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 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

1991 ◽  
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 (1991). 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% 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.

Measurement and Calculation of End Wall Heat Transfer and Aerodynamics on a Nozzle Guide Vane in Annular Cascade

1990 ◽  
Author(s):  
N. W. Harvey ◽  
T. V. Jones
Keyword(s):  
Heat Transfer ◽  
Guide Vane ◽  
Three Dimensional ◽  
Surface Flow ◽  
Wall Heat Transfer ◽  
Transfer Rates ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Annular Cascade ◽  
End Wall

Detailed measurements of surface static pressures and heat transfer rates on the aerofoil and hub end wall of an annular nozzle guide vane (in the absence of a downstream rotor) are presented. Heat transfer rates have been measured using thin film gauges in an annular cascade in the Pyestock Isentropic Light Piston Cascade. Test Mach numbers, Reynolds numbers and cascade geometry are fully representative of engine conditions. The results of 3-D calculations of surface Mach number and 2-D calculations of aerofoil heat transfer are presented and compared with the measurements. A new method of calculating end wall heat transfer using the axisymmetric analogue for three-dimensional boundary layers is described in detail. The method uses a 3-D Euler solver to calculate the inviscid surface streamlines along which heat transfer coefficients are calculated. The metric coefficient which describes the lateral convergence or divergence of the streamlines is used to include three-dimensional effects in the calculation. The calculated heat transfer rates compare well with the measured values. Reference is made to surface flow visualization in the interpretation of the results.

Three-Dimensional Navier-Stokes Prediction of Heat Transfer With Film Cooling

1994 ◽  
Author(s):  
J. M. Fougères ◽  
R. Helder
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Stokes Equations ◽  
Guide Vane ◽  
Three Dimensional ◽  
Hole Injection ◽  
Navier Stokes ◽  
Navier Stokes Equations ◽  
Transfer Rates ◽  
Nozzle Guide

Three-dimensional Navier-Stokes calculations have been performed on various geometries in the presence of discrete-hole injection. The quality of the aerodynamic and thermal predictions of the flow is assessed by comparison to experiments. The code used for the calculations is developed at ONERA and has previously been presented by various authors (Billonnet et al., 1992). It solves the unsteady set of three-dimensional Navier-Stokes equations, completed by a mixing-length turbulence model, using a finite volume technique. The multi-domain approach of the code has facilitated the treatment of this type of geometry. The injection holes are discretized on cylindrical subdomains which overlap the mesh of the main flow. Two applications of the code are presented in this paper. First, a calculation was performed on a row of hot jets injected into a flat plate turbulent boundary layer. Secondly, the code was tested on a plane nozzle guide vane grid with multiple injections. Heat transfer rates, temperature and velocity profiles are compared to experimental data.

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.

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.

scholarly journals Wake Region Measurements of a Highly Three-Dimensional Nozzle Guide Vane Tested at DRA Pyestock Using Particle Image Velocimetry

1994 ◽  
Author(s):  
M. Funes-Gallanzi ◽  
P. J. Bryanston-Cross ◽  
K. S. Chana
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Guide Vane ◽  
Three Dimensional ◽  
Light Sheet ◽  
Test Facility ◽  
Velocity Mapping ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Annular Cascade

The quantitative whole field flow visualization technique of PIV has over the last few years been successfully demonstrated for transonic flow applications. A series of such measurements has been made at DRA Pyestock. Several of the development stages critical to a full engine application of the work have now been achieved using the Isentropic Light Piston Cascade (ILPC) test facility operating with high inlet turbulence levels: • A method of seeding the flow with 0.5μm diameter styrene particles has provided an even coverage of the flow field. • A method of projecting a 1 mm thick high power Nd/YAG laser light sheet within the turbine stator cascade. This has enabled a complete instantaneous intra-blade velocity mapping of the flow field to be visualized, by a specially developed diffraction-limited optics arrangement. • Software has been developed to automatically analyze the data. Due to the sparse nature of the data obtained, a spatial approach to the extraction of the velocity vector data was employed. • Finally, a comparison of the experimental results with those obtained from a three-dimensional viscous flow program of Dawes; using the Baldwin-Lomax model for eddy viscosity and assuming fully turbulent flow. The measurements provide an instantaneous quantitative whole field visualization of a high-speed unsteady region of flow in a highly three-dimensional nozzle guide vane; which has been successfully compared with a full viscous calculation. This work represents the first such measurements to be made in a full-size transonic annular cascade at engine representative conditions.

Conjugate Heat Transfer Analysis on the Interior Surface of Nozzle Guide Vane with Combined Impingement and Film Cooling

2020 ◽  
Vol 37 (4) ◽  
pp. 327-342
Author(s):  
Arun Kumar Pujari ◽  
B. V. S. S. S Prasad ◽  
Nekkanti Sitaram
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Surface Temperature ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Guide Vane ◽  
Internal Surface ◽  
Conducting Material ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

AbstractThe effect of conjugate heat transfer is investigated on a first stage nozzle guide vane (NGV) of a high pressure gas turbine which has both impingement and film cooling holes. The study is carried out computationally by considering a linear cascade domain, having two passages formed between the vanes, with a chord length of 228 mm and spacing of 200 mm. The effect of (i) coolant and mainstream Reynolds numbers, (ii) thermal conductivity (iii) temperature difference between the mainstream and coolant at the internal surface of the nozzle guide vane are investigated under conjugate thermal condition. The results show that, with increasing coolant Reynolds number the lower conducting material shows larger percentage decrease in surface temperature as compared to the higher conducting material. However, the internal surface temperature is nearly independent of mainstream Reynolds number variation but shows significant variation for higher conducting material. Further, the temperature gradient within the solid thickness of NGV is higher for the lower conductivity material.

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.

Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwalls: Part 2 — Thermal Measurements

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

Abstract Flow over gas turbine endwalls is complex and highly three-dimensional. As boundaries for modern engine designs are pushed, this already-complex flow is affected by aggressive application of film cooling flows that actively interact. This two-part study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The approach flow conditions represent flow exiting a low-NOx combustor. The test section includes geometric and cooling details of a combustor-turbine interface in addition to endwall film cooling flows injected upstream of the passage. The first part of this study describes in detail, the passage aerodynamics as affected by injection of cooling flows. It reveals a system of secondary flows, including the newly-discovered Impingement Vortex, which redefines our understanding of the aerodynamics of flow in a modern, film-cooled, first-stage vane row. The second part investigates, through thermal measurements, the distribution, mixing and disruption of cooling flows over the endwall. Measurements are made with and without active endwall film cooling. Descriptions are made through adiabatic surface effectiveness measurements and correlations with in-passage velocity (presented in part one) and thermal fields. Results show that the newly-discovered impingement vortex has a positive effect on coolant distribution through passage vortex suppression and by carrying the coolant to hard-to-cool regions in the passage, including the pressure surface near the endwall.

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.

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