Cavity Purge Research Articles

Time-Averaged Aerodynamics of a High-Speed Low-Pressure Turbine Cascade With Cavity Purge and Unsteady Wakes

Abstract The time-averaged aerodynamics of a high-speed low-pressure turbine (LPT) cascade were investigated under the impact of unsteady wakes and purge flows. The tests were performed at an outlet Mach (M) and Reynolds numbers (Re) of 0.90 and 70 k, respectively. Unsteady wakes were simulated by means of a spoked-wheel type wake generator (WG), and a reduced frequency of 0.95 was achieved. The development of a purge flow system to operate at low-pressure levels is presented alongside guidelines for the operation of the circuit. The new purge system was commissioned in terms of its long-term stability and flow uniformity at the cascade inlet. The impact of varying purge flows from a cavity versus a flat endwall was assessed by means of static pressure measurements on the blade pressure side (PS) and suction side (SS) and traverses performed downstream with a miniaturized multi-hole probe. Differences in the secondary flow structures resultant from injecting purge flow into the flowfield are observed, namely, the intensification of the trailing shed vortex (TSV) and passage vortex (PV) that is also displaced away from the endwall. An increase of the endwall losses occurs as the flat endwall was replaced with a cavity geometry and the cavity purge flowrate increases.

Brush seal design and secondary air system modification of a heavy-duty gas turbine to improve the output power

The overall performance of heavy-duty gas turbines is strongly affected by the mass flow distribution of the secondary air system. Reducing the mass flow of the secondary air is a remarkable method to improve the performance of a gas turbine. The aim of this article is to implement a brush seal as an alternative to labyrinth seal. Based on the experiences gained from the integrity of brush seals into a gas turbine, installing a brush seal would not necessarily improve turbine performance. This is because of SAS mass flow redistribution, which may not lead to overall SAS mass flow reduction. Therefore, some other components in the SAS arrangement should be modified as controlling variables to overcome this problem. The effect of these modifications on the airflow distribution is investigated using an in-house network analysis code. To establish an optimum solution (optimized brush seal geometry and controlling variables), the network analysis code is coupled with an in-house optimization code that benefits from the Genetic Algorithms and Artificial Neural Networks. Some constraints including upstream and downstream cavity purge flow and vane cooling air are considered in the optimization process. The final network results show a 33.27% reduction in overall SAS mass flow. To ensure an improvement in the performance of the new three-stage turbine, a CFD analysis is conducted, which indicates 1.0% more power with respect to the original turbine. In the end, the aerodynamic and mechanical behavior of the brush seal is analyzed using CFD and FEM.

A Sector-Cascade Test Rig for Measurements of Heat Transfer in Turbine Center Frames

Abstract The turbine center frame (TCF) is an inherent component of turbofan aircraft engines and is used for connecting the high-pressure turbine (HPT) to the low-pressure turbine (LPT). Its position immediately downstream of the HPT makes it susceptible to the extremely high temperatures of future engines. Despite this, fundamental knowledge of heat transfer in TCFs and the influencing factors is still missing. This paper presents a new 45-deg sector-cascade test rig specifically designed for fundamental studies of film cooling effectiveness and heat transfer coefficient in TCFs and the development and validation of a measurement technique involving infrared thermography and heating foils. Measurements of heat transfer coefficient in the TCF were taken for two purge-to-mainstream mass flow ratios corresponding to the case of no purge and nominal (to engine operation) purge. The magnitude of the heat transfer coefficients on the hub and strut surfaces was highly influenced by the various flow structures in the passage and by the velocity variation of the mainstream flow due to the “aggressive” design of the TCF. Heat transfer on the surface of the strut was mainly governed by boundary layer behavior (laminar near the leading edge and turbulent for the rest of the strut) augmented by the effect of the secondary flow structures. Measurements of film cooling effectiveness were also taken for the single case of the nominal purge. A region of high film cooling effectiveness was observed, extending from the purge cavity exit to about 40% of the passage axial length. In this region, the effectiveness decreased with increasing axial length. On the surface of the struts and fillet radii, the film cooling effectiveness was found to be zero. This was attributed to the effect of the horseshoe vortex that sweeps the purge flow away from the strut surface and dilutes it by continuously entraining hot mainstream flow.

Experimental Investigation of Turbine Stage Flow Field and Performance at Varying Cavity Purge Rates and Operating Speeds

The aspect of hub cavity purge has been investigated in a high-pressure axial low-reaction turbine stage. The cavity purge is an important part of the secondary air system, used to isolate the cavities below the hub level from the hot main annulus flow. A full-scale cold-flow experimental rig featuring a rotating stage was used in the investigation, quantifying main annulus flow field impact with respect to purge flow rate as it was injected upstream of the rotor. Five operating speeds were investigated of which three with respect to purge flow, namely, a high loading design case, and two high-speed points encompassing the peak efficiency. At each of these operating speeds, the amount of purge flow was varied from 0% to 2%. Observing the effect of the purge rate on measurement plane averaged parameters, a minor flow angle decrease and Mach number increase is seen for the low speed case, while maintaining near constant values for the higher operating speeds. The prominent effect due to purge is seen in the efficiency, showing a linear sensitivity to purge of 1.3%-points for every 1% of added purge flow for the investigated speeds. While spatial average values of flow angle and Mach number are essentially unaffected by purge injection, important spanwise variations are observed and highlighted. The secondary flow structure is strengthened in the hub region, leading to a generally increased over-turning and lowered flow velocity. Meanwhile, the added volume flow through the rotor leads to higher outlet flow velocities visible at higher span, with associated decreased turning. A radial efficiency distribution is utilized, showing negative impact through span heights from 15% to 70%. Pitchwise variation of investigated flow parameters is significantly influenced by purge flow, making this a parameter to include for instance when evaluating benefits of stator clocking positions.

Turbine Blade Surface Phantom Cooling from Upstream Nozzle Trailing-Edge Ejection

This paper presents upstream nozzle trailing-edge coolant ejection on downstream uncooled blades. A pressure-sensitive paint mass transfer technique provides detailed phantom cooling effectiveness distribution on modeled land-based turbine rotor blade surfaces. Without cavity purge and tip leakage flows, a uniform blade inlet temperature is adopted in the current study. Experiments have been completed in a low-speed wind-tunnel facility with a five-blade linear cascade. The inlet Reynolds numbers based on chord length are 100,000 and 200,000. Nozzle trailing-edge coolant ejection on a rotor blade is simulated by a spoked-wheel-type rotating facility with 32 hollow rods equipped with coolant ejection from 128 holes per rod. Coolant to mainstream density ratio maintains at 1.5 to match engine conditions. Nozzle coolant discharge velocity to nozzle mainstream velocity ratio varies from 0.4 to 1.4. Velocity ratios from 0.4–0.6 are closest to typical engine conditions. Coolant to mainstream mass flow rate ratio effect is from 0.67 to 2.94%. Higher phantom cooling effectiveness occurs on suction and pressure surfaces at the velocity ratio of 0.4–0.6 and over 1.0, respectively. Velocity ratio effect impacts phantom cooling effectiveness distribution more than the mass flow rate ratio effect. Most of the trailing-edge coolant migrates toward blade inner and outer spans than the blade midspan.

Experimental Study of Ingestion in the Rotor–Stator Disk Cavity of a Subscale Axial Turbine Stage

This paper describes experiments in a subscale axial turbine stage equipped with an axially overlapping radial-clearance seal at the disk cavity rim and a labyrinth seal radially inboard which divides the disk cavity into a rim cavity and an inner cavity. An orifice model of the rim seal is presented; values of ingestion and egress discharge coefficients based on the model and experimental data are reported for a range of cavity purge flow rate. In the experiments, time-averaged pressure distribution was measured in the main gas annulus and in the disk cavity; also measured was the time-averaged ingestion into the cavity. The pressure and ingestion data were combined to obtain the discharge coefficients. Locations on the vane platform 1 mm upstream of its lip over two vane pitches circumferentially defined the main gas annulus pressure; in the rim cavity, locations at the stator surface in the radially inner part of the “seal region” over one vane pitch defined the cavity pressure. For the sealing effectiveness, two locations in the rim cavity at the stator surface, one in the “mixing region” and the other radially further inward at the beginning of the stator boundary layer were considered. Two corresponding sets of ingestion and egress discharge coefficients are reported. The ingestion discharge coefficient was found to decrease in magnitude as the purge flow rate increased; the egress discharge coefficient increased with purge flow rate. The discharge coefficients embody fluid-mechanical effects in the ingestion and egress flows. Additionally, the minimum purge flow rate required to prevent ingestion was estimated for each experiment set and is reported. It is suggested that the experiments were in the combined ingestion (CI) region with externally induced (EI) ingestion being the dominant contributor.

The Effect of Annulus Performance Parameters on Rotor-Stator Cavity Sealing Flow

The amount of cooling air assigned to seal high pressure (HP) turbine rim cavities is critical for performance as well as component life. Insufficient air leads to excessive hot annulus gas ingestion and its penetration deep into the cavity compromising disk or cover plate life. Excessive purge air, on the other hand, adversely affects performance. This paper is a continuation of the authors' work on ingestion reported by Mirzamoghadam et al. (2008) (“3D CFD Ingestion Evaluation of a High Pressure Turbine Rim Seal Disk Cavity,” ASME Paper No. GT2008-50531), where the main focus of that investigation was to qualitatively describe ingestion driven by annulus circumferential pressure asymmetry under constant annulus conditions and rotational speed. In this paper, the research team investigated the variation of annulus circumferential pressure fluctuation and rotational speed on the double overlap platform rim seal cavity reported in part-1. The outcome from this study was to map out the resulting nondimensional minimum sealing flow (minimum value of Cw or Cw,min) as it relates to entrained ingestion in the absence of cavity cooling flow (Cw,ent). As was done in part-1, the runs were made with 3D computational fluid dynamics (CFD) in setup/run mode option using Fine/Turbo. At two rotational speeds, annulus conditions were varied by reducing turbine inlet pressure (i.e., mass flow) from the baseline operating condition, and the resulting pressure fluctuation was quantified. In addition, an investigation to assess the selected aft-located mixing plane steady state solution for this study as compared to the forward-located steady run was performed using unsteady (nonlinear Harmonics) CFD as the referee. The results yielded the linear decrease in Cw,ent at fixed rotational Reynolds number as annulus Reynolds number was decreased. Moreover, the rate of change in entrained flow sharply increases with increase in rotational Reynolds number. As annulus mass flow is reduced to a critical value defined by annulus-to-rotational Reynolds number ratio, the CFD prediction for Cw,ent converges to the turbulent boundary layer entrainment solution for the rotor, and Cw,min reverts to the rotational Reynolds number dominating region. The results from this study were compared to what has been observed by a previous study for a single overlap platform geometry. The resulting design curve allows insight in relating cavity purge flow requirements versus turbine cycle parameters which could lead to better efficiency.

History of Short-Duration Measurement Programs Related to Gas Turbine Heat Transfer, Aerodynamics, and Aeroperformance at Calspan and The Ohio State University

Short-duration facilities have been used for the past 35 years to obtain measurements of heat transfer, aerodynamic loading, vibratory response, film-cooling influence, purge flow migration, and aeroperformance for full-stage, high-pressure turbines operating at design-corrected conditions of flow function, corrected speed, and stage pressure ratio. This paper traces the development of experimental techniques now in use at The Ohio State University (OSU) Gas Turbine Laboratory (GTL) from initial work in this area at the Cornell Aeronautical Laboratory (CAL, later to become Calspan) from 1975 through to the present. It is intended to summarize the wide range of research that can be performed with a short-duration facility and highlight the types of measurements that are possible. Beginning with heat flux measurements for the vane and blade of a Garrett TFE 731-2 HP turbine stage with vane pressure-surface slot cooling, the challenge of each experimental program has been to provide data to aid turbine designers in understanding the relevant flow physics and help drive the advancement of predictive techniques. Through many different programs, this has involved collaborators at a variety of companies and experiments performed with turbine stages from Garrett, Allison, Teledyne, Pratt and Whitney (P/W), General Electric Aviation (GEA), Rocketdyne, Westinghouse, and Honeywell. The vane/blade interaction measurement and computational fluid dynamics (CFD) program, which ran from the early 1980s until 2000, provided a particularly good example of what can be achieved when experimentalists and computational specialists collaborate closely. Before conclusion of this program in 2000, the heat flux and pressure measurements made for this transonic turbine operated with and without vane trailing edge cooling flow were analyzed and compared to predictive codes in conjunction with engineers at Allison, United Technologies Research Center (UTRC), P/W, and GEA in jointly published papers. When the group moved to OSU in 1995 along with the facility used at Calspan, refined techniques were needed to meet new research challenges, such as investigating blade-damping and forced response, measuring aeroperformance for different configurations, and preparing for advanced cooling experiments that introduced complicating features of an actual engine to further challenge computational predictions. This required conversion of the test-gas heating method from a shock-tunnel approach to a blowdown approach using a combustor emulator to also create inlet temperature profiles, the development of instrumentation techniques to work with a thin-walled airfoil with backside cooling, and the adoption of experimental techniques that could be used to successfully operate fully cooled turbine stages (vane row-cooled, blade row-cooled, and proper cavity purge flow provided). Further, it was necessary to develop techniques for measuring the aeroperformance of these fully cooled machines.

Time-Averaged and Time-Accurate Aerodynamic Effects of Forward Rotor Cavity Purge Flow for a High-Pressure Turbine—Part I: Analytical and Experimental Comparisons

The effect of rotor purge flow on the unsteady aerodynamics of a high-pressure turbine stage operating at design corrected conditions has been investigated, both experimentally and computationally. The experimental configuration consisted of a single-stage high-pressure turbine with a modern film-cooling configuration on the vane airfoil and the inner and outer end wall surfaces. Purge flow was introduced into the cavity located between the high-pressure vane and the high-pressure disk. The high-pressure blades and the downstream low-pressure turbine nozzle row were not cooled. All of the hardware featured an aerodynamic design typical of a commercial high-pressure ratio turbine and the flow path geometry was representative of the actual engine hardware. In addition to instrumentation in the main flow path, the stationary and rotating seals of the purge flow cavity were instrumented with high frequency response flush-mounted pressure transducers and miniature thermocouples in order to measure the flow field parameters above and below the angel wing. Predictions of the time-dependent flow field in the turbine flow path were obtained using FINE/Turbo, a three-dimensional Reynolds-averaged Navier–Stokes computational fluid dynamics CFD code that had the capability to perform both a steady and unsteady analysis. The steady and unsteady flow fields throughout the turbine were predicted using a three blade-row computational model that incorporated the purge flow cavity between the high-pressure vane and disk. The predictions were performed in an effort to mimic the design process with no adjustment of boundary conditions to better match the experimental data. The time-accurate predictions were generated using the harmonic method. Part I of this paper concentrates on the comparison of the time-averaged and time-accurate predictions with measurements in and around the purge flow cavity. The degree of agreement between the measured and predicted parameters is described in detail, providing confidence in the predictions for the flow field analysis that will be provided in Part II.

Numerical investigation of the interaction between upstream cavity purge flow and main flow in low aspect ratio turbine cascade

In modern gas turbines, rim seal located between the stator-disc and rotor-disc is used to prevent hot-gas ingestion into the inner stage-gap of high pressure turbine. However, the purge flow supplied to the cavity through the rim seal interacts with the main flow, producing additional aerodynamic loss due to the mixing process which plays a significant role in the formation, development and evolution of downstream secondary flow. In this paper, a set of cascade representative of low aspect ratio turbine is selected to numerically investigate the influence of upstream cavity purge flow on the hub secondary flow structure and aerodynamic loss. Cascade with/without upstream cavity and four different purge mass flow rates are all taken into account in this simulation. Then, a deep insight into the loss mechanism of interaction between purge flow and main flow is gained. The results show that the presence of cavity and purge flow has a significant impact on the main flow which not only changes the vortex structure in both the passage and upstream cavity, but also alters the cascade exit flow angle distribution along the spanwise. Moreover, aerodynamic loss in the cascade rises with the increase of purge flow rate while the sealing effect is also enhanced. Therefore, the effect of upstream cavity purge flow must be considered in the process of turbine aerodynamic design. What is more, it is necessary to minimize the purge flow rate in order to reduce aerodynamic loss on the premise of satisfying cooling requirements.

Heat Transfer for the Film-Cooled Vane of a 1-1/2 Stage High-Pressure Transonic Turbine—Part I: Experimental Configuration and Data Review With Inlet Temperature Profile Effects

This paper investigates the vane airfoil and inner endwall heat transfer for a full-scale turbine stage operating at design corrected conditions under the influence of different vane inlet temperature profiles and vane cooling flow rates. The turbine stage is a modern 3D design consisting of a cooled high-pressure vane, an un-cooled high-pressure rotor, and a low-pressure vane. Inlet temperature profiles (uniform, radial, and hot streaks) are created by a passive heat exchanger and can be made circumferentially uniform to within ±5% of the bulk average inlet temperature when desired. The high-pressure vane has full cooling coverage on both the airfoil surface and the inner and outer endwalls. Two circuits supply coolant to the vane, and a third circuit supplies coolant to the rotor purge cavity. All of the cooling circuits are independently controlled. Measurements are performed using double-sided heat-flux gauges located at four spans of the vane airfoil surface and throughout the inner endwall region. Analysis of the heat transfer measured for the uncooled downstream blade row has been reported previously. Part I of this paper describes the operating conditions and data reduction techniques utilized in this analysis, including a novel application of a traditional statistical method to assign confidence limits to measurements in the absence of repeat runs. The impact of Stanton number definition is discussed while analyzing inlet temperature profile shape effects. Comparison of the present data (Build 2) to the data obtained for an uncooled vane (Build 1) clearly illustrates the impact of the cooling flow and its relative effects on both the endwall and airfoils. Measurements obtained for the cooled hardware without cooling applied agree well with the solid airfoil for the airfoil pressure surface but not for the suction surface. Differences on the suction surface are due to flow being ingested on the pressure surface and reinjected on the suction surface when coolant is not supplied for Build 2. Part II of the paper continues this discussion by describing the influence of overall cooling level variation and the influence of the vane trailing edge cooling on the vane heat transfer measurements.

Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine—Part I: Influence of Cooling Variation

Heat-flux measurements are presented for a one-and-one-half stage high-pressure turbine operating at design-corrected conditions with modulated cooling flows in the presence of different inlet temperature profiles. Coolant is supplied from a heavily film-cooled vane and the purge cavity (between the rotor disk and the upstream vane) but not from the rotor blades, which are solid metal. Thin-film heat-flux gauges are located on the uncooled blade pressure and suction surface (at multiple span locations), on the blade tip, on the blade platform, and on the disk and vane sides of the purge cavity. These measurements provide a comprehensive picture of the effect of varying cooling flow rates on surface heat transfer to the turbine blade for uniform and radial inlet temperature profiles. Part I of this paper examines the macroscopic influence of varying all cooling flows together, while Part II investigates the individual regions of influence of the vane outer and purge cooling circuits in more detail. The heat-flux gauges are able to track the cooling flow over the suction surface of the airfoil as it wraps upwards along the base of the airfoil for the uniform vane inlet temperature profile. A similar comparison for the radial profile shows the same coolant behavior but with less pronounced changes. From these comparisons, it is clear that cooling impacts each temperature profile similarly. Nearly all of the cooling influence is limited to the blade suction surface, but small changes are observed for the pressure surface. In addition to the cooling study, a novel method of calculating the adiabatic wall temperature is demonstrated. The derived adiabatic wall temperature distribution shows very similar trends to the Stanton number distribution on the blade.

Heat Transfer for the Blade of a Cooled Stage and One-Half High-Pressure Turbine—Part II: Independent Influences of Vane Trailing Edge and Purge Cooling

The independent influences of vane trailing edge and purge cooling are studied in detail for a one-and-one-half stage transonic high-pressure turbine operating at design-corrected conditions. This paper builds on the conclusions of Part I, which investigated the combined influence of all cooling circuits. Heat-flux measurements for the airfoil, platform, tip, and root of the turbine blade, as well as the shroud and the vane side of the purge cavity, are used to track the influence of cooling flow. By independently varying the coolant flow rate through the vane trailing edge or purge circuit, the region of influence of each circuit can be isolated. Vane trailing edge cooling is found to create the largest reductions in blade heat transfer. However, much of the coolant accumulates on the blade suction surface and little influence is observed for the pressure surface. In contrast, the purge cooling is able to cause small reductions in heat transfer on both the suction and pressure surfaces of the airfoil. Its region of influence is limited to near the hub, but given that the purge coolant mass flow rate is 1/8 that of the vane trailing edge, it is impressive that any impact is observed at all. The cooling contributions of these two circuits account for nearly all of the cooling reductions observed for all three circuits in Part I, indicating that the vane inner cooling circuit that feeds most of the vane film-cooling holes has little impact on the downstream blade heat transfer. Time-accurate pressure measurements provide further insight into the complex interactions in the purge region that govern purge coolant injection. While the pressures supplying the purge coolant and the overall coolant flow rate remain fairly constant, the interactions of the vane pressure field and the rotor pressure field create moving regions of high pressure and low pressure at the exit of the cavity. This results in pulsing regions of injection and ingestion.

Temperature Predictions and Comparison With Measurements for the Blade Leading Edge and Platform of a 1 1/2 Stage Transonic HP Turbine

A series of computational predictions generated using FINE/TURBO are compared with data to investigate implementation techniques available for predicting temperature migration through a turbine stage. The experimental results used for comparison are from a one-and-one-half stage turbine operating at design-corrected conditions in a short-duration facility. Measurements of the boundary conditions are used to set up the computational models, and the predicted temperatures are compared with measured fluid temperatures at the blade leading edge and just above the blade platform. Fluid temperature measurements have not previously been available for these locations in a transonic turbine operating at design-corrected conditions, so this represents a novel comparison. Accurate predictions for this short-duration turbine experiment require use of the isothermal wall boundary condition instead of an adiabatic boundary condition and accurate specification of the inlet temperature profile all the way to the wall. Predictions using the harmonic method agree with the temperatures measured for the blade leading edge from 65% to 95% span to within 1% normalized temperature data. Agreement over much of the rest of the leading edge is within 5% of the measured value. Comparisons at 5–10% span and for the blade platform show larger differences up to 10%, which indicates that the flow in this region is not fully captured by the prediction. This is not surprising since the purge cavity and platform leading-edge features present in the experiment are treated as a smooth hub wall in the current simulation. This work represents a step toward the larger goal of accurately predicting surface heat-flux for the complicated environment of an operational engine as it is reproduced in a laboratory setting. The experiment upon which these computations are based includes realistic complications such as one-dimensional and two-dimensional inlet temperature profiles, a heavily film-cooled vane, and purge cooling. While the ultimate goal is to accurately handle all of these features, the current model focuses on the treatment of a subset of experiments performed for a one-dimensional radial inlet temperature profile and no cooling.

Aerodynamics and Heat Transfer for a Cooled One and One-Half Stage High-Pressure Turbine–Part II: Influence of Inlet Temperature Profile on Blade Row and Shroud

Heat flux measurements are presented for the uncooled blades of a one and one-half stage turbine operating at design corrected conditions with a fully cooled upstream vane row and with rotor disk cavity purge flow. This paper highlights the differences in blade heat flux and temperature caused by uniform, radial, and hot streak inlet temperature profiles. A general discussion of temperature profile migration is provided in Part I, and Part III presents data for hot streak magnitudes and alignments. The heat flux and fluid temperature measurements for the blade airfoil, platform, angel wing (near the root), and tip as well as for the stationary outer shroud are influenced by the vane inlet temperature profile. The inlet temperature profile shape can be clearly observed in the blade Stanton number measurements, with the radial and hot streak profiles showing a greater redistribution of energy than the uniform case due to secondary flows. Hot-gas segregation is observed to increase with the strength of the temperature distortion. Measurements for the hot streak profile show a segregation of higher temperature fluid to the pressure surface when compared with a uniform profile. The introduction of vane and purge cooling is found to further accentuate the flow segregation due to coolant migration to the suction surface.

Investigation of the Flow Field on a Transonic Turbine Nozzle Guide Vane With Rim Seal Cavity Flow Ejection

Ensuring an adequate life of high pressure turbines requires efficient cooling methods such as rim seal flow ejection from the stator-rotor wheel space cavity interface, which prevents hot gas ingress into the rotor disk. The present paper addresses the potential to improve the efficiency in transonic turbines at certain rim seal ejection rates. To understand this process, a numerical study was carried out, combining computational fluid dynamic (CFD) simulations and experiments on a single stage axial test turbine. The three dimensional steady CFD analysis was performed, modeling the purge cavity flow ejected downstream of the stator blade row at three flow regimes: subsonic M2=0.73, transonic M2=1.12, and supersonic M2=1.33. Experimental static pressure measurements were used to calibrate the computational model. The main flow field-purge flow interaction is found to be governed by the vane shock structures at the stator hub. The interaction between the vane shocks at the hub and the purge flow has been studied and quantitatively characterized as a function of the purge ejection rate. The ejection of 1% of the core flow from the rim seal cavity leads to an increase in the hub static pressure of approximately 7% at the vane trailing edge. This local reduction of the stator exit Mach number decreases the trailing edge losses in the transonic regime. Finally, a numerically predicted loss breakdown is presented, focusing on the relative importance of the trailing edge losses, boundary layer losses, shock losses, and mixing losses, as a function of the purge rate ejected. Contrary to the experience in subsonic turbines, results in a transonic model demonstrate that ejecting purge flow improves the vane efficiency due to the shock structure modification downstream of the stator.

The Impact of Gas Modeling in the Numerical Analysis of a Multistage Gas Turbine

In this work a numerical investigation of a four stage heavy-duty gas turbine is presented. Fully three-dimensional, multistage, Navier-Stokes analyses are carried out to predict the overall turbine performance. Coolant injections, cavity purge flows, and leakage flows are included in the turbine modeling by means of suitable wall boundary conditions. The main objective is the evaluation of the impact of gas modeling on the prediction of the stage and turbine performance parameters. To this end, four different gas models were used: three models are based on the perfect gas assumption with different values of constant cp, and the fourth is a real gas model which accounts for thermodynamic gas properties variations with temperature and mean fuel∕air ratio distribution in the through-flow direction. For the real gas computations, a numerical model is used which is based on the use of gas property tables, and exploits a local fitting of gas data to compute thermodynamic properties. Experimental measurements are available for comparison purposes in terms of static pressure values at the inlet∕outlet of each row and total temperature at the turbine exit.

Thermal Details in a Rotor–Stator Cavity at Engine Conditions With a Mainstream

This investigation involves a numerical study of enclosed Rotor–Stator cavities of gas turbine engines. The complete elliptic form of the 2-D, axisymmetric Navier–Stokes equations for compressible turbulent flow were solved. Included are the complete fluid and thermal effects of the hot mainstream gas interacting with the cooling cavity purge flow at actual engine flow conditions for generalized geometries. Additional flow conditions above and below those for engine nominal conditions are also considered. The relationships among the important flow parameters are investigated by examining the entire set of computations. The predictions reveal that a small recirculation zone in the stator shroud axial gap region is the primary mechanism for the considerable thermal transport from the mainstream to the turbine blade root/retainer region of the rotor.