Limitations on Gas Turbine Performance Imposed by Large Turbine Cooling Flows

2001 ◽  
Vol 123 (3) ◽  
pp. 487-494 ◽  
Author(s):  
J. H. Horlock ◽  
D. T. Watson ◽  
T. V. Jones
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Loss ◽  
Combustion Temperature ◽  
Pressure Ratio ◽  
Stagnation Temperature ◽  
Optimum Pressure ◽  
Pressure Losses ◽  
Semi Empirical ◽  
Cooling Air

Calculations of the performance of modern gas turbines usually include allowance for cooling air flow rate; assumptions are made for the amount of the cooling air bled from the compressor, as a fraction of the mainstream flow, but this fractional figure is often set in relatively arbitrary fashion. There are two essential effects of turbine blade cooling: (i) the reduction of the gas stagnation temperature at exit from the combustion chamber (entry to the first nozzle row) to a lower stagnation temperature at entry to the first rotor and (ii) a pressure loss resulting from mixing the cooling air with the mainstream. Similar effects occur in the following cooled blade rows. The paper reviews established methods for determining the amount of cooling air required and semi-empirical relations, for film cooled blading with thermal barrier coatings, are derived. Similarly, the pressure losses related to elements of cooling air leaving at various points round the blade surface are integrated over the whole blade. This gives another semi-empirical expression, this time for the complete mixing pressure loss in the blade row, as a function of the total cooling air used. These two relationships are then used in comprehensive calculations of the performance of a simple open-cycle gas turbine. for varying combustion temperature and pressure ratio. These calculations suggest that for maximum plant efficiency there may be a limiting combustion temperature (below that which would be set by stoichiometric combustion). For a given combustion temperature, the optimum pressure ratio is reduced by the effect of cooling air.

Limitations on Gas Turbine Performance Imposed by Large Turbine Cooling Flows

2000 ◽  
Author(s):  
J. H. Horlock ◽  
D. T. Watson ◽  
T. V. Jones
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Loss ◽  
Combustion Temperature ◽  
Pressure Ratio ◽  
Stagnation Temperature ◽  
Optimum Pressure ◽  
Pressure Losses ◽  
Semi Empirical ◽  
Cooling Air

Calculations of the performance of modern gas turbines usually include allowance for cooling air flow rate; assumptions are made for the amount of the cooling air bled from the compressor, as a fraction of the mainstream flow, but this fractional figure is often set in relatively arbitrary fashion. There are two essential effects of turbine blade cooling: [i] the reduction of the gas stagnation temperature at exit from the combustion chamber [entry to the first nozzle row] to a lower stagnation temperature at entry to the first rotor and [ii] a pressure loss resulting from mixing the cooling air with the mainstream. Similar effects occur in the following cooled blade rows. The paper reviews established methods for determining the amount of cooling air required and semi-empirical relations, for film cooled blading with thermal barrier coatings, are derived. Similarly, the pressure losses related to elements of cooling air leaving at various points round the blade surface are integrated over the whole blade. This gives another semi-empirical expression, this time for the complete mixing pressure loss in the blade row, as a function of the total cooling air used. These two relationships are then used in comprehensive calculations of the performance of a simple open-cycle gas turbine, for varying combustion temperature and pressure ratio. These calculations suggest that for maximum plant efficiency there may be a limiting combustion temperature [below that which would be set by stoichiometric combustion]. For a given combustion temperature, the optimum pressure ratio is reduced by the effect of cooling air.

scholarly journals Impact of Inlet Filter Pressure Loss on Single and Two-Spool Gas Turbine Engines for Different Control Modes

2014 ◽  
Vol 136 (9) ◽  
Author(s):  
Uyioghosa Igie ◽  
Orlando Minervino
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Rotational Speed ◽  
Gas Turbines ◽  
Pressure Loss ◽  
Pressure Ratio ◽  
Two Stage ◽  
Pressure Losses ◽  
Stage System ◽  
Pressure Compressor

Inlet filtration systems are designed to protect industrial gas turbines from air borne particles and foreign objects, thereby improving the quality of air for combustion and reducing component fouling. Filtration systems are of varying grades and capture efficiencies, with the higher efficiency systems filters providing better protection but higher pressure losses. For the first time, two gas turbine engine models of different configurations and capacities have been investigated for two modes of operation (constant turbine entry temperature (TET) and load/power) for a two- and three-stage filter system. The main purpose of this is to present an account on factors that could decide the selection of filtration systems by gas turbine operators, solely based on performance. The result demonstrates that the two-spool engine is only slightly more sensitive to intake pressure loss relative to the single-spool. This is attributed to higher pressure ratio of the two-spool as well as the deceleration of the high pressure compressor (HPC)/high pressure turbine (HPT) shaft rotational speed in a constant TET operation. The compressor of the single-spool engine and the low pressure compressor (LPC) of the two-spool shows similar behavior: slight increase in pressure ratio and reduced surge margin at their constant rotational speed operation. Loss in shaft power is observed for both engines, about 2.5% at 1000 Pa loss. For constant power operation there is an increase in fuel flow and TET, and as a result the creep life was estimated. The result obtained indicates earlier operating hours to failure for the three-stage system over the two-stage by only a few thousand hours. However, this excludes any degradation due to fouling that is expected to be more significant in the two-stage system.

scholarly journals Methanol Dissociative Intercooling in Gas Turbines

1982 ◽  
Author(s):  
M. F. Bardon ◽  
J. A. C. Fortin
Keyword(s):  
Carbon Monoxide ◽  
Theoretical Analysis ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Unit Mass ◽  
Cycle Model ◽  
Optimum Pressure ◽  
Heating Value ◽  
Compressor Work

This paper examines the possibility of injecting methanol into the compressor of a gas turbine, then dissociating it to carbon monoxide and hydrogen so as to cool the air and reduce the work of compression, while simultaneously increasing the fuel’s heating value. A theoretical analysis shows that there is a net reduction in compressor work resulting from this dissociative intercooling effect. Furthermore, by means of a computer cycle model, the effects of dissociation on efficiency and work per unit mass of airflow are predicted for both regenerated and unregenerated gas turbines. The effect on optimum pressure ratio is examined and practical difficulties likely to be encountered with such a system are discussed.

Selection of Gas Turbine With Minimum Flue Gas Emission

2007 ◽  
Author(s):  
P. Esna Ashari ◽  
V. Nayyeri ◽  
K. Sarabchee
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Power Plants ◽  
Fossil Fuels ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Effective Parameters ◽  
Combustion Chambers ◽  
Optimum Pressure ◽  
Oil Refineries

Many factories in industry such as petrochemical plants, oil refineries and power plants need heat and power to support their process. This demand can be provided by a combined heat and power cycle (CHP) in the factory site. Some factories use gas turbine cycle to provide heat and power. Emissions from gas turbines, produced by burning fossil fuels in the combustion chambers, have important effects on air pollution. This is a significant problem in many developed and developing countries. Parameters such as inlet temperature and pressure ratio are the most effective parameters in gas turbine emission. By selecting an appropriate gas turbine, emission could be reduced to some extent. Further studies indicate that there is an optimum pressure ratio, which minimizes emissions.

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.

Effects of Air Induct Structure to Flow Uniformity and Resistance for Primary Surface Recuperator of a Micro Gas Turbine

2010 ◽  
Author(s):  
Wei Qu ◽  
Shan Gao
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Loss ◽  
Air Flow ◽  
Flow Distribution ◽  
Structure Design ◽  
Flow Uniformity ◽  
Local Pressure ◽  
Pressure Losses ◽  
Micro Gas Turbine

Primary surface recuperator is important for micro gas turbines, the flow distribution and pressure loss are sensitive to the induct structure design significantly. The air induct structure for one recuperator is modelled and simulated. Several flow fields and pressure losses are obtained for different designs of air induct structure. The air induct structure can affect the flow uniformity, further influence the pressure loss a lot. For several changes of air induct structure, the non-distribution of air flow can be decreased from 67% to 13%, and the pressure loss can be decreased to 50% of the original. Considering the recuperator design and the gas turbine, one optimized structure is recommended, which has less local pressure loss and better flow distribution.

Numerical Investigation of the Effect of an Axial Pre-Swirl Nozzle With a Radial Angle in a Pre-Swirl Rotor-Stator System

2021 ◽  
Author(s):  
Gang Zhao ◽  
Shuiting Ding ◽  
Tian Qiu ◽  
Shenghui Zhang
Keyword(s):  
Gas Turbines ◽  
Pressure Loss ◽  
Total Pressure ◽  
Discharge Coefficient ◽  
Pressure Ratio ◽  
Temperature Drop ◽  
Tangential Velocity ◽  
Total Pressure Loss ◽  
Adiabatic Effectiveness ◽  
Cooling Air

Abstract Pre-swirl nozzles are often used in gas turbines to deliver the cooling air to the turbine blades. The static axial nozzles swirl the cooling air in the direction of rotation of the turbine disk, thereby reducing the relative total temperature of the air. Most studies about nozzles focus on its shape, radial location, tangential angle to reduce the pressure loss and increase the temperature drop of the pre-swirl system, but few of them consider the benefit of a radial angle of nozzles. This paper investigated numerically the performance of a pre-swirl system whose pre-swirl nozzles have a radial angle. Six radial angles are selected to study the flow dynamics of a direct-transfer pre-swirl system in terms of the total pressure loss coefficient of the pre-swirl cavity, the discharge coefficient of the receiver holes, and the adiabatic effectiveness. It is shown that the nozzles with radial angles can adjust the tangential velocity and radial velocity and thus can influence the performance of a pre-swirl system. There is a lowerest value of total pressure loss in pre-swirl cavity, that is α = 90°, which can hardly be influenced by the radial angle of nozzle and pressure ratio π. For a specific swirl ratio β∞, there exists an optimal αopt where the discharge coefficient of receiver hole is maximum. Moreover, αopt decreases as pressure ratio π increases. And so is the adiabatic effectiveness Θad.

Advances in Computer Aided Engineering for Gas Turbine Research and Development

2021 ◽  
pp. 171-197
Author(s):  
Valeriu Vilag ◽  
Jeni Vilag ◽  
Cleopatra Cuciumita
Keyword(s):  
Carbon Dioxide ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Combustion Temperature ◽  
Pressure Ratio ◽  
Thermodynamic Cycle ◽  
Working Fluid ◽  
Atmospheric Air ◽  
Computer Aided ◽  
Main Components

Gas turbines represent energetic machines that operate following the Brayton thermodynamic cycle [6], [16] creating mechanical power and/or thrust. The majority of them use atmospheric air as working fluid but some are working in a closed loop with either air, carbon dioxide or other convenient gas. The main parameters defining the cycle are the combustion temperature and the overall pressure ratio [20]. In principle, the higher the combustion temperature is, the higher the efficiency of the entire gas turbine will be and, for a chosen value of the combustion temperature an optimum overall pressure ratio exists [29]. The main components of the gas turbine are:

A Study of Cycle Performance of Ceramic Gas Turbines

1997 ◽  
Author(s):  
H. Sugishita ◽  
H. Mori ◽  
R. Chikami ◽  
Y. Tsukuda ◽  
S. Yoshino ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Performance Improvement ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Combined Cycle ◽  
Cycle Performance ◽  
Gas Temperature ◽  
Blade Cooling ◽  
Combined Cycles ◽  
Cooling Air

A study has been carried out to assess the performance improvement of a combined cycle used for an industrial power plant when ceramic turbine components are employed. This paper presents the details of this study. Performance improvement is obtained as a result of reduced blade cooling air. In this study four different kinds of combined cycles were investigated and these are listed below: A. Combined cycle with a simple gas turbine. B. Combined cycle with an inter-cooled gas turbine. C. Combined cycle with a reheat gas turbine. D. Combined cycle with an inter-cooled reheat gas turbine. Results of this study indicate that the combined cycle with a simple gas turbine is the most practical of the four cycles studied with an efficiency of higher than 60%. The combined cycle with reheat gas turbine has the highest efficiency if a higher compressor exit air temperature and a high gas temperature (over 1000°C) to reheat the combustion system are used. A higher pressure ratio is required to optimize the cycle performance of the combined cycle with the ceramic turbine components than that with the metal turbine components because of reduced blade cooling air. To minimize leakage air for these higher pressure ratios, advanced seal technology should be applied to the gas turbines.

scholarly journals Feasibility of a Helium Closed-Cycle Gas Turbine for UAV Propulsion

2020 ◽  
Vol 11 (1) ◽  
pp. 28
Author(s):  
Emmanuel O. Osigwe ◽  
Arnold Gad-Briggs ◽  
Theoklis Nikolaidis
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Low Noise ◽  
Closed Cycle ◽  
Propulsion Systems ◽  
Component Design ◽  
Readiness Level ◽  
Aerial Vehicle ◽  
Turbine Design

When selecting a design for an unmanned aerial vehicle, the choice of the propulsion system is vital in terms of mission requirements, sustainability, usability, noise, controllability, reliability and technology readiness level (TRL). This study analyses the various propulsion systems used in unmanned aerial vehicles (UAVs), paying particular focus on the closed-cycle propulsion systems. The study also investigates the feasibility of using helium closed-cycle gas turbines for UAV propulsion, highlighting the merits and demerits of helium closed-cycle gas turbines. Some of the advantages mentioned include high payload, low noise and high altitude mission ability; while the major drawbacks include a heat sink, nuclear hazard radiation and the shield weight. A preliminary assessment of the cycle showed that a pressure ratio of 4, turbine entry temperature (TET) of 800 °C and mass flow of 50 kg/s could be used to achieve a lightweight helium closed-cycle gas turbine design for UAV mission considering component design constraints.

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