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This article explores the application of the Legendre pseudospectral method to spacecraft orbital transfer with finite thrust optimization problem. Firstly, the model of the orbital transfer optimization control problem was established, while equations of motion were simplified based on some hypotheses. The performance was optimized to minimize the cumulative fuel consumption. The control variable was the thrust attack angle, and terminal state variable constraints included path angle, altitude, and velocity constraints. Then, the optimal control problem was transformed into a nonlinear programming problem (NLP) using the Legendre pseudospectral method. The dynamic optimization problem was transformed into a static parameter optimization problem. The state variables and control variables were selected as the optimal parameters at all collocation nodes. Lastly, the parameter optimization problem was solved using the SNOPT (Sparse Nonlinear Optimizer) software package. The SNOPT software package shows high convergence for a nonlinear programming problem. During the simulation, it was noted that the Legendre pseudospectral method is not sensitive to orbital transfer initial conditions. It was also observed that the optimal solutions of the orbital transfer optimization problem are fairly good in robustness. Therefore, the Legendre pseudospectral method is a viable approach to the spacecraft orbital transfer with a finite thrust optimization problem. The orbit optimization method proposed in this paper can also provide reference and guidance for solving other interplanetary orbital transfer optimization problems. [ABSTRACT FROM AUTHOR]

The significance of reliable energy storage systems in spacecraft applications cannot be overstated, since they play a vital role in ensuring continuous power supply and prolonged mission durations. This research deals with the modeling of a hybrid multi-mission radioisotope thermoelectric generator (MMRTG)-lithium-ion (Li-ion) battery integrated energy storage system for spacecraft applications to combine the RTGs' long lifespan and reliability benefits alongside the Li-ion battery's rechargeability and high energy density to achieve a single energy unit. The investigation's main problem was exploring a power unit that improves the limitations of MMRTG and Li-ion batteries to achieve a highly efficient and reliable power supply for autonomous systems, such as a spacecraft. The proposed hybrid system comprises a 110 W/32 V RTG and a 3.6 V/43 Ah Li-ion battery connected to a DC motor through power converters. Results demonstrate the potential of the adopted hybrid energy system in improving the efficiency, reliability, and mission duration of spacecraft missions. The assessment of the hybrid energy system under various load conditions shows that the highest power peak of 3500 W was achieved at a load resistance of 1 Ω. Furthermore, the results show that the hybrid energy system output voltage at temperatures of 253 °K and 293 °K are relatively equal. However, the power cycle was wider and required a long time before dropping. [ABSTRACT FROM AUTHOR]

Recent advancements in space technology and reduced launching cost led companies, defence and government organisations to turn their attention to low Earth orbit (LEO) and very low Earth orbit (VLEO) satellites, for they offer significant advantages over other types of spacecraft and present an attractive solution for observation, communication and other tasks. However, keeping satellites in LEO and VLEO presents a unique set of challenges, in addition to those typically associated with exposure to space environment such as damage from space debris, thermal fluctuations, radiation and thermal management in vacuum. The structural and functional elements of LEO and especially VLEO satellites are significantly affected by residual atmosphere and, in particular, atomic oxygen (AO). At VLEO, the remaining atmosphere is dense enough to create significant drag and quicky de-orbit satellites; thus, thrusters are needed to keep them on a stable orbit. Atomic oxygen-induced material erosion is another key challenge to overcome during the design phase of LEO and VLEO spacecraft. This review covered the corrosion interactions between the satellites and the low orbit environment, and how it can be minimised through the use of carbon-based nanomaterials and their composites. The review also discussed key mechanisms and challenges underpinning material design and fabrication, and it outlined the current research in this area. [ABSTRACT FROM AUTHOR]

A new method is proposed for determining the electron density in rarefied plasma, based on simultaneous measurements of the Interball-2 satellite potential using IESP-2 (electric field instrument) and KM-7 (electron temperature sensor) probe devices. This makes it possible to estimate the photoelectron current density based on a procedure proposed earlier by the authors of this study. The electron concentration was determined only for the positive potential of the spacecraft. The balance equations for the satellite and the probe between the currents of the surrounding plasma electrons and photoelectrons emitted by the illuminated surface were compiled. In the magnetosphere, to bring the probe potential to the potential of the surrounding plasma, a bias current is directed into the probe, which was taken into account in the current balance equation for the probe. The electron energy used in the calculations was kTe = 1 eV. We analyzed data from ~350 orbits in the auroral region of the magnetosphere at altitudes of 2–3 RE from October 1996 to March 1998 during the period of low solar activity at the beginning of the 23rd cycle. Examples of the calculated electron density are given, which is in the range of 1–30 cm–3. [ABSTRACT FROM AUTHOR]

The aim of this paper is to provide information on the radial dependence of turbulence quantities as input for ab initio numerical cosmic-ray modulation models. We present an evaluation and analysis of 48-s resolution Voyager 1 data for the normal component of the magnetic field. These data were preprocessed to fill in data gaps and filter out spurious peaks. The effectiveness of this process is illustrated for a sample data set, for which we show that the underlying spectrum is indeed Kolmogorov. We use a second-order structure function to find correlation times, and also the spectral index and spectral level in the inertial range from expressions derived by Huang et al. (2010). The usefulness of the second-order structure function for data with significant gaps is illustrated by using synthetic data. Gaps up to 68% cause changes of less than 5% in the quantities that we calculate. We show that beyond about 20 au there are significant changes in the behaviour of the magnetic field magnitude and the inertial range spectral index, neither of which has a reasonable explanation. We find that the ratio δ B N / B N remains almost constant out to 20 au, that the correlation times is consistent with results from other studies, and that the spectral index in the inertial range is consistent with the Kolmogorov value. A new result is that we could infer that the break between the energy- and the inertial range has a radial dependence that is consistent with that of the correlation scale. [ABSTRACT FROM AUTHOR]

Currently, managing a group of satellites or robot manipulators requires coordinating their motion and work in a cooperative way to complete complex tasks. The attitude motion coordination and synchronization problems are challenging since attitude motion evolves in non-Euclidean spaces. Moreover, the equation of motions of the rigid body are highly nonlinear. This paper studies the attitude synchronization problem of a group of fully actuated rigid bodies over a directed communication topology. To design the synchronization control law, we exploit the cascade structure of the rigid body's kinematic and dynamic models. First, we propose a kinematic control law that induces attitude synchronization. As a second step, an angular velocity-tracking control law is designed for the dynamic subsystem. We use the exponential coordinates of rotation to describe the body's attitude. Such coordinates are a natural and minimal parametrization of rotation matrices which almost describe every rotation on the Special Orthogonal group S O (3) . We provide simulation results to show the performance of the proposed synchronization controller. [ABSTRACT FROM AUTHOR]

Summary: This paper investigates the attitude tracking control problem of rigid spacecraft with inertia parameter identification. Based on the relative attitude and angular velocity error dynamics, a basic adaptive backstepping based attitude tracking control scheme is firstly designed such that asymptotic attitude tracking can be achieved. However, the parameter identification error cannot decay to zero if the persistent excitation (PE) condition is not satisfied. To solve this issue, a filtering‐based concurrent learning adaptive backstepping control scheme is then proposed, by incorporating torque filtering technique with concurrent learning technique. A more mild rank condition, which consists of some collectable historical data, is provided to guarantee the convergence of parameter identification error. In addition, a valid data collection algorithm is given. It should be mentioned that a distinctive feature of the proposed filtering‐based concurrent learning adaptive control scheme is that the convergence rate of attitude tracking errors can be improved from asymptotical to exponential. Finally, simulation results are provided to illustrate the effectiveness of the proposed control schemes. [ABSTRACT FROM AUTHOR]

The object of this study is the technology of ensuring magnetic cleanliness when applied to small spacecraft (SC) with overall dimensions of ~1 m in terms of minimizing the effect of magnetic interference on the on-board magnetometer. As part of solving the general problem of increasing the reliability of spacecraft, the task of improving the technology of magnetic cleanliness is considered. It is shown that the use of the multi-dipole model for the calculation of ~50 nT magnetic interference is limited in the design of small space vehicles that have an increased installation density. The expediency of using a model of its spherical harmonics instead of a multi-dipole model to represent the magnetic field of satellite components was theoretically justified. For the practical application of the model, the budgeting of the projections of the magnetic induction of hindrance to the onboard magnetometer according to the experimentally determined coefficients of the dipole, quadrupole and octupole spherical harmonics is proposed. An algorithm for calculating the coordinates of the location and projections of the magnetic moment of the dipole source inside the satellite based on the results of measuring the coefficients of spherical harmonics of its components is proposed. The possibility of representing the coefficients of the spherical harmonics of the magnetic field of the entire satellite in the form of the sum of the listed corresponding coefficients obtained during field measurements of the components is theoretically justified. It is proposed to use the difference between the calculated and measured values of the corresponding coefficients as a criterion for the quality of the work to ensure magnetic cleanliness. If the proposed procedure is used in the space industry, the quality of ensuring the magnetic cleanliness of small space vehicles can be increased, thereby improving the reliability of their operation in orbit [ABSTRACT FROM AUTHOR]

• A novel linear model-based controller is designed for the CEACS attitude regulation. • The Takagi-Sugeno fuzzy model with a parallel distribution approach has been introduced. • The Fuzzy-MPC controller achieves the desired CEACS attitude pointing accuracy. • The proposed controller effectively handles the CEACS actuator constraints. • The CEACS attitude actuation consumes a lower on-board power. A Combined Energy and Attitude Control System (CEACS) is a dual system in which flywheels are used as energy storage and attitude control devices. This work is a progress on CEACS for small satellites to improve the attitude accuracy. In this maiden work, the Fuzzy–MPC controller is introduced to regulate the CEACS attitude with both linear and nonlinear initial angles, and in the presence of actuator constraints. The nonlinear attitude model is transformed into a set of linear subsystems using the Takagi–Sugeno fuzzy modeling approach. The subsystem is designed with a MPC controller and a total control action is summed up using the parallel distributed compensation approach. The numerical results have been analyzed in terms of attitude pointing accuracies, actuator constraints, and energy consumed by the actuators. A performance comparison between the Fuzzy–MPC and Fuzzy–LQR controllers has also been done. The results validate that the Fuzzy-MPC controller achieves the desired CEACS attitude pointing accuracy of 0. 0010 ° with a zero steady-state error and keeps the control torques within the actuator constraints. On the other hand, the Fuzzy–LQR controller gives some measurable steady-state error and achieves the CEACS attitude pointing accuracy of 0. 0012 °. Therefore, the Fuzzy–MPC based CEACS architecture not only achieves the desired pointing accuracies but also produces ten times smaller control torques than the Fuzzy–LQR controller, which result in a lower onboard power consumption. [ABSTRACT FROM AUTHOR]

Thermal analysis of spacecraft involves complex heat flux outside space and heat transfer calculation of equipment inside space, which is the basic guarantee of its long-term stable operation. In this paper, the core module of the Chinese Space Station is taken as an example to conduct hybrid modeling of its complex radiation and heat dissipation network. The external heat flux received by spacecraft is affected by many factors, such as its attitude, orbit parameters, and solar orientation. First, based on the defined reference coordinate system, the whole station is divided into eight quadrants. Then, an intelligent occlusion algorithm is proposed to solve the complex occlusion relationship according to the space quadrant method and fuzzy logic, and then the external heat flux is obtained. Finally, the dynamic response of the station under different conditions is studied by combining traditional mechanism modeling and an intelligent algorithm, and the maximum heat dissipation capacity was obtained. The results show that the temperature control valve opening will seriously affect the main flow of the loop, and the maximum difference can be 0.1 kg/s. In addition, the external heat flux changes periodically, so that the flow of each branch and the valve opening also changes periodically. In the sunlight area, the radiator receiving the external heat flux is about 77,000–90,000 W, and the maximum heat dissipation capacity is 12,000 W. Those values are 28,000 W and 22,000 W, respectively, in the shadowed area. Moreover, the hybrid modeling method proposed in this paper has an important reference value for spacecraft thermal analysis and research in other fields. [ABSTRACT FROM AUTHOR]

This paper deals with the nonfragile fuzzy tracking control technique of a spacecraft subject to external disturbances and control gain uncertainties. First, the Takagi–Sugeno (T–S) fuzzy model of the tracking error dynamics is newly derived. Unlike the existing studies, we assume that the controller output is norm bounded and its gain matrix is perturbed because of the aging of the actuator. Next, a fuzzy disturbance observer (DOB) estimating the external disturbance is constructed to enhance the robustness of the control system. Then, a fuzzy controller feeding back the estimated disturbance is designed to attenuate external disturbances. In addition, we establish the linear matrix inequality (LMI) conditions, by which both the H ∞ disturbance attenuating performance and the asymptotic stabilization are guaranteed. From the solution to the LMI conditions, we obtain the controller and DOB gain matrices. Finally, a simulation example validates the disturbance estimation and tracking performance of the proposed method. [ABSTRACT FROM AUTHOR]

This is a high-level study into the validation of a unified electric power and communications architecture for modern spacecraft systems. We leverage from blending the European Cooperation for Space Standardisation (ECSS) SpaceWire and the IEEE 802.3 Power over Ethernet (PoE) standards for a common onboard power and communications interface. The focus of the work has been communication performance. The requirements from the protocols of both standards were analysed and combined to create a full protocol stack for SpaceWire PoE. The stack is implemented on an embedded platform represented as a proof-of-concept SpaceWire PoE system. The results of functional testing demonstrate SpaceWire PoE as a valid solution that meets the requirements of both protocols. For the 100 Mb/s links, the SpaceWire PoE throughput is on a par (97.39%) with Ethernet (97.52%) and SpaceWire (99.2%). Shared communication and power architecture is intended for decentralised operation achieving greater autonomy of the integrated subsystems. This potentially may lead to standardisation of electrical, mechanical and communication interfaces across spacecraft subsystem manufacturers, easing harness complexity and routing during the assembly, integration and test (AIT) phase of the spacecraft as well as saving mass and launch cost. [ABSTRACT FROM AUTHOR]

This paper reports the implementation of two critical technologies used in light detection and ranging for space applications: (1) a microchip Q-switched laser breadboard; (2) a breadboard of an indium gallium arsenide avalanche photodiode working at 292 K with high reverse polarization voltages. Microchip Q-switched lasers are small solid-state back-pumped lasers that can generate high-energy short pulses. The implemented breadboard used an erbium and ytterbium co-doped phosphate glass, a Co:Spinel crystal with 98% initial transparency, and an output coupler with 98% reflectivity. For the sensor test, a system for simultaneous operation in vacuum and a wide range of temperatures was developed. Avalanche photodiodes are reverse-polarized photodiodes with high internal gain due to their multiple layer composition, capable of building up high values of photocurrent from small optical signals by exploiting the avalanche breakdown effects. The test avalanche photodetector was assembled to be operated in two modes: linear and Geiger mode. The produced photocurrent was measured by using: (1) a passive quenching circuit; (2) a transimpedance amplifier circuit. These two technologies are important for mobile light detection and ranging applications due to their low mass and high efficiencies. The paper describes the breadboard's implementation methods and sensor characterization at low and room temperatures with high bias voltages (beyond breakdown voltage). [ABSTRACT FROM AUTHOR]

Fungi biofilms have been found growing on spacecraft surfaces such as windows, piping, cables, etc. The contamination of these surfaces with fungi, although undesirable, is highly difficult to avoid. While several biofilm forming species, including Penicillium rubens, have been identified in spacecraft, the effect of microgravity on fungal biofilm formation is unknown. This study sent seven material surfaces (Stainless Steel 316, Aluminum Alloy, Titanium Alloy, Carbon Fiber, Quartz, Silicone, and Nanograss) inoculated with spores of P. rubens to the International Space Station and allowed biofilms to form for 10, 15, and 20 days to understand the effects of microgravity on biofilm morphology and growth. In general, microgravity did not induce changes in the shape of biofilms, nor did it affect growth in terms of biomass, thickness, and surface area coverage. However, microgravity increased or decreased biofilm formation in some cases, and this was incubation-time- and material-dependent. Nanograss was the material with significantly less biofilm formation, both in microgravity and on Earth, and it could potentially be interfering with hyphal adhesion and/or spore germination. Additionally, a decrease in biofilm formation at 20 days, potentially due to nutrient depletion, was seen in some space and Earth samples and was material-dependent. [ABSTRACT FROM AUTHOR]

Wind turbines have improved in terms of reliability over the past ten years, but they still experience crucial and expensive tribological issues that shorten their intended lifespans. Most turbines require significant repairs or complete replacements of the tribo components in the 5- to 7-year range. The pricey gearboxes have high repair rates, which greatly raises the operating costs for wind turbines and the cost of wind energy. Currently, various bearings in the turbine are lubricated by automatic lubrication systems with the necessary oil filters to ensure cleanliness. The commonly used lubrication methods for wind turbine bearings are centralized lubrication, oil-air lubrication, oil mist lubrication, and oil bath lubrication. There are similarities between wind turbines and spacecraft-moving mechanical systems. Most spacecraft are designed to work for decades in a remote environment without periodic maintenance. Spacecraft use efficient lubrication systems for their moving mechanical systems to provide the required lubricant to the bearings for the required life. Depending on the lubrication requirements, proven active and passive lubrication systems are employed in spacecraft. This paper reviews the lubrication issues of the wind turbine in comparison with the spacecraft and proposes a few lubrication systems that are successfully performing in spacecraft, intending to improve the reliability of wind turbine tribo-systems. [ABSTRACT FROM AUTHOR]

Context is due to the need to analyze the electromagnetic accessibility of terrestrial radio radiation sources at long distances, which is typical for the operation of a spacecraft of radio technical intelligence with a passive synthesis of the antenna aperture. Objective is to calculate the probability indicators of detecting the fact of signal emission of ground-based radar stations using radio-technical intelligence installed on a space carrier. Method. Analytical calculation of the correct detection probability of the signal of the radio radiation sources by the spacecraft of radio technical intelligence based on the determination of the signal-to-noise ratio at the input of the detection device. The analysis of the technical characteristics of the group of ground radars made it possible to calculate the probability of correct detection only for the virtual radar, which generates the minimum number of pulses with the minimum energy during reconnaissance among all the analyzed radars. Results. The fulfillment of the conditions of electromagnetic accessibility is checked and the correct detection probability of the signals of modern radar stations by the space radio-technical intelligence system with a passive synthesis of the antenna aperture was calculated. Conclusion. Proposed use of the correct detection probability of the signal as the resulting indicator of electromagnetic accessibility of the radio radiation source. In the example radar stations of the air defense proved, that the proposed parameters of the space radio technical intelligence system with a passive synthesis of the antenna aperture provide the values of the correct detection probability of the signal, which are quite acceptable for modern monitoring systems. [ABSTRACT FROM AUTHOR]

The paper presents the results of calculating the reliability of the cables of the spacecraft onboard harness made of aluminum wires for space application. The calculation was based on the results of qualification tests of type representatives of cables with soldered and crimped connections of aluminum wires. Similar copper wire connections were also tested to compare results. The results of the performed calculation confirmed the specified requirements for the reliability of cables with aluminum wires. [ABSTRACT FROM AUTHOR]

This work aims at addressing whether a catastrophic failure of an entry, descent, and landing event of a Multimission Radioisotope Thermoelectric Generator-based lander could embed the heat sources into the martian subsurface and create a local environment that (1) would temporarily satisfy the conditions for a martian Special Region and (2) could establish a transport mechanism through which introduced terrestrial organisms could be mobilized to naturally occurring Special Regions elsewhere on Mars. Two models were run, a primary model by researchers at the Lawrence Berkeley National Laboratory and a secondary model by researchers at the Jet Propulsion Laboratory, both of which were based on selected starting conditions for various surface composition cases that establish the worst-case scenario, including geological data collected by the Mars Science Laboratory at Gale Crater. The summary outputs of both modeling efforts showed similar results: that the introduction of the modeled heat source could temporarily create the conditions established for a Special Region, but that there would be no transport mechanism by which an introduced terrestrial microbe, even if it was active during the temporarily induced Special Region conditions, could be transported to a naturally occurring Special Region of Mars.

The article presents the results of the analysis of heating of printed circuit conductors of printed circuit boards mounted on a metal base, on-board device of spacecraft when pulsed current flows through them. Thermal transient and impulse characteristics of a system consisting of a printed conductor and a printed circuit board are considered. The analysis of the dependence of the temperature rise rate of printed conductors and the duration of transient processes on the linear dimensions of the system elements is given. It is noted in the article that the time of transient processes most strongly depends on the thickness of the layer of insulating materials. All the results were obtained in the process of modeling transients in the ANSYS 2019 R1 Transient Thermal module. Practical recommendations on the consideration of thermal transients in printed conductors in the design of printed circuit boards are given.

The authors describe an algorithm that allows calculating the parameters of maneuvers performed on several turns by the low-thrust engine, which ensure the flight of the active spacecraft to the specified vicinity of the target space object. The movement takes place in the vicinity of a circular orbit. Linearized equations of motion are used in solving the problem. The influence of the non-centrality of the gravitational field and the atmosphere is not taken into account. The determination of maneuver parameters takes place in three stages. At the first and third stages, the parameters of the pulse transition and the transition performed by the low-thrust engine, are determined analytically. At the second stage, the distribution of maneuvering between turns, which provides a solution to the meeting problem, is carried out by iterating over one variable. This method of solving the problem provides simplicity and high reliability of determining the parameters of maneuvers, which allows it to be used on board the spacecraft. The paper investigates the dependence of the total characteristic speed of solving the meeting problem on the number of turns of the flight and the magnitude of the engine thrust.

A design feature of the construction of electric heaters for non-pressurized spacecraft using a heating element in the form of a tape of a material with high resistivity, located on a flat base (electrical insulating substrate) in the form of a snake with the organization of the necessary gaps between the tapes, is revealed. A mathematical model has been developed for calculating the main design parameters of electric heaters: power, mass, heated surface area. The nomenclature of efficiency indicators used as optimization criteria has been formed: power utilization factor, filling factor of an electrically insulating substrate tape, surface density, energy-mass efficiency, thermal efficiency of heating, target efficiency. The methodological principles for ranking the requirements for the parameters of electric heaters are determined using the selected optimization criteria: the preferred use of tape materials with high resistivity and minimum thickness. A method is formulated for step-by-step calculation of the parameters of electric heaters in the presence of restrictions on the maximum heat flux density determined from the conditions for the use of electric heaters. General recommendations have been developed for choosing the parameters of electric heaters: use increased voltage, minimize the thickness of the heating element tape and the width of the gap between the tapes, determine the minimum value of the tape length for the maximum heat flux density at a given voltage rating.

The gear assembly of the rotational motion drive of the spacecraft system is considered. Such gear assemblies are one of the main elements of spacecraft systems with the use of large-sized structures that are folded in the transport position. Due to the development of computer technology, it becomes possible to create more accurate simulation models of such structures. The aim of the work is to create a simulation model that allows to analyze emerging loads on a wave transmission, which is part of the gear assembly of the spacecraft drive. Section one discusses the process of creating a finite element model of the drive. One of the features of creating a finite element model is a method of taking into account the preliminary stress-strain state in the structural elements of the wave transmission. Section two analyzes the data obtained during the calculation. The obtained equivalent stresses on the structural parts, contact stresses in the gear engagement are investigated, the torsional rigidity of the gearbox elements is determined, various tooth positions are considered during movement. The influence of loads applied to the output link of the gearbox on the surface of the gears is investigated. The obtained simulation model allows us to consider similar gear assemblies in which wave transmissions are used.

The design and optimization of re-entry spacecraft or its subsystems is a multidisciplinary or multiobjective optimization problem by nature. Multidisciplinary design optimization (MDO) focuses on using numerical optimization in designing systems with several subsystems or disciplines that have interactions and independent actions. In the present paper, the system-level optimizer, trajectory, geometry and shape, aerodynamics, and aerothermodynamics differential equations, are converted to algebraic equations using the Radau pseudospectral method (RPM) since a spacecraft is a nonlinear, extensive, and sparse system. The solution to the problem with the help of MDO is reached by iterating all the disciplines together; one can simultaneously enhance the design, decrease the time and cost of the entire design cycle, and minimize the structural mass of a re-entry spacecraft. Considering various methods presented in earlier research works, a combined and innovative all-at-once (AAO), RPM-based MDO method, including the key subsystems in the design process of a re-entry capsule-shape spacecraft with a low lift-to-drag ratio (L/D), is presented. Considering the applicable state and control variables, various constraints, and parameters applied to several geometric shapes of a blunt capsule and using Apollo’s aerodynamic and aerothermodynamic coefficients, the optimized dimensions for a re-entry spacecraft are presented. The introduced optimization scheme led to a 17% mass reduction compared to the original mass of the Apollo vehicle. Fast computing and simplified models are used together in this method to analyze a wide range of vehicle shapes and entry types during conceptual design.

This paper presents an analysis of various designs of transformable membrane reflector antennas for spacecrafts, considers the dependence of the surface accuracy on the operating frequency of the antenna and the requirements imposed on deviations of the antenna surface from the paraboloid. A classification of membrane transformable antennas is presented depending on the presence or absence of supporting elements and stiffeners. Structural schemes for opening membrane antennas are described, from the classical umbrella opening scheme to original schemes for winding ribs on a central hub, tension supporting elements, winding paraboloid segments on a central hub, and folding the reflector surface using origami methods. The design of an offset membrane transformable antenna is considered. Modern materials for the manufacture of reflectors of membrane antennas, such as polymer composite materials with triaxial weaving of carbon fabric with a silicone matrix, polyamide and polyimide films with and without reinforcement, are considered. Materials with shape memory for opening membrane antennas are considered. Calculations of structures for natural vibration frequencies, calculations for bending of folds are described. A description of ground-based experimental testing, in particular, for the disclosure of various types of membrane antennas, was carried out.

The paper presents the results of computational study of sections of spatial truss structures. They can form large supporting structures on spacecraft to place onboard equipment. The calculations were carried out by the finite element method in FEMAP with Nastran. The chosen section, the object of study, is a set of straight rods rigidly connected at the nodes in such a way that the cross section of the truss structure is a triangle. Calculation models, procedures for calculation and results analysis are presented. The purpose of the calculation is to determine how the relative position of the rods in the structural scheme affects the mechanical properties of the structure. The main criterion for strength estimating was the magnitude of the stresses derived by load. Stiffness was determined by the value of the first natural frequency. The sequential addition of rods and varying their connection allow modifying the structural schemes of truss structures. Next, the mechanical properties of the structure which effected by made modifications were evaluated again. Thus, six structural schemes of sections of the truss structure, obtained from the results of the study, have been developed. At the same time, the mass of the section, its shape and dimensions, the material and shape of the rods section as well as the initial and boundary conditions, remained unchanged. Based on the results of the analysis, someone can notice that each structure has unique mechanical characteristics. Thus, the paper gives recommendations for choosing a specific structural scheme of the section, depending on the required operating conditions and acceptable manufacturing technology. So, the criteria for choosing one of the above various sections of a truss structure or the principle of its building can be the complexity of manufacturing, maximum stiffness and strength, or minimum displacement. Structural schemes of sections with the highest strength and capability have been selected for use as part of spatial truss rods of spacecraft.

Remote sensing satellites require very accurate pointing to specific locations of interest with high resolution and small latency. Therefore, the space imaging systems attempted to achieve the possible highest ground resolution with minimum cost as possible. In this paper, design process of an electro-optical imaging payload is described considering the limited weight and small volume in addition to minimum power consumptions. The most promising prospects for high resolution imaging are connected with passive optical sub-systems, especially push broom one. For design a multiband Space remote sensing payload with 4°7′ field of view that operates in multi spectral bands and a panchromatic band, it requires the high speed Schmidt–cassegrain objective. The most promising prospects for high resolution imaging are connected with passive optical sub-systems, especially push broom one. The Schmidt-Cassegrain objective is a catadioptric objective that combines a Cassegrain reflector’s optical path with a Schmidt corrector plate to make a compact objective that uses simple spherical surfaces only. Because of the limited weight of the camera (22.2 kg), the multispectral consists of only one single optical system large F-Number (F5) and provided with spectrum splitter.In addition, the key points in design of such types of cameras were listed as well as clarifying the optical design of such objective.

Spacecraft or the instruments are directed to certain directions or points in order to successfully complete their mission. Solar sensors are commonly used as they are capable of being used in planetary or interplanetary space missions in the solar system. Mostly, these instrument sense not only the radiation from the sun but also the albedo of any nearby object. These measurements are known to cause erroneous outputs in the attitude and control system when used without considering the albedo model. In this study, the effect of the lunar albedo on the sun sensor and sun direction estimation is evaluated under different simulation conditions.

The concept of solar-hydrogen systems for spacecraft, orbital stations, lunar and Martian bases is currently receiving a new impetus. The supply of solar energy to energy receivers aboard space vehicles is limited. The number of everyday tasks and energy-intensive experiments on board space objects is growing with the development of astronautics. To perform energy-intensive work and experiments, energy consumption exceeds the incoming solar fluxes. The modern solar-hydrogen system ensures the reception, conversion and accumulation of excess incoming energy in the form of chemical energy - hydrogen. Cryogenics makes it possible to miniaturize the solar-hydrogen system. The safety of the cryogenic solar-hydrogen energy system is ensured by hydrogen concentration and leakage sensors. The article proposes a comprehensive solution for miniaturization of Power Conversion Unit (PCU) of Energy Power System (EPS) of spacecrafts (SC), which consists in a joint solution of four key problems of miniaturization of power devices: energy, structural, design and technological, system. Miniaturization of the solar-hydrogen energy system (SHES) is achieved by installing onboard hydrogen and oxygen microcryogenic refrigerators, as well as hydrogen and oxygen cryogenic tanks, water tank, electrolyzer and hydrogen fuel cells (FC). The accumulation of chemical hydrogen energy on board the orbiting spacecraft ensures reliable operation when entering the shadow. Storage of hydrogen in cryogenic form significantly reduces the volume required. A cryocooler based on the Stirling cycle provides the process of liquefying hydrogen after the electrolysis of distilled water. In addition, cryogenic temperatures of 20.2 K can be used to thermostat precision instruments placed on board the spacecraft, For example, to ensure the operation of the SQUID. A 3D - model of a voltage stabilization module (VSM), on the basis of which EEC can be produced with different output power and redundancy depth. We give an example of a complete structural scheme of the EEC, which allows implementing all the fulfillment of all the tasks, assigned to the EEC of the PCU SC. [ABSTRACT FROM AUTHOR]

The rapid increase of space debris poses a risk to space activities, so it is vital to develop countermeasures in terms of space surveillance to prevent possible threats. The current Space Surveillance Network is majorly composed of radar and optical telescopes that regularly observe and track space objects. However, these measures are limited by size, being able to detect only a tiny amount of debris. Hence, alternative solutions are essential for securing the future of space activities. Therefore, this paper proposes the design of a payload camera breadboard for space surveillance to increase the information on debris, particularly for the under-catalogued ones. The device was designed with similar characteristics to star trackers of small satellites and CubeSats. Star trackers are attitude devices usually used in satellites for attitude determination and, therefore, have a wide potential role as a major tool for space debris detection. The breadboard was built with commercial off-the-shelf components, representing the current space-camera resolution and field of view. The image sensor was characterized to compute the sensitivity of the camera and evaluate the detectability performance in several simulated positions. Furthermore, the payload camera concept was tested by taking images of the night sky using satellites as proxies of space debris, and a photometric analysis was performed to validate the simulated detectability performance. [ABSTRACT FROM AUTHOR]

Introduction: Stringent cleaning procedures during spacecraft assembly are critical to maintaining the integrity of life-detection missions. To ensure cleanliness, NASA spacecraft are assembled in cleanroom facilities, where floors are routinely cleansed with Kleenol 30 (K30), an alkaline detergent. Methods: Through metabolomic and cultivation approaches, we show that cultures of spacecraft-associated Acinetobacter tolerate up to 1% v/v K30 and are fully inhibited at ≥2%; in comparison, NASA cleanrooms are cleansed with ~0.8-1.6% K30. Results: For A. johnsonii 2P08AA (isolated from a cleanroom floor), cultivations with 0.1% v/v K30 yield (1) no changes in cell density at late-log phase, (2) modest decreases in growth rate (~17%), (3) negligible lag phase times, (4) limited changes in the intracellular metabolome, and (5) increases in extracellular sugar acids, monosaccharides, organic acids, and fatty acids. For A. radioresistens 50v1 (isolated from a spacecraft surface), cultivations yield (1) ~50% survivals, (2) no changes in growth rate, (3) ~70% decreases in the lag phase time, (4) differential changes in intracellular amino acids, compatible solutes, nucleotide-related metabolites, dicarboxylic acids, and saturated fatty acids, and (5) substantial yet differential impacts to extracellular sugar acids, monosaccharides, and organic acids. Discussion: These combined results suggest that (1) K30 manifests strain-dependent impacts on the intracellular metabolomes, cultivation kinetics, and survivals, (2) K30 influences extracellular trace element acquisition in both strains, and (3) K30 is better tolerated by the floor-associated strain. Hence, this work lends support towards the hypothesis that repeated cleansing during spacecraft assembly serve as selective pressures that promote tolerances towards the cleaning conditions. [ABSTRACT FROM AUTHOR]

The article presents the results of a study of perforated plates effectiveness to ensure safety in case of emergencies on spacecraft and launch pads. The degree of shock wave attenuation by perforated plates with different permeability coefficients has been studied experimentally and numerically. Experiments were carried out with conical shock tube, which was used to produce spherical shock waves of varying intensity, and perforated plate with a permeability coefficient of 41% and. Shock wave reflection and attenuation coefficients were obtained. To study the qualitative picture of a shock wave interaction with perforated plate, a numerical model using Ansys Fluent was created and verified. The numerical model made it possible to qualitatively evaluate the breakup and coupling of a shock wave front after passing perforated plate. The degree of a shock wave attenuation by plates with permeability coefficients of 29% and 20.6% was numerically studied. It was shown that perforated plates can effectively reduce a shock wave overpressure. • Perforated plates effectively reduce shock wave overpressure. • Shock wave attenuation nonlinearly depends on permeability coefficient. • Perforated plates do not change shock wave positive phase duration. • Shock wave breakup and coupling does not depend on characteristic size of hole. [ABSTRACT FROM AUTHOR]

A new technology for replenishing daily nitrogen losses (200 g/day) in the atmosphere of the Russian segment of the International Space Station and for cooling the on-board electronic devices is proposed. This technology can be used to create systems for solid nitrogen storage and for release of gaseous nitrogen to create an atmosphere in the living compartments at space stations and, in the future, at the Lunar Station. [ABSTRACT FROM AUTHOR]

This paper considers the idea of reducing the debris of near-Earth space by releasing the spent parts of a spacecraft on orbits touching the conditional boundary of the Earth's atmosphere. We optimize the spacecraft transfer trajectory from a circular reference orbit of an artificial Earth satellite to a target elliptical orbit in a modified pulse problem statement. The convergence of Newton's method is improved by introducing a series of auxiliary coordinate systems at each point of applying an impulse action. The derivatives in the transversality conditions are calculated using a special numerical-analytical differentiation technique. [ABSTRACT FROM AUTHOR]

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Ground acoustic tests using stationary sound pressure level spectrum have been conducted to verify the spacecraft survivability against acoustic environment acting on a spacecraft during launch, which is a non-stationary and random dynamic load. In general, a stationary spectrum used in ground acoustic test is traditionally determined by a method called maximax spectrum, which is the enveloped spectrum of time varying non-stationary short-time Fourier transform. However, the maximax spectrum is more or less an excessively conservative test condition because this spectrum focuses on processing of a time-varying acoustic signal itself to extract maximum value, rather than on how the vibro-acoustic response of an excited structure is. In this paper, a new method is proposed to specify a stationary spectrum equivalent to a structural vibro-acoustic response under a non-stationary and random acoustic environment based on extreme response spectrum and fatigue damage spectrum. This proposed method was applied to flight telemetry of both liquid- and solid-propellant launch vehicles developed by JAXA, to show its effect to mitigate the acoustic test conditions compared to the maximax spectrum while maintaining the equivalence of the structural vibro-acoustic response. Furthermore, the maximum predicted environment, which is the statistical upper percentiles of the flight telemetry of eight liquid-propellant launch vehicles, by the proposed method achieved a mitigation of about 2.5 and 6.8 dB in the extreme stress and cumulative fatigue, respectively, compared to the that which is calculated by the conventional maximax spectrum. [ABSTRACT FROM AUTHOR]

The paper discusses the development of a lunar navigation satellite system and a lunar orbital spaceport based on high near-circular orbits of an Moon artificial satellite and using Russian launch vehicles. Two tasks are studied within a complex problem of a lunar orbital system construction for the navigation support during future development of the Moon. The results of analysis of a spacecraft orbital evolution while moving in the real field of the Moon under the effect of gravitational perturbations are presented. It is shown that the high polar near-circular orbits around the Moon can retain their orbital shape for a long time. Spacecraft insertion into such orbits using a few scenarios is considered. Possible configuration of the lunar navigation satellite system is analyzed to illustrate the fundamental possibility of constructing such a navigation system and a lunar orbital spaceport based in high near-circular orbits around the Moon and using the launch vehicles of the Soyuz family. [ABSTRACT FROM AUTHOR]

The on-board communication standard adopted in current generation space missions of the European Space Agency, and many other agencies as well, is SpaceWire (SpW). In a SpW network, data are exchanged as well-formed packets, whose structure offers low packet overhead and allows developers to tailor their implementation for SpW applications. The development of SpW-compliant devices requires a specific set of test instruments, namely Electrical Ground Support Equipment (EGSE), to verify the correct functionality of SpW units under test. An example of a SpW EGSE is the SpaceART EGSE Emulator, an EGSE for generation and processing of SpW packets for a time-constrained use-case scenario. This EGSE has been developed to address a mission-related SpW communication between a device and an Instrument Control Unit (ICU). It has allowed the generation and processing of specific SpW packets, which cannot be provided by the mostly used general-purpose SpW EGSEs. In this scope, the SpaceART SpW Sniffer, a SpW link analyser for unobtrusive monitoring of SpW link, has been employed to run a comprehensive set of tests and provide further information on the considered scenario. The SpaceART EGSE Emulator-ICU communication has been thoroughly tested through the SpaceART SpW Sniffer, unobtrusively analysing the exchanged data and allowing to assess the compliance with the defined time constraints from an external point of view. The use of the Sniffer has been crucial for testing the on-board communication, providing important support for the success of the mission employing the tested ICU. [ABSTRACT FROM AUTHOR]