Experimental Evaluation of an Inlet Profile Generator for High-Pressure Turbine Tests

2006 ◽  
Vol 129 (2) ◽  
pp. 382-393 ◽  
Author(s):  
M. D. Barringer ◽  
K. A. Thole ◽  
M. D. Polanka
Keyword(s):  
Heat Transfer ◽  
Air Force ◽  
Flow Distribution ◽  
Turbine Engine ◽  
Turbine Inlet ◽  
Pressure Profiles ◽  
Flow Interactions ◽  
Temperature And Pressure ◽  
Inlet Conditions ◽  
Air Force Research Laboratory

Improving the performance and durability of gas turbine aircraft engines depends highly on achieving a better understanding of the flow interactions between the combustor and turbine sections. The flow exiting the combustor is very complex and it is characterized primarily by elevated turbulence and large variations in temperature and pressure. The heat transfer and aerodynamic losses that occur in the turbine passages are driven primarily by these spatial variations. To better understand these effects, the goal of this work is to benchmark an adjustable turbine inlet profile generator for the Turbine Research Facility (TRF) at the Air Force Research Laboratory. The research objective was to experimentally evaluate the performance of the nonreacting simulator that was designed to provide representative combustor exit profiles to the inlet of the TRF turbine test section. This paper discusses the verification testing that was completed to benchmark the performance of the generator. Results are presented in the form of temperature and pressure profiles as well as turbulence intensity and length scale. This study shows how a single combustor geometry can produce significantly different flow and thermal field conditions entering the turbine. Engine designers should place emphasis on obtaining accurate knowledge of the flow distribution within the combustion chamber. Turbine inlet conditions with significantly different profile shapes can result in altered flow physics that can change local aerodynamics and heat transfer.

Experimental Evaluation of an Inlet Profile Generator for High Pressure Turbine Tests

2006 ◽  
Author(s):  
M. D. Barringer ◽  
K. A. Thole ◽  
M. D. Polanka
Keyword(s):  
Heat Transfer ◽  
Air Force ◽  
Thermal Field ◽  
Flow Distribution ◽  
Aircraft Engines ◽  
Turbine Engine ◽  
Pressure Profiles ◽  
Flow Interactions ◽  
Temperature And Pressure ◽  
Air Force Research Laboratory

Improving the performance and durability of gas turbine aircraft engines depends highly on achieving a better understanding of the flow interactions between the combustor and turbine sections. The flow exiting the combustor is very complex and it is characterized primarily by elevated turbulence and large variations in temperature and pressure. The heat transfer and aerodynamic losses that occur in the turbine passages are driven primarily by these spatial variations. To better understand these effects, the goal of this work is to benchmark an adjustable turbine inlet profile generator for the Turbine Research Facility (TRF) at the Air Force Research Laboratory (AFRL). The research objective was to experimentally evaluate the performance of the non-reacting simulator that was designed to provide representative combustor exit profiles to the inlet of the TRF turbine test section. This paper discusses the verification testing that was completed to benchmark the performance of the generator. Results are presented in the form of temperature and pressure profiles as well as turbulence intensity and length scale. This study shows how one combustor geometry can produce significantly different flow and thermal field conditions entering the turbine. Engine designers should place emphasis on obtaining accurate knowledge of the flow distribution within the combustion chamber as this can result in significantly different inlet profiles to the turbine that can change local aerodynamics and heat transfer within the turbine.

Integration of a Turbine Cascade Facility Into an Undergraduate Thermo-Propulsion Sequence

2013 ◽  
Author(s):  
Kurt P. Rouser ◽  
Caitlin R. Thorn ◽  
Aaron R. Byerley ◽  
Charles F. Wisniewski ◽  
Scott R. Nowlin ◽  
...  
Keyword(s):  
Wind Tunnel ◽  
Gas Turbine ◽  
Air Force ◽  
Research Laboratory ◽  
Closed Loop ◽  
Turbine Cascade ◽  
Engineering Majors ◽  
Turbine Engine ◽  
Aeronautical Engineering ◽  
Air Force Research Laboratory

The Department of Aeronautics at the United States Air Force Academy utilizes a closed-loop, two-dimensional turbine cascade wind tunnel to reinforce a learning-focused undergraduate thermo-propulsion sequence. While previous work presented in the literature outlined the Academy thermo-propulsion sequence and the contextual framework for instruction, this current paper addresses how the Academy turbine cascade facility is integrated into the aeronautical engineering course sequence. Cadets who concentrate in propulsion are to some extent prepared for each successive course through their contact with the cascade, and ultimately they graduate with an exposure to experimental research that enhances their grasp of gas turbine engine fundamentals. Initially, the cascade is used to reinforce airfoil theory to all cadets in the Fundamentals of Aeronautics course. Aeronautical engineering majors take this course during the first semester of their sophomore year. The next semester all aeronautical engineering majors take Introduction to Aero-thermodynamics. In this course, the closed-loop aspect of the cascade facility is used to reinforce concepts of work addition to the flow. Heat transfer is also discussed, using the heat exchanger that regulates test section temperature. Exposure to the cascade also prepares cadets for the ensuing Introduction to Propulsion and Aeronautics Laboratory courses, taken in the junior and senior year, respectively. In the propulsion course, cadets connect thermodynamic principles to component analysis. In the laboratory course, cadets work in pairs on propulsion projects sponsored by the Air Force Research Laboratory, including projects in the cascade wind tunnel. Individual cadets are selected from the cascade research teams for summer internships, working at an Air Force Research Laboratory turbine cascade tunnel. Ultimately, cadet experiences with the Academy turbine cascade help lay the foundation for a two-part senior propulsion capstone sequence in which cadets design a gas turbine engine starting with the overall cycle selection leading to component-level design. The turbine cascade also serves to integrate propulsion principles and fluid mechanics through a senior elective Computational Fluid Dynamics course. In this course, cadets may select a computational project related to the cascade. Cadets who complete the thermo-propulsion sequence graduate with a thorough understanding of turbine engine fundamentals from both conceptual and applied perspectives. Their exposure to the cascade facility is an important part of the process. An assessment of cadet learning is presented to validate the effectiveness of this integrated research-classroom approach.

Effect of Nonuniform Inlet Conditions on Endwall Secondary Flows

2002 ◽  
Vol 124 (4) ◽  
pp. 623-631 ◽  
Author(s):  
K. S. Hermanson ◽  
K. A. Thole
Keyword(s):  
Secondary Flow ◽  
Total Pressure ◽  
Large Scale ◽  
Secondary Flows ◽  
Turbine Engine ◽  
Thermal Fields ◽  
Pressure Profiles ◽  
Stator Vane ◽  
Inlet Conditions ◽  
Computational Predictions

Exit combustor flow and thermal fields entering downstream stator vane passages in a gas turbine engine are highly nonuniform. These flow and thermal fields can significantly affect the development of the secondary flows in the turbine passages contributing to high platform heat transfer and large aerodynamic losses. The flow and thermal fields combine to give nonuniform total pressure profiles entering the turbine passage which, along with the airfoil geometry, dictate the secondary flow field. This paper presents an analysis of the effects of varying total pressure profiles in both the radial and combined radial and circumferential directions on the secondary flowfields in a first-stage stator vane. These inlet conditions used for the first vane simulations are based on the exit conditions predicted for a combustor. Prior to using the predictions, these CFD simulations were benchmarked against flowfield data measured in a large-scale, linear, turbine vane cascade. Good agreement occurred between the computational predictions and experimentally measured secondary flows. Analyses of the results for several different cases indicate variations in the secondary flow pattern from pitch to pitch, which attributes to the rationale as to why some airfoils quickly degrade while others remain intact over time.

scholarly journals Effects of inertial load and countermeasures on the distribution of pulmonary blood flow

2000 ◽  
Vol 89 (2) ◽  
pp. 445-457 ◽  
Author(s):  
Myron A. Chornuk ◽  
Susan L. Bernard ◽  
John W. Burns ◽  
Robb W. Glenny ◽  
Don D. Sheriff ◽  
...  
Keyword(s):  
Blood Flow ◽  
Air Force ◽  
Pulmonary Blood Flow ◽  
Inertial Force ◽  
Flow Distribution ◽  
Pulmonary Vasculature ◽  
Positive Pressure ◽  
Anatomic Structure ◽  
Air Force Base ◽  
Air Force Research Laboratory

We assessed the influence of cranial-to-caudal inertial force (+G z ) and the countermeasures of anti-G suit and positive pressure breathing during G (PBG), specifically during +G z , on regional pulmonary blood flow distribution. Unanesthetized swine were exposed randomly to 0 G z (resting), +3 G z , +6 G z , and +9 G z , with and without anti-G suit and PBG with the use of the Air Force Research Laboratory centrifuge at Brooks Air Force Base (the gravitational force of the Earth, that is, the dorsal-to-ventral inertial force, was present for all runs). Fluorescent microspheres were injected into the pulmonary vasculature as a marker of regional pulmonary blood flow. Lungs were excised, dried, and diced into ∼2-cm3 pieces, and the fluorescence of each piece was measured. As +G z was increased from 0 to +3 G z , blood flow shifted from cranial and hilar regions toward caudal and peripheral regions of the lung. This redistribution shifted back toward cranial and hilar regions as anti-G suit inflation pressure increased at +6 and +9 G z . Perfusion heterogeneity increased with +G z stress and decreased at the higher anti-G suit pressures. The distribution of pulmonary blood flow was not affected by PBG. ANOVA indicated anatomic structure as the major determinant of pulmonary blood flow.

Effects of Combustor Exit Profiles on High Pressure Turbine Vane Aerodynamics and Heat Transfer

2006 ◽  
Author(s):  
M. D. Barringer ◽  
K. A. Thole ◽  
M. D. Polanka
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Total Pressure ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Pressure Profiles ◽  
Turbine Vane ◽  
Baseline Test ◽  
Total Temperature ◽  
Air Force Research Laboratory

The goal of this work was to investigate the effects of different profiles representative of those exiting aero-engine combustors on high pressure turbine vane aerodynamics and heat transfer. The various profiles were produced using the non-reacting, inlet profile generator in the Turbine Research Facility (TRF) located at the Air Force Research Laboratory (AFRL). This paper reports how the pressure loading and heat transfer along the vane surface is affected by different turbine inlet pressure and temperature profiles at several different span locations. The results indicate that the different inlet total pressure profiles affected the aerodynamic loading by as much as 10%. The results also reveal that the combination of different total pressure and total temperature profiles significantly affected the vane heat transfer for a baseline test with relatively uniform inlet total pressure and total temperature profiles. Near the ID endwall, the baseline heat transfer was reduced 30 to 40% over the majority of the vane surface. Near the OD endwall, it was found that certain inlet profiles could increase the baseline heat transfer by 20 to 30%, while other profiles resulted in a decrease of the baseline heat transfer by 30 to 40%.

A Review of the AFRL Turbine Research Facility

2013 ◽  
Author(s):  
Richard J. Anthony ◽  
John P. Clark
Keyword(s):  
Heat Transfer ◽  
Short Duration ◽  
Air Force ◽  
Full Scale ◽  
Research Facility ◽  
Shock Interactions ◽  
Transfer Testing ◽  
Air Force Research Laboratory ◽  
Future Work ◽  
Validation Testing

The Air Force Research Laboratory Turbine Research Facility is a transient blowdown facility that allows simultaneous measurement of unsteady heat transfer and aerodynamics on full scale engine hardware. It is unique for its size and consequent blowdown duration. Compared to engine validation testing, full scale, short duration turbine-rig testing is able to provide very large amounts of rotating turbine flowfield information using far less energy, orders of magnitude lower cost, and greater instrumentation selection. This paper provides an updated review of the facility’s history, operation, and enhancements since its initial construction over two decades ago. Historical connections to pioneering work in short-duration turbine heat transfer testing are highlighted, and an overview of past developments, features, and capabilities is given. More recent experimental and computational integration is described using a suite of in-house developed CFD design and analysis tools. Example test programs include a non-proprietary 1+ 1/2 stage research turbine rig, which is the most heavily instrumented high pressure turbine tested in the facility to date.. Recent data illustrates the character of unsteady airfoil shock interactions that may lead to large levels of resonant stress or turbine high cycle fatigue. The paper ends with a brief discussion of future work.

Effects of Inlet Temperature Uniformity and Nonuniformity on the Tip Leakage Flow and Rotor Blade Tip and Casing Heat Transfer Characteristics

2011 ◽  
Vol 134 (2) ◽  
Author(s):  
Md. Hamidur Rahman ◽  
Sung In Kim ◽  
Ibrahim Hassan
Keyword(s):  
Heat Transfer ◽  
Operating Conditions ◽  
Inlet Temperature ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Turbine Inlet Temperature ◽  
Turbine Engine ◽  
Tip Leakage ◽  
Turbine Inlet ◽  
Blade Tip

High thermal load appears at the blade tip and casing of a gas turbine engine. It becomes a significant design challenge to protect the turbine materials from this severe situation. As a result of geometric complexity and experimental limitations, computational fluid dynamics tools have been used to predict blade tip leakage flow aerodynamics and heat transfer at typical engine operating conditions. In this paper, the effect of turbine inlet temperature on the tip leakage flow structure and heat transfer has been studied numerically. Uniform low (444 K) and high (800 K) inlet temperatures and nonuniform (parabolic) temperature profiles have been considered at a fixed rotor rotation speed (9500 rpm). The results showed that the change of flow properties at a higher inlet temperature yields significant variations in the leakage flow aerodynamics and heat transfer relative to the lower inlet temperature condition. Aerodynamic behavior of the tip leakage flow varies significantly with the distortion of turbine inlet temperature. For more realistic inlet condition, the velocity range is insignificant at all time instants. At a high inlet temperature, reverse secondary flow is strongly opposed by the tip leakage flow and the heat transfer fluctuations are reduced greatly.

Effect of Non-Uniform Inlet Conditions on Endwall Secondary Flows

2002 ◽  
Author(s):  
K. S. Hermanson ◽  
K. A. Thole
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Total Pressure ◽  
Large Scale ◽  
Secondary Flows ◽  
Turbine Engine ◽  
Thermal Fields ◽  
Pressure Profiles ◽  
Stator Vane ◽  
Inlet Conditions

Exit combustor flow and thermal fields entering downstream stator vane passages in a gas turbine engine are highly non-uniform. These flow and thermal fields can significantly affect the development of the secondary flows in the turbine passages contributing to high platform heat transfer and large aerodynamic losses. The flow and thermal fields combine to give non-uniform total pressure profiles entering the turbine passage which, along with the airfoil geometry, dictate the secondary flow field. This paper presents an analysis of the effects of varying total pressure profiles in both the radial and combined radial and circumferential directions on the secondary flow fields in a first stage stator vane. These inlet conditions used for the first vane simulations are based on the exit conditions predicted for a combustor. Prior to using the predictions, these CFD simulations were benchmarked against flow field data measured in a large-scale, linear, turbine vane cascade. Good agreement occurred between the computational predictions and experimentally measured secondary flows. Analyses of the results for several different cases indicate variations in the secondary flow pattern from pitch to pitch, which attributes to the rationale as to why some airfoils quickly degrade while others remain intact over time.

Performance of a Splittered Transonic Rotor With Several Tip Clearances

2015 ◽  
Author(s):  
Anthony J. Gannon ◽  
Garth V. Hobson ◽  
Mark G. Turner ◽  
Wolfgang Sanz
Keyword(s):  
Pressure Ratio ◽  
Tip Clearance ◽  
Test Case ◽  
Data Set ◽  
Turbine Engine ◽  
Blade Height ◽  
Pressure Profiles ◽  
Full Speed ◽  
Inlet Conditions ◽  
Transonic Rotor

The effect of increasing the tip-gap size on the performance of a splittered transonic rotor is presented. Tip clearance has a large influence on the performance and efficiency of compressors and fans during operation. In a gas turbine engine the ratio of tip-gap to blade height or span usually increases in the direction of flow. The front stages usually have a smaller ratio of tip-gap to blade height than the aft core stages. In addition the front stages are usually operating in the transonic regime while the rear stages operate sub-sonically. In order to be representative of these differing flow regimes the results of a range of tests at varying tip-gaps and speeds from subsonic to transonic are presented. A highly loaded transonic axial splittered rotor is used as the test article in this study. Three experiments with cold tip gaps of 0.53 [mm], 0.76 [mm] and 0.99 [mm] are presented. Each experiment was run at six tip-Mach numbers ranging from Mach 0.72 to Mach 1.2 each over a full speed-line from choked to stalling conditions. Exit temperature and pressure profiles at the rotor exit are presented along with performance maps of pressure ratio and efficiency. Significant differences in performance in terms of pressure ratio, efficiency and operating range due to the tip-gap increase were observed and are presented. The full mechanical geometry is available upon request to provide an open test case to evaluate simulation codes. This includes the cold-shape, the cavity between the rotor and stationary exit and the mounting bolts and geometry. The manufacturing steps in the preparation of the material and methods used to form the abradable material used over the rotor are presented. In addition the inlet conditions taking into account the effect of relative humidity on gas properties are presented. The complete data set of experimental results is also available electronically.

Effects of Combustor Exit Profiles on Vane Aerodynamic Loading and Heat Transfer in a High Pressure Turbine

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
M. D. Barringer ◽  
K. A. Thole ◽  
M. D. Polanka
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Total Pressure ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Local Flow ◽  
Aerodynamic Loading ◽  
Turbine Inlet ◽  
Pressure Profiles ◽  
Total Temperature

The flow and thermal fields exiting gas turbine combustors dictate the overall performance of the downstream turbine. The goal of this work was to investigate the effects of engine representative combustor exit profiles on high pressure turbine vane aerodynamics and heat transfer. The various profiles were produced using a nonreacting turbine inlet profile generator in the Turbine Research Facility (TRF) located at the Air Force Research Laboratory (AFRL). This paper reports how the pressure loading and heat transfer along the vane surface was affected by different turbine inlet pressure and temperature profiles at different span locations. The results indicate that the inlet total pressure profiles affected the aerodynamic loading by as much as 10%. The results also reveal that the combination of different total pressure and total temperature profiles significantly affected the vane heat transfer relative to a baseline test with uniform inlet total pressure and total temperature. Near the inner diameter endwall, the baseline heat transfer was reduced 30–40% over the majority of the vane surface. Near the outer dimeter endwall, it was found that certain inlet profiles could increase the baseline heat transfer by 10–20%, while other profiles resulted in a decrease in the baseline heat transfer by 25–35%. This study also shows the importance of knowing an accurate prediction of the local flow driving temperature when determining vane surface heat transfer.

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