The Design, Development and Testing of a Non-Uniform Inlet Temperature Generator for the QinetiQ Transient Turbine Research Facility

2003 ◽  
Author(s):  
Kam S. Chana ◽  
James R. Hurrion ◽  
Terry V. Jones
Keyword(s):  
Hot Spots ◽  
Guide Vane ◽  
Radial Profile ◽  
Inlet Temperature ◽  
Inlet Section ◽  
Design Development ◽  
Research Facility ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Cold Gas

This paper presents the design, development and testing of a non-uniform inlet temperature generator for the QinetiQ Isentropic Light Piston Facility, a short duration turbine research facility. Major modifications were made to the facility to incorporate a non-uniform inlet temperature generator following a pilot study where a two-dimensional inlet section of the facility was replicated. The inlet region was modified to allow injection of cold gas at the hub and casing to generate the radial profile, whereas cold gas was injected from upstream turbulence rods to generate the circumferential variation (hot spots). Two configurations of hot spots were generated and characterised for both temperature and pressure. Heat transfer results at 50% span for the nozzle guide vane aerofoil are presented with and without inlet temperature non-uniformity.

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.

scholarly journals An Enhanced Flow-Thermo-Structural Modeling and Validation for the Integrated Analysis of a Film Cooling Nozzle Guide Vane

Energies ◽  
2019 ◽  
Vol 12 (14) ◽  
pp. 2775 ◽  
Author(s):  
Peng Guan ◽  
Yan-Ting Ai ◽  
Cheng-Wei Fei
Keyword(s):  
Thermal Stress ◽  
Film Cooling ◽  
Guide Vane ◽  
Integrated Design ◽  
Integrated Analysis ◽  
Inlet Temperature ◽  
Temperature Sensitive ◽  
Computational Accuracy ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

The target of this paper is to develop an enhanced flow-thermo-structural (FTS) model with high computational accuracy, to perform the integrated analysis of film cooling nozzle guide vane (NGV). An efficient turbulence model and weak spring approach are utilized in the enhanced FTS model. In respect of the power balance principle of aeroengine rotor shaft and temperature test of a typical combustor, the mean temperature inlet and five normalization temperature curves were confirmed, respectively. The temperature-sensitive paint (TSP) technology was used to verify the numerical simulation. From this study, we find that the predicted temperature caters for the TSP test well, between which the maximum error is less than 6%, and the maximum thermal stress is 758 MPa around the hole edges and the location of stress concentration keeps the consistency with that of the cracks. The maximum thermal stress increases by 10% with the increasing inlet temperature and reduces by about 16% with the shifting of flame peak from the outer to inner hub. The prediction provides general information on the initiation of cracks on a vane segment. The developed enhanced FTS model is validated to be workable and precise in the integrated analysis of film cooling NGV. The efforts of this study provide an integrated analysis approach of film cooling NGV and are promising to provide guidance for the integrated design of film cooling components besides NGV.

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.

Deposition on a Cooled Nozzle Guide Vane With Non-Uniform Inlet Temperatures

2015 ◽  
Author(s):  
Robin Prenter ◽  
Ali Ameri ◽  
Jeffrey P. Bons
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Internal Flow ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Deposition Rates ◽  
Nozzle Guide Vane ◽  
Tracking Model ◽  
Nozzle Guide ◽  
Hot Streak

External deposition on a slot film cooled nozzle guide vane, subjected to non-uniform inlet temperatures, was investigated experimentally and computationally. Experiments were conducted using a four-vane cascade, operating at temperatures up to 1353 K and inlet Mach number of approximately 0.1. Surveys of temperature at the inlet and exit planes were acquired to characterize the form and migration of the hot streak. Film cooling was achieved on one of the vanes using a single span-wise slot. Deposition was produced by injecting sub-bituminous ash particles with a median diameter of 6.48μm upstream of the vane passage. Several deposition tests were conducted, including a baseline case, a hot streak only case, and a hot streak and film cooled case. Results indicate that capture efficiency is strongly related to both the inlet temperature profiles and film cooling. Deposit distribution patterns are also affected by changes in vane surface temperatures. A computational model was developed to simulate the external and internal flow, conjugate heat transfer, and deposition. Temperature profiles measured experimentally at the inlet were applied as thermal boundary conditions to the simulation. For deposition modeling, an Eulerian-Lagrangian particle tracking model was utilized to track the ash particles through the flow. An experimentally tuned version of the critical viscosity sticking model was implemented, with predicted deposition rates matching experimental results well. Comparing overall deposition rates to results from previous studies indicate that the combined effect of non-uniform inlet temperatures and film cooling cannot be accurately simulated by simple superposition of the two independent effects, thus inclusion of both conditions in experiments is necessary for realistic simulation of external deposition.

Deposition on a Cooled Nozzle Guide Vane With Nonuniform Inlet Temperatures

2016 ◽  
Vol 138 (10) ◽  
Author(s):  
Robin Prenter ◽  
Ali Ameri ◽  
Jeffrey P. Bons
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Internal Flow ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Deposition Rates ◽  
Nozzle Guide Vane ◽  
Tracking Model ◽  
Nozzle Guide ◽  
Hot Streak

External deposition on a slot film cooled nozzle guide vane, subjected to nonuniform inlet temperatures, was investigated experimentally and computationally. Experiments were conducted using a four-vane cascade, operating at temperatures up to 1353 K and inlet Mach number of approximately 0.1. Surveys of temperature at the inlet and exit planes were acquired to characterize the form and migration of the hot streak. Film cooling was achieved on one of the vanes using a single spanwise slot. 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, including a baseline case, a hot streak-only case, and a hot streak and film cooled case. Results indicate that capture efficiency is strongly related to both the inlet temperature profiles and film cooling. Deposit distribution patterns are also affected by changes in vane surface temperatures. A computational model was developed to simulate the external and internal flow, conjugate heat transfer, and deposition. Temperature profiles measured experimentally at the inlet were applied as thermal boundary conditions to the simulation. For deposition modeling, an Eulerian–Lagrangian particle tracking model was utilized to track the ash particles through the flow. An experimentally tuned version of the critical viscosity sticking model was implemented, with predicted deposition rates matching experimental results well. Comparing overall deposition rates to results from previous studies indicates that the combined effect of nonuniform inlet temperatures and film cooling cannot be accurately simulated by simple superposition of the two independent effects; thus, inclusion of both conditions in experiments is necessary for realistic simulation of external deposition.

Endwall Film Cooling Using the Staggered Combustor-Turbine Gap Leakage Flow

2014 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Inlet Temperature ◽  
Local Pressure ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Blind Area ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

A key technology of gas turbine performance improvement was the increase in the turbine inlet temperature, which brought high thermal loads to the Nozzle Guide Vane (NGV) components. Strong pressure gradients in the NGVs and the complex secondary flow field had made thermal protection more challenging. As for the endwall surface near the pressure side gill region, the relatively higher local pressure and cross flow apparently decreased the film-cooling effectiveness. The aim of this investigation was to evaluate a new design, improving the film-cooling performance in a cooling blind area with upstream staggered slot, simulating the combustor-turbine leakage gap flow. The test cascades model was manufactured according to the GE-E3 nozzle guide vane scaled model, with a scale ratio of 2.2. The experiment was performed under the inlet Mach number 0.1 and the Reynolds number 3.5×105 based on an axial chord length of 78 mm. The staggered slots were positioned upstream of the cascades to simulate the combustor-turbine gap leakage. The Pressure Sensitive Painting (PSP) technique was used to detect the film cooling effectiveness distribution on the endwall surface. Through the investigation, the following results could be achieved: 1) the film-cooling effectiveness on the endwall surface downstream the slot and along the pitchwise direction increased, with the highest parameter at Z/Pitch = 0.6; 2) a larger cooled region developed towards the suction side as the blowing ratio increased; 3) the advantage of the staggered slot was apparent on the endwall surface near the inlet area, while the coolant film was obviously weakened along the axial chord at a low blowing ratio. The influence of the staggered slots could only be detected in the downstream area of the endwall surface at the higher blowing ratio.

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.

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.

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.

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