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.

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.

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.

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.

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.

scholarly journals Analysis and Improvement of a Two-Stage Centrifugal Compressor Used in an MW-Level Gas Turbine

2018 ◽  
Vol 8 (8) ◽  
pp. 1347 ◽  
Author(s):  
Wei Zhu ◽  
Xiao-Dong Ren ◽  
Xue-Song Li ◽  
Chun-Wei Gu
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Two Stage ◽  
Vaned Diffuser ◽  
Pressure Stage ◽  
Optimization System ◽  
High Pressure Stage

The performance of a low/high-pressure-stage centrifugal compressor in a land-use MW-level gas turbine with a pressure ratio of approximately 11 is analyzed and optimized with a 1D aerodynamic design and modeling optimization system. 1D optimization results indicate that the diameter ratio of the low-pressure-stage centrifugal compressor with a vane-less diffuser, and the divergent angle of the high-pressure-stage centrifugal compressor with a vaned diffuser, are extremely large and result in low efficiency. Through modeling design and optimization system analysis, a tandem vaned diffuser is used in the low-pressure stage, and a tandem vaned diffuser with splitter vanes is adopted in the high-pressure stage. Computational fluid dynamics (CFD) results show that the pressure ratio and efficiency of the optimized low/high-pressure-stage centrifugal compressor are significantly improved. Coupling calculations of the low/high-pressure stage of the original and optimized designs are conducted based on the results of MW-level gas turbine cycles. CFD results show that the pressure ratio and efficiency of the optimized two-stage centrifugal compressor increase by approximately 8% and 4%, respectively, under three typical load conditions of 100%, 90%, and 60%.

MAN’s New Gas Turbines for Mechanical Drive and Power Generation Applications

2013 ◽  
Author(s):  
Emil Aschenbruck ◽  
Michele Cagna ◽  
Volker Langusch ◽  
Ulrich Orth ◽  
Andreas Spiegel ◽  
...  
Keyword(s):  
Power Generation ◽  
Gas Turbine ◽  
Gas Turbines ◽  
High Efficiency ◽  
Pressure Ratio ◽  
High Heat ◽  
Product Line ◽  
Pollutant Emission ◽  
Two Stage ◽  
Exhaust Temperature

MAN Diesel & Turbo recently developed a completely new gas turbine family for the first time in its history. The first product line contains both two-shaft and single-shaft gas turbines in the 6 – 7 MW class. The two-shaft engine was thoroughly tested at MAN’s gas turbine test center, and the first engine has been delivered to a launch customer. For MAN, it constitutes a technology platform that will produce further developments and new models in the coming years. The two-shaft design makes the new gas turbine an ideal mechanical drive for both turbo compressors and pumps. This gas turbine operates to suit the optimum duty point of the driven machine; both in a wide speed and power range. The two stage power turbine design allows for a wide speed range of 45 to 105% while maintaining high efficiency. For power generation a single-shaft version has been created by adding one additional stage to the two stage high pressure turbine. The compressor pressure ratio is 15, which is high enough for achieving the highest potential efficiency for both generator and compressor drive applications. Low pollutant emission levels are achieved with MAN’s DLN combustion technology. The gas turbine exhaust temperature is sufficiently high to reach high heat recovery rates in combined heat and power cycles. Another important feature of the new gas turbine is its unrestricted suitability for taking load quickly and rapid load changes. Service costs have also been significantly improved upon. MAN opted for a sturdy and modular gas turbine construction, while not compromising on efficiency. The objective is to extend service life and shorten down time occurrences. The modular package assembly process helps to reduce routine maintenance and repair time, and ultimately package downtime.

The Experimental Investigations of Centripetal Air Bleed With Tubed Vortex Reducer for Secondary Air System in Gas Turbine

2014 ◽  
Author(s):  
Xiao Chen ◽  
Ye Feng ◽  
Lijun Wu
Keyword(s):  
High Pressure ◽  
Flow Field ◽  
Gas Turbine ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Pressure Loss ◽  
Experimental Investigations ◽  
Pressure Losses ◽  
Nozzle Guide

In a modern gas turbine, the air bled through High Pressure Compressor (HPC) rotor drums from the main flow is transported radially inwards and then transferred to cool the High Pressure Turbine (HPT). The centripetal air flow creates a strong vortex, which results in huge pressure losses. This not only restricts the mass flow rate, but also reduces the cooling air pressure for down-stream hot components. Adding vortex reducer tubes to the centripetal air bleed can reduce the pressure loss and ensure the pressure and mass flow rate of the supply air. Design optimization of the tubed vortex reducer is essential in minimizing the pressure losses. This paper describes experimental investigations of different configurations of tubed vortex reducers at different rotational speeds and mass flow rates. Particular attention is paid to the shape of the drum hole, the length of the tubed vortex reducers at the same installation location, and the angles of the nozzle guide vane outlets. The core section of test rig is comprised of two steel disks, one drum rotor and stationary cases with nozzle guide vanes. It operates at representative engine parameters, such as the turbulent flow parameter, λT(0.2–1.8) and the Rossby number Ro(0.05–0.08). Three conclusions can be drawn based on the experimental results. 1) The shape of the drum hole is a key factor of the bleed system pressure loss. An oval hole configuration has less flow resistance and results in lower pressure losses compared with a circular hole design. 2) The tests prove that tubed vortex reducers are instrumental in minimizing centripetal air flow. These components effectively restrain the free vortex development and decrease the pressure losses in the cavity. 3) Basically, the flow field consists of a free vortex and a forced vortex. The length of the tube influences the flow field and the pressure losses at the inlet and outlet of the tubed vortex reducer. However, the tube length is less important when compared with the shape of drum hole.

Experimental Measurements of Pressure Losses in the Inter-Turbine Duct of a Gas Turbine

2015 ◽  
Vol 789-790 ◽  
pp. 540-548 ◽  
Author(s):  
Cleopatra Florentina Cuciumita ◽  
Daniel Olaru ◽  
Valeriu Vilag ◽  
Ionut Porumbel ◽  
Sergiy Riznyk ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Mass Flow Rate ◽  
Flow Rate ◽  
Rotational Speed ◽  
Pressure Loss ◽  
Total Pressure ◽  
Operating Regime ◽  
Total Pressure Loss ◽  
Experimental Measurements ◽  
Pressure Losses

The paper presents the total pressure experimental measurements carried out at the Romanian Research and Development Institute for Gas Turbines COMOTI in order to determine the total pressure losses in the Inter - Turbine Duct of a two spools gas turbine, as a function of the gas turbine operating regime (mass flow rate) and rotational speed. The Inter - Turbine Duct experimental assembly has been designed, manufactured and tested at COMOTI. The total pressures were measured as a function of the pre-swirling angle, which simulates the influence of the high pressure turbine rotational speed located upstream of the Inter turbine duct in the real gas turbine, as well as for three operational regimes, without the pre-swirlers modules. The results indicate that the total pressure loss along the Inter - Turbine Duct is of maximum 0.9 %. The lowest overall total pressure loss occurs at 0o pre-swirling angle, around 0.8%, while along the ITD struts, the lowest pressure loss is obtained for a 15o pre-swirling, below 0.1%. The influence of the operating regime on the total pressure loss was found to be linearly, the pressure loss increasing with the reduced mass flow rate, between 1% and 1.9% overall, and between about 0.1% and 0.4 % along the struts.

Modeling and Performance Analysis of a Dual-Shaft Counter-Rotating Gas Turbine

2018 ◽  
Author(s):  
Qiao Zhou ◽  
Zhao Yin ◽  
Chun-qing Tan ◽  
Qing Gao ◽  
Yong-sheng Tian
Keyword(s):  
High Pressure ◽  
Performance Analysis ◽  
Gas Turbine ◽  
Rotational Speed ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Modeling Method ◽  
Speed Ratio ◽  
Shaft Speed ◽  
Wide Range

This paper proposes a modeling method for the 1+1/2 vaneless counter rotating turbine (VCRT) and carries out performance analysis of a micro gas turbine (MGT) with VCRT at off design conditions. The first task of this paper is to obtain performance characteristics maps and develop a characteristics modeling method of VCRT. The VCRT characteristics maps are obtained through 3-D CFD calculation. High pressure turbine (HPT) corrected rotational speed, shaft rotating speed ratio defined as the high pressure (HP) shaft rotational speed divided by low pressure (LP) shaft rotational speed, and the VCRT total expansion pressure ratio are selected to model a 3-D low pressure turbine (LPT) characteristics maps. However, the HPT characteristics map modeling method remains the same as the conventional one. An overall performance simulation model is established in Matlab/Simulink and validated by software GasTurb. The VCRT engine consumes fuel at a higher rate when delivering same power compared with the conventional gas turbine due to LPT performance degradation. When the relative LP shaft speed ranges from 100% to 96%, the performance of the VCRT engine and conventional engine is almost equivalent. As the LP shaft speed continues to drop off, however, the VCRT engine performance degrades considerably. The results indicate that it is crucial to design VCRT with a wide range of efficiency, especially the LPT. The VCRT engine control law also must be effectively optimized to ensure the engine performs well at part load working conditions.

Numerical Investigation of Vortex Reducer Flows in the High Pressure Compressor of Modern Aeroengines

2002 ◽  
Author(s):  
Dieter Peitsch ◽  
Manuela Stein ◽  
Stefan Hein ◽  
Reinhard Niehuis ◽  
Ulf Reinmo¨ller
Keyword(s):  
High Pressure ◽  
Gas Turbines ◽  
Efficient Operation ◽  
Pressure Losses ◽  
High Pressure Compressor ◽  
Air System ◽  
The Difference ◽  
Secondary Air ◽  
Set Up ◽  
Pressure Compressor

Modern jet engines require very high cycle temperatures for efficient operation. In turn, cooling air is needed for the turbine, since the materials are not yet capable of taking these temperatures. Air is taken from the compressor for the purpose of cooling and turbine rim sealing, bypassing the main combustion circuit. Since this affects the efficiency of the engine in a negative manner, measures are taken to reduce the amount of air to an absolute minimum. These measures include the investigation of reducing pressure losses within the involved subsystems. One of these subsystems in the BR700 aeroengine series of Rolls-Royce is the vortex reducer device, which delivers bleed air to the secondary air system of the engine. The German government has set up a research project, aiming for an overall improvement of aeroengines. This program, Engine 3E, where 3E reflects Efficiency, Economy and Environment, concentrates on the main components of gas turbines. Programmes for the high pressure turbine and for the combustion chamber have been set up. The high pressure compressor has been identified as key component as well. A new 9-stage compressor is being developed at Rolls-Royce Deutschland to adress the respective needs. From the point of view of the secondary air system, the vortex reducer in this component plays a major role with respect to the efficient use of cooling and sealing air. Rolls-Royce Deutschland has performed CFD studies on the performance of different vortex reducer geometries, which currently are considered for incorporation into the future engine. The results of these investigations wil be converted into more simple design rules for proper reflection of the behaviour of this system for future designs. The paper presents the set up of the geometries, the applied boundary conditions as well as the final results. To tackle the difference between a high pressure compressor rig and a typical two-shaft engine, a dedicated investigation to assess the difference between a pure high pressure core without an internal shaft and a realistic high/low pressure shaft configuration has been carried out and is included in the paper. Recommendations to improve the design with respect to minimized pressure losses will be shown as well.

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