained freestream conditions are applied to radiatively coupled simulations of a 25mm diameter cylinder in the test flow. Grid independent solutions show good agreement with experimentally measured shock detachment and predict a radiative emission spectrum dominated by the CO Fourth-Postive band system.