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Usage of gas turbine engines has increased by day due to rising demand for military and civil applications. This case results in investigating diverse topics related to energy efficiency and irreversibility of these systems. The purpose of this paper is to perform a detailed entropy assessment of turbojet engines for different flight conditions.

In this study, for small turbojet engines used in unmanned aerial vehicles, parametric cycle analysis is carried out at (sea level-zero Mach (hereinafter phase-I)) and (altitude of 9,000 m- Mach of 0.7 (hereinafter phase-II)). Based on this analysis, variation of performance and thermodynamic parameters with respect to change in isentropic efficiency of the compressor (CIE) and turbine (TIE) is examined at both phases. In this context, the examined ranges for CIE is between 0.78 and 0.88 whereas TIE is between 0.85 and 0.95.

Increasing isentropic efficiency decreases entropy production of the small turbojet engine. Moreover, the highest entropy production occurs in the combustor in the comparison of other components. Namely, it decreases from 2.81 to 2.69 kW/K at phase-I and decreases from 1.44 to 1.39 kW/K at phase-II owing to rising CIE.

It is thought that this study helps in understanding the relationship between entropy production and the efficiency of components. Namely, the approach used in the current analysis could help decision-makers or designers to determine the optimum value of design variables.

Due to rising isentropic efficiencies of both components, it is observed that specific fuel consumption (SFC) decreases whereas specific thrust (ST) increases. Also, the isentropic efficiency of a compressor affects relatively SFC and ST higher than that of the turbine.

Sustainable development calls for a larger share of intermittent renewable energy. To mitigate this intermittency, Compressed Air Energy Storage (CAES) technology was introduced. This technology can be made more sustainable by recovering the heat of the compression phase and reusing it during the discharge phase, resulting in an adiabatic CAES without the need for burning of fossil fuels. The key process parameters of CAES are temperature, pressure ratios, and the mass flow rates of air and thermal fluids. The variation in these parameters during the charge and discharge phases significantly influences the performance of CAES plants. In this chapter, the transient thermodynamic behavior of the system under various operating conditions is analyzed and the impact of heat recovery on the discharge phase energy efficiency, power generation, and CO₂ emissions is studied. Simulations are carried out over the air pressure range from 2,500 to 7,000 kPa for a 65 MW system over a five-hour discharge duration. It is also assumed that the heat loss in the air storage and the hot thermal fluid tank is insignificant and standby duration does not impact the status of the system. This result shows that the system exergy and the generated power are more sensitive to pressure change at higher pressures. This work also reveals that every 10°C increase on the temperature of the stored air can lead to a 0.83% improvement in the energy efficiency. The result of the transient thermodynamic model is used to estimate the reduction in CO₂ emissions in CAES systems. According to the obtained result, a 65 MW ACAES plant can reduce about 17,794 tons of CO₂ emission per year compared to a traditional CAES system with the same capacity.

To develop and find the effect of combination of four cycle design variables that minimizes the specific fuel consumption (sfc) of a turbofan engine.

After choosing the four variables, namely bypass ratio (B), fan pressure ratio, overall pressure ratio, and turbine inlet temperature (T₀₄), first the sfc was minimized without a minimum thrust constraint. Then, a minimum specific thrust constraint was introduced.

The unconstrained‐specific thrust is a two‐dimensional optimization problem, whereas the specific thrust constrained problem was found to be a three‐dimensional one.

The variables B and ï are limiting factors to further improvement, as set by their maximum practical values, whereas the other two variables are to be optimized.

A very useful work, in which numerical optimization program was developed, for a turbofan cycle and could be extended to other cycles.

This paper offers a great help to those intending to optimize certain cycles with a number of variables.

The purpose of this paper is to introduce new and modified “staged combustion” cycles in the form of engineering algorithm as a possible propulsion contender for future aerospace vehicle to achieve the highest possible “total impulse” to “mass” of propulsion system.

In this regard, the mathematical cycle model is formed to calculate the engine’s parameters. In addition, flow conditions (pressure, temperature, flow rate, etc). in the chamber, nozzle and turbopump are assessed based on the results of turbo machinery power balance and initial data such as thrust, propellant mixture ratio and specifications. The developed code has been written in the modern, object-oriented C++ programming language.

The results of the developed code are compared with the Russian RD180 engine which demonstrates the superiority and capability of new “thermodynamic diagrams”.

This algorithm is under constraint to control the critical variation of combustion pressure, turbine rpm, pump cavitation and turbine temperature. It is imperative to emphasize that this paper is limited to “oxidizer-rich staged combustion” engines with “single pre-burner”.

This study sheds light on using fuel booster turbopump and the second-stage fuel pump to moderate the effect of cavitation on pumps which reduces tank pressure and, as a consequence, decreases the propulsion system weight.

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.

The purpose of this paper is to investigate the performance of coke oven gas (COG)-combined cooling, heating and power (CCHP) system and to mainly focus on studying the influence of the environmental conditions, operating conditions and gas conditions on the performance of the system and on quantifying the distribution of useful energy loss and the saving potential of the integrated system changing with different parameters.

The working process of COG-CCHP was simulated through the establishment of system flow and thermal analysis mathematical model. Using exergy analysis method, the COG-CCHP system’s energy consumption status and the performance changing rules were analyzed.

The results showed that the combustion chamber has the largest exergy loss among the thermal equipments. Reducing the environmental temperature and pressure can improve the entire system’s reasonable degree of energy. Higher temperature and pressure improved the system’s perfection degree of energy use. Relatively high level of hydrogen and low content of water in COG and an optimal range of CH₄ volume fraction between 35 per cent and 46 per cent are required to ensure high exergy efficiency of this integration system.

This paper proposed a CCHP system with the utilization of coke oven gas (COG) and quantified the distribution of useful energy loss and the saving potential of the integrated system under different environmental, operating and gas conditions. The weak links of energy consumption within the system were analyzed, and the characteristics of COG in this way of using were illustrated. This study can provide certain guiding basis for further research and development of the CCHP system performance.

THE effect of higher turbine inlet temperature on the performance of an aircraft gas turbine engine can be quite dramatic. It can be used to increase the exhaust velocity of a dry turbojet to providea higher specific thrust; to increase the bypass ratio of a turbofan engine to improve its propulsive efficiency; to optimize the thermodynamic cycle at a higher pressure ratio to improve its specific fuel consumption; to reduce the amount of afterburner fuel flow in an augmented turbojet to improve its specific fuel consumption, or to increase the work output of a turboshaft engine. If the thrust or power of the engine is held constant, a size, cost and/or weight reduction can result. If the size of the engine is held constant growth capability can be provided.

An aircraft engine control system consists of a large scale of control parameters and variables because of the complex structure of aero-engine. Monitoring and adjusting control variables and parameters such as detecting, isolating and reconfiguring the system faults/failures depend on the controller design. Developing a robust controller is based on an accurate mathematical model.

In this study, a small-scale turboprop engine is modeled. Simulation is carried out on MATLAB/Simulink for design and off-design operating conditions. Both steady-state and transient conditions (from idle to maximum thrust levels) are tested. The performance parameters of compressor and turbine components are predicted via trained Neuro-Fuzzy model (ANFIS) based on component maps. Temperature, rotational speed, mass flow, pressure and other parameters are generated by using thermodynamic formulas and conservation laws. Considering these calculated values, error calculations are made and compared with the cycle data of the engine at the related simulation conditions.

Simulation results show that the designed engine model’s simulation values have acceptable accuracy for both design and off-design conditions from idle to maximum power operating envelope considering cycle data. The designed engine model can be adapted to other types of gas turbine engines.

Different from other literature studies, in this work, a small-scale turboprop engine is modeled. Furthermore, for performance prediction of compressor and turbine components, ANFIS structure is applied.

Interest is currently being expressed in heat exchanged propulsion gas turbines for a variety of aeroengine applications, and in support of this, the aim of this paper is to evaluate the relevance of experience gained from development testing of several recuperated aeroengines in the USA in the late 1960s.

Technology status, including engine design features, performance, and specific weight of recuperated propulsion gas turbines based on radial and axial turbomachinery, that were development tested in the power range of about 300 to 4,000 hp (224 to 2,984 kW) is discussed in Part I.

A successful flight worthiness test was undertaken in the USA of a helicopter powered solely by a recuperated turboshaft engine and this demonstrated a specific fuel consumption reduction of over 25 percent compared with the simple‐cycle engine. However; in an era of low‐fuel cost, and uncertainty about the long‐term structural integrity of the high‐temperature heat exchanger, further development work was not undertaken.

The gas turbines tested were based on conventional simple‐cycle engines with essentially “bolted‐on” recuperators. Optimum approaches to minimize engine overall weight were needed in which the recuperator was integrated with the engine structure from the onset of design, and these are discussed in Part II.

Based on engine hardware testing, a formidable technology base was established, which although dated, could provide insight into the technical issues likely to be associated with the introduction of future heat exchanged aeroengines. These are projected to have the potential for reduced fuel burn, less emissions, and lower noise, and recuperated and intercooled turboshaft, turboprop, and turbofan variants are discussed in Part III.

The requirements imposed upon advanced short take‐off and vertical landing (ASTOVL) aircraft give rise to challenging demands on their propulsion systems. One possible approach is to have a high‐performance turbofan of traditional design and an additional, but separate, fan to provide a major part of the lift during the take‐off and landing manoeuvres. For such a design, there are several quite‐different choices of layout for providing the power to drive the remote fan by means of the core engine. These include shaft‐driven and bleed‐driven options. The choice will depend on the anatomy and required thermodynamic‐performance of the whole system. In this paper, several pertinent alternative engine‐designs are discussed. Four of these, based on a high‐performance low‐bypass‐ratio core engine, are studied in detail and their behaviours compared. Prima facie, the preferred choice is the engine with the shaft‐driven fan. A slightly less acceptable choice is the high‐pressure turbine exit‐bleed driven remote‐fan.