1. 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
ledebuhrl@llnl.gov
(925)423-1184
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
autonomouslyperform precision maneuvers and can be used
telerobotically in a variety of mission modes. The
capabilitiesunder 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 operationson a dynamicair bearing table.
TABLEOF C 0 ” T S
1. INTRODUCTION
2. POTENTIAL MISSIONS
3. SUPPORTINGTECHNOLOGIES
4. DESCRIPTIONOF VEHICLES
5. DESCRIPTIONOFGROUND TESTCAPABILITIES
6. PROXIMITYAND DOCKINGEXPERIMENTS
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 demonstratinga successful autonomous docking of a
MicroSat with a target satellite. These capabilitieswill open
up a host of potential new missions that could revolutionize
robotic operations in space.However, before the capabilities
can be seriously consideredfor 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-gravityenvironmentthat 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
developmentof a new capabilityfor micro-satellite vehicles.
2. POTENTIALh.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
showsthe 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€erentviewpoints
of an inspection geometry requires precision Microsat
positioning, pointing, tracking, and imaging. A satellite
~~
’0-7803-6599-2/01/$10.0082001IEEE
*UpdatedDecember 15,2000
5-2493

2. 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
illustratesone 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 precisionhorning.
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 surfaceimage 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
systemcapability 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-arttechnologies 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
scenariois shown in Figure 3.
I MicroSat ManeuveringSequence within- .
a 10mradius 2
6
1
4
ProximityManeuvers:
1. Octahedronimaging 1-2-3456
2. Circular stereo imagingC63156-C64126
3. Soft docking 6-7-8-9
4. Formationflying 1.4,6,2
Figure 3 -A conceptual scenario for proximity
inspectionand soft dockingmission
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-
dockingswith the carrier vehicle.
3. SUPPORTINGTECHNOLOGIES
In describing the technologies involved in the ground test
setup for proximity and docking experiments, it should be
noted that two independent autonomousvehicles were used,
one as a target vehicle and the other as the docking vehicle.
These engineering test vehicles (ETV),designated as the
ETV-150and 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 asithe target for the ETV-250,the new docking vehicle.
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3. 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 sensorsmust supportthe variousmission 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.
Experimentsof the type described in this paper have focused
on proximity operations and docking, since the current
indoor setup restricts operationto distancesunder 10meters.
Sensor functionality within the context of these experiments
will be describedbelow.
Star Tracker
A
Focal
Length
9.9"
12.6mm
Monochrome Video
Camera
Color Stereo
I
Laser Ranger
Figure 4 - Multi-function IntegratedProximity
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 generationIPOS system
has been optimized for our close-in docking experiments,
and our long-range acquisition camera and long-range
LADAR system are not incorporatedintothis 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 availablesystem 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
ClementineI Lunar mapping mission [4-61. The M U is the
commerciallyavailable LN200 fromLitton.
Visible Imagers-With features as shown in Table 1below,
the Visible Inspection Camera is used to initiallyacquire 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 detectorused in this camera is capable of operating at a
500 Hz frame rate.
Table 1. VisibleCamera 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
camerainterfaces. 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 incorporatesa
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 oscillationkequency 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
repackagedfor LLNL to reduce mass and volume.
IMU-The LN200 is currently being used in the ground test
vehicles due to its performance and relativelylow cost. This
IMUhas a measured drift rate of approximately 1"hour and
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4. 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.
WFOVStar Tracker-As statedpreviously, 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 degreecoverageusing the existing lens design.
Avionics
Figure 5 -Avionics Architecture
The MicroSat avionics system described here can
accommodatea variety of sensor suites and can make use of
many types COTS U 0 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
CompactPC1bus. 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
communicationscapability. For a space mission, the specific
communications and telemetry capability required would
determine the onboard transceiver selected as well as the
type of ground stationedemployed 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 asthresholding, 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
5-2496

5. 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.
Specificchoices will await futureflight 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 organizationof our mission software.
Services
Figure 8 - Guidance and controlinterfaces
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-rowx
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 coordinateswas
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 experimentswill be
described in a later section.
4. DESCRIPTIONOF 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
elementthat fits around the central liquid manifold. For the
ETV-250, this is a complete sphericalball designed to allow
for larger tilt angles.
ETV-150
In the ETV-150 the 16thrusters 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 mainstructuralbackbone of the vehicle.
5-2497

6. 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 experimentsonly compressed nitrogen is used.
Figure 9 -ETV-150 inspectionvehicle
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
attachedto 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 commercialPowerPC 603e processor
board. The sensorson the ETV-150are 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
generationsensor 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
systemthe gas bottles could be replaced with liquid tanks to
provide larger delta-v maneuvers. The 16ACS thrusters on
the ETV-250 that provide independent fine control of both
translational position and angular orientationare 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 ,assemblyconsists 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
GrapplerDesign
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 fourgrappler
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 chosenbased on the known flexibility of a spherical
trailer hitch assembly. We desired to maximize the angular
range that the target ball could be approached. Previoushard
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
5 - 2 4 9 8

7. 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/pitchACS
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
100MHZ
m none
m none
LN-200
Imager
Software
Stereo-ranging
Srapplercontrol
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
166MHZ
0.03-30
0.5-5
LN-200
MPE
1024x1024
Yes
Yes
Yes
ves
5. DESCRIPTIONOF GROUNDTEST 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
environmentis 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 facilityis
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 facilitiesenable 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 sohare before launch.
Dynamic Air Bearing (DAB)
The concept of a dynamic air bearing (DAB) is to integrate
translationalmotions 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 5DOFmotion: 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 airpucks. 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 accelerationcapability.
Figure 11-An earlyprototype of a DynamicAir Bearing
DynamicAir Table
Three different generations of test vehicles are shown in
Figure 12. For 5DOF experiments conducted on the table,
each vehicle sits on an airbearing 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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8. 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 driftingtarget.
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 translationalmotion, such as in proximity
inspection maneuvers or in collision avoidance or intercept
maneuvers.
3
4
5
6
7
6. PROXIMITYAND DOCKINGEXPERIMENTS
These experiments were performed on the table described in
the previous section. The experiment sequence involved
severalbasic 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
acquireLED target
( 5.5 meters away)
Fly-in to 1.8meters
away (using stereo
ranger)
Align for docking
head-on (lateral
PerformFinal
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-omthe 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 initiallyreleased 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 stabletarget 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 reflectivebackboard, we
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9. obtained range validation measurements at 20 Hz. The
graph in Figure 17a showshow 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 anothercontributorto
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
directhead-on engagementat the 1.8 meter range. That was
done to demonstrate our ability to correct for such
misalignments.
The lateral alignmentsection of the experiment(step4) used
the offset of the LED center versus the center of the docking
ball to estimate the alignmentangle in each of the two VGA
stereoimagers,via the followingequation.
Theta= arcsin( 2 * pixel-offset / ball-diameter(pixe1s) )
The probe’s alignment angle was simply derived as the
averageof the left andright measured alignmentanglesfrom
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 secondsto 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 alignmentstep. The twoyawjets 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-strengthyaw control authoritybrought the LED back
to the center of the tracking camera’sfield 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
lateralthrustswere 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
16showstheprobe’sfinal approachto the target.
I . ..
.-
Figure 16-Probe is executingfinal dockingmaneuver
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. Duringthis 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 fkomthe stereocameras
5-2501

10. IC
e
- 6
Ev
Ww
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-LasedStereorange and difference
Algrmentanglemeasuredw rt doclangsurfaceperpendicular
10
Gv
i
o! i o 20 30 2 50 $0 70 80
Time (sec)
Figure 17c-Alignmentanglefor final approach
Time (sec)
Figure 1%- Vehicle roll rates
GoL ib 20 A. 40 Sb 60 i o ioTime (sec)
Figure 17d-Vehicle pitchrate
Time (sec)
Figure 17e- SeekerAZEL angle
Time (sec)
Figure 17f- Vehicle yaw rate
5-2502

11. 7. SuMh4ARY
To date, LLNL has successfully developed several
functionalvehicles and ground demonstratedseveral critical
on-orbit MicroSat logistics operations. They include:
autonomous target acquisition, tracking, stereo ranging,
homing on to a target, final approachtrajectory calculation,
soft docking, grappling onto a target, and adding a AV to a
target as in a de-orbitingmaneuver.
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-yeeLu at
LLNL for use of the stereo calibration softwarethat 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
Programat LLNL for their supportof 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 sponsoredby the U.S.
Governmentand performed by the University of California
Lawrence Livermore National Laboratory under Contract
W-7405-Eng-48 with the U.S. DepartmentofEnergy.
DISCLAIMER
This document was prepared as an account of work
sponsoredby an agency of the United States Government.
Neither the United States Government nor the University of
Californianor 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. Referenceherein to any specificcommercialproduct,
process, or serviceby 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
Governmentor the Universityof California,and shall not be
used for advertisingor product endorsementpurposes.
REFERENCE
[l] Ledebuhr,A.G., Ng, L.C., Kordas, J.F. et al., “Microsat
Rescue Demonstration Mission: A Feasibility Study”,
UCRL-ID-129880, Jan~ary30,1998.
[2] Ledebuhr, A.G., “Down-to-Earth Testing of
Microsatellites,”Lawrence LivermoreNational Lab Science
& TechnologyReview, September1998.
[3] Whitehead, J.C., “Hydrogen Peroxide Propulsion for
Small Satellites”, SSC98-Vm-1, The 12th Annual Utah
StateUniversitySmall SatelliteConference,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., ‘WiResCamera and
LIDAR Ranging System for the ClementineMission”, Proc.
SPIEVol. 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 vatents1 ,
inprojection displw te;hnologv. He has been at Lchrence
Livermore National Laboratory since 1986 and led the
development of advanced sensorsfor the Brilliant Pebbles
interceptor program and the design of the Clementine
sensorpayload. In 1996 he was the Clementine II Program
Leader and is currently the MicroSat Technologies
ProgramLeader.
Dr. Larry Ng received his B.S. and MS.
degrees in Aeronautics and Astronautics
from the Massachusetts Institute of
Technologyin 1973, and aPhD 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 HanscomAir
Force Base in Bedford, MA. His work experience includes:
four years with GeneralDynamics ElectricBoat Division in
Groton, CT, responsible for the development of the
TRIDENTsubmarine digital control systems; sevenyears 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 isfocusing
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
SigmaXi, Tau Beta Pi, and the National Research Council.
He haspublished numerouspapers in signal estimation and
precision vehicleguidance and control.
Mark Jones h m worked in theJield of
electronicssince 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 TechnologiesProgram.

12. 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
Clementineprojects, and currentlyleads
the software eflort for the IMicroSat
TechnologyProgram.
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-atmosphericKill Vehicle
(EKV. Upon retuming to LLNL in 1997 he assuimed the
lead GNC engineeringposition on theformer 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 (3DOFACS + 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.
5-2504