Astrodynamics最新文献

Space debris have become exceedingly dangerous over the years as the number of objects in orbit continues to increase. Active debris removal (ADR) missions have gained significant interest as effective means of mitigating the risk of collision between objects in space. This study focuses on developing a multi-ADR mission that utilizes controlled reentry and deorbiting. The mission comprises two spacecraft: a Servicer that brings debris to a low altitude and a Shepherd that rendezvous with the debris to later perform a controlled reentry. A preliminary mission design tool (PMDT) was developed to obtain time and fuel optimal trajectories for the proposed mission while considering the effect of J₂, drag, eclipses, and duty cycle. The PMDT can perform such trajectory optimizations for multi-debris missions with computational time under a minute. Three guidance schemes are also studied, taking the PMDT solution as a reference to validate the design methodology and provide guidance solutions to this complex mission profile.

Long-term configuration stability is essential for an interferometric detection constellation (IDC), which is closely related to initial uncertainty. Therefore, it is vital to evaluate the uncertainty and characterize the configuration stability. In this study, an analytical method was developed for the configuration uncertainty propagation of a geocentric triangular IDC. The angular momentum and the argument latitude were found to be significantly affected by the initial uncertainty and were selected as the core variables. By averaging the perturbation in one revolution, an analytical solution was proposed for propagating the core orbital elements in one revolution. Subsequently, the analytical solution of the orbit elements during the mission period is obtained by multiplying the solutions in iterative revolutions. The relationship between the selected orbital elements and the configuration stability parameters was established using an analytical solution. The effects of the initial uncertainty in different directions on the configuration and stable domains were studied. Simulations show that the developed method is highly efficient and accurate in predicting the configuration stability. The relative error with respect to the Monte Carlo simulations was less than 3% with a time consumption of 0.1%. The proposed method can potentially be useful for constellation design and stability analysis.

Relative navigation is a key enabling technology for space missions such as on-orbit servicing and space situational awareness. Given that there are several special advantages of space relative navigation using angles-only measurements from passive optical sensors, angles-only relative navigation is considered as one of the best potential approaches in the field of space relative navigation. However, angles-only relative navigation is well-known for its range observability problem. To overcome this observability problem, many studies have been conducted over the past decades. In this study, we present a comprehensive review of state-of-the-art space relative navigation based on angles-only measurements. The emphasis is on the observability problem and solutions to angles-only relative navigation, where the review of the solutions is categorized into four classes based on the intrinsic principle: complicated dynamics approach, multi-line of sight (multi-LOS) approach, sensor offset center-of-mass approach, and orbit maneuver approach. Then, the flight demonstration results of angles-only relative navigation in the two projects are briefly reviewed. Finally, conclusions of this study and recommendations for further research are presented.

Relative navigation is crucial for spacecraft noncooperative rendezvous, and angles-only navigation using visible and infrared cameras provides a feasible solution. Herein, an angles-only navigation algorithm with multisensor data fusion is proposed to derive the relative motion states between two noncooperative spacecraft. First, the design model of the proposed algorithm is introduced, including the derivation of the state propagation and measurement equations. Subsequently, models for the sensor and actuator are introduced, and the effects of various factors on the sensors and actuators are considered. The square-root unscented Kalman filter is used to design the angles-only navigation filtering scheme. Additionally, the Clohessy—Wiltshire terminal guidance algorithm is introduced to obtain the theoretical relative motion trajectories during the rendezvous operations of two noncooperative spacecraft. Finally, the effectiveness of the proposed angles-only navigation algorithm is verified using a semi-physical simulation platform. The results prove that an optical navigation camera combined with average accelerometers and occasional orbital maneuvers is feasible for spacecraft noncooperative rendezvous using angles-only navigation.

Real-time guidance is critical for the vertical recovery of rockets. However, traditional sequential convex optimization algorithms suffer from shortcomings in terms of their poor real-time performance. This work focuses on applying the deep learning-based closed-loop guidance algorithm and error propagation analysis for powered landing, thereby significantly improving the real-time performance. First, a controller consisting of two deep neural networks is constructed to map the thrust direction and magnitude of the rocket according to the state variables. Thereafter, the analytical transition relationships between different uncertainty sources and the state propagation error in a single guidance period are analyzed by adopting linear covariance analysis. Finally, the accuracy of the proposed methods is verified via a comparison with the indirect method and Monte Carlo simulations. Compared with the traditional sequential convex optimization algorithm, our method reduces the computation time from 75 ms to less than 1 ms. Therefore, it shows potential for online applications.

The dynamics of a spacecraft propelled by a continuous radial thrust resembles that of a nonlinear oscillator. This is analyzed in this work with a novel method that combines the definition of a suitable homotopy with a classical perturbation approach, in which the low thrust is assumed to be a perturbation of the nominal Keplerian motion. The homotopy perturbation method provides the analytical (approximate) solution of the dynamical equations in polar form to estimate the corresponding spacecraft propelled trajectory with a short computational time. The accuracy of the analytical results was tested in an orbital-targeting mission scenario.

Space object observation requirements and the avoidance of specific attitudes produce pointing constraints that increase the complexity of the attitude maneuver path-planning problem. To deal with this issue, a feasible attitude trajectory generation method is proposed that utilizes a multiresolution technique and local attitude node adjustment to obtain sufficient time and quaternion nodes to satisfy the pointing constraints. These nodes are further used to calculate the continuous attitude trajectory based on quaternion polynomial interpolation and the inverse dynamics method. Then, the characteristic parameters of these nodes are extracted to transform the path-planning problem into a parameter optimization problem aimed at minimizing energy consumption. This problem is solved by an improved hierarchical optimization algorithm, in which an adaptive parameter-tuning mechanism is introduced to improve the performance of the original algorithm. A numerical simulation is performed, and the results confirm the feasibility and effectiveness of the proposed method.

This study investigated periodic coupled orbit—attitude motions within the perturbed circular restricted three-body problem (P-CRTBP) concerning the perturbations of a radiated massive primary and an oblate secondary. The radiated massive primary was the Sun, and each planet in the solar system could be considered an oblate secondary. Because the problem has no closed-form solution, numerical methods were employed. Nevertheless, the general response of the problem could be non-periodic or periodic, which is significantly depended on the initial conditions of the orbit-attitude states. Therefore, the simultaneous orbit and attitude initial states correction (SOAISC) algorithm was introduced to achieve precise initial conditions. On the other side, the conventional initial guess vector was essential as the input of the correction algorithm and increased the probability of reaching more precise initial conditions. Thus, a new practical approach was developed in the form of an orbital correction algorithm to obtain the initial conditions for the periodic orbit of the P-CRTBP. This new proposed algorithm may be distinguished from previously presented orbital correction algorithms by its ability to propagate the P-CRTBP family orbits around the Lagrangian points using only one of the periodic orbits of the unperturbed CRTBP (U-CRTBP). In addition, the Poincaré map and Floquet theory search methods were used to recognize the various initial guesses for attitude parameters. Each of these search methods was able to identify different initial guesses for attitude states. Moreover, as a new innovation, these search methods were applied as a powerful tool to select the appropriate inertia ratio for a satellite to deliver periodic responses from the coupled model. Adding the mentioned perturbations to the U-CRTBP could lead to the more accurate modeling of the examination environment and a better understanding of a spacecraft’s natural motion. A comparison between the orbit-attitude natural motions in the unperturbed and perturbed models was also conducted to show this claim.