Effect of the Compressor Discharge Casing Geometry on Combustor Exit Temperature Profiles in a Multi-Can Gas-Turbine Combustor

2014 ◽  
Author(s):  
Krishna Kant Agarwal ◽  
Stefano Gori
Keyword(s):  
Temperature Profile ◽  
Gas Turbine ◽  
Thermal Stresses ◽  
Flow Behavior ◽  
Thermal Design ◽  
Reacting Flow ◽  
Asymmetric Distribution ◽  
Gas Temperature ◽  
Flow Features ◽  
Exit Temperature

Temperature profile variations in gas turbine combustors are important from the considerations of thermal stresses and material fatigue. The specific profile being addressed in this study is the combustor exit gas temperature profile in radial direction at first stage nozzle entry (also called the combustor transition-piece (TP) exit profile). Normally, in multi-can combustor configurations, this profile is assumed to be constant along the circumferential direction or from one can to another. However, field test on one of the GE-MS5002D class machine revealed that the shape of the combustor TP exit temperature profile is varying across the different cans. It is important to assess the reason of this behavior in order to define thermal input for stage 1 nozzle thermal design and define an average temperature profile for turbine bucket verification. For investigating the reasons of varying TP exit profiles across different cans, a reacting flow CFD study is performed for a combined multiple combustor-cans geometry. This is a challenging attempt considering that mesh for a single can liner is itself typically quite large (∼30 million) for capturing all flow features. The present multi-can study was made feasible with judicious simplification of combustor geometry, retaining only important flow features and using adequate mesh to capture system physics. Results indicate that the varying flame shape across different cans is indeed captured in the CFD. Hence, this effect could be something associated with the combustor design. Subsequent detailed post-processing of CFD results revealed the root cause to be associated with the presence of unsymmetrical arrangement of struts in the compressor discharge casing region. This effect is a slight flow-recirculation created much upstream due to the struts, which eventually results in asymmetric distribution of the flow across the combustor dilution holes. This leads to the flame shifting in different orientation for different cans with a systematic reference to the struts position. In conclusion, this paper describes the approach used for multi-can CFD analysis of the combustor, flow behavior in presence of unsymmetrical strut and its impact on the combustor exit temperature profile much downstream.

Thermal Analysis of Cooling System in a Gas Turbine Transition Piece

2011 ◽  
Author(s):  
Jun Su Park ◽  
Namgeon Yun ◽  
Hokyu Moon ◽  
Kyung Min Kim ◽  
Sin-Ho Kang ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Thermal Stresses ◽  
Cooling System ◽  
Structural Information ◽  
Thermal Analyses ◽  
Thermal Design ◽  
Transfer Coefficients ◽  
Transition Piece ◽  
Inlet Conditions

This paper presents thermal analyses of the cooling system of a transition piece, which is one of the primary hot components in a gas turbine engine. The thermal analyses include heat transfer distributions induced by heat and fluid flow, temperature, and thermal stresses. The purpose of this study is to provide basic thermal and structural information on transition piece, to facilitate their maintenance and repair. The study is carried out primarily by numerical methods, using the commercial software, Fluent and ANSYS. First, the combustion field in a combustion liner with nine fuel nozzles is analyzed to determine the inlet conditions of a transition piece. Using the results of this analysis, pressure distributions inside a transition piece are calculated. The outside of the transition piece in a dump diffuser system is also analyzed. Information on the pressure differences is then used to obtain data on cooling channel flow (one of the methods for cooling a transition piece). The cooling channels have exit holes that function as film-cooling holes. Thermal and flow analyses are carried out on the inside of a film-cooled transition piece. The results are used to investigate the adjacent temperatures and wall heat transfer coefficients inside the transition piece. Overall temperature and thermal stress distributions of the transition piece are obtained. These results will provide a direction to improve thermal design of transition piece.

Turbine Split Rings Thermal Design Using Conjugate Numerical Simulation

2013 ◽  
Author(s):  
Aleksei S. Tikhonov ◽  
Andrey A. Shvyrev ◽  
Nikolay Yu. Samokhvalov
Keyword(s):  
Heat Transfer ◽  
Numerical Simulation ◽  
Gas Turbine ◽  
Gas Dynamics ◽  
Thermal Condition ◽  
Fuel Efficiency ◽  
Thermal Design ◽  
Design Practice ◽  
Gas Temperature ◽  
Split Ring

One of the key factors ensuring gas turbine engines (GTE) competitiveness is improvement of life, reliability and fuel efficiency. However fuel efficiency improvement and the required increase of turbine inlet gas temperature (T*g) can result in gas turbine engine life reduction because of hot path components structural properties deterioration. Considering circumferential nonuniformity, local gas temperature T*g can reach 2500 K. Under these conditions the largest attention at designing is paid to reliable cooling of turbine vanes and blades. At present in design practice and scientific publications comparatively little attention is paid to detailed study of turbine split rings thermal condition. At the same time the experience of modern GTE operation shows high possibility of defects occurrence in turbine 1st stage split ring. This work objective is to perform conjugate numerical simulation (gas dynamics + heat transfer) of thermal condition for the turbine 1st stage split ring in a modern GTE. This research main task is to determine the split ring thermal condition by defining the conjugate gas dynamics and heat transfer result in ANSYS CFX 13.0 package. The research subject is the turbine 1st stage split ring. The split ring was simulated together with the cavity of cooling air supply from vanes through the case. Besides turbine 1st stage vanes and blades have been simulated. Patterns of total temperature (T*Max = 2000 °C) and pressure and turbulence level at vanes inlet (19.2 %) have been defined based on results of calculating the 1st stage vanes together with the combustor. The obtained results of numerical simulation are well coherent with various experimental studies (measurements of static pressure and temperature in supply cavity, metallography). Based on the obtained performance of the split ring cooling system and its thermal condition, the split ring design has been considerably modified (one supply cavity has been split into separate cavities, the number and arrangement of perforation holes have been changed etc.). All these made it possible to reduce considerably (by 40…50 °C) the split ring temperature comparing with the initial design. The design practice has been added with the methods which make it possible to define thermal condition of GTE turbine components by conjugating gas dynamics and heat transfer problems and this fact will allow to improve the designing level substantially and to consider the influence of different factors on aerodynamics and thermal state of turbine components in an integrated programming and computing suite.

scholarly journals A Method for Determining the Main Requirements of Outlet Gas Temperature Profile of Marine Gas Turbine Combustors

1982 ◽  
Author(s):  
Tang Chian-ti
Keyword(s):  
Temperature Distribution ◽  
Temperature Profile ◽  
Gas Turbine ◽  
Guide Vane ◽  
Safety Margin ◽  
Gas Temperature ◽  
Distribution Factor ◽  
Radial Temperature ◽  
Blade Height ◽  
Blade Cooling

Taking account of the marine gas turbine operation features, the author has chosen the hot corrosion peak temperature of materials as the guide vane material limiting temperature while evaluating the overall temperature distribution factor. Along with the blade cooling effectiveness a safety margin factor has been introduced during its evaluation. The gas temperature distribution along blade height is assumed to satisfy the condition that approximately equal safety factor in respect of strength prevails along blade height. Once the gas radial temperature profile becomes known, the radial temperature distribution factor can be readily determined.

scholarly journals Early Fault Detection of Gas Turbine Hot Components Based on Exhaust Gas Temperature Profile Continuous Distribution Estimation

Energies ◽  
2020 ◽  
Vol 13 (22) ◽  
pp. 5950
Author(s):  
Jinfu Liu ◽  
Mingliang Bai ◽  
Zhenhua Long ◽  
Jiao Liu ◽  
Yujia Ma ◽  
...  
Keyword(s):  
Fault Detection ◽  
Temperature Profile ◽  
Gas Turbine ◽  
Detection Method ◽  
Continuous Distribution ◽  
Detection Sensitivity ◽  
Exhaust Gas ◽  
Gas Temperature ◽  
Early Fault Detection ◽  
Exhaust Gas Temperature

Failures of the gas turbine hot components often cause catastrophic consequences. Early fault detection can detect the sign of fault occurrence at an early stage, improve availability and prevent serious incidents of the plant. Monitoring the variation of exhaust gas temperature (EGT) is an effective early fault detection method. Thus, a new gas turbine hot components early fault detection method is developed in this paper. By introducing a priori knowledge and quantum particle swarm optimization (QPSO), the exhaust gas temperature profile continuous distribution model is established with finite EGT measuring data. The method eliminates influences of operating and ambient condition changes and especially the gas swirl effect. The experiment reveals the presented method has higher fault detection sensitivity.

Characteristics of Multi-Cavity Trapped Vortex Combustors

2010 ◽  
Author(s):  
Alejandro M. Briones ◽  
Balu Sekar
Keyword(s):  
Heat Release ◽  
Temperature Profile ◽  
Viscous Drag ◽  
Reacting Flow ◽  
Total Drag ◽  
Pressure Drag ◽  
Air Jets ◽  
Exit Temperature ◽  
Trapped Vortex ◽  
Single Cavity

This research is motivated towards improving and optimizing the performance of AFRL’s Inter-Turbine Burner (ITB) in terms of greater combustion efficiency, reduced losses and exit temperature profile requirements. The ITB is a minicombustor concept, situated in between the high and low pressure turbine stages and typically contains multiple fueled and non-fueled Trapped Vortex Combustor (TVC) cavities. The size, placement, and arrangement of these cavities have tremendous effect on the combustor exit temperature profile. The detailed understanding of the effect of these cavities in a three-dimensional ITB configuration would be very difficult and computationally prohibited. Therefore, a simple but somewhat similar conceptual axi-symmetric burner is used here the design variations of Trapped Vortex Combustor (TVC) through modeling and simulation. The TVC can be one single cavity or can be represented by multi-cavity combustor. In this paper, both single cavity TVC and multi-cavity TVCs are studied. The single cavity TVC is divided into multiple cavities while the total volume of the combustor remains constant. Four combustors are studied: Baseline, Staged, Three-Staged, and Interdigitated TVC. An extensive computational investigation on the characteristics of these multi-cavity TVCs is presented. FLUENT is used for modeling the axisymmetric reacting flow past cavities using a global eddy dissipation mechanism for C3H8-air combustion with detailed thermodynamic and transport properties. Calculations are performed using Standard, RNG, and Realizable k-ε RANS turbulence models. The numerical results are validated against experimental temperature measurements on the Base TVC. Results indicate that the pressure drag is the major contributor to total drag in the Base TVC. However, viscous drag is still significant. By adding a concentric cavity in sequential manner (i.e. Staged TVC), the pressure drag decreases, whereas the viscous drag remains nearly constant. Further addition of a secondary concentric cavity (i.e. Three-Staged TVC), the total drag does not further decrease and both pressure and viscous drag contributions do not change. If instead a non-concentric cavity is added to the Base TVC (i.e. Interdigitated TVC), the pressure drag increases while the viscous drag decreases slightly. The effect of adding swirl flow is to increase the fuel-air mixing and as a result, it increases the maximum exit temperature for all the combustors modeled. The jets and heat release contribute to increase pressure drag with the former being greater. The fuel and air jets and heat release also modify the cavity flow structure. By turning off the fuel and air jets in the Staged TVC, lower drag (or pressure loss) and exit temperature are achieved. It is more effective to turn off the fuel and air jets in the upstream (front) cavity in order to reduce pressure losses. Based on these results, recommendations are provided to the engineer/designer/modeler to improve the performance of the ITB.

On Calculation and Optimization of Energy Systems

1996 ◽  
Author(s):  
Peter Ortmann ◽  
George Gyarmathy
Keyword(s):  
Gas Turbine ◽  
Mathematical Optimization ◽  
Optimization Procedure ◽  
Computer Code ◽  
Control Methods ◽  
Gas Temperature ◽  
Exit Temperature ◽  
Parameter Ranges ◽  
Operating Instructions ◽  
Technological Constraints

Most modern plants offer a variety of control methods which can be used one-by-one or in combination. For instance, in case of CC plants, control methods can comprise e. g. fuel control, adjustable compressor guide vanes and various feed preheat options. For the actual part-load operation the knowledge of optimum combinations is of great interest. A computer code based on standard mathematical optimization tools for large non-linear systems of equations has been developed and tested on various gas-turbine and CC power plant configurations. The code can be employed to plants of almost any level of complexity and to any particular plant layout. The range of control action is always limited by technological constraints imposed by mechanical, thermal, chemical or other limits and prescribed by the equipment manufacturer. The optimization code has been devised to respect an arbitrary number of such conditions as e. g. flue gas temperature, steam turbine moisture, gas turbine exit temperature and the like. Such limits and operating instructions concerning the observance of certain parameter ranges (e. g. limits prescribed for sake of life-cycle extension) can be introduced into the optimization procedure.

CFD Modeling of Non-Premixed Combustion in a Gas Turbine Combustor

2002 ◽  
Author(s):  
Kitano Majidi
Keyword(s):  
Gas Turbine ◽  
Transport Model ◽  
Direct Injection ◽  
Turbulence Models ◽  
Cfd Modeling ◽  
Navier Stokes ◽  
Reacting Flow ◽  
Gas Temperature ◽  
Gas Turbine Combustor ◽  
Reynolds Stress Transport Model

In the present study numerical calculations are used to solve reacting flow in a gas turbine combustor. A 3-D Favre-Averaged Navier-Stokes solver for a mixture of chemically reacting gases is applied to predict the flow pattern, gas temperature and fuel and species concentrations in the entire combustor. The complete combustor geometry with all important details such as air swirler vane passages and secondary holes are modeled. The calculations are carried out using three different turbulence models. Comparisons are made between the standard k-ε model, RNG k-ε model and a Reynolds stress transport model. To provide a closure for the chemical source term the Eddy Dissipation model is used. A lean direct injection of a liquid fuel is employed. Furthermore the influence of radiation will be investigated.

scholarly journals Development of 300kW-Class Ceramic Gas Turbine (CGT301) Engine System

1995 ◽  
Author(s):  
Masashi Tatsuzawa ◽  
Tomoki Taoka ◽  
Takeshi Sakida ◽  
Shinya Tanaka
Keyword(s):  
Gas Turbine ◽  
Development Process ◽  
Thermal Stresses ◽  
Turbine Blades ◽  
Axial Flow ◽  
Thermal Conditions ◽  
Fatigue Tests ◽  
Test Results ◽  
Gas Temperature ◽  
Engine System

CGT301 is a recuperated, single-shaft ceramic gas turbine for co-generation use. Ceramic parts are used in the hot section of the engine, such as turbine blades, nozzle vanes, combustor liners, heat exchanger elements and gas path parts. These ceramic parts are designed axi-symmetrically to reduce their sizes and thermal stresses and to avoid their unexpected deformations. The turbine is a two-stage axial flow type. As a primary feature of this turbine, the rotors are composed of ceramic blades inserted into metallic disks. The ceramic parts of the engine system have been tested before installing them in the engine to assure their reliability in the following manner. The ceramic blades have been examined by hot-spin test with the gas temperature of 1100°C and up to 110% of the engine rated speed. The ceramic stationary parts such as nozzle vanes, combustor liners and gas path parts, have been assembled and installed in a test rig with almost the same constraint and thermal conditions as the engine, and thermal fatigue tests of 100 cycles between 1200°C and 300°C have been conducted. After the proof tests of ceramic parts, they have been installed in the engine, step by step. Finally, the engine has been operated with a TIT of 1200°C at the engine rated speed of 56000 rpm. The present paper describes the development process and shows test results of the ceramic gas turbine at a TIT of 1200°C.

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