Explore the vast range of titles available.

A spacecraft power processing unit (PPU) is utilized to convert power from solar arrays or electric batteries to the payload, including electric propulsion, communication equipment, and scientific instruments. Currently, a high-voltage converter is widely applied to the spacecraft PPU to improve power density and save launch weight. However, the high voltage level poses challenges such as high step-down ratios and high power losses. To achieve less conduction loss, a SiC-based T-type three-level (TL) LLC resonant converter is proposed. To further broaden the gain range and achieve high step-down ratios, a variable frequency and adjustable phase-shift (VFAPS) modulation scheme is proposed. Meanwhile, the steady-state time-domain model is established to elaborate the operation principles and boundary conditions for soft switching. Furthermore, the optimal resonant element design considerations have been elaborated to achieve wider gain range and facilitate easier soft switching. Furthermore, the numerical solutions for switching frequency and phase shift (PS) angle under each specific input could be figured out. Finally, the effectiveness of this theoretical analysis is demonstrated via a 500-W experimental prototype with 650∼950-V input and constant output of 48-V/11-A.

This paper proposes a megawatt (MW)-scale high-voltage (HV) electrical power-conversion element for large-spacecraft electric propulsion (EP) systems. The proposed scheme is intended for long-term and crewed missions, and it is driven by a nuclear electric propulsion (NEP) that acts as a heat source. The scheme includes (i) A two-rotor generator (TRG), (ii) A rectification stage, and (iii) An isolated dual output DC-DC (iDC2) converter. The TRG is a high-reliability electric machine with two rotors, a permanent magnet rotor (PMR), and a wound field rotor (WFR). The PMR has a fixed flux and hence back-EMF, while the back-EMF due to the WFR is controlled by injecting a direct current (DC) into the WFR winding. The total TRG output voltage, which is the sum of voltages due to the PMR and WFR, is controlled over a prescribed region of spacecraft operation. The output of the TRG is rectified and connected to the input of the iDC2 converter. The iDC2 converter uses a three-winding transformer, where the primary winding is fed from the rectified output of TRG, the secondary winding processes the propulsion power to an electric thruster via a high-voltage DC (HVDC) link and a tertiary winding that is connected to the spacecraft’s low-voltage DC (LVDC) power system. Three controllers are proposed for the system: an HVDC voltage controller, an HVDC current controller that controls the voltage and current processed to the thruster, and an LVDC controller that adjusts the current to the LVDC system. Detailed analytical models for the TRG, iDC2 converter, and controllers are developed and verified via simulations under different conditions. The analytical studies are further validated via results from a laboratory prototype.

The composite bus system is a new type of satellite power system, which is suitable for high-power SAR satellites. The bus converter in the compound bus system is very important. This paper analyzes the design requirements of the bus converter, selects PSM LLC as the topology of the bus converter, and verifies the rationality of the scheme through simulation.

Spacecraft operate in a harsh environment, with a complex space electromagnetic environment, diverse loads, and high requirements for communication performance and reliability. As an optional communication technology within spacecraft, power line communication (PLC), in addition to facing the above problems, must first solve the problem of modeling channel characteristics of PLC in DC-powered environments. To this end, this paper proposes a channel analysis and modeling method for spacecraft DC PLC, gives the calculation and analysis process of W-element distribution parameters based on transmission line theory, establishes a specific cable and branch connector simulation model based on CST, and carries out a simulation study on the whole link of the cable, branch connector, and the channel of the DC PLC system. According to the simulation system, the corresponding experimental verification platform is built. The simulation and test results verify the effectiveness and accuracy of the proposed modeling method.

A spacecraft power distribution system is a multi-parameter, strongly coupled complex system, and the safety of a spacecraft power distribution system directly affects the working status of the spacecraft. In order to achieve better research results in the safety analysis system of spacecraft energy systems, it is very necessary to establish an assessment method for the safety of spacecraft power distribution systems. Previous research mainly focuses on the safety assessment of power systems or aircraft power systems and has not formed a relatively complete assessment method for the safety of spacecraft power distribution systems. This paper proposes a comprehensive evaluation method for the safety of spacecraft power distribution systems. It combines objective and subjective evaluation methods to comprehensively evaluate the safety of spacecraft power distribution systems. The objective parameters of the spacecraft power distribution system are processed and objectively evaluated; Subjective evaluations are obtained using analytic hierarchy. The proposed method can reasonably evaluate the safety of spacecraft power distribution systems and assist in the selection of different spacecraft power distribution systems.

On 18 November 2023, SpaceX launched the Starship, the tallest and the most powerful rocket ever built. The Super Heavy engine separated from the Starship spacecraft and exploded at 90 km of altitude, while the main core Starship continued to rise up to 149 km and exploded after ∼8 min of flight. In this work, we used data from ground‐based GNSS receivers and we analyzed total electron content (TEC) response to the Starship flight and the two explosions. For the first time, we observed large‐distance northward propagation of intensive 2,000 km V‐shaped ionospheric disturbances from the rocket trajectory. The observed perturbations, most likely, represent shock waves propagating with the cone angle of ∼14° on the North and ∼7° on the South against the flight track that corresponds to the Mach angle of the shock waves in the lower atmosphere. The Starship explosion also produced a non‐chemical depletion in the ionospheric TEC. Plain Language Summary On 18 November 2023, SpaceX launched the Starship, the tallest and the most powerful rocket ever built. About 2 min and 40 s after the liftoff, the Super Heavy engine separated from the Starship spacecraft and exploded at an altitude of 90 km. The main core Starship continued to rise to 149 km and exploded as well. The rocket launch and explosion produced an unexpected response in the ionosphere—the ionized part of the Earth's atmosphere. The Starship flew at a velocity, exceeding the local sound speed, and generated cone‐like atmospheric shock‐acoustic waves. Most unexpectedly, the observed disturbances represented long and intensive multi‐oscillation wave structures that propagated northward, which is unusual for disturbances driven by a rocket launch. The Starship explosion also generated a large‐amplitude total electron content depletion that could have been reinforced by the impact of the spacecraft's fuel exhaust in the lower atmosphere. This study appears to be the first‐time detection of a non‐chemical ionospheric hole produced by a man‐made explosion. Key Points The 18 November 2023 Starship flight and explosions generated large‐scale multi‐oscillation supersonic conic waves in the ionosphere The cone angle of the V‐shaped ionospheric disturbances corresponds to the Mach angle of shock waves propagating in the lower ionosphere The shock waves from the Starship explosion caused a depletion in total electron content (TEC)

With the development of large spacecraft and satellite payloads, the power supply required is increasing, especially in space solar power stations, and the large flexible solar wing is an inevitable trend in the future. Due to their light weight and large area, flexible solar wings may face thermally induced vibration during orbit operation, which will lead to the degradation of spacecraft performance and even structural damage. In this paper, the test method of a flexible solar wing simulator is established, and the thermally induced vibration test of a flexible solar cell wing simulator is carried out. And vibration amplitude value is 25 and 9 microns with a frequency 0.56Hz.The characteristic of thermally induced vibration of flexible solar cell wing is given, which provides a basis for its thermally induced vibration suppression.

Jupiter's moon Europa contains a subsurface ocean whose presence is inferred from magnetic field measurements, the interpretation of which depends on knowledge of Europa's local plasma environment. A recent Juno spacecraft flyby returned new observations of plasma electrons with unprecedented resolution. Specifically, powerful magnetic field‐aligned electron beams were discovered near Europa. These beams, with energies from ∼30 to ∼300 eV, locally enhance electron‐impact‐excited emissions and ionization in Europa's atmosphere by more than a factor three over the local space environment, and are associated with large jumps of the magnetic fields. The beams therefore play an essential role in shaping Europa's plasma and magnetic field environment and thus need to be accounted for electromagnetic sounding of Europa's ocean and plume detection by future missions such as JUICE and Europa Clipper. Plain Language Summary A recent Juno spacecraft close flyby of Jupiter's moon Europa revealed the presence of powerful electrons beams. Based on previous observations and modeling of electron beams at the moon Io, such beams were not expected to be observed so close to Europa. Overall, the proximity of the beams to Europa indicates that the acceleration of these electrons takes place much closer to Europa than anticipated and that these beams, therefore, stem from a new and previously unknown acceleration mechanism. The beams are predicted to have an outsized influence on the ionization of the constituents of Europa's tenuous atmosphere and are accompanied with large magnetic field perturbations. Hence, these electron beams are an important ionization source that modify the moon's ionosphere, the electric current systems, and the magnetic field environment. In particular, the presence of electron beams will affect plasma conditions that are used to infer the extent of a subsurface ocean via the magnetic induction signal. These beams significantly impact the space plasma environment around Europa which needs to be accounted for by future missions such as ESA's (European Space Agency) JUICE (Jupiter Icy Moons Explorer) and NASA's (National Aeronautics and Space Administration) Europa Clipper mission. Key Points Powerful electron beams that significantly shape Europa's space environment are discovered during a Juno flyby The beams enhance electron‐impact‐excited emissions in Europa's atmosphere and are associated with large jumps of the magnetic fields The beams' proximity to Europa and their pitch angle distribution constrain the source acceleration to be near or within the plasma disk

In this paper, the relevance of the use of mathematical modeling with the electromagnetic compatibility requirements in solving problems of space engineering is shown. Approbation of the software prototype developed in TUSUR is performed by designing the power supply network elements of a spacecraft. The mathematical and software features of the prototype are described.

To solve the problems of high functional requirement, complex structure and poor versatility of manned spacecraft power supply and distribution ground test equipment, a modular universal ground power supply system (PSS) is designed. The functions of the system cover the functions of wire instruction opening, parameter measurement and state detection of the interface between the manned spacecraft and the ground equipment, and realize the universal and extensible configuration. This paper provides a general architecture design method for complex manned spacecraft power supply and distribution ground test system.