Electric Propulsion Engines Research Articles

Selection of design parameters of a spacecraft electric propulsion engine ensuring the geostationary orbit raising mission

The problem of selecting suboptimal parameters of an electric propulsion system as part of a spacecraft, ensuring a geostationary orbit raising mission using a Soyuz-5 launch vehicle at the initial launch stage, is investigated. An expression is obtained for the main optimality criterion – the payload relative mass. An algorithm is obtained for solving the problem of parametric optimization of an electric propulsion system as part of a hybrid transportation system. The task of selecting the optimal electric propulsion engine for launching a spacecraft with an electric propulsion system to a geostationary orbit using the launch vehicle under consideration is solved. A design ballistic calculation and a comparative analysis of the obtained data are performed.

Data Science and Machine Learning Processes for IoT Based Pulsed Plasma Thruster Research

Pulsed Plasma Thrusters (PPT) are very high specific impulse but low efficient electric propulsion engine which are being used as primary or secondary propulsion mechanism for spacecrafts (cube, micro, nano or pico satellites). PPTs are used in academic and/or industrial research from decades, but its experimental data collection methods are normally offline, analysis methods are conventional, and data interpretation lies in electromagnetic, physical, and chemical domain. It is inadequate to explain PPT generated data in few specific domains only and it may prone to inadequate explanations which can hide vast insights of data. Actually, PPT generated data are usually big data. It is essential to analyze and explain PPT data in data science and machine learning domain to get keen insights from the data. To this date no such complete solution exists for PPT either in industrial arena or academia to use all these three technologies. To meet the gap, we propose and implement an Internet of Things (IoT) based architecture for PPT experimental facility to collect, accumulate, and process PPT data intelligently. This architecture yields some machine learning (ML) models with prediction accuracy above 93%. This architecture is a complete IoT and big data based Artificial Intelligence (AI) implementation on PPT data through data science and ML processes. This architecture is capable to contribute in both academic as well as industrial research, development and deployment.

Deep reinforcement learning spacecraft guidance with state uncertainty for autonomous shape reconstruction of uncooperative target

In recent years, space research has shifted heavily its focus towards enhanced autonomy on-board spacecrafts for on-orbit servicing activities (OOS). OOS and proximity operations include a variety of activities: the focal point of this work is the autonomous guidance of a chaser spacecraft for the shape reconstruction of an artificial uncooperative object. Adaptive guidance depends on the ability of the system to build a map of the uncertain environment, figuring out its location inside of it and accordingly determining the control law. Thus, autonomous navigation is framed as an active Simultaneous Localization and Mapping (SLAM) problem and modeled as a Partially Observable Markov Decision Process (POMDP). A state-of-the-art Deep Reinforcement Learning (DRL) method, Proximal Policy Optimization (PPO), is investigated to develop an agent capable of cleverly planning the shape reconstruction of the target. Starting from previous research on the topic, this work develops further proposing a continuous action space, such that the agent is no more forced to choose between a predefined set of possible discrete actions, fixed both in magnitude and direction. In this way any combination of the three-dimensional thrust vector components is available. The chaser spacecraft is a small satellite mounting an electric propulsion engine defining the action space range, in linearized eccentric relative motion with the selected uncooperative object. Through a rendered triangular mesh, the agent capabilities of geometry reconstruction and mapping are evaluated, considering the number of quality pictures made for each face. Extensive training tests are performed with random initial conditions to verify the generalizing capability of the DRL agent. The results are then validated in a comprehensive testing campaign, whose primary focus is the introduction of noisy measurements coming from navigation, affecting pose estimation. The sensitivity of the proposed method to this condition is analyzed and the efficiency of a retraining procedure is examined. The applicability of DRL methods and neural networks to support autonomous guidance in a close proximity scenario is corroborated and the technique employed is vastly tested and verified.

Design of a Fast Molten Salt Reactor for Space Nuclear Electric Propulsion

The French National Center for Scientific Research (CNRS) is carrying out design studies on a nuclear electric propulsion (NEP) engine based on a molten salt reactor (MSR). A NEP engine based on liquid nuclear fuel could allow developing a core design with relatively high power densities and temperatures while using simple reactivity control systems and keeping low pressure and temperature gradients in the fuel. Nevertheless, the design work of such an engine poses significant technical challenges and requires the use of advanced numerical simulation tools. Different MSRs for space are currently being studied. In this work, a MSR concept using a fast neutron spectrum is investigated using a multiphysics tool based on a numerical coupling between the OpenFOAM (computational fluid dynamics) and SERPENT 2 (Monte Carlo neutronics) codes. The analysis of this paper is focused on the reactor core coupled neutronic and thermal-hydraulic phenomena. Steady state full-power conditions are calculated for two different fast MSR designs using low-enriched uranium (LEU) and highly enriched uranium. The results show that the proposed core layout and materials allow obtaining a satisfactory temperature distribution in the core (maximal values and gradients) without significant penalization of the reactor operating conditions. A reactivity control strategy excluding the use of control rods is studied for the LEU concept. Transient and safety studies are also performed and show acceptable performance.

Performance evaluation and off-design strategy of a high temperature proton exchange membrane fuel cell turbine-less hybrid system for low pressure ratio and Mach number aircraft

The aviation industry is moving towards low-carbon and environment-friendly electric propulsion. This paper presents for the first time a high temperature proton exchange membrane fuel cell turbine-less hybrid system for low pressure ratio and Mach number aircraft. It is a new electric propulsion engine in which a plate reformer is used to supply hydrogen to the fuel cell. To investigate its performance and off-design strategies, thermodynamic, exergetic and exergoeconomic analyses models are established. Firstly, the thermodynamic performance is described and compared with that of a turbojet. Then, exergetic and exergoeconomic performances are revealed. Finally, the impact and comparison of four off-design strategies are studied in detail. The results show that the system can reduce specific fuel consumption by 24.90%. The burner, core nozzle and high temperature proton exchange membrane fuel cell need further development due to their low exergy efficiencies and high irreversibility ratios. The system has an exergoeconomic factor of 11.32%, relative cost difference of 24.65%, carbon dioxide mass specific emissions of 2.70 kg/(kW·h), and total system cost of 82.50 $/h. Strategy 2 is an economic option as it produces the minimum of specific fuel consumption, carbon dioxide mass specific emissions and total system cost.

Aerodynamic Design Optimization and Analysis of Ducted Fan Blades in DEP UAVs

In order to reduce the dependence on fossil fuels, the concept of distributed electric propulsion (DEP) has been introduced. At present, there are a few studies on the small electric propulsion engines of DEP vertical take-off and landing (VTOL) UAVs. According to the design requirements of DEP VTOL UAV ducted fans, the aerodynamic design of a ducted fan is carried out in this paper. Firstly, the initial model is established by referring to theoretical derivations. Then, Bezier curves are used to control the mid-arc and thickness distributions of the blade profiles at different blade heights in order to realize the parameterization as design variables. Finally, the objective function that can represent the performance of the rotor is selected, and a multi-objective genetic algorithm is used to optimize the performance. The optimal results are obtained, and the aerodynamic characteristics are analyzed. Compared with the initial model, the total pressure ratio and isentropic efficiency of the optimized model are increased by 5.3% and 7.8%, respectively, reaching 1.20 and 90.37%. Meanwhile, the aerodynamic performance of the optimized model is improved, indicating that the performance of the rotor is significantly enhanced and that the design task is completed.

Оптимальные затраты на комбинированный увод крупногабаритных объектов космического мусора с использованием электрореактивной двигательной установки

The article proposes the development of a method for the combined deorbiting large space debris objects from low Earth orbits, performed using an electric propulsion engine and an aerodynamic sailing device. Simulation modeling of the combined deorbiting was carried out for various combinations of parameters of deorbit scheme, such as the altitude of the initial orbit, the phase of solar activity at the moment of deorbiting start, the ballistic coefficient, the time of active operation of the control system, and the time of one battery charge. Analytical dependences of the minimum increment in the velocity of an electric propulsion system, gained in one impulse, and the minimum number of impulses on the parameters of the deorbiting scheme, necessary to ensure the withdrawal for 25 years, are determined. Sectors of the solar activity phase at the moment of the deorbiting start providing optimal energy costs for the withdrawal process, are identified. The results obtained are of practical interest for the problems of designing modern means of deorbiting large space objects from low Earth orbits at enterprises in the rocket and space industry.

Mathematical Modeling of the Dynamics of a Low-Flying Spacecraft with a Ramjet Electric Propulsion Engine

Mathematical Modeling of the Dynamics of a Low-Flying Spacecraft with a Ramjet Electric Propulsion Engine

Intrinsic interfaces between additively manufactured metal and composite structures for use in electric propulsion engines

For electric aviation drive systems engines with significantly increased power density are required. Combining the technologies of additive manufacturing of metals with the technologies of fiber reinforced composites enables hybrid structures with same or increased functionality and lower mass. Rotors or housings in hybrid design have interfaces to adjacent structures are classically made of metal. Areas for remote load transmission or with increasing distance to the rotation axis can be made of carbon fiber reinforced plastic (CFRP). The interfaces between these two materials, required for such a hybrid metal-CFRP design, can be specifically designed by using the design options of additive manufacturing to fulfill the required functions with a the most even distribution of stress possible. However, the large number of degrees of freedom of the material and the geometry leads to great challenges for engineers in the development of such structures. Within the scope of this publication, an approach for the function-oriented design of a metal-CFRP rotor for an aviation electric motor is proposed (global design). A design and pre-dimensioning method for the intrinsic interface between the materials is presented (local design).

Analysis of characteristics of electric propulsion systems intended for carrying out maneuvers of maintenance of low Earth working orbit of small satellites

An analysis of the mass of the working fluid and motor operating time of electric propulsion systems applied as a part of small spacecraft to carry out maneuvers of maintenance of the low Earth working orbit is carried out. The analysis is carried out for the small spacecraft with the weight in the range from 300 to 1000 kg functioning in working orbits with the height in the range from 400 to 600 km. When carrying out the analysis the values of the specific impulse of the propulsion system in the range from 800 to 1600 sec were accepted. Procedural guidelines for assessing the value of the required characteristic speed depending on the aerodynamic drag force, as well as for assessing the value of mass of the working fluid with account for the value of the specific impulse and defining the motor operating time of the propulsion system depending on the exhaust speed of the working fluid were used. The results of calculations given in the article show that the mass of the working fluid and the motor operating time vary depending on the height of the orbit and the mass of the small spacecraft and allow making quick preliminary assessment of the main design characteristics of the electric propulsion engines used to carry out maneuvers of maintenance of the low Earth working orbit of small spacecraft with different weight dimension characteristics during the prescribed term of active existence.

Thermophysical characteristics of gas-power dual-fuel engines w12v50 df for liquefied gas shipping vessels

Maritime transport plays important role in the economic development of society – 90% of goods are transported by ships. At the same time, maritime transport requires a significant amount of fuel resources. Production of liquefied natural gas (LNG) is becoming the fastest growing industry in the modern global energy sector. Today, LNG accounts for 40% of the physical volume of world gas trade, and its share will increase up to 60% by 2040. Currently, natural gas is used on ships in the form of liquefied petroleum gas, compressed natural gas, and liquefied natural gas (LNG). The article deals with the urgent problem of operation of dual-fuel diesel-electric installations of ships. The need to study the heat-engineering parameters of two-fuel diesel generators of the Wartsila company has been substantiated. The authors present the dependencies of main heat-engineering parameters on the load of Wartsila W12V50DF dual-fuel engines used as a generator drive in the main electric propulsion engines on LNG tankers. A comparative assessment of the dependencies of exhaust gas temperature, turbocharger rotation speed, boost pressure and gas pressure on the load of diesel generators on two LNG tankers has been carried out. The article analyzes the presented dependencies. The authors substantiate the need for further improvement of their design and workflow.

Сравнительный анализ эффективности дросселированных электрических ракетных двигателей постоянной мощности для межорбитальных перелетов на геоста

Осуществление перелетов космических аппаратов в околоземном пространстве с промежуточных эллиптических орбит на геостационарную с использованием электрических ракетных двигателей (ЭРД) малой тяги является одной из самых актуальных задач современной космонавтики. ЭРД, скорость истечения реактивной струи которых на порядок больше, чем у химических РД, лучше для межорбитальных околоземных перелетов с максимальной полезной нагрузкой в случае, когда существенное увеличение продолжительности маневра является допустимым. Дросселирование тяги ракетного двигателя традиционно рассматривается как один из способов понижения массы двигательной установки и необходимого запаса топлива для выполнения заданного маневра. Использование концепции идеально регулируемого двигателя обеспечивает получение верхних оценок массы полезной нагрузки межорбитальных перелетов для выбранного уровня мощности. Учет свойств реальных двигателей приводит к необходимости рассмотрения математических моделей с более жесткими ограничениями на управление. Проведено исследование эффективности трех режимов управления тягой ЭРД при выполнении практически интересных перелетов с высокоэллиптических промежуточных околоземных орбит на геостационарные. Построена математическая модель релейного ЭРД постоянной мощности. Дана формулировка вариационной задачи типа Майера о выполнении заданного динамического маневра космического аппарата для дро-сельированного и нерегулируемого двигателей постоянной мощности. С использованием принципа максимума Понтрягина проведен анализ оптимальных управляющих функций, для которых выписаны конечные соотношения, позволившие записать систему дифференциальных уравнений оптимального движения космического аппарата, оснащенного релейным двигателем. Полученные численные и качественные результаты исследования эффективности различных режимов управления тягой для увеличения полезной нагрузки заданного орбитального маневра подтвердили корректность математических моделей дросселированного и релейного двигателей и, в целом, эффективность использования решений усредненных уравнений оптимального движения космического аппарата для численного решения соответствующих краевых задач в точной. постановке.

Obtaining a composite material based on quartz woven filler and pyrolysis matrix of organosilicon resin

The relevance of the study is conditioned by the fact that the most popular and irreplaceable materials that have found wide application in the aerospace industry are composites based on quartz materials. These materials are distinguished by their high mechanical and electrical strength, chemical and corrosion resistance. In this regard, it is of interest to obtain a composite material that combines a low specific gravity, processability of polymers, and thermal stability of ceramics. The aim of this work was to study the effect of the temperature of thermal oxidative destruction of a polymer binder, which is a semi-finished product of a pyrolysis matrix, on the electrophysical parameters of a composite material. The paper investigates a composite material based on woven quartz material with a pyrolysis matrix of an organosilicon binder and functional additives. This composite was considered as a material for creating an electric rocket engine chamber. Thermogravimetric analysis was used to evaluate the effect of the temperature of thermooxidative degradation of the polymer binder on the electro-physical parameters of the obtained material. The tests were carried out according to standard test methods on an Instron 5969 universal testing machine with Bluehill software until the samples failed. In the course of the study, it was found that the processes that occur up to 400°C are mainly associated with the course of the reaction for non-entered functional groups, the telomerisation reaction, intramolecular rearrangement of macromolecules and the removal of low-boiling substances. According to the results of the study, the obtained characteristics of the test material turned out to be suitable for its use in structural elements of electric propulsion engines.

Dynamics and control of an electrodynamic tug: Transfer to the graveyard orbit

Space debris transfer to the graveyard orbit is a promising solution for impact mitigation in geosynchronous orbits. This paper is a development on previous work done on the eponymous technique to capture and transfer the debris to graveyard orbit located at 200 km above geostationary orbit. Here, the focus is kept on studying the dynamics and control of the collector-debris system. The debris is assumed to be spherical in shape and composed of a magnetically susceptible material with properties of copper-aluminum alloy. The capture and transfer phase is realized by low-thrust transfer trajectory using electric propulsion engine as main thruster and two control thrusters for continuous actuation of relative position of debris. A feedback control law is developed to ensure that the debris remains at the desired equilibrium position during the transfer maneuver. Stability of the system is investigated using results drawn from numerical simulation. The results show how the choice of control coefficients affects the stability of the system.

Investigation of the attenuation of the fluxes of the low-energy plasma component of electric rocket engines by the ventilation holes of the unpressurized instrument compartment of the spacecraft

The results of an experimental study of the attenuation of the fluxes of the low-energy component of the plasma formed during the operation of electric propulsion engines (ERE), ventilation holes (VH) of the non-sealed equipment compartment (NSEC) of the spacecraft (SC) are presented. Authors studied the attenuation of plasma fluxes by standard VHs made in honeycomb panels that form the NSEC. A Hall-effect engine of the SPT-70 type was used as a plasma source. The experiment consisted of measuring the plasma concentration at the inlet and outlet of the VH. The concentration at the inlet was measured with a flat Langmuir probe, and at the outlet with a Faraday probe, which allows collecting all ions passing through the VH. The aim of the work was to study the weakening of the fluxes of the lowenergy component of the EJE plasma when passing through the VH in the honeycomb-nels that form the NSEC. Based on the experimental data, a semi-empirical model was constructed that describes the dependence of the attenuation coefficient of plasma flows on the geometric parameters of the vent-holes. It has been established that a vent-holes of this design attenuates the plasma flows by 102 ... 104 times. The largest contribution to the weakening of plasma fluxes is made by the honeycomb filler, which is due to the recombination of ions during their collision with the channel walls. Taking into account the attenuation of the fluxes of the low-energy component of the plasma of electric rocket engines by ventilation holes is a key stage in assessing the effect of plasma on the power on-board equipment of spacecraft and should be used by spacecraft developers when analyzing the resistance to this factor.

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.

POLYMER PRECURSORS FOR CREATING GAS DISCHARGE CHAMBER FOR ELECTRIC ROCKET ENGINE

Electric rocket engines are widely used in space technology. Furthermore, at present, electric propulsion engines are also used as mid-flight engines for flights in interplanetary space. On modern spacecraft, the following types of electric propulsion are mostly used: SPT and grid ion thruster. When using these engines as sustainers, it is important is to increase the total power for obtaining the required thrust and specific impulse. With an increase in total power, the volume of the discharge chamber increases, which leads to technological difficulties in the manufacture of discharge chambers from ceramic materials. Thus, the task of finding alternative ceramic materials is relevant and necessary in the development of high-frequency ion thrusters. The article discusses the issues of creating a composite material based on woven quartz materials and organosilicon binder as a precursor filled with silicon nitride for the manufacture of gas discharge chamber (GDC) of high-frequency ion thruster (RFIT). By thermos-gravimetric analysis, a thermosetting binder, which meets the requirements of vibration resistance and electromagnetic permeability of GDC in the megahertz range, was selected. Based on the binder filled with silicon nitride powder, reinforced by quartz woven fabrics, manufactured GDC. The resulting product was tested as part of the laboratory electric propulsion device with a diameter of 100 mm and power of 200W.

The electromagnetic problems of interplanetary spacecraft communication

The results of the radio frequency emission measurements from the ground-based experiments using electric propulsion engine applied to electromagnetic compatibility evaluation of spacecraft radio communication. The report presents the way of looking at the electromagnetic compatibility problem based on experimental and theoretical investigations of the physical foundation of the electromagnetic field generation by the electric propulsion engines. The approach under review allows evaluating the electromagnetic performance characteristics depending on thruster construction and engine operating modes.

Main engine calculation for ferry Ro-Ro 500 DWT ship using electric propulsion

The development of the current design of electrical system to protect the environment in the world. The purpose of this research is to obtain an electric propulsion system and engine power needed by the Ferry RoRo 500 DWT ship. The analysis begins by calculating the total ship resistance using the software. Next, the calculation of determining the generator power is calculated by a mathematical method. Resistance of 500 DWT RORO ferry at 12 knots 67.7 kN service speed. The increase in ship resistance is affected by the volume of the hull below the waterline. The greater the volume the greater the value resistance ship to get. The power play engine generator required to fight the value of ship resistance at the ferry RORO service speed of 984.49 kW.

Analysis of the effectiveness of the de-orbiting devices for small satellite

At present, the use of various methods is proposed for the de-orbit of small satellite, and the most feasible and promising of them are analyzed. The task of evaluating the effectiveness of the de-orbiting system for small satellite is set, in the framework of which a criterion and basic performance indicators are formed taking into account design features. As a methodological basis for evaluating the effectiveness, the method of relative integral assessment was used. Using the developed algorithm for calculating the coefficients of the integral relative assessment for each de-orbiting method, the most effective option for constructing the system isdetermined for given priority coefficients and taking into account the imposed design restrictions. For the analysis of efficiency, fuel-free de-orbiting devices and three types of propulsion systems were chosen – electric propulsion engine, solid rocket motor and liquid engine. Efficiency analysis was carried out for devices with various mass and target characteristics, the result is the choice of the type of de-orbiting system and the calculation of its parameters. The dependence of the de-orbiting device on the purpose of the spacecraft, the altitude and inclination of the orbit of its functioning, as well as on the requirements for mass, cost and other design parameters is shown.