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.

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.

Comparison of Steady and Unsteady Coupled Heat-Transfer Simulations of a High-Pressure Turbine Blade

2015 ◽  
Author(s):  
Markus Schmidt ◽  
Christoph Starke
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Steady State ◽  
Flow Field ◽  
Film Cooling ◽  
Strong Increase ◽  
Pressure Side ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Time Resolved

This article presents results for the coupled simulation of a high-pressure turbine stage in consideration of unsteady hot gas flows. A semi-unsteady coupling process was developed to solve the conjugate heat transfer problem for turbine components of gas turbines. Time-resolved CFD simulations are coupled to a finite element solver for the steady state heat conduction inside of the blade material. A simplified turbine stage geometry is investigated in this paper to describe the influence of the unsteady flow field onto the time-averaged heat transfer. Comparisons of the time-resolved results to steady state results indicate the importance of a coupled simulation and the consideration of the time-dependent flow-field. Different film-cooling configurations for the turbine NGV are considered, resulting in different temperature and pressure deficits in the vane wake. Their contribution to non-linear effects causing the time-averaged heat load to differ from a steady result is discussed to further highlight the necessity of unsteady design methods for future turbine developments. A strong increase in the pressure side heat transfer coefficients for unsteady simulations is observed in all results. For higher film-cooling mass flows in the upstream row, the preferential migration of hot fluid towards the pressure side of a turbine blade is amplified as well, which leads to a strong increase in material temperature at the pressure side and also in the blade tip region.

Time-Averaged and Time-Accurate Aero-Dynamics for the Recessed Tip Cavity of a High-Pressure Turbine Blade and the Outer Stationary Shroud: Comparison of Computational and Experimental Results

2004 ◽  
Author(s):  
Brian R. Green ◽  
John W. Barter ◽  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Computational Models ◽  
Turbine Rotor ◽  
Single Stage ◽  
Cfd Modeling ◽  
Blade Surface ◽  
High Pressure Turbine ◽  
Pressure Transducers ◽  
Blade Tip

The unsteady aero-dynamics of a single-stage high-pressure turbine blade operating at design corrected conditions has been the subject of a thorough study involving detailed measurements and computations. The experimental configuration consisted of a single-stage high-pressure turbine and the adjacent, downstream, low-pressure turbine nozzle row. All three blade-rows were instrumented at three spanwise locations with flush-mounted, high frequency response pressure transducers. The rotor was also instrumented with the same transducers on the blade tip and platform and the stationary shroud was instrumented with pressure transducers at specific locations above the rotating blade. Predictions of the time-dependent flow field around the rotor were obtained using MSU-TURBO, a 3D, non-linear, computational fluid dynamics (CFD) code. Using an isolated blade-row unsteady analysis method, the unsteady surface pressure for the high-pressure turbine rotor due to the upstream high-pressure turbine nozzle was calculated. The predicted unsteady pressure on the rotor surface was compared to the measurements at selected spanwise locations on the blade, in the recessed cavity, and on the shroud. The rig and computational models included a flat and recessed blade tip geometry and were used for the comparisons presented in the paper. Comparisons of the measured and predicted static pressure loading on the blade surface show excellent correlation from both a time-average and time-accurate standpoint. This paper concentrates on the tip and shroud comparisons between the experiments and the predictions and these results also show good correlation with the time-resolved data. These data comparisons provide confidence in the CFD modeling and its ability to capture unsteady flow physics on the blade surface, in the flat and recessed tip regions of the blade, and on the stationary shroud.

Three Dimensional Heat Transfer Analysis of High Pressure Turbine Blade

2009 ◽  
Author(s):  
A. Sipatov ◽  
L. Gomzikov ◽  
V. Latyshev ◽  
N. Gladysheva
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
High Pressure ◽  
Rotor Blade ◽  
Computational Analysis ◽  
Three Dimensional ◽  
Heat Transfer Analysis ◽  
Aircraft Engines ◽  
Turbine Stage ◽  
High Pressure Turbine

The present tendency of creating new aircraft engines with a higher level of fuel efficiency leads to the necessity to increase gas temperature at a high pressure turbine (HPT) inlet. To design such type of engines, the improvement of accuracy of the computational analysis is required. According to this the numerical analysis methods are constantly developing worldwide. The leading firms in designing aircraft engines carry out investigations in this field. However, this problem has not been resolved completely yet because there are many different factors affecting HPT blade heat conditions. In addition in some cases the numerical methods and approaches require tuning (for example to predict laminar-turbulent transition region or to describe the interaction of boundary layer and shock wave). In this work our advanced approach of blade heat condition numerical estimation based on the three-dimensional computational analysis is presented. The object of investigation is an advanced aircraft engine HPT first stage blade. The given analysis consists of two interrelated parts. The first part is a stator-rotor interaction modeling of the investigated turbine stage (unsteady approach). Solving this task we devoted much attention to modeling unsteady effects of stator-rotor interaction and to describing an influence of applied inlet boundary conditions on the blade heat conditions. In particular, to determine the total pressure, flow angle and total temperature distributions at the stage inlet we performed a numerical modeling of the combustor chamber of the investigated engine. The second part is a flow modeling in the turbine stage using flow parameters averaging on the stator-rotor interface (steady approach). Here we used sufficiently finer grid discretization to model all perforation holes on the stator vane and rotor blade, endwalls films in detail and to apply conjugate heat transfer approach for the rotor blade. Final results were obtained applying the results of steady and unsteady approaches. Experimental data of the investigated blade heat conditions are presented in the paper. These data were obtained during full size experimental testing the core of the engine and were collected using two different type of experimental equipment: thermocouples and thermo-crystals. The comparison of experimental data and final results meets the requirements of our investigation.

Mesh Generation for Conjugate Heat Transfer Analysis of a Cooled High Pressure Turbine Stage

2008 ◽  
Author(s):  
Colinda Goormans-Francke ◽  
Guy Carabin ◽  
Charles Hirsch
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Mesh Generation ◽  
Conjugate Heat Transfer ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Flow Solver ◽  
Rib Turbulators ◽  
Structured Mesh ◽  
Grid Points

The presented work demonstrates the feasibility of quasi-automatic structured mesh generation for all details in the complex cooling system of an industrial high pressure turbine stage, as required by advanced Conjugate Heat Transfer (CHT) simulations. The grid generation software has been adapted in order to quasi-automatically mesh typical cooling configurations such as cooling passages, basins, inserts, solid bodies, cooling holes, slots, and rib turbulators. A multi-domain structured mesh with about 154 million grid points and 12,316 blocks has been generated for the turbine stage. It includes 1,000 cooling holes, over 250 rib turbulators and 150 pin fins for the turbine stage. In order to verify the CFD response to the grid properties, simulations were performed as a first step on the coarse grid level (of 21.8 million grid points) using the 3D flow solver package FINE™/Turbo. The conductivity equation was solved for the solid part of the computational domain using the same temporal discretization scheme as for the flow solver. Parallel, coupled fluid/solid calculations using the k-ε turbulence model were performed on three different configurations: nozzle guide vane alone, rotor-blade alone, and full stage. These results show the feasibility of this approach to mesh generation for use in CHT modeling of the complex configuration of cooled turbine stages.

Averaged and Time-Dependent Aerodynamics of a High Pressure Turbine Blade Tip Cavity and Stationary Shroud: Comparison of Computational and Experimental Results

2004 ◽  
Vol 127 (4) ◽  
pp. 736-746 ◽  
Author(s):  
Brian R. Green ◽  
John W. Barter ◽  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Computational Models ◽  
Single Stage ◽  
Time Dependent ◽  
Cfd Modeling ◽  
Blade Surface ◽  
High Pressure Turbine ◽  
Pressure Transducers ◽  
Blade Tip

The unsteady aero-dynamics of a single-stage high-pressure turbine blade operating at design corrected conditions has been the subject of a thorough study involving detailed measurements and computations. The experimental configuration consisted of a single-stage high-pressure turbine and the adjacent, downstream, low-pressure turbine nozzle row. All three blade-rows were instrumented at three spanwise locations with flush-mounted, high-frequency response pressure transducers. The rotor was also instrumented with the same transducers on the blade tip and platform and the stationary shroud was instrumented with pressure transducers at specific locations above the rotating blade. Predictions of the time-dependent flow field around the rotor were obtained using MSU-TURBO, a three-dimensional (3D), nonlinear, computational fluid dynamics (CFD) code. Using an isolated blade-row unsteady analysis method, the unsteady surface pressure for the high-pressure turbine rotor due to the upstream high-pressure turbine nozzle was calculated. The predicted unsteady pressure on the rotor surface was compared to the measurements at selected spanwise locations on the blade, in the recessed cavity, and on the shroud. The rig and computational models included a flat and recessed blade tip geometry and were used for the comparisons presented in the paper. Comparisons of the measured and predicted static pressure loading on the blade surface show excellent correlation from both a time-average and time-accurate standpoint. This paper concentrates on the tip and shroud comparisons between the experiments and the predictions and these results also show good correlation with the time-resolved data. These data comparisons provide confidence in the CFD modeling and its ability to capture unsteady flow physics on the blade surface, in the flat and recessed tip regions of the blade, and on the stationary shroud.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2011 ◽  
Vol 134 (4) ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer, and film cooling effectiveness of advanced high pressure turbine blade tips and endwalls. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with leading edge and trailing edge cutouts. Both blade tip configurations have pressure side film cooling and cooling air extraction through dust holes, which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9×105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aerothermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the midchord region. However, on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

scholarly journals Computational Modeling and Thermal Paint Verification of Film-Cooling Designs for an Unshrouded High-Pressure Turbine Blade

1999 ◽  
Author(s):  
Steven G. Gegg ◽  
Nathan J. Heidegger ◽  
Ronald A. Mikkelson
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbine Blade ◽  
Film Cooling ◽  
Data Exchange ◽  
Turbine Blades ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Extreme High Temperature ◽  
Blade Tip

High pressure turbine blades are exposed to an extreme high temperature environment due to increasing turbine inlet temperature. High heat fluxes are likely on the blade pressure surface. Other regions, such as the trailing edge and blade tip may be difficult to cool uniformly. Unshrouded blades present an additional challenge due to the pressure driven transport of hot gas across the blade tip. The blade tip region is therefore prone to severe thermal stress, fatigue and oxidation. In order to develop effective cooling methods, designers require detailed flow and heat transfer information. This paper reports on computational aerodynamics and heat transfer studies for an unshrouded high pressure turbine blade. The emphasis is placed on the application of appropriate 3-D models for the prediction of airfoil surface temperatures. Details of the film cooling model, boundary conditions and data exchange with heat transfer models are described. The analysis approach has been refined for design use to provide timely and accurate results. Film cooling designs are to be tailored for the best coverage of the blade tip region. Designs include near-tip pressure side films and blade tip cooling holes. Hole placement and angle are investigated to achieve the best coolant coverage on the blade tip. Analytical results are compared to a thermal paint test on engine hardware. In addition to film cooling strategies, other aerodynamic/heat transfer design approaches are discussed to address the cooling requirements for an unshrouded blade.

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.

Optimization of a high pressure turbine blade tip cavity with conjugate heat transfer analysis

2016 ◽  
Vol 30 (12) ◽  
pp. 5529-5538 ◽  
Author(s):  
Jinuk Kim ◽  
Young Seok Kang ◽  
Dongwha Kim ◽  
Jihyeong Lee ◽  
Bong Jun Cha ◽  
...  
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbine Blade ◽  
Conjugate Heat Transfer ◽  
Heat Transfer Analysis ◽  
High Pressure Turbine ◽  
Blade Tip
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