scholarly journals Experimental Measurement and CFD Prediction of Heat Transfer to a Nozzle Guide Vane

1996 ◽  
Author(s):  
Trevor J. Kirsten ◽  
Andreas M. Lippert ◽  
Glen C. Snedden ◽  
Graham D. J. Smith
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Gas Flow ◽  
Guide Vane ◽  
Surface Heat ◽  
Analytical Prediction ◽  
Surface Heat Transfer ◽  
Short Period ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

The determination of aerothermal loading plays a vital role in any performance or structural investigation of turbine hot-end components during both the design and in-service phases. This investigation considers the measurement and prediction of heat transfer on a nozzle guide vane in a linear cascade. The experimental method is based on a transient approach and requires that the gas and blade have the correct temperature difference. In this study the transient technique is facilitated by maintaining a constant external gas flow in a closed circuit cascade facility and moving a cooled MACOR blade rapidly into the cascade, displacing an aluminium dummy blade. For a short period of time the correct temperature ratio exists between the gas and blade. It is during this period that the change in the thin-film resistance is measured, which is transformed to represent surface heat transfer. The experimental data is compared to an analytical prediction using an Euler code for the bulk flow solution and a boundary layer code for surface heat transfer. The boundary layer code includes the effects of turbulence in the boundary layer, and the transitional nature of the boundary layer flow is also captured. The accuracy associated with the levels of heat transfer is shown to be good for two exit Mach numbers, although some variation is evident towards the trailing edge of the suction surface. The laminar, transitional and turbulent regions of the boundary layer show good correspondence.

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.

Experimental Study of the Endwall Heat Transfer of a Transonic Nozzle Guide Vane With Upstream Jet Purge Cooling: Part 1 — Effect of Density Ratio

2020 ◽  
Author(s):  
Shuo Mao ◽  
Ridge A. Sibold ◽  
Stephen Lash ◽  
Wing F. Ng ◽  
Hongzhou Xu ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Mass Flux ◽  
Guide Vane ◽  
Density Ratio ◽  
Momentum Ratio ◽  
Blowing Ratio ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
The Impact

Abstract Nozzle guide vane platforms often employ complex cooling schemes to mitigate ever-increasing thermal loads on endwall. Understanding the impact of advanced cooling schemes amid the highly complex three-dimensional secondary flow is vital to engine efficiency and durability. This study analyzes and describes the effect of coolant to mainstream blowing ratio, momentum ratio and density ratio for a typical axisymmetric converging nozzle guide vane platform with an upstream doublet staggered, steep-injection, cylindrical hole jet purge cooling scheme. Nominal flow conditions were engine representative and as follows: Maexit = 0.85, Reexit/Cax = 1.5 × 106 and an inlet large-scale freestream turbulence intensity of 16%. Two blowing ratios were investigated, each corresponding to upper and lower engine extrema at M = 3.5 and 2.5, respectively. For each blowing ratio, the coolant to mainstream density ratio was varied between DR = 1.2, representing typical experimental neglect of coolant density, and DR = 1.95, representative of typical engine conditions. An optimal coolant momentum ratio between = 6.3 and 10.2 is identified for in-passage film effectiveness and net heat flux reduction, at which the coolant suppresses and overcomes secondary flows but imparts minimal turbulence and remains attached to endwall. Progression beyond this point leads to cooling effectiveness degradation and increased endwall heat flux. Endwall heat transfer does not scale well with one single parameter; increasing with increasing mass flux for the low density case but decreasing with increasing mass flux of high density coolant. From the results gathered, both coolant to mainstream density ratio and blowing ratio should be considered for accurate testing, analysis and prediction of purge jet cooling scheme performance.

The Effects of Localized Blade Endwall Suction on Secondary Flows and Heat Transfer

2010 ◽  
Author(s):  
Rebecca Hollis ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
High Surface ◽  
Surface Heat ◽  
Secondary Flows ◽  
Surface Shear Stress ◽  
Mean Velocity ◽  
Surface Shear ◽  
Surface Heat Transfer ◽  
High Heat Transfer

Two methods of flow control were designed to mitigate the effects of the horseshoe vortex structure (HV) at an airfoil/endwall junction. An experimental study was conducted to quantify the effects of localized boundary layer removal on surface heat transfer in a low-speed wind tunnel. A transient infrared technique was used to measure the convective heat transfer values along the surface surrounding the juncture. Particle image velocimetry was used to collect the time-mean velocity vectors of the flow field across three planes of interest. Boundary layer suction was applied through a thin slot cut into the leading edge of the airfoil at two locations. The first, referred to as Method 1, was directly along the endwall, the second, Method 2, was located at a height ∼1/3 of the approaching boundary layer height. Five suction rates were tested; 0%, 6.5%, 11%, 15% and 20% of the approaching boundary layer mass flow was removed at a constant rate. Both methods reduced the effects of the HV with increasing suction on the symmetry, 0.5-D and 1-D planes. Method 2 yielded a greater reduction in surface heat transfer but Method 1 outperformed Method 2 aerodynamically by completely removing the HV structure when 11% suction was applied. This method however produced other adverse effects such as high surface shear stress and localized areas of high heat transfer near the slot edges at high suction rates.

Aero-Thermal Performance of a Two-Dimensional Highly Loaded Transonic Turbine Nozzle Guide Vane: A Test Case for Inviscid and Viscous Flow Computations

1992 ◽  
Vol 114 (1) ◽  
pp. 147-154 ◽  
Author(s):  
T. Arts ◽  
M. Lambert de Rouvroit
Keyword(s):  
Heat Transfer ◽  
Turbulence Intensity ◽  
Static Pressure ◽  
Free Stream ◽  
Guide Vane ◽  
Free Stream Turbulence ◽  
Nozzle Guide Vane ◽  
Transonic Turbine ◽  
Nozzle Guide ◽  
Highly Loaded

This contribution deals with an experimental aero-thermal investigation around a highly loaded transonic turbine nozzle guide vane mounted in a linear cascade arrangement. The measurements were performed in the von Karman Institute short duration Isentropic Light Piston Compression Tube facility allowing a correct simulation of Mach and Reynolds numbers as well as of the gas to wall temperature ratio compared to the values currently observed in modern aero engines. The experimental program consisted of flow periodicity checks by means of wall static pressure measurements and Schlieren flow visualizations, blade velocity distribution measurements by means of static pressure tappings, blade convective heat transfer measurements by means of platinum thin films, downstream loss coefficient and exit flow angle determinations by using a new fast traversing mechanism, and free-stream turbulence intensity and spectrum measurements. These different measurements were performed for several combinations of the free-stream flow parameters looking at the relative effects on the aerodynamic blade performance and blade convective heat transfer of Mach number, Reynolds number, and free-stream turbulence intensity.

Boundary Layer Influence on the Unsteady Horseshoe Vortex Flow and Surface Heat Transfer

2007 ◽  
Author(s):  
D. R. Sabatino ◽  
C. R. Smith
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Bluff Body ◽  
Leading Edge ◽  
Horseshoe Vortex ◽  
Surface Heat ◽  
Wall Region ◽  
Thermochromic Liquid Crystal ◽  
Surface Heat Transfer ◽  
End Wall

The spatial-temporal flow-field and associated surface heat transfer within the leading edge, end-wall region of a bluff body were examined using both particle image velocimetry and thermochromic liquid crystal temperature measurements. The horseshoe vortex system in the end-wall region is mechanistically linked to the upstream boundary layer unsteadiness. Hairpin vortex packets, associated with turbulent boundary layer bursting behavior, amalgamate with the horseshoe vortex resulting in unsteady strengthening and streamwise motion. The horseshoe vortex unsteadiness exhibits two different natural frequencies: one associated with the transient motion of the horseshoe vortex, and the other with the transient surface heat transfer. Comparable unsteadiness occurs in the end-wall region of the more complex airfoil geometry of a linear turbine cascade. To directly compare the horseshoe vortex behavior around a turning airfoil to that of a simple bluff body, a length scale based on the maximum airfoil thickness is proposed.

Numerical Prediction of Cooling Performance Sensitivity of 1st Stage Nozzle Guide Vane Under Aerothermal Conditions

2019 ◽  
Author(s):  
Prasert Prapamonthon ◽  
Bo Yin ◽  
Guowei Yang ◽  
Mohan Zhang
Keyword(s):  
Heat Transfer ◽  
Surface Roughness ◽  
High Pressure ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Cooling Performance ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract To obtain high power and thermal efficiency, the 1st stage nozzle guide vanes of a high-pressure turbine need to operate under serious circumstances from burned gas coming out of combustors. This leads to vane suffering from effects of high thermal load, high pressure and turbulence, including flow-separated transition. Therefore, it is necessary to improve vane cooling performance under complex flow and heat transfer phenomena caused by the integration of these effects. In fact, these effects on a high-pressure turbine vane are controlled by several factors such as turbine inlet temperature, pressure ratio, turbulence intensity and length scale, vane curvature and surface roughness. Furthermore, if the vane is cooled by film cooling, hole configuration and blowing ratio are important factors too. These factors can change the aerothermal conditions of the vane operation. The present work aims to numerically predict sensitivity of cooling performances of the 1st stage nozzle guide vane under aerodynamic and thermal variations caused by three parameters i.e. pressure ratio, coolant inlet temperature and height of vane surface roughness using Computational Fluid Dynamics (CFD) with Conjugate Heat Transfer (CHT) approach. Numerical results show that the coolant inlet temperature and the vane surface roughness parameters have significant effects on the vane temperature, thereby affecting the vane cooling performances significantly and sensitively.

Linear Cascade and Wind Tunnel Development for Turbine Blade Tip Heat Transfer Investigations With and Without Film Cooling

2019 ◽  
Author(s):  
Michael Sampson ◽  
Avery Fairbanks ◽  
Jacob Moseley ◽  
Phillip M. Ligrani ◽  
Hongzhou Xu ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Film Cooling ◽  
Boundary Layer Thickness ◽  
Surface Heat ◽  
Heat Transfer Characteristics ◽  
Surface Heat Transfer ◽  
Linear Cascade ◽  
External Cooling ◽  
Blade Tip

Abstract Currently, there is a deficit of experimental data for surface heat transfer characteristics and thermal transport processes associated with tip gap flows, and a lack of understanding of performance and behavior of film cooling as applied to blade tip surfaces. As a result, many avenues of opportunity exist for development of creative tip configurations with innovative external cooling arrangements. Overall goals of the present investigations are to reduce cooling air requirements, and reduce thermal loading, with equivalent improvements of thermal protection and structural integrity. Described is the development of experimental facilities, including a Supersonic/Transonic Wind Tunnel and linear cascade, for investigations of surface heat transfer characteristics of transonic turbine blade tips with unique squealer geometries and innovative film cooling arrangements. Note that data from past investigations are used to illustrate some of the experimental procedures and approaches which will be employed within the investigation. Of interest is development of a two-dimensional linear cascade with appropriate cascade airfoil flow periodicity. Included are boundary layer flow bleed devices, downstream tailboards, and augmented cascade inlet turbulence intensity. The present linear cascade approach allows experimental configuration parameters to be readily varied. Tip gap magnitudes are scaled so that ratios of tip gap to inlet boundary layer thickness, ratios of tip gap to blade axial chord length, and ratios of tip gap magnitudes to blade true chord length match engine hardware configurations. Ratios of inlet boundary layer thickness to tip gap range from 3 to 5. Innovative film cooling configurations are utilized for one blade tip configuration, and scaled engine components are modelled and tested with complete external cooling arrangements. Blade tip and geometry characteristics are also considered, including squealer depth and squealer tip wall thickness. With these experimental components, results will be obtained with engine representative transonic Mach numbers, Reynolds numbers, and film cooling parameters, including density ratios, which are achieved using foreign gas injection with carbon dioxide. Transient, infrared thermography approaches will be employed to measure spatially-resolved distributions of surface heat transfer coefficients, adiabatic surface temperature, and adiabatic film cooling effectiveness.

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