Active Formation Flying Along an Elliptic Orbit

T HE relative motion of a satellite (follower) with respect to the reference satellite (leader) in a circular orbit is described by autonomous nonlinear differential equations. The linearized equations at the origin are known as Hill–Clohessy–Wiltshire (HCW) equations [1]. The in-plane motion and the out-of plane motion are independent. The latter is a simple sinusoidal motion. The former has also periodic solutions under the so-called Clohessy–Wiltshire condition. Periodic solutions are given by sinusoidal functions, and the relative orbits are ellipses. These orbits are parametrized by two parameters representing size and phase. Because of this simple nature, relative periodic orbits are used for formation flying [2]. Periodic orbits of the nonlinear relative dynamics are also characterized by the parameters of the follower’s elliptic orbit and its inclination angle in an inertial reference frame, and initial conditions for periodic orbits are explicitly given [3]. If the reference orbit is elliptic, the equations of relative motion involve the true anomaly and the radius of the orbit, which are periodic functions. The linearized equations of motion at the origin are known as Tschauner–Hempel (TH) equations. The in-plane motion and the out-of plane motion remain independent. However, the derivation of the state transition matrix in this case is not immediate. In earlier studies [4–6], the true anomaly is used as independent variable, because this simplifies the resulting equations and the transition matrix can be obtained for them. Rendezvous problems have been extensively studied using this representation [7,8]. The out-of-plane motion is a sinusoidal motion, and the inplane motion has periodic solutions [2,9]. The condition, which generalizes the CW condition, for periodic solutions is given in [9]. Periodic orbits of the TH equations are also used for formation flying [2,9,10]. However, they are more complicated compared to ellipses of the HCW equations, and their shape changes with eccentricity. Some of them are very much distorted. Hence, it is not intuitively clear how to design relative orbits. The direct study of the relative dynamics using time as independent variable is given in [11], and a series expansion of the state transition matrix in the eccentricity is obtained in [12]. A closed form of the state transition matrix is recently derived in [13],where a brief history of elliptic rendezvous is also found. The characterization of periodic orbits of the nonlinear relative dynamics as functions of time is extended to the elliptic case, and initial conditions for periodic orbits are explicitly given [14]. The general solution of the relative dynamics based on the orbital elements is obtained in [15], and the relative motion invariant manifold, where the solution lies, is determined. This technical note is concernedwith the generation of all periodic solutions of the in-plane motion of the TH equations. For this purpose, the state transition matrix of [16], which is given as a function of the true anomaly, is used. The solution is parametrized by four constants, and the condition for periodic solution follows from this. As in the case of the HCWequations, size and phase parameters are introduced. The main contributions of this note are summarized as follows. First, the region where all periodic solutions lie is determined by two ellipses, whose size depends on the eccentricity of the leader orbit. Given size and phase parameters, a graphical method to determine the initial point of a periodic orbit at a given true anomaly is proposed. Varying the value of the true anomaly, the whole family of periodic orbits is generated. The time domain versions of the region and periodic orbits are determined by multiplying the radius of the leader orbit. This clarifies how the shape of a periodic orbit changes with eccentricity and why the shapes of some orbits are distorted. As the eccentricity goes to zero, the region of periodic motion shrinks to a single ellipse, which is a periodic orbit of the HCW equations. The size parameter gives the distance between the leader and the follower, and the phase determines the position of the follower. As in [15], the maximum and minimum distances to the leader can be calculated. Hence, our method is useful to design relative orbits with size and position specifications. This note is organized as follows. Section II reviews the equations of relativemotion along an elliptic orbit. Section III gives a geometric method to draw periodic orbits of the TH equations. Finally, Section IV gives conclusions.

Rotational Motion Control by Feedback with Minimum L1-Norm

The relative motion of chief and deputy satellites in close proximity with orbits of arbitrary eccentricity can be approximated by linearized time-periodic equations of motion. The linear time-invariant Hill–Clohessy–Wiltshire equations are typically derived from these equations by assuming the chief satellite is in a circular orbit. Two Lyapunov–Floquet transformations and an integral-preserving transformation are here presented which relate the linearized time-varying equations of relative motion to the Hill–Clohessy–Wiltshire equations in a one-to-one manner through time-varying coordinate transformations. These transformations allow the Hill–Clohessy–Wiltshire equations to describe the linearized relative motion for elliptic chief satellites.

Formation flying and mission design for Proba-3

A new second order sliding mode (2nd-SM) control algorithm of MIMO nonlinear systems for low thrust multiple picosatellites formation flying is proposed in this paper. The problem is tackled by applying chatteringfree 2nd-SM control law as a control scheme for the case where the leader satellite is in circular and elliptical reference orbits under J2 perturbation. The asymptotic convergence to the 2nd-SM and the robust stability of the asymptotic 2nd-SM control system are for the first time proved with Lyapunov functions. Several numerical results are presented to demonstrate the efficacy of the proposed controllers in maintaining the satellite formation in presence of initial formation errors and J2 perturbation effects that tend to disperse the formation. Simulation results confirm that the suggested methodologies yield high formation keeping precision and effectiveness in ensuring formation maneuvering with thrusters faults of formation flying systems for picosatellites. Using this 2nd-SM method the fuel consumption is lower than in most other works.

Revisiting the general periodic relative motion in elliptic reference orbits

This paper studies the relative orbit geometry of a leader–follower spacecraft formation flying in unperturbed elliptical reference orbits. The first-order relative position equations, derived using the reference orbital element approach under the condition that the follower and leader spacecraft have equal semimajor axes, are transformed fromtrigonometricformstoparametricandalgebraicforms.Theconditionsforandthenumberofself-intersections oftherelativeorbitprojectedontothethreecoordinateplanesoftheleaderlocal-vertical–local-horizontal frameare obtained. The relative orbit proves to be three-dimensional instead of planar in most cases, and may self-intersect spatially at most once. The collision between the follower and leader spacecraft corresponds to the case where the solution curve to the first-order relative motion equations passes through the origin. The conditions for collision are subsequently determined. For anondegenerate case(in which none of the relative motion in the radial, in-track, and cross-trackdirectionsvanish),threetypesofrelativeorbitarepossible.Mostfrequently,therelativeorbitisonaonesheet hyperboloid. Otherwise, when the relative orbit has a real or finite imaginary self-intersection, it rests on an elliptic cone. In the rest of the cases, including that with an imaginary self-intersection at infinity, the relative orbit is on an elliptic cylinder. The criteria for these three types are given, respectively, followed by examples.

I N RECENT years, many proposed satellite missions have undergone a paradigm shift away from large, expensive spacecraft toward the use of smaller satellites flying in formation. Although formation flying offers a number of advantages, such as robustness of the mission to failure and potential cost savings, the benefits are partially offset by the increased complexity of the guidance, navigation, and control systems and by V-limited mission durations. The development of advantageous relative dynamics between the constituent satellites could help extend mission durations by minimizing fuel usage. In particular, periodic or quasiperiodic relative orbits would reduce the control effort necessary to maintain the formation and significantly prolong the mission’s lifetime. Attempts to identify naturally periodic relative orbits can be found throughout the current literature. Vaddi et al. [1] modified initial conditions from the Hill–Clohessy–Wiltshire (HCW) equations to enforce bounded relative motion in the presence of small eccentricity and second-order differential gravity terms. Kasdin and Kolemen [2] solved the Hamilton–Jacobi equation in terms of epicyclic orbital elements to derive bounded, periodic orbits in the presence of higher-order gravity and certain perturbations. Schaub and Alfriend [3] identified the conditions for J2-invariant relative orbits bymatching the secular drift rates of themean orbital elements of the chief and deputy satellites. In most of these analytical cases, however, the authors established precise periodic motion using approximated dynamical models. When higher fidelity dynamics and realistic orbital parameters are used, these methods break down and exhibit relative orbit drift or a high sensitivity to initial condition errors. Several numerical approaches to the problem of relative orbit periodicity have also been attempted. Sabatini et al. [4] and Izzo and Sabatini [5] used a refined genetic algorithm to optimize relative trajectories for maximum periodicity in the presence of J2 perturbations. Quasi-periodic orbits were obtained for two sets of “magic” inclinations (49.11 and 63.43 deg), at which resonance between the formation’s inand out-of-plane motion results in projected circular orbit (PCO) formations with very small orbital drift. Damaren [6] formulated an iterative shooting approach based on the Newton method to close the relative orbit and achieve “almost” periodic initial conditions. This technique requires a low-level state-feedback control loop to track a trajectory based on the HCW equations, but is robust to initial condition errors and exhibits very low drift characteristics in circular orbits. A similar approach was taken by Becerra et al. [7], who applied a nonlinear Hamiltonian model to the Newton method to obtain quasi-periodic relative initial conditions of the deputy in the presence of the J2 perturbation. A linear quadratic regulator (LQR) controller was used to track reference trajectories that were developed from sinusoidal functions of time and based on a Keplerian orbit. As yet, a method of numerically designing reference trajectories that match or closely match the natural perturbed relative motion of the deputy satellite in a generic orbit, and that are robust to initial condition errors, has not been developed. Such trajectories, with a properly designed low-authority controller, could be tracked for minimal V requirements, thus enabling formation flying missions to be significantly extended. The intent of this study is to design such trajectories by continuing the development of the methods used to search for almost-periodic orbits, begun in [6], and applying them to the J2-invariant trajectories in [3]. The resulting relative orbits will approach true periodicity by capitalizing on advantageous initial conditions, numerically designed reference trajectories, and optimal control strategies.

A computationally efficient algorithm is developed for onboard planning of -impulse fuel-optimal maneuvers for establishment and reconfiguration of spacecraft formations. The method is valid in circular and elliptic orbits and includes first-order secular effects. The dynamics are expressed in terms of differential mean orbital elements, and relations are provided to allow the formation designer to transform these into intuitive geometric quantities for visualization and analysis. The maneuver targeting problem is formulated as an optimal control problem in both continuous and discrete time. The continuous-time formulation cannot be solved directly in an efficient manner, and the discrete-time formulation, which has an analytical solution, does not directly yield the optimal thrust times. Therefore, a new flight-suitable algorithm is designed by iteratively solving the discrete-time formulation while using the continuous-time necessary conditions to refine the thrust times until they converge to the optimal values. Simulation results illustrate the performance for a variety of reconfiguration maneuvers and reference orbits, including examples for the NASA CubeSat Proximity Operations Demonstration mission.

Collision Probability with Gaussian Mixture Orbit Uncertainty

Chapter 4 - Application of time domain collocation in formation flying of satellites

In this paper, the analytical closed-form solution of the Tschauner-Hempel equations that are used for rendezvous in elliptic orbits are studied and exploited. The solution is based on a state transition matrix that is completely explicit in time. Autonomous guidance algorithms to approach, to flyaround, and to depart from a target vehicle in elliptic orbits are exploited based on that solution. The algorithms are general and able to translate the chaser vehicle in any direction, decelerate while approaching the target vehicle, and accelerate when moving away. These guidance algorithms are based on the glideslope used in the past for rendezvous and proximity operations of the Space Shuttle with other vehicles. The algorithms are applied to the problem of terminal rendezvous near any Keplerian elliptic orbit in a Newtonian gravitational field showing the effects of target orbit eccentricity on the relative motion. These algorithms extend and generalize previously published solutions. Numerical simulations confirm the brevity and accuracy of the general solutions developed in the current paper.

Linear Equations of Relative Motion

Introduction E LECTRODYNAMIC tethers are expected to be a viable and efficient means for generating power or controlling the orbit of future spacecraft.1−5 For cases where the tether current is not explicitly controlled for orbital maneuvering, then the long-term dynamic behavior of the tether librations must be controlled. The control of electrodynamic tethers on inclined orbits is important because of the inherent instability provided by the electric current.6 It has been conclusively established that, even for a tether modeled as an inelastic rod, the variation in electrodynamic forces over an orbit causes an influx of energy into the libration dynamics. Although, for the most part, the instability is slow to develop, it must be controlled for longterm operations. There exists a unique libration cycle, however, for which the net energy input per orbit is zero. Such a libration cycle is a periodic solution to the equations of motion for the system.7 The periodic solution is unstable, with instability growing with the electric current.6 Periodic solutions have also been determined for tethers including the lateral dynamics.8 The instability in the lateral dynamics grows at a faster rate than the librational instability. The effect of orbit eccentricity was shown by Pelaez and Andres9 to amplify the unstable nature of the dynamics. Corsi and Iess10 considered a simple control scheme to stabilize the tether librations for deorbiting applications. Hoyt11 has described schemes for stabilizing flexible electrodynamic tethers during deorbit by periodically measuring the motion of several points along the tether to estimate the oscillation energy, and adjusting the control input accordingly. The application of electromagnetic forces for directly controlling the librational motion has also been considered.12−15 Control of the unstable skip-rope motion during retrieval was shown to be effectively suppressed by Williams et al.15 if a movable attachment point is used to damp the lateral oscillations of the electrodynamic tether. In Ref. 16, Pelaez and Lorenzini addressed the issue of controlling the librations of an electrodynamic tether around the periodic solution. Their approach consisted of two linear feedback control schemes. The first was simple proportional feedback for the inand out-of-plane libration rates, and the second was a delayed-feedback version of the same controller. However, the implementations of these controllers were assumed to be independent, where it was suggested that control could be achieved using movement of the

The problem of constructing and classifying elliptic solutions of nonlinear differential equations is studied. An effective method enabling one to find any elliptic solution of an autonomous nonlinear ordinary differential equation is described. The method does not require integrating additional differential equations. Much attention is being paid to the case of elliptic solutions with several poles in a parallelogram of periods. With the help of the method we find elliptic solutions up to the fourth order inclusively of an ordinary differential equation with a number of physical applications. The method admits a natural generalization and can be used to find elliptic solutions satisfying systems of ordinary differential equations.