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.

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.

High-Pressure Turbine Rim Seal Design for Increased Efficiency

2016 ◽  
Author(s):  
M. Chilla ◽  
H. P. Hodson ◽  
G. Pullan ◽  
D. Newman
Keyword(s):  
High Pressure ◽  
Numerical Simulations ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Main Stream ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Disc Space ◽  
Seal Design

In high-pressure turbines, compressor air is used to purge the disc space in an effort to protect the blade roots and the turbine disc from overheating and failure. The purge air exits the disc space through a rim seal at the hub of the main annulus and is subsequently entrained in the rotor hub endwall flows. The introduction of the purge air into the turbine main stream causes additional losses and therefore reduced turbine efficiency. For a given rim sealing mass flow rate, the rim seal geometry has to be designed in a way that reduces the detrimental impact of the sealing flow on turbine performance. In this study, the rim seal of a generic high-pressure turbine, representative of modern large civil aero-engines, is redesigned under consideration of the pressure field upstream of the rotor. Unsteady numerical simulations of the turbine stage are used to compare the aerodynamic impact of three different rim seal designs. The numerical simulations predict an increase in the time-averaged turbine stage efficiency of over 0.2% for the stage configuration with the final redesigned rim seal compared to the configuration with the original baseline rim seal geometry at the nominal sealing mass flow rate.

Heat-Transfer Measurements and Predictions for the Vane and Blade of a Rotating High-Pressure Turbine Stage

2004 ◽  
Vol 126 (1) ◽  
pp. 101-109 ◽  
Author(s):  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Three Dimensional ◽  
Experimental Results ◽  
Thermal Barrier ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Transfer Data ◽  
Commercial Code ◽  
Blade Tip

This paper describes heat-transfer measurements and predictions obtained for the vane and blade of a rotating high-pressure turbine stage. The measurements were obtained with the stage operating at design corrected conditions. A previous paper described the aerodynamics and the blade midspan location heat-transfer data and compared these experimental results with predictions. The intent of the current paper is to concentrate on the measurements and predictions for the 20%, 50%, and 80% span locations on the vane, the vane inner and outer endwall, the 20% and 96% span location on the blade, the blade tip (flat tip), and the stationary blade shroud. Heat-transfer data obtained at midspan for three different thermal-barrier-coated vanes (fine, medium, and coarse) are also presented. Boundary-layer heat-transfer predictions at the off-midspan locations are compared with the measurements for both the vane and the blade. The results of a STAR-CD (a commercial code) three-dimensional prediction are compared with the 20% and 96% span results for the blade surface. Predictions are not available for comparison with the tip and shroud experimental results.

Heat Transfer Measurements and Predictions for the Vane and Blade of a Rotating High-Pressure Turbine Stage

2003 ◽  
Author(s):  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
High Pressure ◽  
Experimental Results ◽  
Current Paper ◽  
Blade Surface ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Transfer Data ◽  
Blade Tip

This paper describes heat-transfer measurements and predictions obtained for the vane and blade of a rotating high-pressure turbine stage. The measurements were obtained with the stage operating at design corrected conditions. A previous paper described the aerodynamics and the blade midspan location heat-transfer data and compared these experimental results with predictions. The intent of the current paper is to concentrate on the measurements and predictions for the 20%, 50%, and 80% span locations on the vane, the vane inner and outer endwall, the 20% and 96% span location on the blade, the blade tip (flat tip), and the stationary blade shroud. Heat-transfer data obtained at midspan for three different TBC coated vanes (fine, medium and coarse) are also presented. Boundary-layer heat transfer predictions at the off-midspan locations are compared with the measurements for both the vane and the blade. The results of a STAR-CD 3D prediction are compared with the 20% and 96% span results for the blade surface. Predictions are not available for comparison with the tip and shroud experimental results.

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.

Unsteady winglet-cavity tip on leakage flow in a high-pressure turbine stage of low-aspect ratio

2017 ◽  
Vol 232 (20) ◽  
pp. 3708-3721 ◽  
Author(s):  
Zhihua Zhou ◽  
Shaowen Chen ◽  
Songtao Wang
Keyword(s):  
High Pressure ◽  
Aspect Ratio ◽  
Flow Characteristics ◽  
Leakage Flow ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Tip Leakage ◽  
Low Aspect Ratio ◽  
Passage Vortex ◽  
Blade Tip

In an unshrouded high-pressure turbine, the upstream vane wake, vane–blade interaction and blade tip leakage flow indicate complex and unsteady flow characteristics. Considering a high-pressure turbine stage of low-aspect ratio, the effects of the flat tip, cavity tip and winglet-cavity tip on the unsteady flow characteristics are investigated by numerical simulation. The exit Mach number and Reynolds number based on the chord of vane are 0.9 and 5.5×105, respectively. The pressure ratio of stage is 2.4. The time-resolved results indicate that the winglet-cavity tip scheme has smaller time variation of the total performance parameters and obtains a smaller tip leakage mass flow and a higher turbine stage efficient than the other cases. The vane–rotor interaction affects the pressure distribution at the region of blade leading edge remarkably and leads to a large time variation of tip leakage mass flow at the front of blade tip. The unsteady aerodynamic performance is analyzed with entropy-increase distribution at stage outlet and the vane–rotor interaction is discussed by the entropy-increase phase–time and phase–phase diagrams at blade outlet. Compared with the flat tip and cavity tip, the upper passage vortex loss is reduced obviously by the winglet-cavity tip. Thus, it is evaluated that the improvement of turbine stage coefficient with winglet-cavity tip results from the reduction of the upper passage vortex loss.

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.

A dynamic probabilistic design method for blade-tip radial running clearance of aeroengine high-pressure turbine

2014 ◽  
Vol 229 (10) ◽  
pp. 1861-1872 ◽  
Author(s):  
Cheng-Wei Fei ◽  
Wen-Zhong Tang ◽  
Guang-chen Bai ◽  
Zhi-Ying Chen
Keyword(s):  
High Pressure ◽  
Response Surface ◽  
Response Surface Method ◽  
Probabilistic Analysis ◽  
High Reliability ◽  
Probabilistic Design ◽  
High Pressure Turbine ◽  
Analysis And Design ◽  
Surface Method ◽  
Blade Tip

Around the engineering background of the probabilistic design of high-pressure turbine (HPT) blade-tip radial running clearance (BTRRC) which conduces to the high-performance and high-reliability of aeroengine, a distributed collaborative extremum response surface method (DCERSM) was proposed for the dynamic probabilistic analysis of turbomachinery. On the basis of investigating extremum response surface method (ERSM), the mathematical model of DCERSM was established. The DCERSM was applied to the dynamic probabilistic analysis of BTRRC. The results show that the blade-tip radial static clearance δ = 1.82 mm is advisable synthetically considering the reliability and efficiency of gas turbine. As revealed by the comparison of three methods (DCERSM, ERSM, and Monte Carlo method), the DCERSM reshapes the possibility of the probabilistic analysis for turbomachinery and improves the computational efficiency while preserving computational accuracy. The DCERSM offers a useful insight for BTRRC dynamic probabilistic analysis and optimization. The present study enrichs mechanical reliability analysis and design theory.

Aerothermal Effect of Cavity Welding Beads on a Transonic Squealer Tip

2020 ◽  
Author(s):  
Joao Vieira ◽  
John Coull ◽  
Peter Ireland ◽  
Eduardo Romero
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Reynolds Numbers ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Blade Tip ◽  
Mass Flow Rates

Abstract High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

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

2012 ◽  
Author(s):  
Brian R. Green ◽  
Randall M. Mathison ◽  
Michael G. Dunn
Keyword(s):  
High Pressure ◽  
Flow Field ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Unsteady Aerodynamics ◽  
Flow Path ◽  
Field Analysis ◽  
High Pressure Turbine ◽  
Purge Flow

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 as well as 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 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 to measure 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 CFD code that had the capability to perform both 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 flow field analysis that will be provided in Part II.

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