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Article
Publication date: 1 May 1992

M.C. MELAAEN

A solution algorithm for the numerical calculation of isothermal fluid flow inside gas turbine combustors is presented. The finite‐volume method together with curvilinear…

Abstract

A solution algorithm for the numerical calculation of isothermal fluid flow inside gas turbine combustors is presented. The finite‐volume method together with curvilinear non‐orthogonal coordinates and a non‐staggered grid arrangement is employed. Cartesian velocity components are chosen as dependent variables in the momentum equations. The turbulent flow inside the combustor is modelled by the k—ε turbulence model. The grid is generated by solving elliptic equations. This solution algorithm, which can be used on both can‐type and annular combustors, is tested on a water model can‐type combustor because of the availability of geometrical and experimental data for comparison.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 2 no. 5
Type: Research Article
ISSN: 0961-5539

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Article
Publication date: 1 September 1971

CF6–50A engines have been selected to power the European A300B aircraft scheduled for certification and airline delivery in 1974.

Abstract

CF6–50A engines have been selected to power the European A300B aircraft scheduled for certification and airline delivery in 1974.

Details

Aircraft Engineering and Aerospace Technology, vol. 43 no. 9
Type: Research Article
ISSN: 0002-2667

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Article
Publication date: 4 July 2008

Colin F. McDonald, Aristide F. Massardo, Colin Rodgers and Aubrey Stone

This paper seeks to evaluate the potential of heat exchanged aeroengines for future Unmanned Aerial Vehicle (UAV), helicopter, and aircraft propulsion, with emphasis…

Abstract

Purpose

This paper seeks to evaluate the potential of heat exchanged aeroengines for future Unmanned Aerial Vehicle (UAV), helicopter, and aircraft propulsion, with emphasis placed on reduced emissions, lower fuel burn, and less noise.

Design/methodology/approach

Aeroengine performance analyses were carried out covering a wide range of parameters for more complex thermodynamic cycles. This led to the identification of major component features and the establishing of preconceptual aeroengine layout concepts for various types of recuperated and ICR variants.

Findings

Novel aeroengine architectures were identified for heat exchanged turboshaft, turboprop, and turbofan variants covering a wide range of applications. While conceptual in nature, the results of the analyses and design studies generally concluded that heat exchanged engines represent a viable solution to meet demanding defence and commercial aeropropulsion needs in the 2015‐2020 timeframe, but they would require extensive development.

Research limitations/implications

As highlighted in Parts I and II, early development work was focused on the use of recuperation, but this is only practical with compressor pressure ratios up to about 10. For today's aeroengines with pressure ratios up to about 50, improvement in SFC can only be realised by incorporating intercooling and recuperation. The new aeroengine concepts presented are clearly in an embryonic stage, but these should enable gas turbine and heat exchanger specialists to advance the technology by conducting more in‐depth analytical and design studies to establish higher efficiency and “greener” gas turbine aviation propulsion systems.

Originality/value

It is recognised that meeting future environmental and economic requirements will have a profound effect on aeroengine design and operation, and near‐term efforts will be focused on improving conventional simple‐cycle engines. This paper has addressed the longer‐term potential of heat exchanged aeroengines and has discussed novel design concepts. A deployment strategy, aimed at gaining confidence with emphasis placed on assuring engine reliability, has been suggested, with the initial development and flight worthiness test of a small recuperated turboprop engine for UAVs, followed by a larger recuperated turboshaft engine for a military helicopter, and then advancement to a larger and far more complex ICR turbofan engine.

Details

Aircraft Engineering and Aerospace Technology, vol. 80 no. 4
Type: Research Article
ISSN: 0002-2667

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Article
Publication date: 9 April 2019

Dongmei Zhao, Yifan Xia, Haiwen Ge, Qizhao Lin, Jianfeng Zou and Gaofeng Wang

Ignition process is a critical issue in combustion systems. It is particularly important for reliability and safety prospects of aero-engine. This paper aims to…

Abstract

Purpose

Ignition process is a critical issue in combustion systems. It is particularly important for reliability and safety prospects of aero-engine. This paper aims to numerically investigate the burner-to-burner propagation during ignition process in a full annular multiple-injector combustor and then validate it by comparing with experimental results.

Design/methodology/approach

The annular multiple-injector experimental setup features 16 swirling injectors and two quartz tubes providing optical accesses to high-speed imaging of flames. A Reynolds averaged Navier–Stokes model, adaptive mesh refinement (AMR) and complete San Diego chemistry are used to predict the ignition process.

Findings

The ignition process shows an overall agreement with experiment. The integrated heat release rate of simulation and the integrated light intensity of experiment is also within reasonable agreement. The flow structure and flame propagation dynamics are carefully analyzed. It is found that the flame fronts propagate symmetrically at an early stage and asymmetrically near merging stage. The flame speed slows down before flame merging. Overall, the numerical results show that the present numerical model can reliably predict the flame propagation during the ignition process.

Originality/value

The dedicated AMR method together with detailed chemistry is used for predicting the unsteady ignition procedure in a laboratory-scale annular combustor for the first time. The validation shows satisfying agreements with the experimental investigations. Some details of flow structures are revealed to explain the characteristics of unsteady flame propagations.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 29 no. 6
Type: Research Article
ISSN: 0961-5539

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Article
Publication date: 6 November 2018

Raja Marudhappan, Chandrasekhar Udayagiri and Koni Hemachandra Reddy

The purpose of this paper is to formulate a structured approach to design an annular diffusion flame combustion chamber for use in the development of a 1,400 kW range aero…

Abstract

Purpose

The purpose of this paper is to formulate a structured approach to design an annular diffusion flame combustion chamber for use in the development of a 1,400 kW range aero turbo shaft engine. The purpose is extended to perform numerical combustion modeling by solving transient Favre Averaged Navier Stokes equations using realizable two equation k-e turbulence model and Discrete Ordinate radiation model. The presumed shape β-Probability Density Function (β-PDF) is used for turbulence chemistry interaction. The experiments are conducted on the real engine to validate the combustion chamber performance.

Design/methodology/approach

The combustor geometry is designed using the reference area method and semi-empirical correlations. The three dimensional combustor model is made using a commercial software. The numerical modeling of the combustion process is performed by following Eulerian approach. The functional testing of combustor was conducted to evaluate the performance.

Findings

The results obtained by the numerical modeling provide a detailed understanding of the combustor internal flow dynamics. The transient flame structures and streamline plots are presented. The velocity profiles obtained at different locations along the combustor by numerical modeling mostly go in-line with the previously published research works. The combustor exit temperature obtained by numerical modeling and experiment are found to be within the acceptable limit. These results form the basis of understanding the design procedure and opens-up avenues for further developments.

Research limitations/implications

Internal flow and combustion dynamics obtained from numerical simulation are not experimented owing to non-availability of adequate research facilities.

Practical implications

This study contributes toward the understanding of basic procedures and firsthand experience in the design aspects of combustors for aero-engine applications. This work also highlights one of the efficient, faster and economical aero gas turbine annular diffusion flame combustion chamber design and development.

Originality/value

The main novelty in this work is the incorporation of scoops in the dilution zone of the numerical model of combustion chamber to augment the effectiveness of cooling of combustion products to obtain the desired combustor exit temperature. The use of polyhedral cells for computational domain discretization in combustion modeling for aero engine application helps in achieving faster convergence and reliable predictions. The methodology and procedures presented in this work provide a basic understanding of the design aspects to the beginners working in the gas turbine combustors particularly meant for turbo shaft engines applications.

Details

Aircraft Engineering and Aerospace Technology, vol. 91 no. 1
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 21 September 2010

Chun‐Hsiang Yang, Di‐Han Wu and Chiun‐Hsun Chen

Utilizing renewable energy and developing new energy sources are practical responses to the shortage of fossil fuels and environmental regulations for carbon dioxide…

Abstract

Purpose

Utilizing renewable energy and developing new energy sources are practical responses to the shortage of fossil fuels and environmental regulations for carbon dioxide emissions. The purpose of this paper is to assess the practicability of using low heating value (LHV) fuel on an annular miniature gas turbine (MGT) via numerical simulations.

Design/methodology/approach

The MGT used in this study is MW‐44 Mark I, whose original fuel is liquid (Jet A1). Its fuel supply system is re‐designed to use biogas fuel with LHV. The simulations, aided by the commercial code CFD‐ACE+, were carried out to investigate the cooling effect in a perforated combustion chamber and combustion behavior in an annular MGT when using LHV gas. In this study, four parameters of rotational speeds are considered. At each specific speed, various mixture ratios of methane (CH4) to carbon dioxide (CO2) including 90, 80, 70, and 60 percent were taken into consideration as simulated LHV fuels.

Findings

The simulation results show the chamber design can create a proper recirculation zone to concentrate the flame at the center of the chamber, and prevent the flame from expanding to cause hot spot. Furthermore, the hot gas exhausted from combustor outlet is cooled down effectively by jet flow discharged from dilution holes, which prevent turbine blade from heat damage.

Originality/value

Simulation results demonstrate that CFD‐ACE+ can simulate flow field performance and combustion behavior in an annular MGT precisely. The results of these CFD analyses confirm that the methane fuel can be used in such small volume of MGT and still have high performance. With the aid of the constructed combustor model, the performance of a methane‐used MGT can be realized before the experiment procedure starts.

Details

International Journal of Numerical Methods for Heat & Fluid Flow, vol. 20 no. 7
Type: Research Article
ISSN: 0961-5539

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Article
Publication date: 8 April 2021

Kirubakaran V. and David Bhatt

The lean blowout (LBO) limit of the combustor is one of the important performance parameters for any gas turbine combustor design. This study aims to predict the LBO…

Abstract

Purpose

The lean blowout (LBO) limit of the combustor is one of the important performance parameters for any gas turbine combustor design. This study aims to predict the LBO limits of an in-house designed swirl stabilized 3kW can-type micro gas turbine combustor.

Design/methodology/approach

The experimental prediction of LBO limits was performed on 3kW swirl stabilized combustor fueled with methane for the combustor inlet velocity ranging from 1.70 m/s to 6.80 m/s. The numerical prediction of LBO limits of combustor was performed on two-dimensional axisymmetric model. The blowout limits of combustor were predicted through calculated average exit gas temperature (AEGT) method and compared with experimental predictions.

Findings

The results show that the predicted LBO equivalence ratio decreases gradually with an increase in combustor inlet velocity.

Practical implications

This LBO limits predictions will use to fix the operating boundary conditions of 3kW can-type micro gas turbine combustor. This methodology will be used in design stage as well as in the testing stage of the combustor.

Originality/value

This is a first effort to predict the LBO limits on micro gas turbine combustor through AEGT method. The maximum uncertainty in LBO limit prediction with AEGT is 6 % in comparison with experimental results.

Details

Aircraft Engineering and Aerospace Technology, vol. 93 no. 4
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 5 March 2018

Yasin Şöhret and T. Hikmet Karakoc

It is essential to develop more environment-friendly energy systems to prevent climate change and minimize environmental impact. Within this scope, many studies are…

Abstract

Purpose

It is essential to develop more environment-friendly energy systems to prevent climate change and minimize environmental impact. Within this scope, many studies are performed on performance and environmental assessments of many types of energy systems. This paper, different from previous studies, aims to prove exergy performance of a low-emission combustor of an aero-engine.

Design/methodology/approach

It is a well-known fact that, with respect to previous exergy analysis, highest exergy destruction occurs in the combustor component of the engine. For this reason, it is required to evaluate a low-emission aero-engine combustor thermodynamically to understand the state of the art according to the authors’ best of knowledge. In this framework, combustor has been operated at numerous conditions (variable engine load) and evaluated.

Findings

As a conclusion of the study, the impact of emission reduction on performance improvement of the aero-engine combustors exergetically is presented. It is stated that exergy efficiency of the low-emission aero-engine combustor is found to be 64.69, 61.95 and 71.97 per cent under various operating conditions.

Practical implications

Results obtained in this paper may be beneficial for researchers who are interested in combustion and propulsion technology and thermal sciences.

Originality/value

Different from former studies, the impact of operating conditions on performance of a combustor is examined from the viewpoint of thermodynamics.

Details

Aircraft Engineering and Aerospace Technology, vol. 90 no. 2
Type: Research Article
ISSN: 1748-8842

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Article
Publication date: 3 October 2016

Chaozhi Cai, Leyao Fan and Bingsheng Wu

This paper aims to understand the outlet temperature distribution of the combustor of a high-temperature, high-speed heat-airflow simulation system.

Abstract

Purpose

This paper aims to understand the outlet temperature distribution of the combustor of a high-temperature, high-speed heat-airflow simulation system.

Design/methodology/approach

The paper uses numerical simulation to study the temperature distribution of the combustor of a high-temperature, high-speed heat-airflow simulation system. First, the geometrical model of the combustor and the combustion model of the fuel are established. Then, the combustion of fuel in the combustor is simulated by using FLUENT under various conditions. Finally, the results are obtained.

Findings

The paper found three conclusions: when the actual fuel–gas ratio is equal to the theoretical fuel–gas ratio, the temperature in the combustor of the high-temperature, high-speed heat-airflow simulation system (HTSAS) can reach its highest and the distribution is the most uniform. Although increases in the total temperature of the inlet air can increase the highest temperature in the combustor of the HTSAS, the average temperature of the combustor outlet will decrease. At the same time, it will lead to an uneven temperature distribution of the combustor outlet. When the spray angle of the kerosene droplet is at 30 degrees, the outlet temperature field of the combustor is more uniform.

Originality/value

The paper presents a method to analyze the combustion performance of fuel and the gas temperature distribution in the combustor. The results will lay the foundation for the gas temperature control of a combustor.

Details

World Journal of Engineering, vol. 13 no. 5
Type: Research Article
ISSN: 1708-5284

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Article
Publication date: 4 September 2017

Ahmet Topal and Onder Turan

The purpose of this study is to perform an exergy analysis of a turbojet engine combustor at different cycle parameters.

Abstract

Purpose

The purpose of this study is to perform an exergy analysis of a turbojet engine combustor at different cycle parameters.

Design/methodology/approach

Base cycle parameters have been defined for the engine, and then differentiation of the combustor exergy efficiencies and destruction rates have been evaluated by changing overall pressure ratio, combustor exit temperature and combustor pressure ratio.

Findings

For the basic engine cycle, combustor unit is found to have lowest exergy efficiency as 62.3 for the sea level static condition. Compressor turbine exhaust and whole engine exergy efficiencies have been calculated as 88.7, 96.5, 68.2 and 69.4, respectively.

Practical implications

Because of the biggest exergy, destruction is seen mainly in combustion system; effect of the combustor inlet pressure (related to the compressor design technology), pressure drop and exit temperature on the exergy efficiencies have been analyzed and combustor second law efficiency have been evaluated.

Social implications

The investigation’s purposes are highly connected with social wellness and targeted at sustainable development of the society. Practical implementation of the obtained scientific results is directed on the improving of combustor for a turbojet engines and decreasing negative influence on the environment.

Originality/value

As a result of this paper, the following are the contribution of this paper in the field of gas turbine exergy subjects: Combustor has been found as the most critical component in respect of the exergy efficiency. Therefore, the effect of the combustor main cycle parameters such as inlet pressure, combustor pressure ratio and exit temperature have been analyzed.

Details

Aircraft Engineering and Aerospace Technology, vol. 89 no. 5
Type: Research Article
ISSN: 1748-8842

Keywords

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