Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade: Part 1—Experimental Technique and Results

1992 ◽  
Vol 114 (4) ◽  
pp. 734-740 ◽  
Author(s):  
S. P. Harasgama ◽  
C. D. Burton
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Suction Side ◽  
Coolant Temperature ◽  
Guide Vanes ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

Heat transfer and aerodynamic measurements have been made on the endwalls of an annular cascade of turbine nozzle guide vanes in the presence of film cooling. The results indicate that high levels of cooling effectiveness can be achieved on the endwalls of turbine nozzle guide vanes (NGV). The NGV were operated at the correct engine nondimensional conditions of Reynolds number, Mach number, gas-to-wall temperature ratio, and gas-to-coolant density ratio. The results show that the secondary flow and horseshoe vortex act on the coolant, which is convected toward the suction side of the NG V endwall passage. Consequently the coolant does not quite reach the pressure side/casing trailing edge, leading to diminished cooling in this region. Increasing the blowing rate from 0.52 to 1.1 results in significant reductions in heat transfer to the endwall. Similar trends are evident when the coolant temperature is reduced. Measured heat transfer rates indicate that over most of the endwall region the film cooling reduces the Nusselt number by 50 to 75 percent.

Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade: Part 1 — Experimental Technique and Results

1991 ◽  
Author(s):  
S. P. Harasgama ◽  
C. D. Burton
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Suction Side ◽  
Coolant Temperature ◽  
Guide Vanes ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

Heat transfer and aerodynamic measurements have been made on the endwalls of an annular cascade of turbine nozzle guide vanes in the presence of film cooling. The results indicate that high levels of cooling effectiveness can be achieved on the endwalls of turbine nozzle guide vanes (NGV). The NGV were operated at the correct engine non-dimensional conditions of Reynolds number, Mach number, gas-to-wall temperature ratio and gas-to-coolant density ratio. The results show that the secondary flow and horse-shoe vortex act on the coolant which is converted towards the suction side of the NGV endwall passage. Consequently the coolant does not quite reach the pressure side/casing trailing edge, leading to diminished cooling in this region. Increasing the blowing rate from 0.52 to 1.1 results in significant reductions in heat transfer to the endwall. Similar trends are evident when the coolant temperature is reduced. Measured heat transfer rates indicate that over most of the endwall region the film cooling reduces the Nusselt number by 50% to 75%.

Heat Transfer Analysis of the Surface of Nonfilm-Cooled and Film-Cooled Nozzle Guide Vanes in Transonic Annular Cascade

2017 ◽  
Author(s):  
Kasem E. Ragab ◽  
Lamyaa El-Gabry
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Guide Vane ◽  
Heat Transfer Analysis ◽  
Guide Vanes ◽  
Cfd Models ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

One of the approaches adopted to improve turbine efficiency and increase power to weight ratio is reducing vane count. In the current study, numerical analysis was performed for the heat transfer over the surface of nozzle guide vanes under the condition of reduced vane count using three dimensional computational fluid dynamics (CFD) models. The investigation has taken place in two stages: the baseline nonfilm-cooled nozzle guide vane, and the film-cooled nozzle guide vane. A finite volume based commercial code (ANSYS CFX 15) was used to build and analyze the CFD models. The investigated annular cascade has no heat transfer measurements available; hence in order to validate the CFD models against experimental data, two standalone studies were carried out on the NASA C3X vanes, one on the nonfilm-cooled C3X vane and the other on the film-cooled C3X vane. Different modelling parameters were investigated including turbulence models in order to obtain good agreement with the C3X experimental data, the same parameters were used afterwards to model the industrial nozzle guide vanes. Three Shear Stress Transport (SST) turbulence model variations were evaluated, the SST with Gamma-Theta transition model was found to yield the best agreement with the experimental results; model capabilities were demonstrated when the laminar to turbulent transition took place.

The Influence of the Boundary Layer State and Reynolds Number on Film Cooling and Heat Transfer on a Cooled Nozzle Guide Vane

2000 ◽  
Author(s):  
Hans Reiss ◽  
Albin Bölcs
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Flow Conditions ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Film cooling and heat transfer measurements were carried out on a cooled nozzle guide vane in a linear cascade, using a transient liquid crystal technique. Three flow conditions were realized: the nominal operating condition of the vane with an exit Reynolds number of 1.47e6, as well as two lower flow conditions: Re2L = 1.0e6 and 7.5e5. The vane model was equipped with a single row of inclined round film cooling holes with compound angle orientation on the suction side. Blowing ratios ranging form 0.3 to 1.5 were covered, all using foreign gas injection (CO2) yielding an engine-representative density ratio of 1.6. Two distinct states of the incoming boundary layer onto the injection station were compared, an undisturbed laminar boundary layer as it forms naturally on the suction side, and a fully turbulent boundary layer which was triggered with a trip wire upstream of injection. The aerodynamic flow field is characterized in terms of profile Mach number distribution, and the associated heat transfer coefficients around the uncooled airfoil are presented. Both detailed and spanwise averaged results of film cooling effectiveness and heat transfer coefficients are shown on the suction side, which indicate considerable influence of the state of the incoming boundary layer on the performance of a film cooling row. The influence of the mainstream flow condition on the film cooling behavior at constant blowing ratio is discussed for three chosen injection regimes.

A Converging Slot-Hole Film-Cooling Geometry—Part 2: Transonic Nozzle Guide Vane Heat Transfer and Loss

2002 ◽  
Vol 124 (3) ◽  
pp. 461-471 ◽  
Author(s):  
J. E. Sargison ◽  
S. M. Guo ◽  
M. L. G. Oldfield ◽  
G. D. Lock ◽  
A. J. Rawlinson
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Guide Vane ◽  
Density Ratio ◽  
Wide Band ◽  
Nozzle Guide Vane ◽  
Cooling Hole ◽  
Hole Geometry ◽  
Temperature Levels ◽  
Nozzle Guide

This paper presents the first experimental measurements on an engine representative nozzle guide vane, of a new film-cooling hole geometry, a con¯vergings¯lot-hole¯ or console. The patented console geometry is designed to improve the heat transfer and aerodynamic performance of turbine vane and rotor blade cooling systems. These experiments follow the successful validation of the console design in low-speed flat-plate tests described in Part 1 of this paper. Stereolithography was used to manufacture a resin model of a transonic, engine representative nozzle guide vane in which seven rows of previously tested fan-shaped film-cooling holes were replaced by four rows of consoles. This vane was mounted in the annular vane ring of the Oxford cold heat transfer tunnel for testing at engine Reynolds numbers, Mach numbers and coolant to mainstream momentum flux ratios using a heavy gas to simulate the correct coolant to mainstream density ratio. Heat transfer data were measured using wide-band thermochromic liquid crystals and a modified analysis technique. Both surface heat transfer coefficient and the adiabatic cooling effectiveness were derived from computer-video records of hue changes during the transient tunnel run. The cooling performance, quantified by the heat flux at engine temperature levels, of the console vane compares favourably with that of the previously tested vane with fan-shaped holes. The new console film-cooling hole geometry offers advantages to the engine designer due to a superior aerodynamic efficiency over the fan-shaped hole geometry. These efficiency measurements are demonstrated by results from midspan traverses of a four-hole pyramid probe downstream of the nozzle guide vane.

scholarly journals Endwall Heat Transfer and Aerodynamic Measurements in an Annular Cascade of Nozzle Guide Vanes

1995 ◽  
Author(s):  
M. C. Spencer ◽  
G. D. Lock ◽  
T. V. Jones ◽  
N. W. Harvey
Keyword(s):  
Heat Transfer ◽  
Mach Number ◽  
Reynolds Numbers ◽  
Flow Visualisation ◽  
Surface Shear ◽  
Guide Vanes ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes ◽  
Transient Liquid Crystal Technique

Aerodynamic and heat transfer measurements have been made on the hub and casing endwalls of an annular cascade of high pressure nozzle guide vanes. The measurements have been made over a range of engine representative Mach and Reynolds numbers and with large levels of freestream turbulence intensity. The transient liquid crystal technique has been employed, which has the advantage of yielding full surface maps of heat transfer coefficient. Computational predictions and aerodynamic measurements of Mach number distributions on the endwall surfaces are also presented, along with surface-shear flow visualisation using oil and dye techniques. The heat transfer results are discussed and interpreted in terms of the secondary flow and Mach number patterns.

scholarly journals Efficiency Measurements of an Annular Nozzle Guide Vane Cascade With Different Film Cooling Geometries

1998 ◽  
Author(s):  
Charles R. B. Day ◽  
Martin L. G. Oldfield ◽  
Gary D. Lock ◽  
Stephen N. Dancer
Keyword(s):  
Film Cooling ◽  
Total Pressure ◽  
Guide Vane ◽  
Cumulative Effect ◽  
Aerodynamic Efficiency ◽  
Nozzle Guide Vane ◽  
Density Ratios ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
Nozzle Guide Vanes

This paper further extends the research reported by Day et al. (1997), which reported aerodynamic efficiency measurements on an annular cascade of engine representative transonic nozzle guide vanes with extensive film cooling. This work compares the measured aerodynamic efficiencies of blades with 14 rows of cylindrical cooling holes with a new geometry in which 8 of the rows have been replaced by holes having a fan-shaped exit geometry. The effects of adding trailing edge slot ejection are also presented. By selectively blocking rows of holes, the cumulative effect on the mid-span efficiency of adding rows of cooling holes has also been determined. A dense foreign gas (SF6/Ar mixture) is used to simulate engine representative coolant-to-mainstream density ratios, momentum ratios and blowing rates under ambient temperature 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. Experimental results are presented as area traverse maps (total pressure, isentropic Mach number and flow angles), from which the incremental changes in efficiency due to film cooling have been calculated. The effects of different assumptions for the coolant total pressure are shown. Experimental data agrees reasonably well with loss predictions using a Hartsel model.

Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade: Part 2—Analysis and Correlation of Results

1992 ◽  
Vol 114 (4) ◽  
pp. 741-746 ◽  
Author(s):  
S. P. Harasgama ◽  
C. D. Burton
Keyword(s):  
Heat Transfer ◽  
Pressure Gradient ◽  
Film Cooling ◽  
Guide Vane ◽  
Three Dimensional ◽  
Companion Paper ◽  
Hole Injection ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Annular Cascade

Results have been presented on the heat transfer characteristics of the film cooled endwall (platform) of a turbine nozzle guide vane in an annular cascade at engine representative conditions in a companion paper by Harasgama and Burton (1992). The present paper reports on the analysis of these measurements. The experimental results are well represented by the superposition theory of film cooling. It is shown that high cooling effectiveness can be achieved when the data are corrected for axial pressure gradients. The data are correlated against both the slot-wall jet parameter and the discrete hole injection function for flat-plate, zero pressure gradient cases. The pressure gradient correction brings the present data to within ± 11 percent of the discrete hole correlation. Preliminary predictions of heat transfer reduction have been carried out using the STANCOOL program. These indicate that the code can predict the magnitude of heat transfer reduction correctly, although the absolute values are not in good agreement. This is attributed to the three-dimensional nature of the flow at the endwall.

scholarly journals Unsteady phenomena at the combustor-turbine interface

2021 ◽  
Vol 5 ◽  
pp. 202-215
Author(s):  
Faisal Shaikh ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
High Speed ◽  
Guide Vane ◽  
Secondary Flows ◽  
Test Facility ◽  
Unsteady Heat Transfer ◽  
Guide Vanes ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

The combustor-turbine interface in a gas turbine is characterised by complex, highly unsteady flows. In a combined experimental and large eddy simulation (LES) study including realistic combustor geometry, the standard model of secondary flows in the nozzle guide vanes (NGV) is found to be oversimplified. A swirl core is created in the combustion chamber which convects into the first vane passages. Four main consequences of this are identified: variation in vane loading; unsteady heat transfer on vane surfaces; unsteadiness at the leading edge horseshoe vortex, and variation in the position of the passage vortex. These phenomena occur at relatively low frequencies, from 50–300 Hz. It seems likely that these unsteady phenomena result in non-optimal film cooling, and that by reducing unsteadiness designs with greater cooling efficiency could be achieved. Measurements were performed in a high speed test facility modelling a large industrial gas turbine with can combustors, including nozzle guide vanes and combustion chambers. Vane surfaces and endwalls of a nozzle guide vane were instrumented with 384 high speed thin film heat flux gauges, to measure unsteady heat transfer. The high resolution of measurements was such to allow direct visualisation in time of large scale turbulent structures over the endwalls and vane surfaces. A matching LES simulation was carried out in a domain matching experimental conditions including upstream swirl generators and transition duct. Data reduction allowed time-varying LES data to be recorded for several cycles of the unsteady phenomena observed. The combination of LES and experimental data allows physical explanation and visualisation of flow events.

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