Developments Leading to an Axial Flow Compressor for a 25 MW Class High Efficiency Gas Turbine

In this paper, the development leading to a 17-stage axial flow compressor (pressure ratio 14.7) for the 25 MW class heavy duty gas turbine H-25 is described. In the course of developing the H-25’s compressor, extensive measurements were carried out on models. Experimental results are compared with predicted values. Aerodynamic experiments covered the measurements of unsteady flows such as rotating stall and surge as well as the steady-state performance of the compressor. Based on the results of these tests, the aerodynamic and mechanical design parameters of the full scale H-25 compressor were finalized on the basis of two model compressors. Detailed measurements of the first unit of the H-25 gas turbine were carried out. Test results on the compressor are presented and show the achievement of the expected design targets.

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Design of the 6C Heavy-Duty Gas Turbine

Volume 2: Turbo Expo 2003 ◽

10.1115/gt2003-38686 ◽

2003 ◽

Author(s):

David Harper ◽

Devin Martin ◽

Harold Miller ◽

Robert Grimley ◽

Fre´de´ric Greiner

Keyword(s):

Power Systems ◽

Gas Turbine ◽

High Efficiency ◽

Pressure Ratio ◽

Axial Flow ◽

Load Test ◽

Advanced Technology ◽

General Electric ◽

Three Stages ◽

Flow Compressor

The MS6001C gas turbine combines the proven reliability of the General Electric gas turbine family with the advanced technology developed for the FA, FB and H machine designs. The engine configuration is a single shaft bolted rotor, driving a 50 or 60 Hz. generator though a cold end mounted load gear. Rated at 42.3 MW, with a thermal efficiency of 36.3%, the MS6001C will provide greater than a four percent increase in efficiency over the MS6001B. This paper is focused on the design and development of the MS6001C gas turbine, highlighting the commonality between this and other General Electric Power Systems (GEPS) and General Electric Aircraft Engines (GEAE) designs, as well as introducing some new and innovative features. The new high efficiency, 12 stage, axial flow compressor, features a 19:1 pressure ratio with three stages of variable guide vanes. The can annular, six chamber, Dry Low NOx (DLN-2.5H) combustion system is scaled from field proven, low emission technology. The turbine incorporates three stages, two cooled blade rows, and operates at a 1327°C firing temperature. After a thorough factory full speed no load test has been conducted, the first MS6001C engine will be shipped to a customer site in Kemalpasalzmir Turkey, where an instrumented full load test will be conducted to validate the design.

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Gas Turbine Performance Simulation Using an Optimized Axial Flow Compressor

Volume 6: Turbomachinery, Parts A and B ◽

10.1115/gt2006-91225 ◽

2006 ◽

Author(s):

Cleverson Bringhenti ◽

Jesui´no Takachi Tomita ◽

Francisco de Sousa Ju´nior ◽

Joa˜o R. Barbosa

Keyword(s):

Computer Program ◽

Gas Turbine ◽

Gas Turbines ◽

Axial Flow ◽

Engine Performance ◽

Design Parameters ◽

Axial Flow Compressor ◽

Surge Margin ◽

Flow Compressor ◽

And Control

Gas turbines need to operate efficiently due to the high specific fuel consumption. In order to reach the best possible efficiency the main gas turbine components, such as compressor and turbine, need to be optimized. This work reports the use of two specially developed computer programs: AFCC [1, 2] and GTAnalysis [3, 4] for such purpose. An axial flow compressor has been designed, using the AFCC computer program based on the stage-stacking technique. Major compressor design parameters are optimized at design point, searching for best efficiency and surge margin. Operation points are calculated and its characteristics maps are generated. The calculated compressor maps are incorporated to the GTAnalysis computer program for the engine performance calculation. Restrictions, like engine complexity, manufacture difficulties and control problems, are not taken into account.

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Development of a High-Pressure-Ratio Axial Flow Compressor for a Medium-Size Gas Turbine

Journal of Turbomachinery ◽

10.1115/1.3262042 ◽

1986 ◽

Vol 108 (2) ◽

pp. 233-239 ◽

Cited By ~ 5

Author(s):

Y. Kashiwabara ◽

Y. Matsuura ◽

Y. Katoh ◽

N. Hagiwara ◽

T. Hattori ◽

...

Keyword(s):

Gas Turbine ◽

Mechanical Design ◽

Pressure Ratio ◽

Axial Compressor ◽

Axial Flow ◽

Medium Size ◽

High Pressure Ratio ◽

Compressor Pressure Ratio ◽

Predicted Values ◽

Flow Compressor

In this paper, the development of a model 17-stage axial compressor (pressure ratio 14.7) for a medium-size gas turbine is described. The aerodynamic and mechanical design features of the compressor are presented. In advance of the full 17-stage test, the first three and nine stages were tested. Measured results confirm the design performance in the first stages of the 17-stage compressor. The details of the construction of the facilities, instrumentation and data acquisition system for the full 17-stage test are described. Test results for the 17-stage compressor are presented. The measured results are in good agreement with the predicted values.

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Numerical Efforts of Aerodynamic Re-Design in a Single-Stage Transonic Axial Compressor: Part 1 — Stator Design

Volume 2A: Turbomachinery ◽

10.1115/gt2017-63031 ◽

2017 ◽

Author(s):

Justin (Jongsik) Oh

Keyword(s):

Pressure Ratio ◽

Axial Compressor ◽

Axial Flow ◽

Single Stage ◽

Design Flow ◽

Design Parameters ◽

Original Design ◽

Circular Arc ◽

Surge Margin ◽

Flow Compressor

In many aerodynamic design parameters for the axial-flow compressor, three variables of tailored blading, blade lean and sweep were considered in the re-design efforts of a transonic single stage which had been designed in 1960’s NASA public domains. As Part 1, the re-design was limited to the stator vane only. For the original MCA (Multiple Circular Arc) blading, which had been applied at all radii, the CDA (Controlled Diffusion Airfoil) blading was introduced at midspan as the first variant, and the endwalls of hub and casing (or tip) were replaced with the DCA (Double Circular Arc) blading for the second variant. Aerodynamic performance was predicted through a series of CFD analysis at design speed, and the best aerodynamic improvement, in terms of pressure ratio/efficiency and operability, was found in the first variant of tailored blading. It was selected as a baseline for the next design efforts with blade lean, sweep and both combined. Among 12 variants, a case of positively and mildly leaned blades was found the most attractive one, relative to the original design, providing benefits of an 1.0% increase of pressure ratio at design flow, an 1.7% increase of efficiency at design flow, a 10.5% increase of the surge margin and a 32.3% increase of the choke margin.

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Design and Development of a 12:1 Pressure Ratio Compressor for the Ruston 6-MW Gas Turbine

Volume 1: Turbomachinery ◽

10.1115/82-gt-20 ◽

1982 ◽

Cited By ~ 1

Author(s):

F. Carchedi ◽

G. R. Wood

Keyword(s):

Gas Turbine ◽

Test Data ◽

Pressure Ratio ◽

Axial Flow ◽

Aerodynamic Design ◽

Design And Development ◽

Surge Line ◽

Flow Compressor ◽

Industrial Gas Turbine ◽

Spanwise Flow

This paper describes the design and development of a 15-stage axial flow compressor for a −6MW industrial gas turbine. Detailed aspects of the aerodynamic design are presented together with rig test data for the complete characteristic including stage data. Predictions of spanwise flow distributions are compared with measured values for the front stages of the compressor. Variable stagger stator blading is used to control the position of the low speed surge line and the effects of the stagger changes are discussed.

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Distributed Control of Rotating Stall in an Axial Flow Compressor With Actuator Constraints

10.1115/imece2000-2379 ◽

2000 ◽

Author(s):

Craig A. Buhr ◽

Matthew A. Franchek ◽

Sanford Fleeter

Keyword(s):

Absolute Stability ◽

Closed Loop ◽

Axial Flow ◽

Gain Control ◽

Rotating Stall ◽

Control Law ◽

Circle Criterion ◽

Axial Flow Compressor ◽

Flow Compressor ◽

Stall Control

Abstract Presented in this paper is an analytical study evaluating the closed loop stability of rotating stall control in an axial flow compressor subject to a nonlinear spatial actuation constraint that limits the amplitude of a spatial mode input. Absolute stability of the rotating stall control system is investigated by applying the circle criterion to a linearized model of an axial compressor in series with the saturation element. This stability analysis is then used to design the gain and phase of the ‘classical’ complex gain feedback control law. Resulting is a systematic method for designing the parameters of the complex gain control law which increases the region of absolute stability guaranteed by the circle criterion for the closed-loop system.

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2MW Class High Efficiency Gas Turbine IM270 for Co-Generation Plants

Volume 4: Heat Transfer; Electric Power; Industrial and Cogeneration ◽

10.1115/96-gt-001 ◽

1996 ◽

Cited By ~ 1

Author(s):

Hideo Kobayashi ◽

Shogo Tsugumi ◽

Yoshio Yonezawa ◽

Riuzou Imamura

Keyword(s):

Gas Turbine ◽

High Efficiency ◽

Mechanical Design ◽

Development Program ◽

Gas Turbine Engine ◽

Maintenance Cost ◽

Turbine Engine ◽

Generation Plant ◽

Emission Regulation ◽

Heavy Duty Gas Turbine

IHI is developing a new heavy duty gas turbine engine for 2MW class co-generation plants, which is called IM270. This engine is a simple cycle and single-spool gas turbine engine. Target thermal efficiency is the higher level in the same class engines. A dry low NOx combustion system has been developed to clear the strictest emission regulation in Japan. All parts of the IM270 are designed with long life for low maintenance cost. It is planned that the IM270 will be applied to a dual fluid system, emergency generation plant, machine drive engine and so on, as shown in Fig.1. The development program of IM270 for the co-generation plant is progress. The first prototype engine test has been started. It has been confirmed that the mechanical design and the dry low NOx system are practical. The component tuning test is being executed. On the other hand, the component test is concurrently in progress. The first production engine is being manufactured to execute the endurance test using a co-generation plant at the IHI Kure factory. This paper provides the conceptual design and status of the IM270 basic engine development program.

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Development of a 6 MW-Class High-Efficiency Gas Turbine M7A-01

Volume 4: Heat Transfer; Electric Power; Industrial and Cogeneration ◽

10.1115/94-gt-066 ◽

1994 ◽

Author(s):

T. Sugimoto ◽

K. Ikesawa ◽

S. Kajita ◽

W. Karasawa ◽

T. Kojima ◽

...

Keyword(s):

Gas Turbine ◽

Thermal Efficiency ◽

High Efficiency ◽

Axial Flow ◽

Inlet Temperature ◽

Gas Temperature ◽

Flow Compressor ◽

Exhaust Gas Temperature ◽

Cycle Efficiency

The M7A-01 gas turbine is a newly developed 6 MW class single-shaft machine. With its high simple-cycle efficiency and high exhaust gas temperature. it is particularly suited for use in electric power generation and co-generation applications. An advanced high efficiency axial-flow compressor, six can-type combustors, and a high inlet temperature turbine has been adopted. This results in a high thermal efficiency of 31.5% at the gas turbine output shaft and a high overall thermal efficiency of co-generation system. In addition, low NOx emissions from the combustors and a long service life permit long-term continuous operation under various environmental limitations. The results of the full load shop test, accelerated cyclic endurance test and extra severity tests verified that the performance, the mechanical characteristics and the emission have satisfied the initial design goals.

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Numerical Simulation of Steady Air Injection Flow Control Effects on a Transonic Axial Flow Compressor

Applied Mechanics and Materials ◽

10.4028/www.scientific.net/amm.224.352 ◽

2012 ◽

Vol 224 ◽

pp. 352-357

Author(s):

Islem Benhegouga ◽

Ce Yang

Keyword(s):

Mass Flow Rate ◽

Mass Flow ◽

Flow Rate ◽

Pressure Ratio ◽

Axial Flow ◽

Leading Edge ◽

Air Injection ◽

Axial Flow Compressor ◽

Blade Tip ◽

Flow Compressor

In this work, steady air injection upstream of the blade leading edge was used in a transonic axial flow compressor, NASA rotor 37. The injectors were placed at 27 % upstream of the axial chord length at blade tip, the injection mass flow rate is 3% of the chock mass flow rate, and 3 yaw angles were used, respectively -20°, -30°, and -40°. Negative yaw angles were measured relative to the compressor face in opposite direction of rotational speeds. To reveal the mechanism, steady numerical simulations were performed using FINE/TURBO software package. The results show that the stall mass flow can be decreased about 2.5 %, and an increase in the total pressure ratio up to 0.5%.

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