Microsat Ground Test Vehicle

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Micro-Satellite Ground Test Vehicle for Proximity and Docking Operations Deve1opmentl.t A.G. Ledebuhr, L.C. Ng, M.S. Jones, B.A. Wilson, R J . Gaughan, E.F. Breitfeller, W.G. Taylor, J.A. Robinson, D.R. Antelman, and D.P. Nielsen Lawrence Livermore National Laboratory P.O. Box 808, L-491, Livermore, CA 94550 [email protected] (925) 423-1 184 Abstract-This paper updates the ongoing effort at Lawrence Livermore National Laboratory (LLNL) to develop autonomous, agile micro-satellites (MicroSats). The objective of this development effort is to develop MicroSats weighing only a few tens of kilograms, that are able to autonomously perform precision maneuvers and can be used telerobotically in a variety of mission modes. The capabilities under development include satellite rendezvous, inspection, proximity-operations,fonnation-flying, docking, and servicing. The focus of this paper is on the development of several ground test vehicles and the demonstration of proximity operations on a dynamic air bearing table. TABLE OF C0”TS 1. INTRODUCTION 2. POTENTIAL MISSIONS 3. SUPPORTING TECHNOLOGIES 4. DESCRIPTION OF VEHICLES 5. DESCRIPTION OF GROUND TEST CAPABILITIES 6. PROXIMITY AND DOCKING EXPERIMENTS 7. SUMMAlzY 1. INTRODUCTION This paper describes a micro-satellite technology development effort, which extends work sponsored by the Air Force three years ago, and which has been continued most recently through Laboratory Directed Research and Development (LDRD), LLNL’s internally funded research. The purpose of this work is to demonstrate critical capabilities and technologies necessary for proximity- operations and formation-flying of micro-satellites, with the goal of demonstrating a successful autonomous docking of a MicroSat with a target satellite. These capabilities will open up a host of potential new missions that could revolutionize robotic operations in space. However, before the capabilities can be seriously considered for future orbital missions, they must be comprehensively developed and demonstrated on the ground. To date, LLNL has developed several unique engineering test-bed vehicles and a dynamic air bearing test capability, that enables four degrees-of-freedom (4DOF) vehicle motion on an air rail and 5DOF motion on an air table. This apparatus allows vehicle maneuvers in a simulated zero-gravity environment that enables high fidelity testing of vehicle hardware and software systems in an integrated manner. Along with the development of this test capability, work has also focused on the development of on- board integrated proximity-operations sensor packages, avionics for system control, and the image processing, guidance, navigation and control software necessary to operate these systems. Initial tests have successfully demonstrated a soft docking capability on a 4DOF air rail using an imaging camera and an active laser ranger sensor. With the addition of a vision-based 3D stereo ranging system, we have also demonstrated successful docking on a 5DOF air table with a moving target. This represents the development of a new capability for micro-satellite vehicles. 2. POTENTIAL h.IISSIONS Potential missions for MicroSats center on space “logistics” missions such as rescue and servicing, that will require vehicles with the ability to perform a variety of functions autonomously or semi-autonomously (i.e., man-in-the-loop mode). These include rendezvous, inspection, proximity- operations (formation flying), docking, and robotic servicing functions (refueling, repowering or repairing). Figure 1 shows the various MicroSat missions of interest. Rendezvous & lnsnection Proximltu-Onerations Dockina Servicina: Refuelina.ReDair. Retrieval Figure 1 - Potential missions of a MicroSat in LEO Each of these mission functions requires key technical capabilities. For example, rendezvous with a space asset by performing orbit matching requires precision maneuvering. Inspection of a space asset by flying to dif€erent viewpoints of an inspection geometry requires precision Microsat positioning, pointing, tracking, and imaging. A satellite ~~ 0-7803-6599-2/01/$10.00 8 2001 IEEE *Updated December 15,2000 5-2493

Transcript of Microsat Ground Test Vehicle

Page 1: Microsat  Ground Test Vehicle

Micro-Satellite Ground Test Vehicle for Proximity and Docking Operations Deve1opmentl.t

A.G. Ledebuhr, L.C. Ng, M.S. Jones, B.A. Wilson, RJ . Gaughan, E.F. Breitfeller, W.G. Taylor, J.A. Robinson, D.R. Antelman, and D.P. Nielsen

Lawrence Livermore National Laboratory P.O. Box 808, L-491, Livermore, CA 94550

[email protected] (925) 423-1 184

Abstract-This paper updates the ongoing effort at Lawrence Livermore National Laboratory (LLNL) to develop autonomous, agile micro-satellites (MicroSats). The objective of this development effort is to develop MicroSats weighing only a few tens of kilograms, that are able to autonomously perform precision maneuvers and can be used telerobotically in a variety of mission modes. The capabilities under development include satellite rendezvous, inspection, proximity-operations, fonnation-flying, docking, and servicing. The focus of this paper is on the development of several ground test vehicles and the demonstration of proximity operations on a dynamic air bearing table.

TABLE OF C 0 ” T S

1. INTRODUCTION 2. POTENTIAL MISSIONS 3. SUPPORTING TECHNOLOGIES 4. DESCRIPTION OF VEHICLES 5. DESCRIPTION OF GROUND TEST CAPABILITIES 6. PROXIMITY AND DOCKING EXPERIMENTS 7. SUMMAlzY

1. INTRODUCTION

This paper describes a micro-satellite technology development effort, which extends work sponsored by the Air Force three years ago, and which has been continued most recently through Laboratory Directed Research and Development (LDRD), LLNL’s internally funded research. The purpose of this work is to demonstrate critical capabilities and technologies necessary for proximity- operations and formation-flying of micro-satellites, with the goal of demonstrating a successful autonomous docking of a MicroSat with a target satellite. These capabilities will open up a host of potential new missions that could revolutionize robotic operations in space. However, before the capabilities can be seriously considered for future orbital missions, they must be comprehensively developed and demonstrated on the ground. To date, LLNL has developed several unique engineering test-bed vehicles and a dynamic air bearing test capability, that enables four degrees-of-freedom (4DOF) vehicle motion on an air rail and 5DOF motion on an air table. This apparatus allows vehicle maneuvers in a simulated zero-gravity environment that enables high fidelity testing of vehicle hardware and software systems in an integrated manner. Along with the development of this test capability, work has also focused on the development of on- board integrated proximity-operations sensor packages,

avionics for system control, and the image processing, guidance, navigation and control software necessary to operate these systems. Initial tests have successfully demonstrated a soft docking capability on a 4DOF air rail using an imaging camera and an active laser ranger sensor. With the addition of a vision-based 3D stereo ranging system, we have also demonstrated successful docking on a 5DOF air table with a moving target. This represents the development of a new capability for micro-satellite vehicles.

2. POTENTIAL h.IISSIONS

Potential missions for MicroSats center on space “logistics” missions such as rescue and servicing, that will require vehicles with the ability to perform a variety of functions autonomously or semi-autonomously (i.e., man-in-the-loop mode). These include rendezvous, inspection, proximity- operations (formation flying), docking, and robotic servicing functions (refueling, repowering or repairing). Figure 1 shows the various MicroSat missions of interest.

Rendezvous & lnsnection Proximltu-Onerations

Dockina Servicina: Refuelina. ReDair. Retrieval

Figure 1 - Potential missions of a MicroSat in LEO

Each of these mission functions requires key technical capabilities. For example, rendezvous with a space asset by performing orbit matching requires precision maneuvering. Inspection of a space asset by flying to dif€erent viewpoints of an inspection geometry requires precision Microsat positioning, pointing, tracking, and imaging. A satellite

~~

’ 0-7803-6599-2/01/$10.00 8 2001 IEEE *Updated December 15,2000

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rescue might involve docking, repairing or refueling the satellite, followed by a departure, and post-rescue inspection. The rescue mission requires the following specific capabilities: precision guidance, navigation, and control; precision ranging; high-resolution imaging; and some type of micro-robotic manipulation. For example, a variety of robotic arms could be used to enable the MicroSat to perform a physical dock with a target satellite. Figure 2 illustrates one such approach.

Figure 2 - Artist conception of a MicroSat docking maneuver

Here a MicroSat deploys four mechanical arms to grapple the launch vehicle interface flange as it lands on the target satellite. Once docked, a precision 6DOF manipulator (actuator) could be used to align and plug an external connector into a targeted satellite’s umbilical connector for data collection, diagnostic measurements or re-powering. This servicing operation could be. performed teleraibotically from the ground, to provide the flexibility and problem solving of a human presence. Other space logistic operations, such as the collection and de-orhiting of hazardous space debris (junk), require precision vehicle guidance, navigation and control, and precision horning.

Formation flying, flying in concert with a space object or another Microsat, requires station keeping, positioning, and precision state vector estimation.

Close-up inspection missions offer a means to remotely determine a satellite’s health and status, and can collect data that cannot be obtained from the ground. For example, a laser vibration sensor can determine bearing wear on moving components. like momentum wheels, control- moment gyros or solar array drives. Infrared sensing can observe thermal nonuniformites and detect le& and differences in thermal insulation. Other missions may involve physically moving or towing a space object to a different orbit, constructing a 3D surface image of the object using stereo vision, estimating the object mass properties, and perhaps even reconstructing the internal structure of the object from 3D computed tomography. There aue many potential missions that will become apparent once the basic system capability becomes routinely available.

A previous study [l] has shown that a Microsat with 300 m / s of velocity change (or AV) is about the ininimum necessary to carry out a basic mission, assurning the

McroSat is placed in the same orbit as the satellite to be inspected. Clearly vehicles with larger AV offer multiple mission capabilities and the ability to change orbits. In addition, if a spaceborne refueling capability is developed, then the mission utility of the MicroSat can be greatly extended by positioning refueling stations at appropriate orbits. In order to demonstrate many of these proximity operation capabilities, LLNL, under the sponsorship of the Air Force Research Laboratory, has developed several MicroSat prototype vehicles for ground testing using a number of state-of-the-art technologies to support future Air Force missions. One mission concept calls for the MicroSat to demonstrate proximity-operations near a space object including the apability to soft dock. A proposed mission scenario is shown in Figure 3.

I MicroSat Maneuvering Sequence within - .

a 10m radius 2

6

1

4 Proximity Maneuvers:

1. Octahedron imaging 1-2-3456

2. Circular stereo imaging C63156-C64126

3. Soft docking 6-7-8-9

4. Formation flying 1.4,6,2

Figure 3 - A conceptual scenario for proximity inspection and soft docking mission

An agile MicroSat will eject from the carrier vehicle to a distance of about 1Om and conduct a series of proximity operations within the lOm-radius sphere. For example, proximily inspection sequence via points at 1-2-3-4-5-6; circular stereo imaging via circles 6-3-1-5-6 and 6-4-1-2-6; soft landing via points 6-7-8-9; and formation flying at points 1, 4, 6, and 2. The mission is completed when the MicroSert has successfully demonstrated repeated soft- dockings with the carrier vehicle.

3. SUPPORTING TECHNOLOGIES In describing the technologies involved in the ground test setup for proximity and docking experiments, it should be noted that two independent autonomous vehicles were used, one as a target vehicle and the other as the docking vehicle. These engineering test vehicles (ETV), designated as the ETV-150 and ETV-250, will be detailed in a later section. They are introduced here simply to explain that there are differences in the technologies between them in terms of propulsion configurations, sensor suites, and avionics architecture. The ETV-150 is the older test vehicle, and is used as ithe target for the ETV-250, the new docking vehicle.

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Propulsion The ETV-150 and ETV-250 ground test vehicles are used in indoor air table and air rail experiments, both divert and attitude control thrusters utilize high pressure cold gas (N2) stored onboard in four small gas bottles. To improve total AV capability, we have built other vehicles using H202 as the propellant [3].

Function Image Field of Pixel Format View Size

Inspection 1024x1024 40°x400 10 Camera Monochrome micron

square Stereo 640x480 2Oox16" 7.9

Cameras Color (VGA) micron square

Sensors

Sensor Functions-For a typical InspectionfDocking mission, the sensors must support the various mission phases that the MicroSat must execute. In the initial Rendezvous phase, the sensors must first acquire the satellite to be inspected and then carry out the necessary proximity operations during the Inspection and Docking Phases. Experiments of the type described in this paper have focused on proximity operations and docking, since the current indoor setup restricts operation to distances under 10 meters. Sensor functionality within the context of these experiments will be described below.

Star Tracker A

Focal Length 9 . 9 "

12.6mm

Monochrome Video Camera

Color Stereo

I Laser Ranger

Figure 4 - Multi-function Integrated Proximity Operations Sensor package

The sensors are incorporated into an Integrated Proximity Operations Sensor (IPOS) package that provides data for guidance, navigation, and control, and also provides images for transmission to the ground. A variety of sensor assemblies could be integrated into this package. The specific configuration used, shown in Figure 4, has two different visible imaging sensor systems, a laser ranger, a star tracker, and an IMU. This first generation IPOS system has been optimized for our close-in docking experiments, and our long-range acquisition camera and long-range LADAR system are not incorporated into this system.

The sensors complement each other, providing detailed information suitable for a variety of proximity mission operations. The visible cameras are based upon commercial CMOS active-pixel sensor (APS) imaging detectors. The laser ranger is a modified commercially available system for close-in ranging. Although not used in the indoor experiments described here, the star tracker is a wide-field system, originally developed and flown on the successful Clementine I Lunar mapping mission [4-61. The M U is the commercially available LN200 from Litton.

Visible Imagers-With features as shown in Table 1 below, the Visible Inspection Camera is used to initially acquire the

target vehicle in the docking experiments described later. After acquisition, this camera is used to track an LED on the docking ball of the target fixture. The Inspection Camera delivers digital data at a frame rate of 10 Hz, although for these experiments all cameras were used at a 3.3 Hz rate (300ms cycle). It should be noted that the megapixel CMOS APS detector used in this camera is capable of operating at a 500 Hz frame rate.

Table 1. Visible Camera Characteristics

Once track is established via the Inspection Camera, the Color Stereo Pair is given a region-of-interest (ROI) within which to find the target. These are also digital cameras that provide data at a frame rate of 30 Hi. The cameras are independent, and may be used individually (in monocular mode) or when used together (in stereo mode), can provide passive ranging information and also provide a telepresence capability, for remote man-in-the-loop operation of the Microsat.

CMOS APS detectors were chosen due to the large reductions in mass, volume, and power possible with these devices in comparison with standard CCDs. In addition to the savings for each single camera, there will be additional savings in the avionics because of the commonality of the camera interfaces. The CMOS sensors are extremely flexible and provide selectable, gain, frame rate, and integration times. In addition, they offer windowing or on-chip pixel sununing allowing more rapid image acquisition or easier downlink of a region of interest.

Ranger-The Proximity Ranger is a laser diode based distance measurement sensor accurate to less than 0.5 cm from 25 m to 0 m (Docking). The laser ranger incorporates a laser diode transmitter whose output is detected by a photo- diode detector system. The ranging system operates by emitting a collimated laser beam which is reflected off the target fixture using retro-reflecting tape and the return beam is measured using the photo-diode detector. The output of the detector is inverted and used to control the output of the laser diode transmitter. When there is no signal, the diode laser is turned on, and when return photons are detected, the laser transmitter is turned off. This forms an oscillating circuit whose period is a function of the time of flight of the laser beam. Measuring the oscillation kequency provides a measure of the target range to relatively high accuracy. The benefit of this approach is its operation down to very short ranges that enable its application as a ranger for docking applications. In addition to distance output, the Unit provides signal strength, ambient light level, and sensor temperature outputs. The ranger is a commercial unit repackaged for LLNL to reduce mass and volume.

IMU-The LN200 is currently being used in the ground test vehicles due to its performance and relatively low cost. This IMU has a measured drift rate of approximately 1 "hour and

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is a ruggedized package that optionally can be ordered suitable for use in space. There are a number of other lMUs that could be utilized in these vehicles including some new MEMS systems that have reduced mass and power, but at the expense of performance. Measured drift rates falr MEMS lMus are currently at the 10°/hour rate, which if periodically updated by a Star Tracker is quite acceptable for MicroSat applications.

WFOV Star Tracker-As stated previously, the Star Tracker, although mounted on each of the test vehicles. was not incorporated into the indoor experiments. It is, however, a key part of the MicroSat sensor package, and will be utilized in nighttime outdoor ground tests. This wide-field-of-view (WFOV) Star Tracker provides inertial orientatiaa of the vehicle and updates for the IMU. The Star Tracker camera in conjunction with Stellar Compass software can provide a quaternion pointing accuracy of 450 prad, 30 in the roll axis, and 90 prad, 30 in pitch and yaw. The Star Tracker field of view is large enough to contain at least five bright stars ( M ~ 4 . 5 ) in any orientation. Single images are processed to identify unique stellar patterns and provide the determination of the inertial orientation of the MicroSat in real-time. The collection aperture of the lens is “ i z e d for the greatest possible light gathering capability. At FA.25, Mp4.5 GO stars provide an integrated star signal that is 15 times the electronic noise fiom the focal plane. This level of signal gathering capability, matched with the wide field of view, ensures a 99.9% probability that five stars above minimum threshold will be available: for the algorithm set for all possible quaternion pointing vectors. This allows the Star Tracker to handle the “lost in space” condition with a single star image fiame and no other apriori knowledge of attitude. A future Star Tracker generation may replace the previous CCD array with the same 1024 X 1024 pixel CMOS APS array used in the visible camera. The modified Star Tracker would have a 32x32 degree coverage using the existing lens design.

Avionics

Figure 5 - Avionics Architecture

The MicroSat avionics system described here can accommodate a variety of sensor suites and can make use of many types COTS U0 modules, depending on the end

application. The avionics architecture is based on a high performance PowerPC processor and CompactPCI bus as illustrated in Figure 5 (architecture shown as configured for indoor experiments). The PowerPC family is widely used in embedded systems for its performance and low power features. In addition, commercial versions of the PowerPC 603e have been tested, demonstrating significant inherent radiation tolerance.

The CompactPCI bus is a high-performance, processor independent U 0 bus, which provides an efficient path for processas upgrades. The system supports modern, real-time embedded software development environments. This design allows rapid code development, debugging, and testing. Its modular design leverages COTS technologies, permitting early integration and test of both hardware and software elements;. The chosen architecture provides a high performance solution for current and future MicroSat missions.

Processor Module-The MicroSat processor module shown in Figure6 contains a high-performance PowerPC 603e RISC ClPU and utilizes a 33 MHz, 32 bit data path CompactPC1 bus. For a flight system, the processor would be a COTS module with modifications for thermal management and radiation tolerant parts as needed.

Figure 6 - Processor Module

Wireless Communications-In the MicroSat ground test vehicles, wireless Ethernet is used to communicate with ground support equipment. A CompactPCI Ethernet module is connected via lObaseT to a wireless transceiver. This provides upload, download, and intervehicle communications capability. For a space mission, the specific communications and telemetry capability required would determine the onboard transceiver selected as well as the type of ground stationed employed for the mission.

Image Acquisition and Processing-The current digital fkme griabber module is a COTS board designed to provide a high-performance image acquisition and data handling interface between the CompactPC1 bus and high-speed digital cameras. It features 8 to 16 bit pixels, pixel clock rates of up to 20Mhz, multiplexed operation for cameras sharing the video channel, Automated Imaging Association (AIA) digital camera compatibility, a Look Up Table (LUT) to allow real-time hardware functions such as thresholding, a 16K by 32 bit FIFO buffer to support DMA over the CompaciPCI bus, and a Region of Interest (ROI) acquisition mode. For additional image processing capability, the next- generation *e buffer module will be a commercial DSP

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board with local storage and an add-on mezzanine designed for hardware-based image compression.

Vehicle Control-The interface modules shown in Figure 6 provide connection, control, and acquisition functions for the Microsat guidance and navigation elements. The current design uses Industry Pack (IP) carrier boards on which IP modules interface with the various subsystem components. Serial IP modules are used to communicate with the IMU and Ranger components. A parallel digital U 0 module controls the valve drivers. Programmable digital U 0 modules are used for camera control and to operate the grappler used in docking. Onboard housekeeping functions are monitored using an analog-to-digital converter IP module.

Power Distribution System-The current Microsat ground test vehicles have used low cost rechargeable Ni-Cad cells. For future developments, a trade study is planned to examine the state-of-the-art in rechargeable battery technologies, lightweight power supplies, and advanced solar cell technologies that can be used for flight vehicles. Specific choices will await future flight design efforts.

Software The MicroSat software development environment uses the VxWorks real-time operating system (RTOS). VxWorks is a commonly used, well-tested, RTOS that provides a rapid development environment for integration of new software modules. It is also portable among many processors, and has been used in space applications including JpL's Mars Pathfinder and the Clementine I spacecraft. Figure 7 shows the hierarchical organization of our mission software.

Services

Figure 8 - Guidance and control interfaces

The stereo ranging code was developed in two steps. First, in the calibration step, a target board was imaged by the VGA cameras mounted in their fixture on the probe. Image pairs were taken at 1, 2, and 2.4 meters range. The 50 cm x 50 cm target board had a white background with a 12-row x 13-column grid of black circles (1.25 cm diameter) spaced 2.54 cm apart (center to center). Using a laser aimed along the centerline (x-axis) of the probe, the target board was kept from shifting sideways at the three range locations. A transformation matrix for mapping left camera and right camera target pixel locations into 3-D world coordinates was then generated by using the same 8 x 8 sub-grid of dots in each of the stereo image pairs and knowing the actual measurements from the probe to those dots. The sub-grid was chosen because that was the area that fit in the field of view at the 1 meter range. Then, a simple transformation function was added to the real-time code, which converted the pixel location of the LED target, in both the left and right VGA cameras, into 3-D world coordinates (leftlight, upldown, nearlfar). The nearlfar element provided stereo range.

Figure 7 - Multi-layered software architecture

Many software modules for GN&C, imaging, target tracking, and other real-time operating codes can be adapted for the representative satellite mission. Figure 8 illustrates the structure of a set of GN&C software interfaces for a given MicroSat architecture (not all elements were used in the indoor experiments).

Stereo Ranging Software-There was stereo image ranging software included in the system, which provided the range measurements during the initial fly-in phase (6 - 1.8 meters). Specific phases of the docking experiments will be described in a later section.

4. DESCRIPTION OF VEHICLES

As previously stated, two autonomously functioning engineering test vehicles were used in this work. Both of these vehicles, the ETV-150 and the ETV-250, utilize compressed nitrogen gas as the reaction propellant and each vehicle utilizes 16 small thrusters arrayed in a coupled arrangement that affords a full 6DOF motion in a micro- gravity environment. Seeker and Avionics payloads are mounted on the front and rear portion of these vehicles. Spherical air bearings are mounted at the center of gravity of each vehicle. For the ETV-150, there is a hemispherical element that fits around the central liquid manifold. For the ETV-250, this is a complete spherical ball designed to allow for larger tilt angles.

ETV-150

In the ETV-150 the 16 thrusters are.mounted in four sets of four thrusters, with a pair of pitch and yaw thrusters at each end of the tank structure. Two pairs of roll and axial thrusters are mounted orthogonal to the pitch and yaw pairs on the sides of the vehicle. The location of these thrusters can be seen in Figure 9. The ETV-150 structural configuration is based on a pair of in-line liquid propellant tanks that form the main structural backbone of the vehicle.

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Four high-pressure bottles are arrayed about the center of the vehicle between the location of four lateral divert thrusters (only two are mounted for these experiments). The liquid propellant tanks were designed to carry rocket grade hydrogen peroxide at 85%-90% concentration level. For the docking experiments only compressed nitrogen is used.

Figure 9 - ETV-150 inspection vehicle

The payloads are balanced on the ends of the liquid tanks, with a Seeker assembly on the forward end and Avionics Assembly at the rear of the vehicle. The target fixture is attached to the Avionics assembly. The fmture is comprised of a backboard covered with retro-reflective tape for the laser ranger and a rod holding a 3 inch diameter docking ball that has an LED mounted on its outer surface, which is on axis with the rod. The ETV-150 uses a VME electronics crate that accommodates an older generation of support electronics including a commercial PowerPC 603e processor board. The sensors on the ETV-150 are a residual hardware from previous program and include a Clementine WNisible camera along an earlier generation UVNisible camera that was a flight spare fiom the Brilliant Pebbles space-based interceptor development program. This earlier generation sensor is used as the main tracking camera on this vehicle. No active ranging sensors are utilized by the target vehicle.

ETV-250

In the design of this test vehicle we wanted to minimize the overall volume of the vehicle in order to enhance fllexibility for hture launch opportunities of this system. Ratlher than the in-line tank design we went to a design with two pairs of propellant tanks along the sides of the vehicle and the fiont and rear payloads in-between these components. This resulted in a squarer design form and required a rethinking of the location and configuration of the ACS thrusters. A line drawing of this vehicle is shown in Figure 10. Ini a flight system the gas bottles could be replaced with liquid tanks to provide larger delta-v maneuvers. The 16 ACS thrusters on the ETV-250 that provide independent fine control of both translational position and angular orientation are mounted at the four corners of the vehicle. There are four pairs of opposing vertical thrusters that can provide pitch and roll along with vertical motion in free space. A thruster in each comer is mounted horizontally outward to provide yaw motion as well as lateral translation while rear-pointing thrusters mounted to the fiont of the vehicle (and canted out

at 30 degrees) are used for forward (axial) translation. A mirrored image pair in the rear of the vehicle enables rear (- axial) m'otion. This configuration of axial thrusters was designed to minimize any gas impingement on the docking target velhicle. A lightweight structural frame was designed to mount the gas bottles and the front and rear payload assemblies. This structure also serves as an attachment point for the spherical air-bearing ball. On each side of the vehicle, between the gas bottles are located larger 5-pound cold-gas thrusters that can be used for large lateral impulses. For these experiments, the larger thrusters were not used.

The nose of this vehicle contains both the Seeker and the grappler assemblies. The Seeker includes the on-axis mega- pixel tracking camera, two stereo VGA color cameras, the short-range laser ranger, a Star Tracker and the IMU. The grappler has four arms that are designed to optimize the capture of the target ball during a docking mission. The avionics ,assembly consists of a CompactPC1 eight slot card box that contains most of the control system for the vehicle. This system is approximately half the volume of the previous VME card box. Table 2 shows a comparison between the properties of the ETV-150 and ETV-250 vehicles.

Awtomcs

Star Tracker N2ws 7 7

Divert thrust

Figure 10 - ETV-250 docking vehicle

Grappler Design

The graplpler was designed to enable a positive measure of a successful docking attempt. It incorporated a linear-drive stepper motor that was linked to each of the four grappler arms. The four armed grappler has spherically shaped end paddles that conform to the shape of the target ball. Each paddle contains a pressure-switch, which provides feedback that the ball has been engaged. The choice of a spherical ball was chosen based on the known flexibility of a spherical trailer hitch assembly. We desired to maximize the angular range that the target ball could be approached. Previous hard docking fixtures tended to utilize a three-point mount spaced relatively far apart. This approach tends to drive up the requirements of both the relative alignment between the two vehicles as well as their angular orientation. Our concept was to nlaximize the angular cone over which the target could be approached and soft grappled. In a hard dock, the vehicles are securely coupled and behave as a new solid body wiih the combined moment of inertia of the two objects. In a soft dock, the two vehicles are loosely coupled and still have some relative angular degrees of fieedom of motion. The requirement for this experimental fixture was to achieve a 90 degree cone angle, within which the target ball

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could be successfully grabbed. This requirement was achieved with this design. Though this fixture was used as a means of scoring the docking experiments, this grappler design could be extended into an operational system. Some proposed changes to this system could provide for a hard dock capability as well as offer electrical andor fluid interconnects.

Table 2. Summary of vehicle characteristics

Qttributes Vehicle

length diameter

mass roll-moment

pitch-moment yaw-moment

Divert yaw/pitch ACS

roll ACS Max. Accel.

Propulsion (N,)

Vehicle Characteristics Unit ETV-150

m 1 m 0.4 kg 23

kg/m2 0.4 kg/m2 2.6 kg/m2 2.6

Nt 5.5 Nt-m 1.6 Nt-m 0.6 m/s2 0.24

Avionics Communication

Processor Speed

Sensors Laserranger Stereo-pairs

IMU

Wireless MVME 1603

100 MHZ

m none m none

LN-200

Imager Software

Stereo-ranging Srappler control

Sofi-docking

Clem2 VIS 384x288

ETV-250

0.6 0.4 20 0.1 1 1

~

11 0.55 0.2 0.55

Wireless CPCI 3603 166 M H Z

0.03-30 0.5-5

LN-200 MPE

1024x1024

Yes Yes Yes ves

5 . DESCRIPTION OF GROUND TEST CAPABILITIES

Ground performance testing is the key to the success of a MicroSat mission. It is crucial to be able to repeatedly practice and test the integrated vehicle's ability to perform precision orientation and translational maneuvers. These tests should include maneuvers to achieve orbit matching, endgame chase, inspection, docking, satellite servicing, and undocking. Ideally, one would l i e to have 6DOF test environment. However, in most cases a 5DOF or 4DOF environment is sufficient.

In order to support the ground testing of integrated MicroSats, LLNL has developed 4DOF and 5DOF dynamic air bearing ground testing facilities [2]. The 4DOF facility is an air rail with three degrees of rotational fieedom and one degree of translational fieedom. The 5DOF facility is an air table with three degrees of rotational fkeedom and two degrees of translational fieedom. These facilities enable low cost repeatable end-to-end performance testing of completely integrated MicroSat test-bed vehicles, and full-

up performance acceptance testing of final flight hardware and s o h a r e before launch.

Dynamic Air Bearing (DAB)

The concept of a dynamic air bearing (DAB) is to integrate translational motions into a traditional 3DOF angular motion air bearing. This enables a vehicle that is equipped with divert engines and attitude control jets, which is sitting on the hemispherical air bearing to achieve 5DOF motion: three rotations and two translations. A fixture of the DAB is shown in Figure 11. The hemispherical air bearing draws air from the three high pressure tanks, which also supply air to the three air pucks. The air pucks provide a cushion of thin air between it and any smooth surfaces such as a thick glass table. The three air pucks, equally distributed on a 7.5" radius circle, can support a total weight of more than 150 kilograms. The DAB itself weighs less than five kilograms. For a 25 kilogram lightweight Microsat, this represents a 20% increase in overall weight or equivalently a 20% reduction in acceleration capability.

Figure 11 - An early prototype of a Dynamic Air Bearing

Dynamic Air Table

Three different generations of test vehicles are shown in Figure 12. For 5DOF experiments conducted on the table, each vehicle sits on an air bearing fixture as described in the preceding section. The hemispherical air bearing allows approximately f15" pitch, f 360" yaw, and f3O"roll while

Figure 12 - Test vehicles on Air Table and Air Rail

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the three linear air bearings enable the vehicle to translate in two dimensions. The table, measuring 1.5m x 7.3m, is covered with plate glass. Supported by a cushion of compressed air, the linear air bearings allow an almost frictionless translation motion to any position on the table. This facility has been operational for over three years and has had over 650 experimental runs performed on it including 3DOF tracking, 5DOF tracking, 360" yaw maneuvers, precision translational maneuvers, and soft docking with a drifting target.

Dynamic Air Rail

The ETV-100, an earlier vehicle built prior to ET'W-150, is also pictured in Figure 12 on the air rail. The 4D0F air rail, as pictured, allowed over 16 meters of translation and was used in this configuration to conduct guided intercept and line-of-sight experiments [7]. The air rail is needed for testing large linear translational motion, such as in proximity inspection maneuvers or in collision avoidance or intercept maneuvers.

3

4

5

6 7

6. PROXIMITY AND DOCKING EXPERIMENTS These experiments were performed on the table described in the previous section. The experiment sequence involved several basic steps, as shown in Table 3 below.

Table 3. Docking experiment scenario and mission timeline

17

35

41

52 56

Step T(sec) * Probe Operation

Open Grappler

Slew (yaw) to acquire LED target ( 5.5 meters away) Fly-in to 1.8 meters away (using stereo ranger) Align for docking head-on (lateral

Perform Final Approach (using laser ranger) Grip the target ball Pull backward while connected, push forward; release Push away fi-om target; returnto "home" orientation

thrusting)

-- Target 'Vehicle

0per:ation Yaw to aim target LED Divert toward comer

Hold attitude

Hold attitude

-~ Hold attitude

Hold attitude Hold attitude --

-~ Hold attitude

Initially, at step 1, the docking vehicle (probe) was located at one end of the table, about six meters away fi-om the target vehicle, and it was held in place by its grappler arms hanging onto a launching fixture, aimed to the side of the table about 90 degrees fi-om the target. Upon receiving the go command, the probe opened its grappler (see Fiigure 13) and initiated a counter-clockwise yaw motion (step 2), while simultaneously, the target vehicle performed its own yaw to direct its LED toward the probe.

Firrure 13 - Vehicle is initially released from stand

While the probe was acquiring the LED in its tracking camera, the target vehicle diverted into the comer of the table, still about six meters from the probe. Here we demonstrated the ability to control two vehicles at the same time, &de still providing a stable target for docking. Figure 14 shows the probe after having completed its 90 degree yaw maneuver. Here the target vehicle is in its stable position. The numbers shown near the overlaid probe path correspond to the steps listed in Table 3.

Figure 14 - Soft docking experiments

After one second of steady aim with the LED centered in the tracking: camera, axial thrusts (step 3) propelled the probe down to 1.8 meters range (as calculated by the stereo ranger code). This fly-in phase lasted about 18 seconds, with a gentle asymptotic final approach to the 1.8 meter goal. While the probe was oriented such that the onboard laser ranger had its spot aimed at the reflective backboard, we

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obtained range validation measurements at 20 Hz. The graph in Figure 17a shows how laser range compared with the calculated stereo range. A sample stereo image pair taken at a 2 meter range is shown in Figure 15. We got excellent agreement around 2 meters, with the largest error

, of about 0.5 meters (10%) at the maximum range. That’s consistent with having used calibration images at 1, 2, and 2.4 meters. Motion of the vehicle was another contributor to the stereo ranger errors. The left and right images were taken 50 msec apart, so rotational or translational movement would cause the LED target to move in one image relative to the other. Taking the image pair at the exact same time would eliminate this effect, but our frame grabber didn’t support multiple unsynchronized image acquisition. The target vehicle was set up to hold its attitude so the line of sight from the probe would be off-angle from a direct head-on engagement at the 1.8 meter range. That was done to demonstrate our ability to correct for such misalignments.

The lateral alignment section of the experiment (step 4) used the offset of the LED center versus the center of the docking ball to estimate the alignment angle in each of the two VGA stereo imagers, via the following equation.

Theta = arcsin( 2 * pixel-offset / ball-diameter(pixe1s) )

The probe’s alignment angle was simply derived as the average of the left and right measured alignment angles from each VGA imager. Lateral thrusting was applied to bring the probe’s alignment angle within +/- 2 degrees for 1.5 seconds. In the run shown here, it took 6 seconds to satisfy that condition from 12 degrees misalignment initially. See Figure 17c. Also, note that while thrusting laterally, the LED target did not stay centered in the tracking camera. That can be seen in the plot of target azimuth (Figure 17e), where it drops from 0.5 down to -4.5 degrees during the lateral alignment step. The two yaw jets on each side of the probe were usurped for lateral thrusting, which cut the yaw control capability in half. Once the lateral thrusts stopped, the full-strength yaw control authority brought the LED back to the center of the tracking camera’s field of view.

In the next sequence (step 5), after achieving alignment, the ranging was handed over to the laser ranger for final closure from 1.8 meters away. That separation was covered in 11 seconds, with primarily axial thruster pulses, however some lateral thrusts were applied to help maintain alignment.

At about 0.6 meters, we stopped getting any stereo range measurements, since the LED disappeared out of the right VGA imager. Target azimuth readings stopped at about 49

seconds, since the tracking camera was tumed off when the laser ranger reading got below 0.5 meters. At that time the grappler was only 10 cm away from being centered around the target ball. At 0.4 meters, the grappler was closed onto the docking ball (step 6). This is the distance between the reflective target backboard and the laser ranger when the grappler arms are set to properly close onto the ball. Figure 16 shows the probe’s final approach to the target.

I . ..

.-

Figure 16 - Probe is executing final docking maneuver

After a couple seconds of pausing, to assure a well closed grappler, we applied an axial pull backward for three seconds (step 7), then coasted together, backward, for another four seconds, and then reversed the thrust, pushing forward to stop the motion. During this time when the probe and target vehicles were connected, they did a shimmying dance as their attitude control systems worked against each other, trying to maintain their own (slightly different) preferred orientations. That can be seen as roll, pitch and yaw rate oscillations starting at about 55 seconds (see Figures 17b, 17d, and 170.

We then opened the grappler, releasing the target vehicle, and pushed away (step 8) with a final one second axial thrust backward. After three seconds, while still coasting away from the target, we reoriented back to the initial attitude for easier placement back into the launching fixture. These final steps can be seen in the plot of laser range (for the pushing away maneuver), and in the yaw omega plot (Figure 170 for the 90 degree yaw back to the initial attitude at 70 seconds.

Figure 15 - Left and right image pairs fkom the stereo cameras

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IC

e

- 6 E v

W w

E 4

2

0

G 6

I I I I i p i 4 4 10 20 30 40 -I' 60 70 80

tlme (sec)

Figure 17a - LasedStereo range and difference

Algrment angle measured w r t doclang surface perpendicular

10

Gv

i o! i o 20 30 2 50 $0 70 80

Time (sec)

Figure 17c - Alignment angle for final approach

Time (sec)

Figure 1% - Vehicle roll rates

GoL ib 20 A. 40 Sb 60 i o i o Time (sec)

Figure 17d - Vehicle pitch rate

Time (sec)

Figure 17e - Seeker AZEL angle Time (sec)

Figure 17f - Vehicle yaw rate

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7. SuMh4ARY

To date, LLNL has successfully developed several functional vehicles and ground demonstrated several critical on-orbit MicroSat logistics operations. They include: autonomous target acquisition, tracking, stereo ranging, homing on to a target, final approach trajectory calculation, soft docking, grappling onto a target, and adding a AV to a target as in a de-orbiting maneuver.

ACKNOWLEDGMENTS The authors would like to thank Drs. John Holzrichter and Rokaya Al-Ayat at the LLNL LDRD ofice for their financial support of this research. We thank Shin-yee Lu at LLNL for use of the stereo calibration software that she had previously developed. For consulting work in the area of image processing, we thank Jamie Markevitch. The authors also would like to thank Ed English, Gloria Purpura, and Donna Bergin of the Missile Defense and Space Logistics Program at LLNL for their support of this work.

The research was part of the MicroSat Technologies Program at LLNL, supported by the U.S. Air Force Research Laboratory. This work was sponsored by the U.S. Government and performed by the University of California Lawrence Livermore National Laboratory under Contract W-7405-Eng-48 with the U.S. Department of Energy.

DISCLAIMER This document was prepared as an account of work sponsored by an agency of the United States Government. Neither the United States Government nor the University of California nor any of their employees, makes any warranty, express or implied, or assumes any legal liability or responsibility for the accuracy, completeness, or usefulness of any information, apparatus, product, or process disclosed, or represents that its use would not infringe privately owned rights. Reference herein to any specific commercial product, process, or service by trade name, trademark, manufacturer, or otherwise, does not necessarily constitute or imply its endorsement, recommendation, or favoring by the United States Government or the University of California. The views and opinions of authors expressed herein do not necessarily state or reflect those of the United States Government or the University of California, and shall not be used for advertising or product endorsement purposes.

REFERENCE [l] Ledebuhr, A.G., Ng, L.C., Kordas, J.F. et al., “Microsat Rescue Demonstration Mission: A Feasibility Study”, UCRL-ID-129880, Jan~ary 30,1998.

[2] Ledebuhr, A.G., “Down-to-Earth Testing of Microsatellites,” Lawrence Livermore National Lab Science & Technology Review, September 1998.

[3] Whitehead, J.C., “Hydrogen Peroxide Propulsion for Small Satellites”, SSC98-Vm-1, The 12th Annual Utah State University Small Satellite Conference, 1998.

141 Kordas, J.F., Priest, RE. et al., “Star Tracker Stellar Compass for the Clementine Mission”, Proc. SPIE Vol. 2466, p70-83, June 1995.

~

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[5] Kordas, J.F., Lewis, I.T. et al., “UVNisible Camera for the Clementine Mission”, Proc. SPIE Vol. 2478, p175-186, June 1995.

[6] Ledebuhr, A.G., Kordas, J.F. et al., ‘WiRes Camera and LIDAR Ranging System for the Clementine Mission”, Proc. SPIE Vol. 2472 p62-81, June 1995.

[7] Ng, L.C. and Ledebuhr, A.G., “Dynamic Air Bearing Guided Intercept and Line-of-sight Experiments”, UCRL- JC-128922, February 28,1998.

BIOGRAPHY Dr. Arno Ledebuhr earned an undergraduate degree in Physics and Math in 1976 jFom the University of Wisconsin and masters and doctorate degrees in Physics from Michigan State University in 1982. Dr. Ledebuhr spent the following four years at the Hughes Aircraft Comvanv and eamed 13 vatents

1 ,

in projection displw te;hnologv. He has been at Lchrence Livermore National Laboratory since 1986 and led the development of advanced sensors for the Brilliant Pebbles interceptor program and the design of the Clementine sensor payload. In 1996 he was the Clementine II Program Leader and is currently the MicroSat Technologies Program Leader.

Dr. Larry Ng received his B.S. and M S . degrees in Aeronautics and Astronautics from the Massachusetts Institute of Technology in 1973, and a PhD degree in Electrical Engineering and Computer Sciences from the University of Connecticut in 1983 under a Naval Undersea Warfare Center NUWC)

Fellowship. In addition, Dr. Ng riceived his commission & an Air Force oflcer in 1973 and served at the Hanscom Air Force Base in Bedford, MA. His work experience includes: four years with General Dynamics Electric Boat Division in Groton, CT, responsible for the development of the TRIDENT submarine digital control systems; seven years at the NUWC where he led the development of the advanced sonar signal processing for the Seawolf submarine. Since 1986, he joined the Lawrence Livermore National Laboratory where he is currently the group leader of the signalhmage processing and control group and is focusing his research in micro-spacecraft guidance and control, integrated ground testing, and ballistic missile defense systems analysis. Dr. Ng is a member of several professional societies, including honorary memberships in Sigma Xi, Tau Beta Pi, and the National Research Council. He has published numerous papers in signal estimation and precision vehicle guidance and control.

Mark Jones h m worked in theJield of electronics since 1977. He holh a B.S. degree in Computer Engineering @om the University of the PaciJc, Stockton, CA. Mark came to LLNL in 1984 where he has worked in electronics engineering, sofhvare development, and system integration on space-related projects including Brilliant Pebbles, MSTI,

Clementine, and Clementine II. Currently, Mark leads the avionics eflort for the MicroSat Technologies Program.

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Bruce Wilson has done computer programming for Lawrence Livermore National Laboratoiy since 1974. He holds a BA in Mathematics from the University of California, Santa Barbara (1973). He has provided sojhvare support for the Brilliant Pebbles and Clementine projects, and currently leads the software eflort for the IMicroSat

Technology Program.

Eric F. Breiqeller has been an electrical engineer at Lawrence Livermore National Laboratory (LLNL) from 1989-1996 and 1997-present. From 1996-1997 he was at Hughes Missile Systems Company. He received his A4S.E.E. in 1988, from Ohio State University, with emphasis on control systems. From 1990-1992 he worked on guidance, navigation, and control (GNC) 6DOF simulations related to the Brilliant Pebbles program. Subsequently, he supported BMDO through the POET, where he developed a 6DOF simulation that was used in trade studies as they related to missile intercept scenarios. While at Hughes (Raytheon) he developed the attitude control system (ACS) for the Exo-atmospheric Kill Vehicle (EKV. Upon retuming to LLNL in 1997 he assuimed the lead GNC engineering position on the former Clementine-11 program (currently the MicroSat Technology program). Precision control and estimation algorithms were designed in a 6DOF environment, and then applied to the fully- functional 5DOF (3DOF ACS + 2DOF translation) hot-gas micro-satellite. FThile not working on missile intercept problems, he has been involved in robotic applications related to the alignment of optical fibers to wave guides, and to beam steering of single-beam linear particle accelerators.

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