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

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.

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.

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.

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

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.

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.

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.

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.

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.

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.

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.

The swirling flow of sudden‐expansion dump combustor with central V‐gutter flameholder and six side‐inlets is studied by employing the SIMPLE‐C algorithm and Jones‐Launder k‐ε two‐equation turbulent model. Both combustion models of one‐step with infinite chemical reaction rate and two‐step with finite chemical reaction rate of eddy‐breakup (EBU) model are used to solve the present problem. The results agreed well with available prediction data in terms of axial‐velocity and total pressure coefficient along combustor centerline. The flowfield structure of combustor considered is strongly affected by swirling, flameholder and side‐inlet flow. For the fixed strength of swirling, the length of central recirculation zone is decreased when the angle of V‐gutter is increased. The outlet velocity of combustor in reacting flow is higher than that in cold flow because the released heat of combustion causes the decrease of density throughout the combustor flowfield. The distribution of mass fraction of various species in reacting process depends on the mixing effect, chemical kinetic and the geometric configuration of combustor.

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.

The purpose of this paper is to review the applications of the chemical reactor network (CRN) approach for modeling the combustion in gas turbine combustors and classify the CRN construction methods that have been frequently used by researchers.

This paper initiates with introducing the CRN approach as a practical tool for precisely predicting the species concentrations in the combustion process with lower computational costs. The structure of the CRN and its elements as the ideal reactors are reviewed in recent studies. Flow field modeling has been identified as the most important input for constructing the CRNs; thus, the flow field modeling methods have been extensively reviewed in previous studies. Network approach, component modeling approach and computational fluid dynamics (CFD), as the main flow field modeling methods, are investigated with a focus on the CRN applications. Then, the CRN construction approaches are reviewed and categorized based on extracting the flow field required data. Finally, the most used kinetics and CRN solvers are reviewed and reported in this paper.

It is concluded that the CRN approach can be a useful tool in the entire process of combustion chamber design. One-dimensional and quasi-dimensional methods of flow field modeling are used in the construction of the simple CRNs without detailed geometry data. This approach requires fewer requirements and is used in the initial combustor designing process. In recent years, using the CFD approach in the construction of CRNs has been increased. The flow field results of the CFD codes processed to create the homogeneous regions based on construction criteria. Over the past years, several practical algorithms have been proposed to automatically extract reactor networks from CFD results. These algorithms have been developed to identify homogeneous regions with a high resolution based on the splitting criteria.

This paper reviews the various flow modeling methods used in the construction of the CRNs, along with an overview of the studies carried out in this field. Also, the usual approaches for creating a CRN and the most significant achievements in this field are addressed in detail.

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.

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.

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.

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

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

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.

This paper aims to study the influences of hydrogen jet pressure on flow features of a strut-based injector in a scramjet combustor under-reacting cases are numerically investigated in this study.

The numerical analysis is carried out using Reynolds Averaged Navier Stokes (RANS) equations with the Shear Stress Transport k-ω turbulence model in contention to comprehend the flow physics during scramjet combustion. The three major parameters such as the shock wave pattern, wall pressures and static temperature across the combustor are validated with the reported experiments. The results comply with the range, indicating the adopted simulation method can be extended for other investigations as well. The supersonic flow characteristics are determined based on the flow properties, combustion efficiency and total pressure loss.

The results revealed that the augmentation of hydrogen jet pressure via variation in flame features increases the static pressure in the vicinity of the strut and destabilize the normal shock wave position. Indeed, the pressure of the mainstream flow drives the shock wave toward the upstream direction. The study perceived that once the hydrogen jet pressure is reached 4 bar, the incoming flow attains a subsonic state due to the movement of normal shock wave ahead of the strut. It is noticed that the increase in hydrogen jet pressure in the supersonic flow field improves the jet penetration rate in the lateral direction of the flow and also increases the total pressure loss as compared with the baseline injection pressure condition.

The outcome of this research provides the influence of fuel injection pressure variations in the supersonic combustion phenomenon of hypersonic vehicles.

This paper substantiates the effect of increasing hydrogen jet pressure in the reacting supersonic airstream on the performance of a scramjet combustor.