The Journal of the Astronautical Sciences

A robust yet affordable autonomous orbit determination system must be in place as the traffic of spacecraft in cislunar space increases. The methods of autonomous orbit determination analyzed in this paper are the use of Global Navigation Satellite System signals and horizon-based optical navigation with the Moon. This paper details a simulation of cislunar space with various non-Keplerian orbits around the Moon, including L1/L2 Halo orbit families, Distant Retrograde Orbits, and Butterfly Orbits. The orbit determination algorithm uses an Extended Kalman Filter and Monte Carlo simulations are conducted to characterize the performance of the algorithm in different orbits and with different measurement types. Results from this paper emphasize the importance of considering the orbit’s geometry for optical navigation-only measurements, although combining the two measurements significantly relieves this limitation in most types of near-Moon orbits with smaller root-mean-square errors of position and clock bias/drift estimates.

We compute periodic orbits in the Artificial Satellite Theory. We fix our attention on the Earth case where we considered a zonal gravitational model for the potential. In a rotating frame attached to the Earth we have found sixteen polar orbits that are periodic after different number of revolutions. The altitude of each orbit remains almost constant and varies from about 263 km for the sixteen-revolutions orbit to about 35,830 km for the one-revolution orbit. All the orbits are linearly stable and their period is exactly 24 hours.

This paper presents a new nonlinear control methodology for slewing spacecraft, which provides both necessary and sufficient conditions for stability by identifying the stability boundaries, rigid body modes, and limit cycles. Conservative Hamiltonian system concepts, which are equivalent to static stability of airplanes, are used to find and deal with the static stability boundaries: rigid body modes. The application of exergy and entropy thermodynamic concepts to the work-rate principle provides a natural partitioning through the second law of thermodynamics of power flows into exergy generator, dissipator, and storage for Hamiltonian systems that is employed to find the dynamic stability boundaries: limit cycles. This partitioning process enables the control system designer to directly evaluate and enhance the stability and performance of the system by balancing the power flowing into versus the power dissipated within the system subject to the Hamiltonian surface (power storage). Relationships are developed between exergy, power flow, static and dynamic stability, and Lyapunov analysis. The methodology is demonstrated with two illustrative examples: (1) a nonlinear oscillator with sinusoidal damping and (2) a multi-input-multi-output three-axis slewing spacecraft that employs proportional-integral-derivative tracking control with numerical simulation results.

This paper presents an on-board guidance and targeting design that enables explicit state and thrust vector control and on-board targeting for planetary descent and landing. These capabilities are developed utilizing a new closed-form solution for the constant thrust arc of the braking phase of the powered descent trajectory. The key elements of proven targeting and guidance architectures, including braking and approach phase quartics, are employed. It is demonstrated that implementation of the proposed solution avoids numerical simulation iterations, thereby facilitating on-board execution of targeting procedures during the descent. It is shown that the shape of the braking phase constant thrust arc is highly dependent on initial mass and propulsion system parameters. The analytic solution process is explicit in terms of targeting and guidance parameters, while remaining generic with respect to planetary body and descent trajectory design. These features increase the feasibility of extending the proposed integrated targeting and guidance design to future cargo and robotic landing missions.

This paper deals with a modified version of the Circular Restricted Three-Body Problem (CR3BP). In this version, the additional effect of a three-body interaction is taken into account. In particular, we examine numerically the result of this interaction on the evolution of the well-known family of Sitnikov motion of CR3BP as well as that on the families of 3D periodic orbits bifurcating from this family.

In recent years, the sustained growth of space launch missions has inevitably led to a rapid increase in space debris, which has attracted extensive attention to the capture and removal of space debris. This paper presents a new strategy for detumbling a flexible tumbling target in the post-capture phase using a flexible-base space robot (also called a chaser), which involves trajectory optimization and composite control of the chaser. Based on the dynamic equation of the combined system, the trajectories of the chaser base and the manipulator’s joints are parameterized by quintic polynomial curves, and then are simultaneously optimized for the collision avoidance of the combined system and the reduction of panel deformation, control energy, and mission duration. The Pareto-optimal solutions of the optimization problem are obtained by the multi-objective particle swarm optimization (MOPSO) algorithm. The composite control scheme, composed of a trajectory tracking controller and a vibration suppression controller, is used for the chaser to detumble the target along the planned trajectories, stabilize its base attitude, and eliminate the residual vibration of the flexible panels. Two representative cases are used to verify the effectiveness and robustness of the proposed detumbling strategy for the target with parameter uncertainties.

A complete and careful foundation is presented for maximum-likelihood attitude estimation and the calculation of measurement sensitivity matrices with the intent of revealing heretofore undisclosed pitfalls associated with unconstrained quaternion estimation. Efficient formulas are developed for computing the measurement sensitivity matrix for any attitude representation for which an efficient formula for the inverse kinematic equation is known. In particular, it is shown that the measurement sensitivity matrix for the quaternion is ambiguous and may take on a wide range of values. Hence, estimates of a quaternion which do not take correct account of the norm constraint will be physically meaningless. It is shown also that within Maximum Likelihood Estimation the form of the Wahba cost function for attitude estimation is incorrect when the attitude constraint is relaxed. A simple physical example is presented for quaternion estimation from noise-free vector measurements which fails when the norm constraint on the quaternion is relaxed. Part I of this work provides the basis for more detailed investigations of unconstrained attitude estimation in Part II [1].

Several commercial organizations have recently launched or plan to launch constellations containing thousands of satellites. Such large constellations potentially adversely affect astronomical observations. This study formulates a set of indicators that assess the impact of light pollution from different constellations on ground-based visible band astronomy. These include the statistically expected number of visible and sunlit satellites above ground-based observers, as well as the number that are also expected to be brighter than the currently recommended limit for constellation satellites. The latter indicator provides a consolidated means to evaluate the potential for a constellation to affect ground-based astronomy too severely, by simultaneously accounting for the effects of constellation population, orbital distribution as well as brightness magnitude and variability. For existing constellations, the evaluation process incorporates actual satellite photometric brightness measurements, which are becoming increasingly available in web-accessible databases and repositories. For proposed constellations, a semi-empirical method allows rough approximations of pre-launch light pollution levels, based on observed brightness distributions observed of currently orbiting analog satellites.