Development of New Single and High-Density Heat-Flux Gauges for Unsteady Heat Transfer Measurements for a Rotating Transonic Turbine

2020 ◽  
Author(s):  
Richard Celestina ◽  
Spencer Sperling ◽  
Louis Christensen ◽  
Randall Mathison ◽  
Hakan Aksoy ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Pressure Ratio ◽  
High Density ◽  
Secondary Flows ◽  
Single Stage ◽  
High Pressure Turbine ◽  
Transonic Turbine ◽  
New Generation

Abstract This paper presents the development and implementation of a new generation of double-sided heat-flux gauges at The Ohio State University Gas Turbine Laboratory (GTL) along with heat transfer measurements for film-cooled airfoils in a single-stage high-pressure transonic turbine operating at design corrected conditions. Double-sided heat flux gauges are a critical part of turbine cooling studies, and the new generation improves upon the durability and stability of previous designs while also introducing high-density layouts that provide better spatial resolution. These new customizable high-density double-sided heat flux gauges allow for multiple heat transfer measurements in a small geometric area such as immediately downstream of a row of cooling holes on an airfoil. Two high-density designs are utilized: Type A consists of 9 gauges laid out within a 5 mm by 2.6 mm (0.20 inch by 0.10 inch) area on the pressure surface of an airfoil, and Type B consists of 7 gauges located at points of predicted interest on the suction surface. Both individual and high-density heat flux gauges are installed on the blades of a transonic turbine experiment for the second build of the High-Pressure Turbine Innovative Cooling program (HPTIC2). Run in a short duration facility, the single-stage high-pressure turbine operated at design-corrected conditions (matching corrected speed, flow function, and pressure ratio) with forward and aft purge flow and film-cooled blades. Gauges are placed at repeated locations across different cooling schemes in a rainbow rotor configuration. Airfoil film-cooling schemes include round, fan, and advanced shaped cooling holes in addition to uncooled airfoils. Both the pressure and suction surfaces of the airfoils are instrumented at multiple wetted distance locations and percent spans from roughly 10% to 90%. Results from these tests are presented as both time-average values and time-accurate ensemble averages in order to capture unsteady motion and heat transfer distribution created by strong secondary flows and cooling flows.

Development of New Single and High-Density Heat-Flux Gauges for Unsteady Heat Transfer Measurements for a Rotating Transonic Turbine

2021 ◽  
Author(s):  
Richard Celestina ◽  
Spencer Sperling ◽  
Louis Christensen ◽  
Randall Mathison ◽  
Hakan Aksoy ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Density ◽  
Unsteady Heat Transfer ◽  
Transonic Turbine

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.

Design and Test Results of a Ultra High Loaded Single Stage High Pressure Turbine

2013 ◽  
Author(s):  
Harjit S. Hura ◽  
Scott Carson ◽  
Rob Saeidi ◽  
Hyoun-Woo Shin ◽  
Paul Giel
Keyword(s):  
High Pressure ◽  
Total Pressure ◽  
High Speed ◽  
Pressure Ratio ◽  
Single Stage ◽  
Test Results ◽  
High Pressure Turbine ◽  
Technology Program ◽  
High Pressure Ratio ◽  
Expected Performance

This paper describes the engine and rig design, and test results of an ultra-highly loaded single stage high pressure turbine. In service aviation single stage HPTs typically operate at a total-to-total pressure ratio of less than 4.0. At higher pressure ratios or energy extraction the nozzle and blade both have regions of supersonic flow and shock structures which, if not mitigated, can result in a large loss in efficiency both in the turbine itself and due to interaction with the downstream component which may be a turbine center frame or a low pressure turbine. Extending the viability of the single stage HPT to higher pressure ratios is attractive as it enables a compact engine with less weight, and lower initial and maintenance costs as compared to a two stage HPT. The present work was performed as part of the NASA UEET (Ultra-Efficient Engine Technology) program from 2002 through 2005. The goal of the program was to design and rig test a cooled single stage HPT with a pressure ratio of 5.5 with an efficiency at least two points higher than the state of the art. Preliminary design tools and a design of experiments approach were used to design the flow path. Stage loading and through-flow were set at appropriate levels based on prior experience on high pressure ratio single stage turbines. Appropriate choices of blade aspect ratio, count, and reaction were made based on comparison with similar HPT designs. A low shock blading design approach was used to minimize the shock strength in the blade during design iterations. CFD calculations were made to assess performance. The HPT aerodynamics and cooling design was replicated and tested in a high speed rig at design point and off-design conditions. The turbine met or exceeded the expected performance level based on both steady state and radial/circumferential traverse data. High frequency dynamic total pressure measurements were made to understand the presence of unsteadiness that persists in the exhaust of a transonic turbine.

Aerodynamics and Heat Transfer for a Cooled One and One-Half Stage High-Pressure Turbine—Part III: Impact of Hot-Streak Characteristics on Blade Row Heat Flux

2011 ◽  
Vol 134 (1) ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Radial Profile ◽  
Leading Edge ◽  
Coolant Temperature ◽  
High Pressure Turbine ◽  
Blade Row ◽  
Hot Streak ◽  
The Impact

The influence of hot-streak magnitude and alignment relative to the vane leading edge on blade row heat flux is investigated for a one and one-half stage high-pressure turbine with a film-cooled vane, purge cooling, and uncooled blades. The full-stage turbine is operated at design-corrected conditions. In addition to investigating the impact of different hot-streak characteristics, this study also looks at the interaction of cooling flow with the hot streaks. This paper builds on the investigation of profile migration utilizing temperature measurements presented in Part I and the heat transfer measurements presented in Part II. Hot streaks aligned with the vane midpitch have a greater impact on blade temperatures and heat-flux values than hot streaks aligned with the vane leading edge. The leading edge hot streaks tend to be mixed out over the surface of the vane. The magnitude of the hot streak is observed to have the largest influence on the temperature and heat flux for the downstream blade. Time-accurate measurements confirm these conclusions and indicate that further analysis of the time-accurate data is warranted. Film cooling is found to impact a hot-streak profile in a way similar to that observed for a radial profile. Differences in core to coolant temperature ratio cause the uniform profile to show different coolant effects, but the overall spread of the cooling appears similar.

Heat Transfer Measurements and Predictions for a Modern, High-Pressure, Transonic Turbine, Including Endwalls

2009 ◽  
Vol 131 (2) ◽  
Author(s):  
James A. Tallman ◽  
Charles W. Haldeman ◽  
Michael G. Dunn ◽  
Anil K. Tolpadi ◽  
Robert F. Bergholz
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbulence Modeling ◽  
Pressure Ratio ◽  
Operating Conditions ◽  
Test Facility ◽  
Navier Stokes ◽  
Local Measurements ◽  
Transonic Turbine ◽  
The Ohio State University

This paper presents both measurements and predictions of the hot-gas-side heat transfer to a modern, 112 stage high-pressure, transonic turbine. Comparisons of the predicted and measured heat transfer are presented for each airfoil at three locations, as well as on the various endwalls and rotor tip. The measurements were performed using the Ohio State University Gas Turbine Laboratory Test Facility (TTF). The research program utilized an uncooled turbine stage at a range of operating conditions representative of the engine: in terms of corrected speed, flow function, stage pressure ratio, and gas-to-metal temperature ratio. All three airfoils were heavily instrumented for both pressure and heat transfer measurements at multiple locations. A 3D, compressible, Reynolds-averaged Navier–Stokes computational fluid dynamics (CFD) solver with k-ω turbulence modeling was used for the CFD predictions. The entire 112 stage turbine was solved using a single computation, at two different Reynolds numbers. The CFD solutions were steady, with tangentially mass-averaged inlet/exit boundary condition profiles exchanged between adjacent airfoil-rows. Overall, the CFD heat transfer predictions compared very favorably with both the global operation of the turbine and with the local measurements of heat transfer. A discussion of the features of the turbine heat transfer distributions, and their association with the corresponding flow-physics, has been included.

scholarly journals High Pressure Turbine Vane Annular Cascade Heat Flux and Aerodynamic Measurements With Comparisons to Predictions

1998 ◽  
Author(s):  
Christopher R. Joe ◽  
Xavier A. Montesdeoca ◽  
Friedrich O. Soechting ◽  
Charles D. MacArthur ◽  
Matthew Meininger
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
High Pressure ◽  
Heat Loss ◽  
Short Duration ◽  
Wall Temperature ◽  
Measured Temperature ◽  
Unique Contribution ◽  
Pressure Measurements ◽  
High Pressure Turbine

Experimental tests were performed at the USAF Turbine Research Facility (TRF) to obtain heat transfer and aerodynamic data on a first stage vane of a modern high pressure turbine. This is a transient blowdown facility that provides data from short duration tests. Data for a matrix of test conditions were obtained to document the effect of inlet Reynolds number, the stage pressure ratio across the vane, and the gas-to-wall temperature ratio. The objectives of these tests were to assess the capability of obtaining accurate aerodynamic total pressure loss measurements and airfoil static pressure measurements as well as determine the heat transfer coefficient distributions on the vanes. Results from these tests were compared to analytical predictions and are presented. The unique contribution of the work presented herein is: 1) demonstration of circumferential traversing temperature and pressure data in a short duration facility test, and 2) heat loss closure during a short duration test using heat flux gauges and the measured temperature loss. The transient heat loss during a short duration test is a fundamental requirement to determine turbine efficiency when work extraction is determined from the temperature drop across the turbine stage. Heat transfer data were acquired from heat flux gauges that were fabricated using thin-film sputtering techniques and placed on the airfoil surfaces. The surface temperature of the gauge was measured and heat flux was determined from a closed form transient semi-infinite solution that included the resistance of the heat flux gauge and the underlying metal substrate. Circumferentially, pressure measurements were obtained on the airfoil surfaces and on traversing rakes at the inlet and exit of the vane test section. Total and differential pressure rake instrumentation was required to obtain accurate aerodynamic loss measurements over a range of gas-to-wall temperature ratios.

scholarly journals Analysis of Unsteady Tip and Endwall Heat Transfer in a Highly Loaded Transonic Turbine Stage

2010 ◽  
Author(s):  
Vikram Shyam ◽  
Ali Ameri ◽  
Jen-Ping Chen
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Reynolds Number ◽  
Rotor Blade ◽  
High Speed ◽  
Pressure Ratio ◽  
Stagnation Temperature ◽  
Turbine Stage ◽  
Transonic Turbine ◽  
Highly Loaded

In a previous study, vane-rotor shock interactions and heat transfer on the rotor blade of a highly loaded transonic turbine stage were simulated. The geometry consists of a high pressure turbine vane and downstream rotor blade. This study focuses on the physics of flow and heat transfer in the rotor tip, casing and hub regions. The simulation was performed using the URANS (Unsteady Reynolds-Averaged Navier-Stokes) code MSU-TURBO. A low Reynolds number k-ε model was utilized to model turbulence. The rotor blade in question has a tip gap height of 2.1% of the blade height. The Reynolds number of the flow is approximately 3×106 per meter. Unsteadiness was observed at the tip surface that results in intermittent ‘hot spots’. It is demonstrated that unsteadiness in the tip gap is governed by inviscid effects due to high speed flow and is not strongly dependent on pressure ratio across the tip gap contrary to published observations that have primarily dealt with subsonic tip flows. The high relative Mach numbers in the tip gap lead to a choking of the leakage flow that translates to a relative attenuation of losses at higher loading. The efficacy of new tip geometry is discussed to minimize heat flux at the tip while maintaining choked conditions. In addition, an explanation is provided that shows the mechanism behind the rise in stagnation temperature on the casing to values above the absolute total temperature at the inlet. It is concluded that even in steady mode, work transfer to the near tip fluid occurs due to relative shearing by the casing. This is believed to be the first such explanation of the work transfer phenomenon in the open literature. The difference in pattern between steady and time-averaged heat flux at the hub is also explained.

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.

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