Optimal State-Space Control of a Gas Turbine Engine

1992 ◽  
Vol 114 (4) ◽  
pp. 763-767 ◽  
Author(s):  
J. W. Watts ◽  
T. E. Dwan ◽  
C. G. Brockus
Keyword(s):  
State Space ◽  
Gas Turbine ◽  
State Space Model ◽  
Gas Turbine Engine ◽  
The State ◽  
Operating Conditions ◽  
Turbine Engine ◽  
Analog Control ◽  
Linear State ◽  
High Level

An analog fuel control for a gas turbine engine was compared with several state-space derived fuel controls. A single-spool, simple cycle gas turbine engine was modeled using ACSL (high level simulation language based on FORTRAN). The model included an analog fuel control representative of existing commercial fuel controls. The ACSL model was stripped of nonessential states to produce an eight-state linear state-space model of the engine. The A, B, and C matrices, derived from rated operating conditions, were used to obtain feedback control gains by the following methods: (1) state feedback; (2) LQR theory; (3) Bellman method; and (4) polygonal search. An off-load transient followed by an on-load transient was run for each of these fuel controls. The transient curves obtained were used to compare the state-space fuel controls with the analog fuel control. The state-space fuel controls did better than the analog control.

Optimal State Space Control of a Gas Turbine Engine

1991 ◽  
Author(s):  
J. W. Watts ◽  
T. E. Dwan ◽  
C. G. Brockus
Keyword(s):  
State Space ◽  
Gas Turbine ◽  
State Space Model ◽  
Gas Turbine Engine ◽  
The State ◽  
Operating Conditions ◽  
Turbine Engine ◽  
Analog Control ◽  
Linear State ◽  
High Level

An analog fuel control for a gas turbine engine was compared with several state space derived fuel controls. A single spool, simple cycle gas turbine engine was modeled using ACSL (high level simulation language based on FORTRAN). The model included an analog fuel control representative of existing commercial fuel controls. The ACSL model was stripped of non-essential states to produce an 8 state linear state space model of the engine. The A, B, and C matrices, derived from rated operating conditions, were used to obtain feedback control gains by the following methods: (1) state feedback; (2) LQR theory; (3) Bellman method; and (4) polygonal search. An off-load transient followed by an on-load transient was run for each of these fuel controls. The transient curves obtained were used to compare the state space fuel controls with the analog fuel control. The state space fuel controls did better than the analog control.

scholarly journals A Model and State-Space Controllers for an Intercooled, Regenerated (ICR) Gas Turbine Engine

1992 ◽  
Author(s):  
J. W. Watts ◽  
T. E. Dwan ◽  
R. W. Garman
Keyword(s):  
State Space ◽  
Gas Turbine ◽  
High Efficiency ◽  
Linear Quadratic Regulator ◽  
Gas Turbine Engine ◽  
Linear Quadratic ◽  
Turbine Engine ◽  
Simulation Language ◽  
Linear State ◽  
High Level

A two-and-one-half spool gas turbine engine was modeled using the Advanced Computer Simulation Language (ACSL), a high level simulation environment based on FORTRAN. A possible future high efficiency engine for powering naval ships is an intercooled, regenerated (ICR) gas turbine engine and these features were incorporated into the model. Utilizing sophisticated instructions available in ACSL linear state-space models for this engine were obtained. A high level engineering computational language, MATLAB, was employed to exercise these models to obtain optimal feedback controllers characterized by the following methods: (1) state feedback; (2) linear quadratic regulator (LQR) theory; and (3) polygonal search. The methods were compared by examining the transient curves for a fixed off-load, and on-load profile.

Quantifying Uncertainty of Gas Turbine Engine Models Generated Using Inverse Solution Methods

2018 ◽  
Author(s):  
Craig Nolen ◽  
Jacob Delimont
Keyword(s):  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Performance Model ◽  
Inverse Solution ◽  
Operating Conditions ◽  
Turbine Engine ◽  
Machine Performance ◽  
Solution Methods ◽  
High Level ◽  
Industrial Gas Turbine

In the current economic and political environment, there is a push for gas turbine operators to achieve higher operating efficiencies, which in turn, reduces emissions and fuel consumption. As these owners and operators seek to increase the efficiency of their machines, they are increasingly turning to physics-based performance modeling. This allows the end user to analyze machine performance, plan for performance upgrades, and evaluate use cases and operating conditions not originally envisioned by the original equipment manufacturers (OEMs). For owners/operators who do not have access to physics-based models provided by the hardware OEM or would like to evaluate modifications to legacy hardware, physics-based models may be developed using measured turbine performance data and high-level knowledge of the turbine architecture. In previous work, a physics-based performance model of an industrial gas turbine engine was created using measured plant operating data and an inverse solution method to allow off-design exploration of its performance. However, this model’s uncertainty was unknown, and knowledge of uncertainty is crucial to understanding a model’s reliability. In the present work, the model’s uncertainty in predicted performance at a particular operating point is investigated using statistical methods. Polynomial regressions of standard deviation are used alongside the performance regressions to describe the uncertainty at various operating points. These regressions are also used to visualize the variation of uncertainty across the performance map. Such knowledge of uncertainty can aid gas turbine operators in decision making with regard to the risks of off-design operation or equipment modifications.

The Effect of Heat Transfer Coefficient Increase on Tip Clearance Control in H.P. Compressors in Gas Turbine Engine

2013 ◽  
Author(s):  
Godwin Ita Ekong ◽  
Christopher A. Long ◽  
Peter R. N. Childs
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Time Constant ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Test Facility ◽  
Turbine Engine

Compressor tip clearance for a gas turbine engine application is the radial gap between the stationary compressor casing and the rotating blades. The gap varies significantly during different operating conditions of the engine due to centrifugal forces on the rotor and differential thermal expansions in the discs and casing. The tip clearance in the axial flow compressor of modern commercial civil aero-engines is of significance in terms of both mechanical integrity and performance. In general, the clearance is of critical importance to civil airline operators and their customers alike because as the clearance between the compressor blade tips and the casing increases, the aerodynamic efficiency will decrease and therefore the specific fuel consumption and operating costs will increase. This paper reports on the development of a range of concepts and their evaluation for the reduction and control of tip clearance in H.P. compressors using an enhanced heat transfer coefficient approach. This would lead to improvement in cruise tip clearances. A test facility has been developed for the study at the University of Sussex, incorporating a rotor and an inner shaft scaled down from a Rolls-Royce Trent aero-engine to a ratio of 0.7:1 with a rotational speed of up to 10000 rpm. The idle and maximum take-off conditions in the square cycle correspond to in-cavity rotational Reynolds numbers of 3.1×106 ≤ Reφ ≤ 1.0×107. The project involved modelling of the experimental facilities, to demonstrate proof of concept. The analysis shows that increasing the thermal response of the high pressure compressor (HPC) drum of a gas turbine engine assembly will reduce the drum time constant, thereby reducing the re-slam characteristics of the drum causing a reduction in the cold build clearance (CBC), and hence the reduction in cruise clearance. A further reduction can be achieved by introducing radial inflow into the drum cavity to further increase the disc heat transfer coefficient in the cavity; hence a further reduction in disc drum time constant.

scholarly journals Comprehensive report on Materials for Gas Turbine Engine Components

2019 ◽  
Vol 9 (2S2) ◽  
pp. 155-158
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Thermal Stresses ◽  
Prolonged Exposure ◽  
Raw Materials ◽  
Turbine Blades ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Turbine Engine ◽  
Engine Component

In the past three decades, it is very challenging for the researchers to design and development a best gas turbine engine component. Engine component has to face different operating conditions at different working environments. Nickel based superalloys are the best material to design turbine components. Inconel 718, Inconel 617, Hastelloy, Monel and Udimet are the common material used for turbine components. Directional solidification is one of the conventional casting routes followed to develop turbine blades. It is also reported that the raw materials are heat treated / age hardened to enrich the desired properties of the material implementation. Accordingly they are highly susceptible to mechanical and thermal stresses while operating. The hot section of the turbine components will experience repeated thermal stress. The halides in the combination of sulfur, chlorides and vanadate are deposited as molten salt on the surface of the turbine blade. On prolonged exposure the surface of the turbine blade starts to peel as an oxide scale. Microscopic images are the supportive results to compare the surface morphology after complete oxidation / corrosion studies. The spectroscopic results are useful to identify the elemental analysis over oxides formed. The predominant oxides observed are NiO, Cr2O3, Fe2O3 and NiCr2O4. These oxides are vulnerable on prolonged exposure and according to PB ratio the passivation are very less. In recent research, the invention on nickel based superalloys turbine blades produced through other advanced manufacturing process is also compared. A summary was made through comparing the conventional material and advanced materials performance of turbine blade material for high temperature performance.

Method for calculating the thrust and specific thrust of a double-flow gas turbine engine based on the results of tests at the laboratory unit «Flame»

2021 ◽  
pp. 58-64
Author(s):  
S.M. Sergeev ◽  
◽  
V.A. Kudriashov ◽  
N.V. Petrukhin ◽  
◽  
...  
Keyword(s):  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Similarity Criteria ◽  
Fuel Quality ◽  
Jet Engines ◽  
Laboratory Unit ◽  
Turbine Engine ◽  
Specific Thrust ◽  
The Cost

The main technical characteristics of jet engines depend on the fuel quality: thrust and fuel consumption. As a rule, the comparative assessment of real engines is carried by specific values. Specific thrust is one of the most important parameters of the gas turbine engine (GTE). The larger it is, the smaller the required air flow rate through the engine at a given thrust and therefore its dimensions and mass. To date, a system for evaluating the performance properties of fuels based on qualification methods has been created. However, these methods do not allow calculating the thrust and specific thrust of the engine and potentially assessing the effect of fuels on these characteristics. Therefore, the issues of efficient use of fuels for GTE are solved almost exclusively on the basis of tests at testing units with full-scale engines, which are carried out repeatedly, which leads to a significant increase in the cost of testing. The article proposes a method for calculating the thrust and specific thrust of a double-flow gas turbine engine according to the results of tests at a constant volume laboratory unit of bypass type “Flame”. The method is based on modeling the engine operating conditions using the similarity criteria of the bench reactor and the real engine and allows reducing significantly the material and time costs for testing. The experimental of the combustion characteristics of hydrocarbon fuels and the rated values of their thrust and specific thrust for a double-flow gas turbine engine are presented.

Cycle Design Exploration Using Multi-Design Point Approach

2012 ◽  
Author(s):  
Jeffrey Schutte ◽  
Jimmy Tai ◽  
Jonathan Sands ◽  
Dimitri Mavris
Keyword(s):  
Gas Turbine ◽  
Design Space ◽  
Traditional Approach ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Feasible Region ◽  
Turbine Engine ◽  
Design Point ◽  
Performance Requirements ◽  
Design Cycle

The focus of this study is to compare the aerothermodynamic cycle design space of a gas turbine engine generated using two on-design approaches. The traditional approach uses a single design point (SDP) for on-design cycle analysis, where off-design cycle analysis must be performed at other operating conditions of interest. A multi-design point (MDP) method performs on-design cycle analysis at all operating conditions where performance requirements are specified. Effects on the topography of the cycle design space as well as the feasibility of the space are examined. The impacts that performance requirements and cycle assumptions have on the bounds and topography of the feasible space are investigated. The deficiencies of a SDP method in determining an optimum gas turbine engine will be shown for a given set of requirements. Analysis will demonstrate that the MDP method, unlike the SDP method, always obtains a properly sized engine for a set of given requirements and cycle design variables, resulting in an increased feasible region of the aerothermodynamic cycle design space from which the optimum performance engine can be obtained.

Development of Condition Monitoring Test Cell Using Micro Gas Turbine Engine

2009 ◽  
Author(s):  
Seonghee Kho ◽  
Jayoung Ki ◽  
Miyoung Park ◽  
Changduk Kong ◽  
Kyungjae Lee
Keyword(s):  
Performance Analysis ◽  
Real Time ◽  
Gas Turbine ◽  
Condition Monitoring ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Test Cell ◽  
Turbine Engine ◽  
Time Performance ◽  
Engine Condition

This study is aim to be programmed the simulation which is available for real-time performance analysis so that is to be developed gas turbine engine’s condition monitoring system with analyzing difference between performance analysis results and measuring data from test cell. In addition, test cell created by this study have been developed to use following applications: to use for learning principals and mechanism of gas turbine engine in school, and to use performance test and its further research for variable operating conditions in associated institutes. The maximum thrust of the micro turbojet engine is 137 N (14 kgf) at 126,000 rpm of rotor rotational speed if the Jet A1 kerosene fuel is used. The air flow rate is measured by the inflow air speed of duct, and the fuel flow is measured by a volumetric fuel flowmeter. Temperatures and pressures are measured at the atmosphere, the compressor inlet and outlet and the turbine outlet. The thrust stand was designed and manufactured to measure accurately the thrust by the load cell. All measuring sensors are connected to a DAQ (Data Acquisition) device, and the logging data are used as function parameters of the program, LabVIEW. The LabVIEW is used to develop the engine condition monitoring program. The proposed program can perform both the reference engine model performance analysis at an input condition and the real-time performance analysis with real-time variables. By comparing two analysis results the engine condition can be monitored. Both engine performance analysis data and monitoring results are displayed by the GUI (Graphic User Interface) platform.

Simulation of a Gas Turbine Engine With Performance Degradation Modeling

2011 ◽  
Author(s):  
Ugo Campora ◽  
Mauro Carretta ◽  
Carlo Cravero
Keyword(s):  
Gas Turbine ◽  
Engine Performance ◽  
Gas Turbine Engine ◽  
Performance Degradation ◽  
Matlab Simulink ◽  
Turbine Engine ◽  
Simulink Model ◽  
Blade Thickness ◽  
High Level ◽  
Throat Section

A simulation of performance degradation for an aeronautical gas turbine engine (Honeywell T55 L712) is presented. The effects of turbine (low and high pressure stages) erosion on the engine performance have been investigated in some detail. The behavior of the engine has been simulated using a dynamic model implemented in Matlab-Simulink. Using a throughflow code the LPT and HPT have been simulated and their performance maps have been obtained with a high level of accuracy. In order to understand the effects of turbine erosion nine degradation levels have been introduced and the LPT and HPT performance have been computed using the abovementioned throughflow code. The degradation levels have been based on stator erosion effects (increase of throat section and blade thickness reduction) only according to the experimental evidence from the engine tests from Piaggio Aero Industries. The introduction of the modified turbine characteristics into the Matlab-Simulink model has allowed the degradation effects on the overall engine performance to be tested and discussed. Finally, using experimental data from the industrial maintenance database, the link of each level of degradation with the number of the engine operational time (hours) has been obtained.

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