Search Results (165)

Small satellites or CubeSats orbiting in low Earth orbit (LEO) have become increasingly popular in Earth Observation missions, where high-resolution imaging is essential. Due to the lower mass of these spacecrafts, they are more sensitive to vibrations, and image quality can be particularly negatively affected by micro-vibrations. These vibrations originate from on-board subsystems, such as the Attitude Determination and Control System (ADCS), which uses reaction wheels to change the orientation of the satellite. The main goal of our research was to analyze these micro-vibrations so that the acquired data could be used for post-correction of camera images. Obuda University, as a participant in a research project, was tasked with designing and building a micro-vibration measuring device for the LEO CubeSat called WREN-1. In the first phase of the project, the satellite was launched into orbit, and test data were collected and analyzed. The results are presented in this article. Based on the data obtained in this way, the next step will be to analyze the images taken at the same time as the vibration measurements and to search for a correlation between the image quality and the vibrations. Based on the results of the entire project, it could be possible to improve the image quality of the onboard cameras of microsatellites. Full article

The interstellar object 3I/ATLAS is expected to arrive at a distance of $53.56(\pm 0.45)$ million $\mathrm{km}$ ($0.358\pm 0.003$ au) from Jupiter on 16 March 2026. We show that applying a total thrust $\Delta$V of $2.6755\mathrm{km}{\mathrm{s}}^{-1}$ to the lower perijove on 9 September 2025 and then executing a Jupiter Oberth Maneuver can bring the Juno spacecraft from its orbit around Jupiter to intercept the path of 3I/ATLAS on 14 March 2026. We further show that it is possible for Juno to come much closer to 3I/ATLAS (~27 million $\mathrm{km}$) with 110 kg of remaining propellant, merely 5.4% of the initial fuel reservoir. We find that for low available $\Delta$V, there is no particular benefit in the application of a double impulse (for example, to reach ~27 million $\mathrm{km}$ from 3I/ATLAS); however, if Juno has a higher $\Delta$V capability, there is a significant advantage of a second impulse, typically saving propellant by a factor of a half. A close fly-by might allow us to probe the nature of 3I/ATLAS far better than telescopes on Earth. Full article

With the continuous advancement of on-orbit services and space intelligence sensing technologies, the efficient and accurate identification of spacecraft components has become increasingly critical. However, complex lighting conditions, background interference, and limited onboard computing resources present significant challenges to existing segmentation algorithms. To address these challenges, this paper proposes a lightweight spacecraft component segmentation framework for on-orbit applications, termed M4MLF-YOLO. Based on the YOLOv5 architecture, we propose a refined lightweight design strategy that aims to balance segmentation accuracy and resource consumption in satellite-based scenarios. MobileNetV4 is adopted as the backbone network to minimize computational overhead. Additionally, a Multi-Scale Fourier Adaptive Calibration Module (MFAC) is designed to enhance multi-scale feature modeling and boundary discrimination capabilities in the frequency domain. We also introduce a Linear Deformable Convolution (LDConv) to explicitly control the spatial sampling span and distribution of the convolution kernel, thereby linearly adjusting the receptive field coverage range to improve feature extraction capabilities while effectively reducing computational costs. Furthermore, the efficient C3-Faster module is integrated to enhance channel interaction and feature fusion efficiency. A high-quality spacecraft image dataset, comprising both real and synthetic images, was constructed, covering various backgrounds and component types, including solar panels, antennas, payload instruments, thrusters, and optical payloads. Environment-aware preprocessing and enhancement strategies were applied to improve model robustness. Experimental results demonstrate that M4MLF-YOLO achieves excellent segmentation performance while maintaining low model complexity, with precision reaching 95.1% and recall reaching 88.3%, representing improvements of 1.9% and 3.9% over YOLOv5s, respectively. The mAP@0.5 also reached 93.4%. In terms of lightweight design, the model parameter count and computational complexity were reduced by 36.5% and 24.6%, respectively. These results validate that the proposed method significantly enhances deployment efficiency while preserving segmentation accuracy, showcasing promising potential for satellite-based visual perception applications. Full article

The rapid development of spacecraft on-orbit services has increased the requirements for docking technology, especially for large-scale payloads that exceed the launch envelope. Docking technology based on astronaut extravehicular activities is one of the most promising directions for on-orbit services. In view of this, this paper designs and characterizes a handheld double-point docking mechanism for assembling large-scale payloads that is suitable for extravehicular activity (EVA) in dual-astronaut collaborative operations. It achieves the functional decoupling of docking, locking, unlocking, and separation throughout the whole process. The mechanism also has excellent design for human factors engineering, allowing astronauts to change hands, operate with one hand, and apply limited force. The mechanism adopts a dual-point probe–drogue configuration, while the misalignment tolerance design guarantees the docking accuracy and the operating range, and forms a rigid structural connection through a force amplification mechanism. Theoretical analysis and numerical simulations are implemented to estimate the dynamics, statics, and kinematics of the docking process. Corresponding experiments of the prototype are also conducted, including high–low temperature dynamics, docking tests, and kinematic tolerance experiments. The experiments validate the finite element analysis and verify the actual performance of the mechanism. The designed handheld dual-point docking mechanism was successfully applied for the first time by the Shenzhou 15 crew on China’s Space Station in March 2023. This paves a new road for spacecraft on-orbit service of large-scale payloads by EVAs, providing guidance as well as a technical foundation for the on-orbit construction of large spacecraft in the future. Full article

Non-cooperative spatial target detection plays a vital role in enabling autonomous on-orbit servicing and maintaining space situational awareness (SSA). However, due to the limited computational resources of onboard embedded systems and the complexity of spaceborne imaging environments, where spacecraft images often contain small targets that are easily obscured by background noise and characterized by low local information entropy, many existing object detection frameworks struggle to achieve high accuracy with low computational cost. To address this challenge, we propose YOLO-GRBI, an enhanced detection network designed to balance accuracy and efficiency. A reparameterized ELAN backbone is adopted to improve feature reuse and facilitate gradient propagation. The BiFormer and C2f-iAFF modules are introduced to enhance attention to salient targets, reducing false positives and false negatives. GSConv and VoV-GSCSP modules are integrated into the neck to reduce convolution operations and computational redundancy while preserving information entropy. YOLO-GRBI employs the focal loss for classification and confidence prediction to address class imbalance. Experiments on a self-constructed spacecraft dataset show that YOLO-GRBI outperforms the baseline YOLOv8n, achieving a 4.9% increase in mAP@0.5 and a 6.0% boost in mAP@0.5:0.95, while further reducing model complexity and inference latency. Full article

This paper aims to present a set of possible transfer trajectories for a rendezvous mission with asteroid 2024 YR4, using a spacecraft propelled by a photonic solar sail. Asteroid 2024 YR4 was discovered in late December 2024 and was briefly classified as Torino Scale 3 for three weeks in early 2025, before being downgraded to zero at the end of February. In this study, rapid Earth-to-asteroid transfers are analyzed by solving a typical optimal control problem, in which the thrust vector generated by the solar sail is modeled using the optical force approach. Numerical simulations are carried out assuming a low-to-medium performance solar sail, considering both a simplified orbit-to-orbit transfer and a more accurate scenario that incorporates the actual ephemerides of the celestial bodies. The numerical results indicate that a medium-performance solar sail can reach asteroid 2024 YR4, achieving the global minimum flight time and arriving before its perihelion passage in late December 2032. Full article

Continuous low thrust is widely used in orbit transfer maneuvers. If the unknown maneuvers are not correctly compensated, the orbiting accuracy will be seriously affected. We propose a rapid method for pre-identifying thrust acceleration based on single-arc orbit determination in order to determine the orbit of non-cooperative continuous low-thrust maneuvering spacecraft. The single-arc orbit determination results of two ground-based radar observations with a certain time interval are used to inversely determine the direction and magnitude of acceleration of the spacecraft under continuous thrust based on their relationship with satellite orbit parameters. The solution error is relatively small when using this method, even over a short period of time when data are sparse. The results can then be applied to the orbital adjustment of a satellite. The results show that when the satellite climbs with maximum tangential acceleration, the interval between the two radar observations is greater than 7 h, and the proposed method can rapidly pre-identify tangential thrust acceleration with a solution error of less than 5%. When the satellite adjusts the orbital plane with the maximum normal acceleration, the average relative measurement error of the normal acceleration is about 20% when the time interval between two observations is 24 h. The longer the observation interval and the greater the thrust acceleration, the smaller the relative error. The calculation results can be used as the initial value for precision orbit determination of continuous low-thrust maneuvering spacecraft. Full article

Since 1957, the near-Earth population of trackable space objects has grown in number to over 36,000. Of these 36,000+ trackable objects now in low Earth orbit, just a few thousand are working spacecraft. The rest are Earth-orbiting objects which are no longer operational and are considered to be space junk. Because this junk can no longer receive maneuvering commands from its Earth-based owners, the survivability of other spacecraft traveling through or operating in Earth orbit can be jeopardized by the impacts of any number of pieces of this space junk, whose origins can usually be traced back to defunct satellites. As a result, a major design parameter for Earth-orbiting spacecraft is the possibility of such high-speed impacts and the damage they can cause. Furthermore, several private companies are now launching several thousand spacecraft into Earth orbit, many of which are satellites built for communication purposes. Other satellites have been launched to expand the reach of the World Wide Web and to provide better tools for disaster management. Two questions quickly become evident, namely, what is the beneficial purpose of these large satellite constellations, and what are some of the deleterious consequences of their proliferation? Numerous topics related to space debris will be discussed in this paper, including issues in space law that concern the growing problem of orbital debris. In the end, several areas of concern will be noted that are vital to the continuing presence of humans in near-Earth space and must be addressed as the near-Earth orbital environment becomes more congested and space traffic management becomes more difficult. Full article

Laser propulsion is a new conceptual technology that drives spacecraft and possesses advantages such as high specific impulse, large payload ratio, and low launch cost. It has potential applications in diverse areas, such as space debris mitigation and removal, microsatellite attitude control, and orbital maneuvering. Liquid pulse laser propulsion has notable advantages among the various laser propulsion systems. We review the concept and the theory of liquid laser propulsion. Then, we categorize the current state of research based on three types of propellants—non-energetic liquids, energetic liquids, and liquid metals—and provide an analysis of the propulsion characteristics arising from the laser ablation of liquids such as water, glycidyl azide polymer (GAP), hydroxylammonium nitrate (HAN), and ammonium dinitramide (ADN). We also discuss future research directions and challenges of pulsed liquid laser propulsion. Although experiments have yielded encouraging outcomes due to the distinctive properties of liquid propellants, continued investigation is essential to ensure that this technology performs reliably in actual aerospace applications. Consistent results under both spatial and ground conditions remain a key research content for fully realizing its potential. Full article

A stable and high-precision autonomous orbit determination scheme for a Low Lunar Orbit (LLO) spacecraft is proposed, leveraging satellite-to-satellite tracking (SST) measurement data and lunar laser ranging data. One satellite orbits around the LLO, while the other satellite orbits around the Distant Retrograde Orbit (DRO). An inter-satellite ranging link is established between the two satellites, while the LLO satellite conducts laser ranging with a Corner Cube Reflector (CCR) on the lunar surface. Both inter-satellite ranging data and lunar laser ranging data are acquired through measurements. By integrating these data with orbital dynamics and employing the Extended Kalman Filter (EKF) method, the position and velocity states of the two formation satellites are estimated. This orbit determination scheme operates independently of ground measurement and control stations, achieving a high degree of autonomy. Simulation results demonstrate that the position accuracy of the LLO satellite can reach 0.1 m, and that of the DRO satellite can reach 10 m. Compared to the autonomous orbit determination scheme relying solely on SST measurement data, this proposed scheme exhibits several advantages, including shorter convergence time, higher convergence accuracy, and enhanced robustness of the navigation system against initial orbit errors and orbital dynamic model errors. It can provide a valuable engineering reference for the autonomous navigation of lunar-orbiting satellites. Full article

The force conditions experienced by spacecraft and astronauts in space are vastly different from those in Earth’s gravitational environment, hence it is necessary to conduct adequate micro-low-gravity environment simulation tests on the ground before launch. In this paper, an overview is provided of the current status of micro-low-gravity simulation test technology for spacecraft based on hydrostatic air-bearing. The paper systematically organizes the application of hydrostatic air-bearing technology in micro-low-gravity simulation tests, such as the deployment of space mechanisms, spacecraft GNC (Guidance, Navigation, and Control), on-orbit operations of space manipulators, and astronaut training. It summarizes the principles of air-flotation micro-low-gravity simulation technology in different scenarios and distills suitable solutions for various requirements. Finally, the paper looks forward to the development trends of air-flotation micro-low-gravity simulation test technology and proposes key technical challenges that need to be overcome in aerostatic bearing. Full article

The use of continuous-thrust propulsion systems allows spacecraft to cover complex space trajectories and to complete missions that would be difficult using chemical thrusters. Among the continuous-thrust propulsion systems proposed in recent decades, solar electric thrusters occupy an important position thanks to the maturity reached by this technology. Technological advances in the miniaturization of spacecraft components allow an electric thruster to be installed even in a small and standardized vehicle such as a CubeSat. In this context, the BIT-3 RF ion thruster is an interesting option that has been recently employed in some space missions for the study of the lunar surface. In the recent literature, the performance of a CubeSat equipped with a propulsion system based on the BIT-3 has been studied considering a simplified model in which the thrust magnitude has a fixed value or varies continuously within a prescribed range. However, the operating levels of a BIT-3 are finite in number. This paper studies the transfer performance of a BIT-3-propelled CubeSat considering the actual operating levels that can be provided by such a thruster. The work analyzes the optimal transfer towards asteroid 2000 SG344 when the electric power is obtained through solar arrays. Full article

Enhanced real-time onboard orbit determination for low-Earth-orbit satellites is essential for autonomous spacecraft operations. However, the accuracy of such systems is often limited by signal propagation delays between GPS satellites and the user spacecraft. These delays, primarily due to Earth’s rotation and ionospheric effects become particularly significant in high-dynamic LEO environments, leading to considerable errors in range and range rate measurements, and consequently, in position and velocity estimation. To mitigate these issues, this paper proposes a real-time orbit determination algorithm that applies Earth rotation correction and dual-frequency (L1 and L2) ionospheric compensation to raw GPS measurements. The enhanced orbit determination method is processed directly in the Earth-centered Earth-fixed frame, eliminating repeated coordinate transformations and improving integration with ground-based systems. The proposed method employs a reduced-dynamic orbit determination strategy to balance model fidelity and computational efficiency. A predictive correction model is also incorporated to compensate for GPS signal delays under dynamic motion, thereby enhancing positional accuracy. The overall algorithm is embedded within an extended Kalman filter framework, which assimilates the corrected GPS observations with a stochastic process noise model to account for dynamic modeling uncertainties. Simulation results using synthetic GPS measurements, including pseudoranges and pseudorange rates from a dual-frequency spaceborne receiver, demonstrate that the proposed method provides a significant improvement in orbit determination accuracy compared to conventional techniques that neglect signal propagation effects. These findings highlight the importance of performing orbit estimation directly in the Earth-centered, Earth-fixed reference frame, utilizing pseudoranges that are corrected for ionospheric errors, applying reduced-dynamic filtering methods, and compensating for signal delays. Together, these enhancements contribute to more reliable and precise satellite orbit determination for missions operating in low Earth orbit. Full article

Accurate localization is a key requirement for deep-space exploration, enabling spacecraft operations with limited ground support. Upcoming commercial and scientific missions to the Moon are designed to extensively use optical measurements during low-altitude orbital phases, descent and landing, and high-risk operations, due to the versatility and suitability of these data for onboard processing. Navigation frameworks based on optical data analysis have been developed to support semi- or fully-autonomous onboard systems, enabling precise relative localization. To achieve high-accuracy navigation, optical data have been combined with complementary measurements using sensor fusion techniques. Absolute localization is further supported by integrating onboard maps of cataloged surface features, enabling position estimation in an inertial reference frame. This study presents a navigation framework for optical image processing aimed at supporting the autonomous operations of lunar orbiters. The primary objective is a comprehensive characterization of the navigation camera’s properties and performance to ensure orbit determination uncertainties remain below 1% of the spacecraft altitude. In addition to an analysis of measurement noise, which accounts for both hardware and software contributions and is evaluated across multiple levels consistent with prior literature, this study emphasizes the impact of process noise on orbit determination accuracy. The mismodeling of orbital dynamics significantly degrades orbit estimation performance, even in scenarios involving high-performing navigation cameras. To evaluate the trade-off between measurement and process noise, representing the relative accuracy of the navigation camera and the onboard orbit propagator, numerical simulations were carried out in a synthetic lunar environment using a near-polar, low-altitude orbital configuration. Under nominal conditions, the optical measurement noise was set to 2.5 px, corresponding to a ground resolution of approximately 160 m based on the focal length, pixel pitch, and altitude of the modeled camera. With a conservative process noise model, position errors of about 200 m are observed in both transverse and normal directions. The results demonstrate the estimation framework’s robustness to modeling uncertainties, adaptability to varying measurement conditions, and potential to support increased onboard autonomy for small spacecraft in deep-space missions. Full article

The electric solar wind sail (E-sail) is a propellantless propulsion system concept based on the use of a system of very long and thin conducting tethers, which create an artificial electric field that is able to deflect the solar-wind-charged particles in order to generate a net propulsive acceleration outside the planetary magnetospheres. The radial rig of conducting tethers is deployed and stretched by rotating the spacecraft about an axis perpendicular to the nominal plane of the sail. This rapid rotation complicates the thrust vectoring of the E-sail-based spacecraft, which is achieved by changing the orientation of the sail nominal plane with respect to an orbital reference frame. For this reason, some interesting steering techniques have recently been proposed which are based, for example, on maintaining the inertial direction of the spacecraft spin axis or on limiting the excursion of the so-called pitch angle, which is defined as the angle formed by the unit vector perpendicular to the sail nominal plane with the (radial) direction of propagation of the solar wind. In this paper, a different control strategy based on maintaining the pitch angle value constant during a typical interplanetary flight is investigated. In this highly constrained configuration, the spacecraft spin axis can rotate freely around the radial direction, performing a sort of conical motion around the Sun-vehicle line. Considering an interplanetary Earth–Venus or Earth–Mars mission scenario, the flight performance is here compared with a typical unconstrained optimal transfer, aiming to quantify the flight time variation due to the pitch angle value constraint. In this regard, simulation results indicate that the proposed control law provides a rather limited (percentage) performance variation in the case where the reference propulsive acceleration of the E-sail-based spacecraft is compatible with a medium- or low-performance propellantless propulsion system. Full article