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.

Heat-Exchanged Propulsion Gas Turbines: A Candidate for Future Lower SFC and Reduced-Emission Military and Civil Aeroengines

2009 ◽  
Author(s):  
Colin F. McDonald ◽  
Colin Rodgers
Keyword(s):  
Gas Turbines ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Thermodynamic Cycle ◽  
Simple Cycle ◽  
Specific Power ◽  
Operational Experience ◽  
Fuel Cost ◽  
Turbofan Engines ◽  
Fuel Burn

After seven decades of service the evolution of simple cycle propulsion gas turbines continues with emphasis now being placed on reduced fuel burn, lower emissions, and less noise. With compressor and turbine efficiencies near plateauing, and turbine inlet temperatures paced by materials and blade cooling technologies, improvements in SFC, specific power and weight for conventional engines (including small turboprop, and turboshaft engines and larger turbofans) will likely be incremental compared with the past. With retention of the simple cycle both evolutionary and revolutionary approaches are being taken by the aeroengine industry to improve performance, particularly reduced fuel burn in an era of high fuel cost. In this paper a further step is suggested, that is in concert with meeting performance, economic, and environmental goals of future aeroengines, namely the use of a more complex thermodynamic cycle involving recuperation for turboprop and turboshaft engines, and intercooling together with recuperation for higher pressure ratio turbofan engines. The idea of heat exchanged propulsion gas turbines is not new, but the many concepts identified from studies done periodically over the last 65 years, including the few engines that were static tested and one test flown, didn’t find acceptance in an era of low fuel cost and concerns about recuperator integrity and reliability. With today’s very high fuel cost there is current interest in this type of engine because of its potential for low SFC and reduced emissions. In this paper potential applications are summarized and the features of various heat exchanged aeroengine design concepts together with projected performance are presented. Included is a discussion on the various issues that must be resolved before they enter service. A postulated deployment scenario is suggested with engines initially developed to meet military aviation needs, such as recuperated turboprop and turbofan engines for extended range UAV’s, followed by a recuperated turboshaft engine for a helicopter. Operational experience and demonstrated reliability from these would pave the way for high efficiency ICR turbofan engines for military and civil aircraft service sometime after the year 2020.

scholarly journals The Constant Volume Gas Turbine: A Thermodynamic Assessment

1974 ◽  
Author(s):  
M. Zockel
Keyword(s):  
Heat Exchanger ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Constant Pressure ◽  
Pressure Ratio ◽  
Constant Volume ◽  
Specific Power ◽  
Inlet Temperature ◽  
Turbine Efficiency ◽  
Compressor Pressure Ratio

A quasi-steady-state analysis is made of the performance of a gas-turbine working with intermittent, constant volume combustion. Variables considered include inlet temperature, compressor pressure ratio, scavenge ratio, combustion time, heat exchanger thermal ratio. Characteristics are computed over a full loading range. Computations are based on turbines having the following behavior: (a) constant turbine efficiency, (b) characteristics of a multistage axial turbine, and (c) characteristics of a single-stage radial turbine. The analysis indicates that the constant volume gas turbine has advantages in thermal efficiency, specific power and part load performance over constant pressure gas turbines operating at the same compressor pressure ratio and turbine inlet temperature. However, the addition of a heat exchanger shows less advantage when applied to a constant volume than to a constant pressure engine.

scholarly journals Advanced Core Technology: Key to Subsonic Propulsion Benefits

1989 ◽  
Author(s):  
Arthur J. Glassman ◽  
Christopher A. Snyder ◽  
Gerald Knip
Keyword(s):  
Thermal Efficiency ◽  
Pressure Ratio ◽  
Constant Current ◽  
Specific Power ◽  
Inlet Temperature ◽  
Turbofan Engines ◽  
Turbine Cooling ◽  
Core Technology ◽  
Specific Power Output ◽  
The Individual

A study was conducted to identify the potential performance benefits and key technology drivers associated with advanced cores for subsonic high-bypass turbofan engines. Investigated first were the individual sensitivities of varying compressor efficiency, pressure ratio and bleed (for turbine cooling); combustor pressure recovery; and turbine efficiency and inlet temperature on thermal efficiency and core specific power output. Then, engine cycle and mission performance benefits were determined for systems incorporating all potentially achievable technology advancements. The individual thermodynamic sensitivities are shown over a range of turbine temperatures (at cruise) from 2900 to 3500 °R and for both constant (current technology) and optimum (maximum thermal efficiency) overall pressure ratios. It is seen that no single parameter by itself will provide a large increase in core thermal efficiency, which is the thermodynamic parameter of most concern for transport propulsion. However, when all potentially achievable advancements are considered, there occurs a synergism that produces significant cycle and mission performance benefits. The nature of these benefits are presented and the technology challenges associated with achieving them are discussed.

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.

The Preliminary CFD Design of a Compressor and Combustor System Towards a 100 kW Hydrogen Fuelled Micro Gas Turbine

2019 ◽  
Author(s):  
C. Devriese ◽  
W. Pennings ◽  
H. de Reuver ◽  
R. Bastiaans ◽  
W. De Paepe
Keyword(s):  
Combustion Chamber ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Thermodynamic Cycle ◽  
Energy System ◽  
Inlet Temperature ◽  
Pure Hydrogen ◽  
Tidal Power ◽  
Large Power

Abstract Within the context of an ever-increasing share of wind, solar and emerging tidal power, the need to store energy, not only on the short term, but also in the medium to long-term to balance out the power grid will become more important in the near future. One of the most promising routes for this mid- to long term storage, is to produce hydrogen through electrolysis using excess electricity and store it. Instead of using this hydrogen then to generate electricity in a conventional, large, power plant, a more efficient route is to use it in a Decentralised Energy System (DES) using micro Gas Turbines (mGTs). Although the mGT presents itself as a promising option to convert pure hydrogen into electricity in this DES framework, several challenges, linked to the necessary increase of Turbine Inlet Temperature (TIT) for efficiency increase to make the unit compatible and the use of pure hydrogen in the combustor, still need to be overcome. In this paper we present the first steps towards a fully hydrogen fuelled mGT. Firstly, a full thermodynamic cycle analysis was performed to determine the optimal operating parameters, such as compressor pressure ratio and mass flow rate, air-to-fuel ratio and TIT. Secondly, a full CFD design and optimisation of the compressor and the combustion chamber was performed (steady and transient RANS and LES). CFD simulations of the compressor and combustion chamber matched the 1D performance calculations and also reached the desired performance goals. This CFD supported validation of the component performance shows that the design of a pure hydrogen combustion chamber for mGT applications is possible.

scholarly journals Performance Benefits for Wave Rotor-Topped Gas Turbine Engines

1996 ◽  
Author(s):  
Scott M. Jones ◽  
Gerard E. Welch
Keyword(s):  
Steady State ◽  
Gas Turbine ◽  
Engine Performance ◽  
Thermodynamic Cycle ◽  
Cycle Performance ◽  
Specific Power ◽  
Inlet Temperature ◽  
Wave Rotor ◽  
Turbine Engine ◽  
One Dimensional

The benefits of wave rotor-topping in turboshaft engines, subsonic high-bypass turbofan engines, auxiliary power units, and ground power units are evaluated. The thermodynamic cycle performance is modeled using a one-dimensional steady-state code; wave rotor performance is modeled using one-dimensional design/analysis codes. Design and off-design engine performance is calculated for baseline engines and wave rotor-topped engines, where the wave rotor acts as a high pressure spool. The wave rotor-enhanced engines are shown to have benefits in specific power and specific fuel flow over the baseline engines without increasing turbine inlet temperature. The off-design steady-state behavior of a wave rotor-topped engine is shown to be similar to a conventional engine. Mission studies are performed to quantify aircraft performance benefits for various wave rotor cycle and weight parameters. Gas turbine engine cycles most likely to benefit from wave rotor-topping are identified. Issues of practical integration and the corresponding technical challenges with various engine types are discussed.

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.

Optimization of Turbofan Propulsion Cycle Using a Genetic Algorithm

2010 ◽  
Author(s):  
Adel Ghenaiet
Keyword(s):  
Genetic Algorithm ◽  
Fuel Consumption ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Optimization Procedure ◽  
Specific Fuel Consumption ◽  
High Mass ◽  
Design Parameters ◽  
Inlet Temperature ◽  
Simplifying Assumptions

This paper presents an evolutionary approach as the optimization framework to design for the optimal performance of a high-bypass unmixed turbofan to match with the power requirements of a commercial aircraft. The parametric analysis had the objective to highlight the effects of the principal design parameters on the propulsive performance in terms of specific fuel consumption and specific thrust. The design optimization procedure based on the genetic algorithm PIKAIA coupled to the developed engine performance analyzer (on-design and off-design) aimed at finding the propulsion cycle parameters minimizing the specific fuel consumption, while meeting the required thrusts in cruise and takeoff and the restrictions of temperatures limits, engine size and weight as well as pollutants emissions. This methodology does not use engine components’ maps and operates on simplifying assumptions which are satisfying the conceptual or early design stages. The predefined requirements and design constraints have resulted in an engine with high mass flow rate, bypass ratio and overall pressure ratio and a moderate turbine inlet temperature. In general, the optimized engine is fairly comparable with available engines of equivalent power range.

Turbocharger-Design Effects on Gasoline-Engine Performance

2005 ◽  
Vol 127 (3) ◽  
pp. 525-530 ◽  
Author(s):  
Theodosios Korakianitis ◽  
T. Sadoi
Keyword(s):  
Steady Flow ◽  
Mass Flow ◽  
Gasoline Engine ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Thermodynamic Cycle ◽  
Final Choice ◽  
Design Point ◽  
Theoretical Considerations ◽  
Flow Experiments

Specification of a turbocharger for a given engine involves matching the turbocharger performance characteristics with those of the piston engine. Theoretical considerations of matching turbocharger pressure ratio and mass flow with engine mass flow and power permits designers to approach a series of potential turbochargers suitable for the engine. Ultimately, the final choice among several candidate turbochargers is made by tests. In this paper two types of steady-flow experiments are used to match three different turbochargers to an automotive turbocharged-intercooled gasoline engine. The first set of tests measures the steady-flow performance of the compressors and turbines of the three turbochargers. The second set of tests measures the steady-flow design-point and off-design-point engine performance with each turbocharger. The test results show the design-point and off-design-point performance of the overall thermodynamic cycle, and this is used to identify which turbocharger is suitable for different types of engine duties.

Pulsed Impingement Turbine Cooling and its Effect on the Efficiency of Gas Turbines With Pressure Gain Combustion

2020 ◽  
Author(s):  
Nicolai Neumann ◽  
Dieter Peitsch ◽  
Arne Berthold ◽  
Frank Haucke ◽  
Panagiotis Stathopoulos
Keyword(s):  
Gas Turbine ◽  
Thermal Efficiency ◽  
Gas Turbines ◽  
Turbine Blades ◽  
Impingement Cooling ◽  
Turbine Cooling ◽  
Pressure Gain Combustion ◽  
Cooling Air ◽  
Gas Turbine Cycle ◽  
Cycle Efficiency

Abstract Performance improvements of conventional gas turbines are becoming increasingly difficult and costly to achieve. Pressure Gain Combustion (PGC) has emerged as a promising technology in this respect, due to the higher thermal efficiency of the respective ideal gas turbine cycle. Previous cycle analyses considering turbine cooling methods have shown that the application of pressure gain combustion may require more turbine cooling air. This has a direct impact on the cycle efficiency and reduces the possible efficiency gain that can potentially be harvested from the new combustion technology. Novel cooling techniques could unlock an existing potential for a further increase in efficiency. Such a novel turbine cooling approach is the application of pulsed impingement jets inside the turbine blades. In the first part of this paper, results of pulsed impingement cooling experiments on a curved plate are presented. The potential of this novel cooling approach to increase the convective heat transfer in the inner side of turbine blades is quantified. The second part of this paper presents a gas turbine cycle analysis where the improved cooling approach is incorporated in the cooling air calculation. The effect of pulsed impingement cooling on the overall cycle efficiency is shown for both Joule and PGC cycles. In contrast to the authors’ anticipation, the results suggest that for relevant thermodynamic cycles pulsed impingement cooling increases the thermal efficiency of Joule cycles more significantly than it does in the case of PGC cycles. Thermal efficiency improvements of 1.0 p.p. for pure convective cooling and 0.5 p.p. for combined convective and film with TBC are observed for Joule cycles. But just up to 0.5 p.p. for pure convective cooling and 0.3 p.p. for combined convective and film cooling with TBC are recorded for PGC cycles.

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