View
214
Download
0
Category
Preview:
Citation preview
Jet Propulsion LaboratoryCalifornia Institute of Technology
Meeting No. 96Aerospace Control and Guidance Systems
Committee
Micro-spacecraft GN&C
Greg Mungas Mission and Systems Architecture Group
Jet Propulsion Laboratory
Oct. 21, 2005
Work Sponsored through Mars Program Office’s Eureka Team and Next Generation Orbiters (MAX – Mars Aeronomy Explorer, Robert Shotwell, Andrew Gray), the Laboratory of Atmospheric and Space Physics’ Inner
Magnetospheric Explorer (IMEX)
2005 Aerospace Control and Guidance Systems Committee
Overview
• Motivation and requirements for micro-spacecraft performing science
• IMEX mission case study– Detumbling and Sun Acquisition system
• MAX mission case study– Low cost spacecraft constellation at Mars
• Conclusions and Wrapup
2005 Aerospace Control and Guidance Systems Committee
Micro Science Spacecraft
• Spacecraft need to address science mission objectives to be able to compete in NASA AO process $350M-$750M– Mission designs are often iteratively designed from
measurements and instrument payload up. • Most measurement approaches have some fundamental
limitations in scalability that limit sizing down to and selection of micro-spacecraft
– Bouy and Constellation missions are one exception• Relatively simple instrumentation with requirement for:
– Low-cost/long life measurements (i.e. aeronomy/space weather)– Simultaneous multi-spatial measurements (i.e. space weather)
• Navigation infrastructure for feed-forward missions (Mars/Lunar NAV/COM)
– Need to address developing low cost multi-spacecraft to fit within program budgets
• Minimizing Processors and Software (i.e. Qualification of Mars Science Lab FFT algorithm ~$300K out of ~$10M instrument budget)
• Solutions typically appear to involve use of additional “non-instantaneous”, non-sensor-based information available to user
2005 Aerospace Control and Guidance Systems Committee
• Inner Magnetospheric Explorer to survive in and study earth’s radiation belts over an extended duration of time.
• Proposed as a SMEX ~$15M hitchhiker spacecraft on last Titan IUS carrying ~$1B military spy satellite
• Extreme radiation environment in radiation belts limited use of conventional processors (i.e. state-machine-based and open loop ground control)
• “Free” hitchhiker ride provided no favors– No active processor– Unfavorable thermal environment (no heat for
10 hour period on spent rocket stage)– Random deployment tumble of up to 5rpm– Upper stage and RCS extreme plume
contamination• i.e. mounted beneath an RCS thruster module
IMEX Mission Summary
2005 Aerospace Control and Guidance Systems Committee
• Needed to deal with early deployment – Large angle (non-linear) spinup, detumbling, and sun acquisition maneuvers
• Developed simple state-machine based controller, Spinning Smallsat Detumbling and Sun Acquisition System [G. Mungas, D. Lawrence AIAA 2000-4143]
• Subsequently discovered other’s with similar design problem (i.e. Carnegie Mellon’s Helio-gyro experiment, MAX mission, etc.)
• Managed to develop concept for single module that can perform deployment operation as a small “slap-on” module.
IMEX GNC Design Theory
InertialSensor
Sun PresenceDiode
SpinupThruster
PrecessionThruster
SunCrossingDiode
y
x
z
InertialSensor
Sun PresenceDiode
SpinupThruster
PrecessionThruster
SunCrossingDiode
y
x
z
2005 Aerospace Control and Guidance Systems Committee
Spinning S/C Stability Theory
• Initial detumbling stabilization based on Lyapunov-like stability theory (not necessarily to single attractor) in one of two regions on body-mounted manifold 2222 HHHH zyx
MOMENTUM SPHERE2222 HHHH zyx
MOMENTUM SPHERE
ENERGY ELLIPSOID
TI
H
I
H
I
H
zz
z
yy
y
xx
x 2222
ENERGY ELLIPSOID
TI
H
I
H
I
H
zz
z
yy
y
xx
x 2222
2005 Aerospace Control and Guidance Systems Committee
Spinup Criteria
+ Z
2
Separatrix
Separatrix
- Z
+ X
- X
mins
0
• Spinup stabilization criteria and sizing of actuators based only on knowledge of spacecraft moment’s of inertia, worst case deployment tumble rate, and effective actuator torque
zzyyzz
xxxxyy
III
III
)(
)(tan
)costan(sin0min s
L
IF
spinup
tspinup
2
max )(
(2) Nutating Spin
z
x
zz
xx
y
y
th
zh
22
22221
2zzttzz IIIhh
2112 zzIT 2
22
22 zztt IIT
22
212 zztt
z
t III
IT
th zt
ztz
I
II
)(
th
th
y
x
h
h
2005 Aerospace Control and Guidance Systems Committee
Effective Torquing Efficiency
• Even with finite thrusting times over large spin angles, , effective applied torquing efficiency, torque remains high
xx
actuator
effective
y
x
2sin
2
torque (4)
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 30 60 90 120 150 180
Torquing Angle ( in deg)
torq
ue
2005 Aerospace Control and Guidance Systems Committee
Denutation
• Simple denutation law applied in body frame.
• Implemented with low res. rate gyro or accelerometer.
th
y
y
x , Torque Axis, Precession Direction
DE
NU
TA
TIO
N
ZO
NE
Torque axis
2005 Aerospace Control and Guidance Systems Committee
Sun Acquisition I• Developed sensor/actuator candidate bang-bang control
hardware solution that should always close distance between inertial angular momentum vector, , and local sun vector, .
• Sun-seaking “Hardware” solution is based only on hemispherical slit photo-diode collocated with actuated transverse torquing axis
H
Optical Slit
x
y
zH
S
H
x
z
S
Inertial Frame Spacecraft Frame
S
2005 Aerospace Control and Guidance Systems Committee
Sun Acquisition II
• Define “Set” of normalized S/C properties (mass and thruster torque) that have sun-tracking solutions with acceptable tolerances on max nutation angle
;;xx
zz
xx
yy
I
I
I
I ;
22
ii
i
s
i I
;2
; ss
ii
tt
Normalized Parameters
Normalized Dynamics And Kinematics
zyxz
yxzy
xzyx
12
12
2
qtd
qd
zyx
zxy
yxz
xyz
~
0
0
0
0~
2005 Aerospace Control and Guidance Systems Committee
Sun Acquisition III
XX
ZZ
I
I
XX
YY
I
I
Major Axis Spinners
• 180° Acquisition
• Thruster On-Time = 1/12 s
• 5° Thruster steps
• 0° Initial nutation
Simulation Parameters
Spherically Symmetric
2005 Aerospace Control and Guidance Systems Committee
Sun Acquisition IV
Minor Axis Spinners
XX
ZZ
I
I
XX
YY
I
I
• 180° Acquisition
• Thruster On-Time = 1/12 s
• 5° Thruster steps
• 0° Initial nutation
Simulation Parameters
Spherically Symmetric
Not Realistic Mass Distributions
2005 Aerospace Control and Guidance Systems Committee
Detumbling/Sun Acquisition Summary
• Developed a simple hardware/control logic solution to a traditionally complex, non-linear spacecraft deployment problem. Solution (which can be packaged as add-on module to experiment) consists of:– Miniature shielded Photo-cell – sun acquisition– Spin-up and Precession Thruster + Fuel– Denutation sensor (MEMS rate gyro or accelerometer)– Simple state-machine + If/then control logic
• No requirement for S/C processor – Significant mission cost savings – flight qualifying software– Suitable for probes in high radiation environments – no
processor requirements. Measurements can be state-machine-based
2005 Aerospace Control and Guidance Systems Committee
• Develop low cost mission architecture for performing multi-node network Mars Aeronomy Explorer mission.
• Simultaneous spatial observations of electric/magnetic fields throughout magnetosphere and low altitude ion/neutrals composition
• Residual NAV/COM network• Package entire mission into a ~$500M cost cap
Group 1b : 4 Magnetometer Spacecraft, 100 x 10,000 km
Group 2 : 2 identical telecommunications spacecraft, 150 x 1,000 km
Group 3 : 2 identical spinner Spacecraft, 200 x 10,000 km
Group 1a : 2 Magnetometer Spacecraft, 250 x 30,000 km
MAX Mission Overview
SPINNERS in Highly Elliptic Orbits
SPINNERS in Highly Elliptic Orbits
2005 Aerospace Control and Guidance Systems Committee
MAX Constellation Spacecraft
SPINNERS in Highly Elliptic Orbits
2005 Aerospace Control and Guidance Systems Committee
MAX Spinners
• Adapt open loop ADCS architecture developed for the Laboratory of Atmospheric and Space Physics (LASP) SNOE spacecraft (~$15M mission including Pegasus launch) to the Mars environment.– SNOE (in sun synchronous orbit) was open loop
magnetorquer precessed ~1/day to track apparent solar motion.
– For open loop control, MAX spinner requires tracking RAAN precession of orbit normal vector within 1 with propulsive torque (no significant Mars magnetic field)
torqueth
h
2005 Aerospace Control and Guidance Systems Committee
MAX Orbital Precession• Maximum RAAN precession rate of <1.5/day
– Consistent with SNOE implementation approach
nRAAN Precession vs. Orbital Inclination for
Decaying 200x10,000 km Orbit
0.0
0.2
0.4
0.6
0.8
1.0
1.2
1.4
1.6
0 10 20 30 40 50 60 70 80 90
Orbital Inclination (deg)
Orb
ita
l Pre
ce
ss
ion
Ra
te
(de
g/d
ay
)
10,000 km
8,000 km
6,000 km
Apogee Altitude
mZ
n
2005 Aerospace Control and Guidance Systems Committee
Active Denutation
• Body frame denutation coupled control law– Restricts Precession torques to
a <50% duty cycle - ~25% for efficient fuel usage (see next slide)
• For MAX Spinner, applied torque momentum capability: – Vacuum arc thruster:
• Precession: >55/day• Min Angular Bit: 0.00002
– Milli-Newton N2H2 thruster:• Precession: >>100/day • Min Angular Bit: 0.0002
• Given <1.5/day Precession Requirement, Active Denutation Control is NOT a Significant Implementation Constraint
th
y
y
x , Torque Axis, Precession Direction D
EN
UT
AT
ION
Z
ON
E
2005 Aerospace Control and Guidance Systems Committee
Aerodynamic Drag Induced Nutation
Zz ˆ,ˆ
Yaeroˆ,
x
y
X
z
INERTIAL FRAME ( ZYX ˆ,ˆ,ˆ ) BODY FRAME ( zyx ˆ,ˆ,ˆ )
Zz ˆ,ˆ
Yaeroˆ,
x
y
z
0
cos)(
sin)(
z
zaerozxzt
tzy
zaerozyzt
ztx
h
thhII
IIh
thhII
IIh
)ˆsinˆcos( ytxtI
Ih zz
zz
aerott
22maxtan
zz
aerot
z
t
I
I
h
h
11
z
aero
z I
Figure 5. Calculating Nutation Excited by Constant Inertial Torque (close approximation of ~inertially fixed aerodynamic torque relative to spin rate)
Zz ˆ,ˆ
Yaeroˆ,
x
y
X
z
INERTIAL FRAME ( ZYX ˆ,ˆ,ˆ ) BODY FRAME ( zyx ˆ,ˆ,ˆ )
Zz ˆ,ˆ
Yaeroˆ,
x
y
z
0
cos)(
sin)(
z
zaerozxzt
tzy
zaerozyzt
ztx
h
thhII
IIh
thhII
IIh
)ˆsinˆcos( ytxtI
Ih zz
zz
aerott
22maxtan
zz
aerot
z
t
I
I
h
h
11
z
aero
z I
Figure 5. Calculating Nutation Excited by Constant Inertial Torque (close approximation of ~inertially fixed aerodynamic torque relative to spin rate)
Zz ˆ,ˆ
Yaeroˆ,
x
y
X
z
ZYX ˆ,ˆ,ˆ zyx ˆ,ˆ,ˆ
Zz ˆ,ˆ
Yaeroˆ,
x
yz
NUTATION ANGLE,
0
cos)(
sin)(
z
zaerozxzt
tzy
zaerozyzt
ztx
h
thhII
IIh
thhII
IIh
Solution is )ˆsinˆcos( ytxtI
Ih zz
zz
aerott
22maxtan
zz
aerot
z
t
I
I
h
h
Valid for spin rates >> precession rates 11
z
aero
z I
Figure 5. Calculating Nutation Excited by Constant Inertial Torque (close approximation of ~inertially fixed aerodynamic torque relative to spin rate)
Zz ˆ,ˆ
Yaeroˆ,
x
y
X
z
ZYX ˆ,ˆ,ˆ zyx ˆ,ˆ,ˆ
Zz ˆ,ˆ
Yaeroˆ,
x
yz
NUTATION ANGLE,
0
cos)(
sin)(
z
zaerozxzt
tzy
zaerozyzt
ztx
h
thhII
IIh
thhII
IIh
Solution is )ˆsinˆcos( ytxtI
Ih zz
zz
aerott
22maxtan
zz
aerot
z
t
I
I
h
h
Valid for spin rates >> precession rates 11
z
aero
z I
Figure 5. Calculating Nutation Excited by Constant Inertial Torque (close approximation of ~inertially fixed aerodynamic torque relative to spin rate)
2005 Aerospace Control and Guidance Systems Committee
Aerodynamic Induced Disturbances
• Assuming 6 rpm, 10 cm Cp/Cg axial offset, 60 minute flight through worst case periapse (200x10,000km orbit), with MAX’s worst case aerodynamic profile– <0.007 Precession of – <0.000003 Excited Nutation
h
2005 Aerospace Control and Guidance Systems Committee
Attitude Determination with HCI’s
HCI Pointing
Roll Angle
HCI Mounting Angle
200 km altitude 1000 km altitude
Projected Horizon Plane Projected Horizon
122.36
126.84
47.46 56.56
Projected Horizon Plane
Figure 7. HCI Spin Angle Measurements for 1 Roll Angle Pointing Error as a Function of Altitude for a 45 HCI Mounting Angle.
HCI Pointing
Roll Angle
HCI Mounting Angle
200 km altitude 1000 km altitude
Projected Horizon Plane Projected Horizon
122.36
126.84
47.46 56.56
Projected Horizon Plane
HCI Pointing
Roll Angle
HCI Mounting Angle
HCI Pointing
Roll Angle
HCI Mounting Angle
200 km altitude 1000 km altitude
Projected Horizon Plane Projected Horizon
122.36
126.84
47.46 56.56
Projected Horizon Plane
200 km altitude 1000 km altitude
Projected Horizon Plane Projected Horizon
122.36
126.84
47.46 56.56
Projected Horizon Plane
Figure 7. HCI Spin Angle Measurements for 1 Roll Angle Pointing Error as a Function of Altitude for a 45 HCI Mounting Angle.
HCI Pointing
Roll Angle
HCI Mounting Angle
-2000 -1000 0 1000 2000
-2000
-1500
-1000
-500
0
500
1000
1500
2000
200 km altitude 1000 km altitude
Projected Horizon Plane Projected Horizon
-2000 -1000 0 1000 2000
-2000
-1500
-1000
-500
0
500
1000
1500
2000
122.36
126.84
47.46 56.56
Projected Horizon Plane
Figure 7. HCI Spin Angle Measurements for 1 Roll Angle Pointing Error as a Function of Altitude for a 45 HCI Mounting Angle.
HCI Pointing
Roll Angle
HCI Mounting Angle
-2000 -1000 0 1000 2000
-2000
-1500
-1000
-500
0
500
1000
1500
2000
200 km altitude 1000 km altitude
Projected Horizon Plane Projected Horizon
-2000 -1000 0 1000 2000
-2000
-1500
-1000
-500
0
500
1000
1500
2000
122.36
126.84
47.46 56.56
Projected Horizon Plane
HCI Pointing
Roll Angle
HCI Mounting Angle
HCI Pointing
Roll Angle
HCI Mounting Angle
-2000 -1000 0 1000 2000
-2000
-1500
-1000
-500
0
500
1000
1500
2000
200 km altitude 1000 km altitude
Projected Horizon Plane Projected Horizon
-2000 -1000 0 1000 2000
-2000
-1500
-1000
-500
0
500
1000
1500
2000
122.36
126.84
47.46 56.56
Projected Horizon Plane
-2000 -1000 0 1000 2000
-2000
-1500
-1000
-500
0
500
1000
1500
2000
200 km altitude 1000 km altitude
Projected Horizon Plane Projected Horizon
-2000 -1000 0 1000 2000
-2000
-1500
-1000
-500
0
500
1000
1500
2000
122.36
126.84
47.46 56.56
Projected Horizon Plane
Figure 7. HCI Spin Angle Measurements for 1 Roll Angle Pointing Error as a Function of Altitude for a 45 HCI Mounting Angle.
2005 Aerospace Control and Guidance Systems Committee
HCI AD Sensitivity
• Over entire altitude range, HCI’s provide better than 0.1 degree instantaneous roll angle knowledge for a perfectly spherical Mars.
2005 Aerospace Control and Guidance Systems Committee
Filtering Body-fixed Measurement Errors
• Given extremely low disturbance environment, observed motion during a single pass is effectively due to body-fixed mounting errors.
h
2005 Aerospace Control and Guidance Systems Committee
Filtering w/ Elliptic Orbit• Similar geometric interpretation as circular orbit with low altitude arc to
remove estimation errors associated with body-fixed sensor errors
Orbital
Geometry
errh
h
errh
h
Sensor Design Constraint
Attitude Estimation
Arc (a)
Arc Traverse
Time
1deg/day RAAN
Precession over Arc
Peak Altitude
Nadir to Horizon Angle
(a)
Max Variation in a over Arc
(deg) (min) (deg) (km) (deg) (deg)0.0 0.0 0 0.000 200.0 70.8 0.0
10.0 2.4 15 0.002 205.0 70.6 0.2
20.0 4.8 29 0.003 220.1 69.9 0.9
30.0 7.3 44 0.005 245.4 68.8 2.040.0 9.8 59 0.007 281.1 67.5 3.450.0 12.4 74 0.009 327.6 65.8 5.060.0 15.0 90 0.010 385.4 63.9 6.970.0 17.7 106 0.012 454.8 61.9 8.980.0 20.5 123 0.014 536.6 59.7 11.190.0 23.5 141 0.016 631.6 57.5 13.3
100.0 26.6 160 0.018 740.6 55.2 15.6110.0 29.9 179 0.021 864.7 52.9 17.9120.0 33.4 200 0.023 1005.1 50.5 20.3
130.0 37.1 223 0.026 1163.0 48.2 22.6
140.0 41.2 247 0.029 1340.1 45.8 25.0
Number of Attitude
Measurements at 6rpm
2005 Aerospace Control and Guidance Systems Committee
Implementation Plan
torqueth
h
RAAN Precession
Pointing Requirement
Daily Permissible Tracking Error
Angular Momentum Vector Maneuver Starting Point
• ~Daily download of HCI vector• Filtered State Update• Upload of Daily Thruster Pulse Train
h
2005 Aerospace Control and Guidance Systems Committee
Hardware/Propellant Summary
• 2 x = ~10’s g.
• Spinup + precession thruster (10’s g)
• Nutation sensor (accelerometer or rate gyro) <10g.
• 0.5-1kg worst case propellant for 2 year mission + storage and plumbing
• Wiring to C&DH
2005 Aerospace Control and Guidance Systems Committee
Conclusions
• Proposed stable alternative open loop control architecture for low altitude velocity-vector-aligned mapping to within 1 for a low cost/low mass network spacecraft mission. Primary control error is orbital RAAN precession.
• HCI Attitude error estimates down near ~0.1 appear possible by triangulating in on h vector at different orbital locations
• Pointing control is very robust with >10-100 fold capability in daily momentum management and <1/1000 of minimum pointing error with thruster minimum impulse bits
2005 Aerospace Control and Guidance Systems Committee
Questions??
Recommended