Electric Jet Engine Research Articles

Design & Development of Hybrid Electric Jet Engine

Hybrid electric jet engines without turbines, in which the compressors are powered by fuel cells rather than the turbines, are a potential and evolving technology for unmanned aerial vehicles. The hybrid electric jet engine combines the benefits of fuel cells and turbo-electric engines. The hybrid engine produces high specific thrust and excellent thermal efficiency.

Digital model of the drag force of the Earth’s upper atmosphere for the design of low-orbit spacecraft

A digital model is described for estimating the drag force of the Earth's upper atmosphere acting on a low-orbit spacecraft. Unlike the classical model of aerodynamic force, the digital model is the result of a computer algorithm. This algorithm calculates the field of aerodynamic force values, taking into account changes in the flight altitude, aerodynamic coefficient and the area of the spacecraft midsection over time. The calculations also use a digital model of the dynamic density of the Earth's upper atmosphere. The digital model of the drag force of the Earth's upper atmosphere is useful in the design of low-orbit spacecraft with a low-thrust jet propulsion system (thrust order – millinewtons). On the basis of the calculated values of the aerodynamic force, it is possible to form requirements for the geometry of the spacecraft, attitude modes and arrangement of low-thrust jet engines (for example, electric jet engines). With the accumulation of big data in the form of tabular functions of aerodynamic force from the space of various design and ballistic parameters of low-orbit spacecraft, it is possible to create a machine model (ML-model). It will make forecasts of the values of the design characteristics for the low-orbit spacecraft being developed according to a limited set of initial requirements.

Electric Analysis of the Maritime Application High-Frequency Magnetohydrodynamic Thruster

A magnetohydrodynamic (MHD) thruster is the next-generation electric jet engine for maritime applications. It eliminates the moving mechanical components that make the noises and reduces physical harm to sea creatures. This paper finds that aluminum electrodes have higher conductivity and less capacitive value in a KCl solution than the 316 stainless steel and zinc in MHD applications. Further, the AC operation can mitigate the power loss during electrolysis and power loss while on the water. The new optimal coil design with the enclosed-type ferrite layout of the MHD thruster is addressed by this simulation study. The AC operation and electric drive with a Lorentz force analysis will be demonstrated. Lastly, a verification experiment that pushes the KCl solution at 3 cm/s will be interpreted by the prototype to display the electric operation detail.

Numerical calculation of the operating parameters of the ION engine for satellite correction

The article provides a description of methods for numerical simulation of physical processes occurring in an electric jet engine. A method for estimating the ionization of the working fluid based on drift-diffusion equations is presented. The possibility of calculating the neutralization coefficient to prevent erosion of an engine with an original design based on a system of nested electrodes is shown. The article describes the sequence of creating a computer model of an ion engine and the main methods of numerical analysis of the dynamics of charged particles.

АНАЛИЗ СОВРЕМЕННОГО СОСТОЯНИЯ И ПЕРСПЕКТИВЫ РАЗВИТИЯ РАКЕТНЫХ ДВИГАТЕЛЕЙ ДЛЯ ИССЛЕДОВАНИЯ ДАЛЬНЕГО КОСМОСА

The article analyzes the current state and development prospects of rocket engines for space exploration. Currently, for flights to other planets, not to mention the stars, the use of liquid-propellant and solid-propellant rocket engines is becoming increasingly unprofitable, although many rocket engines have been developed. Thus, to reach manned planes even the nearest planets, it is necessary to develop rocket launchers on engines operating on principles different from chemical propulsion systems. The most promising in this regard are electric, laser and nuclear rocket engines, as well as hybrid rocket engines such as solid-state nuclear rocket engine (NRE) and electric jet engine (ERE) or gas-phase NRE and ERE.

Radial Magnetic Bearings for Rotor–Shaft Support in Electric Jet Engine

New technologies are being developed to elaborate cutting-edge electrical jet engines to replace classical constructions. These new concepts consider the possibility of using electrical machines both as starters and generators, as well as suspension systems for the turbine shafts of aircraft engines. The paper will present mathematical analysis regarding active magnetic bearing (AMB) implementation for rotor–shaft support. This technology allows the elimination of friction forces between cooperating kinematic pairs (stator and rotor), reduces the adverse effects of classic bearings, and increases operating speed range and an operational susceptibility. The mathematical and numerical analysis of active magnetic suspension systems are presented. Next, a comparison of the theoretical studies using Comsol Multiphysics software and its experimental verification are described. A discussion regarding the mathematical analysis and experimental effects is also provided. The conclusion summarizes the theoretical and experimental features of heteropolar radial active magnetic bearings in new electric aircraft engines.

Combining Macro- and Mesoscale Optimization: A Case Study of the General Electric Jet Engine Bracket

Topology optimization (TO) is a mathematical method that optimizes the material layout in a pre-defined design domain. Its theoretical background is widely known for macro-, meso-, and microscale levels of a structure. The macroscale TO is now available in the majority of commercial TO software, while only a few software packages offer a mesoscale TO with the design and optimization of lattice structures. However, they still lack a practical simultaneous macro–mesoscale TO. It is not clear to the designers how they can combine and apply TO at different levels. In this paper, a two-scale TO is conducted using the homogenization theory at both the macro- and mesoscale structural levels. In this way, the benefits of the existence and optimization of mesoscale structures were researched. For this reason, as a case study, a commercial example of the known jet engine bracket from General Electric (GE bracket) was used. Different optimization workflows were implemented in order to develop alternative design concepts of the same mass. The design concepts were compared with respect to their weight, strength, and simulation time for the given load cases. In addition, the lightest design concept among them was identified.

Magnetic Suspension Technology for Electric Jet Engines

Reducing the emission of harmful compounds such as carbon dioxide and nitrogen oxides has been identified as a priority target in the European Union. Aviation is one of the main sources of pollution. The reduction of pollutant emissions can be achieved by the use of the electric jet engine. This type of a jet engine differs significantly from a kerosene-powered engine. The article presents the concept of an electric jet engine with the rotor that is magnetically suspended. Demonstrators of active and passive magnetic bearing technologies and bearingless electric motors, developed at the Avionics Department, are presented in the paper.

Математическая модель маневрирующего наноспутника с гелиотермической двигательной установкой и маховичной системой ориентации и стабилизации

The paper considers a possibility of using a heliothermal propulsion system for constructing constellations of nanosatellites in the CubeSat format. Such a propulsion system implies direct heating of the working fluid by focused solar radiation. In comparison with electric jet engines, the proposed propulsion has orders of magnitude higher thrust, which makes it possible to build a constellation of nanosatellites in low Earth orbit in less than 6 days. In comparison with electrothermal motors, the presented solution achieves higher efficiency of converting the energy of solar radiation into the thermal energy of the working fluid. The spacecraft motion was simulated taking into account the mutual influence of the propulsion system, the orientation and stabilization system, the power supply system, as well as the passage of the shadow sections of the orbit in which there is a loss of electrical and thermal energy.

Анализ эффективности использования электрореактивных двигателей для поддержания низкой орбиты малого космического аппарата

This paper examines the problem of maintaining the plane parameters of the working orbit of a small spacecraft using an electric propulsion engine. In low working orbits, due to the Earth’s atmosphere, a spacecraft is subjected to aerodynamic drag forces, which results in a decrease in the radius of the orbit and a potential termination of the useful target functioning. The time parameters of the cyclogram for maintaining the working orbit of a small spacecraft with an electric low thrust engine are analyzed taking into account the variability of the atmospheric density. The cyclogram consists of sections of the passive and active movement under the action of the low thrust engine. For the satellite under study, suitable thrust parameters of the electric engine are selected, which allow the correction of the plane parameters of the low orbit. Using the characteristics of the thrust and specific impulse of the electric jet engine, fuel reserves for correction over a long period of time are calculated. The results of the analysis confirm the effectiveness of the electric propulsion engine in terms of fuel consumption for correction.

Determination of the safe operation zone for a turbine-less and solid oxide fuel cell hybrid electric jet engine on unmanned aerial vehicles

Solid oxide fuel cell (SOFC) hybrid electric jet engines without turbines in which compressors are powered by fuel cells instead of the turbines are promising and developmental aero-engines for unmanned aerial vehicles. The hybrid electric jet engine combines the merits of turbo-electric engines and fuel cell engines. To avoid component malfunctions and engine performance deterioration, the determination of the safe operation zone is necessary. In this study, a performance analysis model of the hybrid electric jet engine is built and the main conclusions are as follows. The safe operating zone of the hybrid engine is not restricted by turbine inlet temperature. Under low fuel flow rate and low air flow rate, zones of too low reforming temperature or too low SOFC open voltage exist. An unbalanced energy zone exists under high fuel flow rate and low air flow rate. The power produced by the SOFC is above that consumed by the compressor. The lower boundary of the safe zone is determined by the choke zone of the compressors. In the safe zone, the hybrid engine has properties of high specific thrust (837.6 N/(kg.s−1)) and high thermal efficiency (70.4%) at high fuel low and air flow rate.

Analysis of the cyclogram of maintaining a low working orbit of a space-craft of the AIST-2 class using an electric jet engine

A study was made of the time parameters of the cyclogram for maintaining the low working orbit of a small spacecraft of the AIST-2 class using an electric jet engine. The analysis is made for working orbits with a height in the range from four hundred to five hundred kilometers with a changing upper atmosphere of the Earth, depending on the level of solar activity. The calculations used the thrust of an electric jet engine equal to twenty millinewtons with a service life of not more than a thousand hours. The methodological and software were used: for calculating the level of aerodynamic drag depending on the level of solar activity; for modeling and analysis of the parameters of the orbital motion of the spacecraft under the action of corrective and aerodynamic forces. The results of the analysis showed that the electric jet engine allows maintaining the working orbit in the range of designated heights. If the average altitude of the orbit deviates by no more than three kilometers, correction can be carried out in less than a day. The time of one correction cycle can vary from four to four hundred and seventy-eight days, depending on the level of solar activity and the design and ballistic parameters of the spacecraft. The operating life of an electric jet engine equal to one thousand hours can maintain the working orbit of the spacecraft for more than seven years with low solar activity in the range of the studied heights.

Features of using plasma flows to compensate the aerodynamic drag of mini-satellites moving in the ionosphere

The possibilities of using plasma flows formed in electric jet engines (EJE) are considered. One of the variants of the EJE construction are ion sources with multi-cell ion-optical systems used to compensate the aerodynamic drag of mini-satellites in the Earth’s ionosphere with the aim of ensuring long-term operation of the latter at low flight altitudes (∼180–220 km). The analysis of methods of simulation of the operation of such engines together with air intakes under ground conditions is given. The possibility of using plasma-ion sources producing molecular beams of nitrogen or air ions with low (orbital) flight speeds as a source of the free flow in the simulations is considered. A source with volumetric ionization of the working gas and electron oscillation in the magnetic field in a chamber between the walls held at the cathode potential and a large multi-cell aperture of the ion–optical system is used to create the flow. The results of the studies of the operation of a plasma–ion engine (PIE) together with air intake obtained by the gas-dynamic simulation method are presented. The possibility of a high performance of the engine under non-standard conditions is shown. It is found that a stable operation of the PIE can be realized at pressures of 6⋅10−5–10−4 Torr in its chamber, which is important for matching the engine with the air intake.

On a theorem of impulse and energy to determine parameters of the electric jet engine KA

The report presents the modern physical principles of acceleration of charged particles and plasma in electric and magnetic fields, which are the basis for the classification of electric rocket engines. It is shown that for reactive means of correction of orbit parameters the lifetime depends to a greater extent on the reserves of the working fluid and the corresponding thrust and energy for the propulsion system. It is postulated that the special value of the determination of pulses and energy for electric propulsion (ERD) is that their application makes it possible to obtain an idea of these phenomena on the boundary surface of a certain area, without knowledge of the individual phenomena occurring within the ERD. That is, often in cases where the differential equations of plasma flow can not be composed or at least can not be integrated, the theory of pulses or energy makes it possible to know the nature of the movement in its General features. It is convincingly proved that the theorems of impulses and energy are a convenient means for checking the correctness of the results, including experimental ones.

АНАЛИЗ ЦЕЛЕСООБРАЗНОСТИ ИСПОЛЬЗОВАНИЯ АЭРОДИНАМИЧЕСКОГО КОМПЕНСАТОРА ПРИ БЕСКОНТАКТНОМ УДАЛЕНИИ КОСМИЧЕСКОГО МУСОРА

The article deals with the feasibility of aerodynamic compensator application in the scheme of the so-called noncontact method of space debris removal on the low near-earth orbits without mechanical contact of the spacecraft and space debris. Previously, in relation to one of the technologies for implementing such method of collection, referred to as "the shepherd with an ion beam", its modification was proposed, which consists in replacing the compensating electro jet engine with an aerodynamic compensator. The original compensating engine serves to compensate the reactive power of the main electric jet engine, which torch ion flow has a “braking” effect on the space debris. The idea of modifying the technology was to save the cost of an expensive working jet engine and reduce the starting mass of the shepherd’s spacecraft. However, this assumption was not sufficiently substantiated. The analysis of the feasibility of aerodynamic compensator application is made on the basis of calculating the mass of the working medium and the amount of electrical energy saved due to its application. The above mentioned calculation of the saved mass of the working fluid is made applying a number of simplifying assumptions. As a result of the analysis, it was determined that the expediency of aerodynamic compensator applying is not obvious, considering the mass of the compensator and the complexity of the design of the spacecraft. At the same time, if in one mission of the shepherd’s spacecraft will presuppose to carry out several space debris removal on the orbit, then the application of an aerodynamic compensator become justified. The direction of further research involves a more detailed analysis of the feasibility of aerodynamic compensator application, including the elaboration of the compensator design, its mid-area control method, and simulation of the orbital and relative motion of the shepherd – space debris, consider the disturbing factors and the relative motion control algorithm

On Calculation of the Ignition Voltage for a Back-Arc Discharge in a High-Voltage Thermionic Diode

Development of large spacecraft propulsion units involves the creation of high-temperature radiation-resistant systems for current conversion based on thermionic devices of plasma electric power engineering, such as grid-controlled switches and high-voltage plasma thermionic diodes (HVPTD). The current conversion system (CCS) is required to match the electrical parameters of a thermionic conversion power reactor (with an output voltage of 120–150 V) with the parameters of the electric jet engine (having a working voltage of hundreds or thousands of volts). This brings about an urgent issue of calculating the ignition voltage for the back-arc discharge in the interelectrode gap (IEG) of HVPTDs. A semiempirical correlation is proposed for predicting the back-arc discharge ignition voltage as a function of the cesium vapor pressure in IEG and the anode temperature. It can be used by designers of thermionic plasma power engineering devices.

Облик космического аппарата с электроракетной двигательной установкой на забортном воздухе

Облик космического аппарата с электроракетной двигательной установкой на забортном воздухе

Dynamics of the Formation of a Rotating Orbital Tether System with the Help of Electro-thruster

The dynamics of the process of forming the rotating orbital tether system, which is a bundle of two small satellites, is analyzed. The cases of transition deployed tether system in rotation by means of two electric jet engines located on the end bodies are considered. Engines together form a pair of forces required sign, the control program includes active and passive portions of the motion system. Satellites are material points whose masses are comparable in magnitude, and the tether is weightless. The control program is formed by a relatively simple model in the mobile orbital coordinate system in the case of inextensible tether. The elongation effect of the tether, which leads to its eventual sagging is analyzed by a more complete model in a fixed geocentric coordinate system.

Effective sputtering yields of titanium, titanium nitride, and molybdenum induced by nitrogen ions and ions of nitrogen-oxygen mixtures

This work is dedicated to studying the effects of nitrogen ions and ions of nitrogen and oxygen mixtures on the surface of titanium, titanium nitride, and molybdenum. The usage of magnetron sputtering systems as a model device to study the effect of reactive gases on elements of electric jet engines is proposed and justified. The processes of sputtering of a surface exposed to non-monoenergetic ion beams are studied. The effective sputtering yields of titanium, titanium nitride, and molybdenum induced by argon and nitrogen ions and ions of nitrogen-oxygen mixtures at various intermediate-energy ion beams are determined. It is shown that the sputtering yields of reactive-gas ions are significantly lower than the sputtering yields of inert gases.

Nuclear power propulsion system for spacecraft

The proposed designs of high-power space tugs that utilize solar or nuclear energy to power an electric jet engine are reviewed. The conceptual design of a nuclear power propulsion system (NPPS) is described; its structural diagram, gas circuit, and electric diagram are discussed. The NPPS incorporates a nuclear reactor, a thermal-to-electric energy conversion system, a system for the conversion and distribution of electric energy, and an electric propulsion system. Two criterion parameters were chosen in the considered NPPS design: the temperature of gaseous working medium at the nuclear reactor outlet and the rotor speed of turboalternators. The maintenance of these parameters at a given level guarantees that the needed electric voltage is generated and allows for power mode control. The processes of startup/shutdown and increasing/reducing the power, the principles of distribution of electric energy over loads, and the probable emergencies for the proposed NPPS design are discussed.