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.

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.

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 The Influence of Film-Cooling on the Aerodynamic Performance of a Turbine Nozzle Guide Vane

1997 ◽  
Author(s):  
C. Osnaghi ◽  
A. Perdichizzi ◽  
M. Savini ◽  
P. Harasgama ◽  
E. Lutum
Keyword(s):  
Carbon Dioxide ◽  
Mass Flow ◽  
Film Cooling ◽  
Momentum Flux ◽  
Guide Vane ◽  
Flux Ratio ◽  
Aerodynamic Performance ◽  
Trailing Edge ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

The paper presents the results of an investigation on the aerodynamic performance of a full coverage film-cooled nozzle guide vane. The blading is a typical high pressure turbine vane of advanced design, working in the high subsonic regime. Tests have been carried out for a wide range of conditions, including variations in Mach number, coolant to mainstream mass flow rate ratio and location of the coolant injection. Both air and carbon dioxide at ambient conditions have been utilized, as coolant flow. Measurements have been performed in a plane located at 0.5 axial chord downstream of the trailing edge by means of a miniaturized five-hole pressure probe. Performances, in terms of losses, flow angles and profile pressure distributions, for different cooling mass flow rates are presented and compared to the results of the solid blade tests (i.e. with no cooling holes). The results showed a significant increase of the losses with blowing. Test with air and carbon dioxide provided almost equal losses if carried out at the same global momentum flux ratio; however the density ratio was found to influence slightly the share of the coolant fluid among the injection rows and the local momentum flux ratio as well. In order to define the individual contributions of groups of cooling rows on the performance of the blade, three different modes of injection have been tested, namely full, trailing edge and shower head injection. The main trend observed is that trailing edge injection produces the least amount of additional losses at high blowing rates. Full-coverage film-cooling injection did not lead to marked variations in the blade pressure distribution and/or outlet flow angle.

AERO-THERMAL ASPECTS OF FILM COOLED NOZZLE GUIDE VANE ENDWALLS - PART 2: THERMAL MEASUREMENTS

2021 ◽  
pp. 1-39
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong 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.

Aero-Thermal Aspects of Film Cooled Nozzle Guide Vane Endwalls: Part 1 — Aerodynamics

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

Abstract The first stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics are not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The test section involves an engine-representative combustor-turbine interface geometry, combustor coolant flow and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

Aero-thermal Aspects of Film Cooled Nozzle Guide Vane Endwall - Part 1: Aerodynamics

2021 ◽  
pp. 1-54
Author(s):  
Mahmood H. Alqefl ◽  
Kedar P. Nawathe ◽  
Pingting Chen ◽  
Rui Zhu ◽  
Yong W. Kim ◽  
...  
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Three Dimensional ◽  
Leading Edge ◽  
Secondary Flows ◽  
Protective Film ◽  
Cooling Flows ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Endwall Film Cooling

Abstract The first stage turbine of a modern gas turbine is subjected to high thermal loads which lead to a need for aggressive cooling schemes to protect its components from melting. Endwalls are particularly challenging to cool due to the complex system of secondary flows near them that wash the protective film coolants into the mainstream. This paper shows that without including combustor cooling, the complex secondary flow physics are not representative of modern engines. Aggressive injection of all cooling flows upstream of the passage is expected to interact and change passage aerodynamics and, subsequently, mixing and transport of coolants. This study describes, experimentally, the aero-thermal interaction of cooling flows near the endwall of a first stage nozzle guide vane passage. The test section involves an engine-representative combustor-turbine interface geometry, combustor coolant flow and endwall film cooling flow injected upstream of a linear cascade. The approach flow conditions represent flow exiting a cooled, low-NOx combustor. This first part of this two-part study aims to understand the complex aerodynamics near the endwall through detailed measurements of passage three-dimensional velocity fields with and without endwall film cooling. The aerodynamic measurements reveal a dominant vortex in the passage, named here as the Impingement Vortex, that opposes the passage vortex formed at the airfoil leading edge plane. This Impingement Vortex completely changes our description of flow over a modern film cooled endwall.

Control of Secondary Flows in a Turbine Nozzle Guide Vane by Endwall Contouring

2000 ◽  
Author(s):  
T. I-P. Shih ◽  
Y.-L. Lin ◽  
T. W. Simon
Keyword(s):  
Film Cooling ◽  
Computational Study ◽  
Guide Vane ◽  
Three Dimensional ◽  
Secondary Flows ◽  
Navier Stokes ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
A Cell ◽  
Nozzle Guide

Computations were performed to study the three-dimensional flow and temperature distribution in a nozzle guide vane that has one flat and one contoured endwall with and without film cooling injected from two slots, one on each endwall located just upstream of the airfoil. For the contoured endwall, two locations of the same contouring were investigated, one with all contouring upstream of the airfoil and another with the contouring starting upstream of the airfoil and continuing through the airfoil passage. Results obtained show that when the contouring is all upstream of the airfoil, secondary flows on both the flat and the contoured endwalls are similar in magnitude. When the contouring starts upstream of the airfoil and continues through the airfoil passage, secondary flows on the contoured endwall are markedly weaker than those on the flat endwall. With weaker secondary flows on the contoured endwall, film-cooling effectiveness there is greatly improved. This computational study is based on the ensemble-averaged conservation equations of mass, momentum (compressible Navier-Stokes), and energy. Effects of turbulence were modeled by the low Reynolds number shear-stress transport k-ω model. Solutions were generated by a cell-centered, finite-volume method that uses third-order accurate flux-difference splitting of Roe with limiters and multigrid acceleration of a diagonalized ADI scheme with local time stepping on patched/embedded structured grids.

Numerical Analysis on Effects of the Contoured Endwall on the Three Dimensional Flow Characteristics in an Annular Turbine Nozzle Vane

2007 ◽  
Author(s):  
Wu Sang Lee ◽  
Jin Taek Chung ◽  
Dae Hyun Kim ◽  
Seung Joo Choe
Keyword(s):  
Guide Vane ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Cross Flow ◽  
Leading Edge ◽  
Secondary Flows ◽  
Three Dimensional Flow ◽  
Dimensional Flow ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

The three-dimensional flow in a turbine nozzle guide vane passage causes large secondary loss through the passage and increased heat transfer on the blade surface. In order to reduce or control these secondary flows, a linear turbine with contoured endwall configurations was used and changes in the three-dimensional flow field were analyzed and discussed. Contoured endwalls are installed at a location downstream of the saddle point near the leading edge of the pressure side blade and several positions along the centerline of the passage at constant distance. The objective of this study is to document the development of the three-dimensional flow in a turbine nozzle guide vane cascade with modified endwall. In addition, it proposes and appropriates endwall contouring which shows best overall loss reduction performance among the simulated contoured endwall. The results of this study show that the development of passage vortex and cross flow in the cascade composed of one flat and one contoured endwalls are affected by the acceleration which occurs in contoured endwall side. The overall loss is reduced near the flat endwall rather than contoured endwall, the best performance was shown for the case of 10–15% contoured for span-wise, 40–70% length of chord from trailing edge.

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.

scholarly journals Influence of Swirl Clocking on the Performance of Turbine Stage with Three-Dimensional Nozzle Guide Vane

Energies ◽  
2021 ◽  
Vol 14 (17) ◽  
pp. 5503
Author(s):  
Shinyoung Jeon ◽  
Changmin Son ◽  
Jinuk Kim
Keyword(s):  
Guide Vane ◽  
Influencing Factor ◽  
Three Dimensional ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Flow Angle ◽  
Nozzle Guide Vane ◽  
Nozzle Guide ◽  
Relationship Of ◽  
The Relationship

The effect of the swirl clocking on three-dimensional nozzle guide vane (NGV) is investigated using computational fluid dynamics. The research reports the loss characteristics of leaned and swept NGVs and the influence of swirl clocking. The three-dimensional NGVs are built by stacking the same 2D profile along different linear axes, characterized by different angles with respect to the normal or radial direction: ε = −12° ~ +12° for the leaned and γ = −5° ~ +10° for the swept airfoils. A total of 40 models are analyzed to study the effects of lean and sweep on aerodynamic performance. To investigate the influence of swirl clocking, the analysis cases include the center of the swirl that was positioned at the leading edge as well as the middle of the passage. The prediction results show that the relationship of the changes in mass flow rate and throat area are not monotonic. Further observation confirms the redistribution of loading and flow angle under different lean and sweep angles; thus, three-dimensional design is a key influencing factor on aerodynamic performance. In the presence of swirl clocking, NGV performance is changed significantly and the findings offer new insight and opportunities to improve three-dimensional NGV airfoil design.

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