Simulation of Combustor/NGV Interaction Using Coupled RANS Solvers: Validation and Application to a Realistic Test Case

2014 ◽  
Author(s):  
Massimiliano Insinna ◽  
Simone Salvadori ◽  
Francesco Martelli
Keyword(s):  
Guide Vane ◽  
Flame Stabilization ◽  
Point Of View ◽  
Coupled Analysis ◽  
Test Case ◽  
Complex Flow ◽  
High Turbulence ◽  
Commercial Solver ◽  
Stability And Accuracy ◽  
Nozzle Guide

Numerical techniques are commonly used during both design and analysis processes, mainly considering single components. Technological progress asks for advanced approaches that include real-machine conditions and analyze components interaction, especially considering the combustor/turbine coupling. Modern combustors operate with strong swirl motions in order to obtain an adequate flame stabilization, generating a very complex flow field characterized by high turbulence level. These aspects affect performance of downstream components which are subjected to very aggressive inlet flow conditions: non-uniform total temperature, non-uniform total pressure, swirl and high turbulence intensity. For these reasons coupled analysis of combustor and turbine is necessary to accurately predict aero-thermal aspects that influence performance and reliability of these two components. From a numerical point of view the simulation of a single domain characterized by a reactive flow with very different Mach number regimes (from low-Mach flow in combustion chamber to transonic flow in turbine) is problematic due to the different numerical requirements needed, especially concerning stability and accuracy. These problems could be overcome using coupled methods to simultaneously simulate combustor and turbine in separated domains which are managed by different solvers that communicate with each other. A coupling method for the study of combustor/turbine interaction using the RANS methodology is proposed. In the first part of the paper the method is described and validated. The second part is dedicated to the application of the proposed coupling methodology to a realistic test case consisting of a model annular combustor and the Nozzle Guide Vane (NGV) of the MT1 high-pressure turbine stage. A commercial solver and an in-house code are respectively used for the simulation of combustor and NGV. Results are presented and analyzed highlighting the importance of such type of simulations in understanding aero-thermal phenomena that characterize combustor/vane interaction.

Modeling of Combustor and Turbine Vane Interaction

2019 ◽  
Author(s):  
Ishan Verma ◽  
Laith Zori ◽  
Jaydeep Basani ◽  
Samir Rida
Keyword(s):  
Swirling Flow ◽  
Guide Vane ◽  
Primary Objective ◽  
Point Of View ◽  
Computationally Efficient ◽  
Joint Simulation ◽  
Chemically Reacting ◽  
Combined Simulation ◽  
Aero Engines ◽  
Nozzle Guide

Abstract Modern aero-engines are characterized by compact components (fan, compressor, combustor, and turbine). Such proximity creates a complex interaction between the components and poses a modeling challenge due to the difficulties in identifying a clear interface between components since they are usually modeled separately. From a numerical point of view, the simulation of a complex compact aero-engine system requires interaction between these individual components, especially the combustor-turbine interaction. The combustor is characterized by a subsonic chemically reacting and swirling flow while the high-pressure turbine (HPT) stage has flow which is transonic. Furthermore, the simulation of combustor-turbine interactions is more challenging due to aggressive flow conditions such as non-uniform temperature, non-uniform total-pressure, strong swirl, and high turbulence intensity. The simulation of aero-engines, where combustor-turbine interactions are important, requires a methodology that can be used in a real engine framework while ensuring numerical requirements of accuracy and stability. Conventionally, such a simulation is carried out using one of the two approaches: a combined simulation (or joint-simulation) of the combustor and the HPT geometries, or a co-simulation between the combustor and the turbine with the exchange of boundary conditions between these two separate domains. The primary objective of this paper is to assess the effectiveness of the joint simulation versus the co-simulation and propose a more practical approach for modeling combustor and turbine interactions. First, a detailed grid independence study with hexahedral and polyhedral meshes is performed to select the required polyhedral mesh. Then, an optimal location of the interface between the combustor and the nozzle guide vane (NGV) is identified. Co-simulations are then performed by exchanging information between the combustor and the NGV at the interface, wherein the combustor is solved using LES while the NGV is solved using RANS. The joint combustor-NGV simulations are solved using LES. The effect of the combustor-NGV interaction on the flow field and hot streak migration is analyzed. The results suggest that the joint simulation is computationally efficient and more accurate since both components are modelled together.

The Effects of High Mainstream Turbulence and Turbine Vane Film Cooling on the Dispersion of a Simulated Hot Streak

2003 ◽  
Author(s):  
Sean Jenkins ◽  
Krishnakumar Varadarajam ◽  
David G. Bogard
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Peak Temperature ◽  
Nozzle Guide Vane ◽  
Turbulence Effects ◽  
High Turbulence ◽  
Nozzle Guide ◽  
Hot Streak ◽  
Film Cooling Holes

This paper presents the combined effects of high turbulence and film cooling on the dispersion of a simulated hot streak as it passes over a scaled-up nozzle guide vane. Experimental data demonstrates a considerable decay in the strength of a hot streak due to turbulence effects alone. Film cooling further reduces the peak temperature values resulting in a reduction of the peak temperature in the hot streak on the order of 75% relative to the upstream peak temperature in the hot streak. Comparisons are made between high turbulence (Tu = 20%) and moderate turbulence (Tu = 3.5%) as well as between different blowing conditions for the suction side, showerhead, and pressure side film cooling holes on a simulated nozzle guide vane.

Uncertainty Propagation Analyses of Lean Burn Combustor Exit Conditions for a Robust Nozzle Cooling Design

2020 ◽  
Vol 142 (5) ◽  
Author(s):  
Marius Schneider ◽  
Heinz-Peter Schiffer ◽  
Knut Lehmann
Keyword(s):  
Guide Vane ◽  
Uncertainty Propagation ◽  
Flow Fields ◽  
Secondary Flows ◽  
Complex Flow ◽  
Lean Burn ◽  
Numerical Predictions ◽  
Cooling Design ◽  
Parameterized Model ◽  
Nozzle Guide

Abstract Knowing the flow conditions at the combustor turbine interface is a key asset for an efficient cooling design of high-pressure turbines. However, measurements and numerical predictions of combustor exit conditions are challenging due to the extreme temperatures and complex flow patterns in modern combustors. Even the time-averaged flow fields at the combustor exit which are commonly used as inlet condition for simulations of the turbine are therefore subject to uncertainty. The goal of this paper is to illustrate how aleatory uncertainties in the magnitude and position of residual swirl and hot spots at the combustor exit affect uncertainties in the prediction of cooling and heat load of the first nozzle guide vane. Also, it is identified which of these uncertain parameters have the greatest impact. An iso-thermal test rig and an engine realistic setup with lean burn inflow conditions are investigated. The analysis combines a parameterized model for combustor exit flow fields with uncertainty quantification methods. It is shown that the clocking position of turbine inlet swirl has a large effect on the formation of secondary flows on the vane surface and thus affects the uncertainty of thermal predictions on the hub and vanes.

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.

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.

The Effects of High Mainstream Turbulence and Turbine Vane Film Cooling on the Dispersion of a Simulated Hot Streak

2004 ◽  
Vol 126 (1) ◽  
pp. 203-211 ◽  
Author(s):  
Sean Jenkins ◽  
Krishnakumar Varadarajan ◽  
David G. Bogard
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Suction Side ◽  
Peak Temperature ◽  
Nozzle Guide Vane ◽  
Turbulence Effects ◽  
High Turbulence ◽  
Nozzle Guide ◽  
Hot Streak ◽  
Film Cooling Holes

This paper presents the combined effects of high turbulence and film cooling on the dispersion of a simulated hot streak as it passes over a scaled-up nozzle guide vane. Experimental data demonstrates a considerable decay in the strength of a hot streak due to turbulence effects alone. Film cooling further reduces the peak temperature values resulting in a reduction of the peak temperature in the hot streak on the order of 75% relative to the upstream peak temperature in the hot streak. Comparisons are made between high turbulence Tu=20% and moderate turbulence Tu=3.5% as well as between different blowing conditions for the suction side, showerhead, and pressure side film cooling holes on a simulated nozzle guide vane.

scholarly journals A Simplified Program to Predict the Temperature Distribution in Gas Turbine Components as a Function of the Mission Profile of the Aircraft

1999 ◽  
Author(s):  
Alwyn F. Naudé ◽  
Jan A. Visser
Keyword(s):  
Temperature Distribution ◽  
Gas Turbines ◽  
Thermal Loading ◽  
Guide Vane ◽  
Three Dimensional ◽  
Operating Conditions ◽  
Test Case ◽  
Nozzle Guide ◽  
Mission Profile

One of the main thrusts of the modern aerospace industry is to reduce the operating costs of aircraft. This requires a longer on-wing time for the gas turbines and subsequently reduced maintenance. Such a program can only be effectively implemented if the effects of operating conditions on the aircraft can be evaluated continuously. This paper presents a simplified computer program, operating on a personal computer, to predict the temperature distribution in components as a function of the operating condition of the aircraft. This information is then used to determine the deterioration of the engine under the specified operating conditions. The program consists of a simplified model to calculate the conditions in the different modules of the engine as a function of parameters like throttle position, altitude, speed of the aircraft etc. The detailed heat transfer to components is calculated using simplified analytical formulations accounting for three-dimensional effects. As a test case, the temperature change on a semi-cooled nozzle guide vane (NGV) is shown as the engine accelerates to full load conditions. It can be concluded that this approach produces realistic values for the thermal loading on components that can be used to predict long-term engine deterioration.

Influence of Upstream Flow Conditions on the Heat Transfer to Nozzle Guide Vanes

1988 ◽  
Vol 110 (3) ◽  
pp. 412-416 ◽  
Author(s):  
V. Krishnamoorthy ◽  
B. R. Pai ◽  
S. P. Sukhatme
Keyword(s):  
Guide Vane ◽  
Gas Temperature ◽  
Flow Conditions ◽  
Cold Flow ◽  
Hot Gas ◽  
Nozzle Guide Vane ◽  
High Turbulence ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes ◽  
Upstream Flow

The influence of a combustor located just upstream of a nozzle guide vane cascade on the heat flux distribution to the nozzle guide vane was experimentally investigated. The surface temperature distribution around the convectively cooled vane of the cascade was obtained by locating the cascade, firstly in a low-turbulence uniform hot gas stream, secondly in a high-turbulence, uniform hot gas stream, and thirdly in a high-turbulence, nonuniform hot gas stream present just downstream of the combustor exit. The results indicate that the increased blade surface temperatures observed for the cascade placed just downstream of the combustor can be accounted for by the prevailing turbulence level measured at cascade inlet in cold-flow conditions and the average gas temperature at the cascade inlet.

Aero-Thermal Performance of a Nozzle Vane Cascade With a Generic Non Uniform Inlet Flow Condition: Part I — Influence of Non Uniformity Location

2016 ◽  
Author(s):  
A. Perdichizzi ◽  
H. Abdeh ◽  
G. Barigozzi ◽  
M. Henze ◽  
J. Krueckels
Keyword(s):  
Wind Tunnel ◽  
Stagnation Point ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Inlet Section ◽  
Inlet Flow ◽  
Nozzle Vane ◽  
High Turbulence ◽  
Nozzle Guide

In this paper, the modifications induced by the presence of an inlet flow non uniformity on the aerodynamic performance of a nozzle vane cascade are experimentally assessed. Tests were carried out in a six vane linear cascade whose profile is typical of a first stage nozzle guide vane of a modern heavy duty GT. An obstruction was located in the wind tunnel inlet section to produce a non uniform flow upstream of the leading edge plane. The cascade was tested in an atmospheric wind tunnel at an inlet Mach number Ma1 = 0.12, with a high turbulence intensity (Tu1 = 9%) and variable obstruction tangential and axial positions, as well as tangential extension. The presented results show that an inlet flow non uniformity influences the stagnation point position when it faces the vane leading edge from the suction side. A relevant increase of both 2D and secondary losses are observed when the non uniformity is aligned to the vane leading edge. When it is instead located in between the passage it does not affect the stagnation point location, in the meanwhile allowing a reduction in the secondary loss.

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