Endwall Film Cooling Performance for a First-Stage Guide Vane With Upstream Combustor Walls and Inlet Injection

2018 ◽  
Vol 11 (1) ◽  
Author(s):  
Xing Yang ◽  
Zhao Liu ◽  
Zhansheng Liu ◽  
Terrence Simon ◽  
Zhenping Feng
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Distribution Patterns ◽  
Flat Plates ◽  
Cooling Performance ◽  
Axial Shift ◽  
Endwall Film Cooling ◽  
Film Cooling Holes ◽  
Endwall Cooling

Abstract Effects of an upstream combustor wall on turbine nozzle endwall film cooling performance are numerically examined in a linear cascade in this paper. Film cooling is by two rows of cooling holes at 20% of the axial chord length upstream of the vane leading edge (LE) plane. The combustor walls are modeled as flat plates with square trailing edges (TE) positioned upstream of the endwall film cooling holes. A combustor wall is in line with the LE of every second vane. The influence of the combustor wall, when shifted in the axial and tangential directions, is investigated to determine effects on passage endwall cooling for three representative film cooling blowing ratios. The results show how shed vortices from the combustor wall greatly alter the flow field near the cooling holes and inside the vane passage. Film cooling distribution patterns, particularly in the entry region and along the pressure side of the passage, are affected. The combustor wall leads to an imbalance in film cooling distribution over the endwalls for adjacent vane passages. Results show a larger effect of tangential shift of the combustor wall on endwall cooling effectiveness than the effect of an equal axial shift. The study provides guidance regarding design of combustor-to-turbine transition ducts.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

Adiabatic Effectiveness Measurements of Endwall Film-Cooling for a First Stage Vane

2004 ◽  
Author(s):  
D. G. Knost ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Extreme Environment ◽  
Wall Region ◽  
Pressure Side ◽  
Stagnation Region ◽  
Pressure Surface ◽  
Adiabatic Effectiveness ◽  
Endwall Film Cooling

In gas turbine development, the direction has been towards higher turbine inlet temperatures to increase the work output and thermal efficiency. This extreme environment can significantly impact component life. One means of preventing component burnout in the turbine is to effectively use film-cooling whereby coolant is extracted from the compressor and injected through component surfaces. One such surface is the endwall of the first stage nozzle guide vane. This paper presents measurements of two endwall film-cooling hole patterns combined with cooling from a flush slot that simulates leakage flow between the combustor and turbine sections. Adiabatic effectiveness measurements showed the slot flow adequately cooled portions of the endwall. Measurements also showed two very difficult regions to cool including the leading edge and pressure side-endwall junction. As the momentum flux ratios were increased for the film-cooling jets in the stagnation region, the coolant was shown to impact the vane and wash down onto the endwall surface. Along the pressure side of the vane in the upstream portion of the passage, the jets were shown to separate from the surface rather than penetrate to the pressure surface. In the downstream portion of the passage, the jets along the pressure side of the vane were shown to impact the vane thereby eliminating any uncooled regions at the junction. The measurements were also combined with computations to show the importance of considering the trajectory of the flow in the near-wall region, which can be highly influenced by slot leakage flows.

Adiabatic Effectiveness Measurements of Endwall Film-Cooling for a First-Stage Vane

2005 ◽  
Vol 127 (2) ◽  
pp. 297-305 ◽  
Author(s):  
D. G. Knost ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Extreme Environment ◽  
Wall Region ◽  
Pressure Side ◽  
Stagnation Region ◽  
Pressure Surface ◽  
Adiabatic Effectiveness ◽  
Endwall Film Cooling

In gas turbine development, the direction has been toward higher turbine inlet temperatures to increase the work output and thermal efficiency. This extreme environment can significantly impact component life. One means of preventing component burnout in the turbine is to effectively use film-cooling whereby coolant is extracted from the compressor and injected through component surfaces. One such surface is the endwall of the first-stage nozzle guide vane. This paper presents measurements of two endwall film-cooling hole patterns combined with cooling from a flush slot that simulates leakage flow between the combustor and turbine sections. Adiabatic effectiveness measurements showed the slot flow adequately cooled portions of the endwall. Measurements also showed two very difficult regions to cool, including the leading edge and pressure side-endwall junction. As the momentum flux ratios were increased for the film-cooling jets in the stagnation region, the coolant was shown to impact the vane and wash down onto the endwall surface. Along the pressure side of the vane in the upstream portion of the passage, the jets were shown to separate from the surface rather than penetrate to the pressure surface. In the downstream portion of the passage, the jets along the pressure side of the vane were shown to impact the vane thereby eliminating any uncooled regions at the junction. The measurements were also combined with computations to show the importance of considering the trajectory of the flow in the near-wall region, which can be highly influenced by slot leakage flows.

Experimental Investigation of Innovative Cooling Schemes on an Additively Manufactured Engine Scale Turbine Nozzle Guide Vane

2020 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Jet Impingement ◽  
Development Effort ◽  
Suction Surface ◽  
Trailing Edge ◽  
Pin Fin ◽  
Nozzle Guide ◽  
Film Cooling Holes

Abstract This study includes the design, validation, and fabrication via Direct Metal Laser Sintering (DMLS) of a gas turbine nozzle guide vanes (NGV) that incorporates three innovative cooling schemes specifically enabled by additive manufacturing. The novel NGV design is the culmination of an extensive research and development effort over a period of four years that included low and high speed cascade testing coupled with unsteady CFD for numerous candidate innovative cooling architectures. The final vane design (SJ-vane) consists of sweeping jet (SJ) film cooling holes on the suction surface, sweeping jet impingement holes at the leading edge and double-wall partial length triangular pin-fin with impinging jet at the trailing edge. For comparison purposes, a second DMLS enabled vane (777-vane) was designed and fabricated with prototypical cooling circuits to serve as a baseline. This vane consists of a shaped film cooling holes on the suction surface, circular impingement holes at the leading edge and full length cylindrical pin-fins at the trailing edge. Experiments with the two DMLS enabled vanes were performed at the Ohio State University Turbine Reacting Flow Rig (TuRFR) at engine relevant temperature (1375K) and Mach number conditions. Infrared (IR) thermography was utilized to measure the wall temperature of the pressure and suction surface at several coolant mass flow rates to estimate the overall cooling effectiveness (ϕ). Results showed improved cooling performance for the advanced cooling schemes (sweeping jet film cooling, impingement cooling and triangular pin-fin cooling) compared to the baseline cooling schemes.

scholarly journals The Effects of Combustor Cooling Features on Nozzle Guide Vane Film Cooling Experiments

2018 ◽  
Author(s):  
Nicholas E. Holgate ◽  
Peter T. Ireland ◽  
Eduardo Romero
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Stagnation Region ◽  
Airfoil Surface ◽  
Cooling Performance ◽  
Nozzle Guide Vane ◽  
Effectiveness Measurement ◽  
Extraneous Radiation ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Recent advances in experimental methods have allowed researchers to study nozzle guide vane film cooling in the presence of combustor dilution ports and endwall films. The dilution injection creates nonuniformities in temperature, velocity, and turbulence, and an understanding of the vane film cooling performance is complicated by competing influences. In this study, dilution port temperature profiles have been measured in the absence of vane film cooling and compared to film effectiveness measurements in the presence of both films and dilution, illustrating the effects of the dilution port turbulence on film cooling performance. It is found that dilution port injection can create significant effectiveness benefits at the difficult-to-cool vane stagnation region, due to the more turbulent hot mainstream enhancing the mixing of film coolant jets that have left the airfoil surface. Also explored are the implications of endwall film cooling for infrared vane surface temperature measurements. The reduced endwall temperatures reduce the thermal emissions from this surface, so reducing the amount of extraneous radiation reflected from the vane surface where measurements are being made. The results of a detailed calibration show that the maximum local film effectiveness measurement error could be up to 0.05 if this effect were to go unaccounted for.

Optimal Arrangement of the Film Cooling Holes Considering the Manufacturing Tolerance for High Pressure Turbine Nozzle

2016 ◽  
Author(s):  
Sanga Lee ◽  
Dong-Ho Rhee ◽  
Kwanjung Yee
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Manufacturing Tolerance ◽  
Cooling Hole ◽  
Film Cooling Holes

In spite of a myriad of researches on the optimal shape of film cooling holes, only a few attempts have been made to optimize the hole arrangement for film cooling so far. Moreover, although the general scale of film cooling hole is so small that manufacturing tolerance has substantial effects on the cooling performance of turbine, the researches on this issue are even scarcer. If it is possible to obtain optimal hole arrangement which not only improve the film cooling performance but also is robust to the manufacturing tolerance, then overall cooling performance of a turbine would become more reliable and useful from the practical point of view. To this end, the present study proposed a robust design optimization procedure which takes the manufacturing uncertainties into account. The procedure was subsequently applied to the film cooling holes on high pressure turbine nozzle pressure side to obtain the robust array shape under the uncertainty of the manufacturing tolerance. First, the array of the holes was parameterized by 5 design variables using the newly suggested shape functions, and 2 representative factors were considered for the manufacturing tolerance of the film cooling hole. Probabilistic process that consists of Kriging surrogate model and Monte Carlo Simulation with descriptive sampling method was coupled with the design optimization process using Genetic Algorithm. Through this, film cooling hole array which shows the high performance, yet robust to the manufacturing tolerance was obtained, and the effects of the manufacturing tolerance on the cooling performance was carefully investigated. As a result, the region where the film cooling effectiveness is noticeable, as well as the maximum width of the variation of the film cooling effectiveness were reduced through optimization, and it is also confirmed that the tolerance of the holes near the leading edge is more influential to the cooling performance because the film cooling effectiveness is more sensitive to the manufacturing tolerance of the leading edge than that of the trailing edge.

Experimental Investigation of Innovative Cooling Schemes on an Additively Manufactured Engine Scale Turbine Nozzle Guide Vane

2021 ◽  
pp. 1-39
Author(s):  
Mohammad A. Hossain ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey Bons
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Jet Impingement ◽  
Development Effort ◽  
Suction Surface ◽  
Trailing Edge ◽  
Pin Fin ◽  
Nozzle Guide ◽  
Film Cooling Holes

Abstract This study includes the design, validation, and fabrication via Direct Metal Laser Sintering (DMLS) of a gas turbine nozzle guide vanes (NGV) that incorporates three innovative cooling schemes specifically enabled by additive manufacturing. The novel NGV design is the culmination of an extensive research and development effort over a period of four years that included low and high speed cascade testing coupled with unsteady CFD for numerous candidate innovative cooling architectures. The final vane design (SJ-vane) consists of sweeping jet (SJ) film cooling holes on the suction surface, sweeping jet impingement holes at the leading edge and double-wall partial length triangular pin-fin with impinging jet at the trailing edge. For comparison purposes, a second DMLS enabled vane (777-vane) was designed and fabricated with prototypical cooling circuits to serve as a baseline. This vane consists of a shaped film cooling holes on the suction surface, circular impingement holes at the leading edge and full length cylindrical pin-fins at the trailing edge. Experiments with the two DMLS enabled vanes were performed at the Ohio State University Turbine Reacting Flow Rig (TuRFR) at engine relevant temperature (1375 K) and Mach number conditions. Infrared (IR) thermography was utilized to measure the wall temperature of the pressure and suction surface at several coolant mass flow rates to estimate the overall cooling effectiveness (ϕ). Results showed improved cooling performance for the advanced cooling schemes (sweeping jet film cooling, impingement cooling and triangular pin-fin cooling) compared to the baseline cooling schemes.

Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwalls: Part 1 — Aerodynamics

2020 ◽  
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong W. Kim ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Three Dimensional ◽  
Leading Edge ◽  
Secondary Flows ◽  
Protective Film ◽  
Cooling Flows ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract The first stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics are not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The test section involves an engine-representative combustor-turbine interface geometry, combustor coolant flow and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

Aero-thermal Aspects of Film Cooled Nozzle Guide Vane Endwall - Part 1: Aerodynamics

2021 ◽  
pp. 1-54
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong W. Kim ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Three Dimensional ◽  
Leading Edge ◽  
Secondary Flows ◽  
Protective Film ◽  
Cooling Flows ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract The first stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics are not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The test section involves an engine-representative combustor-turbine interface geometry, combustor coolant flow and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

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