Blowdown Wind Tunnel Research Articles

Direct Numerical Simulation Database of High-Speed Flow over Parameterized Curved Walls

This study presents a direct numerical simulation (DNS) database of high-speed turbulent boundary layers (TBLs) subject to pressure gradients due to parametrically varied backward-facing and forward-facing wall curvatures, with an inflow Mach number of 4.9 and a friction Reynolds number of Reτ≃1000 immediately before the onset of wall curvature. The Mach and Reynolds numbers are significantly higher than those reported in the literature for the DNS of pressure-gradient TBLs. The flow conditions and baseline wall geometries are representative of experiments in the high-speed blowdown wind tunnel at the National Aerothermochemistry Laboratory at Texas A&M University. The wall steepness of the baseline geometry for both the backward-facing and forward-facing walls was systematically varied to cause attached, incipiently separated, and fully separated flows. Precomputed flow statistics, including turbulent kinetic energy budgets, are available on the website of the Turbulence Modeling Resource of the NASA Langley Research Center, allowing other investigators to query any property of interest.

Turbulence Measurements Downstream of a Combustor Simulator Designed for Studies on the Combustor–Turbine Interaction

Turbulence intensity impacts the performance of turbine stages and it is an important inlet boundary condition for CFD computations; the knowledge of its value at the turbine inlet is then of paramount importance. In combustor–turbine interaction experimental studies, combustor simulators replace real combustors and allow for the introduction of flow perturbation at the turbine inlet. Therefore, the turbulence intensity of a combustor simulator used in a wide experimental campaign at Politecnico di Milano is characterized using a hot-wire probe in a blow-down wind tunnel, and the results are compared to URANS CFD computations. This combustor simulator can generate a combination of a swirl profile with a steady/unsteady temperature disturbance. In the cold unsteady disturbance case, hot-wire measurements are phase-averaged at the frequency of the injected perturbation. The combustor simulator turbulence intensity is measured at two different axial positions to understand its evolution.

Influence of deep gaps on laminar–turbulent transition in two-dimensional boundary layers at subsonic Mach numbers

Systematic experimental investigations on the influence of deep gaps on the location of laminar–turbulent transition are reported. The tests were conducted in the Cryogenic Ludwieg–Tube Göttingen, a blow-down wind tunnel with good flow quality, at eight different unit Reynolds numbers ranging from Re_1 = {17.5,,times 10^{6},mathrm{{m}^{-1}}} to 80,times,10^{6},mathrm{m}^{-1}, three Mach numbers, M= 0.35, 0.50 and 0.65, and various pressure gradients. A flat-plate configuration, the extended two-dimensional wind tunnel model PaLASTra was modified in order to allow the installation of gaps with nominal widths of 30 upmu m, 100 upmu m and 200 upmu m and a depth of d = {9,mathrm{mm}}. A maximum Reynolds number based on the gap width Re_{w} = Re_1 cdot w approx {16{,}000} was reached. Transition Reynolds numbers ranging from Re_{tr}approx 1 × 106 to 11 × 106 were measured, as a function of gap width, pressure gradient and Mach and Reynolds number. This systematic investigation facilitates a linear approximation of Re_{tr} dependent on the boundary layer shape factor H_{12} for various flow conditions and gap widths. It was therefore possible to conduct an investigation of Re_{tr} depending on Re_{1} and the relative change of the transition location depending on the gap width w. Incompressible linear stability analysis was used to calculate amplification rates of Tollmien–Schlichting waves and determine critical N-factors by correlation with measured transition locations. The change in the critical N-factor varDelta N by installation of the gap is investigated as a function of w and Re_w. It was found that a gap width of 30 upmu m reduces the critical N-factors in the range of varDelta N approx 0.5 pm 0.25, while gap widths of 100 upmu m and 200 upmu m reduce the critical N-factor in the range of varDelta N approx 1.5 pm 1. Interestingly, an increase in gap width from 100 to 200 upmu m was not found to induce smaller transition Reynolds numbers or reduced N-factors, which might be due to resonance effects.

Mass-Flow Simulation in Air-Intake Models at Subsonic Mach Numbers Using Diffuser of Wind Tunnel As Passive Ejector

In order to ensure adequate mass-flow rate at low Mach numbers through air-intake models in a blow-down wind tunnel, the diffuser of the wind tunnel was used as a passive ejector. The ejector action was accomplished by interconnecting the exit plane of the intake model to a low-pressure region downstream of the choked second throat of the tunnel. Due to the reduced back-pressure, there was a significant acceleration of the entry flow, as indicated by nearly doubling of the inlet velocity. Measurements indicate that for an aircraft air-intake model, as well as a submerged air-intake model, the maximum mass flow rate achievable through the control plug was accomplished even at the lowest Mach number of 0.30, obviating the need for using an active ejector system.

Control of Back-pressure During Wind Tunnel Testing of A Two Dimensional Ramjet Air-intake Model

For optimum utilization of available run-time in blowdown wind tunnels, the controls, instrumentation, data acquisition and various other related systems associated with measurements must be efficiently synchronized. Testing of the air-intake model is a typical example where large number of steady and unsteady pressure measurements need to be combined with control of back pressure at the exit plane of the intake. This paper presents the development of a PC-based control system for testing a ramjet air-intake model at supersonic Mach numbers in the 0.3m trisonic wind tunnel at CSIR-National Aerospace Laboratories, Bangalore. The back pressure of the two-dimensional ramjet air intake model was controlled using a rectangular butterfly valve driven by a stepper motor external to the test section. Commands to vary the valve position were programmed in a PC and the valve position acquired from a potentiometer. The motor control and pressure data acquisition were operated in a hand-shake mode to enable hands-off execution of tests and maximize the data throughput from each blowdown.

Mode Switching in a Compressible Rectangular Cavity Flow

This study determines the mean and fluctuating pressures for flow through a rectangular shallow cavity (ratio between the length and the depth = 2.43, 4.43, and 6.14; ratio between the length and the width = 0.5, 1.0, and 2.0) at a Mach number of 0.64 in a blowdown transonic wind tunnel. A amplitude modulation analysis is used for the post-processing of the fluctuating pressure signals. The spectral analysis (wavelet) shows the intermittent behavior of the discrete Rossiter–Heller modes. A correlation analysis determines that mode switching is more significant between the second and third modes (organized structures that are associated with shear layer vortices), particularly for two-dimensional and shallower cavities.

Integrated flutter test of a horizontal-tail model in the large-scale continuous transonic wind tunnel

Wind tunnel test is a critical way to investigate flutter characteristics of aircraft, particularly in identifying the transonic dip phenomenon in the aeroelastic design. The continuous wind tunnel has advantages over the blow-down wind tunnel in long running time, wide bandwidth of dynamic pressure, and stable flow field. Thus, the continuous wind tunnel is more suitable for transonic flutter tests. Flutter test techniques are studied for FL-62 wind tunnel, the first large-scale continuous transonic wind tunnel in China. A protection system and a real-time signal processing are developed to ensure the safety and efficiency. Flutter boundaries of a horizontal-tail model are investigated in FL-62 wind tunnel in the Mach number regime between 0.65 and 0.98. In the test, structural vibrations are measured via strain gauges, deformations and pressure distributions are measured via non-contact optical techniques, and flutter boundaries are predicted via a modal identification method.

Optimized Film Cooling Flow on a Contoured Endwall Within a Transonic Annular Cascade

AbstractFilm cooling was measured on the endwall of a five-vane annular cascade in a blowdown wind tunnel at an exit Mach number of 0.9. The adiabatic film cooling effectiveness was calculated from the partial pressure of oxygen measured with binary pressure-sensitive paint (BPSP). Cylindrical film cooling holes were located in the upstream and passage regions with the coolant-to-mainstream mass flow ratio (MFR) independently varied for each region. One row was located upstream of the vanes and supplied by an upstream plenum. Two rows were located in the passage between two vanes and supplied by a downstream plenum. Three total MFRs were investigated: 1%, 1.5%, and 2%. For a given total MFR, four combinations of upstream and downstream MFRs were compared to an even split of coolant. Coolant-to-mainstream density ratios (DRs) of 1.0 and 2.0 were investigated. The most efficient use of coolant hinged on balancing the downstream MFR for the second row due to the endwall pressure gradient preventing coolant from exiting the holes or a high jet velocity causing liftoff. For this row, selecting the optimum MFR increased the area-averaged film cooling effectiveness by up to 200% with a reduction in row 1 of less than 25%. At high downstream MFRs, increasing the density ratio delayed liftoff and increased film cooling effectiveness in row 2 by 65%. However, at low MFRs, increasing the density ratio reduced film cooling effectiveness in row 2 by 60%.

High pitch sweep rates in supersonic wind tunnel tests

A typical test in a high-Reynolds-number blowdown wind tunnel, such as the VTI T-38 facility in Belgrade, involves a model sweep in a range of angles of attack. Increasing pitch sweep rates can bring two-fold benefit. The angle-of-attack range can be extended in tests at high stagnation pressures with limited run time, or test time can be shortened for the desired angle-of-attack range in other tests, in both cases bringing significant energy savings. The potential effect of the increased pitch rates on measurement data in the VTI T-38 wind tunnel was investigated in a recent supersonic wind tunnel test campaign of a hypervelocity standard model. The Mach 2 tests were performed at pitch rates of 2, 3, 4, 8 and 12° per second. While the wind tunnel balance measurement data were not seemed to be affected, the base pressure data revealed a high dependency on changing pitch sweep rates. It was found that the high lag in pressure piping affected the measured base pressure and, consequently, the forebody axial force, which is an important fact to know when dealing with ‘base drag corrected’ supersonic data. The conclusions derived from the Mach 2 tests have been verified in the additional Mach 4 tests. The verification tests confirmed that, for the currently employed base pressure measurement system, reliable measurement results can be obtained in wind tunnel tests by applying pitch rates up to 3° per second. The recommendations for modifications of the base pressure measurement system have been given.

Upstream and Passage Endwall Film Cooling of a Transonic Turbine Vane

Adiabatic film cooling effectiveness was measured on the endwall of a turbine vane under transonic flow conditions using binary pressure-sensitive paint. The combined effect of upstream and passage film cooling holes was evaluated. The experiments used a five-vane annular cascade in a blowdown wind tunnel. The mainstream velocity was set to exit isentropic Mach numbers of 0.7 and 0.9. The total coolant-to-mainstream mass flow ratio (MFR) was varied from 0.75 to 2.50%. Coolant-to-mainstream density ratios (DRs) of 1.0 and 2.0 were investigated. The first row was located upstream of the vanes, and the second and third rows were located in the passage between the vanes. Coolant was supplied by a divided plenum, allowing independent control of the MFR to the upstream and the passage holes. The coolant MFR was evenly split between the upstream and passage holes. Generally, increasing the MFR, the DR, and the Mach number increased the film cooling effectiveness. However, increasing the DR at MFRs of 0.625% and below decreased the laterally averaged film cooling effectiveness by as much as 60%, largely because mainstream flow was ingested through the second row. Upstream coolant provided some protection through the passage, with effectiveness as high as 0.2 at the second row.

Effects of Upstream Step Geometries on Endwall Film Cooling and Phantom Cooling Performances of a Transonic Turbine Vane

Abstract In gas turbine engines, the endwall misalignment (step geometries with various step heights) commonly exists between the combustor exit and the first nozzle guide vane endwall, due to assembly error, metal corrosion, and thermal expansion. The step geometries affect the secondary flow field near the endwall and coolant flow behavior, thereby leading to a significant variation in endwall film cooling, vane surface phantom cooling, and vane passage aerodynamic performances. This paper presents a detailed experimental and numerical study on endwall film cooling and vane pressure side surface phantom cooling at simulated gas turbine operating conditions (high inlet freestream turbulence level of 16%, exit Mach number of 0.85, and exit Reynolds number of 1.7 × 106). The experimental measurements were conducted at Virginia Tech’s transonic blowdown wind tunnel for two configurations: baseline cascade (HR = 0) and forward-facing step geometry (HR = −0.1). The numerical predictions were performed by solving the steady-state Reynolds-averaged Navier Stokes (RANS) with realizable k–ɛ turbulence model. Based on a double coolant temperature model, a novel numerical method for the predictions of adiabatic wall film cooling effectiveness was proposed. The qualitative and quantitative comparisons between the numerical prediction results and experimental data are provided in this paper. The results indicate that this method can reliably predict endwall film cooling distributions and vane pressure side surface phantom cooling distributions, and significantly reduce (more than 50%) numerical prediction errors. The effects of upstream step geometries were numerically studied at the design flow conditions (BR = 2.5, DR = 1.2), by quantizing the endwall film cooling effectiveness, endwall net heat flux reduction (NHFR), vane pressure side surface phantom cooling effectiveness, and total pressure loss coefficients (TPLC) for various upstream step heights: three forward-facing step heights (HR = −0.16, −0.1, and −0.06), a baseline cascade (HR = 0), and four backward-facing step heights (HR = 0.06, 0.1, 0.136, and 0.2). Results show the upstream forward-facing step geometry is beneficial for enhancing endwall film cooling (maximum enhancement level of 15%), endwall net heat flux reduction (maximum enhancement level of 10%), and vane pressure side surface phantom cooling (maximum enhancement level of 66.7%). Nevertheless, the upstream backward-facing step geometry is pernicious to the endwall film cooling, endwall heat flux reduction, and phantom cooling, and these pernicious effects have sharply deteriorated when the upstream step height is larger than a critical value (HR > 0.136). To be specific, the film cooling effectiveness (from 30% to 80%), net heat flux reduction (up to 78%), and phantom cooling effectiveness (completely disappeared) significantly decrease by introducing an overlarge backward-facing step height HR = 0.2. The upstream step geometries lead to deteriorated aerodynamic performance, though the reduction level may be slight (0.36% at HR = 0.136). These suggest that the overlarge backward-facing misalignment will increase the risk of thermal failure and must be avoided during the component assembly and the operational processes of a gas turbine.

Shock loss measurements in non-ideal supersonic flows of organic vapors

This paper presents the first ever direct measurements of total pressure losses across shocks in supersonic flows of organic vapors in non-ideal conditions, so in the thermodynamic region close to the liquid–vapor saturation curve and the critical point where the ideal gas law is not applicable. Experiments were carried out with fluid siloxane MM (hexamethyldisiloxane, C_6H_{18}OSi_2), commonly employed in medium-/high-temperature organic Rankine cycles (ORCs), in the Test Rig for Organic VApors (TROVA), a blowdown wind tunnel at the Laboratory of Compressible fluid dynamics for Renewable Energy Applications (CREA lab) of Politecnico di Milano. A total pressure probe was inserted in superheated MM vapor flow at Mach number sim 1.5 with total conditions in the range 215-230~^circ hbox {C} and 2-12~hbox {bar} at varying levels of non-ideality, with a compressibility factor evaluated at total conditions between Z_text {T}=0.68-0.98. These operating conditions are representative of the first-stage stator of ORC turbines. Measured shock losses were compared with those calculated from pre-shock quantities by solving conservation equations across a normal shock, with differences always below 2 % attesting a satisfactory reliability of the implemented experimental procedure. An in-depth analysis was then carried out, highlighting the direct effects of non-ideality on shock intensity. Even at the mildly non-ideal conditions with Z_text {T}gtrsim 0.70 considered here, non-ideality was responsible for a significantly stronger shock compared to the ideal gas limit at same pre-shock Mach number, with differences as large as 6%.Graphical abstract

Numerical and Experimental Validation of a Supersonic Mixing Layer Facility

The design of a supersonic-supersonic mixing layer facility was motivated by the need for a benchmark experimental platform to study the physical phenomena underlying supersonic mixing layers. The facility is an intermittent blowdown wind tunnel characterized by a two-stream design separated by a splitter plate in the middle of the nozzle. The splitter plate ends exactly at the start of the mixing layer test section. The Mach number of the primary stream is M1 = 3 for all nozzles and the secondary streams are M2 = 2, 2.5, and 2.9 to generate different convective Mach numbers of Mc = 0.25, 0.10, and 0.01, respectively. The facility was calibrated by pressure measurements to verify the Mach number and the pressure distribution in the streamwise direction. Large-eddy simulation (LES) was performed to illustrate a full view of the turbulent compressible mixing layer flow field and to compare this against the experimental data. Optical diagnosis measurements, i.e., femtosecond laser-induced electronic excitation tagging velocimetry (FLEET) for velocity measurement and focused laser differential interferometer (FLDI) for the density fluctuation, were also performed on the facility.

Flutter test of rudder with real servo actuator in continuous transonic wind tunnel

Flutter wind tunnel test is an important approach of investigation of transonic flutter characteristics of flight vehicle. Comparing with the blow-down wind tunnel, the long-running and low dynamic pressure capabilities of continuous transonic wind tunnel are very suitable for flutter test. The flutter safety protection and analysis of dynamic signals are developed. The safety protection control software, rapid reduction of Mach number and dynamic pressure, model debris catch screen are integrated, which can provide safety protection of test model and wind tunnel. During the test process, the flutter boundary can be achieved by interpolating the reciprocal of spectrum peak. The flutter tests of rudder are conduct through two methods of step and continuous varying dynamic pressure. It is illustrated that the error of flutter dynamic pressure is relatively small, less than 5% between the two methods. Meanwhile, the feedback effect of the real servo actuator on the flutter characteristics is hard to be obtained via numerical simulation. It is demonstrated that the flutter dynamic pressure has been increased by 10% due to the feedback effect.

Flow in Pillow-Plate Channels for High-Speed Turbomachinery Heat Exchangers

In numerous turbomachinery applications, e.g., in aero-engines with regenerators for improving specific fuel consumption (SFC), heat exchangers with low-pressure loss are required. Pil low-plate heat exchangers (PPHE) are a novel exchanger type and promising candidates for high-speed flow applications due to their smooth profiles avoiding blunt obstacles in the flow path. This work deals with the overall system behavior and gas dynamics of pillow-plate channels. A pillow-plate channel was placed in the test section of a blow-down wind tunnel working with dry air, and compressible flow phenomena were investigated utilizing conventional and focusing schlieren optics; furthermore, static and total pressure measurements were performed. The experiments supported the assumption that the system behavior can be described through a Fanno–Rayleigh flow model. Since only wavy walls with smooth profiles were involved, linearized gas dynamics was able to cover important flow features within the channel. The effects of the wavy wall structures on pressure drop and Mach number distribution within the flow path were investigated, and a good qualitative agreement with theoretical and numerical predictions was found. The present analysis demonstrates that pressure losses in pillow-plate heat exchangers are rather low, although their strong turbulent mixing enables high convective heat transfer coefficients.

Pneumatic system for pressure probe measurements in transient flows of non-ideal vapors subject to line condensation

This paper presents the design, construction and commissioning of a pneumatic system for pressure probe measurements in flows of organic vapors in non-ideal conditions, namely in the thermodynamic region close to the liquid–vapor saturation curve and the critical point where the ideal gas law is not applicable.Experiments were carried out with fluid siloxane MM (hexamethyldisiloxane, C6H18OSi2), commonly employed in medium/high temperature Organic Rankine Cycles (ORCs), in the Test Rig for Organic VApors (TROVA), a blow-down wind tunnel at the Laboratory of Compressible fluid dynamics for Renewable Energy Applications (CREA lab) of Politecnico di Milano. TROVA operation is intrinsically transient due to its batch nature, with a low frequency content (∼1Hz) related to the emptying of the high-pressure reservoir feeding the test section.The challenges linked to possible condensation in pneumatic lines, such as vapor–liquid menisci, hydrostatic head and mass-sink effects, were evaluated by means of theoretical calculation and experiments. To avoid these issues, a nitrogen-flushed pneumatic system for absolute and differential pressure measurements was designed and successfully tested with superheated MM vapor expanding in planar choked converging nozzles characterized by a portion with constant cross-sectional area yielding design Mach numbers of 0.2, 0.5 and 0.7. Commissioning of the complete system including a probe was performed with the first ever testing of an L-shaped Pitot tube in non-ideal subsonic flows of siloxane MM vapor at Mach numbers M=0.2 and 0.5. Measurement delay issues were identified and assessed through a dynamic testing procedure, and were solved by reducing the overall pneumatic lines volume including the one hidden within pressure transducers. The correct performance of the complete system was therefore verified for probe measurements of total, static and dynamic pressure in non-ideal flows of organic vapors. This sets the foundation for future directional pressure probes calibration and use in the characterization of such flows, as in direct measurement of total pressure losses across shocks and in testing of ORC turbine blade cascades.

Attenuation characteristics of Karman vortex passing through the wire mesh

In this study, the attenuation and frequency characteristics of the Karman vortex passing through the wire mesh was investigated. This experiment was conducted by generating the Karman vortex from a cylinder in an acrylic duct of the blowdown type wind tunnel and passing through the wire mesh at the duct exit. The cylinder with a diameter of 6mm was used to generate the Karman vortex, and the Reynolds number was 3000 based on the flow velocity toward the cylinder and the cylinder diameter. The wake of the Karman vortex passing through the wire mesh was measured by hot-wire anemometer, and was visualized by smoke wire method. As a result, it was found that the smaller the wire diameter and open area ratio of the wire mesh, the velocity and turbulence intensity distributions ware more uniform. And, the dominant frequency of the Karman vortex was confirmed that even in the wire mesh that showed uniformity and in the wire mesh that appeared to have collapsed in the photograph by frequency analysis.

Modal Analysis of Cylinder-Induced Transitional Shock-Wave/Boundary-Layer Interaction Unsteadiness

Shock-wave/boundary-layer interactions (SWBLIs) are critical phenomena in the design of high-speed vehicles as they exhibit inherent unsteadiness that can damage airframes and lead to engine unstart. This paper presents a novel characterization of the unsteady dynamics of cylinder-induced SWBLIs at Reynolds numbers of and using proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD). Data sets analyzed in this paper were obtained from experiments conducted in the NASA Langley Research Center’s 20-Inch Mach 6 blowdown wind tunnel. The experiments used a cylinder protuberance model mounted on a 10° half-angle wedge to create the interaction. The lower-Reynolds-number case was observed to have a transitional incoming boundary layer, whereas a swept ramp array trip was used in the higher-Reynolds-number case to generate a turbulent incoming boundary layer. The POD analysis from both cases helped isolate spatial regions in the separation bubble that contained the highest percentage of energy for each case, which correlate to the dominant structures in the flow. A power spectral density analysis on the POD temporal components of the highest-energy POD mode (mode 1) revealed a frequency spectrum with a distinct peak at in the lower-Reynolds-number case and a band of high-energy peaks in the range of 0.05–0.2 for the tripped SWBLI case. Through the DMD analysis, an unsteadiness mode was isolated at and an asymmetric motion of the forward lambda shock was identified in the lower-Reynolds-number case. This dynamic asymmetry was not detected in the higher-Reynolds-number case, suggesting that the asymmetric effect may be unique to interactions in transitional incoming boundary layers. Additional spectral-kernel-based POD analysis identified primary unsteadiness frequencies that were in agreement with the DMD findings. A low-frequency mode at , exhibiting relatively high energy, was obtained for the lower-Reynolds-number (transitional) case, whereas a broadband “bump” with increased energy over a frequency range of was observed for the higher-Reynolds-number (tripped) case. Through these analysis techniques, a mode, likely connected to a shock-breathing mechanism, was identified at . In the assessment of the structural motion occurring in the shock-breathing mode, the phenomenon in the lower-Reynolds-number case was found to traverse a longer upstream distance and compress to a larger degree when compared with the breathing behavior in the higher-Reynolds-number case.

Endwall Heat Transfer and Cooling Performance of a Transonic Turbine Vane With Upstream Injection Flow

Abstract Flow fields near the turbine vane endwall region are very complicated due to the presence of high three-dimensional passage vortices and endwall secondary flows. This makes it challenging for the endwall to be effectively cooled by using traditional endwall cooling methods, such as impingement cooling combined with local film cooling inside the vane passage. One effective endwall cooling scheme: coolant injection flow through discrete holes upstream of the vane leading edge on the endwall, has been considered by many gas turbine companies. The present paper focuses on endwall film cooling effectiveness evaluation with upstream coolant injection through discrete holes. Detailed experimental and numerical studies on endwall heat transfer and cooling performance with coolant injection flow through upstream discrete holes are presented in this paper. High-resolution heat transfer coefficient (HTC) and adiabatic film cooling effectiveness values were measured using a transient infrared thermography technique on an axisymmetric contoured endwall. The endwall tested was a scaled up inner endwall of an industrial transonic turbine vane with double-row discrete cylindrical film cooling holes located 0.39Cx upstream of the vane leading edge. The tests were performed in a transonic linear cascade blowdown wind tunnel facility. Conditions were representative of a land-based power generation turbine with exit Mach number of 0.85 corresponding to exit Reynolds number of 1.5 × 106, based on exit condition and axial chord length. A high turbulence level of 16% with an integral length scale of 3.6%P was generated using inlet turbulence grid to reproduce the typical turbulence conditions in real turbine. Low temperature air was used to simulate the typical coolant-to-mainstream condition by controlling two parameters of the upstream coolant injection flow: mass flowrate to determine the coolant-to-mainstream blowing ratio (BR = 2.5, 3.5), and gas temperature to determine the density ratio (DR = 1.2). To highlight the interactions between the upstream coolant flow and the passage secondary flow combined with the influence on the endwall heat transfer and cooling performance, a comparison of CFD predictions with experimental results was performed by solving steady-state Reynolds-averaged Navier–Stokes (RANS) using the commercial CFD solver ansys fluent V.15. A detailed numerical method validation was performed for four different Reynolds-averaged turbulence models. The realizable κ−ε model was validated to be suitable to obtain reliable numerical solution. The influences of a wide range of coolant-to-mainstream blowing ratios (BR = 1.0, 1.5, 1.9, 2.5, 3.0, 3.5) were numerically studied. Complex interactions between coolant injections and secondary flows in vane passage were presented and discussed. An optimal value of the blowing ratio for the present upstream discrete film hole is also suggested based on the current study. Results indicate that for lower values of BR, the endwall coolant coverage from the upstream double-row discrete holes is strongly controlled by the passage secondary flow, thus the cooling effectiveness is very poor. As the BR increases, the strong secondary flow in vane passage can be suppressed by the coolant injections and begin to be almost eliminated when BR increases to a critical value (BR = 2.5–3.0). Beyond the critical BR, most of the injected coolant begins to lift off from the endwall and penetrate significantly into the mainstream flow, yielding inefficient endwall cooling performance.

Experimental Characterization of Three-Dimensional Scramjet Inlet with Variable Internal Contraction

A compact three-dimensional scramjet intake with a high compression ratio and a movable cowl as a starting device and for the adaption of the internal contraction to respective operating conditions was designed for a parabolic reentry flight experiment. In an extensive measurement campaign, conducted in a blowdown wind tunnel at Mach 7, the operational behavior of this inlet was characterized. The parameters investigated include internal contraction ratio, Reynolds number, angle of attack, and angle of yaw, and the starting behavior was examined. Static and total pressure ratios and mass capture ratios were used to evaluate the performance of the inlet and throttle curves to determine the operational limits. The flowfield was studied by schlieren images, wall pressure distributions, and Mach number profiles in the interface of the isolator and combustion chamber, which were determined with a pressure measurement rake with static and pitot pressure probes. The flow on the external ramp was analyzed by infrared thermography. The inlet showed good performance regarding compression and efficiency and proved to be quite robust at various flight angles. The performance tended to improve, and the inlet operated more stably with increasing internal contraction, with the most important factor being whether the inlet leading-edge shock was captured by the cowl.