An Experimental Study of Combustor Exit Profile Shapes on Endwall Heat Transfer in High Pressure Turbine Vanes

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
M. D. Barringer ◽  
K. A. Thole ◽  
M. D. Polanka
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Elevated Temperatures ◽  
Convection Heat Transfer ◽  
Operating Conditions ◽  
Inlet Temperature ◽  
Transfer Coefficients ◽  
High Pressure Turbine ◽  
Radial Profiles ◽  
Temperature And Pressure

The design and development of current and future gas turbine engines for aircraft propulsion have focused on operating the high pressure turbine at increasingly elevated temperatures and pressures. The drive toward thermal operating conditions near theoretical stoichiometric limits as well as increasingly stringent requirements on reducing harmful emissions both equate to the temperature profiles exiting combustors and entering turbines becoming less peaked than in the past. This drive has placed emphasis on determining how different types of inlet temperature and pressure profiles affect the first stage airfoil endwalls. The goal of the current study was to investigate how different radial profiles of temperature and pressure affect the heat transfer along the vane endwall in a high pressure turbine. Testing was performed in the Turbine Research Facility located at the Air Force Research Laboratory using an inlet profile generator. Results indicate that the convection heat transfer coefficients are influenced by both the inlet pressure profile shape and the location along the endwall. The heat transfer driving temperature for inlet profiles that are nonuniform in temperature is also discussed.

An Experimental Study of Combustor Exit Profile Shapes on Endwall Heat Transfer in High Pressure Turbine Vanes

2007 ◽  
Author(s):  
M. D. Barringer ◽  
M. D. Polanka ◽  
K. A. Thole
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Elevated Temperatures ◽  
Convection Heat Transfer ◽  
Operating Conditions ◽  
Inlet Temperature ◽  
Transfer Coefficients ◽  
High Pressure Turbine ◽  
Radial Profiles ◽  
Temperature And Pressure

The design and development of current and future gas turbine engines for aircraft propulsion have focused on operating the high pressure turbine at increasingly elevated temperatures and pressures. The drive towards thermal operating conditions near theoretical stoichiometric limits as well as increasingly stringent requirements on reducing harmful emissions, both equate to the temperature profiles exiting combustors and entering turbines becoming less peaked than in the past. This drive has placed emphasis on determining how different types of inlet temperature and pressure profiles affect the first stage airfoil endwalls. The goal of the current study was to investigate how different radial profiles of temperature and pressure affect the heat transfer along the vane endwall in a high pressure turbine. Testing was performed in the Turbine Research Facility located at the Air Force Research Laboratory using an inlet profile generator. Results indicate that the convection heat transfer coefficients are influenced by both the inlet pressure profile shape and the location along the endwall. The heat transfer driving temperature for inlet profiles that are non-uniform in temperature is also discussed.

Heat transfer characteristic modelling and the effect of operating conditions on re-cooled cycle for a scramjet

2011 ◽  
Vol 115 (1164) ◽  
pp. 83-90 ◽  
Author(s):  
W. Bao ◽  
J. Qin ◽  
W. X. Zhou
Keyword(s):  
Heat Transfer ◽  
Heat Sink ◽  
Transfer Characteristic ◽  
Cooling Capacity ◽  
Sensitivity Study ◽  
Performance Model ◽  
Operating Conditions ◽  
Inlet Temperature ◽  
Fuel Flow ◽  
Temperature And Pressure

Abstract A re-cooled cycle has been proposed for a regeneratively cooled scramjet to reduce the hydrogen fuel flow for cooling. Upon the completion of the first cooling, fuel can be used for secondary cooling by transferring the enthalpy from fuel to work. Fuel heat sink (cooling capacity) is thus repeatedly used and fuel heat sink is indirectly increased. Instead of carrying excess fuel for cooling or seeking for any new coolant, the cooling fuel flow is reduced, and fuel onboard is adequate to satisfy the cooling requirement for the whole hypersonic vehicle. A performance model considering flow and heat transfer is build. A model sensitivity study of inlet temperature and pressure reveals that, for given exterior heating condition and cooling panel size, fuel heat sink can be obviously increased at moderate inlet temperature and pressure. Simultaneously the low-temperature heat transfer deterioration and Mach number constrains can also be avoided.

THE INFLUENCE OF PURGE FLOW PARAMETERS ON HEAT TRANSFER AND FILM COOLING IN TURBINE CENTER FRAMES

2021 ◽  
pp. 1-26
Author(s):  
Patrick René Jagerhofer ◽  
Marios Patinios ◽  
Tobias Glasenapp ◽  
Emil Goettlich ◽  
Federica Farisco
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Density Ratio ◽  
Constant Heat Flux ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Turbofan Engines ◽  
Blowing Ratio ◽  
Purge Flow

Abstract The imperative improvement in the efficiency of turbofan engines is commonly facilitated by increasing the turbine inlet temperature. This development has reached a point where also components downstream of the high-pressure turbine have to be adequately cooled. Such a component is the turbine center frame (TCF), known for a complex aerodynamic flow highly influenced by purge-mainstream interactions. The purge air, being injected through the wheelspace cavities of the upstream high-pressure turbine, bears a significant cooling potential for the TCF. Despite this, fundamental knowledge of the influencing parameters on heat transfer and film cooling in the TCF is still missing. This paper examines the influence of purge-to-mainstream blowing ratio, density ratio and purge swirl angle on heat transfer and film cooling in the TCF. The experiments are conducted in a sector-cascade test rig specifically designed for such heat transfer studies using infrared thermography and tailor-made flexible heating foils with constant heat flux. Three purge-to-mainstream blowing ratios and an additional no purge case are investigated. The purge flow is injected without swirl and also with engine-similar swirl angles. The purge swirl and blowing ratio significantly impact the magnitude and the spread of film cooling in the TCF. Increasing blowing ratios lead to an intensification of heat transfer. By cooling the purge flow, a moderate variation in purge-to-mainstream density ratio is investigated, and the influence is found to be negligible.

Numerical Prediction of Cooling Performance Sensitivity of 1st Stage Nozzle Guide Vane Under Aerothermal Conditions

2019 ◽  
Author(s):  
Prasert Prapamonthon ◽  
Bo Yin ◽  
Guowei Yang ◽  
Mohan Zhang
Keyword(s):  
Heat Transfer ◽  
Surface Roughness ◽  
High Pressure ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Cooling Performance ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract To obtain high power and thermal efficiency, the 1st stage nozzle guide vanes of a high-pressure turbine need to operate under serious circumstances from burned gas coming out of combustors. This leads to vane suffering from effects of high thermal load, high pressure and turbulence, including flow-separated transition. Therefore, it is necessary to improve vane cooling performance under complex flow and heat transfer phenomena caused by the integration of these effects. In fact, these effects on a high-pressure turbine vane are controlled by several factors such as turbine inlet temperature, pressure ratio, turbulence intensity and length scale, vane curvature and surface roughness. Furthermore, if the vane is cooled by film cooling, hole configuration and blowing ratio are important factors too. These factors can change the aerothermal conditions of the vane operation. The present work aims to numerically predict sensitivity of cooling performances of the 1st stage nozzle guide vane under aerodynamic and thermal variations caused by three parameters i.e. pressure ratio, coolant inlet temperature and height of vane surface roughness using Computational Fluid Dynamics (CFD) with Conjugate Heat Transfer (CHT) approach. Numerical results show that the coolant inlet temperature and the vane surface roughness parameters have significant effects on the vane temperature, thereby affecting the vane cooling performances significantly and sensitively.

Transient Forced Convection Heat Transfer to Helium During a Step in Heat Flux

1983 ◽  
Vol 105 (2) ◽  
pp. 350-357 ◽  
Author(s):  
P. J. Giarratano ◽  
W. G. Steward
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Forced Convection ◽  
Carbon Film ◽  
Convection Heat Transfer ◽  
Fast Response ◽  
Operating Conditions ◽  
Convection Heat ◽  
Transfer Coefficients ◽  
Forced Convection Heat Transfer

Transient forced convection heat transfer coefficients for both subcritical and supercritical helium in a rectangular flow channel heated on one side were measured during the application of a step in heat flux. Zero flow data were also obtained. The heater surface which served simultaneously as a thermometer was a fast response carbon film. Operating conditions covered the following range: Pressure, 1.0 × 105 Pa (1 bar) to 1.0 × 106 Pa (10 bar); Temperature, 4 K–10 K; Heat Flux, 0.1 W/cm2−10 W/cm2; Reynolds number, 0–8 × 105. The experimental data and a predictive correlation are presented.

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.

Aerodynamics and Heat Transfer for a Cooled One and One-Half Stage High-Pressure Turbine—Part I: Vane Inlet Temperature Profile Generation and Migration

2011 ◽  
Vol 134 (1) ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Temperature Profile ◽  
Film Cooling ◽  
Leading Edge ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Fluid Temperature ◽  
High Pressure Turbine ◽  
Wide Range

As controlled laboratory experiments using full-stage turbines are expanded to replicate more of the complicated flow features associated with real engines, it is important to understand the influence of the vane inlet temperature profile on the high-pressure vane and blade heat transfer as well as its interaction with film cooling. The temperature distribution of the incoming fluid governs not only the input conditions to the boundary layer but also the overall fluid migration. Both of these mechanisms have a strong influence on surface heat flux and therefore component life predictions. To better understand the role of the inlet temperature profile, an electrically heated combustor emulator capable of generating uniform, radial, or hot streak temperature profiles at the high-pressure turbine vane inlet has been designed, constructed, and operated over a wide range of conditions. The device is shown to introduce a negligible pressure distortion while generating the inlet temperature conditions for a stage-and-a-half turbine operating at design-corrected conditions. For the measurements described here, the vane is fully cooled and the rotor purge flow is active, but the blades are uncooled. Detailed temperature measurements are obtained at rake locations upstream and downstream of the turbine stage as well as at the leading edge and platform of the blade in order to characterize the inlet temperature profile and its migration. The use of miniature butt-welded thermocouples at the leading edge and on the platform (protruding into the flow) on a rotating blade is a novel method of mapping a temperature profile. These measurements show that the reduction in fluid temperature due to cooling is similar in magnitude for both uniform and radial vane inlet temperature profiles.

Design and Optimization of the Internal Cooling Channels of a High Pressure Turbine Blade—Part I: Methodology

2010 ◽  
Vol 132 (2) ◽  
Author(s):  
Sergio Amaral ◽  
Tom Verstraete ◽  
René Van den Braembussche ◽  
Tony Arts
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Thermal Stress ◽  
Turbine Blade ◽  
Inlet Temperature ◽  
Analysis Tool ◽  
Internal Cooling ◽  
High Pressure Turbine ◽  
Cooling Channels ◽  
The Impact

This first paper describes the conjugate heat transfer (CHT) method and its application to the performance and lifetime prediction of a high pressure turbine blade operating at a very high inlet temperature. It is the analysis tool for the aerothermal optimization described in a second paper. The CHT method uses three separate solvers: a Navier–Stokes solver to predict the nonadiabatic external flow and heat flux, a finite element analysis (FEA) to compute the heat conduction and stress within the solid, and a 1D aerothermal model based on friction and heat transfer correlations for smooth and rib-roughened cooling channels. Special attention is given to the boundary conditions linking these solvers and to the stability of the complete CHT calculation procedure. The Larson–Miller parameter model is used to determine the creep-to-rupture failure lifetime of the blade. This model requires both the temperature and thermal stress inside the blade, calculated by the CHT and FEA. The CHT method is validated on two test cases: a gas turbine rotor blade without cooling and one with five cooling channels evenly distributed along the camber line. The metal temperature and thermal stress distribution in both blades are presented and the impact of the cooling channel geometry on lifetime is discussed.

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

2011 ◽  
Vol 134 (3) ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Suction Surface ◽  
Adiabatic Wall ◽  
Cooling Flows ◽  
Cooling Flow

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.

Evaluation of the External Heat Transfer Coefficients in the High-Pressure Turbine of a Full-Scale Core Engine

1996 ◽  
Author(s):  
A. R. Narcus ◽  
H. R. Przirembel ◽  
F. O. Soechting
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Data Reduction ◽  
Turbine Blades ◽  
External Heat ◽  
Transfer Coefficients ◽  
High Pressure Turbine ◽  
External Heat Transfer ◽  
Heat Transfer Coefficients

The external heat transfer coefficients, necessary for efficient and accurate turbine blade design, have been quantified using three independent methods of data reduction for the high-pressure turbine blades tested in a core engine. Two of the methods utilized external and internal thermocouple data to determine the heat transfer coefficient levels while the third method required the applied heat-flux levels to determine the coefficients. The heat-flux was calculated from the measured potential difference between thermocouple pairs embedded in the external and internal walls of the turbine blades. The instrumented airfoils were calibrated in a laboratory prior to engine testing. The results of the experimental test showed external heat transfer coefficients could be obtained in an engine environment with a ±3.2% minimum absolute uncertainty. All three data reduction methods produced external heat transfer coefficients within a high degree of accuracy and precision for all data locations on the instrumented airfoils. The three data reduction approaches are presented as well as the data for a specific location on a turbine blade for each method of data reduction. In addition, pre-test calibration procedures and data are discussed along with supporting engine instrumentation used to verify the data acquired during the experimental evaluation.

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