Design and Development of a 14-Stage Axial Compressor for Industrial Gas Turbine

2012 ◽  
Author(s):  
Takuya Ikeguchi ◽  
Akinori Matsuoka ◽  
Yusuke Sakai ◽  
Yoshinobu Sakano ◽  
Kenichiro Yoshiura
Keyword(s):  
Gas Turbine ◽  
Structural Reliability ◽  
Scale Effects ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Geometric Optimization ◽  
Cfd Analysis ◽  
Aerodynamic Design ◽  
Stall Margin ◽  
Industrial Gas Turbine

A 14-stage axial flow compressor was newly designed and tested for developing an advanced industrial gas turbine. In order to achieve a high thermal efficiency required for the new gas turbine, the compressor needed to have a significantly higher pressure ratio and higher efficiency than those of existing engines. The new design methodology used to this compressor design was based on an automated airfoil geometric optimization system combined with a 3D-CFD analysis, which resulted in arbitrary shaped airfoil design in most blade rows. A multi-stage CFD analysis was used effectively in order to adjust a loading distribution along stages and to obtain a proper stage matching. Before the full development of the gas turbine, an approximately two-thirds scaled compressor rig tests were conducted to verify the aerodynamic design and the structural reliability. The test results of the first build indicated a satisfactory level of efficiency and mass flow, but with a lack of sufficient stall margin. The second build with the re-staggered vanes was tested and its result showed improvements both in stall margin and in efficiency. The prototype test of developing an industrial gas turbine also had been conducted. The measured performance of the compressor which was scaled up from the second build rig compressor achieved the design target. Consequently, the aerodynamic design which considered the scale effects of the compressor was successful.

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

1982 ◽  
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.

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

1982 ◽  
Vol 104 (4) ◽  
pp. 823-831 ◽  
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

The paper describes the design and development of a 15 stage axial flow compressor for a 6-MW 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.

A New 1.7MW Industrial Gas Turbine: The Ruston Hurricane

1991 ◽  
Author(s):  
A. T. Sanders ◽  
M. H. Tothill ◽  
G. R. Wood
Keyword(s):  
Gas Turbine ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Thermodynamic Efficiency ◽  
The Novel ◽  
Exhaust Temperature ◽  
High Pressure Ratio ◽  
Industrial Gas Turbine ◽  
Novel Design

The paper describes the design of a compact new 1.7MW (2300hp) single shaft industrial gas turbine and package, with high efficiency and exhaust temperature ideal for industrial congeneration applications. These advantages are obtained with a high pressure ratio single stage centrifugal compressor, single high temperature combustor and two-stage axial flow turbine using only one row of cooled blades. The novel design features are described with the associated development testing. A typical installation is also described showing the potential for very high overall thermodynamic efficiency.

Improved Compressor Exit Diffuser for an Industrial Gas Turbine

2001 ◽  
Vol 124 (1) ◽  
pp. 19-26 ◽  
Author(s):  
U. Orth ◽  
H. Ebbing ◽  
H. Krain ◽  
A. Weber ◽  
B. Hoffmann
Keyword(s):  
Gas Turbine ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Aerodynamic Design ◽  
Turbine Efficiency ◽  
Noticeable Improvement ◽  
Manufacturing Procedure ◽  
Multistage Compressor ◽  
Back Stage ◽  
Industrial Gas Turbine

Cycle studies carried out for the medium pressure ratio gas turbine THM 1304 of 10 MW power output manufactured by MAN Turbomaschinen AG GHH BORSIG predicted that the overall efficiency of the multistage compressor, composed of a ten-stage axial and a single-stage centrifugal compressor, would improve by 0.8 percent if the efficiency of the back stage centrifugal unit could be raised by 4 percent. It was expected that this would result in a noticeable improvement of the thermal gas turbine efficiency. The paper describes the aerodynamic design process used for the stage optimization, applying today’s advanced design tools for blade generation and three-dimensional aerodynamic calculation methods. Additionally, it describes the manufacturing procedure for the resulting three-dimensional blades and the experimental verification of the design approach.

scholarly journals Aerodynamic Design and Testing of an Axial Flow Compressor With Pressure Ratio of 23.3:1 for the LM2500+ Gas Turbine

1999 ◽  
Author(s):  
A. R. Wadia ◽  
D. P. Wolf ◽  
F. G. Haaser
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Performance Test ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Flow ◽  
Combined Cycle ◽  
Aerodynamic Design ◽  
Point Selection ◽  
Technical Requirements

The LM2500+ gas turbine, rated between 39,000 to 40,200 shaft horsepower (shp), was introduced for field service in 1998. This growth aero-derivative gas turbine is suitable for a variety of power generation applications, such as co-generation and combined cycle, as well as mechanical drive applications. At the heart of the LM2500+ 25% power increase is an up-rated derivative 17-stage axial compressor. This paper describes the aerodynamic design and development of this high pressure ratio single spool compressor for the LM2500+ gas turbine. The compressor is derived by zero-staging the highly efficient and reliable LM2500 compressor to increase the flow by 23% at a pressure ratio of 23.3:1. The aerodynamic efficiency of the compressor is further improved by using three-dimensional, custom-tailored airfoil designs similar to those used in the CF6-80C2 high pressure compressor. The compressor achieved a peak polytropic efficiency above 91 percent, meeting all its operability objectives. The technical requirements and overall aerodynamic design features of the compressor are presented first. Next, the zero stage match point selection is described and the procedure used to set up the vector diagrams using a through-flow code with secondary flow and mixing is outlined. Detailed design results for the new transonic airfoils in the compressor using three-dimensional viscous analysis are presented. The compressor instrumentation and performance test results are discussed. The performance of the zero stage is separated from that of the baseline compressor with the CF6-80C2 airfoils to show the improvement in efficiency with the new airfoils.

Aerodynamic Design and Testing of an Axial Flow Compressor With Pressure Ratio of 23.3:1 for the LM2500+ Gas Turbine

2002 ◽  
Vol 124 (3) ◽  
pp. 331-340 ◽  
Author(s):  
A. R. Wadia ◽  
D. P. Wolf ◽  
F. G. Haaser
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Performance Test ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Flow ◽  
Combined Cycle ◽  
Aerodynamic Design ◽  
Point Selection ◽  
Technical Requirements

The LM2500+ gas turbine, rated between 39,000–40,200 shaft horsepower (shp), was introduced for field service in 1998. This growth aero-derivative gas turbine is suitable for a variety of power generation applications, such as co-generation and combined cycle, as well as mechanical drive applications. At the heart of the LM2500+ 25% power increase is an up-rated derivative 17-stage axial compressor. This paper describes the aerodynamic design and development of this high-pressure ratio single-spool compressor for the LM2500+ gas turbine. The compressor is derived by zero-staging the highly efficient and reliable LM2500 compressor to increase the flow by 23% at a pressure ratio of 23.3:1. The aerodynamic efficiency of the compressor is further improved by using three-dimensional, custom-tailored airfoil designs similar to those used in the CF6-80C2 high-pressure compressor. The compressor achieved a peak polytropic efficiency above 91%, meeting all its operability objectives. The technical requirements and overall aerodynamic design features of the compressor are presented first. Next, the zero stage match point selection is described and the procedure used to set up the vector diagrams using a through-flow code with secondary flow and mixing is outlined. Detailed design results for the new transonic airfoils in the compressor using three-dimensional viscous analysis are presented. The compressor instrumentation and performance test results are discussed. The performance of the zero stage is separated from that of the baseline compressor with the CF6-80C2 airfoils to show the improvement in efficiency with the new airfoils.

scholarly journals Axial Flow Compressor Stability Enhancement: Circumferential T-Shape Grooves Performance Investigation

Aerospace ◽  
2021 ◽  
Vol 8 (1) ◽  
pp. 12
Author(s):  
Marco Porro ◽  
Richard Jefferson-Loveday ◽  
Ernesto Benini
Keyword(s):  
Vortex Breakdown ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Suction Side ◽  
Treatment Technique ◽  
Tip Leakage Flow ◽  
Stall Inception ◽  
Stall Margin ◽  
Casing Treatment ◽  
Stall Margin Improvement

This work focuses its attention on possibilities to enhance the stability of an axial compressor using a casing treatment technique. Circumferential grooves machined into the case are considered and their performances evaluated using three-dimensional steady state computational simulations. The effects of rectangular and new T-shape grooves on NASA Rotor 37 performances are investigated, resolving in detail the flow field near the blade tip in order to understand the stall inception delay mechanism produced by the casing treatment. First, a validation of the computational model was carried out analysing a smooth wall case without grooves. The comparisons of the total pressure ratio, total temperature ratio and adiabatic efficiency profiles with experimental data highlighted the accuracy and validity of the model. Then, the results for a rectangular groove chosen as the baseline case demonstrated that the groove interacts with the tip leakage flow, weakening the vortex breakdown and reducing the separation at the blade suction side. These effects delay stall inception, improving compressor stability. New T-shape grooves were designed keeping the volume as a constant parameter and their performances were evaluated in terms of stall margin improvement and efficiency variation. All the configurations showed a common efficiency loss near the peak condition and some of them revealed a stall margin improvement with respect to the baseline. Due to their reduced depth, these new configurations are interesting because they enable the use of a thinner light-weight compressor case as is desirable in aerospace applications.

scholarly journals The Design and Evaluation of a 14 Stage, 16:1 Pressure Ratio Compressor for an Industrial Gas Turbine

1996 ◽  
Author(s):  
Mohammad R. Saadatmand
Keyword(s):  
Experimental Data ◽  
Gas Turbine ◽  
Rotor Blade ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Operating Conditions ◽  
Suction Surface ◽  
Design Intent ◽  
Flow Through ◽  
Industrial Gas Turbine

The aerodynamic design process leading to the production configuration of a 14 stage, 16:1 pressure ratio compressor for the Taurus 70 gas turbine is described. The performance of the compressor is measured and compared to the design intent. Overall compressor performance at the design condition was found to be close to design intent. Flow profiles measured by vane mounted instrumentation are presented and discussed. The flow through the first rotor blade has been modeled at different operating conditions using the Dawes (1987) three-dimensional viscous code and the results are compared to the experimental data. The CFD prediction agreed well with the experimental data across the blade span, including the pile up of the boundary layer on the corner of the hub and the suction surface. The rotor blade was also analyzed with different grid refinement and the results were compared with the test data.

scholarly journals Industrial Gas Turbine Uprates: Tips, Tricks and Traps

1997 ◽  
Author(s):  
T. L. Ragland
Keyword(s):  
Gas Turbine ◽  
Output Power ◽  
Gas Turbines ◽  
Customer Value ◽  
Low Cost ◽  
Pressure Ratio ◽  
Original Design ◽  
Advantages And Disadvantages ◽  
Industrial Gas Turbine ◽  
Cycle Efficiency

After industrial gas turbines have been in production for some amount of time, there is often an opportunity to improve or “uprate” the engine’s output power or cycle efficiency or both. In most cases, the manufacturer would like to provide these uprates without compromising the proven reliability and durability of the product. Further, the manufacturer would like the development of this “Uprate” to be low cost, low risk and result in an improvement in “customer value” over that of the original design. This paper describes several options available for enhancing the performance of an existing industrial gas turbine engine and discusses the implications for each option. Advantages and disadvantages of each option are given along with considerations that should be taken into account in selecting one option over another. Specific options discussed include dimensional scaling, improving component efficiencies, increasing massflow, compressor zero staging, increasing firing temperature (thermal uprate), adding a recuperator, increasing cycle pressure ratio, and converting to a single shaft design. The implications on output power, cycle efficiency, off-design performance engine life or time between overhaul (TBO), engine cost, development time and cost, auxiliary requirements and product support issues are discussed. Several examples are provided where these options have been successfully implemented in industrial gas turbine engines.

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