Evaluation of Performance Gain by Interstage Injection in a Four-Stage Axial Compressor

2021 ◽  
Author(s):  
Tobias Doerr ◽  
Sebastian Schuster ◽  
Dieter Brillert
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Injection Rate ◽  
Local Injection ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Axial Compressors ◽  
Evaluation Of Performance ◽  
Mass Flow Rates

Abstract Recently, the energy market has seen a shift towards renewable energies due to changing demands. Gas turbines are used as a transitional technology to cope with grid fluctuations. The changing conditions have increased the interest in applying Wet Compression in order to increase the power output during peak demands. The novelty of this paper arises from the experimental results of Interstage Injection by analysing the stage and overall pressure ratios at different operating points in the four stage axial compressor “eco.MAC” (“evaporative cooling Multiphase Axial Compressor”). An innovative injection design is realized with twin jet nozzles in the trailing edge of SLM printed stator blades. A variation of water mass fraction, inlet temperature and rotational speed is performed and shows a gain in pressure ratio up to 1.5 %. Moreover, a polynomial approach is used for the dry data to compare wet and dry results at equal air mass flow rates. For the first time, a linear dependency of the pressure gain on the compressor’s gas temperature is experimentally found. It can be concluded that Interstage Injection is an effective technology to be applied in later stages of axial compressors due to the strong influence of local gas temperatures on the evaporation rate and thus the pressure gain. Furthermore, reducing the local injection rate decreases aerodynamic losses between the liquid and gas phase. Hence, a multiple injection and reduced local injection rates should be targeted.

scholarly journals New Diagram and Equation for Thermodynamic Calculations for Gas Turbines

1984 ◽  
Author(s):  
B. Herrmann
Keyword(s):  
Gas Turbines ◽  
Liquid Fuel ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Iso Standard ◽  
Compressor Pressure Ratio ◽  
German Standard ◽  
Exhaust Gas Temperature ◽  
Acceptance Tests

On basis of ISO-Standard 2314, the German Standard Organisation (DIN) has prepared the German Standard DIN 4341, which deals with acceptance tests for gas turbines. Sample calculations have been included. In connection with the development of the sample calculations a new diagram for thermodynamic properties of air and products of combustion was developed on basis of -humid air as per ISO standard 2314 -standard gaseous fuel -standard liquid fuel This diagram allows exact calculation of performance data. Further, a simplified but relatively acurate formula is presented for calculating the turbine inlet temperature on basis of -compressor pressure ratio -exhaust gas temperature -thermal efficiency Development and limitation of this formula is presented.

Application of Ultra-Low NOx Combustor to the MHPS Existing Gas Turbine

2021 ◽  
Author(s):  
Takashi Nishiumi ◽  
Hirofumi Ohara ◽  
Kotaro Miyauchi ◽  
Sosuke Nakamura ◽  
Toshishige Ai ◽  
...  
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Combined Cycle ◽  
Inlet Temperature ◽  
Operational Experience ◽  
Emission Rates ◽  
Gas Temperature ◽  
Design Development ◽  
And Control

Abstract In recent years, MHPS achieved a NET M501J gas turbine combined cycle (GTCC) efficiency in excess of 62% operating at 1,600°C, while maintaining NOx under 25ppm. Taking advantage of our gas turbine combustion design, development and operational experience, retrofits of earlier generation gas turbines have been successfully applied and will be described in this paper. One example of the latest J-Series technologies, a conventional pilot nozzle was changed to a premix type pilot nozzle for low emission. The technology was retrofitted to the existing F-Series gas turbines, which resulted in emission rates of lower than 9ppm NOx(15%O2) while maintaining the same Turbine Inlet Temperature (TIT: Average Gas Temperature at the exit of the transition piece). After performing retrofitting design, high pressure rig tests, the field test prior to commercial operation was conducted on January 2019. This paper describes the Ultra-Low NOx combustor design features, retrofit design, high pressure rig test and verification test results of the upgraded M501F gas turbine. In addition, it describes another upgrade of turbine to improve efficiency and of combustion control system to achieve low emissions. Furthermore it describes the trouble-free upgrade of seven (7) units, which was completed by utilizing MHPS integration capabilities, including handling all the design, construction and service work of the main equipment, plant and control systems.

Off-Design Prediction of Transonic Axial Compressors: Part 1 — Mean-Line Code and Tuning Factors

2018 ◽  
Author(s):  
John Kidikian ◽  
Marcelo Reggio
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Design Tool ◽  
Information Storage ◽  
Axial Compressors ◽  
Design Performance ◽  
Engineering Software ◽  
Line Loss ◽  
Loss Modelling ◽  
Line Analysis

With yearly advances in CFD techniques and methodologies, and the increased capacity and capabilities of computer CPU, GPU, and information storage, CFD has become a powerful design tool. However, despite its vast strengths, a CFD analysis is still based on the sound development of the 1D mean-line analysis methodology. This paper (part 1 of 2) describes an off-design axial compressor mean-line code, tested in a specialized engineering software for the development and analysis of a whole gas turbine engine, and the various tuning factors used to obtain an off-design performance match. It will be shown that, to obtain a proper match of the off-design performance of single-stage transonic axial compressors, both the rotor and stage pressure ratio, and the rotor temperature ratio are required to be converged upon. To do so, the off-design mean-line analysis requires the incorporation of a set of inlet & exit blockage factors and deviation angles that vary with the compressor performance conditions. This approach differs from the literature-based procedural assumptions (or rule-of-thumb) of fixed inlet and exit blockage factors of approximately “0.98”, and the use of a unique deviation angle based on Carter’s rule. The results obtained in this paper are then used to develop a generalized off-design mean-line loss modelling methodology (part 2 of 2) capable of predicting the off-design performance of four well documented NASA transonic axial compressors.

Improving Steady CFD to Capture the Effects of Radial Mixing in Axial Compressors

2019 ◽  
Author(s):  
Lorenzo Cozzi ◽  
Filippo Rubechini ◽  
Andrea Arnone ◽  
Andrea Schneider ◽  
Pio Astrua
Keyword(s):  
Steady State ◽  
Gas Turbines ◽  
Turbulent Viscosity ◽  
Computational Cost ◽  
Axial Compressor ◽  
Secondary Flows ◽  
Tip Clearance ◽  
Axial Compressors ◽  
Rans Model ◽  
Aspect Ratios

Abstract The axial compressors of power-generation gas turbines have a high stage count, blades with low aspect ratios and relatively large clearances in the rear section. These features promote the development of strong secondary flows. An important outcome deriving from the convection of intense secondary flows is the enhanced span-wise transport of fluid properties mainly involving the rear stages, generally referred to as “radial mixing”. An incorrect prediction of this key phenomenon may result in inaccurate performance evaluation and could mislead the designers during the compressor design phase. As shown in a previous work, in the rear stages of an axial compressor the stream-wise vorticity associated with tip clearance flows is one of the main drivers of the overall span-wise transport phenomenon. Limiting it by circumferentially averaging the flow at row interfaces is the reason why a steady-state analysis strongly under-predicts radial mixing. To properly forecast the span-wise transport within the flow-path, an unsteady analysis should be adopted. However, due to the high blade count, this approach has a computational cost not yet suitable for industrial purposes. Currently, only the steady-state full-compressor simulation can fit in a lean industrial design chain and any model upgrade improving its radial mixing prediction would be highly beneficial for the daily design practice. To attain some progresses in RANS model, its inherent lack of convection of stream-wise vorticity must be addressed. This can be done by acting on another mixing driver, able to provide the same outcome, that is turbulent diffusion. In particular, by enhancing turbulent viscosity one can promote span-wise diffusion, thus improving the radial mixing prediction of the steady approach. In this paper, this strategy to update the RANS model and its application in simulations on a compressor of the Ansaldo Energia fleet is presented, together with the model tuning that has been performed using the results of unsteady simulations as the target.

The Influence of Fan Root Flow on the Aerodynamics of a Low-Pressure Compressor Transition Duct

2019 ◽  
Author(s):  
A. D. Walker ◽  
I. Mariah ◽  
D. Tsakmakidou ◽  
H. Vadhvana ◽  
C. Hall
Keyword(s):  
Gas Turbines ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Profile ◽  
Low Pressure ◽  
Secondary Flows ◽  
Test Facility ◽  
Inlet Flow ◽  
Flat Profile ◽  
Bypass Ratio

Abstract To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the higher required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. Additionally, the desire for shorter, lighter nacelles means that the engine intake may be unable provide a uniform inlet flow to the fan when the aircraft is at an angle of attack or subject to cross winds. Any inlet distortion this generates can also further degrade the quality of the flow entering the core of the engine. This paper uses a combination of experiments and CFD to examine the effects of the inlet flow on the aerodynamics of an engine section splitter and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A fully annular test facility incorporating a 1½ stage axial compressor was used to compare the system performance of a rotor that produced a nominally flat profile with one that had a notably hub deficient flow. A RANS CFD model, employing a mixing plane between the rotor and Engine Section Stator (ESS) and a Reynolds Stress turbulence model, was then validated and used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. Overall, changes to the inlet condition were seen to have a surprisingly small effect on the flow at duct exit — i.e. the flow presented to the downstream compressor. Changes to the inlet did, however, generate increased secondary flows and degrade the performance of the ESS. This resulted in notably increased total pressure loss; in excess of 12% for the hub-low inlet and in excess of 30% at high inlet swirl where the flow in the ESS separated. However, the increased ESS wake structures, and the enhanced mixing, delayed separation in the duct suggesting that, overall the design was reasonably robust, albeit with a significant penalty in system loss.

Off-Design Performance Investigation of a Low Calorific Value Gas Fired Generic-Type Single-Shaft Gas Turbine

2008 ◽  
Vol 130 (3) ◽  
Author(s):  
Raik C. Orbay ◽  
Magnus Genrup ◽  
Pontus Eriksson ◽  
Jens Klingmann
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Calorific Value ◽  
Working Fluid ◽  
Inlet Temperature ◽  
Flammability Limits ◽  
Generic Type ◽  
Wobbe Index ◽  
The Impact

When low calorific value gases are fired, the performance and stability of gas turbines may deteriorate due to a large amount of inertballast and changes in working fluid properties. Since it is rather rare to have custom-built gas turbines for low lower heating value (LHV) operation, the engine will be forced to operate outside its design envelope. This, in turn, poses limitations to usable fuel choices. Typical restraints are decrease in Wobbe index and surge and flutter margins for turbomachinery. In this study, an advanced performance deck has been used to quantify the impact of firing low-LHV gases in a generic-type recuperated as well as unrecuperated gas turbine. A single-shaft gas turbine characterized by a compressor and an expander map is considered. Emphasis has been put on predicting the off-design behavior. The combustor is discussed and related to previous experiments that include investigation of flammability limits, Wobbe index, flame position, etc. The computations show that at constant turbine inlet temperature, the shaft power and the pressure ratio will increase; however, the surge margin will decrease. Possible design changes in the component level are also discussed. Aerodynamic issues (and necessary modifications) that can pose severe limitations on the gas turbine compressor and turbine sections are discussed. Typical methods for axial turbine capacity adjustment are presented and discussed.

Comprehensive Application of a First Principles Based Methodology for Design of Axial Compressor Configurations

2012 ◽  
Vol 134 (6) ◽  
Author(s):  
Vishwas Iyengar ◽  
Lakshmi N. Sankar
Keyword(s):  
First Principles ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Aerodynamic Characteristics ◽  
Design Parameters ◽  
Axial Compressors ◽  
Design Variables ◽  
Improved Accuracy ◽  
Total Pressure Ratio

Axial compressors are widely used in many aerodynamic applications. The design of an axial compressor configuration presents many challenges. It is necessary to retool the design methodologies to take advantage of the improved accuracy and physical fidelity of these advanced methods. Here, a first-principles based multiobjective technique for designing single stage compressors is described. The study accounts for stage aerodynamic characteristics and rotor-stator interactions. The proposed methodology provides a way to systematically screen through the plethora of design variables. This method has been applied to a rotor-stator stage similar to NASA Stage 35. By selecting the most influential design parameters and by optimizing the blade leading edge and trailing edge mean camber line angles, phenomena such as tip blockages, blade-to-blade shock structures and other loss mechanisms can be weakened or alleviated. It is found that these changes to the configuration can have a beneficial effect on total pressure ratio and stage adiabatic efficiency, thereby improving the performance of the axial compression system.

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.

Thermo-Economic Analysis of an Intercooled, Reheat and Recuperated Gas Turbine for Cogeneration Applications: Part I — Base Load Operation

2000 ◽  
Author(s):  
R. Bhargava ◽  
M. Bianchi ◽  
G. Negri di Montenegro ◽  
A. Peretto
Keyword(s):  
Economic Analysis ◽  
Energy Saving ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Waste Heat ◽  
Pressure Ratio ◽  
Gas Temperature ◽  
Minimum Value ◽  
Maximum Value ◽  
Cogeneration System

This paper presents a thermo-economic analysis of an intercooled, reheat (ICRH) gas turbine, with and without recuperation, for cogeneration applications. The optimization analyses of thermodynamic parameters have permitted to calculate variables, such as low-pressure compressor pressure ratio, high-pressure turbine pressure ratio and gas temperature at the waste heat recovery unit inlet while maximizing electric efficiency and “Energy Saving Index”. Subsequently, the economic analyses have allowed to evaluate return on the investment, and the minimum value of gross payout period, for the cycle configurations of highest thermodynamic performance. In the present study three sizes (100 MW, 20 MW and 5 MW) of gas turbines have been examined. The performed investigation reveals that the maximum value of electric efficiency and “Energy Saving Index” is achieved for a large size (100 MW) recuperated ICRH gas turbine based cogeneration system. However, a non-recuperated ICRH gas turbine (of 100 MW) based cogeneration system provides maximum value of return on the investment and the minimum value of gross payout period compared to the other gas turbine cycles, of the same size and with same power to heat ratio, investigated in the present study. A comprehensive thermo-economic analysis methodology, presented in this paper, should provide useful guidelines for preliminary sizing and selection of gas turbine cycle for cogeneration applications.

Thermodynamic Evaluation of Gas Turbine Cogeneration Cycles: Part I—Heat Balance Method Analysis

1987 ◽  
Vol 109 (1) ◽  
pp. 1-7 ◽  
Author(s):  
I. G. Rice
Keyword(s):  
Gas Turbine ◽  
Heat Balance ◽  
Gas Turbines ◽  
Heat Recovery ◽  
Pressure Ratio ◽  
Turbine Rotor ◽  
Balance Method ◽  
Inlet Temperature ◽  
Thermodynamic Evaluation ◽  
Exhaust Temperature

This paper presents a heat balance method of evaluating various open-cycle gas turbines and heat recovery systems based on the first law of thermodynamics. A useful graphic solution is presented that can be readily applied to various gas turbine cogeneration configurations. An analysis of seven commercially available gas turbines is made showing the effect of pressure ratio, exhaust temperature, intercooling, regeneration, and turbine rotor inlet temperature in regard to power output, heat recovery, and overall cycle efficiency. The method presented can be readily programmed in a computer, for any given gaseous or liquid fuel, to yield accurate evaluations. An X–Y plotter can be utilized to present the results.

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