Autonomy is needed for future spacecraft to solve the problems of human operator overload and transmission delay. This paper describes the autonomous spacecraft executive for rendezvous and docking. It is an onboard expert system and has decision making capability for mission planning of nominal and contingency cases. The executive has been developed and verified using a hardware motion based simulator

Anderson, Robert L.

Furuya, Masatoshi

Komura, Fuminobu

Sasaki, Toshiro

Tsugawa, Roy K.

English

NASA Technical Reports Server

N95- 23719
Autonomous Spacecraft Executive and Its Application to
Rendezvous and Docking
Fuminobu Komura*, Masatoshi Furuya*, Toshiro Sasaki*
Robert L. Anderson**, Roy K. Tsugawa**
*Systems Development Laboratory, Hitachi, Ltd.
1099 Ohzenji, Asao-ku, Kawasaki 215 Japan
Phone:+81-44-966-9 ! 11, Fax:+81-44-966-6823
**Civil and International Systems Division,
TRW Space and Electronics Group
One Space Park, Redondo Beach, CA 90278 USA
Phone:(310)812-2608, Fax :(310)812-7296
KEY WORDS AND PHRASES
autonomy, mission planning, rendezvous and
docking
ABSTRACT
Autonomy is needed for future spacecraft to
solve the problems of human operator overload
and transmission delay. This paper describes the
autonomous spacecraft executive for rendezvous
and docking. It is an onboard expert system and
has decision making capability for mission
planning of nominal and contingency cases. The
executive has been developed and verified using
a hardware motion based simulator.
INTRODUCTION
Research activities have been done to develop
autonomous space systems.[ i ] Spacecraft
autonomy is needed to avoid the overload of
human operators and to overcome the delay or
loss of command link. Spacecraft rendezvous
and docking is a typical mission which needs
autonomous operations. [2] [3]
Spacecraft autonomy is attained by realizing
mission planning and contingency management
functions in onboard computers. The product of
mission planning or contingency management is
a sequence of commands to the conventional
control systems of the spacecraft.[3]
AUTONOMOUS SPACECRAFT
EXECUTIVE
PRECEDINg,; F -_.'.i _'_:.;,_:_ i'; T_ FILMED
235
Fig. 1 shows the architecture of an autonomous
spacecraft.[3] The Autonomous Spacecraft
Executive is an expert system implemented on an
onboard computer that makes decisions needed
for the spacecraft mission. The Executive is
interfaced to the GN&C (Guidance, Navigation
& Control system) and the SM (System
Manager), and receives state and status from the
GN&C and SM, and generates control
commands and sends them to the GN&C and
SM.
This architecture has the following
characteristics.
( 1) It is a universal modular architecture and is
applicable to any spacecraft.
(2) The modules that receive the control
commands don't need to know whether the
commands are sent from the Executive or from a
ground controller.
(3) The Executive has a vehicle dynamics
simulator as a mission planning tool.
EXECUTIVE FOR AUTONOMOUS
RENDEZVOUS AND DOCKING
Requirements
We consider rendezvous and docking missions
where the target vehicle is a cooperative passive
vehicle which is holding its attitude in a LVLH
(local vertical - local horizontal) frame and has a
receiver tbr differential GPS and reflectors on the
target for a docking sensor on the chaser vehicle.
The active chaser vehicle has the architecture of
Fig. 1.
To complete a rendezvous and docking mission
https://ntrs.nasa.gov/search.jsp?R=19950017299 2020-06-16T08:44:51+00:00Z
manydecisionsmustbemade.Themost
essentialdecisionis to plana flightpathor a
velocityprofile to attainthemissiongoalsunder
safety,timing andconsumablesconstraints.The
plansmustbemadefor bothnominaland
contingencysituations.Theyvarydependingon
the phasesof flight, i.e.,approachfrom a
parkingorbit,proximity, dock,separation,etc.
To accomplishtherendezvousanddocking
missionautonomouslytheExecutiveisrequired
to createtheseflight plans.[3][4]
For thefinal stageof proximitye.g.from 1000
fi to 0 fi, therequirementsfor theExecutivewill
beasfollows.
modes:
- nominalapproachplan
- contingency:lossof GPSlockor lossof
proximitysensorlock
- replanandtry again,or
- abortthemission
constraints:
- safevelocityprofile
- safeapproachcorridor
- timeof arrival (for lightingcontrol,crew
schedule,communicationsavailability,etc.)
Executive Functions
The Executive has the following functions to
meet the above requirements.
( 1) input
- mission goals from the ground controller
- spacecraft state and status from the GN&C
and SM
(2) monitor
- status of sensors
- position and velocity of chaser relative to
target:
determine whether within control volume and
safety limits, and if mission requirements are
attainable
(3) plan
Depending on the output of ( 1) and (2), either
of the following plans is generated from the
rules.
- nominal approach based on the time of arrival
requirements
- contingency plan based on the spacecraft state
and status
- abort
(4) output
- control commands to the GN&C and SM
Monitoring and Planning Rules
The Executive functions of monitoring and
planning can be realized by a set of decision rules
which are expressed in the following form.
IF
(current_control_state)(relative_position)
(vehicle_status)(mission_requirements)
THEN
(create new plan
or continue
or create contingency plan
or station keep
or back away
or abort)
The IF part represents the monitoring, and the
THEN part represents the planning. By these
rules the control state of the vehicle is
determined. Fig.2 shows a state transition
diagram for the proximity operation.
The generation of the nominal plan "create new
plan" consists of the following processes.
1. Design velocity profile for each phase
The proximity operations consist of a number of
phases separated by station keeping positions.
For example, station keeping positions are set at
-1000ft, -300 ft, -35 fl, and -20 ft. They are
needed for changing the vehicle control modes
and for adjusting the arrival time at the target. A
transfer is usually used from -1000 ft to -300 fi
to save fuel, and an LVLH approach is preferable
within -300 fi for safety. The velocity profile is
computed by using mission planning tools, e.g. a
vehicle dynamics simulator.
2. Select the earliest possible docking window
3. Allocate duration for each station keeping
position
4. Abort if no window is attainable
With these rules the Executive can make
decisions needed for the nominal and
contingency operations in the proximity stage.
Other set of rules are used for the autonomous
operations in other stages.
VERIFICATION TESTS USING A
HARDWARE SIMULATOR
236
Simulator Configuration
The configuration of the verification test facility
is shown in Fig. 3. The Executive was
implemented on a PC and it was connected via an
RS422 link to the 6 DOF (Degree of Freedom)
dynamics simulator and the GN&C system
installed on a VAX at the NASA Marshall Space
Flight Center astrionics laboratory. The mockup
of a chaser vehicle with the actual VGS (Video
Guidance Sensor) was mounted on the floor and
the VGS was connected to the 6 DOF simulator.
The GPS was simulated in the 6 DOF simulator.
The DOTS (Dynamic Overhead Target Simulator)
on a VAX moves a crane arm based on the output
state of the 6 DOF simulator. The mockup of a
target vehicle is attached on the arm end. The
reflectors for the VGS are attached to the back
face of the target vehicle.
With this Configuration the motion of the target
vehicle relative to the chaser vehicle can be
simulated. The range of simulated flight covers
the final approach from 50 ft to 0 ft station
keeping position where the three point docking
mechanism can be activated to complete the
docking.
In addition to the simulations using the above
setup, the software simulations were done using
only the Executive on the PC and the VAX
simulator. The range of flight in these software
simulations are from 1000 fl to 0 ft.
Test Results
Test runs of the chaser approach were made
both in hardware simulations and software
simulations by changing the initial conditions and
the docking windows. The contingencies were
brought about by either physically disabling the
VGS hardware or simulating the loss of GPS
lock at an arbitrary time during approach. In all
of the cases it was verified that the Executive can
start the mission replanning and generate a new
approach or abort profile based on ground
supplied mission rules.
Fig. 4 shows an example test result of a case
where VGS lock was lost and regained during
the final approach. While station keeping at x = -
35 ft, the Executive generated a flight plan,
PLAN 1, for the nominal approach. The plan
drives the chaser first to the next station keeping
point at x = -20 ft, and the vehicle stays for the
period needed to check the vehicle status, and the
vehicle resumes the approach to x = 0 ft to meet
the docking window #2. But during the approach
the VGS lock was lost at t = 190 sec. When the
Executive detected the loss it generated the
contingency plan, PLAN2. The plan forces the
vehicle to back up to the safe station keeping
position at x = -35 ft, and let it wait until sensor
lock is regained. Because the lock is regained
during this back up, the Executive generated a
new plan, PLAN3, similar to PLAN 1, to resume
a nominal approach, but this time the earliest
window available is window #3. Tables 1. and
2. show the control commands for PLAN 1 and
PLAN2.
CONCLUSIONS
The autonomous spacecraft executive has been
developed for autonomous on board mission
planning for rendezvous and docking. Its
decision making capability for nominal and
contingency cases has been verified by
simulations.
The executive is also applicable to other
spacecraft missions which need autonomous
onboard decision making.
REFERENCES
[1] M.L. Montemerlo, The AI Programs at the
National Aeronautics & Space Administration,
AI Magazine, Winter 1992.
[2] H. Nakamura, et al., Study of Autonomous
Rendezvous and Docking, J. Space Technology
and Science, vol.4, No. 2, 1989.
[3] R.L. Anderson, et al., Demonstration of
Automated Proximity and Docking Technology,
In Proceedings of AIAA Space Programs and
Technologies Conference, March 1992,
Huntsville, AL.
[4] M. Furuya, et al., A Study on High Level
Decision Making Function in Onboard Computer
System for Rendezvous and Docking, In
Proceedings of the 37th Space Sciences and
Technology Conference, October 1993,
Kita-Kyushu, Japan.
237
AUTONOMO_
.,_o ,Nn'u, uzA no_ °N_"_ S CECRAFT _... G,SYSTEM_GUR 'nON PA N&C MOOE $ELEC_ON
_AI S SPACECRAFT SENSOR
...... .o_\ t sYstems / /
OF.AC'ITVA'nON,ANO EFFECTOR
Fig. l Architecture of Autonomous Spacecraft
GENCY ]
Fig. 2 State Transition Diagram for Proximity
DOTS (on VAX)
Chaser
i _6DnO:rnS_Ma(nO;GVNA_C_-[ F-'XEO(°n PC)
x (It)
-40 -
35,
-30.
-20
-I0
...9
EXAMPLE
SENSOR LOSS AND MISSION REPLAN
T - fgO T = 251
VGS VGS
LOSS REGAIN
WINDOW (_)
Fig. 3 Configuration of Verification Test Facility Fig. 4 Sample Test Result
Table 1. Control Commands for PLAN 1 Table 2. Control Commands for PLAN2
T(sec) X(ft) EVENT
0
0
0
97.3
263.0
273.0
273.0
313.0
919.1
-35 SET LVLH FRAME
-35 SET TARGET POINTING
-35 START STATION KEEPING
-35 START APPROACH
-20 START STATION KEEPING
-20 START TARGET BODY FRAME
-20 START ATTITUDE HOLD
-20 START APPROACH
0 START STATION KEEPING
T(sec) X(ft) EVENT
191.0 -31.8
191.6 -31.8
191.6 -31.8
199.1 -31.8
240.8 -35
START STATION KEEPING
SET LVLH FRAME
SET TARGET POINTING
START SEPARATION
START STATION KEEPING
238


Autonomous spacecraft executive and its application to rendezvous and docking

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