The aerothermodynamic cycle optimal design of a turbofan engine

To ensure that the aerothermodynamic cycle design of a turbofan engine is more accurate, efficient, and provide a reliable decision-making basis for engine designers, the multi-objective particle swarm optimization (MOPSO) method was used to optimize the aerothermodynamic performance parameters of the turbofan engine at multiple design points (MDPs). Fuel consumption rate and the specific thrust were considered as optimization targets. The thrust requirements and cycle parameter constraints under each working state were comprehensively considered to obtain the optimal performance boundary of the engine, the corresponding cycle parameters, and the correlations between different requirements and constraints. The results showed that the MOPSO algorithm could accurately and completely obtain the optimal performance boundary surface of the engine in the feasible region and the corresponding cycle parameter value. The feasible region obtained by the aerothermodynamic cycle design at MDPs was more accurate and effective than the design at a single design point.

A novel full-electric aircraft propulsion based on the DEA compressor

This article introduces a novel full-electric aircraft propulsion designed based on the DEA compressor, named DEAThruster that is supplied by a PEMFC. The thruster designed and modeled in this study is a compact DEAThruster to operate in altitudes up to 15,000 m and in the subsonic/transonic region with a specific thrust around 8300 N s/kg. The results show the thruster can satisfy all expectations, and it can generate up to 8300 N s/kg specific thrust for flight conditions encompassing static condition at sea level up to flight Mach number 0.95 in altitude 15,000 m. The DEAThruster can potentially be a practical alternative for gas turbine propulsions in all aspects when all available options for full-electric propulsions are not competitive for conventional aircraft propulsions in performance, size, and weight.

Design and Performance of a Compact Air-Breathing Jet Hybrid-Electric Engine Coupled With Solid Oxide Fuel Cells

A compact air-breathing jet hybrid-electric engine coupled with solid oxide fuel cells (SOFC) is proposed to develop the propulsion system with high power-weight ratios and specific thrust. The heat exchanger for preheating air is integrated with nozzles. Therefore, the exhaust in the nozzle expands during the heat exchange with compressed air. The nozzle inlet temperature is obviously improved. SOFCs can directly utilize the fuel of liquid natural gas after being heated. The performance parameters of the engine are acquired according to the built thermodynamic and mass models. The main conclusions are as follows. 1) The specific thrust of the engine is improved by 20.25% compared with that of the traditional jet engine. As pressure ratios rise, the specific thrust increases up to 1.7 kN/(kg·s−1). Meanwhile, the nozzle inlet temperature decreases. However, the temperature increases for the traditional combustion engine. 2) The power-weight ratio of the engine is superior to that of internal combustion engines and inferior to that of turbine engines when the power density of SOFC would be assumed to be that predicted for 2030. 3) The total pressure recovery coefficients of SOFCs, combustors, and preheaters have an obvious influence on the specific thrust of the engine, and the power-weight ratio of the engine is strongly affected by the power density of SOFCs.

Method for calculating the thrust and specific thrust of a double-flow gas turbine engine based on the results of tests at the laboratory unit «Flame»

The main technical characteristics of jet engines depend on the fuel quality: thrust and fuel consumption. As a rule, the comparative assessment of real engines is carried by specific values. Specific thrust is one of the most important parameters of the gas turbine engine (GTE). The larger it is, the smaller the required air flow rate through the engine at a given thrust and therefore its dimensions and mass. To date, a system for evaluating the performance properties of fuels based on qualification methods has been created. However, these methods do not allow calculating the thrust and specific thrust of the engine and potentially assessing the effect of fuels on these characteristics. Therefore, the issues of efficient use of fuels for GTE are solved almost exclusively on the basis of tests at testing units with full-scale engines, which are carried out repeatedly, which leads to a significant increase in the cost of testing. The article proposes a method for calculating the thrust and specific thrust of a double-flow gas turbine engine according to the results of tests at a constant volume laboratory unit of bypass type “Flame”. The method is based on modeling the engine operating conditions using the similarity criteria of the bench reactor and the real engine and allows reducing significantly the material and time costs for testing. The experimental of the combustion characteristics of hydrocarbon fuels and the rated values of their thrust and specific thrust for a double-flow gas turbine engine are presented.

Tte stage deorbiting with a deceleration pulse detonation engine

The article discusses the issues related to reducing the amount of space debris from rocket stages. The main ways to remove the separable part of a rocket from a space orbit are: the usе of a deceleration detonation propulsion system; gasification of fuel residues and the use of a gas-reactive deceleration pulse system; continuation of the work of the main propulsion system after the separation of stages; the use of a harpoon to capture the rocket stage and the use of sail for its further braking; the use of anti-missile or combat lasers to destroy a stage on the orbit followed by the stage fragments’ burning in the Earth’s atmosphere. To select the optimal method for removing from the orbit the separated part of a space rocket, the arithmetic progression method was applied. It has certain advantages over the classical hierarchy analysis method and has no inherent disadvantages of this method. A ranked row of solutions was obtained according to the five most significant performance criteria, and its stability was proved. A new deceleration detonation propulsion design scheme is proposed. Detonation burning of residual fuel components provides the maximum possible value of the deceleration thrust impulse. Using the example of the second stage of the “Zenit” launch vehicle, we analyzed the nature of the dependence of the entry angle into the atmosphere on the important characteristic parameters: the deceleration speed impulse, the entry speed into the Earth’s atmosphere of the separated launch vehicle stage, the required value of the specific thrust impulse of the deceleration propulsion system. A new analytical formula has been obtained, which connects the thrust and specific thrust impulse values of the detonation engine with the determined detonation process parameters. The results of the computational experiment were compared with the results of calculating the specific thrust impulse using the new formula for oxygen-based fuel compositions, known experimental data, and numerical simulation data of other authors. The data obtained in this study make it possible to evaluate the design parameters of the deceleration detonation engine at the stage of analyzing technical proposals.

The impact of thermodynamics parameters of turbofan engine with ITB on its performance

Presented paper is focused on the influence of additional combustor chamber named inter turbine burner on turbofan engine unit parameters. Investigation has been made how changing selected engine parameters affect its performance. A comparison has been made between the baseline turbofan engine and the engine with ITB. Engine thermodynamics model was prepared in MATLAB software. Main combustion chamber was fueled by kerosene, commmonly used in aviation transport, while inter turbine burner by alternative fuel. As an alternative fuel were choose liquid hydrogen and methane. Numerical researches were carried out for take-off conditions. Engine specific thrust and specific fuel consumption were obtained as a function of bypass ratio, turbine inlet temperature, fan pressure ratio, HPC and LPC pressure ratio. The results of the study indicate that hybrid engine with additional combustion chamber fueled by hydrogen fuel is more efficient than other studied cases.

Robust Aerodynamic Design of Nacelles for Future Civil Aero-Engines

As the growth of aviation continues it is necessary to minimise the impact on the environment, through reducing NOx emissions, fuel-burn and noise. In order to achieve these goals, the next generation of Ultra-High Bypass Ratio engines are expected to increase propulsive efficiency through operating at reduced specific thrust. Consequently, there is an expected increase in fan diameter and the associated potential penalties of nacelle drag and weight. In order to ensure that these penalties do not negate the benefits obtained from the new engine cycles, it is envisaged that future civil aero-engines will be mounted in compact nacelles. While nacelle design has traditionally been tackled by multi-objective optimisation at different flight conditions within the cruise segment, it is anticipated that compact configurations will present larger sensitivity to off-design conditions. Therefore, a design method that considers the different operating conditions that are met within the full flight envelope is required for the new nacelle design challenge. The method is employed to carry out multi-point multi-objective optimisation of axisymmetric aero-lines at different transonic and subsonic operating conditions. It considers mid-cruise conditions, end-of-cruise conditions, the sensitivity to changes in flight Mach number, windmilling conditions with a cruise engine-out case and an engine-out diversion scenario. Optimisation routines were conducted for a conventional nacelle and a future aero-engine architecture, upon which the aerodynamic trade-offs between the different flight conditions are discussed. Subsequently, the tool has been employed to identify the viable nacelle design space for future compact civil aero-engines for a range of nacelle lengths.

Theoretical Assessment of Turbocharged Pulse Detonation Engine Performance at Various Flight Conditions

A typical Pulse Detonation Engine (PDE) cycle of operation includes three basic processes: initiation and propagation of detonation wave in the Detonation Chamber (DC); a quasi-steady exhaust of detonation products from the DC at varying pressure through the supersonic nozzle; and a steady exhaust of remained detonation products at constant pressure through the nozzle while filling the DC with fresh air. In the present work, a novel method of Turbo-charging is proposed to increase the inlet pressure/density of fresh air fed into the DC in each cycle so as to increase the thrust developed per unit area of DC. The thermodynamic cycle of operation of Turbocharged Pulse Detonation Engine (TPDE) is analyzed based on quasi-steady state one dimensional formulation, and a computer code is developed in MATLAB to simulate the cycle performance at different compressor pressure ratios. Thrust per unit area of DC, the specific thrust and the fuel-based specific impulse are estimated at various flight conditions at different pressure ratios by considering C2H4/air as the fuel-oxidizer. The net thrust developed per unit area of DC increases with an increase in compressor pressure ratio, up to the pressure ratio of 4.0, at all flight conditions. The compressor pressure ratio of about 2.0 is observed to be optimum pressure ratio as TPDE develops nearly the same air-based specific thrust at this pressure ratio irrespective of flight operating conditions.

Exergo-economic cost accounting for PW4000 turbofan engine and its components

You The several series of PW4000 high bypass turbofan engine have used so far in many aircrafts. These commercial engines have played a crucial role on passenger and freight transportations. Namely, these engines are closely related to the environment impacts and security of energy supply. In this article, exergoeconomic analysis which is useful tool to investigate existing potential for improvement of the a system efficiency were carried out. The assesment, design and optimization of energy consuming systems are performed by means of these analyses. Therefore, thermo-economic costs were assigned to existing exergetic values of PW400 engine. Also exergo-economic performance parameters were evaluated. Finally, exergoeconomic deputy parameters were examined to understand relations with exergo-economic parameters. Based on the results of exergo-economics analysis, for Fan and exhaust, specific thrust costs are estimated 5.7051 $/hkN and 68.45$/hkN respectively. Also exergo-economics factor of PW4000 is found 7.958 % , while relative cost difference is determined at highest rate with 24.458 % for combustion chamber . With examination relations between economic variables and exergo-economic performance parameters, the change between 0.6 and 1.2 $/kg in the fuel price leads to increase the exhaust and fan specific thrust costs with 82.4701 $/hkN and 5.4332 $/hkN respectively. It is expected that conclusions of this study are helpful to notify exergo-economic impact of PW4000 engine Also, it may be benchmarking for similar gas turbine engines.