Heat Transfer and Aerodynamics of an Intermediate Pressure Nozzle Guide Vane With and Without Inlet Temperature Non-Uniformity

2003 ◽  
Author(s):  
Kam S. Chana ◽  
T. Povey ◽  
Terry V. Jones
Keyword(s):  
Heat Transfer ◽  
Temperature Profile ◽  
Guide Vane ◽  
Turbine Blades ◽  
Inlet Temperature ◽  
Uniform Temperature ◽  
Intermediate Pressure ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Pressure Nozzle

In modern gas turbine engines the combustor exit flow has a non-uniform temperature profile because of the discrete nature of the injection of fuel and dilution air, and the wall cooling flows. The affect of this non-uniform temperature profile on the aerodynamics and heat transfer rate of nozzle guide vanes and turbine blades is difficult to predict, and knowledge of this is important for estimating turbine component life and efficiency. Measurements of heat transfer have been conducted on an annular transonic intermediate pressure nozzle guide vane operating downstream of a high pressure rotating turbine stage. Measurements were made with and without a radial and circumferential inlet temperature profile. The experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ, a short duration engine size turbine facility with 1.5 turbine stages, in which Mach number, Reynolds number and gas-to-wall temperature ratios are correctly modelled. Experimental results are compared to predictions performed using boundary layer methods.

Heat transfer measurements on an intermediate-pressure nozzle guide vane tested in a rotating annular turbine facility, and the modifying effects of a non-uniform inlet temperature profile

2003 ◽  
Vol 217 (4) ◽  
pp. 421-431 ◽  
Author(s):  
T Povey ◽  
K. S. Chana ◽  
T. V. Jones
Keyword(s):  
Heat Transfer ◽  
Temperature Profile ◽  
Heat Transfer Rate ◽  
Transfer Rate ◽  
Guide Vane ◽  
Temperature Gradients ◽  
Inlet Temperature ◽  
Uniform Temperature ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

In modern gas turbine engines there exist significant temperature gradients in the combustor exit flow. These gradients arise because both fuel and dilution air are introduced within the combustor as discrete jets. The effects of this non-uniform temperature field on the aerodynamics and heat transfer rate distributions of nozzle guide vanes and turbine blades is difficult to predict, although an increased understanding of the effects of temperature gradients would enhance the accuracy of estimates of turbine component life and efficiency. Low-frequency measurements of heat transfer rate have been conducted on an annular transonic intermediate-pressure (IP) nozzle guide vane operating downstream of a high-pressure (HP) rotating turbine stage. Measurements were conducted with both uniform and non-uniform inlet temperature profiles. The non-uniform temperature profile included both radial and circumferential gradients of temperature. Experiments were conducted in the isentropic light piston facility at QinetiQ Pyestock, a short-duration engine-size turbine facility with 1.5 turbine stages, in which Mach number, Reynolds number and gas—wall temperature ratios are correctly modelled. Experimental heat transfer results are compared with predictions performed using boundary layer methods.

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.

Effect of Simulated Combustor Temperature Nonuniformity on HP Vane and Endwall Heat Transfer: An Experimental and Computational Investigation

2010 ◽  
Author(s):  
Imran Qureshi ◽  
Arrigo Beretta ◽  
Thomas Povey
Keyword(s):  
Heat Transfer ◽  
Guide Vane ◽  
Research Programme ◽  
Test Facility ◽  
Inlet Temperature ◽  
Computational Investigation ◽  
Nozzle Guide Vane ◽  
Inlet Conditions ◽  
Nozzle Guide ◽  
Computational Predictions

This paper presents experimental measurements and computational predictions of surface and endwall heat transfer for a high-pressure (HP) nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular rotating turbine facility, with and without inlet temperature distortion (hot-streaks). A detailed aerodynamic survey of the vane surface is also presented. The test turbine was the unshrouded MT1 turbine, installed in the Turbine Test Facility (previously called Isentropic Light Piston Facility) at QinetiQ, Farnborough UK. This is a short duration facility, which simulates engine representative M, Re, non-dimensional speed and gas-to-wall temperature ratio at the turbine inlet. The facility has recently been upgraded to incorporate an advanced second-generation combustor simulator, capable of simulating well-defined, aggressive temperature profiles in both the radial and circumferential directions. This work forms part of the pan-European research programme, TATEF II. Measurements of HP vane and endwall heat transfer obtained with inlet temperature distortion are compared with results for uniform inlet conditions. Steady and unsteady CFD predictions have also been conducted on vane and endwall surfaces, using the Rolls-Royce CFD code HYDRA to complement the analysis of experimental results. The heat transfer measurements presented in this paper are the first of their kind in the respect that the temperature distortion is representative of an extreme cycle point measured in the engine situation, and was simulated with good periodicity and with well defined boundary conditions in the test turbine.

Effect of Nozzle Guide Vane Lean Under Influence of Inlet Temperature Traverse

2013 ◽  
Vol 136 (7) ◽  
Author(s):  
A. Rahim ◽  
B. Khanal ◽  
L. He ◽  
E. Romero
Keyword(s):  
Heat Transfer ◽  
Guide Vane ◽  
Tangential Displacement ◽  
Inlet Temperature ◽  
Heat Transfer Characteristics ◽  
Transfer Characteristics ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Hot Streak ◽  
The Impact

One of the most widely studied parameters in turbine blade shaping is blade lean, i.e., the tangential displacement of spanwise sections. However, there is a lack of published research that investigates the effect of blade lean under nonuniform temperature conditions (commonly referred to as a “hot-streak”) that are present at the combustor exit. Of particular interest is the impact of such an inflow temperature profile on heat transfer when the nozzle guide vane (NGV) blades are shaped. In the present work, a computational study has been carried out for a transonic turbine stage using an efficient unsteady Navier–Stokes solver (HYDRA). The configurations with a nominal vane and a compound leaned vane under uniform and hot-streak inlet conditions are analyzed. After confirming the typical NGV loading and aeroloss redistributions as seen in previous literature on blade lean, the focus has been directed to the rotor aerothermal behavior. While the overall stage efficiencies for the configurations are largely comparable, the results show strikingly different rotor heat transfer characteristics. For a uniform inlet, a leaned NGV has a detrimental effect on the rotor heat transfer. However, once the hot-streak is introduced, the trend is reversed; the leaned NGV leads to favorable heat transfer characteristics in general and for the rotor tip region in particular. The possible causal links for the observed aerothermal features are discussed. The present findings also highlight the significance of evaluating NGV shaping designs under properly conditioned inflow profiles, rather than extrapolating the wisdom derived from uniform inlet cases. The results also underline the importance of including rotor heat transfer and coolability during the NGV design process.

scholarly journals Behaviour of a Two Rows of Holes Coolant Film Along the Pressure Side of a High Pressure Nozzle Guide Vane

1989 ◽  
Author(s):  
T. Arts ◽  
A. E. Bourguignon
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Guide Vane ◽  
Internal Flow ◽  
Pressure Side ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Film Cooling Efficiency ◽  
Pressure Nozzle

The purpose of this paper is to quantify the influence on external convective heat transfer of a coolant film whose position varies along the pressure side of a high pressure turbine nozzle guide vane. The measurements were performed in the short duration Isentropic Light Piston Compression Tube facility of the von Karman Institute. The effects of external and internal flow are considered in terms of Mach number, Reynolds number, freestream turbulence intensity, blowing rate and coolant to freestream temperature ratio. The way to evaluate these results in terms of film cooling efficiency and heat transfer coefficient is finally discussed.

scholarly journals Thermal-Fluid–Solid Coupling—Parametrical Numerical Analysis of Hot Turbine Nozzle Guide Vane

Materials ◽  
2021 ◽  
Vol 14 (23) ◽  
pp. 7313
Author(s):  
Marcin Froissart ◽  
Tomasz Ochrymiuk
Keyword(s):  
Failure Modes ◽  
Guide Vane ◽  
Turbine Blades ◽  
Coolant Temperature ◽  
Viscous Heating ◽  
Inlet Temperature ◽  
Fine Mesh ◽  
Nozzle Guide Vane ◽  
Thermally Driven ◽  
Nozzle Guide

The cooling technology of hot turbine components has been a subject of continuous improvement for decades. In high-pressure turbine blades, the regions most affected by the excessive corrosion are the leading and trailing edges. In addition, high Kt regions at the hot gas path are exposed to cracking due to the low and high cycle fatigue failure modes. Especially in the case of a nozzle guide vane, the ability to predict thermally driven loads is crucial to assess its life and robustness. The difficulties in measuring thermal properties in hot conditions considerably limit the number of experimental results available in the literature. One of the most popular test cases is a NASA C3X vane, but coolant temperature is not explicitly revealed in the test report. As a result of that, numerous scientific works validated against that vane are potentially inconsistent. To address that ambiguity, the presented work was performed on a fully structural and a very fine mesh assuming room inlet temperature on every cooling channel. Special attention was paid to the options of the SST (shear-stress transport) viscosity model, such as Viscous heating (VH), Curvature correction (CC), Production Kato-Launder (KT), and Production limiter (PL). The strongest impact was from the Viscous heating, as it increases local vane temperature by as much as 40 deg. The significance of turbulent Prandtl number impact was also investigated. The default option used in the commercial CFD code is set to 0.85. Presented study modifies that value using equations proposed by Wassel/Catton and Kays/Crawford. Additionally, the comparison between four, two, and one-equation viscosity models was performed.

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.

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.

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