Exergy analysis is a method of evaluating the proportion of each process on the transmission of system availability and determining how the useful energy is loosed. In this research، a turboprop engine is simulated using GASTURB 10 commercial software which is developed based on zero-dimensional approach. The performance of turboprop engine is calculated at design point and some off-design points and thermodynamic properties of engine components at different flight altitudes and Mach numbers are calculated. After that، by defining exergy terms and using related exergy equilibrium equations، the necessary conceptual basis for exergy analysis is introduced. Then thermodynamic data obtained from GASTURB software are used to exergy analysis of engine components. Results show that combustion process has the dominant portion of the system irreversibility.

In this research, an accurate modeling of the thermodynamic behavior of a gas turbine for a trainer jet engine has been developed, considering volume dynamics, shaft dynamics, and the effects of altitude and Mach number changes. In order to maintain the engine within the acceptable operational range, a two-level fuzzy-PID controller has been designed to control the turbojet engine in a software environment aimed at examining the effectiveness of the designed method. This controller effectively controls all nonlinear thermodynamic behaviors and Mach/altitude variations. Additionally, a protective loop has been implemented to safeguard against engine shutdown, extreme temperature increase, and surge using a Min-Max strategy with a paired controller that communicates the controller response to the engine and prevents engine damage. This model has the capability to simulate engine performance in both transient and steady-state conditions. After validating the model using Gasturb 13 software, with a maximum error of 8. 76%, simulation results indicate the capability of the hybrid two-level controller in different operation conditions, resulting in an average 26% shorter settling time, 21. 5% shorter rise time, and no permanent error compared to PID control.

In this study, the conceptual design and optimization of an air gas turbine combustor were investigated to reduce the liner temperature and simultaneously decrease the overall weight, including the casing and liner. The combustor under study belongs to the CFM56 engine, with initial data obtained from GasTurb software and similar studies. The NSGA-II multi-objective optimization algorithm and MATLAB software were used for optimization. At the beginning of the study, design equations and fundamental principles were examined. Heat transfer within the combustor was considered one-dimensional and without additional cooling to optimize the worst-case scenario. The optimization results show a 16% reduction in thermal stresses on the liner and a 10% decrease in combustor weight. This study also highlights the importance of selecting proper performance constraints, as an inappropriate constraint range can lead to a non-physical or limited Pareto front. The proposed optimization process can accelerate the design of gas turbine combustors, reducing excessive iterations needed to resolve design conflicts that are often difficult to assess during conventional design phases.

In this paper, the thermodynamic modeling and analysis of various components of a turboprop engine and their relation in the steady-state design and off-design performances are addressed. An algorithm is presented and a computer code is developed for the analysis of the cycle in design and off-design conditions. Then off-design performance curves are plotted for different turbine inlet temperatures, flight elevations and Mach numbers and their results are interpreted. The results are compared and validated with the results of GasTurb software. In addition, the effects of the combustion chamber defects on the performance of the engine are investigated. The results of this paper can be used as the first step in the study of the performance and upgrading of these engines. The results of this study show that the specific fuel consumption will rise by 14% for a 10% increase in the combustion chamber pressure drop. Furthermore, the specific fuel consumption increases by 13% and the engine power reduces by 17. 5% for a 10% reduction in the combustion efficiency

Thrust and specific fuel consumption are two essential parameters in a gas turbine engine. The fundamental design of various engine components is required to enhance thrust and provide optimum specific fuel consumption. This paper studies the effects of the nozzle exit area on the operating lines of fan, compressor, and turbine in a single-spool, mixed-flow, and low bypass ratio turbofan engine. The zero-dimensional modeling of the engine is examined in GasTurb, and the nozzle area is changed in 6 cases by ±5, ±10, and ±15 percent relative to design point. The inlet temperature of turbine rotor and the spool speed of the engine are limited to their maximum amount. The results show that any change in the nozzle exit area will reduce the net thrust of the engine. On the other hand, with the reduction of the nozzle area, the operating line of the fan and compressor becomes closer to the surge line, reducing the surge margin, while with the increase of the nozzle area, the surge margin of the fan and compressor increases. Also, the changes in specific fuel consumption are inversely proportional to the changes in the nozzle outlet area.

Modeling the behavior of a gas turbine system and designing its control has always been of interest to researchers in this field. Proper modeling allows the implementation of a suitable controller on a dynamic system, and using suitable control, the system can be controlled in the best and safest way possible. In the design of gas turbine control, due to the existence of irreparable risks to the system, protective constraints must be observed. These risks include surge, turbine overheating, and flame extinction. In this study, a non-linear model of the J85 engine was initially built, and its results were validated with the Gasturb software. The validation results show that the maximum error for this engine in the compressor pressure ratio and turbine inlet temperature in transient conditions is 5 and 5. 8 percent, respectively. Then, using the Narma L-2 neural network and the Min-Max protection constraints, a controller was designed for this system. The designed controller is able to maintain the surge limit at the idle to maximum speed maneuver above 5% and prevent turbine over temprature and flame-out, and its constant error value at the design point is zero.

The main objective of this study is to analyze the performance parameters and simulation of PW615F turbofan engines both with and without secondary combustion chamber. For validation a commercial software called GasTurb & GSP is used. The purpose of the the propulsion system plays an important role in determining of mission and flight regime of an aircraft In this study, the first choice and then simulate a turbine engine thrust which is equivalent to 6. 35 KN produce. And more the method for analyzing on-design and off-design analysis of twin-spool unmixed flow turbofan engines and also we present scaling method to produce the characteristics performance maps of compressor and fans. Mathematical modeling of engine performance has been conducted by serial nested loops algorithm. Performance of components have been modeled by performance diagrams, thermodynamics and gas dynamic equations. Mathematical modeling of a turbofan engine is somewhat complex. Therefore, initially, the off-design performance of a twin spool power generating gas generator and twin spool turbofan are analyzed. Finally, the off-design performance of a twin spool unmixed flow turbofan engine is simulated and the generated performance diagrams have been analyzed. The results show that the cycles that Advantage the secondary combustion chamber dramatically increase the specific thrust, reduce the main secondary combustion chamber outlet emissions and increase the output power of the low pressure turbine.

In recent years, many turbofan engines that lack flight certification have been repurposed for other applications, such as power generation or providing air mass flow. This article investigates the feasibility and characteristics of using a standard turbofan combustion chamber in a ramjet application. One can of the combustion chamber from the D30K engine, which is of a can-annular type, was selected as the basis for this study. A redesign process was undertaken to determine the geometry of the combustion chamber based on existing scientific and technical literature. The geometric ratios of the D30K engine were used for validation, demonstrating satisfactory adaptation.The Gas Turb software was employed to determine the flow entrance conditions for a combustion chamber applied in a ramjet engine. Subsequently, the Fluent software was used to perform simulations of combustion using a non-premixed liquid phase method. The results indicated that, in this configuration, the induced vortices were not strong enough to stabilize the flame, making using a flame holder essential.Additionally, a parametric study was conducted to investigate the effects of the flame holder's size, number, position, and distance on parameters such as pressure loss, maximum temperature, Mach number, and flame stability. The results demonstrate that a multi-flame holder setup positively impacts performance, while more oversized flame holders are not recommended due to undesirable effects.