Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade

2012 ◽  
Author(s):  
Lamyaa A. El-Gabry ◽  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Co2 Concentration ◽  
Pressure Side ◽  
Trailing Edge ◽  
Nozzle Guide

An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge can be used to visualize the mixing of the coolant flow with the mainstream. Flow field measurements are performed in the downstream plane with a 5-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side of the vane trailing edge where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge does not reach the pressure side endwall, potentially creating a local hotspot.

Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade

2015 ◽  
Vol 137 (8) ◽  
Author(s):  
Lamyaa A. El-Gabry ◽  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten Fransson
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Co2 Concentration ◽  
Horseshoe Vortex ◽  
Mainstream Flow ◽  
Nozzle Guide

An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes (NGVs) to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge (TE) can be used to visualize the mixing of the coolant flow with the mainstream. Flow field measurements are performed in the downstream plane with a five-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side (PS) of the vane TE where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge (LE) does not reach the PS endwall, potentially creating a local hotspot.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Nozzle Guide ◽  
Cooling Air ◽  
Momentum Boundary Layer

Tests on a specifically designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of laser Doppler measurements. Cases with and without coolant ejection allowed us to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic liquid crystals (TLC) were used to map the adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

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.

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.

Flow Field Measurements on a Large Scale Turbine Cascade With Leading Edge Film Cooling by Two Rows of Holes

1997 ◽  
Author(s):  
Sabine Ardey ◽  
Leonhard Fottner
Keyword(s):  
Flow Field ◽  
Film Cooling ◽  
Gas Turbines ◽  
Large Scale ◽  
Vortex Pair ◽  
Leading Edge ◽  
Field Measurements ◽  
Suction Side ◽  
Pressure Side ◽  
Turbine Cascade

To increase the understanding of the aerodynamic processes dominating the flow field of turbine bladings with leading edge film cooling, isothermal investigations were carried out on a large scale high pressure turbine cascade. Near the stagnation point the blades are equipped with one row of film cooling holes on the suction side and one on the pressure side. Blowing ratio, turbulence intensity, Mach number, and Reynolds number are set to values typically found in modern gas turbines. Experimental data of the cascade flow were obtained by pneumatic probes and static pressure tappings. The flow field was visualized by Schlieren and oil flow techniques. For detailed investigations near the blowing holes the Laser Transit Velocimetry and the three dimensional Hot Wire Anemometry were used. The flow field measurements in the near hole region of the suction side show the typical kidney shaped vortex pair. A local suction peak on the pressure side causes a large recirculation area behind the holes on the pressure side and induces separation bubbles in between the pressure side holes. This leads to the generation of two pairs of vortices: The kidney-vortex is located on top of a second vortex pair and a trough flow that fills up the deficit of the recirculation. Thus the film cooling air is detached from the pressure side surface. In addition to the mean flow vectors Reynolds stress components are a good means to judge the propagation of the jet. In spite of the complex flow pattern occurring on each single jet, the surveyed loss-increase due to the leading edge blowing can be predicted by the mixing layer model.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

2011 ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Nozzle Guide ◽  
Cooling Air ◽  
Momentum Boundary Layer

Tests on a specific designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of Laser Doppler measurements. Cases with and without coolant ejection allowed to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic Liquid Crystals (TLC) were used to map adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

scholarly journals Aerothermal Investigations of Mixing Flow Phenomena in Case of Radially Inclined Ejection Holes at the Leading Edge

1999 ◽  
Author(s):  
Dieter E. Bohn ◽  
Karsten A. Kusterer
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Gas Turbines ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Fluid ◽  
Flow Phenomena ◽  
Vortex Systems

A leading edge cooling configuration is investigated numerically by application of a 3-D conjugate fluid flow and heat transfer solver, CHT-Flow. The code has been developed at the Institute of Steam and Gas Turbines, Aachen University of Technology. It works on the basis of an implicit finite volume method combined with a multi-block technique. The cooling configuration is an axial turbine blade cascade with leading edge ejection through two rows of cooling holes. The rows are located in the vicinity of the stagnation line, one row is on the suction side, the other row is on the pressure side. The cooling holes have a radial ejection angle of 45°. This configuration has been investigated experimentally by other authors and the results have been documented as a test case for numerical calculations of ejection flow phenomena. The numerical domain includes the internal cooling fluid supply, the radially inclined holes and the complete external flow field of the turbine vane in a high resolution grid. Periodic boundary conditions have been used in the radial direction. Thus, end wall effects have been excluded. The numerical investigations focus on the aerothermal mixing process in the cooling jets and the impact on the temperature distribution on the blade surface. The radial ejection angles lead to a fully three dimensional and asymmetric jet flow field. Within a secondary flow analysis it can be shown that complex vortex systems are formed in the ejection holes and in the cooling fluid jets. The secondary flow fields include asymmetric kidney vortex systems with one dominating vortex on the back side of the jets. The numerical and experimental data show a good agreement concerning the vortex development. The phenomena on the suction side and the pressure side are principally the same. It can be found that the jets are barely touching the blade surface as the dominating vortex transports hot gas under the jets. Thus, the cooling efficiency is reduced.

Computational Design and Experimental Evaluation of Using a Leading Edge Fillet on a Gas Turbine Vane

2001 ◽  
Author(s):  
G. A. Zess ◽  
K. A. Thole
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Gradient ◽  
Gas Turbine ◽  
Total Pressure ◽  
Guide Vane ◽  
Leading Edge ◽  
Field Measurements ◽  
Computational Design ◽  
Horseshoe Vortex

With the desire for increased power output for a gas turbine engine comes the continual push to achieve higher turbine inlet temperatures. Higher temperatures result in large thermal and mechanical stresses particularly along the nozzle guide vane. One critical region along a vane is the leading edge-endwall juncture. Based on the assumption that the approaching flow to this juncture is similar to a two-dimensional boundary layer, previous studies have shown that a horseshoe vortex forms. This vortex forms because of a radial total pressure gradient from the approaching boundary layer. This paper documents the computational design and experimental validation of a fillet placed at the leading edge-endwall juncture of a guide vane to eliminate the horseshoe vortex. The fillet design effectively accelerated the incoming boundary layer thereby mitigating the effect of the total pressure gradient. To verify the CFD studies used to design the leading edge fillet, flow field measurements were performed in a large-scale, linear, vane cascade. The flow field measurements were performed with a laser Doppler velocimeter in four planes orientated orthogonal to the vane. Good agreement between the CFD predictions and the experimental measurements verified the effectiveness of the leading edge fillet at eliminating the horseshoe vortex. The flowfield results showed that the turbulent kinetic energy levels were significantly reduced in the endwall region because of the absence of the unsteady horseshoe vortex.

scholarly journals Thermal Fatigue Life Prediction of Thermal Barrier Coat on Nozzle Guide Vane via Master–Slave Model

2019 ◽  
Vol 9 (20) ◽  
pp. 4357 ◽  
Author(s):  
Peng Guan ◽  
Yanting Ai ◽  
Chengwei Fei ◽  
Yudong Yao
Keyword(s):  
Fatigue Life ◽  
Thermal Fatigue ◽  
Life Prediction ◽  
Guide Vane ◽  
Leading Edge ◽  
Fatigue Life Prediction ◽  
Trailing Edge ◽  
Thermal Fatigue Life ◽  
Barrier Coat ◽  
Nozzle Guide

The aim of this paper was to develop a master–slave model with fluid-thermo-structure (FTS) interaction for the thermal fatigue life prediction of a thermal barrier coat (TBC) in a nozzle guide vane (NGV). The master–slave model integrates the phenomenological life model, multilinear kinematic hardening model, fully coupling thermal-elastic element model, and volume element intersection mapping algorithm to improve the prediction precision and efficiency of thermal fatigue life. The simulation results based on the developed model were validated by temperature-sensitive paint (TSP) technology. It was demonstrated that the predicted temperature well catered for the TSP tests with a maximum error of less than 6%, and the maximum thermal life of TBC was 1558 cycles around the trailing edge, which is consistent with the spallation life cycle of the ceramic top coat at 1323 K. With the increase of pre-oxidation time, the life of TBC declined from 1892 cycles to 895 cycles for the leading edge, and 1558 cycles to 536 cycles for the trailing edge. The predicted life of the key points at the leading edge was longer by 17.7–40.1% than the trailing edge. The developed master–slave model was validated to be feasible and accurate in the thermal fatigue life prediction of TBC on NGV. The efforts of this study provide a framework for the thermal fatigue life prediction of NGV with TBC.

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.

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