Showing papers on "Freestream published in 2007"

Direct numerical simulation of supersonic turbulent boundary layer over a compression ramp

Abstract: A direct numerical simulation of shock wave and turbulent boundary layer interaction for a 24 deg compression ramp configuration at Mach 2.9 and Re θ 2300 is performed. A modified weighted, essentially nonoscillatory scheme is used. The direct numerical simulation results are compared with the experiments of Bookey et al. at the same flow conditions. The upstream boundary layer, the mean wall-pressure distribution, the size of the separation bubble, and the velocity profile downstream of the interaction are predicted within the experimental uncertainty. The change of the mean and fluctuating properties throughout the interaction region is studied. The low frequency motion of the shock is inferred from the wall-pressure signal and freestream mass-flux measurement.

Transition Mechanisms in Separation Bubbles Under Low and Elevated Freestream Turbulence

Abstract: Through numerical simulations, this paper examines the nature of instability mechanisms leading to transition in separation bubbles. The results of two direct numerical simulations are presented in which separation of a laminar boundary layer occurs over a flat surface in the presence of an adverse pressure gradient. The primary difference in the flow conditions between the two simulations is the level of freestream turbulence with intensities of 0.1% and 1.45% at separation. In the first part of the paper, transition under a low-disturbance environment is examined, and the development of the Kelvin-Helmholtz instability in the separated shear layer is compared to the well-established instability characteristics of free shear layers. The study examines the role of the velocity-profile shape on the instability characteristics and the nature of the large-scale vortical structures shed downstream of the bubble. The second part of the paper examines transition in a high-disturbance environment, where the above-mentioned mechanism is bypassed as a result of elevated freestream turbulence. Filtering of the freestream turbulence into the laminar boundary layer results in streamwise streaks which provide conditions under which turbulent spots are produced in the separated shear layer, grow, and then merge to form a turbulent boundary layer. The results allow identification of the structure of the instability mechanism and the characteristic structure of the resultant turbulent spots. Recovery of the reattached turbulent boundary layer is then examined for both cases. The large-scale flow structures associated with transition are noted to remain coherent far downstream of reattachment, delaying recovery of the turbulent boundary layer to an equilibrium state.Copyright © 2007 by ASME

Plasma Actuator Force Measurements

Abstract: In previous work at the U.S. Air Force Academy, the phenomenology and behavior of the aerodynamic plasma actuator, a dielectric barrier discharge plasma, was investigated. To provide insight into the phenomenology associated with the transfer of momentum to air by a plasma 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 (5-10-kV amplitude). The body forces on the air were calculated using a control volume momentum balance. In a second experiment, time-averaged results were also obtained by measuring the reaction force using a pendulum. A third experiment uses an accelerometer to gain insight into the time-dependent forces or, more specifically, the direction of the forces. The results show that the body force acts within the first 4 mm above the surface of the actuator (within the boundary layer). For a constant peak-to-peak 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 time-dependent measurements support the theory that the body force of the actuator consists of one large push followed by one small pull during each cycle.

Dual-Pump Coherent Anti-Stokes Raman Scattering Measurements in a Supersonic Combustor

Abstract: The dual-pump coherent anti-Stokes Raman scattering (CARS) method was used to measure temperature and the mole fractions of N 2 and O 2 in a supersonic combustor. Experiments were conducted in NASA Langley Research Center's Direct-Connect Supersonic Combustion Test Facility. In this facility, H 2 - and oxygen-enriched air burn to increase the enthalpy of the simulated air test gas. This gas is expanded through a Mach 2 nozzle and into a combustor model consisting of a short constant-area section followed by a small rearward-facing step and another constant-area section. At the end of this straight section, H 2 fuel is injected at Mach 2 and at a 30-deg angle with respect to the freestream. One wall of the duct then expands at a 3-deg angle for over 1 m. The ensuing combustion is probed optically through ports in the side of the combustor. Dual-pump CARS measurements were performed at the facility nozzle exit and at four planes downstream of fuel injection. Maps are presented of the mean temperature, as well as N 2 and O 2 mean mole-fraction fields. Correlations between fluctuations of the different measured parameters are also presented.

Comparative Investigation of Three Highly Loaded LP Turbine Airfoils: Part I — Measured Profile and Secondary Losses at Design Incidence

Abstract: At the 2006 ASME-IGTI Turbo-Expo, low-speed cascade results were presented for the midspan aerodynamic behaviour of a family of three highly loaded low-pressure (LP) turbine airfoils operating over a wide range of Reynolds numbers (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%. All three airfoils have the same design inlet and outlet flow angles. The baseline cascade has a Zweifel coefficient of 1.08 and the two additional blade rows have values of 1.37. The new, more highly-loaded blade rows differ mainly in their loading distributions: one is front-loaded while the other is aft-loaded. The new front-loaded airfoil was found to have particularly attractive profile performance. Despite its exceptionally high value of Zweifel coefficient, it was found to be free of a separation bubble on its suction side at Reynolds numbers as low as 50,000, and this was reflected in very good profile loss behaviour. However, it was also noted in the earlier paper that the choice of a particular loading level and loading distribution would be influenced by more than its profile performance at design incidence. The present two-part paper extends the midspan aerodynamic comparison of the three airfoils to the secondary flow performance. The first part of the paper discusses both the profile and secondary flow performance of the three cascades at their design Reynolds number of 80,000 (or ∼ 125,000 based on exit velocity) for two freestream turbulence intensities of 1.5% and 4%. The secondary flow behaviour was determined from detailed flowfield measurements made at 40% axial chord downstream of the trailing edge using a seven-hole pressure probe. In addition to providing total pressure losses, the seven-hole probe measurements were also processed to give the downstream vorticity distributions. As has been found in other secondary flow investigations in turbine cascades, the present front-loaded airfoil showed higher secondary losses than the aft-loaded airfoil with the same value of Zweifel coefficient.Copyright © 2007 by ASME

Large-Eddy Simulation of Compressible Transitional Flows in a Low-Pressure Turbine Cascade

Abstract: Large-eddy simulation of compressible transitional flows in a low-pressure turbine cascade is performed by using sixth-order compact difference and a 10th-order filtering method. Numerical results without freestream turbulence and those with about 5 % of freestream turbulence are compared. In these simulations, separated flows in the turbine cascade accompanied by laminar-turbulent transition are realized, and the present results closely agree with past experimental measurements in terms of the static pressure distribution around the blade. In the case where no freestream turbulence is taken into account, the unsteady pressure field essentially differs from that with strong freestream turbulence. In the no freestream turbulence case, pressure waves that propagate from the blade's wake region have noticeable effects on the separated-boundary layer near the trailing edge and on the neighboring blade. Also, based on the snapshot proper orthogonal decomposition analysis, dominant behaviors of the transitional boundary layers are investigated.

The Effects of Freestream Turbulence, Turbulence Length Scale and Exit Reynolds Number on Turbine Blade Heat Transfer in a Transonic Cascade

Abstract: This paper experimentally investigates the effect of high freestream turbulence intensity, turbulence length scale, and exit Reynolds number on the surface heat transfer distribution of a turbine blade at realistic engine Mach numbers. Passive turbulence grids were used to generate freestream turbulence levels of 2%, 12%, and 14% at the cascade inlet. The turbulence grids produced length scales normalized by the blade pitches of 0.02, 0.26, and 0.41, respectively. Surface heat transfer measurements were made at the midspan of the blade using thin film gauges. Experiments were performed at the exit Mach numbers of 0.55, 0.78, and 1.03, which represent flow conditions below, near, and above nominal conditions. The exit Mach numbers tested correspond to exit Reynolds numbers of 6 × 10 5 , 8 × 10 5 , and 11 × 10 5 , based on true chord. The experimental results showed that the high freestream turbulence augmented the heat transfer on both the pressure and suction sides of the blade as compared with the low freestream turbulence case. At nominal conditions, exit Mach 0.78, average heat transfer augmentations of 23% and 35% were observed on the pressure side and suction side of the blade, respectively.

Control of the shear layer above a supersonic cavity using energy deposition

Abstract: An experimental study was conducted to investigate the effectiveness of pulsed energy deposition as a means of active flow control for the shear layer above a supersonic cavity in the open configuration. The excitation pulse was generated with a Q-switched Nd:YAG laser and was applied as a spanwise oriented line along the leading edge of a cavity. The study was conducted at a freestream Mach number of 1.4 and for a cavity length-to-depth ratio of 5.29. The flowfield was analyzed over a range of delay times from the excitation laser pulse using schlieren photography and particle image velocimetry. Analysis of phase-averaged schlieren images suggested the formation and growth of a coherent large-scale structure (consisting of two adjoining vortices) in the wake of the generated disturbance. This result was confirmed through two-component velocity field data obtained from particle image velocimetry measurements. The velocity information was also used to determine the instantaneous convective velocity and define characteristic scales for the large-scale structure.

Attenuation of Low-Speed Flow-Induced Cavity Tones Using Plasma Actuators

Abstract: The aeroacoustic environment of an open cavity with a length-to-depth (LID) ratio of 1.0 was studied with and without plasma actuators. The study was conducted through low-speed wind-tunnel experiments at freestream velocities between 10 and 20 m/s, corresponding to Reynolds numbers based on the depth of the cavity of 3.6 x 104 to 7.1 x 10 4 . The fluid flow inside the cavity was studied using a range of measurement techniques, which include oil flow, particle imaging velocimetry, and surface-mounted-microphone measurements. For acoustic control, an array of plasma actuators were located on the approaching surface to the cavity, aligned with the direction of the oncoming flow. Results show that the plasma actuators lead to a significant attenuation of the dominant cavity mode. The particle imaging velocimetry surveys around the electrode elements reveal vortical structures produced by the plasma actuators. These structures convect downstream with the mean flow and produce spanwise variations in the flow over the cavity that affect the spanwise coherence of the shear layer and hence the corresponding Rossiter mode.

Calculation of Thermal Response of Ablator Under Arcjet Flow Condition

Abstract: An integrated computational method is developed to calculate thermal response of ablator under an arcjet flow condition. In the method, the arcjet freestream condition in the test section is evaluated by calculating the flows in the arcjet wind tunnel fully theoretically. The thermal response of the ablator is calculated by loosely coupling the shock layer computational fluid dynamics code and the 2-D version of ablation code using the arcjet freestream condition so evaluated. The method is applied to the heating tests conducted in the 1 MW arcjet wind tunnel for one operating condition. The influence of catalytic conditions of ablating surface and the effect of nitridation reaction and surface roughness on the thermal response of the ablator are investigated. Comparison of the temperature profile at the ablating surface between calculation and measurement suggests that the measured temperature profile can be reproduced with a low catalytic efficiency of the surface. It is found that the nitridation reaction increase the surface temperature moderately, and that the effect of the roughness on the surface were small for the present operating condition.

Effects of upstream coherent structures on low-frequency motion of shock-induced turbulent separation

Abstract: x Streamwise direction z Spanwise direction y Wall-normal direction U Streamwise velocity (m/s) U∞ Streamwise freestream velocity (m/s) uτ Skin friction velocity M Mach number δ Boundary layer thickness θ Momentum thickness δ∗ Displacement thickness Um Mean streamwise velocity at any given location in the boundary layer σu R.M.S streamwise velocity at any given location in the boundary layer xsep Separation location (located using a threshold technique) Ul Mean streamwise velocity along a streamwise line upstream of the separation location T Static temperature Cp Specific heat at constant pressure

Reacting and Nonreacting Flowfields of a V-Gutter Stabilized Flame

Abstract: Particle image velocimetry measurements have been performed to determine the flowfield around a confined V-gutter bluff body in both nonreacting and reacting environments. The incoming flow to the V-gutter bluff body is a vitiated mixture with secondary liquid fuel being injected upstream through a fuel spray bar to mimic realistic conditions. The measurements capture instantaneous and mean flow structures formed in the wake region of a three-dimensional bluff body. In addition to the flow structures, the mean velocity components, mean out-of-plane vorticity, and the turbulent kinetic energy have been compared between the nonreacting and reacting test cases. The results show significant differences in the instantaneous flow structures. The nonreacting case results in asymmetric shedding of large-scale vortical structures which span the entire wake region, whereas the reacting case results in both symmetric and asymmetric shedding of smaller scale vortical structures which are flattened out within the shear layer. A comparison of the mean velocity components clearly shows that the reacting case results in a larger region of reversed flow, experiences an acceleration of the freestream flow due to combustion, and results in a slower dissipation of the wake region.

Measuring Nitric Oxide Freestream Velocity Using a Quantum Cascade Laser at CUBRC

Abstract: velocity measurements This work reports on freestream velocity measurements for a number of flows for velocities from 2,060 m/s to 5,350 m/s The measured velocity was compared to facility calculations that are based on the measured shock speed in the driven tube and measured reservoir conditions The calculations agree with the velocity measurements up to 10 MJ/kg At higher enthalpies the calculations under predict the actual velocity of the flow However, the velocity measurement provides a new capability to operate the tunnel in non-tailored mode to obtain velocities greater than 5,300 m/s for at least 2 milliseconds of steady flow Under most circumstances the LENS I and LENS II shock tunnels produce fully-duplicated conditions for various flight trajectories in the upper atmosphere At high enthalpy some of the dissociated products are frozen during the expansion into the test section, so non equilibrium chemistry models have been developed to predict the Nitric Oxide (NO) concentration in the free stream gas Since the presence of this contaminant in the flow changes the nature of the test gas, however slightly, computational fluid dynamic (CFD) calculations are used to compensate for these effects 2 The actual concentration of the gas must be known Also it is very desirable to measure the free stream velocity since it is a direct measure of the gas kinetic energy and thus connected to the total enthalpy of the flow, a basic operating parameter of the facility At CUBRC we have constructed a continuous wave Quantum

Zonal-Detached-Eddy Simulation of Projectiles in the Subsonic and Transonic Regimes

Abstract: This paper presents nonspinning-projectile computations with advanced turbulence modeling to demonstrate the relevance of using a hybrid method in projectile simulations. A zonal-detached-eddy simulation methodology is used to improve the base-flow prediction without deteriorating the incoming attached flow upstream of the base in the subsonic and transonic regimes. The numerical results are found to compare fairly well with the available experimental wall-pressure data. Both time-averaged and unsteady features are then discussed. Particular attention was paid to the near-wake flowfield and its dependency upon the freestream Mach number. In these calculations, some classical features of massively separated flows, such as the pressure evolution along the base or the wake centerline characteristics, are poorly predicted with the Spalart-Allmaras model. The use of a hybrid method leads to promising results and allows for a physical analysis of the separated flowfield. The flow appears to exhibit self-similar properties, even in the recirculation area, independently of the Mach number value in the range of freestream conditions investigated. Moreover, the instability process leading to the shear-layer growth is found to be similar in the range of parameters investigated and is in accordance with previous results concerning the compressibility effects in free shear flows. Finally, base-pressure spectra are reported and compared with axisymmetric base-flow data.

Shape Optimization Methodology for Reducing the Sonic Boom Initial Pressure Rise

Abstract: A shape optimization methodology for reducing the initial shock pressure rise on the ground of a supersonic aircraft is presented. This methodology combines elements from the linearized aerodynamic theory, such as the Whitham F function, with elements from the nonlinear aerodynamic theory, such as the prediction of lift distribution by an Euler or a Navier-Stokes flow solver. It is applied to the optimization of two different airplane concepts developed by Reno Aeronautical and Lockheed Martin, respectively, for the Defense Advanced Research Projects Agency's Quiet Supersonic Platform program. For Reno Aeronautical's laminar-flow supersonic aircraft, the initial shock pressure rise on the ground is reduced by a factor close to 2, from 1.224 psf (58.605 N/m 2 ) at a freestream Mach number of 1.5 to 0.671 psf (32.127 N/m 2 ), while maintaining constant lift For Lockheed Martin's point of departure aircraft, a tenfold reduction of the initial shock pressure rise on the ground is demonstrated, from 1.623 psf (77.71 N/m 2 ) at a freestream Mach number of 1.5 to 0.152 psf (7.278 N/m 2 ), also while maintaining constant lift.

Measurements in Film Cooling Flows With Periodic Wakes

Abstract: Film cooling flows subject to periodic wakes were studied experimentally. The wakes were generated with a spoked wheel upstream of a flat plate. Cases with a single row of cylindrical film cooling holes inclined at 35 degrees to the surface were considered at blowing ratios of 0.25, 0.50, and 1.0 with a steady freestream and with wake Strouhal numbers of 0.15, 0.30, and 0.60. Temperature measurements were made using an infrared camera, thermocouples, and constant current (cold wire) anemometry. Hot wire anemometry was used for velocity measurements. The local film cooling effectiveness and heat transfer coefficient were determined from the measured temperatures. Phase locked flow temperature fields were determined from cold wire surveys. Wakes decreased the film cooling effectiveness for blowing ratios of 0.25 and 0.50 when compared to steady freestream cases. In contrast, effectiveness increased with Strouhal number for the 1.0 blowing ratio cases, as the wakes helped mitigate the effects of jet liftoff. Heat transfer coefficients increased with wake passing frequency, with nearly the same percentage increase in cases with and without film cooling. The time resolved flow measurements show the interaction of the wakes with the film cooling jets. Near-wall flow measurements are used to infer the instantaneous film cooling effectiveness as it changes during the wake passing cycle.

Effect of Flow Parameter Variations on Full Coverage Film-Cooling Effectiveness for a Gas Turbine Blade

Abstract: The effect of film cooling holes placed along the span of a fully-cooled high pressure turbine blade in a stationary, linear cascade on film cooling effectiveness is studied using the Pressure Sensitive Paint (PSP) technique. Effect of showerhead injection at the leading edge and the presence of compound angled, diffusing holes on the pressure and suction side are also examined. Six rows of compound angled shaped film cooling holes are provided on the pressure side while four such rows are provided on the suction side of the blade. The holes have a laidback and fan-shaped diffusing cross-section. Another three rows of cylindrical holes are drilled at a typical angle on the leading edge to capture the effect of showerhead film coolant injection. The film cooling hole arrangement simulates a typical film cooled blade design used in stage 1 rotor blades for gas turbines used for power generation. A typical blowing ratio is defined for each film hole row and tests are performed for 100%, 150% and 200% of this typical value. Tests are performed for inlet Mach numbers of 0.36 and 0.45 with corresponding exit Mach numbers of 0.51 and 0.68 respectively. The flow remains subsonic in the throat region for both Mach numbers. The corresponding free stream Reynolds number, based on the axial chord length and the exit velocity, are 1.3 million and 1.74 million respectively. Freestream turbulence intensity level at the cascade inlet is 6%. Results show that varying blowing ratios can have a significant impact on film-cooling effectiveness distribution. Large spanwise variations in effectiveness distributions are also observed. Similar distributions were observed for both Mach numbers.© 2007 ASME

Aerodynamics of a Low-Pressure Turbine Airfoil at Low Reynolds Numbers—Part I: Steady Flow Measurements

Abstract: This two-part paper presents a detailed experimental investigation of the laminar separation and transition phenomena on the suction surface of a high-lift low-pressure (LP) turbine airfoil, PakB. The first part describes the influence of Reynolds number, freestream turbulence intensity and turbulence length scale on the PakB airfoil under steady inflow conditions. The present measurements are distinctive in that a closely-spaced array of hot-film sensors has allowed a very detailed examination to be made of both the steady and unsteady behaviour of the suction surface boundary layer. In addition, this paper presents a technique for interpreting the transition process in steady, and periodically unsteady, separated flows based on dynamic and statistical properties of the hot-film measurements. Measurements were made at Reynolds number varying from 25,000 to 150,000 and for freestream turbulence intensities of 0.4%, 2% and 4%. Two separate grids were used to generate turbulence intensity of 4% with integral length scales of about 10% and 40% of the airfoil axial chord length. The first is comparable with the turbulence length scales expected in the engine and the second is considerably larger. While the higher levels of freestream turbulence intensity promoted earlier transition and a shorter separation bubble, the varying turbulence length scale did not have a noticeable effect on the transition process. The size of the separation bubble increased with decreasing Reynolds number, and under low freestream turbulence levels the bubble failed to reattach at low Reynolds numbers. As expected, the losses increased with the length of the separation bubble on the suction side of the airfoil, and increased significantly when the bubble failed to reattach.Copyright © 2007 by ASME

Calculation of High-Enthalpy Aerothermal Environment in an Arcjet Facility

Abstract: S EGMENTED-CONSTRICTOR type arc-heated wind tunnels are used to test the heat shield materials for spacecraft thermal protection systems. The arc-heated wind tunnel consists of an upstream electrode (anode) chamber, constrictor section, downstream electrode (cathode) chamber, and a diverging–converging nozzle connecting to a test chamber. In the test section, the heat shield materials are exposed to a high-enthalpy flow environment produced by the facility. The high-enthalpy environment is often such that the flowdoes not reach equilibrium condition at the edge of the boundary layer over the tested material. In such a case, we need to calculate the flow properties at the surface of the material using a computational fluid dynamics approach to understand the thermal response of the material [1]. For this purpose, the arcjet freestream conditions must be known accurately. To calculate an arcjet freestream condition, two important physical processes occurring in the arcjet wind tunnel should be accounted for: the heating process in the arc heater region upstream of the nozzle throat and the relaxing process in the expanding nozzle region downstream of the nozzle throat. The ARCFLO3 code has been developed recently to calculate the flowfield in the segmentedconstrictor type of arc heaters [2]. Unlike the arc heater flowfield code named ARCFLO developed in the 1970s [3], which is able to calculate the flow in the constrictor section, this new code calculates the flow from the anode chamber to the nozzle throat [2]. Arcjet freestream conditions can be calculated fully theoretically if a nonequilibrium expanding nozzle calculation is made with the calculated flow properties at the nozzle throat obtained by using the ARCFLO3 code. In addition, because the radial distribution of the flow properties at the nozzle throat is calculated with the ARCFLO3 code, the unified computational method can give the radial flow properties in the arcjet freestream at the test section. We tried to make such a unified calculation very recently for one operating condition in an arcjet facility [1]. However, the question remains as to how well such a computational approach predicts the flow properties in an arcjet freestream. It is the purpose of the present work to test the validity of the unified method. The method is applied to calculate the flowfield in a 0.75-MW arcjet wind-tunnel facility at the Institute of Aerospace Technology of the Japan Aerospace Exploration Agency (IAT/JAXA) in Japan. This facility was chosen for the following reasons: 1) In the recent measurement [4] in the IAT/JAXA arcjet facility, the operational characteristic parameters for a wide range of conditions were obtained. The experimental data offer an opportunity to test the current state of the computational modeling in the proposed method. 2) In our previous work, the ARCFLO3 code was applied for the arc heater flowfield calculation only in a high-power-level arc heater, such as the 20or 60-MW arcjet facility at NASA Ames Research Center [2]. The applicability of the ARCFLO3 code to submegawatt class facilities is unknown.

A Computational Analysis of Thermochemical Studies in the LENS Facilities

Abstract: A review has been presented detailing the comparisons between numerical prediction and experimental data collected in the LENS facilities for a program focusing on thermochemical modeling of the flow in the reflected shock tunnel facility at high enthalpy. Comparisons have been provided for several studies made in the LENS-I facility including fundamental laser diode measurements of freestream nitric oxide concentration, temperature, and velocity; measurements on a two-dimensional cylinder; and measurements of laminar shock-wave/boundary layer interaction on a double cone geometry. The freestream velocity of the facility is found to be predicted to an accuracy of 2.5% or better for well-tailored conditions at all enthalpy levels, but the static temperature is significantly over-predicted for high enthalpy flows. Despite this, good agreement is obtained with all measurements for a two-dimensional cylinder. For the double cone model, the effect of vibration-dissociation coupling with the T-TV and CVDV models has been investigated. The choice of coupling model has impact for the lower enthalpy case, but in the high enthalpy case, it is unclear whether the discrepancy with the measurements is attributable to the coupling in the interaction region or the understanding of the freestream conditions.

Jet-into-Crossflow Boundary-Layer Control: Innovation in Gas Turbine Blade Cooling

Abstract: Jet into crossflow has numerous technological applications, such as in film cooling of gas turbine blades. It has been more than half a century that people have been studying this problem and research is still underway due to its importance and its complexities. This paper is a computational study concerned with film cooling of gas turbine blades. A novel near-wall flow control technique of using staggered arrangement of small injection ports near a film-cooling hole (combined triple jet) is introduced. The fluid injected from the small ports changes the flow pattern downstream, resulting in a considerable enhancement of cooling efficiency. The flowfield computations, governed by the Reynolds-averaged Navier-Stokes equations (incorporating the Reynolds stress turbulence model), were performed using the SIMPLE algorithm on a nonuniform staggered grid. The results show that, due to the introduction of new counter-rotating vortex pairs, this approach provides considerable improvement in 1) film-cooling efficiency, 2) uniform distribution of the coolant film, 3) reduction of the mixing strength between the freestream and the coolant jets, and 4) reduction of skin friction drag. In addition, qualitative comparison of our results with those of regular staggered holes arrangement indicated that this new technique has a considerably higher film-cooling performance.

Direct numerical simulation of supersonic boundary-layer stabilization by porous coatings

Abstract: Two-dimensional direct numerical simulation (DNS) of the porous coating stabilization effect is carried out for near-wall flows over a flat plate, sharp cone and compression corner at freestream Mach numbers 5-6. Numerical data obtained for disturbances propagating in the boundary layer on a flat plate agree satisfactory with the linear stability theory (LST). The coating end effects, which are associated with discontinuity of boundary conditions at the juncture between solid and porous walls, are considered. It is found that the end effects are localized over 2-3 disturbance wavelength and can be neglected in calculations of the integral performance of porous coatings. Receptivity of supersonic boundary layer to fast and slow acoustic waves is modeled. It is shown that the porous coating weakly affects acoustic disturbances and initial amplitudes of the boundary-layer modes, while it strongly suppresses the second-mode amplification. For the compression corner flow, DNS shows that the porous coating weakly affects high-frequency disturbances in the separation region and strongly stabilizes them in the reattached boundary layer downstream from the separation bubble. These numerical studies confirm robustness of the porous coating stabilization concept.

Measurements of Fluctuating Pitot Pressure, "Tunnel Noise," in the AEDC Hypervelocity Wind Tunnel No. 9

Abstract: The Arnold Engineering Development Center (AEDC) Hypervelocity Wind Tunnel No. 9 facility has played a key role in the development of hypersonic vehicles for over 30 years, providing high-quality aerodynamic and aerothermal test data covering high Mach number and high Reynolds number flight simulations. Although Tunnel 9 can achieve flight level Reynolds numbers and naturally transitioning boundary layers on most test articles, the presence of “tunnel noise” can complicate the understanding of the boundary-layer transition phenomenon. In an attempt to better characterize the freestream disturbances described as “tunnel noise” a set of data was collected using a flush-mounted Pitot acoustic probe. The data quantify the relative Pitot acoustic noise of the freestream flow for the Mach 8, 10, and 14 nozzles at AEDC Tunnel 9. The percent noise level for each nozzle varied on the basis of Reynolds number from approximately 2 to 3.5 percent at Mach 8, 2.5 to 4 percent at Mach 10, and 3.75 to 6.25 percent at Mach 14.

Numerical study of transitional separated-reattached flow over surface-mounted obstacles using large-eddy simulation

Abstract: Large-eddy simulation (LES) of transitional separatingreattaching flow on a square surface mounted obstacle has been performed. The Reynolds number based on the uniform inlet velocity and the obstacle height is 4.5 103. A dynamic subgrid-scale model is employed in this work. The mean LES results compare favourably with the available experimental and direct numerical simulation (DNS) data. Extensive analysis of the time series signals of the velocity and pressure fields at different locations including positions close to solid surfaces, at the centre and edge of the separatedreattached boundary layer using the windowed Fourier transform (WFT) and the wavelet transform was performed. The spectra analysis revealed the nature of the amplified frequencies at all the important locations of the flow field. Excited modes that could be due to the movement (shedding) of large-scale structures and pairing of such types of structures are identified. A clear frequency peak was captured just upstream of the separation line. The value of the frequency peak and the low percentage of the back flow velocity compared to the freestream velocity in the current case strongly support the idea that this amplified frequency is most likely due to the KelvinHelmholtz (KH) instability mechanism of the shear layer forming in the boundary of the small upstream separated region rather than being attributed to the flapping of the shear layer

Vitiation Effects on Scramjet Engine Performance in Mach 6 Flight Conditions

Abstract: Quasi-one-dimensional analyses with either chemical equilibrium or finite-rate reaction were conducted to evaluate the effects of contamination due to flow vitiation on engine performances, namely, thrust production. The incoming flow state was calculated based on measurements, and a higher total enthalpy for the freestream through a vitiation heater compared to that through a storage heater was found, the difference in the flow condition being responsible for two-thirds of the reduction in the thrust production with vitiation observed in the engine tests. With the finite-rate reaction calculation, the possibility of a further reduction due to the contamination effect on the combustion mechanism was found, which was responsible for one-third of the measured reduction. The one-dimensional analyses were further pursued to find matched test conditions with the storage heater to that with the vitiation heater in view of the thrust production and pressure distribution within the engine, and both freestream enthalpy and fuel equivalence ratio should be adjusted to attain the matched conditions.

Aerodynamics of a low-pressure turbine airfoil at low reynolds numbers-part ii: Blade-wake interaction

Abstract: The relative motion of rotor and stator blade rows causes periodically unsteady flows that influence the performance of airfoils through their effects on the boundary layer development. Part 1 of this two-part paper described the influence of Reynolds number, freestream turbulence intensity and turbulence length scales on a low-pressure (LP) high-lift turbine airfoil, PakB, under steady inlet flow conditions. The aerodynamic behaviour of the same airfoil under the influence of incoming wakes is presented in Part 2. The unsteady effects of wakes from a single upstream blade-row were measured in a low-speed linear cascade facility at Reynolds numbers of 25000, 50000 and 100000 and at two freestream turbulence intensity levels of 0.4% and 4%. In addition, eight reduced frequencies between 0.53 and 3.2, at three flow coefficients of 0.5, 0.7 and 1.0 were examined. The complex wake-induced transition, flow separation and reattachment on the suction surface boundary layer was determined from an array of closely-spaced surface hot-film sensors. The wake-induced transition caused the separated boundary layer to reattach to the suction surface at all conditions examined. The time-varying profile losses were measured downstream of the trailing edge. Profile losses increase with decreasing Reynolds number and the influence of increased freestream turbulence intensity is only evident in between wake-passing events at low reduced frequencies. At higher values of reduced frequency, the losses increase slightly and for the cases examined here, losses were slightly larger at lower flow coefficients than the higher flow coefficients. An optimum wake-passing frequency was observed at which the profile losses were a minimum.Copyright © 2007 by ASME

Effect of Turbulence Intensity and Periodic Unsteady Wake Flow Condition on Boundary Layer Development, Separation, and Reattachment Along the Suction Surface of a Low-Pressure Turbine Blade

Abstract: This paper experimentally investigates the individual and combined effects of periodic unsteady wake flows and freestream turbulence intensity (FSTI) onflow separation along the suction surface of a low-pressure turbine blade. The experiments were carried out at a Reynolds number of 110,000 based on the suction surface length and the cascade exit velocity. The experimental matrix includes freestream turbulence intensities of 7.9%, 3.0%, 8.0%, and 13.0%, and three different unsteady wake frequencies with the steady inlet flow as the reference configuration. Detailed boundary layer measurements are performed along the suction surface of a highly loaded turbine blade with a separation zone. Particular attention is paid to the aerodynamic behavior of the reparation zone at different FSTIs at steady and periodic unsteady flow conditions. The objective of the research is (i) to quantify the effect of FSTIs on the dynamics of the separation bubble at steady inlet flow conditions and (ii) to investigate the combined effects of Tu in and the unsteady wake flow on the behavior of the separation bubble.