Full-Annulus URANS Study on the Transportation of Combustion Inhomogeneity in a Four-Stage Cooled Turbine

2019 ◽  
Author(s):  
Zhongran Chi ◽  
Haiqing Liu ◽  
Shusheng Zang ◽  
Chengxiong Pan ◽  
Guangyun Jiao
Keyword(s):  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Rotor Blades ◽  
Inlet Temperature ◽  
Computational Domain ◽  
Temperature Inhomogeneity ◽  
Unsteady Temperature ◽  
Turbine Engine ◽  
Exhaust Temperature ◽  
The Time Domain

Abstract The inhomogeneity of temperature in a turbine is related to the nonuniform heat release and air injections in combustors. In addition, it is influenced by the interactions between turbine cascades and coolant injections. Temperature inhomogeneity results in nonuniform flow temperature at turbine outlets, which is commonly measured by multiple thermal couples arranged in the azimuthal direction to monitor the operation of a gas turbine engine. Therefore, the investigation of temperature inhomogeneity transportation in a multistage gas turbine should help in detecting and quantifying the over-temperature or flameout of combustors using turbine exhaust temperature. Here the transportation of temperature inhomogeneity inside the four-stage turbine of a 300-MW gas turbine engine was numerically investigated using 3D CFD. The computational domain included all four stages of the turbine, consisting of more than 500 blades and vanes. Realistic components (N2, O2, CO2, and H2O) with variable heat capacities were considered for hot gas and cooling air. Coolants were added to the computational domain through more than 19,000 mass and momentum source terms. his was simple compared to realistic cooling structures. A URANS CFD run with over-temperature/flameout at 6 selected combustors out of 24 was carried out. The temperature distributions at rotor–stator interfaces and the turbine outlet were quantified and characterized by Fourier transformations in the time domain and space domain. It is found that the transport process from the hot-streaks/cold-streaks at the inlet to the outlet is relatively stable. The cold and hot fluid is redistributed in time and space due to the stator and rotor blades, in the region with a large parameter gradient at the inlet, strong unsteady temperature field and composition field appear. The distribution of the exhaust gas composition has a stronger correlation with the inlet temperature distribution and is less susceptible to interference.

Full-Annulus URANS Study on the Transportation of Combustion Inhomogeneity in a Four-Stage Cooled Turbine

2019 ◽  
Vol 141 (11) ◽  
Author(s):  
Zhongran Chi ◽  
Haiqing Liu ◽  
Shusheng Zang ◽  
Chengxiong Pan ◽  
Guangyun Jiao
Keyword(s):  
Gas Turbine ◽  
Rotor Blades ◽  
Navier Stokes ◽  
Inlet Temperature ◽  
Computational Domain ◽  
Temperature Inhomogeneity ◽  
Unsteady Temperature ◽  
Turbine Engine ◽  
Exhaust Temperature ◽  
Unsteady Temperature Field

Abstract The inhomogeneity of temperature in a turbine is related to the nonuniform heat release and air injections in combustors, which is commonly measured by thermal couples at the turbine exit. Investigation of temperature inhomogeneity transportation in a multistage gas turbine should help in detecting and quantifying the over-temperature or flameout of combustors using turbine exhaust temperature. Here, the transportation of combustion inhomogeneity inside the four-stage turbine of a 300-MW gas turbine engine was numerically investigated. The computational domain included four turbine stages, consisting of more than 500 blades and vanes. Realistic components (N2, O2, CO2, and H2O) with variable heat capacities were considered for hot gas and cooling air. Coolants were added to the computational domain through more than 19,000 mass and momentum source terms. An unsteady Reynolds averaged Navier–Stokes computational fluid dynamics (URANS CFD) run with over-temperature/flameout at 6 selected combustors out of 24 was carried out. The temperature distributions at rotor–stator interfaces and the turbine outlet were quantified and characterized using Fourier transformations in the space domain. It is found that the transport process from the hot streaks/cold streaks at the inlet to the outlet is relatively stable. The cold and hot fluid is redistributed in time and space due to the stator and rotor blades, and in the region with a large parameter gradient at the inlet, a strong unsteady temperature field and a composition field appear. The distribution of the exhaust-gas composition has a stronger correlation with the inlet temperature distribution and is less susceptible to interactions.

New Model Based Gas Turbine Fault Diagnostics Using 1D Engine Model and Nonlinear Identification Algorithms

2009 ◽  
Author(s):  
Seyyed Hamid Reza Hosseini ◽  
Hiwa Khaledi ◽  
Mohsen Reza Soltani
Keyword(s):  
Gas Turbine ◽  
Diagnostic System ◽  
Gas Turbine Engine ◽  
Fault Model ◽  
Fault Identification ◽  
Identification System ◽  
Turbine Engine ◽  
Engine Model ◽  
Exhaust Temperature ◽  
Model Based

Gas turbine fault identification has been used worldwide in many aero and land engines. Model based techniques have improved isolation of faults in components and stages’ fault trend monitoring. In this paper a powerful nonlinear fault identification system is developed in order to predict the location and trend of faults in two major components: compressor and turbine. For this purpose Siemens V94.2 gas turbine engine is modeled one dimensionally. The compressor is simulated using stage stacking technique, while a stage by stage blade cooling model has been used in simulation of the turbine. New fault model has been used for turbine, in which a degradation distribution has been considered for turbine stages’ performance. In order to validate the identification system with a real case, a combined fault model (a combination of existing faults models) for compressor is used. Also the first stage of the turbine is degraded alone while keeping the other stages healthy. The target was to identify the faulty stages not faulty components. The imposed faults are one of the most common faults in a gas turbine engine and the problem is one of the most difficult cases. Results show that the fault diagnostic system could isolate faults between compressor and turbine. It also predicts the location of faulty stages of each component. The most interesting result is that the fault is predicted only in the first stage (faulty stage) of the turbine while other stages are identified as healthy. Also combined fault of compressor is well identified. However, the magnitude of degradation could not be well predicted but, using more detailed models as well as better data from gas turbine exhaust temperature, will enhance diagnostic results.

Design, Analysis, and Manufacture of an Axial Length-Saving Disk-Oriented Engine

2020 ◽  
Vol 143 (1) ◽  
Author(s):  
Bennett M. Staton ◽  
Brian T. Bohan ◽  
Marc D. Polanka ◽  
Larry P. Goss
Keyword(s):  
Gas Turbine ◽  
Axial Length ◽  
Electric Propulsion ◽  
Guide Vane ◽  
Combustion Process ◽  
Gas Turbine Engine ◽  
Inlet Temperature ◽  
Turbine Engine ◽  
Power Extraction ◽  
Turbine Inlet

Abstract A disk-oriented engine was designed to reduce the overall length of a gas turbine engine, combining a single-stage centrifugal compressor and radial in-flow turbine (RIT) in a back-to-back configuration. The focus of this research was to understand how this unique flow path impacted the combustion process. Computational analysis was accomplished to determine the feasibility of reducing the axial length of a gas turbine engine utilizing circumferential combustion. The desire was to maintain circumferential swirl from the compressor through a U-bend combustion path. The U-bend reverses the outboard flow from the compressor into an integrated turbine guide vane in preparation for power extraction by the RIT. The computational targets for this design were a turbine inlet temperature of 1300 K, operating with a 3% total pressure drop across the combustor, and a turbine inlet pattern factor (PF) of 0.24 to produce a cycle capable of creating 668 N of thrust. By wrapping the combustion chamber about the circumference of the turbomachinery, the axial length of the entire engine was reduced. Reallocating the combustor volume from the axial to radial orientation reduced the overall length of the system up to 40%, improving the mobility and modularity of gas turbine power in specific applications. This reduction in axial length could be applied to electric power generation for both ground power and airborne distributive electric propulsion. Computational results were further compared to experimental velocity measurements on custom fuel–air swirl injectors at mass flow conditions representative of 668 N of thrust, providing qualitative and quantitative insight into the stability of the flame anchoring system. From this design, a full-scale physical model of the disk-oriented engine was designed for combustion analysis.

The AGT101 Advanced Automotive Gas Turbine

1982 ◽  
Author(s):  
R. A. Rackley ◽  
J. R. Kidwell
Keyword(s):  
United States ◽  
Gas Turbine ◽  
The United States ◽  
Gas Turbine Engine ◽  
Development Project ◽  
Single Stage ◽  
Department Of Energy ◽  
Inlet Temperature ◽  
United States Department ◽  
Turbine Engine

The Garrett/Ford Advanced Gas Turbine Powertrain System Development Project, authorized under NASA Contract DEN3-167, is sponsored by and is part of the United States Department of Energy Gas Turbine Highway Vehicle System Program. Program effort is oriented at providing the United States automotive industry the technology base necessary to produce gas turbine powertrains competitive for automotive applications having: (1) reduced fuel consumption, (2) multi-fuel capability, and (3) low emissions. The AGT101 powertrain is a 74.6 kW (100 hp), regenerated single-shaft gas turbine engine operating at a maximum turbine inlet temperature of 1644 K (2500 °F), coupled to a split differential gearbox and Ford automatic overdrive production transmission. The gas turbine engine has a single-stage centrifugal compressor and a single-stage radial inflow turbine mounted on a common shaft. Maximum rotor speed is 10,472 rad/sec (100,000 rpm). All high-temperature components, including the turbine rotor, are ceramic. AGT101 powertrain development has been initiated, with testing completed on many aerothermodynamic components in dedicated test rigs and start of Mod I, Build 1 engine testing.

Development of the Centaur Type H Gas Turbine Engine

1985 ◽  
Author(s):  
G. L. Padgett ◽  
W. W. Davis
Keyword(s):  
Gas Turbine ◽  
High Efficiency ◽  
State Of The Art ◽  
Gas Turbine Engine ◽  
Development Plan ◽  
Inlet Temperature ◽  
Market Place ◽  
Turbine Inlet Temperature ◽  
Turbine Engine ◽  
Computer Aided

In response to the needs of the market place for turbines in the 5000 to 6000 hp class, Solar Turbines Incorporated has responded with an uprate of their Centaur engine. Discussed in this paper are the features of the uprated engine, the Development Plan and the methodology for incorporating into the design the advanced aerodynamic and mechanical technology of the Mars engine. The Mars engine is a high efficiency 12,500 hp engine which operates at a turbine inlet temperature of 1935°F. State-of-the-art computer aided methods have been applied to produce the design, and the results from this approach are displayed.

Design of a State Space Controller for an Intercooled, Regenerated (ICR) Gas Turbine Engine

1994 ◽  
Author(s):  
Karl F. Prigge ◽  
Jerry W. Watts ◽  
Terrence E. Dwan
Keyword(s):  
Gas Turbine ◽  
Time Constant ◽  
Gas Turbine Engine ◽  
The Other ◽  
Pole Placement ◽  
Inlet Temperature ◽  
Fast Time ◽  
Turbine Engine ◽  
Power Turbine ◽  
Input Control

A multi-input, multi-output (MIMO) controller for an advanced gas turbine has been developed and tested using a computer simulation. The engine modeled is a two-and-one half spool gas turbine with both an intercooler and a regenerator. In addition, variable stator vanes are present in the free-power turbine. This advanced engine is proposed for future naval propulsion for both mechanical drive ships and electrical drive ships. The designed controller controls free-power turbine speed and turbine inlet temperature using fuel flow and angle of the stator vanes. The controller will also have four modes of operation to deal with over temperature and over speed conditions. An eight state reduced order controller was used with pole placement and LQR to arrive at control gains. Both these methods required considerable insight into the problem. This insight was provided by previous experience with controller design for a less complicated engine, and also by use of a polyhedral search model of the gas turbine engine. The difficulty with a MIMO controller was that both inputs affect both of the control variables. The classical resolution of this problem is to have one input control one variable at a fast time constant and the other input control the other variable at a slow time constant. The “optimal” resolution of this problem is analyzed using the transient curves and basic control theory.

Effect of long service on the fine structure of the surface of gas turbine engine rotor blades

1977 ◽  
Vol 9 (10) ◽  
pp. 1257-1261
Author(s):  
O. I. Marusii ◽  
Yu. I. Koval' ◽  
E. N. Kaspruk ◽  
V. N. Torgov
Keyword(s):  
Fine Structure ◽  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Rotor Blades ◽  
Turbine Engine ◽  
Engine Rotor

Rational Design of Gas Turbine Engine Compressor to Provide the Required Dynamic Strength Level of Rotor Blades

2016 ◽  
Author(s):  
Daria Kolmakova ◽  
Grigorii Popov ◽  
Aleksandr Shklovets ◽  
Aleksandr Ermakov
Keyword(s):  
Gas Turbine ◽  
Rational Design ◽  
Gas Flow ◽  
Dynamic Strength ◽  
Gas Turbine Engine ◽  
Rotor Blades ◽  
Dynamic Stresses ◽  
Turbine Engine ◽  
Acceptable Method ◽  
Flow Oscillations

Compressors of gas turbine engines often operate under the conditions of uneven gas flow. Oscillations of the blades occur under the influence of circumferential flow unevenness. The goal of the research was to find an acceptable method of reducing the level of dynamic stresses in the rotor blades. Motivation for the study was the problem of destruction of rotor blades of the last stage of Intermediate Pressure Compressor (IPC) which has been designed and produced at JSC “Kuznetsov” (Russia). The source of circumferential flow unevenness was middle annular frame located downstream the IPC. Seven struts with different maximum thickness are arranged irregularly in support passage. The first approaches propose to reduce the dynamic stresses in blades by detuning the blades from the dangerous harmonic due to changes in their natural frequency. The next two consist in reducing the circumferential unevenness of flow. Thus, this study gives the ideas about methods of improving the dynamic strength of rotor blades of gas turbine engine compressors. On the basis of existing conditions (under development or existing compressor) it allows the selection of the most suitable method for reducing dynamic stresses.

Modeling the Gas Turbine Engine Under Its Dynamic Heating Conditions

2014 ◽  
Vol 137 (3) ◽  
Author(s):  
Sergiy V. Yepifanov ◽  
Roman L. Zelenskyi ◽  
Igor Loboda
Keyword(s):  
Gas Turbine ◽  
Mutual Effect ◽  
Gas Turbine Engine ◽  
Path Model ◽  
Structural Elements ◽  
Dynamics Model ◽  
Unsteady Temperature ◽  
Turbine Engine ◽  
Operating Modes ◽  
Dynamic Heating

A modern gas turbine engine (GTE) is a complex nonlinear dynamic system with the mutual effect of gas-dynamic and thermal processes in its components. The engine development requires the precise real-time simulation of all main operating modes. One of the most complex operating modes for modeling is “cold stabilization,” which is the rotors acceleration without completely heated up the turbine elements. The dynamic heating problem is a topical practical issue. Solving the problem requires coordinating a gas-path model with heat and stress models, which is also a significant scientific problem. The phenomenon of interest is the radial clearances change during engines operation and its influence on engines static and dynamic performances. To consider the clearance change, it is necessary to synthesize the quick proceeding stress-state models (QPSSM) of a rotor and a casing for the initial temperature and dynamic heating. The unique feature of the QPSSM of GTEs is separate equation sets, which allow the heat exchange between structure elements and the gas (air) and the displacements of the turbine rotor and the casing. This ability appears as a result of determining the effect of each factor on different structural elements of the engine. The presented method significantly simplifies the model identification, which can be performed based on a precise calculation of the unsteady temperature fields of the structural elements and the variation of the radial clearance. Thus, the present paper addresses a new method to model the engine dynamics considering its heating up. The method is based on the integration of three models: the gas-path dynamics model, the clearance dynamics model, and the model of the clearance effect on the efficiency. The paper also comprises the program implementation of the models. The method was tested by applying to a particular turbofan engine.

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