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The purpose of this paper is to comprehensively evaluate the progress in the development of trapped vortex combustors (TVCs) in the past three decades. The review aims to identify the needs, predict the scope and discuss the challenges of numerical simulations in TVCs applied to gas turbines.

TVC is an emerging combustion technology for achieving low emissions in gas turbine combustors. The overall operation of such TVCs can be on very lean mixture ratio and hence it helps in achieving high combustion efficiency and low overall emission levels. This review introduces the TVC concept and the evolution of this technology in the past three decades. Various geometries that were explored in TVC research are listed and their operating principles are explained. The review then categorically arranges the progress in computational studies applied to TVCs.

Analyzing extensive literature on TVCs the review discusses results of numerical simulations of various TVC geometries. Numerical simulations that were used to optimize TVC geometry and to enhance mixing are discussed. Reactive flow studies to comprehend flame stability and emission characteristics are then listed for different TVC geometries.

To the best of the authors’ knowledge, this review is the first of its kind to discuss extensively the computational progress in TVC development specific to gas turbine engines. Earlier review on TVC covers a wide variety of applications including land-based gas turbines, supersonic Ramjets, incinerators and hence compromise on the depth of analysis given to gas turbine engine applications. This review also comprehensively group the numerical studies based on geometry, flow and operating conditions.

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.

The purpose of this study is to have semiempirical correlations for carbon monoxide (CO), unburned hydrocarbon (UHC) and nitrogen oxide (NO_x) emissions that are collected and calibrated by using experimental data of a tubular-type combustor.

Combustor uses a coflow radial-type air-blast atomizer and is especially designed for the empirical correlation issues. Air mass flow rate, air inlet temperature and air-to-fuel ratio parameters have been changed and different inlet conditions have been created for combustor tests. Six different inlet temperatures from 475 to 350 K have been set for each air mass flow rate. Air mass flow rate values from 0.035 to 0.050 kg/s have been used to create varied combustor aerodynamic loadings.

Increasing combustor inlet temperature decreases the CO and UHC emissions. However, it has an adverse effect in NO_x emissions. Moreover, CO and UHC emissions have an increasing trend by the mass flow rate rise that results an extra aerodynamic loading.

It is difficult to obtain real operating parameters for the combustor. Therefore, as a different approach in respect of the literature, rig test parameters have been used for thermodynamic calculations. Additionally, emission calculations of the combustor design point have been performed based on a conditioned test environment. Moreover, combustor outlet temperature and emission values have been scanned and mean values used for the analysis.

To perform preliminary calculations for these pollutants, designers need experimentally calibrated correlations for the similar combustors.

If the application area of the designed engine is a civil aircraft, emissions are one of the most important issues because of the strict regulations of International Civil Aviation Organization. Therefore, aviation companies are continuously working on reducing of emissions.

A comprehensive study for the preliminary emission calculation of newly designed gas turbine combustors was performed to investigate semiempirical correlations in the atmospheric test rig.

A combustor design is a particularly important and difficult task in the development of gas turbine engines. During studies for accurate and easy combustor design, reasonable design methodologies have been established and used in engine development. The purpose of this paper is to review the design methodology for combustor in development of advanced gas turbine engines. The advanced combustor development task can be successfully achieved in less time and at lower cost by adopting new and superior design methodologies.

The review considers the main technical problems (combustion, cooling, fuel injection and ignition technology) in the development of modern combustor design and deals with combustor design methods by dividing it into preliminary design, performance evaluation, optimization and experiment. The advanced combustion and cooling technologies mainly used in combustor design are mentioned in detail. In accordance with the modern combustor design method, the design mechanisms are considered and the methods used in every stage of the design are reviewed technically.

The improved performances and strict emission limits of gas turbine engines require the application of advanced technologies when designing combustors. The optimized design mechanism and reasonable performance evaluation methods are very important in reducing experiments and increasing the effectiveness of the design.

This paper provides a comprehensive review of the design methodology for the advanced gas turbine engine combustor.

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

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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.

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

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.

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.

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.

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.

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.

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.

This paper aims to predict the effect of combustor inlet area ratio (CIAR) on the lean blowout limit (LBO) of a swirl stabilized can-type micro gas turbine combustor having a thermal capacity of 3 kW.

The blowout limits of the combustor were predicted predominantly from numerical simulations by using the average exit gas temperature (AEGT) method. In this method, the blowout limit is determined from characteristics of the average exit gas temperature of the combustion products for varying equivalence. The CIAR value considered in this study ranges from 0.2 to 0.4 and combustor inlet velocities range from 1.70 to 6.80 m/s.

The LBO equivalence ratio decreases gradually with an increase in inlet velocity. On the other hand, the LBO equivalence ratio decreases significantly especially at low inlet velocities with a decrease in CIAR. These results were backed by experimental results for a case of CIAR equal to 0.2.

Gas turbine combustors are vulnerable to operate on lean equivalence ratios at cruise flight to avoid high thermal stresses. A flame blowout is the main issue faced in lean operations. Based on literature and studies, the combustor lean blowout performance significantly depends on the primary zone mass flow rate. By incorporating variable area snout in the combustor will alter the primary zone mass flow rates by which the combustor will experience extended lean blowout limit characteristics.

This is a first effort to predict the lean blowout performance on the variation of combustor inlet area ratio on gas turbine combustor. This would help to extend the flame stability region for the gas turbine combustor.