Showing papers on "Freestream published in 2006"

Plasma Actuators for Separation Control of Low-Pressure Turbine Blades

Abstract: This work involves the documentation and control of flow separation that occurs over turbine blades in the low-pressure-turbine (LPT) stage at the low Reynolds numbers typical of high-altitude cruise. We utilize a specially constructed linear cascade that is designed to study the flowfield over a generic LPT cascade consisting of Pratt and Whitney Pak B shaped blades. The center blade in the cascade is instrumented to measure the surface-pressure coefficient distribution. Optical access allows laser-Doppler-velocimetry measurements for boundary-layer profiles. Experimental conditions were chosen to give a range of chord Reynolds numbers from 10 4 to 10 5 , and a range of freestream turbulence levels from u'/U∞ = 0.08 to 2.85%. The surface-pressure measurements were used to define a region of separation and reattachment that depends on the freestream conditions

Time evolution and mixing characteristics of hydrogen and ethylene transverse jets in supersonic crossflows

Abstract: We report an experimental investigation that reveals significant differences in the near-flowfield properties of hydrogen and ethylene jets injected into a supersonic crossflow at a similar jet-to-freestream momentum flux ratio. Previously, the momentum flux ratio was found to be the main controlling parameter of the jet’s penetration. Current experiments, however, demonstrate that the transverse penetration of the ethylene jet was altered, penetrating deeper into the freestream than the hydrogen jet even for similar jet-to-freestream momentum flux ratios. Increased penetration depths of ethylene jets were attributed to the significant differences in the development of large-scale coherent structures present in the jet shear layer. In the hydrogen case, the periodically formed eddies persist long distances downstream, while for ethylene injection, these eddies lose their coherence as the jet bends downstream. The large velocity difference between the ethylene jet and the freestream induces enhanced mixing at the jet shear layer as a result of the velocity induced stretching-tilting-tearing mechanism. These new observations became possible by the realization of high velocity and high temperature freestream conditions which could not be achieved in conventional facilities as have been widely used in previous studies. The freestream flow replicates a realistic supersonic combustor environment associated with a hypersonic airbreathing engine flying at Mach 10. The temporal evolution, the penetration, and the convection characteristics of both jets were observed using a fast-framing-rate (up to 100 MHz) camera acquiring eight consecutive schlieren images, while OH planar laser-induced fluorescence was performed to verify the molecular mixing.

Experiment of the Two-Dimensional Roughness Effect on Hypersonic Boundary-Layer Transition

Abstract: An experimental investigation of the effects of two-dimensional roughness on hypersonic boundary layers was carried out at the JAXA 0.5 m hypersonic wind tunnel using a 5 deg half-angle sharp cone at a freestream Mach number of 7.1 and a wide range of stagnation conditions. Aerodynamic heating distributions and surface pressure fluctuations were measured with and without roughness elements applied to the cone. A wavy wall roughness with a wavelength of 2� located well upstream of the breakdown region had the effect of delaying transition. Surface pressure spectrum densities indicate that disturbances of the roughness wavelength are amplified compared to the smooth wall case. At a lower stagnation temperature, however, wavy wall roughness configurations had the same effect as regular roughness, and no discernible differences between wavy wall roughness and spherical roughness were observed, either in the transition point or in the pressure fluctuation spectrum.

Experimental Study on Thermal Protection System by Opposing Jet in Supersonic Flow

Abstract: Introduction C URRENTLY, developments of reusable launch vehicle (RLV) for a low-cost space transportation system are in progress. In the development of RLV, one of the most important problems is the severe aerodynamic heating at the nose and leading edges of the vehicle. In such supersonic and hypersonic flights, prediction of aerodynamic heating and construction of proper thermal protection system are especially important. Heat-resistant tiles and ablators are currently used for thermal protection systems. However, those thermal protection systems are not reusable. In the present study, the method using an opposing jet is proposed for fully reusable thermal protection system of RLV. The method can be considered to have almost the same effect of heat reduction at nose region as the method with mechanical spike.1 The opposing jet works as an aerodynamic spike to move the detached shock wave away from the nose and form a recirculation region, which is quite effective to reduce aerodynamic heating at the nose region. The schematic diagram of supersonic flowfields with opposing jet injected at the nose of a blunt body is shown in Fig. 1. In the flowfield, the opposing jet forms a Mach disk and contact surface with freestream. The jet layer reattaches to the body surface and forms a recirculation region between the nozzle exit and reattachment point of the jet layer. The recompression shock wave is formed near the reattachment point of the jet layer. Many studies on opposing jet flow have been conducted in order to reveal the flow mechanism.2−7 However, most of those studies are related to the stability of flowfield and oscillations of shock waves. Except for Warren,6 not much study has been conducted to reveal the effects of opposing jet on reduction of aerodynamic heating. In the present study, geometric ratio of diameters and Mach number are fixed. The flow stability is determined by the total pressure ratio of freestream to opposing jet. We define the total pressure ratio

Momentum Transfer for an Aerodynamic Plasma Actuator with an Imposed Boundary Layer

Abstract: In previous work at the United States Air Force Academy, the phenomenology and behavior of the aerodynamic plasma actuator (a dielectric barrier discharge (DBD)) has been investigated. In order to provide additional insight into the phenomenology associated with the transfer of momentum to air by a plasma flow actuator, the velocity distributions upstream and downstream of a plasma actuator with an induced boundary layer were measured using freestream velocities of approximately 4.6 and 6.8 m/s for a range of frequencies (5-20 kHz) and voltages (7.5-10 kV amplitude). The body forces on the air were calculated using a control volume momentum balance. The results show that the body force acts in the sub-boundary layer region. For constant voltage, the body force is proportional to frequency producing a constant impulse per cycle, and the energy dissipation per cycle and efficiency are independent of frequency. The body forces are not affected by the freestream velocity.

Large-Eddy Simulation of Transition in a Separation Bubble

Abstract: In this paper, large-eddy simulation of the transition process in a separation bubble is compared to experimental results. The measurements and simulations are conducted under low freestream turbulence conditions over a flat plate with a streamwise pressure distribution typical of those encountered on the suction side of turbine airfoils. The computational grid is refined to the extent that the simulation qualifies as a "coarse" direct numerical simulation. The simulations are shown to accurately capture the transition process in the separated shear layer. The results of these simulations are used to gain further insight into the breakdown mechanisms in transitioning separation bubbles.

Partially-averaged Navier-Stokes method for turbulent flows: k-w model implementation

Abstract: Kolmogorov's k-ω model lacked a production term in the equation for ω making the model flawed. Also the model lacked a molecular diffusion term making the model strictly applicable to high Reynolds number flows and unable to be integrated through the viscous sub-layer. Wilcox 2 proposed the k-ω model with the production term included in the ω equation. This two equation model has proven to be numerically more stable that the standard k-e model primarily in the viscous sub-layer near the wall. The model does not require explicit wall-damping functions compared to the k-e model as the specific dissipation rate, ω is large in the wall region. In a numerical computation, specifying the wall boundary condition requires only the specification of the distance from the wall to the first point off the wall without any viscous corrections. In the logarithmic region, the model gives good agreement with experimental results for adverse pressure gradient flows due to the lack of a cross diffusion term in the ω equation. The model predicts the turbulent kinetic energy behavior close to the solid boundary with good accuracy and even describes the boundary-layer transition reasonably well. However, the lack of a cross diffusion term causes the model to be sensitive to small freestream values of ω, adversely affecting the performance of the model in free shear flows. Even though, the sensitivity of the model is reduced for complex flows 3 such as flow past a circular cylinder, the presence of small freestream ω in the wake imparts ambiguity in the predicted results. In the present work, we consider the revised k-ω model 4

Control of Laminar Separation Bubbles Using Instability Waves

Abstract: This paper presents detailed investigations related to active transition control in laminar separation bubbles. The investigations rely on direct numerical simulations based on the complete Navier-Stokes equations for a flat-plate boundary layer. A laminar separation bubble is created by imposing a streamwise adverse pressure gradient at the freestream boundary of the integration domain. Different steady and unsteady boundary layer disturbances are then introduced at a disturbance strip upstream of separation and their effects on the separation bubble are studied. It is shown that the size of the separated region can be controlled most efficiently by very small periodic oscillations, which lead to traveling instability waves that grow to large levels by the hydrodynamic instability of the flow. Indications for the preferred frequency of these waves can be obtained from linear stability theory, but since the problem is nonlinear, only direct numerical simulations can really qualify or disqualify the predictions. Overall, it turns out that unsteady two- or three-dimensional disturbances have a stronger impact on the size of the bubble than steady disturbances, because they directly provide initial amplitudes for the laminar-turbulent transition mechanism.

A Comparison of Cylindrical and Fan-Shaped Film-Cooling Holes on a Vane Endwall at Low and High Freestream Turbulence Levels

Abstract: Fan-shaped film-cooling holes have been shown to provide superior cooling performance to cylindrical holes along flat-plates and turbine airfoils over a large range of different conditions. Benefits of fan-shaped holes include less required cooling air for the same performance, increased part lifetime, and fewer required holes. The major drawback however, is increased manufacturing cost and manufacturing difficulty, particularly for the vane platform region. To this point, there have only been extremely limited comparisons between cylindrical and shaped holes on a turbine endwall at either low or high freestream turbulence conditions. This study presents film-cooling effectiveness measurements on an endwall surface in a large-scale, low-speed, two-passage, linear vane cascade. Results showed that film-cooling effectiveness decreased with increasing blowing rate for the cylindrical holes, indicating jet lift-off. However, the fan-shaped passage showed increased film-cooling effectiveness with increasing blowing ratio. Overall, fan-shaped holes increased film-cooling effectiveness by an average of 75% over cylindrical holes for constant cooling flow.Copyright © 2006 by ASME

Effects of Free-Stream Turbulence on Wing-Tip Vortex Formation and Near Field

Abstract: The formation and near-field development of a wing-tip vortex under the influence of freestream turbulence were examined using flow visualization and hot-wire anemometry. A low turbulence freestream as well as two cases of grid turbulence with different intensities and length scales were considered. In all cases, the tip vortex was found to form from three smaller vortices, but the turbulence in its core was found to intensify with increasing freestream turbulence. The vortex trajectory was found to be unaffected by freestream turbulence, but the wing wake that was rolling up around the vortex was observed to have a curvature that decreased as freestream turbulence increased. The mean axial velocity distribution in the low-turbulence case was neither jetlike nor wakelike but had an annular shape. Time-averaged velocity profiles measured in the turbulent freestream cases were wakelike, and it was inferred that the instantaneous profiles would be significantly affected by vortex meandering. Mean circumferential velocity distributions in the vortex core displayed self-similar developments in all cases examined. Finally, it was found that the apparent diffusion in the shear layer shed from the wing increased with increasing freestream turbulence.

Flow Measurements of Vortex Generator Jets in Separating Boundary Layer

Abstract: Vortex generator jets (VGJs) have been found to be an effective method of active separation control on the suction side of a low-pressure turbine (LPT) blade at low Reynolds numbers. The flow mechanisms responsible for this control were studied and documented to provide a basis for future improvements in LPT design. Data were collected using a stereo particle-image-velocimetry system that enabled all three components of velocity to be measured. First, steady VGJs were injected into a laminar boundary layer on a flat plate (nonseparating boundary layer) to more fully understand the characteristics and behavior of the produced vortices. Jets injected normal to the surface created vortices of lesser strength that migrated out farther from the boundary layer. The vortices produced by angled jets (injected at 30-deg pitch and 90-deg skew angles to the freestream) remained closer to the wall and maintained their structure for a longer distance. The angled jets also produced vortices that were more effective at sweeping low momentum fluid up from the boundary layer while transporting high-momentum freestream fluid down towards the wall. Second, pulsed VGJs were injected on a flat plate with an applied adverse pressure gradient equivalent to that experienced by an LPT blade. This configuration was used to study the effectiveness of the flow control exhibited by both jet configurations on a separating boundary layer. Time-averaged results showed similar boundary-layer separation reduction for both normal and angled jets; however, individual characteristics of the control differed. Normal jets created a disturbance that provided flow control earlier during the pulsing cycle, whereas the disturbance produced by the angled jets was more effective in reducing the boundary-layer separation.

Numerical Assessment of Data in Catalytic and Transitional Flows for Martian Entry

Abstract: The conditions for a typical run from the MSL phase two study of transition that was performed in the LENS facility have been analyzed to understand the sensitivity to the freestream conditions of the facility. A simplified analysis technique has been used based on energy accounting to freeze specified portions of the chemical or vibrational energy during the expansion process in the nozzle. The effect of freezing this energy results in increased shock standoff distance that better matches the measured shock shape. Based on several cases, it was found that freezing approximately 42% of the total enthalpy of the flow in the vibration mode results in the best agreement with the measured shock shape. This modified condition also results in significantly better agreement with the measured surface heat transfer at the stagnation point and with the measured pressure at the shoulders of the model. Based on this adjusted freestream condition, the surface heat transfer data shows behavior generally consistent with fully-catalytic recombination on the cold wall. This behavior is consistent with previous results obtained in shock tunnel facilities in carbon dioxide, air, and nitrogen. Although the mechanism causing this frozen energy in the flow has not been identified, the sensitivity of the transition onset point of the flowfield to this phenomenon has been estimated to be less than 10% based on a simple transition criterion.

Mach 6 Boundary-Layer Stability Experiments on Sharp and Blunted Cones

Abstract: Recent studies of hypersonic boundary-layer stability and transition on cones with sharp and blunted nosetips are presented. The experiments were carried out on a 7-deg half-angle cone at freestream Mach number 5.95. Laminar‐turbulent transition locations are measured for various flow parameters and model nose bluntness. Mean and fluctuation characteristics of the flow are obtained using constant-temperature hot-wire anemometry. The spectral content and amplification rates of natural disturbances are obtained. The method of artificial wave packets is applied to obtain detailed information on the disturbances. Data on development of both natural and artificial finite-amplitude disturbances are compared. It is experimentally shown that the wave vector of the most unstable waves of the first mode have an inclination angle of 40‐49 deg. In the frequency range of the second mode, plane waves appear to be dominant and have the highest amplification. The bluntness of the cone nosetip results in an increase of the disturbance amplification rate downstream of the entropy layer swallowing point. At the same time, nose bluntness dramatically increases the transition Reynolds number because of the strong damping of initial disturbances.

Pressure Gradient Effects on Hypersonic Cavity Flow Heating

Abstract: The effect of a pressure gradient on the local heating disturbance of rectangular cavities tested at hypersonic freestream conditions has been globally assessed using the two-color phosphor thermography method. These experiments were conducted in the Langley 31-Inch Mach 10 Tunnel and were initiated in support of the Space Shuttle Return-To-Flight Program. Two blunted-nose test surface geometries were developed, including an expansion plate test surface with nearly constant negative pressure gradient and a flat plate surface with nearly zero pressure gradient. The test surface designs and flow characterizations were performed using two-dimensional laminar computational methods, while the experimental boundary layer state conditions were inferred using the measured heating distributions. Three-dimensional computational predictions of the entire model geometry were used as a check on the design process. Both open-flow and closed-flow cavities were tested on each test surface. The cavity design parameters and the test condition matrix were established using the computational predictions. Preliminary conclusions based on an analysis of only the cavity centerline data indicate that the presence of the pressure gradient did not alter the open cavity heating for laminar-entry/laminar-exit flows, but did raise the average floor heating for closed cavities. The results of these risk-reduction studies will be used to formulate a heating assessment of potential damage scenarios occurring during future Space Shuttle flights.

Time-series and time-averaged characteristics of subsonic to supersonic base flows

Abstract: Physics of cylindrical base flows ranging from subsonic to supersonic speeds at zero angle of attack are computationally investigated. Time-series and time-averaged investigations of base flows show distinctive characteristics at subsonic, transonic, and supersonic regimes. Normalized time-averaged base pressure decreases proportionally with respect to increasing freestream dynamic pressure in the subsonic regime of M ∞ 1.5. Normalized base pressure fluctuations sharply increase at transonic speeds, whereas they decrease with increasing freestream Mach number at subsonic and supersonic speeds. Appearance of unsteady local shock waves change the characteristics of base pressure distinctively at the transonic speeds. Spectra of the base pressure show one clear peak at subsonic speeds (related to the shear layer dynamics), two clear peaks at transonic speeds (related to the shear layer dynamics and its subharmonic), and three major peaks at supersonic speeds (related to the shear layer dynamics, its subharmonic, and an additional mechanism). Instability of the free shear layers has dominant influence on the overall base flowfield over a wide range of Mach numbers ranging from subsonic to supersonic speeds. However, at supersonic speeds, an additional mechanism of instability within the recirculating region is possibly at work and has dominant influence on the flowfield. The dominant mechanisms significantly cause the strong Mach number dependence of the high-pressure region which is strongly related to the base pressure. The spectrum analysis suggests that the substantial base pressure fluctuations are caused by the pulsing of the flow inside the recirculating region.

Assessment of Synthetic Inflow Generation for Simulating Injection Into a Supersonic Crossflow

Abstract: Detached eddy simulation is used to simulate the flow field around gaseous jets injected into supersonic crossflows. The simulations are done with and without the use of synthetic inflow generation, a method used to provide a realistic, time-varying boundary layer for the inflow condition of the simulations. The goal is to assess the ability of DES to simulate the complicated flow field around the injection location, and to assess the ability of synthetic inflow generation to capture the contribution of the preinjection boundary layer to the flow field. Reynolds stresses and turbulent kinetic energy predicted by the simulations are compared to experimentally measured values in one test case of normal injection of air into a Mach 1.6 freestream. A second test case compares predicted mean injectant mole fractions with measured values for a pair of staged injectors that inject air into a Mach 2 freestream. Unlike previous simulations of jets in supersonic crossflows done by the authors, the DES provides highly unsteady jet plumes, even without the synthetic inflow boundary layer. This is due partially to improved mesh resolution, but is mostly due to large and highly energetic separation regions upstream of the normal injectors, compared to smaller, less energetic regions upstream of the angled injectors simulated previously. As a result of the natural instability of these configurations, the role of the boundary layer structures is found to be small. However, the results show that DES reproduces the mean and fluctuating quantities very well compared to the measured values.

Comparisons between Measurements in Regions of Laminar Shock Wave Boundary Layer Interaction in Hypersonic Flows with Navier-Stokes and DSMC Solutions

Abstract: : Experimental studies were conducted in conjunction with computations in a code validation exercise to examine the ability of DSMC and Navier-Stokes techniques to predict the complex characteristics of regions of shock/shock and shock/ boundary layer interactions in hypervelocity flows In the experimental program, detailed heat transfer and pressure measurements in laminar regions of shock wave/boundary layer interaction, and shock/shock interaction, over hollow cylinder/flare and double cone configurations in hypersonic flow The experimental studies were conducted for a Mach number range from 10 to 12 with Reynolds numbers from 1 x 104 to 5 x 105 and stagnation temperatures from 2,000 R to 5,000 R Miniature high-frequency thin-film and piezoelectric instrumentation were employed to obtain the high spatial resolution required to accurately define the distribution of heat transfer and pressure in the strong gradients which occur in regions of shear layer reattachment and shock/shock interaction The program reported here was conducted in two phases In the first phase of the code validation study, measurements and blind computations were made over complete hollow cylinder/flare and double cone configurations in high temperature flows at Mach 10 and 12 for a range of freestream Reynolds numbers In the second phase of the program, detailed heat transfer and pressure measurements were made over an extensive range of Reynolds number and total enthalpy conditions using only the hollow cylinder and the 25 conical segment of the models tested earlier in phase I The selection of the freestream conditions employed in this second phase of the program was performed in conjunction with computations of the contoured nozzle flows and the flows over the two simple model configurations Based on the results of the latter studies, we validated the computational schemes used to predict the properties of the freestream developed in the test section of the tunnel

Comparative Study of Plasma Spray Flow Fields and Particle Behavior Near to Flat Inclined Substrates

Abstract: Numerical models have been developed using computational fluid dynamics (CFD) analysis program FLUENT V6.02© to investigate the effect of the substrate on the behavior of the plasma flow fields and in-flight particles. Simulations are performed for cases where flat substrates are either present or absent, for the former, the substrate is oriented perpendicularly or inclined to the torch axis. It is shown that although the presence of perpendicular or inclined substrate significantly influences the plasma flow fields at the vicinity of the substrate, the particle behavior remain relatively unaffected. The insignificant effect of the substrate on particle behavior is qualitatively verified by experimental observation using SprayWatch© imaging diagnostics equipment. Images captured by the equipment confirm that the particles travel through the plasma plume with high momentum and show no sudden change in theirtrajectories right before impacting the substrate. Both the numerical and experimental findings show that the freestream model is sufficiently detailed for future work of this nature.

Jet in Supersonic Crossflow on a Flat Plate

Abstract: A three dimensional full Navier-Stokes CFD model with the Smith κ κ κ κ-κl turbulence model is used to compute the interaction between a transverse jet and a supersonic crossflow. A new flow model for jets in supersonic crossflow is presented. Details of shock structure, vortex origination and separation/reattachment zones are described. Amplification coefficients are defined and calculated to quantify the obstruction component of the jet interaction forces and moments. The amplification coefficients increase nearly linear with freestream Mach number.

Aerodynamic Losses of a Cambered Turbine Vane: Influences of Surface Roughness and Freestream Turbulence Intensity

Abstract: The effects of surface roughness and freestream turbulence level on the aerodynamic performance of a turbine vane are experimentally investigated. Wake profiles are measured with three different free-Stream turbulence intensity levels (1.1%, 5.4%, and 7.7%) at two different locations downstream of the test vane trailing edge (1 and 0.25 axial chord lengths). Chord Reynolds number based on exit flow conditions is 0.9 X 10 6 . The Mach number distribution and the test vane configuration both match arrangements employed in an industrial application. Four cambered vanes with different surface roughness levels are employed in this study. Effects of surface roughness on the vane pressure side on the profile losses are relatively small compared to suction side roughness. Overall effects of turbulence on local wake deficits of total pressure, Mach number, and kinetic energy are almost negligible in most parts of the wake produced by the smooth test vane, except that higher freestream losses are present at higher turbulence intensity levels. Profiles produced by test vanes with rough surfaces show apparent lower peak values in the center of the wake. Integrated aerodynamic losses and area-averaged loss coefficient Y A are also presented and compared to results from other research groups.

Aerodynamics of a Family of Three Highly Loaded Low-Pressure Turbine Airfoils: Measured Effects of Reynolds Number and Turbulence Intensity in Steady Flow

Abstract: The steady, midspan aerodynamic performance of a family of three low pressure (LP) turbine airfoils has been investigated in a low-speed cascade wind tunnel. The baseline profile has a Zweifel coefficient of 1.08. To examine the influence of increased loading as well as the loading distribution, two additional airfoils were designed, each with 25% higher loading than the baseline version. All three airfoils have the same design inlet and outlet flow angles. The aerodynamic performance was investigated for Reynolds numbers ranging from 25,000 to 150,000 (based on the axial chord and inlet velocity) and for values of freestream turbulence intensity of 1.5% and 4%. The flow field was measured with a three-hole pressure probe. Also, detailed loading distributions were obtained for all three airfoils using surface static pressure taps. The baseline airfoil and the new aft-loaded airfoil showed a separation bubble on the suction side of the airfoil under most of the conditions examined. In addition, a sudden and intermittent stall was observed at low Reynolds numbers for the new aft-loaded airfoil. The relatively short separation bubble would abruptly “burst” and fail to reattach. As the Reynolds number was decreased over a narrow range, the percentage of time that the flow was fully-separated increased to 100%. By comparison, the separation bubble on the baseline airfoil gradually increased in size in an orderly way as the Reynolds number was decreased. The new front-loaded airfoil provided the most encouraging performance: no separation bubble was present except at the very lowest Reynolds numbers. The absence of a separation bubble also had a favourable effect on the loss behaviour of this airfoil: despite its much higher aerodynamic loading, it exhibited very similar midspan losses to those observed for the baseline airfoil.Copyright © 2006 by ASME

Effects of Temperature and Heat Transfer on Shock Train Structures inside Constant-Area Isolators

Abstract: Effects of temperature and heat transfer on shock train structures and isolator performance were investigated both experimentally and numerically. Two heat-sink isolators, one with a rectangular configuration and one with a round configuration, with identical cross-sectional areas and a length of 25", were operated with Mach 1.8 and 2.2 flows. Pressure profiles inside the shock train and temperatures on the isolator walls were measured. It was found that heat addition to a low Mach number flow inside an isolator can choke the flow and potentially decrease the isolator performance. On the other hand, heat removal from the flow can enhance the isolator performance by retarding flow choking. The numerical analysis shows that heat addition to a supersonic flow can increase boundary layer thickness and decrease both flow Mach number and the amount of heat required to choke the flow. Shock trains generated in a high-temperature flow are relatively long, and therefore may require a long isolator to prevent engine unstart. NOMENCLATURE H = isolator duct height L0.8 = shock train length M = Mach number P = pressure PR = pressure ration across the isolator, Pb/P1 T = temperature t = time u = freestream velocity x = freestream direction xs = shock train leading edge y = transverse direction z = spanwise direction γ = specific heat ratio θ = boundary layer momentum thickness Superscript

Modeling of Wake-Induced Transition in Linear Low-Pressure Turbine Cascades

Abstract: An unsteady Reynolds-averaged Navier-Stokes (URANS) strategy is applied to the problem of wake-induced transition at high freestream turbulence on the suction side of two blades representative of those used in low-pressure turbines. Experimentally, the blades are arranged in high-aspect-ratio linear cascades, with upstream circular bars generating passing wakes, and two-dimensional flow conditions are, thus, assumed. The strategy combines an explicit algebraic Reynolds-stress turbulence model with transition-specific modifications targeted at capturing the effects of high freestream turbulence and of pretransitional laminar fluctuations. Close attention is paid to numerical accuracy, and grids of up to 140,000 cells are used in combination with 800 time steps per pitchwise traverse to resolve small-scale features in the blade boundary layers that are associated with the unsteady interaction. The computational results demonstrate that the combined model returns a good representation of the response of the suction-side boundary layer to the passing wakes in both blades. Specifically, in the boundary layer of one of the two blades, the wakes are observed to cause a periodic upstream shift in the transition onset and, thus, correspondingly periodic attachment and calming. In the other, no separation occurs, and the wakes are shown to produce a significant periodic reduction in shape factor and increase in skin friction in the blade boundary layer, again as a consequence of the upstream shift in the transition location.

Transverse supersonic controlled swirling jet in a supersonic cross stream

Abstract: The spreading rate and mixing of a transverse jet in high-speed crossflow were modified using a swirling injector with a central control jet. The controlled supersonic swirling injector (CSSI) could be used to affect mixing both in the core and the shear layer of the jet. Rayleigh/Mie scattering from flowfield ice crystals and planar laser-induced fluorescence of the NO molecules were used to characterize penetration and mixing of the CSSI for six different cases. Instantaneous images were used to study the dynamical structures in the jet, whereas ensemble images provided information regarding the jet trajectory. Standard deviation images revealed information about the large-scale mixing/entrainment. Probability density functions were used to evaluate the probability and location of freestream, mixed, and jet fluid. They were also used to track the centerline and jet boundary on a dynamic scale. Side- (streamwise)-view images showed that the injector was capable of providing high penetration when compared to circular and swirling baseline injectors. An increase of 16% in mixing area was observed with the optimal case as compared with the other control cases. End- (spanwise)-view images show a maximum of 78 % increase in total area contained within the jet boundary for the optimal case when compared to the circular injector. Higher spanwise extent of the jet boundary was also observed with controlled cases, which could provide higher interfacial area for better mixing between the jet and the cross stream when compared to their baseline counterparts.

DES investigations of transverse injection into supersonic crossflow using a hybrid unstructured solver

Abstract: Numerical simulations of transverse injection through low-angled injector ports into a supersonic freestream are performed using a hybrid, unstructured solver. Two cases are investigated: air injected into a M=2.9 air freestream with a 25◦ injection angle, and heated helium injected into a M=4.0 air freestream with a 30◦ injection angle. Simulations were run in RANS and DES modes. A set of the DES simulations were run with an unsteady inflow boundary layer, where the perturbations are a statistically meaningful representation of a series of randomly placed hairpin eddies. This boundary condition was fed periodically into the domain, and was reused multiple times over the course of a simulation. The RANS and DES simulations are found to capture the salient features of the flow, though discrepancies with experimental data are found. While the DES simulations were found to give steady-state solutions for the flow fields, the addition of the unsteady inflow boundary layer was found to greatly impact the downstream flow field and to improve the overall agreement with experiment. In the case of the helium injection, it was found that the predicted mass fraction distributions of the DES simulations with the unsteady inflow boundary layer was far less dependent on the value of the turbulent Schmidt number. This result shows that the mixing found in DES simulation with the unsteady inflow boundary layer is a result of the large-scale turbulent motion of the flow, rather than because of the gradient diffusion term. However, the ‘box of eddies’ used to create the unsteady inflow boundary layers were not long enough to ensure that no bias was introduced into the flow field, and future simulations will be run with larger boxes. The results of the study show a great deal of promise for the use of DES simulations in conjunction with unsteady inflow boundary layers for simulation of SCRAMjet fuel injection.

DES Investigations of Transverse Injection into Supersonic Crossflow using a Hybrid

Abstract: Numerical simulations of transverse injection through low-angled injector ports into a supersonic freestream are performed using a hybrid, unstructured solver. Two cases are investigated: air injected into a M=2.9 air freestream with a 25 injection angle, and heated helium injected into a M=4.0 air freestream with a 30 injection angle. Simulations were run in RANS and DES modes. A set of the DES simulations were run with an unsteady inflow boundary layer, where the perturbations are a statistically meaningful representation of a series of randomly placed hairpin eddies. This boundary condition was fed periodically into the domain, and was reused multiple times over the course of a simulation. The RANS and DES simulations are found to capture the salient features of the flow, though discrepancies with experimental data are found. While the DES simulations were found to give steady-state solutions for the flow fields, the addition of the unsteady inflow boundary layer was found to greatly impact the downstream flow field and to improve the overall agreement with experiment. In the case of the helium injection, it was found that the predicted mass fraction distributions of the DES simulations with the unsteady inflow boundary layer was far less dependent on the value of the turbulent Schmidt number. This result shows that the mixing found in DES simulation with the unsteady inflow boundary layer is a result of the large-scale turbulent motion of the flow, rather than because of the gradient diusion term. However, the ‘box of eddies’ used to create the unsteady inflow boundary layers were not long enough to ensure that no bias was introduced into the flow field, and future simulations will be run with larger boxes. The results of the study show a great deal of promise for the use of DES simulations in conjunction with unsteady inflow boundary layers for simulation of SCRAMjet fuel injection.

New Scheme for the Computation of Compressible Flows

Abstract: A new approach for the computation of unsteady compressible flows has been developed. The new scheme employs upwinding of the convective flux based on particle velocity and has been termed the particle velocity upwinding (PVU) scheme. The PVU scheme is an explicit two-step predictor-corrector scheme, in which the convective fluxes are evaluated on cell faces using a first-order upwinding method. The scheme is accurate and stable, giving solutions free from oscillations near the discontinuities without any explicit addition of artificial viscosity. The PVU scheme has an edge over state-of-the-art high-resolution schemes in terms of simplicity of implementation in multidimensional flows and problems involving complex domains. The numerical scheme is validated for both Euler and Navier-Stokes equations. Furthermore, the PVU scheme is used to investigate laminar supersonic viscous flow over a forward-facing step. The results are obtained for M ∞ = 1.5-3.5 in steps of 0.5 and for Re ∞ = 10 4 . Step heights H s of 10 and 20% of the characteristic length of the problem are considered. The effect of step height and the incoming freestream Mach number on the spatial flow structure and on the important design parameters such as wall pressure, skin friction, heat transfer, and length of separated region are investigated.

On the Optimisation of Road Vehicle Leading Edge Radius in Varying Levels of Freestream Turbulence

Abstract: It has been recognised that the ideal flow conditions that exist in the modern automotive wind tunnel do not accurately simulate the environment experienced by vehicles on the road. This paper investigates the effect of varying one flow parameter, freestream turbulence, and a single shape parameter, leading edge radius, on aerodynamic drag. The tests were carried out at model scale in the Loughborough University Wind Tunnel, using a very simple 2-box shape, and in the MIRA Full Scale Wind Tunnel using the MIRA squareback Reference Car. Turbulence intensities up to 5% were generated by grids and had a strong effect on transcritical Reynolds number and Reynolds sensitivity at both model scale and full scale. There was a good correlation between the results in both tunnels.