scholarly journals The Influence of Film-Cooling on the Aerodynamic Performance of a Turbine Nozzle Guide Vane

1997 ◽  
Author(s):  
C. Osnaghi ◽  
A. Perdichizzi ◽  
M. Savini ◽  
P. Harasgama ◽  
E. Lutum
Keyword(s):  
Carbon Dioxide ◽  
Mass Flow ◽  
Film Cooling ◽  
Momentum Flux ◽  
Guide Vane ◽  
Flux Ratio ◽  
Aerodynamic Performance ◽  
Trailing Edge ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

The paper presents the results of an investigation on the aerodynamic performance of a full coverage film-cooled nozzle guide vane. The blading is a typical high pressure turbine vane of advanced design, working in the high subsonic regime. Tests have been carried out for a wide range of conditions, including variations in Mach number, coolant to mainstream mass flow rate ratio and location of the coolant injection. Both air and carbon dioxide at ambient conditions have been utilized, as coolant flow. Measurements have been performed in a plane located at 0.5 axial chord downstream of the trailing edge by means of a miniaturized five-hole pressure probe. Performances, in terms of losses, flow angles and profile pressure distributions, for different cooling mass flow rates are presented and compared to the results of the solid blade tests (i.e. with no cooling holes). The results showed a significant increase of the losses with blowing. Test with air and carbon dioxide provided almost equal losses if carried out at the same global momentum flux ratio; however the density ratio was found to influence slightly the share of the coolant fluid among the injection rows and the local momentum flux ratio as well. In order to define the individual contributions of groups of cooling rows on the performance of the blade, three different modes of injection have been tested, namely full, trailing edge and shower head injection. The main trend observed is that trailing edge injection produces the least amount of additional losses at high blowing rates. Full-coverage film-cooling injection did not lead to marked variations in the blade pressure distribution and/or outlet flow angle.

Experimental and Computational Study of the Effect of Momentum-Flux Ratio on High-Pressure Nozzle Guide Vane Endwall Cooling Systems

2017 ◽  
Vol 139 (12) ◽  
Author(s):  
Francesco Ornano ◽  
Thomas Povey
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Guide Vane ◽  
Flux Ratio ◽  
Secondary Flows ◽  
Cooling Systems ◽  
Momentum Flux Ratio ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Cooling

High-pressure (HP) nozzle guide vane (NGV) endwalls are often characterized by highly three-dimensional (3D) flows. The flow structure depends on the incoming boundary layer state (inlet total pressure profile) and the (static) pressure gradients within the vane passage. In many engine applications, this can lead to strong secondary flows. The prediction and design of optimized endwall film cooling systems is therefore challenging and is a topic of current research interest. A detailed experimental investigation of the film effectiveness distribution on an engine-realistic endwall geometry is presented in this paper. The film cooling system was a fairly conventional axisymmetric double-row configuration. The study was performed on a large-scale, low-speed wind tunnel using infrared (IR) thermography. Adiabatic film effectiveness distributions were measured using IR cameras, and tests were performed across a wide range of coolant-to-mainstream momentum-flux and mass flow ratios (MFRs). Complex interactions between coolant film and vane secondary flows are presented and discussed. A particular feature of interest is the suppression of secondary flows (and associated improved adiabatic film effectiveness) beyond a critical momentum flux ratio. Jet liftoff effects are also observed and discussed in the context of sensitivity to local momentum flux ratio. Full coverage experimental results are also compared to 3D, steady-state computational fluid dynamics (CFD) simulations. This paper provides insights into the effects of momentum flux ratio in establishing similarity between cascade conditions and engine conditions and gives design guidelines for engine designers in relation to minimum endwall cooling momentum flux requirements to suppress endwall secondary flows.

Application of Unsteady CFD Methods to Trailing Edge Cutback Film Cooling

2014 ◽  
Author(s):  
S. Ravelli ◽  
G. Barigozzi
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Navier Stokes ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Unsteady Rans ◽  
Exit Mach Number ◽  
Nozzle Guide

The main purpose of this numerical investigation is to overcome the limitations of the steady modeling in predicting the cooling efficiency over the cutback surface in a high pressure turbine nozzle guide vane. Since discrepancy between Reynolds-averaged Navier–Stokes (RANS) predictions and measured thermal coverage at the trailing edge was attributable to unsteadiness, Unsteady RANS (URANS) modeling was implemented to evaluate improvements in simulating the mixing between the mainstream and the coolant exiting the cutback slot. With the aim of reducing the computation effort, only a portion of the airfoil along the span was simulated at an exit Mach number of Ma2is = 0.2. Three values of the coolant-to-mainstream mass flow ratio were considered: MFR = 0.66%, 1.05%, and 1.44%. Nevertheless the inherent vortex shedding from the cutback lip was somehow captured by the URANS method, the computed mixing was not enough to reproduce the measured drop in adiabatic effectiveness η along the streamwise direction, over the cutback surface. So modeling was taken a step further by using the Scale Adaptive Simulation (SAS) method at MFR = 1.05%. Results from the SAS approach were found to have potential to mimic the experimental measurements. Vortices shedding from the cutback lip were well predicted in shape and magnitude, but with a lower frequency, as compared to PIV data and flow visualizations. Moreover, the simulated reduction in film cooling effectiveness toward the trailing edge was similar to that observed experimentally.

Research of Aerodynamic Performance of HP-Turbine with Coolant Injections for Variable Cycle Engine

2011 ◽  
Vol 110-116 ◽  
pp. 1047-1053
Author(s):  
Zhi Gang Liu ◽  
Xiang Jun Fang ◽  
Si Yong Liu ◽  
Ping Wang ◽  
Zhao Yin
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Supersonic Nozzle ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Thermodynamic Cycle ◽  
Aerodynamic Performance ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Transonic Rotor

A highly loaded high-pressure turbine with a supersonic nozzle guide vane and a transonic rotor for a Variable Cycle Engine (VCE) has been investigated. Film cooling strategies were designed for the whole stage, during which the positions, injection orientations and arrangements of cooling holes were confirmed. Three-dimensional steady numerical simulations have been performed in the two operation modes of low and high bypass ratio with different thermodynamic cycle parameters according to the VCE and the coolant injections have been simulated by means of additional source term method. The influences of coolant injections in the fully cooled turbine stage on aerodynamic performance and flow characteristics have been analyzed. The results indicate that, the supersonic nozzle guide vane, over-expansion degree of main flows, fluctuations of static pressure and intensity of corner vortex are lessened or alleviated. In the transonic rotor, expansion and doing work capabilities in the mixed fluid are strengthened. Proper coolants injections are beneficial to the flow characteristics in the blade passage.

The Effects of Slot Film Cooling on Deposition on a Nozzle Guide Vane

2014 ◽  
Author(s):  
Robin Prenter ◽  
Steven M. Whitaker ◽  
Ali Ameri ◽  
Jeffrey P. Bons
Keyword(s):  
Surface Temperature ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Guide Vane ◽  
Capture Efficiency ◽  
Coolant Flow Rate ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Slot Film Cooling

The effects of slot film cooling on deposition in a high pressure nozzle guide vane passage were investigated experimentally and computationally. Experiments were conducted in Ohio State’s Turbine Reaction Flow Rig, using a four-vane cascade, operating at temperatures up to 1353 K. Film cooling was achieved on one of the vanes using a span-wise slot, located at approximately 30% chord on the pressure surface. The coolant’s effect on vane surface temperature was characterized by taking infrared images at various cooling levels. Deposition was produced by injecting sub-bituminous ash particles with a median diameter of 6.48 μm upstream of the vane passage. Several deposition tests were conducted with varying coolant levels. Results exhibit a strong relationship between the coolant flow rate and the amount of ash that deposits on the cooled vane. Capture efficiency was reduced by 70% at the highest coolant flow rate (1.27% of the mass flow rate in the passage). Capture efficiency reduction was compared to that achieved using discrete hole film cooling in other studies. The slot scheme showed similar or larger reductions in capture efficiency at lower coolant mass flow rates. Deposit distribution patterns are affected by regions of cooler temperature, both downstream of the slot where film effects dominate, and slightly upstream of the slot which is cooled by conduction. A computational simulation was conducted to model both the flow and deposition. The solid vane was also discretized to allow for conjugate heat transfer calculations, which produced results that were qualitatively similar to IR measurements, but over predicted the effectiveness of the coolant. An Eulerian-Lagrangian particle tracking model was utilized to track the ash particles through the flow. A sticking model was implemented to determine whether particles stick upon impacting the vane surface, from which deposition rates and distributions are obtained. The computational model under predicted the baseline capture efficiency and the capture efficiency reduction factors for each cooling level, suggesting that the model is not sufficiently sensitive to the temperature changes between tests. Inclusion of surface temperature and local shear dependencies was suggested as an improvement to the sticking model.

Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance

2014 ◽  
Vol 136 (11) ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten H. Fransson ◽  
Boris I. Mamaev ◽  
Mats Annerfeldt ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Trailing Edge ◽  
Aerodynamic Loss ◽  
Nozzle Guide Vane ◽  
Turbine Vane ◽  
Engine Component ◽  
Annular Sector ◽  
Exit Mach Number ◽  
Nozzle Guide

An experimental investigation on a cooled nozzle guide vane (NGV) has been conducted in an annular sector to quantify aerodynamic influences of shower head (SH) and trailing edge (TE) cooling. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at a reference exit Mach number of 0.89. The investigations have been performed for various coolant-to-mainstream mass–flux ratios. New loss equations are derived and implemented regarding coolant aerodynamic losses. Results lead to a conclusion that both TE cooling and SH film cooling increase the aerodynamic loss compared to an uncooled case. In addition, the TE cooling has higher aerodynamic loss compared to the SH cooling. Secondary losses decrease with inserting SH film cooling compared to the uncooled case. The TE cooling appears to have less impact on the secondary loss compared to the SH cooling. Area-averaged exit flow angles around midspan increase for the TE cooling.

Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance

2014 ◽  
Author(s):  
Ranjan Saha ◽  
Jens Fridh ◽  
Torsten H. Fransson ◽  
Boris I. Mamaev ◽  
Mats Annerfeldt ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Trailing Edge ◽  
Aerodynamic Loss ◽  
Head Cooling ◽  
Nozzle Guide Vane ◽  
Turbine Vane ◽  
Annular Sector ◽  
Exit Mach Number ◽  
Nozzle Guide

An experimental investigation on a cooled nozzle guide vane has been conducted in an annular sector to quantify aerodynamic influences of shower head and trailing edge cooling. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at a reference exit Mach number of 0.89. The investigations have been performed for various coolant-to-mainstream mass-flux ratios. New loss equations are derived and implemented regarding coolant aerodynamic losses. Results lead to a conclusion that both trailing edge cooling and shower head film cooling increase the aerodynamic loss compared to an uncooled case. In addition, the trailing edge cooling has higher aerodynamic loss compared to the shower head cooling. Secondary losses decrease with inserting shower head film cooling compared to the uncooled case. The trailing edge cooling appears to have less impact on the secondary loss compared to the shower head cooling. Area-averaged exit flow angles around midspan increase for the trailing edge cooling.

The Influence of Film Cooling on the Efficiency of an Annular Nozzle Guide Vane Cascade

1997 ◽  
Author(s):  
C. R. B. Day ◽  
M. L. G. Oldfield ◽  
G. D. Lock
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Guide Vane ◽  
Surface Film ◽  
Flux Ratio ◽  
Ambient Conditions ◽  
Measure Concentration ◽  
Momentum Flux Ratio ◽  
Design Value ◽  
Nozzle Guide

This paper examines the effect of aerofoil surface film cooling on the aerodynamic efficiency of an annular cascade of transonic nozzle guide vanes. A dense foreign gas (SF6/Ar mixture) is used to simulate engine representative coolant-to-mainstream density ratios under ambient conditions. The flowfield measurements have been obtained using a four-hole pyramid probe in a short duration blowdown facility which correctly models engine Reynolds and Mach numbers, as well as the inlet turbulence intensity. The use of foreign gas coolant poses specific challenges not present in an air-cooled cascade, and this paper addresses two. Firstly, a novel method for the determination of mass flow from pneumatic probe data in a heterogeneous gas environment is presented which eliminates the need to measure concentration in order to determine loss. Secondly, the authors argue on the grounds of dimensionless similarity that momentum flux ratio is to be preferred to blowing rate for the correct parametrisation of film cooling studies with varying coolant densities. Experimental results are presented as area traverse maps, from which values for loss have been calculated. It is shown that air and foreign gas at the same momentum flux ratio give very similar results, and that the main difference between cooled and uncooled configurations is an increase in wake width. Interestingly, it is shown that an increase in the momentum flux ratio above the design value with foreign gas coolant reduces the overall loss compared with the design value. The data has been obtained both for purposes of design and for CFD code validation.

Contoured Endwall Flow and Heat Transfer Experiments With Combustor Coolant and Gap Leakage Flows for a Turbine Nozzle Guide Vane

2016 ◽  
Author(s):  
Reema Saxena ◽  
Mahmood H. Alqefl ◽  
Zhao Liu ◽  
Hee-Koo Moon ◽  
Luzeng Zhang ◽  
...  
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Mass Flow ◽  
Film Cooling ◽  
Thermal Loading ◽  
Guide Vane ◽  
Nozzle Guide Vane ◽  
Adiabatic Effectiveness ◽  
Interfacial Gap ◽  
Nozzle Guide

Flow in a high pressure gas turbine passage is complex, involving systems of secondary vortex flows and strong transverse pressure gradients. This complexity causes difficulty in providing film cooling coverage to the hub endwall region, which is subjected to high thermal loading due to combustor exit hot core gases. Therefore, an improved understanding of these flow features and their effects on endwall film cooling is needed to assist designers in developing efficient cooling schemes. The experimental study presented in this paper is performed on a linear, stationary, two-passage cascade representing the first stage nozzle guide vane of a high-pressure gas turbine. The sources of film cooling flows are the upstream combustor liner coolant and the leakage flow from the combustor-nozzle guide vane interfacial gap. Measurements are performed on an axisymmetrically-contoured endwall passage under conditions of various leakage mass flow rates to mainstream flow ratios (MFR= 0.5%, 1.0%, 1.5%). Flow migration and mixing are documented by measuring passage thermal fields and adiabatic effectiveness values over the endwall. It is found that, compared to our previous studies with a rotor inlet leakage slot geometry, the thin slot geometry of the nozzle leakage path gives a more uniform coolant spread over the endwall with significant coverage reaching the downstream and pressure-side regions of the passage. Interestingly, the coverage is seen to be only weakly dependent on the leakage mass low ratio and even reduce slightly with an increase in mass flow ratio above 1%, as indicated by lowered endwall adiabatic effectiveness values.

The Influence of Film Cooling on the Efficiency of an Annular Nozzle Guide Vane Cascade

1999 ◽  
Vol 121 (1) ◽  
pp. 145-151 ◽  
Author(s):  
C. R. B. Day ◽  
M. L. G. Oldfield ◽  
G. D. Lock
Keyword(s):  
Film Cooling ◽  
Momentum Flux ◽  
Guide Vane ◽  
Surface Film ◽  
Flux Ratio ◽  
Ambient Conditions ◽  
Measure Concentration ◽  
Momentum Flux Ratio ◽  
Design Value ◽  
Nozzle Guide

This paper examines the effect of aerofoil surface film cooling on the aerodynamic efficiency of an annular cascade of transonic nozzle guide vanes. A dense foreign gas (SF6/Ar mixture) is used to simulate engine representative coolant-to-mainstream density ratios under ambient conditions. The flowfield measurements have been obtained using a four-hole pyramid probe in a short duration blowdown facility that correctly models engine Reynolds and Mach numbers, as well as the inlet turbulence intensity. The use of foreign gas coolant poses specific challenges not present in an air-cooled cascade, and this paper addresses two. First, a novel method for the determination of mass flow from pneumatic probe data in a heterogeneous gas environment is presented that eliminates the need to measure concentration in order to determine loss. Second, the authors argue on the grounds of dimensionless similarity that momentum flux ratio is to be preferred to blowing rate for the correct parameterization of film cooling studies with varying coolant densities. Experimental results are presented as area traverse maps, from which values for loss have been calculated. It is shown that air and foreign gas at the same momentum flux ratio give very similar results, and that the main difference between cooled and uncooled configurations is an increase in wake width. Interestingly, it is shown that an increase in the momentum flux ratio above the design value with foreign gas coolant reduces the overall loss compared with the design value. The data have been obtained both for purposes of design and for CFD code Validation.

A Method for Correlating the Influence of External Crossflow on the Discharge Coefficients of Film Cooling Holes

2000 ◽  
Vol 123 (2) ◽  
pp. 258-265 ◽  
Author(s):  
D. A. Rowbury ◽  
M. L. G. Oldfield ◽  
G. D. Lock
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Discharge Coefficient ◽  
Guide Vane ◽  
Discharge Coefficients ◽  
Nozzle Guide Vane ◽  
Aero Engine ◽  
Nozzle Guide ◽  
Film Cooling Holes ◽  
Asme Paper

An empirical means of predicting the discharge coefficients of film cooling holes in an operating engine has been developed. The method quantifies the influence of the major dimensionless parameters, namely hole geometry, pressure ratio across the hole, coolant Reynolds number, and the freestream Mach number. The method utilizes discharge coefficient data measured on both a first-stage high-pressure nozzle guide vane from a modern aero-engine and a scale (1.4 times) replica of the vane. The vane has over 300 film cooling holes, arranged in 14 rows. Data was collected for both vanes in the absence of external flow. These noncrossflow experiments were conducted in a pressurized vessel in order to cover the wide range of pressure ratios and coolant Reynolds numbers found in the engine. Regrettably, the proprietary nature of the data collected on the engine vane prevents its publication, although its input to the derived correlation is discussed. Experiments were also conducted using the replica vanes in an annular blowdown cascade which models the external flow patterns found in the engine. The coolant system used a heavy foreign gas (SF6 /Ar mixture) at ambient temperatures which allowed the coolant-to-mainstream density ratio and blowing parameters to be matched to engine values. These experiments matched the mainstream Reynolds and Mach numbers and the coolant Mach number to engine values, but the coolant Reynolds number was not engine representative (Rowbury, D. A., Oldfield, M. L. G., and Lock, G. D., 1997, “Engine-Representative Discharge Coefficients Measured in an Annular Nozzle Guide Vane Cascade,” ASME Paper No. 97-GT-99, International Gas Turbine and Aero-Engine Congress & Exhibition, Orlando, Florida, June 1997; Rowbury, D. A., Oldfield, M. L. G., Lock, G. D., and Dancer, S. N., 1998, “Scaling of Film Cooling Discharge Coefficient Measurements to Engine Conditions,” ASME Paper No. 98-GT-79, International Gas Turbine and Aero-Engine Congress & Exhibition, Stockholm, Sweden, June 1998). A correlation for discharge coefficients in the absence of external crossflow has been derived from this data and other published data. An additive loss coefficient method is subsequently applied to the cascade data in order to assess the effect of the external crossflow. The correlation is used successfully to reconstruct the experimental data. It is further validated by successfully predicting data published by other researchers. The work presented is of considerable value to gas turbine design engineers as it provides an improved means of predicting the discharge coefficients of engine film cooling holes.

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