X-57 power and command system design

Most of marine electric propulsion ships are propelled by large electric motors. The speed control of the propulsion motor on the electric propulsion vessel is operated by adjusting the frequency. The power system has the inverter installed semiconductor element, such as IGBT or IGCT, for high-speed switching. However, the inverter for power conversion unit of the electric power propulsion on large vessels has fundamentally thermal problems of installed semiconductor, such problems could influence to considerable loss by switching and conducting the inverter in the power system transferring large current with high voltage. In order to calculate power loss and thermal loss of the inverter, this paper sets actual parameters of the electric propulsion system of the large ship in the simulation program considering condition of ship operation. The simulation proves that it is possible to analyze both power loss and thermal loss in the diodes and transistor during switching of the inverter in comparison with field oriented control and direct torque control. As a result, large power loss and thermal loss of switching and conducting the inverter by field oriented control considerably occurs and can also impact damages toward the power electronic devices by the thermal loss generated.

The mission targets of the Chinese asteroid exploration are an earth co-orbiting celestial body 2016HO3 and the main-belt comet in the asteroid belt. Compared with Lunar Exploration and Mars Exploration missions, asteroid exploration mission is special. The light intensity varies widely during the mission, resulting in a large change in the output power range of the solar array, which seriously affects the power system design and the load design during the mission; the spacecraft uses a high-power electric propulsion system; the power system has strict weight requirements, the existing power supply system design cannot meet the requirements. Aiming at the design constraints of the large variation of solar wing light intensity, the large difference in power requirements between the electric propulsion and the platform, and the lightweight design of the spacecraft, a dual-bus power supply system is proposed. High-voltage bus power the electric propulsion system. The solar cell array directly converges to form a high-voltage bus, which supplies power for the electric propulsion system. The low-voltage bus adopts a semi-adjustable bus, and the solar cell array adopts a circular flexible solar wing. The main detector and return cabin are equipped with lithium-ion battery pack and lithium fluorocarbon battery pack respectively. Aiming at the requirement of long-term on-orbit operation of electric propulsion system, a solar wing output power measurement system is designed. Product test results show that the power system design can meet the working requirements of electric propulsion and platform loads. The power system design can provide reference for the other deep space exploration power system.

A brief status review of the advanced launch system (ALS) and its associated advanced development program are presented to demonstrate maturation of those technologies that will help meet the overall operability and cost goals. The electric power and actuation systems are highlighted as a specific technology ready not only to meet the stringent ALS goals (cryogenic fuel valves and thrust vector controls with peak power demands to 75 hp), but also those of other launch vehicles, military and civilian aircraft, lunar/martian vehicles, and a multitude of commercial applications. >

Near-Earth Objects (NEOs) are exciting targets for exploration; they are relatively easy to reach but relatively little is known about them. With solar electric propulsion, a vast number of interesting NEOs can be reached within a few years and with extensive flexibility in launch date. An additional advantage of electric propulsion for these missions is that a spacecraft can be small, enabling a fleet of explorers launched on a single vehicle or as secondary payloads. Commercial, flight-proven Hall thruster systems have great appeal based on their performance and low cost risk, but one issue with these systems is that the power processing units (PPUs) are designed for regulated spacecraft power architectures which are not attractive for small NEO missions. In this study we consider the integrated design of power and propulsion systems that utilize the capabilities of existing PPUs in an unregulated power architecture. Models for solar array and engine performance are combined with low-thrust trajectory analyses to bound spacecraft design parameters for a large class of NEO missions, then detailed array performance models are used to examine the array output voltage and current over a bounded mission set. Operational relationships between the power and electric propulsion systems are discussed, and it is shown that both the SPT-100 and BPT-4000 PPUs can perform missions over a solar range of 0.7 AU to 1.5 AU - encompassing NEOs, Venus, and Mars - within their operable input voltage ranges. A number of design trades to control the array voltage are available, including cell string layout, array offpointing during mission operations, and power draw by the Hall thruster system.

This paper examines the role of the TOPAZ II and the 40 kWe thermionic reactors in powering electric propulsion systems for orbital transfer and on‐orbit repositioning applications. Specifically, the feasibility of substituting a TOPAZ II power system and the Russian Stationary Plasma Thruster (SPT) for the conventional power and propulsion systems that might be used on typical Air Force navigation, surveillance, and communications satellites were examined. Additionally, two advanced high‐power missions were analyzed. Four major conclusions were drawn from this top‐level investigation: (1) Stepdown to smaller launch vehicles is possible for both TOPAZ II and 40 kWe systems even though the TOPAZ II at 6 kWe and 1100 kg, is not an extremely effective system for orbital transfer applications; (2) On‐orbit repositioning is an area that profits by the utilization of electric propulsion systems; (3) Advanced high‐power missions are arguably the greatest benefactor of nuclear electric propulsion; and (4) Placing minimum operational altitude requirements on reactor‐propelled systems will significantly diminish their payload capability to the destination orbit.

This paper presents a study on the feasibility of using electric propulsion (EP) for interplanetary transfer in the context of an Earth–Venus mission. Extensive research has been conducted on EP modules for Mars missions, demonstrating their potential for faster mission times and greater flexibility compared to all-chemical propulsion (CP) systems. However, there is a notable gap in our understanding regarding EP systems for Venus missions, despite the unique and challenging environment that Venus presents. This study aims to bridge that gap by evaluating the performance of an EP module compared to a conventional all-CP spacecraft, which serves as the backup mission for the EnVision mission by the European Space Agency. The objectives of the study include identifying key drivers for the EP module’s preliminary design, analyzing tradeoffs and challenges associated with EP systems, and evaluating the feasibility of achieving payload mass delivery. The results reveal that the all-EP spacecraft delivers a higher payload mass compared to the all-CP EnVision mission, highlighting the superiority of EP over CP for interplanetary missions. This research provides valuable insights into the potential advantages of EP systems for future interplanetary missions and emphasizes the necessary design considerations when utilizing EP technology.

Abstract: In this paper, describes about advancement in solar-based power system in satellite. A thrust which provides a back force to satellites that move in upward direction and get satellites help to generate power in space. This power transmits to the earth via wireless power transmission and such type of satellite called space based solar power satellite. In recent time, electrical propulsion system has used mostly in place of chemical propulsion system for satellite. Solar energy collected from the space and converts into DC then transmit power from the space to earth wireless and then send to electric grid system. Solar energy is also used by the thruster elements in DC. Keywords: WPT (wireless power transmission), SBSP, propellant, propulsion Cite this Article Divyanshi Srivastava, S.K. Srivastava. Solar-based Satellite Power System: An Overview. Journal of Thermal Engineering and Applications. 2019; 6(1): 20–23p.

Hybrid Aircraft Propulsion System Synthesis (HAPSS) is a computer program that sizes and analyzes pure-series hybrid electric propulsion systems for aircraft. The development of this program began during a NASA SBIR contract, in conjunction with Empirical Systems Aerospace (ESAero), with the creation of a propulsion fan design tool. Since the completion of this contract in July 2010, the HAPSS program has been expanded to combine the many aspects of a hybrid propulsion system such as the propulsive fans, electric motors, generators, and controllers, and the internal combustion engines. This thesis describes the benefits and drawbacks of aircraft hybrid propulsion systems to reveal the usefulness of a program of this nature. The methodology behind HAPSS, the creation of the program, its operation, and its many applications are also discussed in detail. Finally, this thesis includes a brief example in which HAPSS is used to analyze a hybrid propulsion system for a commercial transport aircraft. This example demonstrated the usefulness of the program and revealed interesting behavior and trends unique to hybrid propulsion. To date, the HAPSS program has been utilized on several different contract projects in which an aircraft hybrid propulsion system was designed. In the summer of 2012, a government organization in conjunction with ESAero will begin funding a contract to continue the development of HAPSS by adding functionality and improving accuracy while making the tool available to other government agencies.

This paper presents a comprehensive approach to the sizing of parallel hybrid propulsion systems for multi-megawatt aircraft, with a specific focus on baseline turboprop aircraft. By exploring key performance variables such as power distribution, system efficiency, or fuels and aircraft weights. The study provides a detailed analysis of how these factors influence overall hybrid aircraft performance. Special attention is given to the impact of these variables on range and fuel consumption, aiming to quantify potential fuel savings and range improvements under different operating conditions. The paper investigates the level of hybridization at various phases of flight, such as take-off, climb, cruise, and descent, seeking to optimize the distribution of power between the electric power train (fuel cell sizing) and gas turbine propulsion systems. By evaluating different scenarios and performance sensitivities, the study aims to identify optimal design strategies for enhancing both efficiency and sustainability in future aircraft propulsion systems. The findings contribute to a better understanding of how hybrid-electric technologies can be integrated into next-generation turboprop aircraft, with implications for reducing emissions and quantify impact on top level aircraft requirement.

THE PURPOSE. Consider the issue of the presence of distortions in the form of sinusoidal voltage and current in autonomous electrical power systems with electrical propulsion systems, built on the principle of the unified electric power system. Compare the results of the study of voltage distortion for electrical power systems with electrical propulsion systems of various structures and give recommendations for their application.METHODS. For research, unified electric power systems with DC electrical propulsion systems on the Yeysk ferry and alternating current on the asymmetric icebreaker Baltika are considered. The possibilities of programming with frequency control of modern drives of rudder propellers with AC motors are analyzed.RESULTS. Oscillograms of voltages and currents of generators were obtained using thyristor converters and inverters to control electrical propulsion systems in various modes. Significant ripples and distortions of the sinusoidal voltage and generator currents were noted when thyristor converters are used to power DC propulsion electric motors. Also, small deviations from the sinusoidal form of the ship's network voltage were recorded when inverters were used to control AC propulsion motors as part of rudder propellers.CONCLUSION. The issue of ensuring the quality of electricity produced in autonomous electrical power systems of sea vessels is of great relevance and importance. To ensure the best results, it is advisable to use azimuthing podded drive with AC electric motors, the rotation speed of which is implemented by inverter frequency converters with a DC link.

ABSTRACTThe increasing environmental requirements in the air transport sector pose great challenges to the aviation industry and are key drivers for innovation. Besides various approaches for increasing the efficiency of conventional gas turbine engines, electric propulsion systems have moved into the focus of aviation research. The first electric concepts are already in service in general aviation. This study analyses the potentials of electric and turbo hybrid propulsion systems for commercial aviation. Its purpose is to compare various architectures of electrical powertrains with a conventional turboprop on a regional aircraft, similar to the ATR 72, on engine and flight mission levels. The considered architectures include a turbo-electric (power controlled and direct driven), hybrid-electric (serial and parallel) and a pure electric concept. Their system weights are determined using today’s technology assumptions. With the help of performance models and flight mission calculations the impact on fuel consumption, CO ${}_{2}$ emissions and aircraft performance is evaluated.

Electric motor performance directly affects the overall performance of an electric vehicle (EV). The electric motor windings heats up when operated under extreme load and hence, reducing the efficiency of the vehicle. As such, EV propulsion system requires cooling systems, not only to the electric motor but also to the battery banks in order to ensure efficient operation and maximizing the electric components and vehicle lifetime. The project focuses on the design and development of a liquid cooling system for an electric vehicle propulsion system to determine the optimum size and cooling capacity by thermodynamic analysis. This paper described the analysis of energy transfer (heat removal) by using thermal resistance network approach in determining the suitable design for the cooling system of an AC-50 induction electric motor. The CAD drawing was generated to suit the dimension constrain and the cooling system was completed using CATIA software. The results showed that the optimal design for the water jacket for the maximum heat removal of 5500 W from the electric motor, should have a wall thickness of 0.5 cm with an annular thickness of 2.4 cm. Both copper and aluminium exhibited similar cooling capabilities with the latter having lower cost and ease of manufacturing.

Power processing units (PPUs) in an electric propulsion system provide many challenging integration issues. The PPU must provide power to the electric thruster while maintaining compatibility with all of the spacecraft power and data systems. Inefficiencies in the power processor produce heat, which must be radiated to the environment in order to ensure reliable operation. Although PPU efficiencies are generally greater than 0.9, heat loads are often substantial. This heat must be rejected by thermal control systems which generally have specific masses of 15-30 kg/kW. PPUs also represent a large fraction of the electric propulsion system dry mass. Simplification or elimination of power processing in a propulsion system would reduce the electric propulsion system specific mass and improve the overall reliability and performance. A direct drive system would eliminate all or some of the power supplies required to operate a thruster by directly connecting the various thruster loads to the solar array. The development of concentrator solar arrays has enabled power bus voltages in excess of 300 V which is high enough for direct drive applications for Hall thrusters such as the Stationary Plasma Thruster (SPT). The option of solar array direct drive for SPTs is explored to provide a comparison between conventional and direct drive system mass.

This review deals with the selection of the electric propulsion system (EPS) for the internationally developed and designed, primary nuclear-electric space tug International Nuclear Power and Propulsion System (INPPS). INPPS is scheduled for interplanetary missions to Mars and Jupiter moon Europa missions by the end of decade 2020. Regarding specific technical and mission parameters preselected electric thruster (ET) types, developed by international companies and institutions, are analysed, evaluated and investigated for a possible application as propulsion system (PS), the so-called CET (Cluster of Electric Thrusters). It is analysed whether solely electric thrusters, combined in an adequate CET, enable the envisaged interplanetary missions—robotic and astronautic/crewed with the INPPS flagship.Thruster clusters with strategic consortium considerations are analysed as a feasible PS of the INPPS. The studied CET consists of the following: (a) only European ETs, (b) combination of German and European ETs, (c) Japanese and European ETs or at least (d) Japanese, European and US thrusters. The main results are (1) Robotic and crewed INPPS mission to Mars/Europa are realizable with EPS only (no chemical propulsion is needed), (2) that every CET, except (c) of only Japanese and part of European thrusters, is capable to perform the main part of envisaged INPPS flagship mission orbit to Mars, back to Earth and to Jupiter/Europa moon.

In this thesis a detailed study on the battery technologies used in aircrafts in the last five decades is done. A general background of the battery system is studied and the performance of the batteries based on energy densities and low temperature capabilities is evaluated for the deployment in the more electric aircraft (MEA). Evolution of MEA with its power system architecture and load profile is analysed to understand the requirements of the battery system. The electrical load profile of the MEA “Aircraft X” is designed and modeled. Different types of electrical loads such as primary/conventional loads, the environmental control system (ECS) load, anti-icing (wing ice protection) loads, various electrical actuators (aileron, spoilers, rudder) loads and signal communication (RADAR etc.) loads are designed and modeled. Furthermore, starting profile of engine starter generator (ESG), transient stability profile of military standard (MIL-STD-704F) and emergency load profile of the MEA “Aircraft X” are also studied and analysed. Thus, all these electrical loads are combined and an algorithm is successfully obtained to separate the constant and transient loads. An effective battery system is designed for supplying the transient loads by considering the weight saving and cost analysis factors between Li-ion and Ni-Cd batteries. Based on the analysis, Li-ion battery with Lithium ferrous phosphate (LFP) chemistry is designed and proposed for the MEA “Aircraft X”. The designed LFP based battery system is also modeled using the modified shepherd equation wherein the voltage polarization term is added to have a lower complexity and more proximity with the battery model block in MATLAB®. After the design and modeling of the proposed Li-ion battery system, the battery charger (phase shifted high power bidirectional dc-dc “PSHPBD” converter) is designed and modeled. The design of the components of PSHPBD converter is done keeping in view the requirement of maximum power transfer in the MEA environment. Zero voltage switching (ZVS) operation in wide range is also considered for design of the PSHPBD converter for MEA operating parameters. The PSHPBD converter is modeled using the Fourier series based harmonic analysis wherein an optimal harmonic number (h=7) is chosen by considering the proximity and complexity between the modeled converter and switch model. The battery energy storage system “BESS” (consisting of the battery system and the PSHPBD converter) is successfully integrated with the power distribution 270V dc bus architecture of the MEA and tested under different dynamic conditions using three different control techniques. Firstly, a direct estimation based technique is proposed for determining the phase shift of the PSHPBD converter (used as a charger for the batteries). Secondly, a current mode control for the single phase shift control technique for the PSHPBD converter is proposed. The proposed control technique has an inherent capability of current limit. Thirdly, a peak current controller is proposed which provides a limit on the peak current and provides a fast transient response. The advantages of using the proposed peak current control in avoiding the saturation of the transformer core is also reported in the work. In integration, finally the study and analysis of zero current switching (ZCS) inverter for integration of BESS to ESG set in MEA “Aircraft X” is discussed. Furthermore, a control method which can improve the life of the BESS in MEA is discussed in which three different modes of control based on the load estimation are given. Depending upon the load on the BESS, the corresponding control structure is implemented/activated leading to reduction in RMS current flow in the charger circuit and average current flow in the battery system thereby reducing the losses and temperature rise in the BESS, which eventually leads to an increase in the life of the BESS. Finally, a multi-functional battery charger (MFBC) is proposed for integration of BESS to the power distribution network in MEA. A novel control method is devised for suitable transfer and sharing of power between different buses. Based on the placement and applicability of the MFBC, many new dc power distribution architectures for MEA are proposed wherein energy storage system can be integrated into aircraft network leading towards fault tolerant operation and weight optimization in the MEA.