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Direct optimization of many-revolution spacecraft trajectories is performed using an unconstrained formulation with many short-arc, embedded Lambert problems. Each Lambert problem shares its terminal positions with neighboring segments to implicitly enforce position continuity. Use of embedded boundary value problems (EBVPs) is not new to spacecraft trajectory optimization, including extensive use in primer vector theory, flyby tour design, and direct impulsive maneuver optimization. Several obstacles have prevented their use on problems with more than a few dozen segments, including computationally expensive solvers, lack of fast and accurate partial derivatives, unguaranteed convergence, and a non-smooth solution space. Here, these problems are overcome through the use of short-arc segments and a recently developed Lambert solver, complete with the necessary fast and accurate partial derivatives. These short arcs guarantee existence and uniqueness for the Lambert solutions when transfer angles are limited to less than a half revolution. Furthermore, the use of many short segments simultaneously approximates low-thrust and eliminates the need to specify impulsive maneuver quantity or location. For preliminary trajectory optimization, the EBVP technique is simple to implement, benefiting from an unconstrained formulation, the well-known Broyden–Fletcher–Goldfarb–Shanno line search direction, and Cartesian coordinates. Moreover, the technique is naturally parallelizable via the independence of each segment's EBVP. This new, many-rev EBVP technique is scalable and reliable for trajectories with thousands of segments. Several minimum fuel and energy examples are demonstrated, including a problem with 6143 segments for 256 revolutions, found within 5.5 h on a single processor. Smaller problems with only hundreds of segments take minutes. [ABSTRACT FROM AUTHOR]

The purpose of this work is to determine the conditions for the rational use of tubular quartz incandescent lamps with a tungsten spiral, in particular, halogen lamps, as part of infrared simulators of external thermal loads on the surface of a spacecraft. The paper represents the results of a study on the feasibility of using tubular incandescent lamps with a tungsten coil as radiation sources in simulators of external thermal loads on the surface of a spacecraft. Tribology of the elements of infrared simulator ensures the reliable operation of mechanisms, which can be broadly grouped into one-shot devices, and continuously and intermittently operating devices. In large and medium-sized thermal vacuum installations intended for conducting thermal vacuum tests of spacecraft (SC) or their individual fragments, among the systems simulating the main factors of the space environment that affect the thermal state of the SC, there are almost always simplified means of reproducing some components of external thermal loads, and sometimes total loads. It is shown that if an acceptable value of the discreteness of the placement of emitters (at least 20) is implemented in the design of the simulator, the density value of the heat flux leaving the simulator, commensurate with the value of the solar constant, can be achieved only at a relatively high temperature of the spiral - more than 1500 K. Based on the study, it was concluded that incandescent halogen lamps can be used in infrared simulators to simulate external thermal loads on spacecraft with uniform and gray radiation-optical characteristics of their surfaces. [ABSTRACT FROM AUTHOR]

• A method to predict the FEM capability to represent the dynamic characteristics. • Method is well suited for base excitation problems such as a spacecraft. • Correlation is demonstrated using flightworthy spacecraft structure. • Correlation can be done as part of base excited qualification test. The accuracy of the finite element model is an important aspect, especially in the aerospace domain, where design margins are barely minimum. Hence, standard test-analysis model correlation methods apply to assess the quality of the analytical model. However, it is observed that these standard correlation techniques are inadequate to assure the model capability to predict the forced response characteristics of complex structures. The base force assurance criterion is found to be a good tool to overcome this inadequacy. During the dynamic test of the structure, the force measurement device can be used to measure the transmitted force to the base. Thus measured base forces, and the moments computed from the measured forces are vector correlated with the analytical force and moments to predict qualitatively the analytical model capability to represent the design critical peak acceleration responses and the global modes. The procedure is demonstrated using a set of experiments conducted on a real spacecraft under the sine excitations in two mutually perpendicular directions. The sensitivity of the base force assurance criterion to the inaccuracies that may be present in the finite element model and/or in the experimental results is evaluated using simplified Monte Carlo simulations. [ABSTRACT FROM AUTHOR]

The problem of determining the range and the radial velocity of space crafts in the middle and deep space by a monostatic radar is considered. The tested determination method is based on the calculations concerning the relayed signal reception moment and its delay relative to both the emission moment and the Doppler shift of the carrier frequency from the emitted signal frequency (active mode). The results of the determination of the radial velocity only by calculating the Doppler shift of the onboard generator carrier frequency (passive mode) are also examined. The analysis of the space-time relationships during the curvilinear movement of the radar in an inertial geocentric system is conducted and it allows testing the radar as an autonomous measuring device that measures the range to the spacecraft in a certain coordinate system and then matches the measurement result with the estimated time suggested by this system. It is shown that practically the same result is obtained in the station coordinate system for the mean time between the signal emission and the reception relative to the calculated point that does not coincide with the radar location but changes its position with the measured signal delay depending on the time of the day and time of the year. It is established that when the range and the velocity in the station coordinate system (active operation mode) are measured simultaneously, the radial velocity is calculated relative to the target point, while the measured value depends on the angular direction set for the spacecraft. In the passive mode of operation (one-sided emission), to calculate the direction and the components of the velocity vector accurately, one needs a knowledge of the spacecraft coordinates and the scalar of orbital velocity. [ABSTRACT FROM AUTHOR]

Results of solutions to problems of determining the parameters of anisogrid lattice structures of spacecraft bodies are presented. These problems include analyzing the deformability of a body loaded with transverse inertial forces, determining the fundamental frequency of transverse vibrations of a cantilevered body, studying the deformability of a body loaded with axial compressive force, and analyzing the deformability of a body with an internal fuel tank loaded with transverse inertial forces. The problems under consideration are solved using a continuous model of the lattice structure of a cylindrical body and the finite-element method. [ABSTRACT FROM AUTHOR]

Many fields, including the aerospace industry, have shown increased interest in the use of plastics to lower the mass of systems. However, the use of plastics in space can be challenging for a number of reasons. Ultraviolet radiation, atomic oxygen, and other phenomena specifically associated with space cause the degradation of polymers. Here we show a path toward creation of space-grade components by combining additive manufacturing (AM) and atomic layer deposition (ALD). Our method produced ALD Al2O3 coated thermoplastic parts suitable for space applications. The highlight of this work is a significant reduction in outgassing, demonstrated using residual gas analyzer (RGA) sampling. Compared to uncoated parts, the ALD Al2O3 coating decreased the outgassing of polyether ether ketone (PEEK), acrylonitrile butadiene styrene (ABS), polycarbonate (PC), and nanodiamond-doped polylactide (ND-PLA) by 46%, 49%, 58%, and 65%, respectively. The manufacturing method used in this work enables the use of topology optimization already in the early concept creation phase. The method is ideally suited for spacecraft applications, in which the volume and mass of parts is critical, and could also be adapted for in-space manufacturing. [ABSTRACT FROM AUTHOR]

The Mars flux has great influence on the solar panel temperature. This paper establishes a mathematical model of the Mars albedo flux and analyzes the variation of the solar panel temperature in the near Mars area and the far Mars area when the Mars infrared radiation and the Mars albedo change. The research indicates that the influence are different when the spacecraft is at different orbital heights. There exists a critical orbit height. When the orbit height is less than the critical height, the influence is very large, and when it is great than this critical height, the influence could be ignored. This research provids a reference for the future Mars exploration. [ABSTRACT FROM AUTHOR]

This paper is concerned with the transfer trajectory design among multiple planets using multiple flybys. The proposed trajectory design method consists of a semi-analytical tuning strategy and iterative patched-conic method. The tuning strategy refers to a method realizing the smooth connection between a hyperbolic trajectory near a planet and a trajectory obtained by Lambert's problem between the planet and another one. The proposed method iteratively solves Lambert's problem and the tuning strategy so that the arrival or departure velocity obtained by Lambert's problem coincides with the velocity at the corresponding point in the sphere of influence (SOI), assigned by the tuning strategy. Consequently, a smooth and continuous trajectory with multiple flybys can be obtained, where the SOI of each flyby planet and the flyby time are explicitly taken into account. Moreover, analytical conditions with respect to the orbital elements of the arrival and departure hyperbolas are provided to minimize velocity mismatch at the periapsis of a flyby planet. Through numerical simulation using the Cassini example, it is shown that the resultant trajectory by the proposed method is close to the optimal trajectory obtained by the n -body problem. Thus, a suitable approximate trajectory can be effectively generated without solving the n -body problem. • Interplanetary trajectories with multiple flybys are designed. • The proposed method consists of analytical tuning strategy and iterative patched-conic. • The SOI and flyby time for each flyby planet are explicitly taken into account. • Analytical conditions of the orbital elements are provided to minimize velocity mismatch. • A suitable approximate trajectory is effectively generated without solving the N-body problem. [ABSTRACT FROM AUTHOR]

In the quaternionic setting, we consider the problem of optimal control of a spatial attitude maneuver of a spacecraft viewed as a rigid body with one symmetry axis without a constraint on the control function. For the optimality criterion we use a functional combining the time and energy spent on turning the spacecraft. Based on the Pontryagin maximum principle for this problem, new analytical solutions are obtained in the classes of conical and generalized conical motions. Numerical examples are given. [ABSTRACT FROM AUTHOR]

Introduction/purpose: At altitudes of 80 to 40 km, while the spacecraft made of duralumin without a thermal-protective coating was descending from the flight orbit at the first and second cosmic velocities, data were obtained on the increase in density, pressure, and temperature behind the shock front, as well as on the backout of the shock wave from the surface of the descending spacecraft. Methods: Calculations were made of the energy fluxes on the surface of the spacecraft for every 10 km, for convective and radiative heat transfer, as well as for the impact of electrons produced due to ionization of negative ions. Results: At the first cosmic velocity, the greatest energy flux is realized under the influence of an electron flux, and at the second cosmic velocity, radiative heat transfer occurs. In the shock-compressed gas at all the considered altitudes, pressure increases instantly to a value of 109 ÷ 1011 Pa, which leads to a sequential explosion with increasing power resulting in shock waves in the surrounding atmosphere and compression waves in the entire aircraft structure. The last most powerful explosion occurs at an altitude of approx. 40 km. Conclusion: The descending aircraft is destroyed into separate small parts to the size of small dust particles. [ABSTRACT FROM AUTHOR]

In view of at the problem that the traditional STK simulation algorithm for tracking angle of relay antenna cannot he used in real - time testing processing for spacecraft AIT, a real - time tracking algorithm for relay antenna is proposed and implemented in software. In this method, the orbit of spacecraft, antenna beam and coordinate system are applied to establish the equivalent model of the direction angle for relay antenna. Combining the spacecraft flight altitude and real - time telemetry, the tracking angle algorithm, real - time test result verification and closed - loop control strategy for relay antenna are realized. The experimental results show that the proposed method not only achieve real - Lime comparison of relay antenna angle tracking during the AIT test process, but also achieves the real - Lime prediction of TIC coverage under various flight altitude on orbit. The maximum TIC coverage of a single arc is increased by 13.3%. The method described in the study has been already applied to a certain type of spacecraft sub - system test and in orbit flight missions. [ABSTRACT FROM AUTHOR]

Background: spacecrafts rotate around the Earth every 90 minutes, so the 24- hour cycle turns to 90 minutes. Retinoic acid, an active metabolite of vitamin A, plays a role in regulating the circadian rhythm and its deficiency can impair the biological clock function and consequently impair the circadian rhythm of locomotor activity . The goal of the study was to assay the effects of simulated spaceship conditions on serum and hippocampal retinoic acid levels in rats . Methods: The tail-hanging, hindleg-weightlessness rat model was applied to mimic the microgravity conditions from June to September 2017 in the zoo of the Aja University of Medical Sciences. Thirty-two male Wistar rats weighing 200- 250 gr/in an experimental study were randomly assigned into 4 groups (n=8): 1) control animals with 12/12 h day-night rotation, 2) animals with 12/12 h daynight rotation under 14 days unloading, 3) control animals with 45/45 min daynight cycle, and 4) animals with 45/45 min day-night cycle under 14 days unloading. At the termination of the simulation step, rats were anesthetized, and blood samples were collected and the entire brain was removed and each brain was segmented into Hippocampus. ELISA assays were used to detect retinoic acid in the serum and homogenized hippocampus. Data were statistically assayed using two-way ANOVA and Bonferroni as a post hoc test. P<0.05 was considered significant. Results: The circadian cycle of 45 /45 min compared to the 12 /12 h dark cycle increased the levels of retinoic acid in serum (F=4.901, P=0.037) and hippocampus (F=5.793, P=0.029) in rats. But simulated weightlessness had no effect on retinoic acid content in serum (F=0.149, P=0.703) and hippocampus (F=0.004, P=0.952) . Conclusion: It seems that retinoic acid content in serum and hippocampus is affected by circadian cycle change, and weightlessness has not affect. [ABSTRACT FROM AUTHOR]

Lithium-ion battery is generally used as energy storage device in spacecraft and other applications. Spacecraftbattery modeling and energy balance estimation is critical. The proposed genetic algorithm driven generic estimation - energy balance (GAGE-EB) model uses cell open circuit voltage characteristics, internal resistance (rint) characteristics, and coulomb counting. The cell characteristics are obtained by offline tests. For series-parallel or parallel-series configured battery, the model estimates deliverable capacity and internal resistance by accurately matching the terminal voltage. Estimated parameters are used to check the energy balance for new load and generation profiles. The GAGE-EB model is validated for electrical performance with various cell types and battery configurations with respect to experimental data and is within 1% and 2% accuracy for voltage and capacity estimation, respectively. [ABSTRACT FROM AUTHOR]

As the amount of debris in orbit increases, so does the risk of collisions and their seriousness. All nations involved with space operations acknowledge this growing threat. One solution receiving increased attention is active debris removal. The first step in a debris removal mission would be to approach the debris. In this phase, it is important to ensure passive abort safety and to guarantee the robustness against collisions in the case of off-nominal thruster burns, that may be caused by spacecraft anomalies such as navigation sensor or actuator failures. This paper compares two types of passive abort safe trajectories -the V-bar hopping and spiral approaches-considering the ΔV budget, the duration of operations, and variation in the line-of-sight vector to the target. This paper also proposes design strategies for determining the parameters in the two candidate trajectories, considering passive abort safety. The robustness of the trajectories against collisions due to off-nominal thruster burns is also demonstrated through Monte Carlo simulations. The paper investigates which trajectories are suitable for an active debris removal mission to a non-cooperative target. [ABSTRACT FROM AUTHOR]

This paper considers the spacecraft formation flying (SFF) with a desired attitude. Two control problems are studied. One is the formation control of multiple spacecraft and the other is the attitude synchronization of multiple spacecraft. To solve the two problems, a 6-DOF model is adopted for each spacecraft. In the 6-DOF model, it is assumed that the inertia parameters are not well known and there are perturbations of orbitals, gravity, etc. With the aid of the backstepping technique and the algebraic graph theory, distributed robust control laws are proposed for the formation flying and attitude synchronization problems, respectively. The effectiveness of the proposed controllers is shown by simulation. [ABSTRACT FROM AUTHOR]

This paper studies the attitude-tracking control problem of spacecraft considering on-orbit refuelling. A time-varying inertia model is developed for spacecraft on-orbit refuelling, which actually includes two processes: fuel in the transfer pipe and fuel in the tank. Based upon the inertia model, an adaptive attitude-tracking controller is derived to guarantee the stability of the resulted closed-loop system, as well as asymptotic convergence of the attitude-tracking errors, despite performing refuelling operations. Finally, numerical simulations illustrate the effectiveness and performance of the proposed control scheme. [ABSTRACT FROM AUTHOR]

For CubeSats requiring high pointing accuracy and slewing agility, ground-based hardware-in-the-loop simulations are strongly demanded to test and validate spacecraft subsystems and guidance, navigation, and control algorithms. In this article, a magnetic attitude control system (MACS) testbed for a CubeSat is developed utilizing a spherical air bearing and a Helmholtz cage. The design, development, and verification procedure of MACS is presented together with different test scenarios. To generate enough torque with the magnetorquer system in the dynamic testbed, the Helmholtz coil system of the testbed has driven to provide an augmented magnetic field. As an example of experimentation, the B-dot control algorithm was implemented to dissipate the angular momentum of the dynamic MACS testbed. The experimental results were compared with those of the numerical simulations. [ABSTRACT FROM AUTHOR]

The problem of attitude takeover control of spacecraft by using cellular satellites with limited communication, actuator faults and input saturation is investigated. In order to lighten the communication burden of cellular satellites, an event-triggered control strategy is adopted. The filtered attitude information needs to be transmitted only when the defined measurement error reaches the event-triggered threshold in this strategy. Then, to deal with the unknown inertia matrix, actuator faults, external disturbances and the errors caused by event-triggered scheme, fuzzy logic systems is introduced to estimate the uncertainties directly. Combining fuzzy logic control strategy and the event-triggered method, the first event-triggered adaptive fuzzy control law is developed. Then, torque saturation of cellular satellites is further considered in the second control law, where the upper bound of the uncertainties is estimated by fuzzy logic systems. The resulting closed-loop systems under the two control laws are guaranteed to be bounded. Finally, the effectiveness of two proposed control laws is verified by the numerical simulations. [ABSTRACT FROM AUTHOR]

China's Chang'e-4 (CE-4) mission is the first human lander/rover mission on the far side of the Moon. Its probe is composed of a lander, rover, and the Queqiao relay satellite. Queqiao was successfully launched on May 21, 2018, and entered the halo orbit of the L2 point on June 14, becoming the first satellite connecting the Earth and the Moon's far side. The lander carrying Yutu-2 was successfully launched on December 8, 2018, and landed in the Von Kármán crater (45.5° S, 177.6° E) at 10:26 (UTC+8) on January 3, 2019. The CE-4 probe carried nine science instruments. Four instruments are on the lander: a landing camera (LCAM), a terrain camera (TCAM), a low-frequency radio spectrometer (LFRS), and a lunar lander neutrons and dosimetry (LND) provided by Germany. Four instruments are on the rover: a panoramic camera (PCAM), a visible and near-infrared imaging spectrometer (VNIS), a lunar penetrating radar (LPR), and an advanced small analyzer for neutrals (ASAN) provided by Sweden. The instrument on the relay satellite is the Netherlands-China Low-Frequency Explorer (NCLE). The scientific objectives of the CE-4 mission include (1) performing low-frequency radio-astronomical observations; (2) investigating the geomorphology, mineral compositions and shallow subsurface structure of the landing and roving sites; and (3) detecting the Earth-Moon space environment at the lunar far side. As of February 1, 2020, CE-4 has completed 14 lunar days of scientific exploration after one year of operation. The components, fight, scientific objectives and investigation of CE-4 are introduced in this paper. We also describe the accessibility of the initial archived science data and their preliminary analysis results. [ABSTRACT FROM AUTHOR]

• First review of vapor compression refrigeration for space applications. • Summarizing more than 40 years of development. • Covering gas-liquid two-phase flow, thermal gravitational scaling and system design considerations. • Suggesting a general evaluation criterion to select cooling technologies for spacecraft. Using a vapor compression cycle for cooling in microgravity environments was already suggested in the 1970s to leverage the high coefficient of performance. Since then, only a few systems have operated in microgravity with scarce documentation of these flights. The lack of measured data and detailed documentation makes identifying the necessary next steps difficult for researchers entering the field of refrigeration in space. This paper provides a review of available literature for vapor compression systems flown in microgravity by outlining the history of vapor compression devices in space and presenting performance data. Moreover, gaps in the literature are highlighted and open questions are posed based on the reviewed material. Next steps of research are suggested to support and ultimately achieve reliable vapor compression refrigeration in space. Calculating equivalent masses for a fair comparison of different microgravity cooling technologies is proposed by capturing both energy consumption and used volume. [ABSTRACT FROM AUTHOR]

Highly porous aluminum alloys and honeycomb structures may be utilized as single-use shock absorbers in landing gear and other elements of spacecraft thus ensuring safety of the landing process. This paper presents the results of studies that make it possible to estimate mechanical properties and shock-absorbing characteristics of highly porous aluminum alloys and honeycomb aluminum alloy structures, the initial average densities of said alloys and structures ranging from 330 to 2250 kg/m³. We list analogs that correspond with different stages found on stress-strain curves of highly porous aluminum alloys and honeycomb structures. The paper also provides the foundations of calculating the specific and effective energies of shock absorption. • The deformation diagrams of cellular aluminum are close to the idealized one. • Specific and effective impact absorption energy calculated by the deformation diagram. • Some honeycomb structures have a deformation diagram with hardening. • The presence of hardening worsens the shock-absorbing characteristics. • The deformation diagram depends on the porosity, matrix material and pore structure. [ABSTRACT FROM AUTHOR]

This paper presents a neural-aided Guidance, Navigation & Control algorithm for reconfiguration of distributed space systems. The guidance algorithm is based on Artificial Potential Fields (APF) in the Relative Orbital Elements (ROE) space. Since the relative orbit determination measurements are typically referred to the Cartesian metrics (e.g. range or range rate), a linear mapping between the set of ROE and the Cartesian coordinates expressed in the Local-Vertical-Local-Horizontal (LVLH) reference frame is derived. The navigation and control algorithms rely on the relative dynamics expressed in the same ROE set of coordinates. To cope with uncertainties and nonlinearities of the system, a Radial Basis Function Neural Network (RBFNN) is employed to reconstruct the perturbed dynamics. The Artificial Neural Network (ANN) is coupled with an adaptive Extended Kalman Filter for state estimation. A feedback control is designed to track the desired state, whose stability is analyzed using Lyapunov theory. The guidance, navigation and control algorithms are tested in a high-fidelity numerical orbit propagator. Moreover, the algorithm is tested in relevant Processor-In-the-Loop (PIL) simulations using a TI C2000-Delfino MCU F28379D. The results demonstrate the effectiveness of the algorithm for relative reconfiguration maneuvers involving relative distances ~ 10 2 m with limited fuel consumption and constrained available thrust (⩽ 1 mN). In particular along-track maneuvers, relative plane change and formation enlargement are analysed in the paper, showing the comparison between the proposed algorithm and the legacy one without the neural network. The benefit of implementing a neural network is particularly highlighted when the nonlinearities or unmodelled terms in the on-board dynamics become prominent. [ABSTRACT FROM AUTHOR]

Internal charging effect is one of the main factors affecting on-orbit performance of spacecraft. Current measurement can directly obtain the firsthand data of the internal charging effect. It also can obtain electron environment status through current measurement. Based on the existing current data and the possible exposure of energetic electron storm, the current of the internal charging effect is 10 fA—500 pA on the medium and high earth orbits. For the requirement of the current monitoring for the future earth orbit, the Jupiter and Mars exploring mission, a pico current measurement method has been proposed. A pico current monitor unit has been developed and calibrated in the laboratory. The calibration results have shown that the linearity between input current and output voltage is good from 10 fA to 500 pA. [ABSTRACT FROM AUTHOR]

In this article, a novel distributed attitude coordination control scheme is proposed for a group of rigid spacecraft with unknown disturbances. The communication topology is described by a general directed graph, and the leader's attitude is available to only a small fraction of followers. Different from existing attitude coordination control schemes, here an event-triggered communication strategy without Zeno phenomenon is designed, which avoids continuous communication between neighboring spacecraft and considerably reduces the communication burden. Moreover, a nonlinear observer is constructed for each follower, which enables us to estimate and compensate for a large class of disturbances exactly. The proposed scheme is able to ensure the boundedness of all closed-loop signals and steer the attitude tracking errors into an arbitrarily small residual set, regardless of the intermittent communication. Simulation results illustrate the effectiveness of the proposed scheme. [ABSTRACT FROM AUTHOR]

Summary: This article is devoted to the problem of attitude control for rigid spacecraft subject to multiple disturbances. Due to the limited communication and storage resource, an event‐triggered anti‐disturbance attitude control approach is proposed for the attitude control system (ACS) with multiple disturbances. Different from some existing event‐triggered attitude control methods, the multiple disturbances are considered and can be divided into an uncertain modeled disturbance and a norm bounded equivalent disturbance. A disturbance observer is designed to estimate the modeled disturbance. Based on the framework of disturbance observer‐based control, the proposed event‐triggered anti‐disturbance controller can guarantee that the ACS converges toward a small invariant set, which successfully avoids continuous communication and Zeno phenomenon. Finally, simulation results are given to demonstrate the efficiency of the proposed approach. [ABSTRACT FROM AUTHOR]

Photocatalysis has the potential to solve the contamination problem for spacecrafts in orbit. The atmosphere surrounding photocatalysts affects their photocatalytic activity. In this paper, decomposition tests of methyl red (MR) under vacuum condition were conducted to investigate the effects of vacuum on basic photocatalytic activity of titanium dioxide (TiO 2). TiO 2 thin films were fabricated on quartz glass and indium-tin oxide (ITO) substrates by a sol–gel method. The TiO 2 samples, coated with MR, were irradiated with UV rays in vacuum, and their absorbance spectra were measured. Moreover, mass change of MR and gas chromatography mass spectrometry were measured using TiO 2 powder to investigate intermediate materials (IMs) generated from decomposed MR. In vacuum, TiO 2 /quartz glass could decompose MR to IMs but could not further decompose the IMs. To investigate which roles of H 2 O and oxygen molecules lead to this result, decomposition experiments, using TiO 2 /ITO and TiO 2 /ITO filled with oxygen vacancies, were performed. The results indicated that TiO 2 cannot decompose IMs in vacuum due to the lack of active oxygen species and/or radical chain reaction of oxygen molecules. • The effects of vacuum on photocatalytic activity of TiO 2 were investigated. • TiO 2 can decompose methyl red to intermediate materials even in vacuum. • TiO 2 cannot further decompose the intermediate materials in vacuum. • This results from lack of active oxygen species and/or radical chain reaction. [ABSTRACT FROM AUTHOR]

Electron microscopic studies of the surface of the spacecraft descent vehicle (DV) have revealed faceted particles of various shape and size, which are formed when the DV travels through the Earth's atmosphere at an ultrasonic velocity. The particles are no more than 5 µm in size. An X-ray diffraction analysis of the phase composition of the precipitate on the DVs porthole after their landing showed that the test samples contain the following minerals: сryptohalite—(NH4)2[SiF6]), sal ammoniac—NH4Cl, mullite—Al2SiO5, and other minerals that rarely occur in nature. The paper presents results of electron microscopic and X-ray studies of samples taken from 4 descent vehicles. [ABSTRACT FROM AUTHOR]

The fuel mixing mechanism is a critical aspect of the propulsion system of hypersonic vehicles. This study aimed to comprehensively investigate a novel jet configuration that enhances fuel mixing in the combustion chamber. The primary objective was to analyze the significance of vortexes in fuel distribution within the combustion chamber. Computational fluid dynamics was used to investigate the three-dimensional structure of the transverse fuel jet released from an extruded nozzle in a supersonic free stream. The influence of jet type (circular and lobe structure nozzles) and nozzle height was fully explored in this research. The free stream air jet was supersonic (Mach = 4), and the hydrogen jet was injected using various annular nozzles in sonic condition. The results obtained show that the use of an annular circular nozzle is efficient in fuel mixing behind the extruded nozzle. Additionally, fuel mixing improves by approximately 90% when the height of the extruded 3-lobe injector is increased by twice its original length. • Effect of injector type on mixing performance of single annular extruded nozzle at supersonic flow is studied. • Flow analysis around the extruded nozzle is done to disclose the shock effects on mixing mechanism. • Computational modeling is applied to visualize the compressible flow near extruded nozzle. • Circulation and mixing are compared downstream of various extruded nozzles. [ABSTRACT FROM AUTHOR]

• A thermal surrogate model for spacecraft systems was proposed. • Physics-informed machine learning was successfully applied to the thermal mathematical model of spacecraft systems. • POD data reduction enabled us to predict full model temperature distribution. • Physics-informed machine learning decreased the total training cost by 26% to 81% compared to other methods. Thermal analysis of spacecraft systems is one of the most critical processes in the design and operation of space missions. However, it is computationally costly to predict the temperature under all conditions including attitude, orbit, or all combinations of equipment use cases. This study proposes a novel surrogate model for full-scale thermal models using data reduction via proper orthogonal decomposition and physics-informed machine learning. The prediction accuracy of the proposed method is evaluated using two types of thermal mathematical models to investigate the dependency of the training dataset and the size of the models. In addition, the prediction accuracy and training cost were compared between data-driven machine learning and physics-informed machine learning. As a result, the proposed method successfully predicted the full model temperature for both models under various heat input conditions. Moreover, the proposed method decreased the total training cost by 26% to 81% compared to data-driven machine learning. Thermal analysis of spacecraft systems is one of the most critical processes in the design and operation of space missions. However, it is computationally costly to predict the temperature under all conditions including attitude, orbit, or all combinations of equipment use cases. This study proposes a novel surrogate model for full-scale thermal models using data reduction via proper orthogonal decomposition and physicsinformed machine learning. The prediction accuracy of the proposed method is evaluated using two types of thermal mathematical models to investigate the dependency of the training dataset and the size of the models. In addition, the prediction accuracy and training cost were compared between data-driven machine learning and physics-informed machine learning. As a result, the proposed method successfully predicted the full model temperature for both models under various heat input conditions. Moreover, the proposed method decreased the total training cost by 26% to 81% compared to data-driven machine learning. [ABSTRACT FROM AUTHOR]

The problems of autonomous digital control of a spacecraft in the initial orientation modes are considered. The developed methods, algorithms and simulation results for the information satellites and space robots in a sun-synchronous orbit are briefly presented. [ABSTRACT FROM AUTHOR]

During ground-based thermal vacuum tests of spacecraft, it is necessary to control and monitor the change in dust content in closed and semiclosed hollows of the spacecraft, near the optical surfaces of scientific and service hardware, mirrors, solar arrays, multilayer insulation, as well as in the volume of the thermal-vacuum chamber itself during its evacuation etc. It is known that the scientific reliability and significance of the results from instruments in space to ground-based research centers are greatly affected by dusty optics. The article provides a description of a new method for automated dust control of the external atmosphere of a spacecraft and its modules during thermal vacuum tests. A patent has been issued for this method; the patent holder is the Russian Federation, on behalf of which the State Space Corporation Roscosmos acts (RU). In addition to the rocket and space and aviation industries, this method can be used in electronic and microelectronic technology, nuclear technology, medicine, the food industry, etc. [ABSTRACT FROM AUTHOR]

The construction of the trajectories for spacecraft return from the Moon to the Earth for delivering lunar soil to the polygon specified by the corresponding geographical coordinates (φΠλΠ) considered in this paper. A special technique for this purpose is developed. The spacecraft departs from the polar near-Moon orbit with a height of 100 km. The analysis of the energy characteristics of the spacecraft return trajectories is carried out for the year 2024. The most favorable dates in terms of the minimum value of acceleration impulse are selected. The satellite flight from the Moon to the Earth is simulated taking account of the perturbations of gravitational fields of the Earth, the Moon, and the Sun in the geocentric geoequatorial nonrotating coordinate system. [ABSTRACT FROM AUTHOR]

The paper describes a modified smooth approximation method for smoothing the dependence of the equivalent fluence of charged particles of the Earth's Van Allen radiation belts on time and position in near-Earth space. It is shown that this method can be effectively used to solve the problems of minimizing radiation-induced degradation of spacecraft solar cell arrays. The latter problems involve the variation of the trajectory that the spacecraft follows during an electric propulsion transfer from an initial orbit to a geostationary one. [ABSTRACT FROM AUTHOR]

This paper investigates the attitude maneuver problem of a spacecraft containing internal movable masses with zero angular momentum. The spacecraft dynamics in one dimension is first studied using the proposed moving mass control system with predefined mass motion. Next, a small-angle attitude maneuver method based on the sliding mode control technique is designed to accomplish the reorientation when the required maneuver angle is within the control capability of single-direction movable mass motion. Furthermore, a novel mass-shifting procedure is proposed for achieving full-domain attitude reorientation. The procedure includes a rest-to-rest maneuver method together with the trajectory optimization. Which, the rest-to-rest maneuver takes advantage of the attitude difference caused by sequential mass motion and the trajectory optimization is carried out based on time sub-optimal motion planning methods. The procedure is next combined with the small-angle maneuver method to achieve a high precision reorientation. Finally, numerical simulation results show that high precision full-domain attitude reorientation can be realized using the proposed control system with three internal movable masses. • Feasibility of full-domain attitude reorientation using internal movable masses is investigated for the first time. • A novel control procedure using internal movable masses to achieve full-domain attitude reorientation is proposed. • A high precision reorientation method in a combination of sequence rotation with small angle maneuver is presented. [ABSTRACT FROM AUTHOR]

A new fire-extinguishing technology in an inhabited pressurized compartment of manned spacecraft with simulated gravity created for prevention of negative influence of zero-gravity on cosmonauts during a long-term spaceflight is presented. The new fire-extinguishing technology in an inhabited pressurized compartment of a manned spacecraft involves selection of polymeric materials, which are not combustible in simulated gravity conditions and generation of momentum of reaction force in the direction opposite to the direction of its rotation for the time, during which the angular velocity of rotation of the spacecraft module is reduced to the value, under which gravity acceleration is decreased up to the value of combustion limit by gravity acceleration of the most combustible polymeric materials at all the points of the inhabited pressurized compartment during fire occurring inside the pressurized compartment with simulated gravity simultaneously with the activation of automatic fire-extinguishing system, which operation is based on the immediate decrease of ventilation flow velocity lower that combustion limits by to gravity acceleration of the most combustible polymeric materials. • Research results of limiting combustion conditions under decreased gravity accelerations are considered. • Design of fire-extinguishing technology in INPC SC during the flight with simulated gravity is studied. • Scientific basis for design of fire extinguishing means in InPC SC in simulated gravity mode gravity is analyzed. • Arrangement of experimental installations recommended for calculation of G lim value of materials is presented. [ABSTRACT FROM AUTHOR]

• The dynamical model of a three-axis stabilized spacecraft with lateral solar arrays is established. • Comparisons of assumed modes and global modes are presented for the multi-rigid-flexible body coupling modes. • The thermal stress of flexible structures is considered and the finite difference method is adopted. • Thermal response analysis of the spacecraft is given under the thermal alternation. The rigid-flexible-thermal coupling dynamic analysis for a spacecraft in orbit is studied in this paper. The spacecraft consists of a central rigid platform and two groups of lateral solar arrays. There exists the relative motion between the rigid platform and solar arrays, thus the spacecraft is a multi-rigid-flexible bodies coupling system. As the spacecraft in orbit experience different light areas, alternations of the heat flux on solar arrays can result in changes of dynamic characteristics. Considering thermal stress effects of solar arrays, the dynamical model of the spacecraft is established by using Hamiltonian principle. Further, multi-rigid-flexible coupling modes of the system are obtained. The finite difference method is developed to obtained the responses of the spacecraft and the variation of temperature gradients under the different solar radiation. Results of natural characteristics illustrate that constrained modes can be used to discrete the system directly and efficiently. Modal shapes and parameters analysis reveal the rigid-flexible coupling effects of such spacecraft. The thermal-structural analysis demonstrates the thermal alternation may induce the vibration and even change the original vibration of the spacecraft. [ABSTRACT FROM AUTHOR]

We consider the methodology of intelligent analysis of telemetry data based on modern mathematical methods of information processing. We define urgent tasks of analysis and forecasting, taking into account the 20-year experience of operating the Russian segment of the International Space Station (ISS RS). We give a basic description of these methods and present the results of testing on real data obtained from ISS RS telemetry information. [ABSTRACT FROM AUTHOR]

A novel state estimation method is proposed for target tracking in the boost stage using space-based infrared cameras (SBIRC) whose measurements are essentially corrupted by both Gaussian noise and quantization noise. As the quantization noise has non-Gaussian properties, conventional extended Kalman filtering (EKF) suffers from poor performance. The quantization noise of SBIRC is modelled using the mid-riser quantizer, which is usually adopted in the digital signal processing field. A novel minimum square error (MMSE) state estimation algorithm with quantized measurements, named the quantized extended Kalman filtering (QEKF), is then proposed. The time update is given based on first-order linearization of the nonlinearities, and the measurement update is derived based on the conditional mean estimate given the quantized measurements. As the multidimensional integrals in the measurement update derived doesn't have analytical solutions, a numerical integration method is proposed by combining Genz's transformation and quasi-Monte Carlo (QMC) method, which can avoid the curse of dimensionality. To further improve the tracking accuracy, quantized high-degree cubature Kalman filtering (QHCKF) is developed by integrating the fifth-degree cubature rule into the framework of the QEKF. Numerical simulation results illustrate the superiority of the proposed QEKF and QHCKF methods. [ABSTRACT FROM AUTHOR]

Alpha Magnetic Spectrometer (AMS-02) is a particle-physics detector from a module attached to the outside of the International Space Station (ISS). The temperature of the components in AMS-02 must be kept within their different operational ranges but also must be stable over both time and volume. Thermal modeling and simulations for the radiator and the electronic crates of AMS-02 were carried out by applying the Crank–Nicholson implicit solution. Based on reducing the temperature gradient of the radiator and the mass of the thermal control system, the layout of the heat pipes in the radiator was optimized to solve the overheating issue of the electronic crates. The non-operational and operational temperature dissipation for the thermal control system was calculated. The analysis results for the radiator and the electronic crates can meet the running requirements of AMS-02 on the ISS. [ABSTRACT FROM AUTHOR]

Materials have a primary purpose in the design of space vehicles, such as fuels, walls, racks, windows, etc. Additionally, each will also effect space radiation protection. The shielding capabilities of 39 materials and nine layering configurations are evaluated for deep space travel in terms of whole-body effective dose equivalent (ED). Polymer and composite materials are also evaluated in terms of # o f N u c l e o n s V o l u m e . It is clear that a "magic" material or layering configuration is not possible; however, polymers and composites should be used instead of metals if they can serve their primary purpose. Polyethylene is shown to be the best feasible material from this material sample. Thermal neutron absorbers 6Li and 10B do not have a significant effect on ED as homogeneous shields or in layering configurations. Alloying of materials such as aluminum for strengthening purposes does not increase ED. Tanking liquid hydrogen within aluminum does significantly reduce ED when compared to aluminum. Ultimately, a space vehicle is a system of systems and radiation protection must be one of them. [ABSTRACT FROM AUTHOR]

Aluminum corrosion has become a major obstacle in spacecraft construction given that aluminum is used extensively throughout the construction process. Despite its many attributes in strength and durability, aluminum is susceptible to corrosion, in particular, corrosion due to microbial contamination. Scientists have encountered a number of problems with microbial aluminum corrosion within spacecraft components. Here, we summarize recent findings with regard to the phenomenon of microbiologically influenced corrosion (MIC) on space stations in the context of microbial strains isolated from the Mir space station (Mir) and the International Space Station (ISS). Given that strains found on spacecraft are of terrestrial origin, an understanding of the contribution of Al-corrosive microbes to corrosion and related risks to space travel and astronaut health is essential for implementation of prevention strategies. Accordingly, an efficient rapid identification method of microbes with the capability to degrade aluminum is proposed. In particular, onboard implementation of a matrix-assisted laser desorption/ionization–time of flight mass spectrometer (MALDI-TOF MS) is addressed. The use of a MALDI-TOF MS on board spacecraft will be crucial to future successes in space travel given that traditional methods of identifying corrosive species are far more time-consuming. Identification of microbes by way of a MALDI-TOF MS may also aid in the study of microbial corrosion and be a valuable asset for MIC prevention. [ABSTRACT FROM AUTHOR]

Other papers in this special issue have discussed the diversity of planetary atmospheres and some of the key science questions for giant planet atmospheres to be addressed in the future. There are crucial measurements that can only be made by orbiters of giant planets and probes dropped into their atmospheres. To help the community be more effective developers of missions and users of data products, we summarize how NASA and ESA categorize their planetary space missions, and the restrictions and requirements placed on each category. We then discuss the atmospheric goals to be addressed by currently approved giant-planet missions as well as missions likely to be considered in the next few years, such as a joint NASA/ESA Ice Giant orbiter with atmospheric probe. Our focus is on interplanetary spacecraft, but we acknowledge the crucial role to be played by ground-based and near-Earth telescopes, as well as theoretical and laboratory work. [ABSTRACT FROM AUTHOR]