1. Introduction
ISAS/JAXA is currently studying on the solar sail propulsion for future applications to deep space explorations. The solar sail is a means of propulsion utilizing the momentum of photons from the sun. In September 23, 2006, the small solar sail demonstration satellite (SSSAT) was launched by the 7th M-V rocket as the sub-payload. The main purpose of this satellite was electric power supply by the thin film solar cell. For the next step of the solar sail project, we are planning inter planetary space mission to demonstrate and evaluate the photon propulsion under the air drag free condition.
This demonstration spacecraft is launched by H2A rocket as a piggyback of the main satellite whose orbit is geostationary transfer orbit (GTO). In order to leave from the orbit to interplanetary space, this spacecraft has the deorbit motor. For the orbital transformation, this mission has three staged sequence: (1) separation from H2A and injection into GTO, (2) delta-V by the deorbit motor, and (3)
deployment of the membrane and verification of the photon propulsion. The configuration of the spacecraft varies with stage, so the attitude control strategy needs to be designed for each stage.(See Fig.1) Flat Spin‡™VSeparation of theDeorbit MotorH2A rocketSeparationDeployment of The MembraneSun AcquisitionManeuver GTOSun AcquisitionManeuver Maneuver Before ‡™VRollSpinFlat VRollSpin
Fig.1: Mission Sequence
In this paper, the attitude control system for this spacecraft is designed.
First, the system alignment of the spacecraft and the attitude control device are introduced.
Next, we design the attitude control strategy for each stage. The attitude control system is consists of the spin rate control, active nutation control, and spin axis reorientation control. The mission sequence is designed using these control methods.
Finally, results of the analysis on the attitude control are shown. The spacecraft is spin stabilized type. To make the attitude stable, proper spin rate is estimated for each stage of the mission. The amount of the propellant consumed through the sequence is also estimated. The attitude control sequence must be designed within the limits of the amount of the propellant. Under this restriction, the feasibility of this mission is discussed.
2. System Configuration
2-1. Configuration
H2A-piggy solar sail spacecraft consists of the probe, the deorbit motor, and others such as ignition timing controller. (See Fig.2) ProbeDeorbitMotorIgnition ControllerMount Surface to RocketSeparationPlaneProbeDeorbitMotorIgnition RocketSeparationPlane
Fig.2: System Configuration
The probe is the modification of SSSAT. This has cylindrical body of 350mm diameter and 150mm height, and its weight is restricted under 10kg. The reaction control system and attitude sensors are mounted on the probe. The solar sail membrane is wound around the side surface of the probe, and deployed by the centrifugal force of the spinning probe.
The deorbit motor is originally developed for the penetrator of Lunar-A. After the injection of the probe into the interplanetary space, the motor and other parts are separated from the probe.
Fig.3 shows the coordinate axes used through this paper. The origin is set to the center of mass.
Before the SeparationAfter the SeparationX (Roll)Y (Pitch)Z (Yaw)X (Roll)Y (Pitch)Z (Yaw)
Fig.3 : The Coordinate Axes
Table.1 shows the estimation of the moment of inertia of the spacecraft. Before the separation, the spacecraft has the maximum principal axis of inertia on z-axis. After the separation, the probe has the maximum principal axis of inertia on x-axis.
Table1 : Moment of Inertia
Before the separation
[kgEmm2]
After the separation
[kgEmm2]
Ixx
376659
124107
Iyy
2058304
78386
Izz
2103358
78386
Ixy
3336
10
Iyx
-1961
10
Ixz
2756
10
2-2. Reaction Control System
The probe has the gas-liquid equilibrium thruster, which uses LPG as propellant. This system can bring the propellant as liquid in the tank and eject the propellant as gas using the gas liquid equilibrium pressure to produces the thrust. This thruster system is simple and light weight [1].
For this spacecraft, gHFC-134ah is used as propellant. The thruster force is about 0.8N, and Isp is about 30sec.
This spacecraft has four thrusters for the reaction control system.
2-3. Attitude Sensors
For the attitude determination of the spacecraft, sun sensors and gyro sensor are used. In addition, we use the attitude determination method using RF-doppler measurements. In the method, the attitude of the spacecraft is determined by analyzing Doppler-frequency of RF signal from the probe received at ground station [2].
To monitor the sun direction continuously during spin period, we use two-dimensional position sensitive detector (2D-PSD, sun sensor). Two 2D-PSDs are mounted on the spacecraft. The sun direction is monitored for the spin around both X-axis and Z-axis. Fig.3 shows the alignment of 2D-PSD and thrusters.
Gyro sensor is mounted on the electrical circuit of CPU, and measures the spin rate around three axes.
2D-PSDXYZThruster2D-PSDXYZThruster
Fig.3: The alignment of thrusters and sensors
3. Attitude Control Strategy
For the spin solar sail satellite, we use mainly three attitude control logic as follow:
? Spin Rate Control
? Active Nutation Control (ANC)
? Spin Axis Reorientation
These control laws are effectively utilized for each stage.
3-1. Separation and Maneuver
The spacecraft is injected to GTO by H2A rocket, and then goes around that orbit several times. While the orbiting, z-axis of the spacecraft must be kept to face the sun so as to generate sufficient electricity. To keep such attitude stably, the spacecraft is put into a flat spin around z-axis.
In this mission, the following sequence is taken to make the spacecraft flat spin stabilization.
? S rocket espin ar
eparation from H2A
? D ound x-axis
pin up around z-axi
ind the sun directi
the
eorientation of the
To find the sun direction
one is direction. To
o
After s leaves
the orbi motor.
B
x-axis up to 0.1Hz
ough reorientation e
rbit direction
? S nd x-axis up to 1.0Hz
recise attitude m
elta-V
? S f the deorbit motor
S eorientation o n rate o
need
co is
performed in two stages a
To re e spin axis to
w
o stages again. The following
sequence ent of
membran
ent of mem
While th ent of m
is perfor e
of th
4. Mission A
4-1. Proper Spin Rate during the Flat Spin
? Ss up to 0.1Hz(flat spin)
? Fon using 2D-PSD through flat spin
? R spin axis towards the sun direction.
easily during a spin period, 2D-PSD is mounted on x-ax
rient the spin axis towards the sun easily, another 2D-PSD mounted on z-axis direction. 3-2. Delta-V by the Deorbit Motor
everal periods on GTO, the spacecraftt to interplanetary space by the deorbit
efore the delta-V, precise attitude maneuver must be performed. This maneuver is performed in following sequence. ? Despin of the flat spin ? Spin up around
? R of the spin axis to th
deopin up arou
? Paneuver to the deorbit direction
? Deparation o
pin axis rn high spif 1.0Hz s much propellant. To reduce the large ption of the propreorientation
nsumellant, the
s above. orient th the deorbit direction,
e need another attitude determination method which is different from using sun sensor. For this reason, the attitude determination method using RF Doppler measurements is introduced, though the d
etail of the method is beyond the scope of this paper. The deorbit motor keeps the thrust about 20sec for the injection of the probe to interplanetary space, and then the spacecraft transfer to ANC mode. After the nutation is removed, the probe is separated from the deorbit motor.
3-3. Deployment of Membrane The probe reorients the spin axis to the sun before the deployment of the membrane. The reorientation is performed in tw
is taken before the deployme.
? Spin around x-axis down to 0.1Hz ? Rough reorientation of the spin axis towards the sun ? Spin around x-axis up to 1.0Hz ? ANC eploym
? Dbrane e deploymembrane, support control med in the sam
way of SSSAT. The detail e control strategy is covered in [3].
nalysis
target spin rate o ft must
sp s of the stability of the attitude.
Here we estimate gravity-gradient torque,
aerodynamic torq diation torque
based on the pacecraft. The
su lat
sp is
s
The f the spacecrabe ecified in term
The target rate of the flat spin on first stage is specified to enable the spin axis towards the sun under the disturbance torque.
ue, and solar raconfiguration of the s
mmation of disturbance torque act to the finning spacecraft while a period of GTO
hown in Fig.4.
perigeeapogeeperigeeapogee
Fig.4 : Disturance torque while a period of GTO
In the GTO, the major disturbance torque is
gravity-gradient torque except near the perigee,
where the major disturbance torque is aerodynamic
torque. The maximum torque is about 2.6e-4Nm. The
integration of the torque for the orbital period
becomes as follow:
?
? ?
= ?
Nx = 6.8e [Nm ? s]
-3
-6
elta-V
The target ra cond stage
is specified p the target
direction duri disturbance
torq
s generated about an
axis nce
torq r, it
i
l thrust force.
?s][Nm1.9eNzThis integrated torque is enough small to keep the direction of the flat spin of 0.1Hz. It follow that this spin rate is appropriate for the stable spin to keep z-axis face to the sun direction. 4-2. Proper Spi
??=s][Nm2.1eNy2-
n Rate before the Dte of the roll spin on the seto enable the spacecraft keeng the delta-V against the
ue of the deorbit motorf s thrust. Fig.5 shows the prospective time history of the thrust force of the deorbit motor. If there is a misalignment between the thrust force vector and x-axis, large torque i
orthogonal to the thrust vector. This disturbaue would raise the nutation motion. Howeve
s known that if the changing rate of the thrust force is enough small compared to the spin rate, the declination of the angular momentum vector after the delta-V depends only on the initia
051015202501000
20003000400050006000Thrust [
N]
Time [s]
Fig.5 : Prospective thrust pattern of the deorbit
motor
We simulate numerically the attitude motion of
the spacecraft under the assumption that there is a
offset of 4mm between the thrust force vector and
x-axis.
Fig.6 shows time history of the nutation angle and
the thrust force while the delta-V. The spin rate
ab es
depend on disturbanc e generated by the
misalignment.
W
tion angle converges.
out x-axis is 1.0Hz. The nutation angle change torqu
hen the spin rate is enough high compare to the changing rate of the thrust force, the disturbance torque is averaged during a spin period. As a result, the nuta
010200
00 10 20
1000
2000300040005000600080
0
20
4060
Œo‰ßŽžŠÔ [sec]
ƒjƒ…[ƒe
„—Í [N][ƒVƒ‡ƒ“Šp[deg]„—Í—š—ð
ƒjƒ…[ƒe[ƒVƒ‡ƒ“Šp
Time[sec]
Nutati [
Thrust Force[N]onAngledeg]Thrust Force
Nutation Angle
Fig.6 : Time History of Nutation Angle while
Delta-V
Fig.7 shows the inclease of the velocity and the
declination angle of the angular momentum vector
after the delta-V for each spin rate.
If the spin rate is lower than 0.8Hz, the efficiency
of the delta-V is drop significantly. For such low
spin rate, the nutation motion diverges and the
spa
In conclusion, the x should be upped to
ot less than 1.0Hz just before the delta-V.
cecraft turns over. -axis spin
n
0.511.520
10
202000
0
000 1
spin rate [Hz]
|ઠm/
ƒÆv [degV| []s]ƒÆ|‡™V|
Fig.7 : The efficiency of the ‡™V for each spin rate
4-3. Propellant Analysis
The amount of propellant is restricted under 300[g]
by
ach attitude control laws, and design the attitude
control seque
The requir up/down is
c
the design condition of the spacecraft. Westimate the amount of the propellant consumed in
ee
nce properly. ed propellant in spin
alculated as follow: gIsplIM??Ģ=ą (1) where I is the moment of inertia about spin axis,
, the required propellant in
reorientation i ow:
l is the length of the moment arm. On the other hand
s calculated as foll
gIspl??
M I ? Ģ
=
ą Į
(2)
where ĮĢ is the amount of inclined angle of the
a
n based on the
equation (1) and (2).
Table2 :Est quired propellant
ngular momentum vector. Table 2 shows the result of estimatio
imation of the re
Required Propellant Contrmode Spin up/dowol n Reori at 1.0Hz spin
Flat Spin 156.7[g/Hz] 4.48[g/deg] Roll Spin 45.97[g/Hz] 2.00[g/deg]
Probe Sp
in 15.15[g/Hz] 0.66[g/deg]
ion control lant in
pro to h
a cons ant is r me
c the ontr the
As n ch the
orientation control is performed with the spin rate
c
of the propellant.
ission sequence is
Reorientatportion
requires much propel
the spin rate. T
is denotes that the
mount of
umed propell
educed in so
ases via
spin rate c
ol before
reorientation.
shown i
apter 3,
re
ontrol to reduce the large consumption of the propellant. To estimate the amount of required propellant through the sequence designed at chapter 3, we use also the numerical simulation. On the simulation, the spacecraft control the attitude in line with the designed sequence, and calculates the expected consumption
Fig.8 shows the result of time history of the amount of propellant consumption. In addition to the eventual amount of the propellant, extra propellant is required to keep the tank pressure. Consequently, the total amount of required propellant is 285g. This result denotes that the all m
feasible with the propellant stored in the tank. However, in case the reorientation control is performed without spin rate control, the total amount of required propellant may over 300g. Fig.8: Consumption of the Propellant 5. Conclusion This paper shows the design of attitude controlsystem t. The
lignment of the attitude control and determination
device is designed mission sequence
of 3 stages. The ol laws are also
d
t Mechanics,
pp.117-120
[2] Fuyuhiko Kiku no, Hideo Handa,
nth, Proceedings of 15 Workshop on
for H2A piggy solar sail spacecraf
a
in line with theattitude contr
esigned, and the feasibility of the mission is examined by numerical simulations. 6. References [1] Takayuki Yamamoto, Osamu Mori, Maki Shida, gDevelopment of Gas-Liquid Equilibrium Thruster for Small Satelliteh, Proceedings of 16th Workshop on JAXA, Astrodynamics and Fligh
chi, Yusuke Ko
Takahiro Iwata, Takayuki Ono, and Nobuyuki Kawano, gAttitude Estimation for a Spin Stabilized Spacecraft from Doppler Shifth, IEICE Transactions on communications, Vol. J86-B, No.6, 2003, pp. 959-968
[3] Takanao Saiki, gAttitude Control System of Spinning Solar Sail Sub-Payload Satelliteh, Proceedings of 16th Workshop on JAXA, Astrodynamics and Flight Mechanics, pp.237-241 [4] Takanao Saiki, gReport of the Spin Axis Control Experimeth
JAXA, Astrodynamics and Flight Mechanics, pp.198-203