Joseph Shoer

Flux-Pinned Interfaces for the Assembly, Manipulation, and Reconfiguration of Modular Space Systems

Non-contacting interactions between permanent magnets and superconductors known as “flux pinning” provide a novel way to fix many modular space systems in desired relative positions and orientations, from space stations to close-proximity formations. When cooled appropriately, these flux-pinned interfaces require no power or active control and very little mass but provide very high mechanical stiffness (>200 N/m for a few hundred grams of material) and damping (2% of critical) between modules, making the technology ideal for in-orbit assembly applications. We describe new measurements and simulations to characterize these values for spacecraft applications. Flux-pinned interfaces have so far achieved inter-module separations in the 8–10 cm range with ∼100 g of mass on each module, with the prospect of larger separations. We also discuss several means to actuate the noncontacting couplers, which is a first step toward the development of devices for the noncontacting manipulation and reconfiguration of modular space systems.

Reconfigurable Spacecraft as Kinematic Mechanisms Based on Flux-Pinning Interactions

T HE challenges of the space environment require that spacecraft exhibit a high degree of mission assurance, which often takes the form of autonomous fault tolerance [1,2]. However, the technology for spacecraft repair and reconfigurationmissions has not yet matured to the point where autonomous operations are also robust. Many completed and envisioned spacecraft reconfiguration or repair techniques involve substantial human-in-the-loop activity, including Advanced X-Ray Astrophysics Facility servicing activities and Hubble Space Telescope repair and expansion missions [3]. Another example is the construction of the International Space Station, which astronauts have assembled and reconfigured piecemeal during many hours of extravehicular activity. Fully autonomous repair and reconfiguration tasks have been achieved only recently, such as on the Orbital Express mission [4], with extensive sensing and active control solutions [5]. We propose to depart significantly from traditional approaches to reconfiguration by treating modular, reconfigurable spacecraft as kinematic mechanisms. This proposal addresses the need for robust reconfiguration techniques in space without treating the problem of reconfiguration as one of docking or formation flight. In so doing, its approach incorporates passively stable physics, involving little to no active control at the level of the interface between modules and focusing on architectural control of the system start and end states. This methodology blurs the distinction between deployments and reconfigurations, where “reconfiguration” refers to changes in spacecraft structure, and possibly function, at the system level. There are many reasons why reconfigurable spacecraft architectures are desirable: this capability may allow advantageous changes to a spacecraft’s mass properties, adapt the spacecraft to changing mission requirements, or enable autonomous self-repair. The reconfiguration of modular or swarm spacecraft is currently studied in this context [6,7]. Traditional approaches view these operations as problems in active control [8], particularly as extensions of rendezvous and docking tasks [7,9]. However, we suggest that spacecraft achieve reconfiguration without detaching from one another, by forming appropriate kinematicmechanisms. The degrees of freedom of these mechanisms lead to deterministic reconfiguration behaviors in the sense that the spacecraft can be engineered such that the possiblemotions and configurations of the system are limited and a desired subset of possible configurations becomes a stable reconfiguration sequence. This mitigates the risk of instability from active control, as actively controlled equilibria may not be trivial to define, create, identify, or exploit, and the stable subset of configurations may not be as susceptible to unmodeled dynamics or interactions between the controller and system dynamics. These systems also offer the potential advantage of low power and fuel consumption, saving spacecraft mass. One enabling technology that provides an opportunity for spacecraft reconfiguration via this kinematic-mechanism paradigm is the flux-pinned interface (FPI) [10,11]. This interface links spacecraft modules with high stiffness and damping. It does so through an action-at-a-distance force provided by interactions between magnetic fields and type II superconductors such as yttrium barium copper oxide (YBCO). The flux-pinning interaction is passively stable and acts in multiple degrees of freedom (DOF). Two features of an FPI are central to the formation of flux-pinned mechanisms. First, FPIs can allow certain DOF to remain unconstrained by pinning forces [12], enabling the formation of fluxpinned joints. Second, the operation of an FPI depends on the applied field and the thermal state of the superconductor; manipulating the temperature of the YBCO (for instance, by extending or retracting a sunshade) or temporarily superimposing another magnetic field therefore toggles the interface on and off. Concurrent work is in progress to characterize and design flux-pinned interface systems [13,14]. Section II of this Note identifies some mathematical tools for treating reconfigurable spacecraft systems as kinematicmechanisms. This Note develops its treatment in the particular context of a reconfigurable system using FPI technology, but the kinematicmechanism concept may be extended to other architectures as well. Section III then provides some examples of spacecraft reconfiguration through kinematic-mechanism techniques. The examples include detailed descriptions of two reconfiguration sequences for changing the order and relative orientation of a line of modules, as well as an air-table demonstration of a simple FPI-based reconfiguration.

Microgravity Demonstrations of Flux Pinning for Station-Keeping and Reconfiguration of CubeSat-Sized Spacecraft

M AGNETIC flux pinning, a noncontacting interaction between Type II superconductors andmagneticfields, has been studied at length by the scientific community for its applications to levitating objects in a 1g environment [1–3]. However, due to the unpowered passive stability that flux pinning can provide, it also has many potential applications for the assembly and reconfiguration of modular space structures [4] and spacecraft formations [5]. Current approaches to autonomous docking of space vehicles [6,7], aswell as spacecraft reconfiguration and formation flying [8–10], rely heavily on active controllers. However, a permanent magnet flux-pinned to a superconductor experiences a passive restoring force that attracts it to the position and orientation it held when the superconductor first cooled below its critical temperature. Previous work and laboratory experiments have suggested that this passively stable effect provides sufficient stiffness and damping to bind modular spacecraft together over separation distances up to about 10 cm [11]. Several possible means for actuation of a flux-pinning interface may be superimposed on the passive stability of this interaction, which requires no power to the superconductor except that required for cooling. In addition to actuation by time-varyingmagnetic fields such as those from electromagnet coils [12], aflux-pinned space systemcan exploit symmetries in the pinned magnetic field to form a noncontacting kinematic mechanism in which the modular components do not touch one another but have some specified kinematic degrees of freedom (DOF) [13]. A simple noncontacting mechanism consisting of a single revolute joint on an air-table testbed has been demonstrated in a laboratory setting [14]. This note reports the results of two demonstrations of magnetic flux-pinning technologies implemented onCubeSat-sized spacecraft during microgravity flights as part of the NASA Glenn Research Center Facilitated Access to the Space Environment for Technology Development and Training (FAST) program in August 2009. In the first experiment, a CubeSat mockup was flux pinned to a CubeSatscale vehicle carrying superconductors and was expected to demonstrate low-stiffness, noncontacting, passive station-keeping in 6 degrees of freedom (6 DOF). The second experiment studied the reconfiguration of two CubeSat mockups between equilibrium configurations via a revolute joint formed by a flux-pinned noncontacting kinematic mechanism. It was expected that the spacecraft would move about an axis defined by the flux-pinned interface rather than their respective centers of mass. These microgravity flight results highlight the role magnetic flux pinning might play in future small satellite operations. Each experiment was performed on a microgravity aircraft with two free-floating modules: one containing an array of magnets appropriate to the experiment, and the other containing superconductors in a Dewar of liquid nitrogen. Three experimenters participated in each flight, two equipment managers to monitor the position of the free-floating modules at all times, and one data collector who operated the motion-capture camera. Figure 1 is a diagram of the test setup.

Demonstration of a Magnetic Locking Flux-Pinned Revolute Joint for Use on CubeSat-Standard Spacecraft

Magnetic flux pinning is an interaction between strong magnets and certain superconductors that causes a damped, non-contacting equilibrium to form, connecting the flux-pinned objects. This interaction has been proposed for use in establishing a stable formation of spacecraft modules that is resistant to disturbances. Although flux pinning can exert forces in all six degrees of freedom, a flux-pinned interface can be designed to constrain only certain degrees of freedom so that it functions as a non-contacting kinematic joint. One such joint consists of a superconductor flux pinned to a cylindrical magnet and free to move around the magnet’s axis of symmetry. Such an interface would serve as a revolute joint that allows two modular spacecraft to reconfigure. This paper explores the development of one such joint compatible with the CubeSat standard. We extend the functionality of the revolute joint by introducing electromagnets that create two stable equilibrium states in a system of two modules. The electromagnets also provide the means of reconfiguration between the two states, eliminating the need for reaction wheels, thrusters, or other conventional actuators for this maneuver. Finally, this paper discusses future testing for flux-pinned joints and ongoing work on an in-orbit demonstration.

A Flux-Pinned Magnet-Superconductor Pair for Close- Proximity Station Keeping and Self-Assembly of Spacecraft

We propose that spacecraft modules be constructed with interfaces consisting of combinations of magnets and Type II superconductors, establishing a non-contacting interaction between the modules thanks to magnetic flux pinning. This stable action-at-adistance interaction overcomes the limit Earnshaw’s Theorem places on other spacecraft positioning strategies involving electromagnetic fields, allowing fractionated or modular spacecraft to fix their relative positions and orientations without any mechanical connection, active control, or power expenditure. We report two experiments investigating the mechanical properties of the magnetostatic interaction to evaluate its utility. First is an experiment to find the 6DOF linear restoring forces and torques on a flux-pinned magnet and superconductor for small displacements from a set of static equilibria. Second is an transient experiment that permits the use of system identification techniques to characterize the modal damping and stiffness. Our results indicate that flux pinning is promising for modular spacecraft assembly and station-keeping applications, providing mechanical stiffnesses over 200 N/m at small (5 mm) magnet-superconductor separations and potentially useful nonzero stiffnesses at larger (over 3 cm) separations, with significant damping. We find that increasing the magnetic flux density at the superconductor surface strengthens the flux pinning forces, suggesting the possibility that higher stiffness can be obtained over larger distances by increasing or focusing the magnetostatic field.

Simulation of Multibody Spacecraft Reconfiguration through Sequential Dynamic Equilibria

This paper describes a novel hybrid-control strategy to reconfigure multibody spacecraft from one shape to another in such a way that passively stable system dynamics enable both low control effort and a high degree of robustness. This approach treats reconfigurable spacecraft systems as multibody kinematic mechanisms with controllable kinematics and takes advantage of ambient force fields in the space environment (gravity gradient, magnetism, etc.) along with passively generated, non-contacting forces on the spacecraft (such as those from permanent magnets) to drive the reconfiguration maneuver to one stable dynamic equilibrium after another, in sequence. The use of kinematic constraints and passive dynamics adds robustness, while the stepwise nature of the reconfiguration maneuver provides many safe-hold points for verification regardless of transient dynamics. The focus on kinematic constraints lends itself well to Udwadia and Kalaba’s technique for generating equations of motion. This work details the augmentation of the Udwadia-Kalaba equations with quaternion states and Euler’s equation for fully 3D rigid body motions, as well as the development of a simulation environment and computational tools for exploring sequential-equilibrium reconfigurations.

Sequences of Passively Stable Dynamic Equilibria for Hybrid Control of Reconfigurable Spacecraft

This paper proposes a novel hybrid-control strategy to reconfigure modular spacecraft. This strategy utilizes passively stable system dynamics, which have the benefits of low control effort and a high degree of robustness. This approach treats reconfigurable spacecraft systems according to the theory of multibody kinematic mechanisms. It employs ambient force fields in the space environment (gravity gradient, magnetism, etc.), along with passively generated, non-contacting force fields on the spacecraft (such as those from permanent magnets), to drive the reconfiguration maneuver. The control strategy for reconfiguration in this paradigm consists of a selection of body incidences, joint Jacobians, and applied force fields that cause the multibody spacecraft system to evolve through passive dynamics to a new configuration. Many of these passive dynamical evolutions, chained together in stepwise fashion, allow the system to reach many possible desired configurations. All possible equilibrium configurations can be computed offline and uploaded to the spacecraft in a graph structure. The use of kinematic constraints and passive dynamics adds robustness to the system, while the stepwise nature of the reconfiguration maneuver provides many safe-hold points for verification regardless of transient dynamics.

Dual exploration architectures for breaking the decades-long cycles of planetary science

The Dual Exploration Architecture is a mission concept that combines remote sensing and in-situ observations into a single mission to answer planetary science questions that can only be answered with both types of data. Adoption of dual exploration architectures may short circuit the long, slow cycle of missions to inaccessible bodies by eliminating the need for separate precursor and follow-up missions. Additionally, the dual architecture possesses inherent flexibility that enables the design of adaptive, event-driven missions. Five key observations about the state and trends of planetary science exploration lead us to the dual architecture: increasing complexity of observations, scarcity of future missions, desire to capture transitory events, continued miniaturization of components, and the Mars exploration cycle. A study of the planetary science the decadal survey reveals broad applicability of dual missions to solve mysteries that cannot be answered with a traditional mission architecture. These missions fall into three classes: choosing a local target from a global survey, dynamic/reactive science, and global in-situ networks. These mission classes further reveal four technology development needs that must be addressed for dual missions: passive landers, guided atmospheric probes, robust sensing packages, and small, precise orbital instruments. This study pursues a specific focus on two examples of such technologies: the ChipSat and cold atom gravimetry. Using these technologies, we study an example dual architecture mission to both characterize and sample the subsurface oceans at Europa. The identification of regions with thin ice precedes the selection of surface targets and dispatching of probes to those targets.

Attitude determination and control system design for the CYGNSS microsatellite

This paper presents the development of the attitude determination and control system design of the Cyclone Global Navigation Satellite System spacecraft. The CYGNSS constellation consists of eight small satellite observatories in 500 km circular orbits at an inclination of 35 deg released from a single launch platform. Each CYGNSS spacecraft will make frequent and accurate measurements of ocean surface winds throughout the life cycle of tropical storms and hurricanes with the objective to fundamentally improve gap-free coverage for hurricane forecast and monitoring. Realising this objective requires the spacecraft to accurately and reliably point its signal collection antennae in desired directions and hold its Earth relative attitude over long time durations to prescribed knowledge and point requirements. Indeed, the microsatellite ADCS regulates all spacecraft estimation and control functionality spanning detumbling; Sun acquisition and hold; pointing control and momentum-management over the micro-satellite lifetime to design requirements. This paper presents the ADCS hardware, software and algorithms used to control the spacecraft in all phases of CYGNSS operations and presents simulation based performance results of the closed-loop estimation and control systems.

Conical scanning approach for Sun pointing on the CYGNSS microsatellite

The small scales of area, volume, and power of small spacecraft, such as NASAs 25-kg Cyclone Global Navigation Satellite System (CYGNSS) satellites, constrain the number of independent subsystems that they can support. Consequently, small satellites often require novel approaches to execute the same mission functions that a larger satellite can easily perform with familiar sensor, actuator and algorithm options. In the case of CYGNSS, the spacecraft must execute a Sun acquisition and pointing phase but the actuator suite does not include 2-axis sun sensors or rate gyros; two measurements that seem like obvious inclusions for the Sun acquisition task. Instead, during Sun acquisition, the CYGNSS attitude control system uses a limited actuator and sensor set consisting of three magnetic torque rods, a three-axis magnetometer, and Sun incidence-angle measurements from three solar panel faces. This paper describes the sensing and control algorithms implemented in CYGNSS flight software to acquire and maintain Sun pointing with the available measurements and actuators. The Sun pointing algorithm uses a conical scanning approach based on traditional RF pointing and target-tracking systems, which consists of two key control loops: (1) a rate loop, which initiates a body spin about the solar-array face axis, and (2) a slower angle controller that tracks the array power gradients measured over the course of the fast spin. A slew toward the peak power eventually drives the solar panel face normal to spin in a cone centered about the Sun vector. The Sun acquisition process has a large convergence basin, is stable in the Lyapunov sense, and demonstrates excellent performance behavior in simulation.