How Might One Qualify Formation Flight (Alvar)

Developing and Maturing Micro-Satellite Formation Flight Technology

Alvar Saenz-Otero, SPHERES Lead Scientist, MIT Space Systems Laboratory, MIT Room 37-381, 70 Vassar St, Cambridge, MA 02139, [email protected] Simon Nolet, Post-Doctoral Associate, MIT Space Systems Laboratory MIT Room 37-356, 70 Vassar St, Cambridge, MA 02139, [email protected] Raymond Sedwick, Assistant Professor, Department of Aerospace Engineering 3146 Martin Hall, University of Maryland, College Park, MD 20742 David W. Miller, Director, MIT Space Systems Laboratory, Professor of Aeronautics and Astronautics MIT Room 37-327, 70 Vassar St, Cambridge, MA 02139

Introduction

This chapter addresses some of the challenges of synthesizing the functionality of a larger satellites through the formation flight of multiple micro-satellites. While the motivation for satellite formation flight can be quite compelling, the challenges of developing and maturing the technology can be equally daunting. This chapter will focus on the efforts of the MIT Aeronautics and Astronautics Department’s Space Systems Laboratory (MIT-SSL) to answer the following questions:

 What current and future missions will require formation flight?

 What are the challenges in formation flight?

 How does one mature formation-flight technology?

 Can propellant-less techniques be employed to reduce fuel consumption during formation flight?

The following sections address the motivation, the challenges, a technique for propellant-less satellite formation flight, an approach to in-space technology maturation and verification, and sample results of maturing micro-satellite technologies including estimation, docking, and formation flight.

Motivation for satellite formation flight

Astronomy has entered a golden age. We are starting to answer the age-old questions of how did it all begin, how will it end, and is there life beyond Earth. Discovering the answers to these questions raises daunting engineering challenges. Space telescopes — our premier investigatory tools — are becoming ever larger, exceedingly precise, and more exotic. Whether it is the Mount Wilson telescope in California or the Hubble Space Telescope in Earth orbit,

1 dramatic vistas of the universe have been opened to us. These have been enhanced through periodic upgrades of the instruments located at the focus. The next generation of astronomical telescopes will operate 10 million miles from

Earth at the second Earth-Sun Lagrangian point (L2). To improve angular resolution (the ability to distinguish between two closely-spaced objects), the Space Interferometry Mission1 and Terrestrial Planet Finder – Interferometer (TPF-I)2 exploit multiple telescopes that are spread apart. In the case of the latter, these individual telescopes lie on separate spacecraft (S/C) that are flown in formation. Our challenge is to engineer a telescope effectively the size of a football field, that operates in an environment only slightly warmer than absolute zero, and orbits autonomously 10 million miles from Earth to a precision less than the diameter of a hydrogen atom.

What could be one of the most complex spacecraft ever built will be expected to work in an operating environment it has never run in. It is not even possible to have an end-to-end functional test of such a space telescope on Earth. Further, since launch costs are so high and opportunities so rare, it is only natural for program managers to include as much functionality as possible into each S/C, increasing the complexity to test such a system. These facts result in the aerospace industry converging into an operating mode with multiple drawbacks: high design costs, complexity of system integration and validation, risk of large deployment, low levels of maturation due to design customization, and limited design heritage and legacy. Most of the industry lacks the ability to incrementally develop and test integrated space systems.

Therefore, the MIT-SSL is developing an innovative approach to satellite design that addresses several of those drawbacks. By modularizing typical satellite functions (e.g., propulsion, power, attitude control) and achieving subsystem interconnectivity through genderless docking ports and wireless command and data handling, we seek to simplify assembly and test prior to launch, as well as enable self-assembly and reconfiguration once on orbit. The goal is to reduce the cost, risk, and time required for the deployment of new S/C by changing the fundamental methodology used in their development.

Coordinating the use of multiple satellites to facilitate assembly, servicing, upgrade, and operation of future space-based telescopes is an emerging field. Distributing mission functionality across multiple satellites has the promise of revolutionizing space exploration in general, while also presenting unique challenges.

2 Technology Development Challenges

Many development challenges exist: relative state sensing; autonomous operations; staged coarse and fine control (meters-to-nanometers); on-line path planning; fault detection, isolation and recovery (FDIR); etc. These challenges are receiving considerable attention throughout the formation flight community. To illustrate the MIT-

SSL’s approach to addressing the challenges associated with formation flight technology development, this section focuses on propellant consumption and docking.

Satellites use propellant to maneuver from one orbit to another, much like an automobile uses gasoline.

Formation flying satellites will also need to maneuver in order to keep formation and retarget. Unlike automobiles, these satellites do not have the advantage of maneuvering into a gas station as their propellant runs low. Instead, they must bring with them, at launch, all of the propellant they will need over their lifetime. This makes propellant a precious commodity. Electromagnetic Formation Flight (EMFF), using renewable solar energy, replaces the need for propellant in performing formation flight.

As is presented in the sections below, it has been shown in theory3,4 and practice5,6 that a combination of electromagnetic (EM) dipoles and reaction wheels can control all relative degrees of freedom among a cluster of vehicles without the use of propellant. Using current and future state-of-the-art high temperature superconducting materials/wire to generate the EM fields, low power and lightweight systems can be realized that are competitive with current high specific impulse propulsion, but are not life-limited by propellant and power consumption. Because of the low power requirements, and lack of consumables, much more aggressive maneuvers can be performed continuously over the lifetime of the mission. Any mission that can be satisfied by controlling only relative degrees of freedom is a potential application for EMFF. Potential future applications include: cluster formation flying missions, including those that may need formation reconfiguration; formation keeping missions which maintain precise relative states (fighting perturbations such as differential drag, gravitational variations, and solar pressure); close proximity operations such as

S/C inspection; and in-space assembly and docking of new large spacecraft (such as inter-planetary stacks).

Autonomous docking is another technical challenge. In the event that satellite servicing becomes commonplace, docking with partially failed and fuel-depleted satellites will occur. In fact, some of these satellites may be so degraded that they are tumbling when docking occurs. The dynamics of such docking scenarios is inherently three-dimensional in nature, necessitating maturation of this technology under microgravity conditions. Maturing docking technologies opens the possibility of inspection and robotic assembly. The deployed geometry of a large

3 satellite can be completely de-coupled from its stowed geometry, allowing systems to be packaged most efficiently for launch and then assembled to create apertures many times the size of the fairing. In Figure 1 we see three steps of an in- space assembly sequence concept. A large telescope, here modeled as a ~10m primary, may be launched as a compact stack (a). Autonomous docking maneuvers retrieve smaller mirrors (e.g., 2m) from the stack and assemble (b) the large primary (c). A formation flight satellite can be used to hold the secondary mirror in place (c).

Technology Maturation Challenges

Multi-vehicle guidance, navigation and control (GN&C), a technology that is at the heart of formation flight architectures, must be matured in a systems context. Specifically, all of the other sub-systems must be developed and integrated (sensors, propulsion, avionics, communications, power, etc.) before the GN&C sub-system can undergo hardware-in-the-loop testing. Furthermore, formation flight is an inherently three-dimensional operation. Ground testbeds suffer from the inertia, complexity, and constraints imposed by suspension systems (e.g., air carriages, robotic arms). As a result, GN&C, which is perhaps the highest risk technology because its failure can rapidly lead to the loss of the entire mission, trails other related technology when it comes to hardware-in-the-loop maturation under a representative operational environment.

The MIT SSL approach to this problem is best introduced with an equivalent every-day scenario: consider that we formation “fly” every day on the interstate with surprisingly few collisions. However, we do not toss the keys to our expensive car to our 16-year-olds. Instead, we have them practice in a less expensive car in a risk-tolerant environment

(e.g., parking lots) until handling nominal and off-nominal conditions becomes second nature. Only then do you allow use of the expensive car.

The Synchronized Position Hold Engage Reorient Experimental Satellites (SPHERES, http://ssl.mit.edu/spheres) program recreates this learning environment for satellite GN&C. Formation flight algorithms are the “16-year-olds” which must mature to handle nominal and off-nominal conditions. The SPHERES satellites are substantially less expensive than any traditional demonstration mission, and operate in the risk-tolerant environment of the International Space Station (i.e., the ISS is our “parking lot”), shown in Figure 2. By operating in this environment,

SPHERES can demonstrate that satellite formation flight is not only feasible but also robust. Further, by exploiting platforming concepts, where a common chassis with standardized interfaces allows modular components to be added, it is extensible in both hardware and software to accommodate a myriad of diverse research objectives. Uplinking

4 software, downlinking data and attaching payloads to the SPHERES expansion ports facilitates spiral algorithm development and hardware extensibility.

The following sections describe how the MIT-SSL is addressing these satellite formation flight challenges.

First, the development of propellant-less formation flight is described with focus on electromagnet design, dynamics and controls, and application to the TPF-I mission. Second, the philosophy behind and the resulting design of the

SPHERES approach to testing and maturing GN&C algorithms is discussed. Finally, results of SPHERES experiments to mature GN&C algorithms are shown.

Electromagnetic Formation Flight (EMFF)

Most proposed methods of actuating S/C in sparse aperture arrays use propellant as a reaction mass. For formation flying systems, propellant can become a critical consumable which is quickly exhausted while maintaining relative separation and orientation. Furthermore, the total required propellant mass increases exponentially with ∆V.

Additional problems posed by propellant include optical contamination, plume impingement, thermal emission, and vibration excitation. For those missions where control of relative degrees of freedom (DOFs) is important, an alternative is to use a system of electromagnets, in concert with reaction wheels, to replace the consumables. A system of electromagnets, powered by solar energy, does not rely on consumables such as propellant.7,8

A cluster of S/C in Earth orbit will experience geo-potential disturbance forces that will cause the S/C to drift apart without constant formation keeping. The ∆V required in LEO is about one centimeter per second per meter of separation per orbit (scaling linearly with orbital period and separation distance) causing the total ∆V over the mission lifetime to be substantial. If propellant is used to provide this ∆V, the large mass fractions will ultimately limit the advantage that formation flight can deliver as mission lifetimes are extended. EMFF can formation-keep indefinitely as well as dump any accumulated angular momentum into Earth’s magnetic field.9,10

EMFF is particularly beneficial during close proximity operations (within 10 meters), where thruster firing directions are severely constrained. This is particularly true when creating positive separation rates for collision avoidance. If thruster firing directions are not properly constrained, plume impingement can pose a contamination and damage threat to neighboring S/C. EMFF does not have this problem. In fact, EMFF thrives in close proximity where the coupling between magnetic fields on neighboring S/C can be stronger, increasing control authority.

5 How Does EMFF Work?

EMFF replaces thruster-based propulsion with EM forces for controlling the relative DOFs of a S/C cluster.11

By creating attraction and repulsion forces between S/C (Figure 3-a), as well as torques and shear (off axis) forces

(Figure 3-b), EMFF coils and reaction wheels can control all DOFs except the three translations of the center of mass of the entire cluster.

MIT-SSL’s EMFFs come in two basic flavors: HTS-EMFF uses high temperature superconducting (HTS) wire to greatly amplify the inter-S/C force, or extend it over much further distances. A typical HTS-EMFF system could produce an EM dipole with micro-propulsion level thrust (tens of milli-newtons) at a hundred meters of separation using commercially available superconducting wire. The MIT-SSL prototype creates this force by driving up to 100 amps of current through on each S/C, at a power consumption of about 10 Watts. As of 2007, the state-of-the art in commercial HTS wire will carry about 150 amps through a cross-section of 0.5 mm2. Micro-EMFF, still in the early stages of development, uses conventional conductors to provide close-proximity (up to one meter in ground operations) perturbation compensation, collision avoidance or slow maneuvering.

Force is the weakest interaction between the S/C. So, to illustrate the effectiveness of EMFF in controlling the geometry of an array, force is analyzed. The far-field force (F) between two identical EM coils is given by the product

of the permeability of free space (o), the square of their total amp-turns (IT) and the fourth power of the ratio of coil radius to their separation (R/d), as shown in Equation 1. The key limitation of EMFF is the rapid reduction of the dipole force with distance. However, the key strength is the fact that the dipole nature of the field permits both force and torque to be created, allowing control of not only separation but also shear and rotation between the S/C.

4  2 3 2 R  3 7 I c 2 1 F ~ o IT   ~ 1 0   M R  4 (1) 2 d  2   d Substituting for the properties of the super-conducting wire, this force relationship can be rewritten revealing two important terms. The first term (I /) is a technology term associated with the HTS. The maximum, or critical,  c current density (Ic) supported by the HTS limits the achievable force. However, this critical amperage density can be increased by lowering the temperature of the superconductor. The volumetric mass density () accounts for the density of the coil and its insulation. This technology term can be altered by selecting different HTS products, changing the operating temperature, and modifying the insulation. The second term is a resource allocation term that shows the

6 impact of adding coil mass (M) with a major radius (R). The equation shows that improvements in HTS amperage density (increase) and insulation mass (decrease) can increase performance substantially, as they have a quadratic effect on the resulting force. Figure 4 plots force versus separation for different products of commercially available coil mass and radius (MR).

Since power consumption is not an issue, HTS-EMFF allows continuous precision positioning of S/C cluster elements. Using reaction wheels, all relative DOFs within a cluster of S/C can be independently controlled. The challenges are thermal control to the required temperature of commercial HTS (~77 degrees Kelvin) and the coupled, nonlinear dynamics, both of which have been experimentally studied in depthError: Reference source not found,12.

Micro-EMFF produces an EM dipole using conventional wire, providing collision avoidance and cluster maintenance capabilities at distances up to 30 meters. The HTS-EMFF system is dominated by the power dissipation in the current controller, while the wire dissipates minimal power. Micro-EMFF does not require thermal control, however substantial power is dissipated in the wire, in addition to that used by the current controller. Therefore pulsed operation is used. The challenge is to synchronize the pulses to ensure adequate electromagnetic coupling between the

Dynamics and Controls

The benefits of EMFF come at the cost of coupled and nonlinear dynamics of the formation, making the control problem a challenging one. The dynamics for a general N-satellite EM formation have been derived for both deep space missions and Low Earth Orbit (LEO) formations.13 Nonlinear control laws using adaptive techniques have been derived for general formations in LEO.14 Angular momentum management in LEO is a problem for EMFF due to interaction of the magnetic dipoles with the Earth’s magnetic field. A solution of this problem for general EM formations takes the form of a dipole polarity switching control lawError: Reference source not found,15. This switching law preserves the forces between S/C but periodically reverses the torques induced by Earth’s magnetic field. This periodic switching allows accumulated angular momentum stored in the reaction wheels to be drained into the Earth’s magnetic field thereby avoiding saturation of the wheels.

For EMFF, the formation reconfiguration problem is a nonlinear, constrained optimal time control problem, as fuel cost for EMFF is zero. Two different methods of trajectory generation, namely feedback motion planning using the

Artificial Potential Function Method (APFM) and optimal trajectory generation using the Legendre Pseudospectral

7 method, have been derived for general EM formationsError: Reference source not found. The results of these methods have been compared for up to ten S/C EM formations. This comparison shows that the APFM is a promising technique for solving the real-time motion-planning problem for nonlinear and constrained systems, such as EMFF, with low computational cost. Specifically, a fully actuated N-satellite EM formation can be stabilized and controlled under fairly general assumptions, therefore showing the viability of this novel approach for satellite formation flight from a dynamics and controls perspective Error: Reference source not found,16.

A two-vehicle HTS-EMFF controls testbed is operational on the MIT-SSL’s flat floor (Figure 5-a). Each vehicle has two HTS coils, cooled by liquid nitrogen, that allow the dipole to be steered in the horizontal plane. The coils, designed for close proximity operations in the laboratory, carry up to 95 amps and produce up to 0.2 tesla. The vehicles, which float on air carriages, also contain one reaction wheel that provides torque about the vertical axis. This testbed has been used to test position regulation in the presence of perturbations, as well as trajectory following. In addition, processors, power systems (alkaline and rechargeable batteries), radio frequency (RF) communication channels, reaction wheel controllers, and Global Positioning System (GPS) receivers have been exposed to the magnetic fields. No interference or degradation in performance, due to the magnetic field, has been detected. While liquid nitrogen boil-off is used to cool the HTS in the laboratory, it is not a good choice for space. Therefore, there also exists a thermal-vacuum chamber that is used to test thermal control techniques for the coils.

A Micro-EMFF testbed is also operational (Figure 5-b). A single circular coil is attached to the end of a swing-arm that is mounted to an air bearing. A second coil is mounted on a pivot. When current is simultaneously pulsed through both coils (at up to 30 amps), this swing-arm rotates due to the generated force. This testbed has been used to demonstrate axial forcing, shearing, collision avoidance and dead-band control in open-loop and closed-loop configurations.

TPF-I Mission Application

The TPF-IError: Reference source not found mission, a 5-vehicle rotating interferometer at the Earth-Sun L2 point under development by NASA, was studied to understand the reductions in S/C mass that can be realized using

EMFF. The goal is to analyze the design of TPF-I with sufficient subsystem detail to fairly compare EMFF with micro- propulsion systems. The micro-propulsion techniques analyzed included pulsed plasma thrusters (PPTs), field emission electric propulsion (FEEPs), and two models of HTS-EMFF (named EM1x and EM3x). The 5-vehicle TPF-I system

8 consists of two outer collectors, two inner collectors and one central combiner. The mass breakdown for the outer collector, inner collector and combiner S/C for the various sub-systems are shown in Figure 6. The additional mass for the EMFF subsystem due to reaction wheels and power storage is captured either within the Attitude Control System

(ACS) or the power subsystem. Note that for the propellant-based options, the outer collector is the most massive since it has the highest centripetal load, while the combiner does not have any propulsion mass since it only needs to rotate in place. The propellant-based options also have a relatively high amount of mass (excluding the dry mass) allocated for propellant. This shows that while the ACS mass increases when EMFF is used, the dramatic reduction in propellant mass makes EMFF quite favorable.

Developing formation flight technology (such as EMFF), demonstrating it in the laboratory, and analyzing its benefits in the context of future missions is only part of the challenge. Maturing the technology so that it can be used in operational systems can be equally difficult. The next section describes the MIT-SSL’s approach to addressing technology maturation.

Maturing Formation Flight GN&C

Space technology maturation is a challenging process. Substantial amounts of money, time, and human resources go into the development of new S/C. At every point in the design life of a new S/C, there are substantial risks involved, especially as the complexity of new design increases. Over a decade ago NASA developed the Technology

Readiness levels (TRL)17 to determine where in the design process a specific technology stands. The levels attempt to divide the design process into incremental steps. As a technology gets to a higher level, the requirements for its demonstration environment increases in complexity, requiring the use of special testing facilities and eventually testing in an environment representative of space.

Microgravity experimental facilities vary in their fidelity, cost, and operational limitations. While not necessarily exhaustive, the list presented in Table 1 shows a wide range of facilities that can provide an environment to reproduce or simulate microgravity conditions in support of research. The table lists fifteen different environments under which microgravity research can be emulated/conducted. The first column shows facilities which can be housed within the individual researcher’s institution, but which don’t necessarily simulate full 6-DOF microgravity. The second column lists facilities which have full 6-DOF capabilities, but which are usually managed by a third party. The

9 third column lists the existing facilities which provide full microgravity conditions, but which present the largest development challenges.

Table 1. Sample of available facilities for -g research.

In-house 3rd Party / Full -g Space Robot Helicopters Reduced Gravity Aircraft Free Flyer 6-DOF Robot Arms Neutral Buoyancy Tank ISS Helium Balloons Drop Towers Shuttle Payload Robot Cars Shuttle Middeck Flat Floor Sounding Rocket Air table Simulation

Traditional space system development, especially for single-satellite systems, goes only through simulation and ground-based tests. The complexity and high costs of creating flight prototypes prevents in-space flight qualification/demonstration of most high value science missions; many demonstration missions turn into high cost multi-year missions themselves ($100M+18 such as DART19 and Orbital Express20). Ground-based tests and simulations have important shortcomings for technology maturation: limited DOFs, non-representative dynamics, gravity and suspension systems change the dynamics, limited operation time, etc.

The MIT-SSL conceived an innovative way to use the existing microgravity environment aboard ISS: use small satellites in conjunction with the station to create a laboratory environment for space technology maturation. The use of micro-satellites, and careful use of the resources which have already been built by others to accommodate research on the ISS (such as NASA, the DoD, and the ISS partners), overcomes the majority of the limitations of full space demonstrations: the facility is low-cost (<$2M), low-risk, flexible, enables human intervention/supervision, and reduces the profile of the mission to that of standard laboratory research. At the same time, the facility allows operations in a 6-DOF microgravity environment with highly representative dynamics. The use of multiple small satellites, which emulate standard satellite-bus craft, enables testing of multiple research areas including formation flight and docking systems.

The Philosophy Behind SPHERES

The SPHERES formation flight and docking laboratory is the culmination of a decade of dynamics and controls research laboratories conducted by the MIT-SSL and flown on the Shuttle, MIR (Russian space station) and

ISS. The MODE (STS-48 & 62), DSL (STS-62, MIR), and MACE (STS-67, ISS Expedition 1 & 2) experiments took

10 place from 1991 through 2001. Based on the experiences of short term (STS) and long term (MIR and ISS) missions, the use of modularity, and the risk-tolerant environment with human supervision, the MIT SSL developed a design philosophy21 to aid the development of SPHERES. The philosophy formalized the lessons learned during those missions and allowed a structured approach to implement a true laboratory environment for GN&C aboard the ISS.

The first step in maturing a technology through hardware-in-the-loop testing is to clearly identify the objectives that must be achieved. To develop a GN&C technology maturation laboratory, the MIT-SSL identified the following hardware-in-the-loop testing objectives:

 Demonstration and Validation in the correct environment such that the results clearly show that the technology

achieves the required performance.

 Repeatability and Reliability of each test by ensuring that they provide the same results more than once under

similar operating conditions with representative system noise.

 Determination of Simulation Accuracy by providing results to calibrate and validate simulations.

 Identification of Performance Limitations by pushing the limits of new technologies or algorithms until failure;

i.e., learning the limitations of an algorithm in a relevant environment before deployment in a full mission.

 Identification of Operational Drivers, or design parameters, which enable engineers to create design trade

studies and system designs based on the most relevant parameters for GN&C.

 Identification of New Physical Phenomena should be possible by allowing observation of the physical system

behavior; the phenomena must be captured well enough to allow creation of models for future designs.

Based on these objectives and the unique resources made available by the ISS, the MIT-SSL conceived the following design principles which designers of microgravity experiments for technology maturation can use in the development of space technologies maturation laboratories operated aboard the ISS. By generalizing the GN&C objectives into principles, these encompass a wider range of technology maturation experiments, beyond the dynamics and controls scope of the MIT-SSL.

Principle of Iterative Research

A successful laboratory environment supports the established scientific process that a hypothesis be tested and modified as experiments are performed. A designer should consider the following steps:

 Conception: initial hypothesis definition and scientific objectives of the laboratory

11  Facility Development: engineering design and manufacturing of the facility

 Technology Maturation Tests

 Individual test development

 Experiment run and data capture

 Data analysis versus expected results

 Hypothesis reformulation as needed

The principle of iterative research separates these steps into science time (hypothesis formulation & modification, test runs, and data analysis) and overhead time (facility design, implementation of tests to operate in the specific facility, and data collection/transfer). While overhead time should clearly be minimized, it is not as simple for science time. Science time must be flexible: sufficient to allow ample time for data analysis and hypothesis

(re)formulation, while not so long that a scientist looses focus on the task at hand.

Principle of Enabling a Field of Study

In order to provide experimentation in a field-of-study, a laboratory must allow for experiments within the different research areas of the field. Given the intent to use the ISS (a highly valuable resource) and the large overhead in the development of facilities to operate there, scientists must use the facility as efficiently as possible. This principle calls for each program aboard the ISS to develop laboratories which aid a field of study. To do so, a facility must:

 Support multiple investigators to work on individual topics to cover the whole field of study.

 Facilitate bringing together the knowledge from the specific areas to mature understanding of the field as a whole.

Principle of Optimized Utilization

The ISS offers a wide range of unique resources that make it ideal for the maturation of space technologies.

Rather than thinking about using the least resources possible, this principle guides the researchers to maximize the value of their use with respect to the research goals. The specific resources of the ISS to consider are:

 Crew The crew can provide valuable feedback, even interpretation of results, based on observations. Further, the

crew can supervise experiments to minimize risks as scientists push the limits of algorithms.

 Consumables The ISS provided electrical power, pressurized gases, and liquids can greatly reduce the cost of a

facility and increase the researcher’s focus on the technology instead of hardware development and maintenance.

12  Data telemetry The ISS communications and data management system allows scientists to obtain their data

within hours of the experiments and enables upload of new software. In the future it is expected to enable real-time

video and other teleconferencing options as part of daily research operations.

 Long-term experimentation Long-term microgravity experimentation enables flexible operating schedules,

iterative research, controlled test conditions, and ample science time.

 Risk-Tolerant Environment / Benign Atmosphere The availability of a risk-tolerant environment enables tests

which push the limits of the technology while, at the same time, reducing complexity and costs.

Principle of Focused Modularity

Since experiments almost always contain basic elements that can support other similar experiments (e.g., power, communications, data storage, processing), the design phase of a facility should identify these common elements. These generic parts should be made available for future experiments as long as it does not compromise the objectives of the original experiment. Just like any terrestrial laboratory, there should exist a suite of generic capabilities that can be used (and possibly reconfigured) by future researchers. The call for focused modularity is to prevent a "do-everything" system. The modular part on the design should focus on those common features identified as generic after the initial facility design.

Principle of Remote Operation & Usability

Remote laboratories are based in remote locations because they offer a scarce environment that researchers cannot emulate in their home institutions. The ISS offers a unique microgravity environment. The design of remotely operated laboratories for the ISS must account for the following facts about the operation:

 Operators: are usually not experts in the specific field and are a limited resource. Their time spent supervising a

test must provide useful feedback to the scientist.

 Research Scientists: have limited or no experience in microgravity operations and the ISS environment, therefore

what may seem easy to them in the ground might not be for the operators. While the scientists can modify

experiments in their facility, it is not possible to do so remotely. Therefore, scientists must constantly be made

aware of the differences between their operations and the remote facility operations.

The goal of a remote facility is to allow for a virtual presence of the research scientist in the operational environment by combining the abilities of the operator (astronaut) and scientist (on the ground).

13 Principle of Incremental Technology Maturation

The goal of incremental technology maturation is to make the complexity, risk, and cost increase smoothly as one climbs through the TRL levels, while being cognizant of the changes in the environment required. The ISS provides an environment that can closely satisfy the requirements for a space environment (space radiation, illumination, vacuum, and thermal conditions can be replicated in terrestrial facilities); yet the presence of humans in the ISS can greatly reduce the risks involved, and the existence of the ISS itself can reduce the costs. Further, successful tests in the ISS may lead to less complexity when moving to even higher TRL levels by providing researchers with a better understanding of the system and its true risks.

Principle of Requirements Balance

By the time of the Critical Design Review, hard requirements are usually set to determine the goals that must be met; they are mostly quantitative. Soft requirements are features desired by the researchers but which do not necessarily have a specific value or which are not essential for the success of the mission (and usually creep up after reviews). A successful design creates a realistic set of requirements, while taking into account the other principles presented herein. This principle does not call for all the requirements to be perfectly balanced or to necessarily eliminate the soft requirements. Rather, this principle calls for the scientist to pro-actively pursue a realistic justification for each requirement and to ensure that a substantial part of the effort spent on the development of the facility goes towards clearly defined needs.

SPHERES Design

The SPHERES22 laboratory for distributed satellite systems consists of three nano-satellites, a custom metrology system, communications hardware, consumables (propellant tanks and batteries), and an astronaut interface aboard the ISS. Figure 7 shows the SPHERES satellites being operated aboard the ISS and identifies the different elements of the facility. The ground-based setup consists of another set of micro-satellites, a more streamlined interface, and a “SPHERES Guest Scientist Program” to allow multiple researchers to use the facility (in response to the principle of a field of study).

The SPHERES satellites were designed to provide the best traceability to upcoming formation flight missions by implementing all the features of a standard thruster-based satellite bus. The satellites have fully functional

14 propulsion, guidance, communications, and power subsystems. These enable the satellites to maneuver in six DOFs, to communicate with each other and with the laptop control station, and to identify their position with respect to each other and to the experiment reference frame (two millimeters in position and one degree attitude resolution). The laptop control station is used to collect and store data as well as to upload control algorithms to the satellites. Figure 8 shows a

CAD drawing of the open satellites as well as a picture of an assembled SPHERES satellite and identifies its main features. Physical properties of the satellites are listed in Table 2. The following list presents basic descriptions of the subsystems, further details are available in the appendices of [Error: Reference source not found] and the SPHERES

Critical Design Review 23.

Table 2. SPHERES satellite properties.

Diameter 0.22 m Mass (w/tank & batteries) 4.3 kg Max linear acceleration 0.051 m/s2 Max angular acceleration 0.944 rad/s2 Power consumption 13 W Battery lifetime 2 h Communications data rate 20 kbps

 Propulsion – The satellites are propelled by pressurized carbon dioxide propellant. The selection of the propellant

was part of a large trade study between multiple inert cold gases and on-board air compressors. CO 2 provided the

best performance for our needs, as it exists in liquid form at room temperature and is reasonably stable without the

need for cryogenics or other power-consuming support systems. The trade-off was the requirement to operate only

when the Carbon-Dioxide Removal Assembly (CDRA) of the ISS is active. A manually adjusted regulator sets the

pressure to 25 psig. Twelve thruster assemblies, distributed among the three orthogonal axes (four per axis, in

opposite pairs), allow 6-DOF maneuvers enabling both attitude and station keeping control. Each thruster provides

approximately 0.10N of force. The system is primarily used with pulse-width modulation (at up to 20Hz) to

achieve variable forces and torques. The propulsion system may be easily replenished; 96 gas tanks were delivered

to the ISS for SPHERES operations.

 Guidance – The SPHERES metrology system closely resembles a GPS system, but it operates inside the ISS

(where GSP does not) at resolutions of up to 2mm. The goal of the metrology system was to create a system

which can use software signal processing to provide state information equivalent of a wide range of space systems.

For example, scientists can be provided only with attitude information to simulate a star tracker. The metrology

15 system consists of a local 6-DOF inertial measurement unit (three accelerometers and three gyroscopes operating

at 50Hz) and a global pseudo-GPS system which uses ultrasonic time-of-flight measurements from transmitters

(equivalent of GPS satellites) placed at known locations in the ISS (or SSL laboratory). These time-of-flight

measurements are converted to ranges and used to derive position and attitude at a frequency up to 5Hz.

 Communications – Each SPHERES unit uses two separate frequency communications channels to simulate the

expected operations of future formation flying missions: one channel is used for satellite-to-satellite (STS)

communications; the other channel enables satellite-to-ground (STG) communications. The system uses COTS

916MHz and 868MHz close range ASK type RF transceivers. The effective data rate of each channel, after

removing the bandwidth needed for packet management and error checking is 20kbps. The system does not use the

higher bandwidth 802.11 standard because the ISS has its own wireless LAN; use of the frequency would have

held SPHERES to safety reviews of any communications between the satellites and the LAN. The goal of effective

iterative research took precedence in this design choice.

 Power – Each SPHERES satellite is powered by two packs of 8-AA alkaline batteries each. The packs provide up

to two hours of operations. The power system regulates the battery voltage to six voltages used by the internal

electronics. The choice of AA batteries was driven by requirements balance. While the better implementation

would have been to select rechargeable batteries (and therefore use ISS resources better), the substantial amount of

time, effort, and money required to approve rechargeable batteries for the ISS would have delayed the project

substantially and consume too many funds. Therefore, while rechargeable batteries are under development in

parallel with operations, the use of AA batteries allowed faster and cheaper deployment to the ISS.

 Avionics and Software – A Texas Instruments C6701 Digital Signal Processor (DSP) provides the computational

power. These DSP’s have been used in other space missions and ensure real-time operations required for GN&C.

Further, the DSP processors include all support functions of a standard processor, allowing it to control the whole

unit. The primary driver in the selection of a DSP was the high ratio of processing to energy consumption; the

selected DSP carrier board provides between 166-1000 MFLOPS (depending on optimization) and requires under

7W. The processor is supported by 16MB of RAM and 256KB of FLASH memory (the avionics’ most stringiest

limitation). The software is based on the TI DSP/BIOS Kernel, with a custom configuration of the real-time

elements to allow easy access to the metrology (inertial and global) data, control interrupt (up to 1kHz), propulsion

16 driver, and communications data. The software enables access to high priority periodic processes (e.g. data

capture, control) and low priority aperiodic background processes (e.g., high level estimation routines).

 ISS Crew Interface - The graphical user interface (GUI) for use aboard the ISS was developed with the idea that

the ISS crew members may not have any contact with the SPHERES researchers while they run tests. Therefore,

the GUI not only guides the crew on how to run tests, but also to capture the observations of the crew - the crew

becomes an extension of the scientist by providing interpretation of the tests runs in presence of the hardware. By

doing so, the GUI became an Electronic Lab Notebook24. Together with multiple angles of video capture and the

possibility to use live audio and video during test sessions, SPHERES creates a virtual presence of the scientist

aboard the ISS.

Guest Scientist Program

The Guest Scientist Program was an integral part of the development of SPHERES. The high-priority objective to enable multiple scientists to use the laboratory heavily guided the development of the software and a set of simulations. The main SPHERES software, called SPHERES Core, creates a generic GN&C software environment with high-level interfaces to the hardware (all other subsystems and the micro-processor), as well as generic algorithms developed at MIT for estimation, control, and utility functions. This allows scientists to concentrate on their science, rather than the specific implementation on SPHERES or science not in their field (e.g., an estimation researcher does not need to develop their own control laws, they can use existing ones developed and tested by the SPHERES team).

This required careful balance between the development of generic tools (which adds overhead to the software and DSP requirements) and the level at which scientists prefer to interface to the software. To achieve this balance the

SPHERES team held a Science Critical Design Review25.

The Guest Scientists Program also included the development of a simulation for the use of the hardware.

These enable scientists to create initial implementations of their algorithms. However, due to resource limitations, the simulations are not end-to-end. Therefore the program also joins an MIT graduate research assistant with remote scientists so that tests can be conducted with ground-based satellites.

Verification and Validation of the GN&C Flight Software

Like any other flight software, the SPHERES software goes under an extensive verification and validation

(V&V) process in order to increase the level of confidence prior to flight. This process includes both simulations and

17 hardware-in-the-loop tests at various levels of software integration. By being smaller and cheaper, micro-satellites can have the advantage of being more accessible for hardware-in-the-loop tests for the integrated GN&C system, which further increase the level of confidence. This section describes the different components of the V&V process used when preparing experiments on the SPHERES micro-satellites.

Verification and Validation Process Overview

The process is illustrated in Figure 9. It involves software-in-the-loop simulations (right column) and hardware-in-the-loop experiments (left column). The simulations are performed in MATLAB. Hardware-in-the-loop experiments are typically performed at the MIT SSL 2-D testing facilities (air table and/or flat floor). A series of five gates (diamond shaped boxes) are used to make a decision as to move on to the next testing phase (the ISS), or to iterate in case the systems engineer’s expectations (requirements) are not met.

The process is initiated with the guest scientist integrating a GN&C software in the MATLAB simulation.

While integrating the software, the guest scientist ensures that all parameters proper to the simulation environment

(mass properties, controller gains, dynamics coefficients, etc.) are contained in one file that can be later easily replaced by another one with flight parameters. Once the software simulation ensures that all expectations are met (Gate #1), the guest scientist integrates the flight software in the C programming language, using as many validated modules as possible to benefit from code heritage. With all parameters set for 2-D (ground) testing, experiments are performed on the MIT SSL 2-D facilities. The ones perturbed by extensive ground effects (gravity, bumps, drag) have to be restarted.

Once the expectations are met at Gate #2, a model of for the external perturbations is extracted whenever possible.

Implementing that model in the simulation environment, along with the parameters used for 2-D testing, allows a direct comparison between the simulations with 2-D parameters and the experiments at the MIT ground facilities. An agreement between the two (Gate #3) validates the simulation with the 2-D parameters, which increases the confidence that the first simulation with the 3-D (microgravity) parameters is representative of the reality.

The last two gates are a mean to further increase the level of confidence in the hardware prior to the flight experiments. After inserting back the parameters for 3-D experimentation in the same flight software used at Gate #2 and #3, more ground testing is performed at the MIT SSL ground facilities. Although the ground dynamics and perturbations are likely to prevent the tests from completing successfully, the goal of Gate #4 is to ensure that the onboard computer can support the flight software running in real time with all sensors and thrusters being activated,

18 and with the 3-D parameters being used. After successfully going through Gate #4, the software is uploaded to the ISS.

At the beginning of every ISS test session, a checkout test (Gate #5) is performed to ensure the hardware operates nominally and all subsystems are go.

The verification process presented here has been developed through multiple ISS test sessions. Like in any flight program, it played a crucial role in increasing the confidence of the onboard software to a level acceptable for flight. By ISS Test Sessions #5 and #6, the process described above allowed for 100% success rate of the 15 experiments, from a total of seven guests scientists who participated in these two test sessions.

Software Simulation

The initial development and verification of a software module is performed in a MATLAB-based simulation.

Created for the SPHERES Guest Scientist Program, the simulation facilitates the development of GN&C software from a distant site. It also benefits from a large array of visualization tools made available by MATLAB, which facilitates the verification of the algorithms embedded in the software.

To create a realistic simulation of the dynamic response of the system, the simulation engine replicates the interrupts that occur in the hardware, executes the different primary interface functions in the same order as they occur on the satellites, and propagates the states of the satellites between the interrupts. The hardware models used in the simulation are based on data acquired from calibration of the SPHERES hardware in the ground and data measured during microgravity operations aboard the ISS. Sensor noise, biases, and blackout behavior has been captured; the actuator models include thruster strength and pressure drop as multiple thrusters open. This simulation does not model communications losses.

Hardware-in-the-Loop Ground Experiments

Three testing facilities were used to mature the software modules which populate the GN&C architecture. The

MIT-SSL 2-D facilities have enabled repeatable testing conditions. NASA's reduced-gravity aircraft (RGA) allowed short duration testing in a microgravity environment. The first test sessions in the ISS, even with only a reduced set of hardware available, were utilized to incrementally test the different GN&C modules in the final operational environment.

Hardware-in-the-loop experiments typically lead to the discovery of problems at an early stage of the software development process, when it is relatively easy and inexpensive to make changes to the flight software. Remedies can

19 be designed early on, thereby improving the reliability of follow-on experiments. Important lessons were learned by the

SPHERES team after discovering the nature of each problem encountered through hardware-in-the-loop experiments:

 When verifying a GN&C software module, it is important to experimentally check every assumption made

associated with the algorithm. Numerical errors caused by the software implementation of the algorithm might

invalidate some assumptions.

 Regular jumps in the state estimates at the measurement update phase of the navigation filter are likely to be

caused by unmodeled biases in sensor measurements. Knowing that the hardware is identical between all

SPHERES micro-satellites, these unmodeled biases are likely to be common to all of them. If not, the problem can

be quickly traced to a specific sensor onboard the satellite.

 Although hardware-in-the-loop experiments at a component level are invaluable, it is crucial to clearly understand

the behavior of the navigation sensors after they are integrated into the system. The integrated sensors might show

a behavior that was not observed by the manufacturer in standalone tests, and therefore not documented in the

specifications.

Software Integration

Because verification and validation can be performed at various levels of software integration, it is important to understand the role and the effect the integration process can have on the overall functionality of the system. The integration process must ensure that the correct information is properly communicated between the software modules, as well as the peripheral devices such as sensors and thrusters, and that the processor and memory are not overloaded by the various software processes. Because of the complexity associated with the integrated software, software integration can sometime lead to undesired behaviors. Therefore, it is crucial not to overlook it, and to perform verifications whenever possible.

A modular GN&C architecture greatly facilitates the integration process. GN&C modules with different capabilities can be quickly integrated to allow an investigator to test different combinations. The selection of the right combination of modules is largely dependent on the available hardware and on the experiment to be performed. For example, the use of a certain module can be restricted when the hardware providing essential inputs is unavailable, or when the required resources exceed the hardware capability (e.g., computational processing, sensing and communication bandwidth).

20 Real-time hardware-in-the-loop experimentation was key in identifying problems related to system integration on the SPHERES facility. It allowed the team to:

 identify a proper sequence of maneuvers and GN&C modes leading to close-proximity operations like docking

with cooperative and tumbling targets;

 establish requirements regarding the transition between the maneuvers;

 confirm the capability of each software module to fulfill its corresponding requirements, when integrated into a

GN&C architecture.

Technology Maturation in Action using Small Satellites

The use of small satellites is not limited to simple missions or Earth observations. By following the MIT-SSL space technology maturation design philosophy, small satellites provide a unique ability to help develop and mature technologies for future missions that may use small or large satellites. Since May 2006, the SPHERES facility has demonstrated these concepts by allowing scientists at MIT and elsewhere to mature a large array of algorithmic technologies for future formation flight and docking missions. Table 3 presents a summary of the research iterations which have taken place for three main areas of study: metrology, docking, and formation flight. These iterations started to occur as early as the software development phase, between 2003 and 2006, as well as during the first five SPHERES test sessions (TS), between May 2006 and November 2006. The following subsections provide more details on the research iterations and technology maturation for each of these areas.

Table 3. Research iterations from Test Sessions 1 to 5.

Area TS Iterations Description Metrology 1,2,3,4,5 4 The metrology system began with a limited set of beacons, which resulted in the development of innovative filters. The lessons from that development and the capture of data in TS4 resulted in no need for further development starting with TS5. Docking 2,3,4,5 4 The basic maneuvers for docking began development during TS2. As each session demonstrated individual parts of a full docking mission, these parts were put together to achieve the final demonstrations of TS5. More iterations follow in future sessions. Formation 3,4,5 2 Starting with attitude-only formation flight, the latest iteration of Flight demonstrations aboard the ISS shows the basic maneuvers that are essential for missions such as TPF-I. Further iterations will occur.

21 Metrology and State Estimation

The core metrology estimation algorithm has seen important iterative improvements to its design. The low- risk, reconfigurable, and long-duration nature of SPHERES enabled the research team to address filtering of noise in raw data and tuning of the extended Kalman filter (EKF) during the core software development process and between

ISS test sessions. As the estimator base code became stable, the team continued incremental development with the introduction of fault detection and isolation (FDI) algorithms to improve estimator convergence.

The most common navigation problem encountered during the hardware-in-the-loop experiments was instability in the state estimates (position, velocity, attitude and rotational rate) output by the EKF 26. The different forms of instabilities included jumps of the state estimates (Figure 10-a) and temporary divergence (Figure 10-b). The jumps, typically occurring in the measurement update phase of the EKF, were caused by multi-path effects (reflections of the ultrasonic signals on the walls and surrounding objects) and unmodeled biases in the ultrasonic sensor measurements. The temporary divergence was caused by the covariance matrix of the navigation filter being driven non-positive definite during the state propagation phase of the EKF. In response to these instabilities, the controllers produced erratic behavior. This proved to be problematic during close-proximity. Moreover, problems with the navigation system occurred frequently in recent unmanned missions involving close-proximity operations (listed in

Table 4 below). Consequently, the resolution of the instability problems with the state estimators quickly became a high priority. The solution, refined through multiple static experiments where the satellites were constrained to remain stationary, included:

 a pre-filter to remove outliers in the ultrasonic measurements before they are passed to the navigation filter;

 an online fault detection algorithm preventing multi-path from affecting the state estimates;

 a reduced covariance matrix improving the robustness of the navigation filter.

Docking

Close-proximity operations, such as docking, require multiple advancements of current GN&C technologies.

Because of the communication delays with the ground controllers and the amount of information to be monitored, the onboard computers have to be granted significantly more autonomy than in common satellite missions. In the past decade, multiple missions were launched to test autonomous close-proximity operations, as illustrated in Table 4. They all involved either autonomous docking or inspection. It is important to emphasize that anomalies occurred for each of

22 these missions, with the possible exception being the XSS-11 mission, although no public report has been issued on its results. At least one of these missions failed, and all others were faced with anomalies that required ground intervention. Therefore, there is a need to mature technologies necessary to perform autonomous docking.

Table 4. Review of automatic and autonomous missions involving close-proximity operations.

Mission Year Status Type Anomalies Cosmos-186, 1967 Success Docking - Misaligned capture Cosmos-18827 28 Progress29 1978- Success Docking - On Progress 18, ground station problems prevented ongoing (multiple) the command uplink for automatic docking ETS-VII30 31 1998 Success Docking - Thruster anomaly SNAP-132 2000 Partial Inspection - Atmospheric drag larger than expected - Ran out of fuel before completing the rendezvous XSS-1033 2003 Success Inspection - Navigation problem (confused star tracker) - Communication interruptions XSS-1134 35 2005 Unknown Inspection - Unknown DART Error: 2005 Failed Inspection - Navigation problem (biased velocity measurements) Reference source - Collision not found 36 37 Orbital Express 2007 Success Docking - Navigation problem (GPS system initialization Error: Reference problem) source not found 38 - Inverted reaction wheel polarity - Navigation software error

In response to this need, the MIT-SSL initiated research into autonomous close proximity operations through the SPHERES project. The first step in this research was to analyze the state-of-the-art in GN&C architectures used in past unmanned missions involving close-proximity operations, and to identify the needed key improvements. A GN&C architecture is defined here as an abstract description of the entities of a GN&C system and the relationships between those entities. This definition is adapted from a broader system architecture definition proposed by the MIT

Engineering Systems Division Committee.39 GN&C architectures found in the literature have common characteristics supporting the definition above:40,41,42,43,44,45

 a decomposition in simple entities;

 a hierarchy between the entities;

 some interactions between the entities.

In his book entitled Automated Rendezvous and Docking of Spacecraft46, Fehse presents a typical GN&C architecture traditionally used for automated rendezvous and docking. The schematic he provides is simplified and shows only the levels of authority, not the actual functional relations within the system. This GN&C architecture

23 involves humans performing high level monitoring and making mission-critical decisions, such as the triggering of a collision avoidance maneuver (CAM) when anomalies occur. However, for complex operations, like docking to a tumbling target (especially if it were to be in Lunar or Mars orbit), the short reaction time required in the event of anomalies needs an increased level of autonomy, which is not supported by the GN&C architecture portrayed by Fehse.

Figure 11 shows a hierarchical view of the baseline GN&C architecture used on the SPHERES micro-satellites for autonomous docking. It is largely based on Fehse’s architecture, but has been extended to enable complex autonomous operations. The components common to Fehse's architecture are grayed. The box labeled autonomous onboard docking control systems contains the software modules that run on a satellite. Each software module is an implementation of an algorithm accomplishing a certain task (e.g., state estimation, trajectory control, thruster actuation). External interfaces include the plant (satellite), the operators and the other satellite. The software modules are grouped into FDIR, solver, mission & vehicle management (MVM) and GN&C. The extended architecture accounts for the idea that communications with human operators may be intermittent or non-existent (illustrated by dashed lines), potentially preventing ground operators from taking direct control of the satellite. This forces the integration of FDIR capability at all levels47,48. In this extended model, the onboard computer is granted the authority, through the main FDIR module, to trigger a CAM if a problem is detected. Furthermore, a solver module allows for online planning and scheduling. The solver is monitored by the FDIR module to prevent infeasible solutions or other failures. Since knowledge of the target's states is required, remote sensing and/or communications between the satellites are needed, as shown by the link with the target satellite.

So far, the team has focused on the implementation of low-level GN&C modules49,Error: Reference source not found on the

SPHERES hardware. The navigation EKFs process the ultrasonic time-of-flight signals to compute the position, velocity, attitude and rotational rate of the satellites. Controllers (based on PID, phase-plane, and glideslope algorithms) compute desired forces and torques from the tracking error at 1 Hz when following a trajectory. Thruster management modules transform the desired forces and torques into thruster ON/OFF times using a pulse-width modulation scheme.

Each module has the potential to be used in either docking or formation flight experiments, and can be called at various frequencies.

The higher-level GN&C modules are still at an early stage of development. The MVM module consists of the partition of a sequence (e.g., a docking sequence) into multiple short closed-loop maneuvers. 50 Each maneuver is validated step-wise, and then carefully put together until a complex demonstration can be achieved. The fault detection

24 system currently implemented rejects ultrasonic time-of-flight measurements incoherent with the state estimates, and causing the residual of the EKF to exceed a threshold determined through simulations. A solver for typical Linear-

Quadratic Regulator (LQR) problems has also been recently implemented.

On August 19, 2006 (TS4), all the necessary SPHERES hardware for performing autonomous docking maneuvers between two satellites was confirmed ready for experimentation onboard the ISS. A series of four autonomous docking experiments, with cooperative and uncooperative (drifting) targets, took place that same date.Error: Reference source not found,51,52 Experiments with fixed and drifting targets were successful (the Velcro patches on satellites, see Figure 8, were joined). Following this test session, the addition of a fault detection module using the navigation filter innovation (residual) significantly improved the robustness of the global estimator to bad measurements caused by multi-path.

On the following test session (TS5 on November 11, 2006), five docking experiments were attempted.Error:

Reference source not found,Error: Reference source not found,51 They were all successful and resulted in closely aligned contact. Two of these experiments were performed using a safe approach trajectory, which is robust to failures occurring a few seconds prior to contact, when safety is critical (passive abort).53 Two others were performed with a tumbling target, resulting in the first autonomous docking with a tumbling target ever achieved in microgravity (Figure 12). All docking experiments were successful in spite of unexpected measurement errors, induced by the camera flash interfering with the infrared components of the navigation system when taking pictures of the satellites. These measurement errors were successfully detected and rejected online by the autonomous fault detection system. Therefore, these experiments went beyond their intended objectives and demonstrated the capability of the GN&C architecture implemented on SPHERES to perform well even in the presence of navigation sensor errors. These accomplishments illustrate the benefits of maturing GN&C technologies on-orbit in a risk-tolerant environment (inside of the ISS). They were possible due to the use of small satellites.

Many iterations were performed in order to obtain an appropriate sequence of GN&C modes (Figure 13). The exit conditions for each mode were the result of multiple iterations to ensure timely transitions while maintaining the satellites along an appropriate trajectory. This experimentation permitted the tuning of different parameters in each

GN&C module (controller gains, estimator process noise, minimum impulse bit) to ensure good performance. The following lessons were learned:

25  A holding point (berthing) is necessary at a safe but close proximity to the target to make precise alignment

corrections prior to capture.

 The closer to the target, the more constraining the exit conditions must be on each GN&C mode, to ensure that the

tracking errors are within the tolerance of the capture mechanism prior to contact.

 The controllers must be tuned together with the exit conditions to ensure a smooth and timely transition between

modes at a reasonable fuel cost.

Formation flight

The prohibitive costs of free-flyer demonstration missions have prevented any three satellite formation flight demonstrations in space, but the availability of small satellites in the risk-tolerant environment of ISS now enables such tests in microgravity. Formation flight (FF) algorithms have been demonstrated with two and three satellites. 54 The iterative nature of SPHERES algorithm development has been exploited at each step. Initial FF tests controlled the 3-

DOF attitude between two satellites. These tests allowed the team to identify the operational drivers for formation flight with SPHERES: communications and synchronization. Two-satellite 6-DOF formation control followed with tests to demonstrate the use of acknowledgement packets with timing information to synchronize the satellites. The next incremental tests demonstrated formation flight of three satellites. The tests demonstrated triangular formations prescribing a circle. These initial tests utilized pre-planned trajectories with classical PID-style low-level control laws to maintain the formation; each satellite operated independently of each other.

The next maneuver, illustrated in Figure 14, involved changing the plane orientation of the three satellites; that is, while the satellites initially prescribed a circle parallel to the deck/overhead walls, the formation reconfigured to a plane perpendicular to the deck/overhead walls of the ISS. This “plane change” maneuver was demonstrated with multiple high-level algorithms: simulation of a rigid body (resulted in low fuel efficiency), optimal fuel maneuver

(resulted in a high-risk path), and a safer non-optimal transfer. Figure 14-a shows a picture of a run of the optimal fuel maneuver (a six minute experiment); Figure 14-b and Figure 14-c compare the desired and measured path respectively.

The tests were performed to begin research on the high-level commands of FF systems. These tests already give a new insight into the differences in high-level control maneuvers between large rigid-body satellites and distributed formation flight satellites: while the rigid-body fuel optimal maneuver is to rotate about a primary axis (usually using a reaction wheel or gyroscope), a formation flight system’s optimal fuel maneuver involves complex paths which exit

26 and re-enter the formation. Future tests will involve the use of relative sensing between the satellites, collision avoidance maneuvers, and on-line path planning, until SPHERES has been used to mature an essential range of GN&C control algorithms for future space telescopes.

Summary

This chapter provides an overview of some of the research that the MIT-SSL has been performing in the development and maturation of satellite formation flight technology. These research activities span the satellite formation flight engineering development lifecycle from mission architecture, to development of enabling technology, to on-orbit technology maturation. Through the use of micro-satellites in the ISS, modular software and a modular

GN&C architecture, the MIT-SSL performs maturation of technologies necessary for autonomous formation flight and close-proximity operations. As a result, the MIT-SSL has developed some perspectives on the answers to the questions posed in the introduction:

 What current and future missions will require formation flight?

The MIT-SSL is specifically addressing the formation flight needs of sparse aperture arrays ranging

from radio to X-ray wavelengths. In addition, formation flight is a key component of inspection,

docking, assembly and repair missions.

 What are the challenges in formation flight?

While the challenges are numerous, the MIT-SSL has chosen to specifically address propellant-less

formation flight, autonomous docking, and cost-effective technology maturation. As others work to

develop new hardware (sensors, computer, etc), the MIT-SSL has started work to mature robust

estimation, control, and autonomy algorithms that will enable micro-satellites to synthesize large

spacecraft, counter to the normal process where these algorithms are the last to be tested.

 How does one mature formation-flight technology?

The MIT-SSL has found that the ISS can provide a risk-tolerant, long-duration, microgravity research

environment that permits the maturing of formation flight technology at a cost that is comparable to

ground-based facilities. Scientists can change the notion that space research cannot take risks by

continuing to push the envelope in the use of this important new facility like the MIT-SSL did.

27  Can propellant-less techniques be employed to reduce fuel consumption during formation flight?

The research points to yes. EMFF is definitely a promising method, although it does remain a

technology challenge. Other methods also being studied include tethers, electro-statics, and the

efficient use of natural orbital motion.

The goal of the satellite formation flight program at the MIT-SSL is to develop new engineering practices that help the next generation of space systems to keep pace with the emerging needs of new missions. If developed and matured properly, these technologies will shape future space missions that will revolutionize the way we explore our universe.

Acknowledgements

The MIT Space Systems Laboratory and the authors acknowledge the support of the DoD/USAF throughout the development of the SPHERES and EMFF programs. We thank the DoD Space Technologies Program, which is the

SPHERES Payload Integration Manager with NASA Payload Operations. Multiple research centers, including Goddard

Space Flight Center, Marshall Space Flight Center, Ames Research Center, the Jet Propulsion Laboratory, and the

Naval Postgraduate School have been contributors as part of the SPHERES Guest Scientist Program and as funding agents. We acknowledge the support of the Lockheed Martin Advanced Technology Center in the development of

EMFF and docking technologies. We also thank the NASA ISS support teams at JSC, KSC, and MSFC and the many

ISS crew members who have and will operate SPHERES.

Keywords for the index

SPHERES, EMFF, GN&C architecture, ISS experiments, maturation, GN&C mode, autonomous docking, formation flight, close-proximity operations, HTS-EMFF, Micro-EMFF, principles, on-orbit laboratory

28 List of Figures

Figure 1. Concept for robotic assembly of a segmented telescope using EMFF. 4

Figure 2. US astronaut Michael Lopez-Alegria operates SPHERES aboard the ISS during March 2007 (Picture by

NASA). 5

Figure 3. EMFF a) attraction and repulsion, and b) shear and torque. 7

Figure 4. HTS-EMFF can provide precision control at distances up to 100 meters. 8

Figure 5. a) HTS-EMFF and b) Micro-EMFF testbeds. 10

Figure 6. Mass breakdown for various propulsion systems, four hour rotation period, 75 m baseline. 11

Figure 7. SPHERES hardware aboard the ISS with US astronaut Jeff Williams in November 2006 (Picture by NASA).

Figure 8. SPHERES CAD view and assembled satellite with subsystems. 18

Figure 9. SPHERES flight software verification process.52 21

Figure 10. Representative results of the SPHERES Extended Kalman Filter: a) Jumps in the state estimates and b) temporary divergence of the state estimates. 26

Figure 11. The SPHERES GN&C architecture for autonomous close-proximity operations.50 52 29

Figure 12. Footage of the first autonomous docking to a tumbling target in microgravity. 31

Figure 13. Mode sequencing for an autonomous docking experiment to a tumbling target. 32

Figure 14. The three SPHERES satellites during a formation flight initialization and plane-reconfiguration test inside the ISS: a) NASA photograph, b) desired path, and c) telemetry data. 34

29 List of Tables

Table 1. Sample of available facilities for -g research...... 12

Table 2. SPHERES satellite properties...... 18

Table 3. Research iterations from Test Sessions 1 to 5...... 25

Table 4. Review of automatic and autonomous missions involving close-proximity operations...... 27

30 31 1 S. Unwin, “Searching for planets with the Space Interferometry Mission”, NASA report, Jet Propulsion Laboratory,

Doc ID 20000056082 (2000).

2 C. A. Beichman, N. J. Woolf, and C. A. Lindensmith, ed., “The Terrestrial Planet Finder (TPF): A NASA Origins

Program to Search for Habitable Planets”, Jet Propulsion Laboratory Publication 99-003 (1999).

3 E. M. C. Kong, “Spacecraft Formation Flight Exploiting Potential Fields”, MIT Department of Aeronautics and

Astronautics, Ph.D. Thesis, SSL Report #02-02 (June 2002).

4 S. A. Schweighart, “Electromagnetic Formation Flight Dipole Solution Planning”, MIT Department of Aeronautics

and Astronautics, Ph.D. Thesis, SSL Report #09-05 (June 2005).

5 M. Neave, “Dynamic and Thermal Control of an Electromagnetic Formation Flight Testbed”, MIT Department of

Aeronautics and Astronautics, S.M. Thesis, SSL Report #07-05 (June 2005).

6 U. Ahsun, “Dynamics and Control of Electromagnetic Satellite Formations”, MIT Department of Aeronautics and

Astronautics, Ph.D. Thesis, SSL Report #12-07 (June 2007).

7 R. J. Sedwick and S. A. Schweighart, “Propellantless Spin-Up and Reorientation of Tethered or Electromagnetically

Coupled Spacecraft,” paper 4849-26 presented at the SPIE Highly Innovative Space Telescope Concepts

Conference (August 2002).

8 S. A. Schweighart and R. J. Sedwick, “Propellantless Formation Flight Operations in Low Earth Orbit Using

EMFF,” paper AIAA-2006-5579 presented at the Space Ops 2006 Conference (June 2006).

9 R. J. Sedwick and S.A. Schweighart, “Electromagnetic Formation Flight,” paper AAS 03-005 presented at the 26 th

Rocky Mt. Guidance and Control Conference (February 2003).

10 D. W. Kwon and D. W. Miller, “Electromagnetic Formation Flight for the Terrestrial Planet Finder,” paper 5905-41

presented at the SPIE International Symposium on Optics & Photonics (July 2005).

11 E. M. C Kong, D. W. Kwon, S. A. Schweighart, L. M. Elias, R. J. Sedwick, and D. W. Miller, “Electromagnetic

Formation Flight for Multi-Satellite Arrays,” J. of Spacecraft and Rockets 41, 4 (2004).

12 D. W. Kwon, “Electromagnetic Formation Flight System Design,” presented at the 6th IAA Symposium on Small

Satellites for Earth Observation, IAA-B6-0602, Berlin, Germany (April 23-26, 2007).

13 L. M. Elias, D. W. Miller, and D. W. Kwon, “Electromagnetic Formation Flight Dynamics including Reaction

Wheel Gyroscopic Stiffening Effects,” accepted to J. of Guidance, Control and Dynamics (2006).

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