Search Results (111)

A novel satellite architecture and deployment method is proposed to reduce the logistical cost and complexity of launching and dispersing satellite constellations. The architecture consists of a primary satellite that separates into multiple smaller units, which are subsequently dispersed using differential drag. An algorithm is developed to determine the required disengagement velocities and optimal timing for separation maneuvers. Two case studies with orbital simulations demonstrate the feasibility of this approach for constellation deployment and phasing. The results indicate that while mission-specific factors influence deployment dynamics, informed selection of the disengagement velocities is crucial for minimizing phase times and mitigating potential delays. The findings confirm the feasibility of the proposed architecture and dispersal method, offering a cost-effective alternative to traditional deployment strategies for future satellite constellations. Full article

This paper investigates the design of optimal low-thrust transfers between relative planar and spatial quasi-satellite orbits (QSOs) in the Earth–Moon system under the Circular Restricted Three-Body Problem (CR3BP). A key contribution is the adaptation of a trajectory optimization framework, previously applied to halo orbit transfers, to accommodate the unique challenges of QSO families, especially the transition between planar and spatial configurations. The method employs a refined beam search strategy to construct diverse initial guess chains, which are then optimized via a successive convexification algorithm tailored for the spatial dynamics of QSOs. Additionally, a linear–quadratic regulator (LQR)-based control scheme is implemented to ensure long-term station-keeping of the final 3D-QSO. Simulation results demonstrate the feasibility of connecting planar and spatial QSOs with minimum-fuel trajectories while maintaining bounded terminal deviations, offering new tools for future Earth–Moon logistics and navigation infrastructure. Key findings include the successful design of low-thrust transfer trajectories between planar QSOs and 1:5 3D-QSOs, with a minimum total $\mathsf{\Delta }V$ of 195.576 m/s over a time of flight (ToF) of 261 days, and a minimum ToF of 41 days with a total $\mathsf{\Delta }V$ of 270.507 m/s. Additionally, the application of LQR control demonstrated the ability to maintain 1:5 3D-QSO families around the Moon with less than 12 mm/s $\mathsf{\Delta }V$ over two months. This research provides valuable insights into the optimization of low-thrust transfer trajectories and the application of advanced control techniques for space missions, particularly those targeting lunar and planetary satellite exploration. Full article

Fuel cells, propulsion systems, and reaction control systems (RCSs) are just a few of the space applications that depend on pressure vessels (PVs) to safely hold high-pressure fluids while enduring extreme environmental conditions both during launch and in orbit. Under these challenging circumstances, PVs must be lightweight while retaining structural integrity in order to increase the efficiency and lower the launch costs. PVs have significant challenges in space conditions, such as extreme vibrations during launch, the complete vacuum of space, and sudden temperature changes based on their location within the satellite and orbit types. Determining the operational temperature limits and endurance of PVs in space applications requires assessing the combined effects of these factors. As the main propellant for satellites and rockets, hydrogen has great promise for use in future space missions. This study aimed to assess the structural integrity and determine the thermal operating limits of different types of hydrogen pressure vessels using finite element analysis (FEA) with Ansys 2019 R3 Workbench. The impact of extreme space conditions on the performances of various kinds of hydrogen pressure vessels was analyzed numerically in this work. This study determined the safe operating temperature ranges for Type 4, Type 3, and Type 1 PVs at an operating hydrogen storage pressure of 35 MPa in an absolute vacuum. Additionally, the dynamic performance was assessed through modal and random vibration analyses. Various aspects of Ansys Workbench were explored, including the influence of the mesh element size, composite modeling methods, and their combined impact on the result accuracy. In terms of the survival temperature limits, the Type 4 PVs, which consisted of a Nylon 6 liner and a carbon fiber-reinforced epoxy (CFRE) prepreg composite shell, offered the optimal balance between the weight (56.2 kg) and a relatively narrow operating temperature range of 10–100 °C. The Type 3 PVs, which featured an Aluminum 6061-T6 liner, provided a broader operational temperature range of 0–145 °C but at a higher weight of 63.7 kg. Meanwhile, the Type 1 PVs demonstrated a superior cryogenic performance, with an operating range of −55–54 °C, though they were nearly twice as heavy as the Type 4 PVs, with a weight of 106 kg. The absolute vacuum environment had a negligible effect on the mechanical performance of all the PVs. Additionally, all the analyzed PV types maintained structural integrity and safety under launch-induced vibration loads. This study provided critical insights for selecting the most suitable pressure vessel type for space applications by considering operational temperature constraints and weight limitations, thereby ensuring an optimal mechanical–thermal performance and structural efficiency. Full article

With the development of commercial space technology and the proposal of concepts such as “Black Jack”, the Space Transport Layer (STL), and the commercial space-based Internet, large-scale low-orbit satellite constellations have become a research hotspot in the aerospace field. Large-scale low-orbit satellite constellations consist of a huge number of satellites, which makes the networking control and operation management of the constellations more complicated. It also increases the difficulty of achieving the economical and efficient networking of the constellations as well as ensuring their safe and stable operation. In this study, aiming at the problem of large-scale constellation phase control in low orbit, strategies for constellation station holding were examined. First, aiming at the problem of station keeping of large-scale constellations in low orbit, the characteristics of satellite phase drift and phase keeping were analyzed, and absolute and relative station-keeping strategies were proposed. Second, a phase-holding loop control method combining semi-major axis overshoot control and passive control was proposed, and a relative phase-maintenance scheme based on a dynamic reference satellite was designed. Then, the absolute and relative station controls of different low-orbit constellations were simulated. The simulation results showed that in order for all satellites in the constellation to maintain a phase angle deviation within ±0.1° in a low-solar-activity year, about 13 days were required on average to adjust the semi-major axis of the satellites by about 71 m. The relative position control of small-scale constellations was simulated, and only four orbital maneuvers were needed to achieve the phase angle maintenance within the threshold of ±5° for all satellites in the constellation within 300 days. Finally, it was concluded that absolute control was suitable for large-scale constellation phase preservation, and relative control was more suitable for small-scale constellation phase preservation. This paper can provide a reference and suggestions for future large-scale constellation deployment and maintenance control strategies of low-orbit constellations. Full article

This paper investigates the fly-around formation between the servicer spacecraft and the target spacecraft. Inspired by the spacecraft orbital motion under the Earth’s gravity, an intuitive, analytical guidance law for spatial fly-around formation design with the low-thrust maneuver is proposed. Beginning with the relative translational dynamics based on relative position and velocity, the control input of the guidance law is designed to contain two parts. The first part is the feed-forward term, which makes the relative dynamics a second-order integration model. The second part is the artificial gravity term, which has similar expressions to the Earth’s gravity, and includes the artificial gravitational coefficient and the vector of the artificial gravity center. The above two parameters can be designed to determine the size, shape, and period of the fly-around trajectory. Specifically, three kinds of fly-around trajectories are discussed in detail. The first two are the spatial ellipses with the target spacecraft locating at the focus and the center of the ellipses, respectively. The third is the spatial circle. The proposed method can be easily extended to the design of planar fly-around formation, which is very systematic and comprehensive, and the fuel consumption of the control input is specifically discussed. Numerical simulations are conducted to demonstrate the efficiency of the proposed method. Full article

During the last 15–20 years, the experimental methods for autonomous navigation and inter-satellite links have been developing rapidly in order to ensure navigation control and data processing without commands from Earth stations. Inter-satellite links are related to relative ranging between the satellites from one constellation or different constellations and measuring the distances between them with the precision of at least 1 $\mathsf{\mu }$m micrometer (=${10}^{-6}$ m), which should account for the bending of the light (radio or laser) signals due to the action of the Earth’s gravitational field. Thus, the theoretical calculation of the propagation time of a signal should be described in the framework of general relativity theory and the s.c. null cone equation. This review paper summarizes the latest achievements in calculating the propagation time of a signal, emitted by a GPS satellite, moving along a plane elliptical orbit or a space-oriented orbit, described by the full set of six Kepler parameters. It has been proved that for the case of plane elliptical orbit, the propagation time is expressed by a sum of elliptic integrals of the first, the second and the third kind, while for the second case (assuming that only the true anomaly angle is the dynamical parameter), the propagation time is expressed by a sum of elliptic integrals of the second and of the fourth order. For both cases, it has been proved that the propagation time represents a real-valued expression and not an imaginary one, as it should be. For the typical parameters of a GPS orbit, numerical calculations for the first case give acceptable values of the propagation time and, especially, the Shapiro delay term of the order of nanoseconds, thus confirming that this is a propagation time for the signal and not for the time of motion of the satellite. Theoretical arguments, related to general relativity and differential geometry have also been presented in favor of this conclusion. A new analytical method has been developed for transforming an elliptic integral in the Legendre form into an integral in the Weierstrass form. Two different representations have been found, one of them based on the method of four-dimensional uniformization, exposed in the monograph of Whittaker and Watson. The result of this approach is a new formulae for the Weierstrass invariants, depending in a complicated manner on the modulus parameter q of the elliptic integral in the Legendre form. Full article

Light curve analysis of defunct satellites is critical for characterizing their rotational motion. An accurate understanding of this aspect will benefit active debris removal and on-orbit servicing missions as part of the solution to the space debris issue. In this study, we explored the attitude behavior of inactive GLONASS satellites, specifically a repeating pattern observed in their spin period evolution. We utilized a large amount of data available in the light curve database maintained by the Astronomical Institute of the University of Bern (AIUB). The morphology of the inactive GLONASS light curves typically features four peaks in two pairs and is presumably attributed to the presence of four evenly distributed thermal control flaps or radiators on the satellite bus. The analysis of the periods extracted from the light curves shows that nearly all of the inactive GLONASS satellites are rotating and exhibit a periodic oscillating pattern in their spin period evolution with an increasing or decreasing secular trend. Through modeling and simulation, we found that the periodic pattern is likely a result of canted solar panels that provide an asymmetry in the satellite model and enable a wind wheel or fan-like mechanism to operate. The secular trend is a consequence of differing values of the specular reflection coefficients of the front and back sides of the solar panels. Assuming an empirical model describing the spin period evolution of 18 selected objects, we found significant variations in the average spin period and amplitude of the oscillations, which range from 8.11 s to 469.58 s and 1.10 s to 513.24 s, respectively. However, the average oscillation period remains relatively constant at around 1 year. Notably, the average spin period correlates well with the average amplitude. The empirical model can be used to extrapolate the spin period in the future, assuming that the oscillating pattern is preserved and roughly shows a linear trend. Full article

This paper introduces a nonlinear Lyapunov-based Finite-Time State-Dependent Differential Riccati Equation (FT-SDDRE) control scheme, considering actuator saturation constraints and ensuring that the control system operates within safe operational limits designed for satellite reconfiguration and formation-keeping in low Earth orbit (LEO) missions. This control approach addresses the challenges of reaching the relative position and velocity vectors within a defined timeframe amid various orbital perturbations. The proposed approach guarantees precise formation control by utilizing a high-fidelity relative motion model that incorporates all zonal harmonics and atmospheric drag, which are the primary environmental disturbances in LEO. Additionally, the article presents an optimization methodology to determine the most efficient State-Dependent Coefficient (SDC) form regarding fuel consumption. This optimization process minimizes energy usage through a hybrid genetic algorithm and simulated annealing (HGASA), resulting in improved performance. In addition, this paper includes a sensitivity analysis to identify the optimized SDC parameterization for different satellite reconfiguration maneuvers. These maneuvers encompass radial, along-track, and cross-track adjustments, each with varying baseline distances. The analysis provides insights into how different parameterizations affect reconfiguration performance, ensuring precise and efficient control for each type of maneuver. The finite-time controller proposed here is benchmarked against other forms of SDRE controllers, showing reduced error margins. To further assess the control system’s effectiveness, an input saturation constraint is integrated, ensuring that the control system operates within safe operational limits, ultimately leading to the successful execution of the mission. Full article

This research addresses minimum-fuel pinpoint lunar landing at the South Pole, focusing on trajectory design and near-optimal guidance aimed at driving a spacecraft from a circular low lunar orbit (LLO) to an instantaneous hovering state above the lunar surface. Orbit dynamics is propagated in a high-fidelity ephemeris-based framework, which employs spherical coordinates as the state variables and includes several harmonics of the selenopotential, as well as third-body gravitational perturbations due to the Earth and Sun. Minimum-fuel two-impulse descent transfers are identified using Lambert problem solutions as initial guesses, followed by refinement in the high-fidelity model, for a range of initial LLO inclinations. Then, a feedback Lambert-based impulsive guidance algorithm is designed and tested through a Monte Carlo campaign to assess the effectiveness under non-nominal conditions related to injection and actuation errors. Because the last braking maneuver is relatively large, a finite-thrust, locally flat, near-optimal guidance is introduced and applied. Simplified dynamics is assumed for the purpose of defining a minimum-time optimal control problem along the last thrust arc. This admits a closed-form solution, which is iteratively used until the desired instantaneous hovering condition is reached. The numerical results in non-nominal flight conditions testify to the effectiveness of the guidance approach at hand in terms of propellant consumption and precision at landing. Full article

The design of a continuous control architecture for faster-than-natural satellite circumnavigation missions, where the desired relative orbital period is shorter than the natural orbital period, is investigated. In such cases, satellites must complete orbits around a target faster than would be possible through natural motion alone, requiring advanced control systems. This paper considers the case of continuous low-thrust control because of its low fuel expenditure. One challenge of continuous control for faster-than-natural relative motion is its susceptibility to noise due to the large nominal thrust that is required. The deleterious effect of noise has not been addressed in the literature for this problem. This work develops a feedforward control that helps to eliminate unnecessary control effort and jitter in the actuation due to noise. When implemented in conjunction with existing feedback control schemes, the resulting control system is shown to perform well and shows promise as an implementable option with current technology. Full article

This review paper identifies key stability and control screening parameters needed to design low-risk, general-purpose high-speed aircraft. These derive from MIL-STD-8785C, MIL-STD-1797, and older AGARD reports, and are suitable for assessing conceptual high-speed vehicles. We demonstrate their applicability using published ground test, computation, and flight test data from the Bell X-2, North American X-15, Martin X-24A, Northrop HL-10, Lockheed Blackbird (YF-12/SR-71), and North American XB-70 as well as the Rockwell Space Shuttle Orbiter. The relative success of the X-15 and Blackbird and the performance limitations of the others indicate the need to scrutinize lateral-directional stability at the preliminary design phase. Our work reveals the need for strong bare-airframe static directional stability to obtain favorable flying qualities. Full article

The increasing need for autonomy in space exploration missions is becoming more and more relevant in the design of missions to small bodies. The long communication latencies and sensitivity of the system to unplanned environmental perturbations mean autonomous methods could be a key design block for this type of mission. In this work, a fully autonomous Guidance, Navigation, and Control (GNC) methodology is introduced. This methodology relies on published CNN-based techniques for surface recognition and pose estimation and also on existing MPC-based techniques for the design of a trajectory to perform a soft landing on an asteroid. Combining Hazard Detection and Avoidance (HDA) with relative navigation systems, a Global Safety Map (GSM) is built on the fly as images are acquired. These GSMs provide the GNC system with information about feasible landing spots and populate a longitude–latitude map with safe/hazardous labels that are later processed to find an optimal landing spot based on mission requirements and a distance-fromhazard metric. The methodology is exemplified using Bennu as the body of interest, and a GSM is built for an arbitrary reconnaissance orbit. Full article

Current-induced field-free magnetic switching using spin–orbit torque has been an important topic for decades due to both academic and industrial interest. Most research has focused on introducing symmetry breakers, such as geometrical and compositional variation, pinned layers, and symmetry-broken crystal structures, which add complexity to the magnetic structure and fabrication process. We designed a relatively simple magnetic structure, composed of a [Co/Pt] multilayer and a Co layer with perpendicular and in-plane magnetic anisotropy, respectively, with a Cu layer between them. Current-induced deterministic magnetic switching was observed in this magnetic system. The system is advantageous due to its easy control of the parameters to achieve the optimal condition for magnetic switching. The balance between magnetic anisotropic strength and interlayer coupling strength is found to provide the optimal condition. This simple design and easy adjustability open various possibilities for magnetic structures in spin-based electronics applications using spin–orbit torque. Full article

This paper proposes a pose control scheme of for proximity maneuvering for tracking and observing noncooperative targets with unknown acceleration, which is an important prerequisite for on-orbit operations in space. It mainly consists of a finite-time extended state observer and constraint processing procedures. Firstly, relative pose-coupled kinematics and dynamics models with unknown integrated disturbances are established based on dual quaternion representations. Then, a finite-time extended state observer is designed using the super-twisting algorithm to estimate the integrated disturbances. Both observation field of view and collision avoidance pose-constrained models are constructed to ensure that the service spacecraft continuously and safely observes the target during proximity maneuvering. And the constraint models are further incorporated into the design of artificial potential function with a unique minimum. After that, the proportional–derivative-like pose-constrained tracking control law is proposed based on the estimated disturbances and the gradient of the artificial potential function. Finally, the effectiveness of the control scheme is verified through numerical simulations. Full article

Control tissue is essential for ensuring the precision of semiquantitative analysis in back-table fluorescence imaging. However, there remains a lack of agreement on the appropriate selection of control tissues. To evaluate the back-table fluorescence imaging performance of different normal tissues and identify the optimal normal tissue, a cohort of 39 patients with orbital tumors were enrolled in the study. Prior to surgery, these patients received indocyanine green (ICG) and following resection, 43 normal control tissues (34 adipose tissues, 3 skin tissues, 3 periosteal tissues, and 3 muscle tissues) were examined using back-table fluorescence imaging. The skin tissue demonstrated significantly elevated fluorescence intensity in comparison to the diseased tissue, whereas the muscle tissue exhibited a broad range and standard deviation of fluorescence signal intensity. Conversely, the adipose and periosteum displayed weak fluorescence signals with a relatively consistent distribution. Additionally, no significant correlations were found between the signal-to-background ratio (SBR) of adipose tissue and patients’ ages, genders, weights, disease duration, tumor origins, dosing of administration of ICG infusion, and the time interval between ICG infusion and surgery. However, a positive correlation was observed between the SBR of adipose tissue and its size, with larger adipose tissues (>1 cm) showing an average SBR 27% higher than smaller adipose tissues (≤1 cm). In conclusion, the findings of this study demonstrated that adipose tissue consistently exhibited homogeneous hypofluorescence during back-table fluorescence imaging, regardless of patient clinical variables or imaging parameters. The size of the adipose tissue was identified as the primary factor influencing its fluorescence imaging characteristics, supporting its utility as an ideal control tissue for back-table fluorescence imaging. Full article