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This study seeks the effect on static thrust, thrust specific energy consumption (TSEC) and exhaust emissions of euro diesel-hydrogen dual-fuel combustion in a small turbojet engine.

Experimental studies are performed in a JetCat P80-SE type small turbojet engine. Euro diesel and hydrogen is fed through two different inlets in a common rail distributing fuel to the nozzles. Euro diesel fuel is fed by a liquid fuel pump to the engine, while hydrogen is fed by a fuel-line with a pressure of 5 bars from a gas cylinder with a pressure of approximately 200 bars.

At different engine speeds, it is found that there is a decrease at the TSEC between a range of 1% and 4.8% by different hydrogen energy fractions (HEF).

The amount of hydrogen is adjusted corresponding to a range of 0–20% of the total heat energy of the euro diesel and hydrogen fuels. The small turbojet engine is operated between a range of 35,000 and 95,000 rpm engine speeds.

On the other hand, remarkable improvements in exhaust emissions (i.e. CO, CO₂, HC and NO_x) are observed with HEFs.

This is through providing improvements in performance and exhaust emissions using hydrogen as an alternative to conventional jet fuel in gas turbine engines.

Reducing fuel consumption of unmanned aerial vehicles (UAVs) during transient operation is a cornerstone to achieve environment-friendly operations. The purpose of this paper is to develop a control scheme that improves the fuel economy of a turbojet in its full operating envelope.

A novel direct-thrust linear quadratic integral (LQI) approach, comprised by an optimal observer/controller satisfying specified performance parameters, is presented. The thrust estimator, based in a Wiener model, is validated with the experimental data of a micro-turbojet. Model uncertainty is characterized by analyzing variations between the identified model and measured data. The resulting uncertainty range is used to verify closed-loop stability with the circle criterion. The proposed controller provides stable responses with the specified performance in the whole operating range, even with after considering plant nonlinearities. Finally, the direct-thrust LQI is compared with a standard thrust controller to assess fuel economy and performance.

The direct-thrust LQI approach reduced the fuel consumption by 2.1090% in the most realistic scenario. The controllers were also evaluated using the environmental effect parameter (EEP) and transient-thrust-specific fuel consumption (T-TSFC). These novel metrics are proposed to evaluate the environmental impact during transient-thrust operations. The direct-thrust LQI approach has a more efficient fuel consumption according to these metrics. The results also show that isolating the thrust dynamics within the feedback loop has an important impact in fuel economy. Controllers were also evaluated using the EEP and T-TSFC. These novel metrics are proposed to evaluate the environmental impact during transient-thrust operations. The direct-thrust LQI approach has a more efficient fuel consumption according to these metrics. The results also show that isolating the thrust dynamics within the feedback loop has an important impact in fuel economy.

This study shows the design of an effective direct-thrust control approach that minimizes fuel consumption, ensures stable responses for the full operation range, allows isolating the thrust dynamics when designing the controller and is compatible with classical robustness and performance metrics. Finally, the study shows that a simple controller can reduce the fuel consumption of the turbojet during transient operation in scenarios that approximate realistic operating conditions.

THE effect of the propulsive efficiency is analysed for all speed regions and methods for obtaining optimum propulsive efficiency for any speed and environmental conditions are investigated for different type vehicles. Also, the importance of the propulsive efficiency is compared to other factors such as weight of power plant, specific fuel consumption, specific power ratio, specific thrust ratio, etc. Finally, on the basis of considering all power plant factors, it is shown how to achieve optimum propulsion for any vehicle at required operating conditions.

In this paper, with the goal of reducing the fuel consumption of UAV, the engine performance optimization is studied and on the basis of aircraft/engine integrated control, the minimum fuel consumption optimization method of engine given thrust is proposed. In the case of keeping the given thrust of the engine unchanged, the main fuel flow of the engine without being connected to the afterburner is optimally controlled so as to minimize the fuel consumption.

In this study, the reference model real-time optimization control method is adopted. The engine reference model uses a nonlinear real-time mathematical model of a certain engine component method. The quasi-Newton method is adopted in the optimization algorithm. According to the optimization variable nozzle area, the turbine drop-pressure ratio corresponding to the optimized nozzle area is calculated, which is superimposed with the difference of the drop-pressure ratio of the conventional control plan and output to the conventional nozzle controller of the engine. The nozzle area is controlled by the conventional nozzle controller.

The engine real-time minimum fuel consumption optimization control method studied in this study can significantly reduce the engine fuel consumption rate under a given thrust. At the work point, this is a low-altitude large Mach work point, which is relatively close to the edge of the flight envelope. Before turning on the optimization controller, the fuel consumption is 0.8124 kg/s. After turning on the optimization controller, you can see that the fuel supply has decreased by about 4%. At this time, the speed of the high-pressure rotor is about 94% and the temperature after the turbine can remain stable all the time.

The optimal control method of minimum fuel consumption for the given thrust of UAV is proposed in this paper and the optimal control is carried out for the nozzle area of the engine. At the same time, a method is proposed to indirectly control the nozzle area by changing the turbine pressure ratio. The relevant UAV and its power plant designers and developers may consider the results of this study to reach a feasible solution to reduce the fuel consumption of UAV.

Fuel consumption optimization can save fuel consumption during aircraft cruising, increase the economy of commercial aircraft and improve the combat radius of military aircraft. With the increasingly wide application of UAVs in military and civilian fields, the demand for energy-saving and emission reduction will promote the UAV industry to improve the awareness of environmental protection and reduce the cost of UAV use and operation.

THE rocket motor is a form of jet propulsion which is characterized by independence of the external atmosphere for combustion, relative independence of altitude and flight velocity upon thrust, small frontal area for high thrusts, simple construction and low weight, and high rate of fuel consumption. Its use was greatly developed during the war years and many applications are now familiar to all. Most of the work on rocket missiles, such as the anti‐aircraft barrages, fighter armament, etc., was performed with solid fuel rockets, but liquid fuels were developed by the Germans for the well‐known V.2, for the Me. 163 aircraft, the Henschel glide bomb and various other applications. They concentrated a great deal of effort on this work and considerable technical progress had been made with different systems. Three main systems emerged and these were distinguished by the oxygen bearing fluids they used. The fluids were:

The purpose of this paper is to verify the feasibility of lunar capture braking through three methods based on particle swarm optimization (PSO) and compare the advantages and disadvantages of the three strategies by analyzing the results of the simulation.

The paper proposes three methods to verify capture braking based on PSO. The constraints of the method are the final lunar orbit eccentricity and the height of the final orbit around the Moon. At the same time, fuel consumption is used as a performance indicator. Then, the PSO algorithm is used to optimize the track of the capture process and simulate the entire capture braking process.

The three proposed braking strategies under the framework of PSO algorithm are very effective for solving the problem of lunar capture braking. The simulation results show that the orbit in the opposite direction of the trajectory has the most serious attenuation at perilune, and it should consume the least amount of fuel in theoretical analysis. The methods based on the fixed thrust direction braking and thrust uniform rotation braking can better ensure the final perilune control accuracy and fuel consumption. As for practice, the fixed thrust direction braking method is better realized among the three strategies.

The process of lunar capture is a complicated process, involving effective coordination between multiple subsystems. In this article, the main focus is on the correctness of the algorithm, and a simplified dynamic model is adopted. At the same time, because the capture time is short, the lunar curvature can be omitted. Furthermore, to better compare the pros and cons of different braking modes, some influence factors and perturbative forces are not considered, such as the Earth’s flatness, light pressure and system noise and errors.

This paper presents three braking strategies that can satisfy all the constraints well and optimize the fuel consumption to make the lunar capture more effective. The results of comparative analysis demonstrate that the three strategies have their own superiority, and the fixed thrust direction braking is beneficial to engineering realization and has certain engineering practicability, which can also provide reference for lunar exploration orbit design.

The proposed capture braking strategies based on PSO enable effective capture of the lunar module. During the lunar exploration, the capture braking phase determines whether the mission will be successful or not, and it is essential to control fuel consumption on the premise of accuracy. The three methods in this paper can be used to provide a study reference for the optimization of lunar capture braking.

THE engines used for the jet propulsion of aircraft have been described in detail on many occasions in the last year or so. Most of these descriptions, however, have been purely factual, and little has been written about the theoretical aspects. In view of the unconventionality of these power units, it is hoped that the following notes, in which a simple analysis of the theory is given, may be of interest.

The aim of the paper is to establish the COst-Specific Air Range (COSAR) as a new figure-of-merit based on the cost of energy to optimise the flight profile of a hybrid energy aircraft.

After reviewing the expression and the application of the specific air range (SAR) and of the energy-specific air range (ESAR), the need of a new figure-of-merit for flight technique optimisation of hybrid energy aircraft is motivated. Based on the specific cost of the energies consumed, the mathematical expression of COSAR is derived. To enable optimum economics operations, a cost index (CI) derivation is introduced for a variety of hybrid-electric concepts to consider the additional time-related cost. The application of COSAR and of the CI is demonstrated for cruise optimisation of a hybrid-electric retrofit aircraft concept.

As a consequence of the consumption of multiple energy sources in a hybrid aircraft, optimisation according to the objective functions SAR and ESAR leads to minimum in-flight CO2 emissions and minimum energy consumption for a given stage length. While the optimisation of a single energy source aircraft according to these figures-of-merit directly results in minimum energy cost for a given unit range, this statement is no longer true for hybrid-energy aircraft. Consequently, introducing a new figure-of-merit established on the specific cost of the energies consumed enables flight technique optimisation for minimum energy cost of hybrid-energy aircraft. Additionally, the related time-cost is taken into account by means of a CI definition for minimum operating cost.

COSAR may serve as an alternative to SAR used today as the standard figure-of-merit for fuel optimised flight profile. Using COSAR and the CI allow airlines to adapt the flight profiles of hybrid-energy aircraft fleets according to the energy market price and their related cost of time to determine optimum economical flight profile.

Using COSAR as a figure-of-merit, the flight profile of hybrid energy aircraft can be optimised for minimum energy cost. Time-related costs are considered for optimum operating economics by utilisation of the CI definition for hybrid energy aircraft.

The increasing cost of hydrocarbon fuel inevitably intensifies the quest for lower engine SFC on the conventional aero gas turbine. This should not obscure the fact that minimising all aircraft direct operating costs is the ultimate yardstick for the engine designer. Higher fuel prices may change priorities in engine design and justify more complex and expensive engines particularly for longer range operations. Optimum engine design for shorter range can be significantly different because of the implication of cyclic life on air cooled turbine blades. Lower specific thrust engines are worthy of close consideration particularly for short haul operation. More effort should be devoted to comprehensive studies of new powerplant concepts.

The purpose of this paper is to examine the effects of heat capacity and density of biofuels on aircraft engine performance indicated by thrust and fuel consumption.

The influence of heat capacity and density was examined by simulating biofuels in a two-spool high-bypass turbofan engine running at cruise condition using a Cranfield in-house engine performance computer tool (PYTHIA). The effect of heat capacity and density on engine performance was evaluated through a comparison between kerosene and biofuels. Two types of biofuels were considered: Jatropha Bio-synthetic Paraffinic Kerosene (JSPK) and Camelina Bio-synthetic Paraffinic Kerosene (CSPK).

Results show an increase in engine thrust and a reduction in fuel consumption as the percentage of biofuel in the kerosene/biofuel mixture increases. Besides a low heating value, an effect of heat capacity on increasing engine thrust and an effect of density on reducing engine fuel consumption are observed.

The utilisation of biofuel in aircraft engines may result in reducing over-dependency on crude oil.

This paper observes secondary factors (heat capacity and density) that may influence aircraft engine performance which should be taken into consideration when selecting new fuel for new engine designs.