Health Monitoring of Variable Geometry Gas Turbines for the Canadian Navy

1989 ◽  
Vol 111 (2) ◽  
pp. 244-250 ◽  
Author(s):  
D. E. Muir ◽  
H. I. H. Saravanamuttoo ◽  
D. J. Marshall
Keyword(s):  
Health Monitoring ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Variable Geometry ◽  
Axial Compressors ◽  
High Pressure Ratio ◽  
Department Of National Defence ◽  
Canadian Navy ◽  
General Method

The Canadian Department of National Defence has identified a need for improved Engine Health Monitoring procedures for the new Canadian Patrol Frigate (CPF). The CPF propulsion system includes two General Electric LM2500 gas turbines, a high-pressure-ratio engine with multiple stages of compressor variable geometry. A general method for predicting the thermodynamic performance of variable geometry axial compressors has been developed. The new modeling technique is based on a meanline stage-stacking analysis and relies only on the limited performance data typically made available by engine manufacturers. The method has been applied to the LM2500-30 marine gas turbine and the variations in engine performance that can result from a malfunction of the variable geometry system in service have been estimated.

scholarly journals Impact of compressed air energy storage demands on gas turbine performance

2020 ◽  
pp. 095765092090627 ◽  
Author(s):  
Uyioghosa Igie ◽  
Marco Abbondanza ◽  
Artur Szymański ◽  
Theoklis Nikolaidis
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Compressed Air ◽  
Design Stage ◽  
Simulation Code ◽  
Ratio Distribution ◽  
Stall Margin ◽  
Industrial Gas Turbines

Industrial gas turbines are now required to operate more flexibly as a result of incentives and priorities given to renewable forms of energy. This study considers the extraction of compressed air from the gas turbine; it is implemented to store heat energy at periods of a surplus power supply and the reinjection at peak demand. Using an in-house engine performance simulation code, extractions and injections are investigated for a range of flows and for varied rear stage bleeding locations. Inter-stage bleeding is seen to unload the stage of extraction towards choke, while loading the subsequent stages, pushing them towards stall. Extracting after the last stage is shown to be appropriate for a wider range of flows: up to 15% of the compressor inlet flow. Injecting in this location at high flows pushes the closest stage towards stall. The same effect is observed in all the stages but to a lesser magnitude. Up to 17.5% injection seems allowable before compressor stalls; however, a more conservative estimate is expected with higher fidelity models. The study also shows an increase in performance with a rise in flow injection. Varying the design stage pressure ratio distribution brought about an improvement in the stall margin utilized, only for high extraction.

Design Study of an Advanced Concept Simple Gas Turbine for Possible Use in Low Emission Automobiles

1972 ◽  
Author(s):  
H. C. Eatock ◽  
M. D. Stoten
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Simple Cycle ◽  
Turbine Engines ◽  
Nox Emissions ◽  
Preliminary Design ◽  
Gas Turbine Engines ◽  
High Pressure Ratio

United Aircraft Corporation studied the potential costs of various possible gas turbine engines which might be used to reduce automobile exhaust emissions. As part of that study, United Aircraft of Canada undertook the preliminary design and performance analysis of high-pressure-ratio nonregenerated (simple cycle) gas turbine engines. For the first time, high levels of single-stage component efficiency are available extending from a pressure ratio less than 4 up to 10 or 12 to 1. As a result, the study showed that the simple-cycle engine may provide satisfactory running costs with significantly lower manufacturing costs and NOx emissions than a regenerated engine. In this paper some features of the preliminary design of both single-shaft and a free power turbine version of this engine are examined. The major component technology assumptions, in particular the high pressure ratio centrifugal compressor, employed for performance extrapolation are explained and compared with current technology. The potential low NOx emissions of the simple-cycle gas turbine compared to regenerative or recuperative gas turbines is discussed. Finally, some of the problems which might be encountered in using this totally different power plant for the conventional automobile are identified.

scholarly journals Development of a Flexible Turbine Cooling Prediction Tool for Preliminary Design of Gas Turbines

2018 ◽  
Vol 140 (9) ◽  
Author(s):  
Feijia Yin ◽  
Floris S. Tiemstra ◽  
Arvind G. Rao
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Thermodynamic Cycle ◽  
Specific Power ◽  
Inlet Temperature ◽  
Turbine Cooling ◽  
Engine Efficiency ◽  
Semi Empirical ◽  
Cooling Air

As the overall pressure ratio (OPR) and turbine inlet temperature (TIT) of modern gas turbines are constantly being increased in the pursuit of increasing efficiency and specific power, the effect of bleed cooling air on the engine performance is increasingly becoming important. During the thermodynamic cycle analysis and optimization phase, the cooling bleed air requirement is either neglected or is modeled by simplified correlations, which can lead to erroneous results. In this current research, a physics-based turbine cooling prediction model, based on semi-empirical correlations for heat transfer and pressure drop, is developed and verified with turbine cooling data available in the open literature. Based on the validated model, a parametric analysis is performed to understand the variation of turbine cooling requirement with variation in TIT and OPR of future advanced engine cycles. It is found that the existing method of calculating turbine cooling air mass flow with simplified correlation underpredicts the amount of turbine cooling air for higher OPR and TIT, thus overpredicting the estimated engine efficiency.

scholarly journals Advanced Labyrinth Seal Design Performance for High Pressure Ratio Gas Turbines

1975 ◽  
Author(s):  
H. L. Stocker
Keyword(s):  
High Pressure ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Labyrinth Seal ◽  
Internal Cavity ◽  
Water Tunnel ◽  
High Pressure Ratio ◽  
Test And Evaluation ◽  
Seal Design ◽  
Dynamic Seal

Labyrinth seal air leakage performance in current and advanced high pressure ratio gas turbines is directly related to the limitations of current available sealing technology. Sea design technology has not kept pace with the gas turbine major component advances. Therefore, an investigation was undertaken to design, fabricate and test several unique labyrinth seal concepts intended to decrease leakage through higher efficiency. The approach used in the unique designs for improving the efficiency of labyrinth seals involved increasing the internal cavity turbulence of the seal. The program involved three test and evaluation phases: (a) water tunnel studies; (b) static air rig tests; and (c) dynamic air rig tests. The water tunnel rig provided an economical method of screening the unique candidate designs. The most promising configurations from the water rig were fabricated and tested in the static air rig. Those configurations demonstrating a significant reduction in seal leakage over current designs were tested dynamically up to 786 ft/sec in an air rig to assess the effects of rotation. The results of this program effort show that each of the unique seal designs achieved lower leakage rates than a standard baseline step seal. In addition the dynamic seal test results show minimal effect on leakage due to rotation up to 786 ft/sec.

Analytical Investigation of the Rules of Component Matching in Turbojet Engines

1991 ◽  
Author(s):  
Dao-Zhi Liu
Keyword(s):  
High Pressure ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Analytical Investigation ◽  
High Pressure Ratio ◽  
Parameter Variations ◽  
Physical Essence ◽  
Component Matching ◽  
Analytical Formulae

In turbojet engines, phenomena under off–design conditions are related to the variations of incidence angles at different compressor stages. In the present paper, analytical formulae about the off–design incidence variations of inlet and exit compressor stages are derived. Using these fomulae, the detailed rules about engine parameter variations under off–design conditions are obtained; a series of phenomena about engine performance for low and high pressure ratio engines, engines with axial, centrifugal and combined compressors, as well as single–shaft and two–spool engines can then be explained systematically, their physical essence will be revealed clearly; and a variety of new situations and problems about engine off–design behavior can be predicted in advance.

scholarly journals Optimum Power Boosting of Gas Turbine Cycles With Refrigerated Inlet Air and Compressor Intercooling

1996 ◽  
Author(s):  
Mohand A. Ait-Ali
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Turbine Blade Cooling ◽  
High Pressure Ratio ◽  
Blade Cooling ◽  
Air Inlet ◽  
Turbine Inlet ◽  
Compressor Work

With or without turbine blade cooling, gas turbine cycles have consistently higher turbine inlet temperatures than steam turbine cycles. But this advantage is more than offset by the excessive compressor work induced by warm inlet temperatures, particularly during operation on hot summer days. Instead of seeking still higher turbine inlet temperatures by means of sophisticated blade cooling technology and high temperature-resistant blade materials, it is proposed to greatly increase the cycle net work and also improve thermal efficiency by decreasing the compressor work. This is obtained by using refrigerated inlet air and compressor intercooling to an extent which optimizes the refrigerated air inlet temperature and consequently the gas turbine compression ratio with respect to maximum specific net power. The cost effectiveness of this conceptual cycle, which also includes regeneration, has not been examined in this paper as it requires unusually high pressure ratio gas turbines and compressors, as well as high volumetric air flow rate and low temperature refrigeration equipment for which reliable cost data is not easily available.

Design and Experimental Investigation of a Mixed Flow Supersonic Compressor Stage

1995 ◽  
Author(s):  
Wolfgang Elmendorf ◽  
Harald Kurz ◽  
Heinz E. Gallus
Keyword(s):  
Experimental Investigation ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Strong Shock ◽  
Mixed Flow ◽  
High Pressure Ratio ◽  
Flow Deceleration ◽  
Stage Performance ◽  
Future Demands ◽  
Inlet Conditions

Highly loaded transonic and supersonic compressors appear capable of meeting the future demands of small gas turbines and jet engines. Particularly mixed flow compressors, taking advantage of the increasing circumferential blade speed between rotor inlet and exit, represent a good compromise with regard to high pressure ratio and mass flow on the one hand, and favorable performance characteristics and efficiency on the other. However, operating a supersonic rotor as part of a stage involves a stator characterized by high turning angles, supersonic inlet conditions and a strong flow deceleration. In fact, the stator can be identified as the critical component regarding overall stage performance. Based on experimentally determined rotor exit flow conditions, the first part of this paper describes the design of a tandem stator with a strong shock in the stator entrance region, followed by subsonic flow turning and diffusion. The main thrust of this paper is to present the analytical results obtained in connection with the experimental investigation of the complete stage at design and off-design conditions. Rotor and stator flow as well as the overall stage performance are discussed in detail. The concept of the tandem stator proves to be suitable for managing the extremely high aerodynamic loading in the Stator. Experimental results reveal the design goals to be met in general.

Compressor Variable Geometry Schedule Optimisation Using Genetic Algorithms

2009 ◽  
Author(s):  
L. Gallar ◽  
M. Arias ◽  
V. Pachidis ◽  
P. Pilidis
Keyword(s):  
High Efficiency ◽  
Engine Performance ◽  
Variable Geometry ◽  
Electric Generator ◽  
Axial Compressors ◽  
Turbine Efficiency ◽  
Surge Margin ◽  
Compressor Performance ◽  
Compressor Map ◽  
Generator Frequency

Variable geometry blade rows in axial compressors are devised to fulfil different requirements. Main objectives include their role as a “part speed crutch” to push the front stages out of surge at low spool speeds, modulation of the power output in industrial machines — given the fact that the spool needs to run at synchronous speed with the electric generator frequency — and they can also be re-staggered to attain a modified capacity (usually upflowed) of the same baseline compressor. The operating schedule of the variable vanes is typically obtained from expensive and time consuming performance rig tests in which a large number of possible combinations are compared. In principle, the final choice is dictated by the pursuit of high efficiency at high rotational speeds and increased surge margin at low speeds where large excursions away from the design point are expected. The aim of this work is to integrate a validated genetic algorithm optimiser within an industry proprietary mean line compressor performance prediction code to maximise the machine efficiency while keeping an adequate user-defined value of the surge margin. In so doing, an optimised variable geometry schedule is derived, together with a modified range of rotational speeds for each given operating point. Nevertheless, aware of the detrimental consequences to the whole engine performance that the new arrangement can cause, the whole engine response for the new settings has been investigated. In this regard and to a first order, the working line on the compressor map is considered unaffected by the setting of the variable vanes and the effect of the spool speed variation on the turbine operation is accounted for by a reduction in turbine efficiency proportional to any fall in the shaft speed. Results for a state of the art eight stage compressor show a marked improvement for the coupled compressor-turbine efficiency particularly at low spool speeds for a sensible value of the surge margin. Free from the surge margin constraint the efficiency is further increased at the expense of a hindered compressor operational stability. The work is intended to continue with the incorporation of bleeds and power off take in the calculations for the sake of a greater applicability of the tool.

scholarly journals The Mach 2 Axial Flow Compressor Stage

1975 ◽  
Author(s):  
F. A. E. Breugelmans
Keyword(s):  
High Pressure ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Inlet Guide Vane ◽  
Variable Geometry ◽  
Compressor Stage ◽  
High Pressure Ratio ◽  
Flow Compressor ◽  
Original Configuration

A supersonic compressor stage has been designed for a high pressure ratio at a tip relative inlet Mach number of 2. The stage was operated in the original configuration, but serious inlet stall occurred at part-speed operation. An inlet blockage ring, a bleed system and a variable geometry inlet guide vane have been analyzed and applied to this configuration. The results obtained with the bleed system in the complete stage are presented. The rotor performance is discussed and compared with the stage performance.

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