Showing papers on "Freestream published in 2014"

Adiabatic Effectiveness Measurements for a Baseline Shaped Film Cooling Hole

Abstract: Film cooling on airfoils is a crucial cooling method as the gas turbine industry seeks higher turbine inlet temperatures. Shaped film cooling holes are widely used in many designs given the improved performance over that of cylindrical holes. Although there have been numerous studies of shaped holes, there is no established baseline shaped hole to which new cooling hole designs can be compared. The goal of this study is to offer the community a shaped hole design, representative of proprietary and open literature holes that serves as a baseline for comparison purposes. The baseline shaped cooling hole design includes the following features: hole inclination angle of 30° with a 7° expansion in the forward and lateral directions; hole length of 6 diameters; hole exit-to-inlet area ratio of 2.5; and lateral hole spacing of 6 diameters. Adiabatic effectiveness was measured with this new shaped hole and was found to peak near a blowing ratio of 1.5 at density ratios of 1.2 and 1.5 as well as at both low and moderate freestream turbulence of 5%. Reductions in area-averaged effectiveness due to freestream turbulence at low blowing ratios were as high as 10%.Copyright © 2014 by ASME

Invariant Formulation for the Minimum Induced Drag Conditions of Nonplanar Wing Systems

Abstract: Under the hypotheses of linear potential flow and rigid wake aligned with the freestream, a configuration-invariant analytical formulation for the induced drag minimization of single-wing nonplanar systems is presented. Following a variational approach, the resulting Euler–Lagrange integral equation in the unknown circulation distribution is obtained. The kernel presents a singularity of the first order, and an efficient computational method, ideal for the early conceptual phases of the design, is proposed. Munk’s theorem on the normalwash and its relation with the geometry of the wing under optimal conditions is naturally obtained with the present method. Moreover, Munk’s constant of proportionality, not provided in his original work, is demonstrated to be the ratio between the freestream velocity and the optimal aerodynamic efficiency. The augmented Munk’s minimum induced drag theorem is then formulated. Additional induced drag theorems are demonstrated following the derivations of this invariant proced...

Modal Analysis of Inclined Film Cooling Jet Flow

Abstract: Thermal and hydrodynamic flow field over a flat surface cooled with a single round inclined film cooling jet and fed by a plenum chamber is numerically investigated using large eddy simulation (LES) and validated with published measurements. The calculations are done for a freestream Reynolds number Re = 16,000, density ratio of coolant to freestream fluid ρj/ρ∞=2.0, and blowing ratio BR=ρjV/ρ∞V=1.0. A short delivery tube with aspect ratio l/D=1.75 and 35 deg inclination is considered. The evolution of the Kelvin–Helmholtz (K-H), hairpin and counterrotating vortex pair (CVP) vortical structures are discussed to identify their origins. Modal analysis of the complete 3D flow and temperature field is carried out using a dynamic mode decomposition (DMD) technique. The modal frequencies are identified, and the specific modal contribution toward the cooling wall temperature fluctuation is estimated on the film cooling wall. The low and intermediate frequency modes associated with streamwise and hairpin flow structures are found to have the largest contribution (in-excess of 28%) toward the wall temperature (or cooling effectiveness) fluctuations. The high frequency Kelvin–Helmholtz mode contributes toward initial mixing in the region of film cooling hole away from the wall. The individual modal temperature fluctuations on the wall and their corresponding hydrodynamic flow structures are presented and discussed.

Nonlinear Feedback Mechanisms Inside a Rectangular Cavity

Abstract: An examination of a rectangular cavity with a length-to-depth ratio of 5.67 was tested at Mach 0.7 and 1.5 with corresponding Reynolds numbers of 6.49×106 and 7.35×106/m, respectively. High-speed shadowgraph movies were simultaneously sampled with dynamic pressure sensors at 75 kHz. From the high-speed shadowgraph movies, observations of the cavity flowfield indicate that linear models like Rossiter’s equation and Helmholtz resonance may be too simplistic to correctly model rectangular cavity flow physics. Some of the observations are as follows. In the cavity’s shear layer, large-scale vortices have been found to be nonperiodic. The shear-layer convection velocity was determined to be 0.6U∞ at both Mach numbers, which is significantly higher than Rossiter’s equation predicts for correct peak frequency matching. Shear-layer entrainment of freestream flow starts and feeds the acoustic cycle inside the cavity. There are more acoustic wave in the cavity at any one time than has been considered before, which ...

Turbulence Measurements in High-speed Wind Tunnels Using Focusing Laser Differential Interferometry

Abstract: Characterization of freestream disturbances and their effect on laminar boundary layer transition is of great importance in high-speed wind tunnel testing, where significant differences between the behavior of scale-model and free-flight transition have long been noted. However, the methods traditionally used to perform this characterization in low-speed flows present significant difficulties when applied to supersonic and especially hypersonic wind tunnels. The design and theory of a focusing laser differential interferometer (FLDI) instrument, originally invented by Smeets at the Institut Saint-Louis in the 1970s and used recently by Parziale in the CalTech T5 shock tunnel, is presented. It is a relatively-simple, non-imaging common-path interferometer for measuring refractive signals from transition and turbulence, and it has a unique ability to look through facility windows, ignore sidewall boundary-layers and vibration, and concentrate only on the refractive signal near a pair of sharp beam foci in the core flow. The instrument’s low cost and ease of implementation make it a promising alternative to traditional hot-wire anemometry and particle-based methods for turbulence characterization. Benchtop experiments using a turbulent supersonic air jet demonstrate its focusing ability, frequency response, unwanted signal rejection, and ease of use. The instrument is used to optically interrogate the flow in the Penn State University Supersonic Wind Tunnel and USAF AEDC Hypervelocity Tunnel 9 for measurement of the overall intensity and spectra of freestream disturbances. Precise characterization of the strength and spectral content of the disturbances provides insight into their nature and potential effect upon boundary layer transition. A special feature of the FLDI instrument used here is the replacement of traditional fixed Wollaston prisms with variable Sanderson prisms for laser-beam separation and recombination.

Control of Incident Shock/Boundary-Layer Interaction by a Two-Dimensional Bump

Abstract: Separation due to the shock-wave/boundary-layer interaction in a supersonic/hypersonic inlet always brings negative effects on its performance. A new control method of the shock-wave/boundary-layer interaction based on a two-dimensional bump is brought forward and investigated by both experimental and computational methods. First, the uncontrolled case is studied. The flow turning angle of the incident shock is 12 deg, and the freestream Mach number is 3.5. The results show that the shock-wave/boundary-layer interaction generates complex three-dimensional flow structures with significant swirling nature, which will substantially deteriorate the performance of a supersonic/hypersonic inlet. Then, the controlled case demonstrates that the shock-induced flow separation can be effectively reduced by the coupling of the precompression effect and the acceleration effect of the bump with a height of 0.33 times of the boundary-layer thickness. In addition, the efficiency of the control method for different shock-...

Boundary-Layer Transition on a Slender Cone in Hypervelocity Flow with Real Gas Effects

Abstract: The laminar to turbulent transition process in boundary layer flows in thermochemical nonequilibrium at high enthalpy is measured and characterized. Experiments are performed in the T5 Hypervelocity Reflected Shock Tunnel at Caltech, using a 1 m length 5-degree half angle axisymmetric cone instrumented with 80 fast-response annular thermocouples, complemented by boundary layer stability computations using the STABL software suite. A new mixing tank is added to the shock tube fill apparatus for premixed freestream gas experiments, and a new cleaning procedure results in more consistent transition measurements. Transition location is nondimensionalized using a scaling with the boundary layer thickness, which is correlated with the acoustic properties of the boundary layer, and compared with parabolized stability equation (PSE) analysis. In these nondimensionalized terms, transition delay with increasing CO2 concentration is observed: tests in 100% and 50% CO2, by mass, transition up to 25% and 15% later, respectively, than air experiments. These results are consistent with previous work indicating that CO2 molecules at elevated temperatures absorb acoustic instabilities in the MHz range, which is the expected frequency of the Mack second-mode instability at these conditions, and also consistent with predictions from PSE analysis. A strong unit Reynolds number effect is observed, which is believed to arise from tunnel noise. NTr for air from 5.4 to 13.2 is computed, substantially higher than previously reported for noisy facilities. Time- and spatially-resolved heat transfer traces are used to track the propagation of turbulent spots, and convection rates at 90%, 76%, and 63% of the boundary layer edge velocity, respectively, are observed for the leading edge, centroid, and trailing edge of the spots. A model constructed with these spot propagation parameters is used to infer spot generation rates from measured transition onset to completion distance. Finally, a novel method to control transition location with boundary layer gas injection is investigated. An appropriate porous-metal injector section for the cone is designed and fabricated, and the efficacy of injected CO2 for delaying transition is gauged at various mass flow rates, and compared with both no injection and chemically inert argon injection cases. While CO2 injection seems to delay transition, and argon injection seems to promote it, the experimental results are inconclusive and matching computations do not predict a reduction in N factor from any CO2 injection condition computed.

Investigation of Freestream Plasma Flow Produced by Inductively Coupled Plasma Wind Tunnel

Abstract: The freestream air plasma jet produced by the Plasmatron facility at the von Karman Institute for Fluid Dynamics is investigated. Intrusive and nonintrusive measurement techniques are combined to characterize the plasma flow at pressures between 1500 and 20,000 Pa and electrical power between 120 and 300 KW, covering a broad range of atmospheric entry conditions relevant for thermal protection systems testing. The high-speed camera imaging technique is applied to investigate the unsteadiness features of the plasma jet, and a novel frequency-spatial-domain elaboration procedure is proposed. The results show that the power supply ripple and the test chamber static pressure are the main driving parameters of the plasma jet unsteadiness. The optical emission spectroscopy measurement allow the conclusion that the freestream flow is in local thermochemical equilibrium in nearly the whole operating envelope investigated. Comparison between spectroscopy measurement and freestream temperature calculations based on...

Simulation of Second-Mode Instability in a Real-Gas Hypersonic Flow with Graphite Ablation

Abstract: A new high-order shock-fitting method with thermochemical nonequilibrium and finite-rate chemistry boundary conditions for graphite ablation is presented. The method is suitable for direct numerical simulation of boundary-layer transition in a hypersonic real-gas flow with graphite ablation. The new method is validated by comparison with three computational datasets and one set of experimental data. Direct numerical simulations were run for a 7 deg half-angle blunt cone at Mach 15.99 to find how graphite ablation and thermochemical nonequilibrium affect boundary-layer receptivity and instability. The real-gas simulation is compared with ideal-gas simulations that set their wall temperature and wall blowing from the real-gas simulation. Weak planar fast-acoustic waves in the freestream are used to perturb the steady base flow. A 525 kHz second-mode wave was found to be significantly unstable for the real-gas simulation, whereas in the ideal-gas simulations, no significant flow instability was seen. For the...

Penetration Characteristics of Air, Carbon Dioxide and Helium Transverse Sonic Jets in Mach 5 Cross Flow

Abstract: An experimental investigation of sonic air, CO2 and Helium transverse jets in Mach 5 cross flow was carried out over a flat plate. The jet to freestream momentum flux ratio, J, was kept the same for all gases. The unsteady flow topology was examined using high speed schlieren visualisation and PIV. Schlieren visualisation provided information regarding oscillating jet shear layer structures and bow shock, Mach disc and barrel shocks. Two-component PIV measurements at the centreline, provided information regarding jet penetration trajectories. Barrel shocks and Mach disc forming the jet boundary were visualised/quantified also jet penetration boundaries were determined. Even though J is kept the same for all gases, the penetration patterns were found to be remarkably different both at the nearfield and the farfield. Air and CO2 jet resulted similar nearfield and farfield penetration pattern whereas Helium jet spread minimal in the nearfield.

Role of data uncertainties in identifying the logarithmic region of turbulent boundary layers

Abstract: Composite expansions based on the log-law and the power-law were used to generate synthetic velocity profiles of zero pressure gradient turbulent boundary layers (TBLs) in the range of Reynolds number $$800 \le Re_{\theta } \le 860{,}000,$$ based on displacement thickness and freestream velocity. Several artificial errors were added to the velocity profiles to simulate typical measurement uncertainties. The effects of the simulated errors were studied by extracting log-law and power-law parameters from all these pseudo-experimental profiles. Various techniques were used to establish a measure of the deviations in the overlap region. When parameters extracted for the log-law and the power-law are associated with similar levels of deviations with respect to their expected values, we consider that the profile leads to ambiguous conclusions. This ambiguity was observed up to $$Re_{\theta }=16{,}000$$ for a 4 % dispersion in the velocity measurements, up to $$Re_{\theta }=8.6 \times 10^{5}$$ for a 400 $$\upmu$$ m uncertainty in probe position (in air flow at atmospheric pressure), and up to $$Re_{\theta }=32{,}000$$ for 3 % uncertainty in the determination of $$u_{\tau }.$$ In addition, a new method for the determination of the log-law limits is proposed. The results clearly serve as a further note for caution when identifying either a log or a power-law in TBLs. Together with a number of available studies in the literature, the present results can be seen as a additional reconfirmation of the log-law.

Numerical Investigation of a Supersonic Cavity Flameholder

Abstract: Simulation results are presented for non-reacting flow within a supersonic cavity flameholder. The freestream is air at Mach 2. A case is simulated with no fuel injection, and two cases are simulated with different rates of ethylene fuel injected through holes located on the back face of the cavity. The simulations correspond to a series of experiments for which particle image velocimetry measurements of two velocity components were made within the cavity. The flow within the cavity is computed using unsteady hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation, as well as steady-state Reynolds-averaged NavierStokes simulations. A thorough grid resolution study is presented in which the resolution required to resolve the flow within the cavity is determined. The effects of wall temperature and the thickness of the oncoming turbulent boundary on the solutions is examined and found not to affect the mean velocity and turbulence, and to affect mean mixing only slightly. For steady-state Reynolds-Averaged Navier-Stokes simulations, several turbulence models are used and compared to the hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation results, and the effects of the turbulent Schmidt number, are investigated. The influence of the side walls are investigated by comparing simulations of the full-width duct to simulations of a partial-width duct that uses periodic boundary conditions. The results of the simulations are also compared to the velocity measurements from the experiments, and the hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation results are found to compare well with the experiment in most locations.

Transient Growth in Hypersonic Boundary Layers

Abstract: This paper investigates the relative importance of modal and non-modal growth mechanisms in flat-plate, hypersonic boundary layers as well as the effects of Mach number and wall cooling on these processes. Optimal disturbances are calculated in both the spatial and temporal frameworks using an eigenvector decomposition of the locally-parallel, linearized Navier-Stokes equations. It is found that for every Mach number there is an optimal level of wall cooling that minimizes transient growth; at this condition the wall temperature is slightly below the freestream temperature, with lower wall temperatures needed as the Mach number increases. The competition between modal and non-modal growth mechanisms is examined over a range of Reynolds numbers by calculating N factor curves for both processes. For conditions relevant to high enthalpy flows (high Mach number, cold wall), transient growth is rapidly overtaken by modal instabilities while the level of amplification remains small. At lower Mach numbers or adiabatic conditions, the transient growth is overtaken more slowly. For low supersonic Mach numbers and cold walls there are no modal instabilities, but the level of non-modal amplification is increased such that the initiation of transition by infinitesimal perturbations is plausible despite the absence of modal instabilities.

Numerical Study of Hypersonic Boundary-Layer Receptivity with Freestream Hotspot Perturbations

Abstract: This paper presents a numerical-simulation study of transient flow over a blunt compression cone under the effect of freestream hotspot perturbations. This study is motivated by concurrent wind-tunnel laser-spot experiments carried out at Purdue University. The flow conditions used in the simulation are based on the experimental conditions. The simulation is performed using a high-order shock-fitting finite-difference scheme. The simulation results show that the hotspot is able to excite second-mode instability, where the instability growth is found to be dominant in the boundary layer. The receptivity mechanism is investigated by comparing the simulated results with linear-stability theory. Fast acoustic waves generated by hotspot–shock interaction excite the boundary-layer disturbances. Also, the synchronization of mode F and mode S leads to the dominance of boundary-layer disturbances by the growing second mode.

Influence of Coolant Density on Turbine Platform Film-Cooling With Stator–Rotor Purge Flow and Compound-Angle Holes

Abstract: A detailed parametric study of film-cooling effectiveness was carried out on a turbine blade platform. The platform was cooled by purge flow from a simulated stator-rotor seal combined with discrete hole film-cooling. The cylindrical holes and laidback fan-shaped holes were accessed in terms of film-cooling effectiveness. This paper focuses on the effect of coolant-to-mainstream density ratio on platform film-cooling (DR = 1 to 2). Other fundamental parameters were also examined in this study — a fixed purge flow of 0.5%, three discrete-hole film-cooling blowing ratios between 1.0 and 2.0, and two freestream turbulence intensities of 4.2% and 10.5%. Experiments were done in a five-blade linear cascade with inlet and exit Mach number of 0.27 and 0.44, respectively. Reynolds number of the mainstream flow was 750,000 and wad based on the exit velocity and chord length of the blade. The measurement technique adopted was conduction-free pressure sensitive paint (PSP) technique. Results indicated that with the same density ratio, shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The optimum blowing ratio of 1.5 exists for the cylindrical holes, whereas the effectiveness for the shaped holes increases with increase of blowing ratio. Results also show that the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity.Copyright © 2013 by ASME

Improved Understanding of Aerodynamic Damping Through the Hilbert Transform

Abstract: The equation of motion used to derive the aerodynamic damping coefficient for a single-degree-of-freedom airfoil oscillating in pitch about its quarter-chord is rewritten in analytic signal form through application of the Hilbert transform. The results yield a mathematical framework that can be used to estimate the aerodynamic damping coefficient throughout the entire pitch cycle. The analysis is then applied to experimental data from attached, light, and deep dynamic stall conditions at freestream Mach numbers ranging from 0.2 to 0.6 and Reynolds numbers up to 3.5×106. The Hilbert-transform-based approach is used to demonstrate that the cycle-integrated aerodynamic damping coefficient masks the physics underlying the stabilizing and destabilizing mechanisms of the dynamic stall process. In particular, conditions that exhibit positive cycle-integrated aerodynamic damping may include time intervals of negative aerodynamic damping during the pitch cycle.

Supersonic Combustion Processes in a Premixed Three-Dimensional Nonuniform-Compression Scramjet Engine

Abstract: A numerical study was undertaken using the commercial computational-fluid-dynamics code CFD++ to analyze the combustion behavior of a three-dimensional nonuniform-compression scramjet. Work here was inspired by Ferri’s concept of thermal compression and extended this to include the influence of three-dimensional viscous flow phenomena on the combustion behavior. A premixed H2/air mixture with an equivalence ratio of 1 was used to eliminate the influence of the fuel-injection method on the combustion characteristics. The freestream properties corresponded to a Mach 10 flight condition that gave a flow enthalpy, total temperature, and total pressure of 4.5 MJ/kg, 4676 K, and 29 MPa, respectively. The combustion flame was found to propagate throughout the entire bottom wall of the combustor. Three flame propagation processes were identified: three-dimensional flow structures that provide ignition sources within the boundary layer, radical transport within a three-dimensional shock-induced boundary-layer sep...

Dynamic Store Release of Ice Models from a Cavity into Mach 2.9 Flow

Abstract: An investigation was conducted into store separation from a cavity in a Mach 2.94 freestream using both experimental and computational methods. Both approaches used an open cavity with a length-to-depth ratio of 4.5, and for the sake of simplicity, release of a spherical model was analyzed. The experimental process used a piezoresistive pressure transducer to collect the time-varying content of the pressure signal, while schlieren visualization and high-speed photography capture the dynamic response of a store released from the cavity. Computationally, the OVERFLOW solver was applied with higher-order numerical methods, Chimera grids, and the delayed detached-eddy simulation/shear-stress transport hybrid turbulence model. Tests were performed in a blowdown tunnel exhausting to a vacuum, which enabled robust control of the total pressure, and computational conditions were selected to match the experiment. The studies demonstrated that the shock wave formed on the bottom surface of the sphere led to the los...

Passive Laminar Flow Control at Low Turbulence Levels

Abstract: S PANWISE arrays of discrete roughness elements (DREs) are used in wind tunnel and flight experiments to excite stationary crossflow vortices (a recent example is [1]). Because flight turbulence intensities are low, the stationary crossflow mode dominates the traveling mode in that environment and leaves surface roughness as the important initiator of stationary crossflow vortices [2]. A precursor to using arrays of DREs, the experiments by Radeztsky et al. [3] examined the role of isolated roughness elements on steady crossflow mode-packet initiation and growth. Small circular roughness elements (height k of 6 μm and diameter d less than half of the most unstable crossflow mode wavelength) applied near the neutral stability point effectively initiated crossflow vortices in a swept-wing boundary layer. As a result, the average transition location across the span x∕cjtr was advanced upstream by as much as 0.30c. Reibert et al. [4] used arrays of DREs to excite uniform series of stationary crossflow vortices. In addition to nonlinear saturation of the disturbance amplitudes, several important observations regarding theDRE arrayweremade.When the roughness spacing λkwas that of the most unstable or critical crossflow wavelength (λk λcr 12 mm in those experiments), the strong growth of disturbances modulatedwith λcr and λcr∕2was observed. Increasing the roughness spacing to λk 36 mm (3λcr) produced a crossflow disturbance spectrally composed of λk and its first eight harmonics (18–4 mm). No subharmonics were observed in the wavelength spectrum. Saric et al. [5] recognized the importance of the absent subharmonics and used a DRE array with λk 8 mm to excite a stationary crossflow disturbance that was eventually less unstable than the naturally arising λcr disturbance. In doing so, the growth of λcr was suppressed. The remarkable effect of these so-called control DREs in this experiment was to extend laminar flow past the pressure minimum. The potential for passive control of the crossflow instability stimulated a series of wind-tunnel and flight experiments examining this prospect. To extend the results of Saric et al. [5] to higherRec, the flight experiments of Carpenter et al. [6] investigated the effectiveness of DREs in a 30 deg swept-wing boundary layer at Rec 7.5 × 10. Applique DREs and variable-height DREs spaced at λcr (4.5 mm, in this case) were used. Although measurements from a spanwise array of hot-film sensors confirmed that the roughness did excite disturbances of the expected wavelength, x∕cjtr was unchanged below a critical roughness height. Continuing these experiments, Saric et al. [7] delayed the transition using 24-μm-high DREs spaced at λk 2.25 mm (half of the most unstable wavelength) with a painted leading edge; the extent of laminar flow was increased to 0.60c. In recent wind-tunnel experiments, Hunt and Saric [8] demonstrated that the amplitude of disturbances created by the DREs increases linearlywith k. In these experiments, DREs spaced at λcr 12 mm moved the transition forward via stationary crossflow mode growth. However, the expected control using DREs with λk 6 mm was not achieved. Reduced freestream turbulence in the Klebanoff–Saric wind tunnel (KSWT) might be responsible for rendering the control DREs ineffective in these experiments. This possibility prompted Downs [9] to examine the effect of moderate freestream turbulence on crossflow instability using one option for control roughness with λk 6 mm. Although increasing the turbulence intensity Tu was shown to destabilize the boundary layer in the baseline and critical roughness configurations, little change to x∕cjtr was observed with control roughness at Rec 2.8 × 10. The role of higher freestream turbulence was also examined through experimentation by Muller and Bippes [10]. Although control was achieved at the Arizona State University (ASU) unsteady wind tunnel (UWT) by Saric et al. [5], it has proven difficult to realize both in flight (successful control was shown only once by Saric et al. [7]) and at the Texas AM the levels measured in flight and at the KSWT are lower than at the ASU UWT. The objective of this work is to determine if laminar flow control can be achieved through the use of DREs at Tu < 0.04%. To answer this question, the transition locations behind DRE arrays of various wavelengths and heights are measured and compared to baseline roughness cases. These experiments are conducted in the KSWT, which has freestream turbulence intensity (vortical component) of approximately 0.02% [11] (u 0 rms values quoted in [11] are the complete signal; this value is computed using a sound/turbulence separation technique). Previous wind-tunnel experiments in which DRE transition control was effective were conducted in a freestream turbulence intensity of approximately 0.04% [12].

Scaling of Film Cooling Performance From Ambient to Engine Temperatures

Abstract: The present study employs Computational Fluid Dynamics (CFD) to explore the complexities of scaling film cooling performance measurements from ambient laboratory conditions to high temperature engine conditions. In this investigation, a single shaped hole is examined computationally at both engine and near ambient temperatures to understand the impact of temperature dependent properties on scaling film cooling performance. By varying select flow and thermal parameters for the low temperature cases and comparing the results to high temperature flow, the parameters which must be matched to scale film cooling performance are determined.The results show that only matching the density and mass flux ratios is insufficient for scaling to high temperatures. In accordance with convective heat transfer fundamentals, freestream and coolant Reynolds numbers and Prandtl numbers must also be matched to obtain scalable results. By virtue of the Prandtl number for air remaining nearly constant with temperature, the Prandtl number at ambient conditions is sufficiently matched to engine temperatures. However, laboratory limitations can prevent matching both the freestream and coolant Reynolds numbers simultaneously.By examining this trade-off, it is determined that matching the coolant Reynolds number produces the best scalability. It is also found that by averaging the adiabatic effectiveness of two experiments in which the freestream and coolant Reynolds number are matched respectively results in significantly better scalability for cases with a separated coolant jet.

Roughness-Induced Transient Growth on a Hypersonic Blunt Cone

Abstract: The effects of surface roughness on boundary-layer disturbance growth and laminar-toturbulent transition are not well understood, especially in hypersonic boundary layers. The transient growth mechanism that produces algebraic growth of streamwise streaks via decaying streamwise vortices may play a key role in roughness-induced transition but has not previously been deliberately observed in hypersonic flow. To make such measurements, the present work studies the boundary layer of a 5° half-angle smooth cone paired with a slightly blunted nosetip and a ring of periodically-spaced discrete roughness elements. Measurements are made in the low-disturbance Texas A&M Mach 6 Quiet Tunnel. The roughness element height is approximately equal to the boundary-layer thickness, yet no transition to turbulence is observed for freestream unit Reynolds numbers between 8.7 × 10 6

Passively controlled supersonic cavity flow using a spanwise cylinder

Abstract: An experimental investigation of a passively controlled open cavity with a length to depth ratio of six and freestream Mach number of 1.4 was conducted to investigate the mechanisms responsible for the observed surface pressure reductions. The passive control comes from placing a spanwise aligned cylinder in the boundary layer near the leading edge of the cavity. The two control configurations were isolated from previous experiments of the fluctuating surface pressure and correspond to a larger diameter rod near the top of the boundary layer and a smaller diameter rod placed near the wall. These were further analyzed using particle image velocimetry in an attempt to elicit the responsible mechanism for the flow control. The use of two-point statistics revealed the wall normal turbulent velocity correlation’s evolution became elongated in the wall normal direction. This suggests that the shear layer may be less-organized and consists of smaller-scale structures. The disturbance of the feedback receptivity loop is clearly demonstrated for the controlled configurations evidenced by weakened correlation signals between the aft wall sensor and positions on the cavity floor. The presence of the rod is shown to decrease the mean shear gradient, more effectively for the large rod placed at the top of the boundary layer, throughout the shear layer. The efficacy of the control leads to an initially thicker shear layer which spreads more rapidly and is clearly demonstrated by vorticity growth rates, mean, and turbulent flowfield statistics.

Effects of dimples on laminar boundary layers

Abstract: A series of direct numerical simulations of the flow past a flat plate with two and eight rows of dimples in a staggered arrangement is carried out. The Reynolds number based on the boundary layer thickness and freestream velocity near the inflow plane is 1000 and the dimples are spherical with a depth to diameter ratio of 0.1. The incoming flow is laminar and the boundary layer thickness before the dimples is half the dimple depth. At this low Reynolds number the flow is expected to remain laminar over a smooth flat plate. The presence of the dimples triggers instabilities that cause significant momentum transport. It is shown that the shear layer that forms as the flow separates over the first two rows of dimple becomes unstable and sheds coherent vortex sheets. The vortex sheets become unstable and are transformed into packets of horseshoe vortices. When these vortices evolve over a flat plate or over a series of dimples the flow dynamics are very different with important changes in momentum transport ...

Conceptual Analysis of Electron Transpiration Cooling for the Leading Edges of Hypersonic Vehicles

Abstract: Recent progress is presented in an ongoing effort to perform a conceptual analysis of possible electron transpiration cooling using thermo-electric materials at the leading edges of hypersonic vehicles. The implementation of a new boundary condition in the CFD code LeMANS to model the thermionic emission of electrons from the leading edges of hypersonic vehicles is described. A parametric study is performed to understand the effects of the material work function, the freestream velocity, and the leading edge geometry on this cooling effect. The numerical results reveal that lower material work functions, higher freestream velocities, and smaller leading edges can increase the cooling effect due to larger emission current densities. The numerical results also show that the electric field produced by the electron emission may not have a significant effect on the predicted properties. Future work recommendations are provided that may improve the physical accuracy of the modeling capabilities used in this study.

Experimental Study of Supersonic Entrainment Using a Cavity

Abstract: The effect of wall-mounted three-dimensional cavities on the entrainment of gaseous injectant into a supersonic stream of Mach number 1.7 is studied. Entrainment induced by flow unsteadiness, turbulence, and exchange of fluid between the cavity and mainstream flow has been investigated. To clarify the influence of the ratio of cavity length to width L/W on the fluid dynamic behavior, cavities of L/W ratios between two and four were applied at a constant length-to-depth ratio L/D. Acoustic oscillations generated from the cavity were observed to have profound impact on mass exchange between the cavity and the freestream flow. These acoustic oscillations were in turn found to be dependent on the L/W ratio of the cavity. Shift in dominant acoustic mode was observed as the L/W ratio was changed from three to four. Spillage of the shear layer over the cavity walls and turbulence induced in the cavity were probed by measuring velocity using laser Doppler velocimetry. Entrainment of secondary gaseous injection in...