Transonic Turbine Blade Research Articles

Shock wave structures and vortex unsteadiness in the tip region of a transonic turbine cascade under different conditions

The tip region of transonic turbine blades (exit Mach number of 0.95) exhibits complex flow characteristics with the coexistence of multiple shock wave systems and multiscale vortices. Based on a validated detached Eddy simulation method, unsteady flow features such as the spatial-temporal dynamic evolution of tip leakage vortices (TLV) and the periodic buffet/oscillation of shock trains inside the tip gap are revealed and discussed systematically under different incidence angles (i) and heights of tip clearance. Moreover, the distinction of tip flow structures between the subsonic and transonic conditions is also manifested. Results indicate that the wandering behavior of the TLV is influenced by both the swirling strength of the vortex itself and the interaction of adjacent secondary vortices. The TLV under a tip gap of 5% blade height (h) exhibits “binary and bimodal” wandering characteristics both along the pitchwise and spanwise direction, whereas under the 1%h case, only the pitchwise wandering is prominent. The dominant characteristic frequency of the TLV wandering under different conditions falls within the spectrum range of 2.2–2.4 kHz. As for the shock trains inside the tip clearance (τ), coherently movement back and forth along the pitchwise direction with varying amplitudes can be observed, where the magnitude within the τ=1%h exceeds that observed in the τ=5%h, depending on the intensity of the shock waves. Notably, significant shock wave oscillations are present throughout the range of the chord length (c) within the τ=1%h, whereas within the tip gap of 5%h, shock wave systems exhibit more pronounced oscillations predominantly near the trailing edge.

Tip gap effects on film cooling performance of a transonic turbine blade with a distinctive coolant supply arrangement

Considered are the influences of tip gap on film cooling performance variations near to and along the tip or end of a turbine blade with a squealer rim contained within a linear cascade with five airfoils. The film coolant is supplied by two plenums which are positioned at forward and aft locations within the blade. The aft blowing ratio varies between 0.88 and 1.48, the forward blowing ratio is set to be approximately constant in the vicinity of 2.75 to 3.0, and a value of about 1.3 is utilized for the ratio of coolant density relative to the main flow density. Measured results include linearly-averaged and spatially-resolved variations of adiabatic effectiveness, which are presented both for the top portion of the pressure surface and end surface of the blade with the squealer rim. Local effectiveness variations along the top pressure side of the blade show significant effectiveness values downstream of aft pressure side hole outlet locations mostly because film lift-off downstream of these holes is less substantial than observed downstream of the forward holes. Spatially-resolved effectiveness data for the end of the blade with a squealer rim show that the largest values of effectiveness are evident near to and immediately downstream of dusting hole outlet locations. Associated linearly-averaged effectiveness data for the tip of the blade recess region with the 1.2 mm tip gap are as much as about two times higher than values for the 2.0 mm tip gap when compared at each surface location. Local and linearly-averaged surface effectiveness variations for the downstream edge region of the blade squealer tip are most strongly influenced by the film cooling holes which are aft and by the dusting hole which is positioned along the tip of the blade rim near the downstream edge. Here, effectiveness magnitudes are also often higher along the downstream edge rim region with the 1.2 mm tip gap, compared to data associated with the 2.0 mm tip gap.

Interaction Mechanism of Transonic Squealer Tip Cooling With the Effect of High-Speed Relative Casing Motion

Abstract The relative casing motion can significantly influence the turbine blade tip aerothermal performance. In this study, experimental investigation was conducted in a newly developed high-speed disk rotor rig which can mimic engine realistic high-speed casing relative motion while enabling full optical access to a transonic turbine blade tip surface. Spatially-resolved tip heat transfer data, including heat transfer coefficient and film cooling effectiveness, were obtained for a cooled transonic squealer tip by infrared transient thermal measurement. Combined with closely coupled Reynolds-averaged Navier–Stokes computational fluid dynamics (CFD) analysis, this paper reveals an interesting interaction mechanism between the cooling injections from the pressure side and the cavity floor with and without the effect of relative casing motion. Both experimental data and CFD results show a consistent trend in both heat transfer and cooling performance. With cavity cooling only, the cooling performance reduces with the effect of relative casing motion. However, with additional cooling injection from the pressure side, a significant improvement in the combined cooling performance with the relative casing motion can be observed. Such opposite trend highlights the importance of relative casing motion when ranking different tip cooling designs. With the consideration of relative casing motion, extra tip cooling benefit can be obtained by combining cooling injections from two different locations.

Effects of pressure side film cooling hole placement and condition on surface heat transfer coefficients along a transonic turbine blade tip

Considered is thermal performance for the squealer blade tip of four film cooling configurations, B1, B2, B3 and B4, with hole exits for each arrangement positioned along a different portion of the upper pressure side of a transonic turbine blade. Each upper pressure side configuration includes five film cooling holes, such that each hole has a compound angle of 45° with respect to the circumferential/axial plane, also with an angle orientation of 40° relative to the plane tangent to the exit cross sectional area of each hole. Results are given for blowing ratios ranging from 0.42 to 3.20 in the form of spatially-resolved and spatially-averaged surface distributions of heat transfer coefficients and heat transfer coefficient ratios. In regard to tip surface variations, surface heat transfer coefficient ratios distributions are a consequence of the manner in which the film coolant collects in a substantial manner along the pressure side rim and then advects above and across the squealer recess with little to no coolant collecting within the cavity, or along the recess surface, to then collect in a substantial manner further downstream along the suction side rim. Such flow characteristics affect heat transfer coefficient ratio variations for all four film cooling configurations, but are especially influential for the B3 arrangement. Considering local heat transfer coefficient ratio distributions along the upper pressure side surface, the most significant local variations in the vicinity of a film cooling hole exit locations. Here, heat transfer coefficient ratio distributions for the B1 and the B4 configurations, in particular, show evidence of a horseshoe-shaped vortex which forms around each emerging coolant concentration. The two downstream legs of each vortex are associated with a pair of locally augmented heat transfer coefficient streaks, often with a streak of locally-lower coefficient ratios positioned between. Local and line-averaged heat transfer coefficient ratios along the upper pressure side also vary significantly with blowing ratio, with additional periodic variations as the normalized circumferential/axial coordinate varies. Within the resulting distributions, local heat transfer coefficient ratio increases are associated with locally augmented mixing and turbulent transport, which are especially present near the exits of different film cooling hole locations.

The Influence of Boundary Layer State and Trailing Edge Wedge Angle on the Aerodynamic Performance of Transonic Turbine Blades

Abstract Trailing edge loss typically accounts for a third of the profile loss of a transonic turbine blade. In this paper, two dominant influencing factors—boundary layer state and trailing edge wedge angle—are quantified using a combined experimental (single airfoil transonic tests) and computational (wall-resolved LES) approach. Three key findings are reported. First, the switch between two flow regimes—detached vortex shedding and transonic vortex shedding—as the Reynolds number is increased, is identified and explained. It is shown that the transition of the pressure surface boundary layer drives the switch from the former to the latter; this change is associated with a jump in loss of close to 100%. The switch is also shown to affect the base pressure distribution around the trailing edge, changing it from uniform, with an isobaric dead-air region, to non-uniform. Second, the influence of the trailing edge wedge angle is explored. It is found that higher wedge angles can offer performance benefits in both loss and base pressure—with a change from 8 deg to 14 deg geometry providing up to a 29% reduction in an overall loss for the geometry tested. Finally, as these findings are the result of changes to the unsteady flow phenomena around the trailing edge, they would not be picked up by current low-fidelity, steady design tools. A method is proposed to use data, such as that presented in this paper, to improve low-fidelity modeling.

Dusting Hole Film Cooling Heat Transfer on a Transonic Turbine Blade Tip

Investigated is a transonic turbine blade tip with a squealer rim and a squealer recess, with a single dusting film cooling hole contained within the leading edge region of the squealer recess. Data are provided for transonic flow conditions for a range of film cooling blowing ratios for two tip gap values, using a linear cascade, with no relative motion between the blade and the casing. Surface heat transfer characteristics are measured using the transient impulse-response measurement approach, employed with infrared thermography. Line-averaged adiabatic film cooling effectiveness values, for the 1.4 mm tip gap, are generally very small along the pressure side rim, with only small, locally increased values along the suction side rim. For the 0.8 mm tip gap, line-averaged adiabatic film cooling effectiveness values are generally somewhat higher along the pressure side rim and along the suction side rim. In general, effectiveness values for both tip gap values, for these locations, and for the recess region, increase as the blowing ratio increases. As the tip gap decreases from 1.4 mm to 0.8 mm, line-averaged adiabatic film cooling effectiveness generally increases on the rims and downstream regions of the recess, with increased magnitudes which are spread over larger spatial surface areas. For tip gaps of 0.8 mm and 1.4 mm, for regions where the line-averaged heat transfer coefficient ratio deviates significantly from 1.00, values generally decrease as the blowing ratio increases. Across every region of the blade, line-averaged heat transfer coefficient ratios either decrease or remain approximately invariant, as the tip gap value decreases from 1.4 mm to 0.8 mm.

Effects of pressure side film cooling hole placement and condition on adiabatic film cooling effectiveness characteristics of a transonic turbine blade tip

The effects of film cooling hole placement location along the upper pressure side of a transonic squealer are considered. The thermal performance of four different film cooling configurations; B1, B2, B3 and B4, are considered using the University of Alabama in Huntsville's SS/TS/WT (supersonic/transonic/wind tunnel) experimental facility and a simulated turbine blade row using a linear cascade. Surface-varying results are provided for both the squealer blade tip surface, and for the upper pressure side of the squealer blade. These results are given for blowing ratios ranging from 0.42 to 3.20 in the form of spatially-resolved and spatially-averaged adiabatic film cooling effectiveness distributions. Because local static pressure variations and gradients vary in a significant manner at different blade locations, coolant accumulations near surfaces (and the associated surface thermal protection and film effectiveness distributions) produced by the different film cooling configurations are vastly different along the squealer blade tip surface. The B1 and B3 film cooling configurations show that substantial magnitudes of adiabatic film cooling effectiveness are present along the pressure side rim and the suction side rim. Also present is a considerable region of locally increased film cooling effectiveness and a wider region of surface protection within the squealer recess region for the B2 film cooling configuration. The contributions are unique because of a deficit of experimental data for film cooled squealer blades which operate with transonic flow, and because the present investigation is the first to consider the effects of hole placement location for upper pressure side film cooling arrangements with transonic flow conditions.

Surface film cooling characteristics of rectangular holes along the trailing edge of a transonic turbine blade tip

The present study investigates film cooling in the trailing edge region of a transonic turbine blade tip, with an arrangement which is significantly different from all previous blade tip investigations. In particular, five rectangular film cooling holes, located along the tip near the blade trailing edge and positioned downstream of a squealer recess region, are employed. Each film cooling hole is oriented at a 45° angle relative to the blade tip surface. Heat transfer coefficient ratio data, and adiabatic film cooling effectiveness data are provided for two tip gaps over a range of blowing ratios. Tip gap magnitudes are 1.4 mm and 0.8 mm, which correspond to 1.16% and 0.66% of the true blade span, respectively. Surface heat transfer characteristics are measured using a transient impulse-response approach with infrared thermography. Along most of the trailing edge tip surface, adiabatic film cooling effectiveness values are locally augmented for all blowing ratios BR and for both tip gaps considered, which evidences substantial film coolant concentrations near the blade tip surface. Local values of adiabatic film cooling effectiveness often increase with increasing blowing ratio and are consistently higher for the smaller tip gap. In general, only minor differences between the 0.8 mm and 1.4 mm tip gap heat transfer coefficient ratio data sets are present, with marginally higher values of heat transfer coefficient ratios for the smaller tip gap. Together, heat transfer coefficient ratio and adiabatic film cooling effectiveness data suggest that the thermal protection offered by the 0.8 mm tip gap along the trailing edge tip region is superior to that of the 1.4 mm tip gap for a range of blowing ratios.

Measurement and determination of local film cooling performance along a transonic turbine blade tip with viscous dissipation

Presented are experimental and analytic procedures for the measurement and determination of film cooling performance parameters for a transonic flow environment along a turbine blade tip, which account for the influences of viscous dissipation. Such viscous dissipation magnitudes are vital to ascertain appropriate driving temperatures for convective heat transfer within high velocity, compressible environments. A key step in this procedure is the separation of adiabatic surface temperature magnitudes due to the thermal field from adiabatic surface temperature values associated with flow effects related to viscous dissipation. These latter adiabatic surface temperature magnitudes, from flow effects only, are determined with no film cooling, which, when considered relative to flow stagnation temperature, are directly related to local Mach number values within the tip gap flow along the blade tip surface. After the separation is implemented, adiabatic surface temperature variations from film cooling thermal effects only are provided relative to local baseline values with no film cooling. Resulting local adiabatic film cooling effectiveness and local heat transfer coefficients are subsequently determined which are based entirely upon locally measured temperatures (without the use of global temperature parameters), which is the most appropriate perspective for presentation of spatially-resolved surface heat transfer characteristics. Associated data are provided for the surface of a blade tip with a squealer rim and recess, with and without film cooling. An array of five film cooling holes, located along the upper pressure side of a two-dimensional turbine blade, are employed to provide the film coolant, with a local blowing ratio BR of 3.02. Ratios of tip gap to true blade span are 0.66%, and 1.16%. Experimental procedures employed to obtain the spatially-resolved surface data include infrared thermography, transient testing techniques, and the impulse response method for transient data analysis.

Measurement and determination of local film cooling performance along a transonic turbine blade tip with viscous dissipation

Abstract Presented are experimental and analytic procedures for the measurement and determination of film cooling performance parameters for a transonic flow environment along a turbine blade tip, which account for the influences of viscous dissipation. Such viscous dissipation magnitudes are vital to ascertain appropriate driving temperatures for convective heat transfer within high velocity, compressible environments. A key step in this procedure is the separation of adiabatic surface temperature magnitudes due to the thermal field from adiabatic surface temperature values associated with flow effects related to viscous dissipation. These latter adiabatic surface temperature magnitudes, from flow effects only, are determined with no film cooling, which, when considered relative to flow stagnation temperature, are directly related to local Mach number values within the tip gap flow along the blade tip surface. After the separation is implemented, adiabatic surface temperature variations from film cooling thermal effects only are provided relative to local baseline values with no film cooling. Resulting local adiabatic film cooling effectiveness and local heat transfer coefficients are subsequently determined which are based entirely upon locally measured temperatures (without the use of global temperature parameters), which is the most appropriate perspective for presentation of spatially-resolved surface heat transfer characteristics. Associated data are provided for the surface of a blade tip with a squealer rim and recess, with and without film cooling. An array of five film cooling holes, located along the upper pressure side of a two-dimensional turbine blade, are employed to provide the film coolant, with a local blowing ratio BR of 3.02. Ratios of tip gap to true blade span are 0.66 percent, and 1.16 percent. Experimental procedures employed to obtain the spatially-resolved surface data include infrared thermography, transient testing techniques, and the impulse response method for transient data analysis.

Transonic Turbine Blade Tip Heat Transfer with Pressure Side Film Cooling

Investigated are spatially resolved distributions of surface adiabatic film cooling effectiveness and surface heat transfer coefficients for a transonic turbine blade tip, with a unique and innovative film cooling arrangement, wherein a single row of five film cooling holes, located on the pressure side of the blade very near to the blade tip, provides thermal protection to the blade tip. The blade tip contains a squealer rim, and a squealer recess region. The present tip gap magnitude is 1.2% of the true blade span. When blade upper pressure side data are considered, heat transfer coefficient ratio variations are due, in part, to a horseshoe-shaped vortex structure, which forms around each film cooling jet, and then advects downstream. Spatially resolved distributions of surface adiabatic film cooling effectiveness show that the coolant advects to the upper edge of the blade pressure surface, along the pressure side rim, over the squealer recess region, and then sometimes collects in a substantial manner along portions of the suction side rim. As a result, local effectiveness values are generally higher on the suction side rim relative to pressure side rim values when compared at a particular value of blowing ratio.

Effects of tip gap on transonic turbine blade heat transfer characteristics with pressure side film cooling

Investigated are local, spatially-resolved and line-averaged distributions of surface adiabatic film cooling effectiveness and surface heat transfer coefficient ratios for a transonic turbine blade tip. The tip contains a squealer rim, and a single row of film cooling holes is located on the pressure-side of the blade very near to the blade tip. A two-dimensional linear cascade, with four flow passages and five complete blades, is utilized for the investigation. Associated data are provided for different film cooling flow conditions for two different tip gaps, which correspond to 1.16 and 0.66% of true blade span. Experimentally-measured, spatially-resolved surface data demonstrate that improved surface thermal protection on the squealer tip surface, from pressure side film cooling, is often present with a smaller tip gap. Such a conclusion is illustrated by heat transfer coefficient values at many surface locations which are generally lower, and by adiabatic film cooling effectiveness magnitudes which are often substantially larger for the smaller tip gap, relative to values associated with the larger tip gap. Associated effectiveness values, along the upper pressure side, vary in a significant manner as blowing ratio changes. Along the tip region of the blade, for both tip gap values, line-averaged adiabatic film cooling effectiveness generally increases with increasing blowing ratio for the pressure side rim. Effectiveness data for the suction side rim with the smaller tip gap, show more complicated variations as blowing ratio changes. Heat transfer coefficient data along the squealer tip surface and on the upper pressure side of the blade surface generally have variations which are similar to the baseline blade with no film cooling for many surface locations. The local deviations, which are present, are due to locally augmented mixing and shear, and increased turbulent transport from the advective presence of the film coolant.

Large-Eddy Simulation of Variable Speed Power Turbine Cascade With Inflow Turbulence

Abstract Numerical results are presented from the National Aeronautics and Space Administration (NASA) Glenn Research Center's in-house turbomachinery code Glenn-HT applied to the variable-speed power turbine (VSPT) experiment at the NASA Transonic Turbine Blade Cascade Facility. The main goal of this paper is to implement a digital filtering method to generate turbulence upstream and a subgrid model (localized dynamic k-equation model (LDKM)) in the framework of large-eddy simulation (LES) in order to investigate the effect of inflow turbulence on the transition seen in the VSPT experimental data at the cruise condition (incidence angle of 40 deg and Tu = 0.5%, 5%, 10%, and 15%). Although the boundary layer on the suction side and pressure side of the blades is initially laminar due to favorable pressure gradient, the laminar flow can transition to turbulent flow past a separation zone on the suction side or by natural or bypass transition. This process determines the total pressure losses in the wake. Therefore, it is desirable to develop a reliable prediction tool to accurately capture the transition mechanism in blade rows operated under the conditions of low Reynolds number and at a variety of freestream turbulence conditions. Our numerical studies reveal that the location of separation is rather insensitive to the level of Tu; however, the effect of increasing Tu seems to be in reducing the size and ultimately suppressing the separation bubble. In addition, we performed spectral analysis to identify the peak frequencies in the region where the separation bubble is formed, which provides valuable insights into the transition/separation mechanism.

The effects of tip gap variation on transonic turbine blade tip leakage flow based on VLES approach

Transonic turbine blade Over-Tip Leakage (OTL) flow has significant effects on turbine efficiency and blade thermal load. At present, most of the numerical studies on OTL flow are based on Reynolds-Averaged Navier-Stokes (RANS) method. However, it is difficult to characterize the physical mechanisms of transonic OTL flow accurately due to the inherent shortcomings of RANS method, especially for the unsteady characteristics. To understand the underlying loss mechanisms and aerothermal characteristics of the transonic OTL flow with tip gap variations, the present study applies a Very Large Eddy Simulation (VLES) method based on Wilcox k–ω model to study a transonic cascade OTL flow. The feasibility and accuracy of the VLES turbulence model in OTL flow simulations are confirmed by numerical assessments comparison with the experimental results. The detailed flow structure and its effect on the losses of tip leakage flow are studied for different cases with varying tip gap heights. It is observed that the flow structures including separation vortex and shock waves in the gap become more complex with increasing the tip gap. The mass flow rate, pressure loss, entropy generation, and aerothermal performance for different tip gap heights are further discussed.

New Supersonic Loss Model for the Preliminary Design of Transonic Turbine Blades and the Influence of Pitch

Abstract In order to produce a more efficient design of a compact turbine driving a cryogenic engine turbo-pump for a satellite delivering rocket, a new supersonic loss model is proposed. The new model was constructed based on high-quality published data, composed of Schlieren photographs and experimental measurements, that combined provided a unique insight into the mechanisms driving supersonic losses. Using this as a cornerstone, model equations were formulated that predict the critical Mach number and shock loss and shock-induced mixing loss as functions of geometrical (i.e., blade outlet and uncovered turning angle and trailing edge thickness) and operational parameters (i.e., exit Mach number). A series of highly resolved CFD numerical simulations were conducted on an in-house designed state-of-the-art transonic turbine rotor row (around unity aspect ratio (AR)) to better understand changes in the shock system for varying parameters. The main outcome showed that pitch to chord ratio has a powerful impact on the shock system, and thus on the manner by which shock loss and shock-induced mixing loss is distributed to compose the overall supersonic losses. The numerical loss estimates for two pitch to chord ratios—t⁄c = 0.70 and t⁄c = 0.98—were compared with absolute loss data of a previously published similar blade with satisfactory agreement. Calibrated equations are provided to allow hands-on integration into existing overall turbine loss models, where supersonic losses play a key role, for further enhancement of preliminary turbine design.

Interaction mechanisms of shock waves with the boundary layer and wakes in a highly-loaded NGV using hybrid RANS/LES

Accurate predictions of Shock Waves and Boundary Layer Interaction (SWBLI) and strong Shock Waves and Wake Vortices Interaction (SWWVI) in a highly-loaded turbine propose challenges to the currently widely used Reynolds-Averaged Navier-Stokes (RANS) model. In this work, the SWBLI and the SWWVI in a highly-loaded Nozzle Guide Vane (NGV) are studied using a hybrid RANS/LES strategy. The Turbulence Kinetic Energy (TKE) budget and the Proper Orthogonal Decomposition (POD) method are used to analyze flow mechanisms. Results show that this hybrid RANS/LES method can obtain detailed flow structures for flow mechanisms analysis. Strong shock waves induce boundary layer separation, while the presence of a separation bubble can in turn lead to a Mach reflection phenomenon. The shock waves cause trailing-edge vortices to break clearly, and the wakes, in turn, can change the shocks intensity and direction. Furthermore, the Entropy Generation Rate (EGR) is used to analyze the irreversible loss. It turns out that the SWWVI can reduce the flow field loss. There are several weak shock waves in the NGV flow field, which can increase the irreversible loss. This work offers flow mechanisms analysis and presents the EGR distribution in SWBLI and SWWVI areas in a transonic turbine blade.

The Role of Vortex Shedding in the Trailing Edge Loss of Transonic Turbine Blades

The loss of square, round, and elliptical turbine trailing edge geometries, and the mechanisms responsible, is assessed using a two-part experimental program. In the first part, a single blade experiment, in a channel with contoured walls, allowed rapid testing of a range of trailing edge sizes and shapes. In the second part, turbine blade cascades with a subset of sizes of the trailing edge geometries tested in part one were evaluated in a closed-loop variable density facility, at exit Mach numbers from 0.40 to 0.97, and exit Reynolds numbers from 1.5 × 105 to 2.5 × 106. Throughout the test campaign, detailed instantaneous Schlieren images of the trailing edge flows have been obtained to identify the underlying unsteady mechanisms in the base region. The experiments reveal the importance of suppressing transonic vortex shedding, and quantify the influence of this mechanism on loss. The state and thickness of the blade boundary layers immediately upstream of the trailing edge are of critical importance in determining the onset of transonic vortex shedding. Elliptical trailing edge geometries have also been found to be effective at suppressing transonic vortex shedding. For trailing edges that exhibit transonic vortex shedding, a mechanism is identified whereby reflected shed shockwaves encourage or discourage vortex shedding depending on the phase with which the shocks return to the trailing edge, capable of modifying the loss generated.

Computational investigation of heat load and secondary flows near tip region in a transonic turbine rotor with moving shroud

The purpose of this paper is to investigate numerically the effect of the relative shroud motion on both the heat transfer and secondary flows near the tip of a transonic turbine blade. The simulations are done by solving the compressible Reynolds averaged Navier–Stokes (RANS) equations using the finite volume method implemented in the Fluent CFD software. The investigated rotor blade geometry was designed by SNECMA for modern aero-engines. The computational grid is created using multi-block topology and each block support a structured mesh. Calculations are done under a transonic flow regime for both stationary and moving shroud cases. The aerothermal properties and flow topology at the blade tip, the shroud and the blade suction surface are compared with and without shroud movement. The relative motion has significantly changed the distribution and the level of the heat transfer coefficient, the pressure coefficient, the isentropic Mach number and the flow topology.

Heat Transfer Performance of a Transonic Turbine Blade Passage in the Presence of Leakage Flow Through Upstream Slot and Mateface Gap With Endwall Contouring

Comparison of heat transfer performance of a nonaxisymmetric contoured endwall to a planar baseline endwall in the presence of leakage flow through stator–rotor rim seal interface and mateface gap is reported in this paper. Heat transfer experiments were performed on a high turning turbine airfoil passage at Virginia Tech's transonic blow down cascade facility under design conditions for two leakage flow configurations—(1) mateface blowing only, (2) simultaneous coolant injection from the upstream slot and mateface gap. Coolant to mainstream mass flow ratios (MFRs) were 0.35% for mateface blowing only, whereas for combination blowing, a 1.0% MFR was chosen from upstream slot and 0.35% MFR from mateface. A common source of coolant supply to the upstream slot and mateface plenum made sure the coolant temperatures were identical at both upstream slot and mateface gap at the injection location. The contoured endwall geometry was generated to minimize secondary aerodynamic losses. Transient infrared thermography technique was used to measure endwall surface temperature and a linear regression method was developed for simultaneous calculation of heat transfer coefficient (HTC) and adiabatic cooling effectiveness, assuming a one-dimensional (1D) semi-infinite transient conduction. Results indicate reduction in local hot spot regions near suction side as well as area averaged HTC using the contoured endwall compared to baseline endwall for all coolant blowing cases. Contoured geometry also shows better coolant coverage further along the passage. Detailed interpretation of the heat transfer results along with near endwall flow physics has also been discussed.

Influence of cooling intensity on shock wave boundary layer interaction region in turbine cascade

The shock wave boundary layer interaction on the suction side of a transonic turbine blade was one of the main objectives of the TFAST project. For this purpose a model of a turbine passage was designed, manufactured and assembled in a transonic wind tunnel. The paper presents the experimental investigations concerning the flow structure on the transonic turbine blade. A clean case (without a cooling system) with a normal shock wave interacting with a laminar boundary layer and also the influence of the blade cooling system with three different coolant blowing intensities on the laminar interaction region were investigated.