Gas Turbine Engine Durability Impacts of High Fuel-Air Ratio Combustors: Part 2 — Near Wall Reaction Effects on Film-Cooled Heat Transfer

2002 ◽  
Author(s):  
Daniel R. Kirk ◽  
Gerald R. Guenette ◽  
Stephen P. Lukachko ◽  
Ian A. Waitz
Keyword(s):  
Heat Flux ◽  
Gas Turbine ◽  
Chemical Reactions ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Gas Turbine Engine ◽  
Surface Heat ◽  
Turbine Engine ◽  
Secondary Reactions ◽  
Near Wall

As commercial and military aircraft engines approach higher total temperatures and increasing overall fuel-to-air ratios, the potential for significant chemical reactions on a film-cooled surface is enhanced. Currently there is little basis for understanding the effects on aero-performance and durability due to such secondary reactions. A shock tube experiment was employed to generate short duration, high temperature (1000–2800 K) and pressure (6 atm.) flows over a film-cooled flat plate. The test plate contained two sets of 35° film cooling holes that could be supplied with different gases, one side using air and the other nitrogen. A mixture of ethylene and argon provided a fuel rich freestream that reacted with the air film resulting in near wall reactions. The relative increase in surface heat flux due to near wall reactions was investigated over a range of fuel levels, momentum blowing ratios (0.5–2.0), and Damko¨hler numbers (ratio of flow to chemical time scales) from near zero to 30. For high Damko¨hler numbers, reactions had sufficient time to occur and increased the surface heat flux by 30 percent over the inert cooling side. When these results are appropriately scaled, it is shown that in some situations of interest for gas turbine engine environments significant increases in surface heat flux can be produced due to chemical reactions in the film-cooling layer. It is also shown that the non-dimensional parameters Damko¨hler number (Da), blowing ratio (B), heat release potential (H*), and scaled heat flux (Qs) are the appropriate quantities to predict the augmentation in surface heat flux that arises due to secondary reactions.

Gas Turbine Engine Durability Impacts of High Fuel-Air Ratio Combustors—Part II: Near-Wall Reaction Effects on Film-Cooled Heat Transfer

2003 ◽  
Vol 125 (3) ◽  
pp. 751-759 ◽  
Author(s):  
D. R. Kirk ◽  
G. R. Guenette ◽  
S. P. Lukachko ◽  
I. A. Waitz
Keyword(s):  
Heat Flux ◽  
Gas Turbine ◽  
Chemical Reactions ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Gas Turbine Engine ◽  
Surface Heat ◽  
Turbine Engine ◽  
Secondary Reactions ◽  
Near Wall

As commercial and military aircraft engines approach higher total temperatures and increasing overall fuel-to-air ratios, the potential for significant chemical reactions on a film-cooled surface is enhanced. Currently, there is little basis for understanding the effects on aero-performance and durability due to such secondary reactions. A shock tube experiment was employed to generate short duration, high temperature (1000–2800 K) and pressure (6 atm) flows over a film-cooled flat plate. The test plate contained two sets of 35 deg film cooling holes that could be supplied with different gases, one side using air and the other nitrogen. A mixture of ethylene and argon provided a fuel rich freestream that reacted with the air film resulting in near wall reactions. The relative increase in surface heat flux due to near wall reactions was investigated over a range of fuel levels, momentum blowing ratios (0.5–2.0), and Damko¨hler numbers (ratio of flow to chemical time scales) from near zero to 30. For high Damko¨hler numbers, reactions had sufficient time to occur and increased the surface heat flux by 30 percent over the inert cooling side. When these results are appropriately scaled, it is shown that in some situations of interest for gas turbine engine environments significant increases in surface heat flux can be produced due to chemical reactions in the film-cooling layer. It is also shown that the non-dimensional parameters Damko¨hler number (Da), blowing ratio (B), heat release potential (H*), and scaled heat flux Qs are the appropriate quantities to predict the augmentation in surface heat flux that arises due to secondary reactions.

Gas Turbine Engine Durability Impacts of High Fuel-Air Ratio Combustors: Near Wall Reaction Effects on Film-Cooled Backward-Facing Step Heat Transfer

2004 ◽  
Vol 128 (2) ◽  
pp. 318-325 ◽  
Author(s):  
David W. Milanes ◽  
Daniel R. Kirk ◽  
Krzysztof J. Fidkowski ◽  
Ian A. Waitz
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Chemical Reactions ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Surface Heat ◽  
Backward Facing Step ◽  
Blowing Ratio ◽  
Film Cooling Holes ◽  
Near Wall

As commercial and military aircraft engines approach higher total temperatures and increasing overall fuel-to-air ratios, the potential for significant chemical reactions to occur downstream of the combustor is increased. This may take place when partially reacted species leave the combustor and encounter film-cooled surfaces. One common feature on turbine endwalls is a step between various engine components and seals. Such step features produce recirculating flows which when in the vicinity of film-cooled surfaces may lead to particularly severe reaction zones due to long fluid residence times. The objective of this paper is to study and quantify the surface heat transfer implications of such reacting regions. A shock tube experiment was employed to generate short duration, high temperature (1000–2800 K) and pressure (6 atm) flows over a film-cooled backward-facing step. The test article contained two sets of 35 deg film cooling holes located downstream of a step. The film-cooling holes could be supplied with different gases, one side using air and the other nitrogen allowing for simultaneous testing of reacting and inert cooling gases. A mixture of ethylene and argon provided a fuel-rich free stream that reacted with the air film resulting in near wall reactions. The relative increase in surface heat flux due to near wall reactions was investigated over a range of fuel levels, momentum blowing ratios (0.5–2.0), and Damköhler numbers (ratio of characteristic flow time to chemical time) from near zero to 30. The experimental results show that for conditions relevant for future engine technology, adiabatic flame temperatures can be approached along the wall downstream of the step leading to potentially significant increases in surface heat flux. A computational study was also performed to investigate the effects of cooling-jet blowing ratio on chemical reactions behind the film-cooled step. The blowing ratio was found to be an important parameter governing the flow structure behind the backward-facing step, and controlling the characteristics of chemical-reactions by altering the local equivalence ratio.

Gas Turbine Engine Durability Impacts of High Fuel-Air Ratio Combustors: Near Wall Reaction Effects on Film-Cooled Backward-Facing Step Heat Transfer

2004 ◽  
Author(s):  
David W. Milanes ◽  
Daniel R. Kirk ◽  
Krzysztof J. Fidkowski ◽  
Ian A. Waitz
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Chemical Reactions ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Surface Heat ◽  
Backward Facing Step ◽  
Blowing Ratio ◽  
Film Cooling Holes ◽  
Near Wall

As commercial and military aircraft engines approach higher total temperatures and increasing overall fuel-to-air ratios, the potential for significant chemical reactions to occur downstream of the combustor is increased. This may take place when partially-reacted species leave the combustor and encounter film-cooled surfaces. One common feature on turbine endwalls is a step between various engine components and seals. Such step features produce recirculating flows which when in the vicinity of film-cooled surfaces may lead to particularly severe reaction zones due to long fluid residence times. The objective of this paper is to study and quantify the surface heat transfer implications of such reacting regions. A shock tube experiment was employed to generate short duration, high temperature (1000–2800K) and pressure (6 atm.) flows over a film-cooled backward-facing step. The test article contained two sets of 35° film cooling holes located downstream of a step. The film-cooling holes could be supplied with different gases, one side using air and the other nitrogen allowing for simultaneous testing of reacting and inert cooling gases. A mixture of ethylene and argon provided a fuel rich freestream that reacted with the air film resulting in near wall reactions. The relative increase in surface heat flux due to near wall reactions was investigated over a range of fuel levels, momentum blowing ratios (0.5–2.0), and Damko¨hler numbers (ratio of characteristic flow time to chemical time) from near zero to 30. The experimental results show that for conditions relevant for future engine technology, adiabatic flame temperatures can be approached along the wall downstream of the step leading to potentially significant increases in surface heat flux. A computational study was also performed to investigate the effects of cooling-jet blowing ratio on chemical reactions behind the film-cooled step. The blowing ratio was found to be an important parameter governing the flow structure behind the backward-facing step, and controlling the characteristics of chemical-reactions by altering the local equivalence ratio.

3D Heat Transfer Assessment of Full-Scale Inlet Vanes With Surface-Optimized Film Cooling: Part 1 — Experimental Results

2019 ◽  
Author(s):  
R. J. Anthony ◽  
J. P. Clark ◽  
J. Finnegan ◽  
J. J. Johnson
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Frequency ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Surface Heat ◽  
Full Scale ◽  
Pressure Side ◽  
Flux Measurements ◽  
Real Hardware

Abstract Full-scale annular experimental evaluation of two different high pressure turbine first stage vane cooling designs was carried out using high frequency surface heat-flux measurements in the Turbine Research Facility at the Air Force Research Laboratory. A baseline film cooling geometry was tested simultaneously with a genetically optimized vane aimed to improve efficiency and part life. Part 1 of this two-part paper describes the experimental instrumentation, test facility, and surface heat flux measurements used to evaluate both cooling schemes. Part 2 of this paper describes the result of companion conjugate heat transfer posttest predictions, and gives numerical background on the design and modelling of both film cooling geometries. Time-resolved surface heat flux data is captured at multiple airfoil span and chord locations for each cooling design. Area based assessment of surface flux data verifies the genetic optimization redistributes excessive cooling away from midspan areas to improve efficiency. Results further reveal key discrepancies between design intent and real hardware behavior. Elevated heat flux above intent in some areas led to investigation of backflow margin and unsteady hot gas ingestion at certain film holes. Analysis shows areas toward the vane inner and outer endwalls of the aft pressure side were more sensitive to reduced aft cavity backflow margin. In addition, temporal analysis shows film cooled heat flux having large high frequency fluctuations that can vary across nearly the full range of film cooling effectiveness at some locations. Velocity and acceleration of these large unsteady heat flux events moving near the endwall of the vane pressure side is reported for the first time. The temporal nature of the unsteady 3-D film cooling features are a large factor in determining average local heat flux levels. This study determined this effect to be particularly important in areas on real hardware along the HPT vane pressure side endwalls towards the trailing edge, where numerical assumptions are often challenged. Better understanding of the physics of the highly unsteady 3D film cooled flow features occurring in real hardware is necessary to accurately predict distress progression in localized areas, prevent unforeseen part failures, and enable improvements to turbine engine efficiency. The results of this two-part paper are relevant to engines in extended service today.

Assessment of LES and RANS-LES Hybrid Models for Heat Transfer Predictions in Effusion Cooled Combustor Liners

2021 ◽  
Author(s):  
Vishwas Verma ◽  
Kiran Manoharan ◽  
Jaydeep Basani ◽  
Dustin Brandt
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Surface Heat Flux ◽  
Hybrid Models ◽  
Surface Heat ◽  
Regions Of Interest ◽  
Wall Region ◽  
Wall Heat Transfer ◽  
Numerical Predictions ◽  
Near Wall

Abstract Accurate numerical predictions of surface heat flux on combustor liners in the presence of effusion cooling involve appropriate resolution of turbulent boundary layers and mixing of two different streams. Precise surface heat flux and wall temperature predictions are necessary for the optimal design of combustor liners to avoid burnout and damage to the combustor. Reynolds Average Navier Stokes (RANS) model has shown superior wall heat transfer predictions for steady flows; however, in combustor liners involving complex effusion jet mixing patterns, it fails. On the other hand, Large Eddy Simulation (LES) can capture to a good extent core flow mixing in such situations, but it requires very high-resolution near-wall meshes for accurate surface heat flux predictions. To overcome these issues, a hybrid model using RANS in the near-wall region and LES in the core region have been proposed for better wall heat transfer predictions. In this study, a numerical analysis is carried out to test the capability of RANS, LES and hybrid models (SBES, WMLES) for wall heat transfer predictions. The computational setup is a flat plate where freestream high-speed flow approaches a thirty-five degree inclined jet. The study is divided into two regions of interest, one before the jet freestream interaction and another post-interaction. We demonstrate with the SBES approach, surface heat flux can be predicted to much better agreement with the test data in both the regions of interest. Also, it is shown that such results can be obtained with much coarser mesh resolution, hence less computational cost, with hybrid models than pure LES.

PHANTOM COOLING EFFECTS ON ROTOR BLADE SURFACE HEAT FLUX IN A TRANSONIC FULL SCALE 1+1/2 STAGE ROTATING TURBINE

2021 ◽  
pp. 1-13
Author(s):  
Richard J. Anthony ◽  
John Finnegan ◽  
John Clark
Keyword(s):  
Heat Flux ◽  
Air Force ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Rotor Blade ◽  
Surface Heat ◽  
Pressure Side ◽  
Blade Surface ◽  
Leakage Flows ◽  
Cooling Effects

Abstract An experimental and numerical investigation of phantom cooling effects on cooled and uncooled rotating high pressure turbine blades in a full scale 1+1/2 stage turbine test is carried out. Objectives set to capture, separate, and quantify the effects of upstream vane film-cooling and leakage flows on the downstream rotor blade surface heat flux. Multiple series of tests were carried out in the Air Force Research Laboratory, Turbine Research Facility, at Wright-Patterson Air Force Base, Ohio. A non-proprietary research turbine test article is uniquely instrumented with high frequency double-sided thin film heat flux gauges custom made at AFRL. High bandwidth, time resolved surface heat flux is measured on multiple film-cooled and non-film-cooled HPT rotor blades downstream of both film-cooled and non-film-cooled vane sectors. Upstream wake passing and heat flux is characterized on both rotor pressure and suction side surfaces, along with quantifying rotor phantom cooling effects from non-uniform 1st stage vane film cooling and leakage flows. Fast response heat flux measurements quantify how rotor phantom cooling impacts the blade pressure side greatest; increasing along the pressure side towards the trailing edge. It is discovered upstream vane film-cooling alone can account for 50% of the rotor blade cooling effect, and even outweigh the rotor blade film cooling effect far from the blade showerhead holes. Added unsteady numerical simulation demonstrates how variations in inlet total temperature and incidence angle can also contribute to circumferentially non-uniform rotor heat flux.

scholarly journals Heat Transfer Phenomena in Gas Turbines

1980 ◽  
Author(s):  
J. R. Taylor
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Gas Turbines ◽  
Turbine Blades ◽  
Gas Turbine Engine ◽  
Transfer Coefficients ◽  
Turbine Engine ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients

A discussion of the problems encountered in prediction of heat transfer in the turbine section of a gas turbine engine is presented. Areas of current gas turbine engine is presented. Areas of current concern to designers where knowledge is deficient or lacking are elucidated. Consideration is given to methods and problems associated with determination of heat transfer coefficients, external gas temperatures, and, where applicable, film cooling effectiveness. The paper is divided into parts dealing with turbine airfoil heat transfer, endwall heat transfer, and heat transfer in the internal cavities of cooled turbine blades. Recent literature dealing with these topics is listed.

Effects of Bulk Flow Pulsations on Film Cooling With Compound Angle Holes: Heat Transfer Coefficient Ratio and Heat Flux Ratio

2001 ◽  
Vol 124 (1) ◽  
pp. 142-151 ◽  
Author(s):  
In Sung Jung ◽  
Joon Sik Lee ◽  
P. M. Ligrani
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Surface Heat ◽  
Surface Heat Transfer ◽  
Spatially Resolved ◽  
Compound Angle

Experiments are conducted to investigate the effects of bulk flow pulsations on film cooling from compound angle holes. A row of five film cooling holes is employed with orientation angles of 0, 30, 60, and 90 deg at a fixed inclination angle of 35 deg. Static pressure pulsations are generated using an array of six rotating shutter blades, which extend across the span of the exit of the wind tunnel test section. Pulsation frequencies of 0 Hz, 8 Hz, and 36 Hz, and time-averaged blowing ratios of 0.5, 1.0, and 2.0 are employed. Corresponding coolant Strouhal numbers based on these values then range from 0.20 to 3.6. Spatially resolved surface heat transfer coefficient distributions are measured (with the film and freestream at the same temperature) using thermochromic liquid crystals. Presented are ratios of surface heat transfer coefficients with and without film cooling, as well as ratios of surface heat flux with and without film cooling. These results, for compound angle injection, indicate that the pulsations cause the film to be spread more uniformly over the test surface than when no pulsations are employed. This is because the pulsations cause the film from compound angle holes to oscillate in both the normal and spanwise directions after it leaves the holes. As a result, the pulsations produce important changes to spatially resolved distributions of surface heat flux ratios, and surface heat transfer coefficient ratios. In spite of these alterations, only small changes to spatially averaged heat transfer coefficient ratios are produced by the pulsations. Spatially averaged surface heat flux ratios, on the other hand, increase considerably at coolant Strouhal numbers larger than unity, with higher rates of increase at larger orientation angles.

Phantom Cooling Effects on Rotor Blade Surface Heat Flux in a Transonic Full Scale 1+1/2 Stage Rotating Turbine

2020 ◽  
Author(s):  
Richard J. Anthony ◽  
John Finnegan ◽  
John P. Clark
Keyword(s):  
Heat Flux ◽  
High Pressure ◽  
Air Force ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Rotor Blade ◽  
Surface Heat ◽  
High Pressure Turbine ◽  
Leakage Flows ◽  
Cooling Effects

Abstract An experimental and numerical investigation of phantom cooling effects on cooled and uncooled rotating high pressure turbine blades in a full scale 1+1/2 stage turbine test is carried out. Objectives set to capture, separate, and quantify the effects of upstream vane film-cooling and leakage flows on the downstream rotor blade surface heat flux. Multiple series of 1+1/2 stage rotating high pressure turbine tests were carried out in the Air Force Research Laboratory, Turbine Research Facility, at Wright-Patterson Air Force Base, Ohio. A non-proprietary research turbine test article is uniquely instrumented with high frequency double-sided thin film heat flux gauges custom made at AFRL. High bandwidth, time resolved surface heat flux is measured on multiple film-cooled and non-film-cooled HPT rotor blades downstream of both film-cooled and non-film-cooled vane sectors. Upstream wake passing and heat flux is characterized on both rotor pressure and suction side surfaces, along with quantifying rotor phantom cooling effects from nonuniform 1st stage vane film cooling and leakage flows. Fast response heat flux measurements quantify how rotor phantom cooling impacts the blade pressure side greatest; increasing along the pressure side towards the trailing edge. It is discovered upstream vane film-cooling alone can account for 50% of the rotor blade cooling effect, and even outweigh the rotor blade film cooling effect far from the blade showerhead holes. Added unsteady aero numerical simulation demonstrate how variations in inlet total temperature and incidence angle can also contribute to circumferentially non-uniform rotor heat flux. Better understanding from this investigation aids modelling and design efforts in optimizing film cooling performance in real high pressure turbine flow fields. Understanding the behavior of such non-uniform circumferential rotor phantom cooling effects can be critical to optimize the efficiency, fuel consumption, range, and durability of advanced turbomachines.

Effects of Bulk Flow Pulsations on Film Cooling With Compound Angle Holes: Heat Transfer Coefficient Ratio and Heat Flux Ratio

2001 ◽  
Author(s):  
In Sung Jung ◽  
Joon Sik Lee ◽  
P. M. Ligrani
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Surface Heat Flux ◽  
Surface Heat ◽  
Surface Heat Transfer ◽  
Spatially Resolved ◽  
Compound Angle

Experiments are conducted to investigate the effects of bulk flow pulsations on film cooling from compound angle holes. A row of five film cooling holes is employed with orientation angles of 0°, 30°, 60°, and 90° at a fixed inclination angle of 35°. Static pressure pulsations are generated using an array of six rotating shutter blades, which extend across the span of the exit of the wind tunnel test section. Pulsation frequencies of 0 Hz, 8 Hz, and 36 Hz, and time-averaged blowing ratios of 0.5, 1.0, and 2.0 are employed. Corresponding coolant Strouhal numbers based on these values then range from 0.20 to 3.6. Spatially-resolved surface heat transfer coefficient distributions are measured (with the film and freestream at the same temperature) using thermochromic liquid crystals. Presented are ratios of surface heat transfer coefficients with and without film cooling, as well as ratios of surface heat flux with and without film cooling. These results, for compound angle injection, indicate that the pulsations cause the film to be spread more uniformly over the test surface than when no pulsations are employed. This is because the pulsations cause the film from compound angle holes to oscillate in both the normal and spanwise directions after it leaves the holes. As a result, the pulsations produce important changes to spatially-resolved distributions of surface heat flux ratios, and surface heat transfer coefficient ratios. In spite of these alterations, only small changes to spatially-averaged heat transfer coefficient ratios are produced by the pulsations. Spatially-averaged surface heat flux ratios, on the other hand, increase considerably at coolant Strouhal numbers larger than unity, with higher rates of increase at larger orientation angles.

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