Time-Accurate Predictions for a Fully Cooled High-Pressure Turbine Stage—Part II: Methodology for Quantifications of Prediction Quality

2009 ◽  
Vol 131 (3) ◽  
Author(s):  
C. W. Haldeman ◽  
M. G. Dunn ◽  
S. A. Southworth ◽  
J.-P. Chen ◽  
G. Heitland ◽  
...  
Keyword(s):  
High Pressure ◽  
Design Methodology ◽  
Pressure Ratio ◽  
Turbulence Models ◽  
Test Facility ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Prediction Quality ◽  
The Difference ◽  
The Ohio State University

The aerodynamics of a fully cooled, axial, single stage high-pressure turbine operating at design corrected conditions of corrected speed, flow function, and stage pressure ratio has been investigated experimentally and computationally and presented in Part I of this paper. In that portion of the paper, flow-field predictions obtained using the computational fluid dynamics codes Numeca’s FINE/TURBO and the code TURBO were obtained using different design methodologies that approximated the fully-cooled turbine stage in different ways. These predictions were compared to measurements obtained using the Ohio State University Gas Turbine Laboratory Turbine Test Facility, in a process that was essentially a design methodology validation study, instead of a computational methodology optimization study. The difference between the two is that the designers were given one chance to use their codes (as a designer would normally do) instead of using the existing data to fine-tune their grids/methodologies by doing grid studies and changes in the turbulence models employed. Part I of this paper showed differing results from the two solvers, which appeared to be mainly dependent on the differences in grid resolution and/or modeling features selected by the code users. Examining these occurrences points to places where the design methodology could be improved, but it became clear that metrics were needed to compare overall performance of each approach. In this part of the paper, three criteria are proposed for measuring overall prediction quality of the unsteady predictions, which include the unsteady envelope size, envelope shape, and power spectrum. These measures capture the main characteristics of the unsteady data and allow designers to use the criteria of most interest to them. In addition, these can be used to track how well predictions improve over time as grid resolutions and modeling techniques change.

scholarly journals Calculation and Correlation of the Unsteady Flowfield in a High Pressure Turbine

2002 ◽  
Author(s):  
Milind A. Bakhle ◽  
Jong S. Liu ◽  
Josef Panovsky ◽  
Theo G. Keith ◽  
Oral Mehmed
Keyword(s):  
High Pressure ◽  
Three Dimensional ◽  
Unsteady Aerodynamics ◽  
Forced Vibrations ◽  
Forced Response ◽  
Operating Conditions ◽  
Test Facility ◽  
Navier Stokes ◽  
Turbine Stage ◽  
High Pressure Turbine

Forced vibrations in turbomachinery components can cause blades to crack or fail due to high-cycle fatigue. Such forced response problems will become more pronounced in newer engines with higher pressure ratios and smaller axial gap between blade rows. An accurate numerical prediction of the unsteady aerodynamics phenomena that cause resonant forced vibrations is increasingly important to designers. Validation of the computational fluid dynamics (CFD) codes used to model the unsteady aerodynamic excitations is necessary before these codes can be used with confidence. Recently published benchmark data, including unsteady pressures and vibratory strains, for a high-pressure turbine stage makes such code validation possible. In the present work, a three dimensional, unsteady, multi blade-row, Reynolds-Averaged Navier Stokes code is applied to a turbine stage that was recently tested in a short duration test facility. Two configurations with three operating conditions corresponding to modes 2, 3, and 4 crossings on the Campbell diagram are analyzed. Unsteady pressures on the rotor surface are compared with data.

Effect of Low-NOX Combustor Swirl Clocking on Intermediate Turbine Duct Vane Aerodynamics With an Upstream High Pressure Turbine Stage—An Experimental and Computational Study

2016 ◽  
Vol 139 (1) ◽  
Author(s):  
Martin Johansson ◽  
Thomas Povey ◽  
Kam Chana ◽  
Hans Abrahamsson
Keyword(s):  
High Pressure ◽  
Test Facility ◽  
Flow Structures ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Inlet Condition ◽  
Intermediate Turbine Duct ◽  
Turbine Vane ◽  
Flow Through ◽  
Inlet Conditions

Flow in an intermediate turbine duct (ITD) is highly complex, influenced by the upstream turbine stage flow structures, which include tip leakage flow and nonuniformities originating from the upstream high pressure turbine (HPT) vane and rotor. The complexity of the flow structures makes predicting them using numerical methods difficult, hence there exists a need for experimental validation. To evaluate the flow through an intermediate turbine duct including a turning vane, experiments were conducted in the Oxford Turbine Research Facility (OTRF). This is a short duration high speed test facility with a 3/4 engine-sized turbine, operating at the correct nondimensional parameters for aerodynamic and heat transfer measurements. The current configuration consists of a high pressure turbine stage and a downstream duct including a turning vane, for use in a counter-rotating turbine configuration. The facility has the ability to simulate low-NOx combustor swirl at the inlet to the turbine stage. This paper presents experimental aerodynamic results taken with three different turbine stage inlet conditions: a uniform inlet flow and two low-NOx swirl profiles (different clocking positions relative to the high pressure turbine vane). To further explain the flow through the 1.5 stage turbine, results from unsteady computational fluid dynamics (CFD) are included. The effect of varying the high pressure turbine vane inlet condition on the total pressure field through the 1.5 stage turbine, the intermediate turbine duct vane loading, and intermediate turbine duct exit condition are discussed and CFD results are compared with experimental data. The different inlet conditions are found to alter the flow exiting the high pressure turbine rotor. This is seen to have local effects on the intermediate turbine duct vane. With the current stator–stator vane count of 32-24, the effect of relative clocking between the two is found to have a larger effect on the aerodynamics in the intermediate turbine duct than the change in the high pressure turbine stage inlet condition. Given the severity of the low-NOx swirl profiles, this is perhaps surprising.

Heat Transfer Measurements and Predictions for a Modern, High-Pressure, Transonic Turbine, Including Endwalls

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
James A. Tallman ◽  
Charles W. Haldeman ◽  
Michael G. Dunn ◽  
Anil K. Tolpadi ◽  
Robert F. Bergholz
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbulence Modeling ◽  
Pressure Ratio ◽  
Operating Conditions ◽  
Test Facility ◽  
Navier Stokes ◽  
Local Measurements ◽  
Transonic Turbine ◽  
The Ohio State University

This paper presents both measurements and predictions of the hot-gas-side heat transfer to a modern, 112 stage high-pressure, transonic turbine. Comparisons of the predicted and measured heat transfer are presented for each airfoil at three locations, as well as on the various endwalls and rotor tip. The measurements were performed using the Ohio State University Gas Turbine Laboratory Test Facility (TTF). The research program utilized an uncooled turbine stage at a range of operating conditions representative of the engine: in terms of corrected speed, flow function, stage pressure ratio, and gas-to-metal temperature ratio. All three airfoils were heavily instrumented for both pressure and heat transfer measurements at multiple locations. A 3D, compressible, Reynolds-averaged Navier–Stokes computational fluid dynamics (CFD) solver with k-ω turbulence modeling was used for the CFD predictions. The entire 112 stage turbine was solved using a single computation, at two different Reynolds numbers. The CFD solutions were steady, with tangentially mass-averaged inlet/exit boundary condition profiles exchanged between adjacent airfoil-rows. Overall, the CFD heat transfer predictions compared very favorably with both the global operation of the turbine and with the local measurements of heat transfer. A discussion of the features of the turbine heat transfer distributions, and their association with the corresponding flow-physics, has been included.

Time-Accurate Predictions for a Fully Cooled High-Pressure Turbine Stage—Part I: Comparison of Predictions With Data

2009 ◽  
Vol 131 (3) ◽  
Author(s):  
S. A. Southworth ◽  
M. G. Dunn ◽  
C. W. Haldeman ◽  
J.-P. Chen ◽  
G. Heitland ◽  
...  
Keyword(s):  
High Pressure ◽  
Power Spectrum ◽  
Surface Pressure ◽  
Pressure Ratio ◽  
Computing Time ◽  
Flow Simulation ◽  
Operating Conditions ◽  
Test Facility ◽  
High Pressure Turbine ◽  
Cooling Flow

The aerodynamics of a fully cooled axial single stage high-pressure turbine operating at design corrected conditions of corrected speed, flow function, and stage pressure ratio has been investigated. This paper focuses on flow field predictions obtained from the viewpoint of a turbine designer using the computational fluid dynamics (CFD) codes Numeca’s FINE/TURBO and the code TURBO. The predictions were all performed with only knowledge of the stage operating conditions, but without knowledge of the surface pressure measurements. Predictions were obtained with and without distributed cooling flow simulation. The FINE/TURBO model was run in 3-D viscous steady and time-accurate modes; the TURBO model was used to provide only 3-D viscous time-accurate results. Both FINE/TURBO and TURBO utilized phase-lagged boundary conditions to simplify the time-accurate model and to significantly reduce the computing time and resources. The time-accurate surface pressure loadings and steady state predictions are compared to measurements for the blade, vane, and shroud as time-averaged, time series, and power spectrum data. The measurements were obtained using The Ohio State University Gas Turbine Laboratory Turbine Test Facility. The time-average and steady comparisons of measurements and predictions are presented for 50% span on the vane and blade. Comparisons are also presented for several locations along the blade to illustrate local differences in the CFD behavior. The comparisons for the shroud are made across the blade passage at axial blade chord locations corresponding to the pressure transducer locations. The power spectrum decompositions of individual transducers (based on the fast Fourier transform (FFT)) are also included to lend insight into the unsteady nature of the flow. The comparisons show that both computational tools are capable of providing reasonable aerodynamic predictions for the vane, blade, and stationary shroud. The CFD model predictions show the encouraging trend of improved matching to the experimental data with increasing model fidelity from mass averaged to distributed cooling flow inclusion and as the codes change from steady to time-accurate modes.

Aeroperformance Measurements for a Fully Cooled High-Pressure Turbine Stage

2013 ◽  
Vol 136 (3) ◽  
Author(s):  
Charles Haldeman ◽  
Michael Dunn ◽  
Randall Mathison ◽  
William Troha ◽  
Timothy Vander Hoek ◽  
...  
Keyword(s):  
High Pressure ◽  
Mass Flow ◽  
Free Stream ◽  
Pressure Ratio ◽  
Specific Work ◽  
Gas Temperature ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Blade Tip ◽  
Tip Shroud

A detailed aero performance measurement program utilizing fully cooled engine hardware (high-pressure turbine stage) supplied by Honeywell Aerospace Advanced Technology Engines is described. The primary focus of this work was obtaining relevant aerodynamic data for a small turbine stage operating at a variety of conditions, including changes in operating conditions, geometry, and cooling parameters. The work extraction and the overall stage performance for each of these conditions can be determined using the measured acceleration rate of the turbine disk, the previously measured moment of inertia of the rotating system, and the mass flow through the turbine stage. Measurements were performed for two different values of tip/shroud clearance and two different blade tip configurations. The vane and blade cooling mass flow could be adjusted independently and set to any desired value, including totally off. A wide range of stage pressure ratios, coolant to free stream temperature ratios, and corrected speeds were used during the course of the investigation. A combustor emulator controlled the free stream inlet gas temperature, enabling variation of the temperature ratios and investigation of their effects on aero performance. The influence of the tip/shroud gap is clearly seen in this experiment. Improvements in specific work and efficiency achieved by reducing the tip/shroud clearance depend upon the specific values of stage pressure ratio and corrected speed. The maximum change of 3%–4% occurs at a stage pressure ratio and corrected speed greater than the initial design point intent. The specific work extraction and efficiency for two different blade tip sets (one damaged from a rub and one original) were compared in detail. In general, the tip damage only had a very small effect on the work extraction for comparable conditions. The specific work extraction and efficiency were influenced by the presence of cooling gas and by the temperature of the cooling gas relative to the free stream gas temperature and the metal temperature. These same parameters were influenced by the magnitude of the vane inlet gas total temperature relative to the vane metal temperature and the coolant gas temperature.

Aeroperformance Measurements for a Fully Cooled High-Pressure Turbine Stage

2012 ◽  
Author(s):  
Charles Haldeman ◽  
Michael Dunn ◽  
Randall Mathison ◽  
William Troha ◽  
Timothy Vander Hoek ◽  
...  
Keyword(s):  
High Pressure ◽  
Mass Flow ◽  
Free Stream ◽  
Pressure Ratio ◽  
Specific Work ◽  
Gas Temperature ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Blade Tip ◽  
Tip Shroud

A detailed aero performance measurement program utilizing fully cooled engine hardware (high-pressure turbine stage) supplied by Honeywell Aerospace Advanced Technology Engines is described. The primary focus of this work was obtaining relevant aerodynamic data for a small turbine stage operating at a variety of conditions, including changes in operating conditions, geometry, and cooling parameters. The work extraction and the overall stage performance for each of these conditions can be determined using the measured acceleration rate of the turbine disk, the previously measured moment of inertia of the rotating system, and the mass flow through the turbine stage. Measurements were performed for two different values of tip/shroud clearance and two different blade tip configurations. The vane and blade cooling mass flow could be adjusted independently and set to any desired value including totally off. A wide range of stage pressure ratios, coolant to freestream temperature ratios, and corrected speeds were used during the course of the investigation. A combustor emulator controlled the free stream inlet gas temperature, enabling variation of the temperature ratios and investigation of their effects on aero performance. The influence of tip/shroud gap is clearly seen in this experiment. Improvements in specific work and efficiency achieved by reducing the tip/shroud clearance depend upon the specific values of stage pressure ratio and corrected speed. The maximum change of 3% to 4% occurs at a stage pressure ratio and corrected speed greater than the initial design point intent. The specific work extraction and efficiency for two different blade tip sets (one damaged from a rub and one original) were compared in detail. In general, the tip damage only had a very small effect on the work extraction for comparable conditions. The specific work extraction and efficiency were influenced by the presence of cooling gas and by the temperature of the cooling gas relative to the free stream gas temperature and the metal temperature. These same parameters were influenced by the magnitude of the vane inlet gas total temperature relative to the vane metal temperature and the coolant gas temperature.

Comparison Between Unstructured and Structured Meshes With Different Turbulence Models for a High Pressure Turbine Application

2012 ◽  
Author(s):  
Jesuino Takachi Tomita ◽  
Lucilene Moraes da Silva ◽  
Diego Thomas da Silva
Keyword(s):  
High Pressure ◽  
Mesh Generation ◽  
Numerical Solutions ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Turbulence Models ◽  
Navier Stokes ◽  
Tip Leakage Flow ◽  
High Pressure Turbine ◽  
Navier Stokes Equations

For the CFD community the mesh generation is still one of the most important stages to obtain a good flow solution based on the full Navier-Stokes equations. For turbomachinery blade passages this task is not straightforward mainly due to the 3D domain and the complex geometries involved. The mesh quality and and elements distribution, orthogonality, smoothing, aspect ratio and angles are very important to guarantee a good numerical stability and solution accuracy. Moreover, the structure of the mesh inside the boundary-layer should be built carefully mainly in the regions where there are horseshoe vortices and tip leakage flow. In this work, the 3D turbulent flow is calculated and compared for structured and unstructured meshes including two equation models and Reynolds stress models. A high pressure turbine with 4.0 total-to-total pressure ratio is used in this study. A commercial software is used for mesh generation and flow calculation. The results are presented comparing the pressure ratio and efficiency from numerical solutions and experimental data and flow properties distributions along the blade span.

CFD Heat Transfer Predictions for a High-Pressure Turbine Stage

2004 ◽  
Author(s):  
James A. Tallman
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbulence Modeling ◽  
Pressure Ratio ◽  
Numerical Procedure ◽  
Navier Stokes ◽  
Turbine Stage ◽  
Temperature Ratio ◽  
High Pressure Turbine ◽  
Computational Fluid Dynamics Cfd

Computational Fluid Dynamics (CFD) was used to predict the turbine airfoil heat transfer for the high-pressure vane and high-pressure blade of a modern, one and one half stage turbine at its correct scale. Airfoil pressure and heat transfer measurements were recently obtained for the turbine in a transient shock tunnel facility, which allows for the replication of the actual engine turbine’s design corrected speed, pressure ratio, and gas-to-metal temperature ratio. A 3-D, compressible, Reynolds-averaged Navier-Stokes CFD solver with k-ω turbulence modeling was used for the CFD predictions. The turbulence model’s implementation into the numerical procedure was modified slightly, in order to better capture the model’s intended near-wall behavior and resolve the heat transfer prediction. Both the high-pressure vane and high-pressure blade were computed as steady-state flows and for two different turbine Reynolds number settings. Overall, the predictions compare very favorably with the measurement for both pressure and heat transfer at the mid-span location. A discussion of the features of the airfoil heat transfer distribution is included.

Developing a Combustor Simulator for Investigating High Pressure Turbine Aerodynamics and Heat Transfer

2004 ◽  
Author(s):  
M. D. Barringer ◽  
K. A. Thole ◽  
M. D. Polanka
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Test Facility ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Blow Down ◽  
Turbine Engine ◽  
The Impact

Within a gas turbine engine, the high pressure turbine vanes are subjected to very harsh conditions from the highly turbulent and hot gases exiting the combustor. The temperature and pressure fields exiting the combustor dictate the heat transfer and aero losses that occur in the turbine passages. To better understand these effects, the goal of this work is to develop an adjustable combustor exit profile simulator for the Turbine Research Facility (TRF) at the Air Force Research Laboratory (AFRL). The TRF is a high temperature, high pressure, short duration blow-down test facility that is capable of matching several aerodynamic and thermal non-dimensional engine parameters including Reynolds number, Mach number, pressure ratio, corrected mass flow, gas-to-metal temperature ratio, and corrected speed. The research objective was to design, install, and verify a non-reacting simulator device that provides representative combustor exit total pressure and temperature profiles to the inlet of the TRF turbine test section. This required the upstream section of the facility to be redesigned into multiple concentric annuli that serve the purpose of injecting high momentum dilution jets and low momentum film cooling jets into a central annular chamber, similar to a turbine engine combustor. The design of the simulator allows for variations in injection levels to generate turbulence and pressure profiles. It also can vary the dilution and film cooling temperatures to create a variety of temperature profiles consistent with real combustors. To date, the design and construction of the simulator device has been completed. All of the hardware has been trial fitted and the flow control shutter systems have been successfully installed and tested. Currently, verification testing is being performed to investigate the impact of the generated temperature, pressure, and turbulence profiles on turbine heat transfer and secondary flow development.

Heat Transfer Measurements and Predictions for a Modern, High-Pressure, Transonic Turbine, Including Endwalls

2006 ◽  
Author(s):  
James A. Tallman ◽  
Charles W. Haldeman ◽  
Michael G. Dunn ◽  
Anil K. Tolpadi ◽  
Robert F. Bergholz
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbulence Modeling ◽  
Pressure Ratio ◽  
Operating Conditions ◽  
Test Facility ◽  
Navier Stokes ◽  
Local Measurements ◽  
Transonic Turbine ◽  
The Ohio State University

This paper presents both measurements and predictions of the hot-gas-side heat transfer to a modern, one and 1/2 stage high-pressure, transonic turbine. Comparisons of the predicted and measured heat transfer are presented for each airfoil at three locations, as well as on the various endwalls and rotor tip. The measurements were performed using the Ohio State University Gas Turbine Laboratory Test Facility (TTF). The research program utilized an uncooled turbine stage at a range of operating conditions representative of the engine: in terms of corrected speed, flow function, stage pressure ratio, and gas-to-metal temperature ratio. All three airfoils were heavily instrumented for both pressure and heat transfer measurements at multiple locations. A 3-D, compressible, Reynolds-averaged Navier-Stokes CFD solver with k-ω turbulence modeling was used for the CFD predictions. The entire, 1-1/2 stage turbine was solved using a single computation, at two different Reynolds numbers. The CFD solutions were steady, with tangentially mass-averaged inlet/exit boundary condition profiles exchanged between adjacent airfoil-rows. Overall, the CFD heat transfer predictions compared very favorably with both the global operation of the turbine and with the local measurements of heat transfer. A discussion of the features of the turbine heat transfer distributions, and their association with the corresponding flow-physics, has been included.

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