INSTRUMENT_HOST_DESC |
Instrument Host Overview
========================
The HAYABUSA spacecraft is box-shaped with approximate dimensions of
1.5 x 1.2 x 1.2 meters. Two solar panels with a combined area of 12
square meters provide power for the spacecraft including the four ion
engines. The spacecraft launch mass was 530 kg including 60 kg of xenon
fuel and 70 kg of hydrazine fuel for the attitude control jets. Each of
the four ion engines uses microwave ionization of the xenon fuel and
electrostatic acceleration of the ionized fuel to generate 5.2 - 23.6 mN
of thrust (Isp = 3200 s). Up to 3 of the 4 ion engines operate at one
time. The sample return capsule is 40 cm in diameter, has a height of 20
cm and has a mass of 16.5 kg. The surface hopper (MINERVA) is nearly
cylindrical (10 x 12 cm) and by itself has a mass of 591 g. The
spacecraft was designed to be three axis stabilized with 3 momentum
wheels operating in orthogonal directions.
Upon return to Earth in June 2010, the return capsule will separate
from the mother spacecraft. The return capsule will separate from the
mother spacecraft with a velocity of 0.2 m/s. Its spin rate will be 0.2
Hz. At 20 G, a timer is armed to jettison the parachute cover at an
altitude of 12 to 8 km. The capsule recovery will be near Woomera,
Australia, the landing velocity will be about 6 m/s and the landing
footprint is estimated to be 150 x 20 km. To assist recovery, the sample
capsule has a 242 MHz beacon radio signal. An introduction to the
spacecraft description and mission results can be found in
[FUJIWARAETAL2006].
Fujiwara, A., J. Kawaguchi, D.K. Yeomans, M. Abe, T. Mukai, and 17
others, The rubble-pile asteroid Itokawa as observed by Hayabusa,
Science 312, 1330-1334.
System Description
==================
The spacecraft coordinate system has the high gain antenna (HGA)
pointing in the +Z direction with the AMICA camera, NIRS, LIDAR, XRS,
along with the low gain X and S band antenna pointing in the -Z
direction. The ion engines point in the +X direction with the star
tracker pointing in the -X direction. The solar panel nearest the star
tracker is in the +Y direction (Y = X x Z) with the second solar panel
pointing in the -Y direction. Bi-propellant attitude control thrusters
are located on each corner of the spacecraft. The power output of the
solar panels is 2.6 kw at 1 AU.
Along with the solar panels and HGA, the science instruments are hard
mounted to the spacecraft. The LIDAR, XRS, AMICA and sample horn are
attached to the outside of the -Z deck with the NIRS mounted on the -X
deck but pointing in the -Z direction. MINERVA is located on the -Z axis
near the sampling horn. The AMICA, NIRS,and LIDAR have co-aligned
fields-of-view. The interior of the spacecraft contains the propulsion
module.
The ionized plasma is accelerated by high-voltage electrodes through
four thruster heads which protrude from one side of the spacecraft body
to provide ion thrusting of the spacecraft. A nitrogen
tetroxide/hydrazine propulsion system with a peak thrust of 22 N was used
for maneuvering until late December 2005 when this fuel was exhausted.
The spacecraft is powered by gallium-arsenide solar cells and a 15 A-hr
rechargeable nickel-metal hydride (Ni-MH) battery. The battery suffered
degredation during the close proximity spacecraft activities in late
2005.
Because of the sun, Earth and asteroid geometry at encounter and the
spacecraft design, most often the spacecraft hovered above the asteroid
along the Sun-asteroid line so that radio communications with Earth
could be maintained at the same time the solar panels were directed
toward the sun. Three optical cameras were carried on board the
spacecraft. In addition to the AMICA science camera, two wide angle
cameras were used for optical navigation. They were equipped with
500 x 500 CCDs with a resolution of 216 arcsec/px, 13 mm focal lengths
and with a f ratio of 9.6. The spacecraft was also equipped with a star
tracker (266 x 384 CCD, focal length = 12 mm, f/1.4, resolution = 375
arc seconds). Of the four laser range finders (LRF) on board, three were
used to measure the distance to the asteroid's surface when the the
spacecraft was close (120 to 7 meters) while the fourth LRF was to be
used to detect the deflection of the sampling horn when it touched the
surface; this would then trigger the pellet firing to generate surface
sample ejecta to capture.
The coordinates and some of the larger elements of the spacecraft are
shown in the schematic figure below.
+Z
/\ HGA
-------------------
\ /
\ /
--------------------------
| |
| |_
| |_] ion engines
| |_] +X
_| |
NIRS | | |
----| |--------------| |-
| | LIDAR XRS
| |
| | -Z
| |
| |
| | Sample horn
/ \
Communications Subsystem
========================
The communications subsystem includes the 1.6 meter two axis gimbaled,
parabolic high gain antenna (HGA), two medium gain horn antennae (MGA)
located on the HGA deck, four micro strip low gain antennae (LGA), an
X-band transmitter (XTX), and two X-band receivers (XRX). The uplink
frequency is 7156.533 MHz and the downlink frequency is 8408.210 MHz.
The signal has right hand circular polarization. The XRX receives and
demodulates command signals and receives and filters range signals. The
maximum transmitter power is 20 W and the bit rates can be set at 8, 16,
32, 64, 128, 256, 512, 1024, 2048, 4096, or 8192 bps with a maximum of
about 5 MB per day.
Command and Data Handling Subsystem
===================================
The Data Handling Subsystem (DHS) comprises the command and telemetry
processors, a solid state Data Recorder (DR), and an interface to a
redundant data bus for communicating with other processor-controlled
subsystems. The functions provided by the DHS are command management,
telemetry management, and autonomous operations.
The command function operates on inputs from the two command receivers
at one of three rates: 15.625 bps, 125 bps, or 1000 bps. The format of
the uplinked commands is Consultative Committee for Space Data Systems
(CCSDS) compliant, with only a single virtual channel. Four types of
commands are supported: discrete commands are used to turn on or off a
function provided by a spacecraft subsystem; serial magnitude commands
are used to set values to one or more parameters of a subsystem; user
data commands are used to send a block of user data to a subsystem; and
memory write commands are used to write a block of data to a portion of a
memory installed in a subsystem. The DHS relays these commands as they
are to intelligent subsystems (or components) using the data bus, or
sends discrete and serial magnitude commands as physical signals to
non-intelligent subsystems (or components) on serial interfaces. A series
of commands that perform a specific function can be stored as a command
macro, which can be executed with a command sent from the ground or by
one of the methods described below.
The DHS can also store a sequence of time-tagged commands or command
macros, which is called a timeline, for later execution at specified
spacecraft times. When the DHS detects an anomaly, it can automatically
execute a command or command macro. Intelligent subsystems can also
request the DHS to issue a command or command macro to another subsystem.
During normal operations, most of commands are executed from a timeline,
which have been pre-stored from the ground. In this way, most operations
are carried out when the spacecraft is out of ground contact.
The telemetry function collects engineering status and science data
using dedicated serial interfaces from non-intelligent subsystems (or
components), or using the data bus from intelligent subsystems (or
components). These data are always packetized either by the DHS or by
intelligent subsystems (or components), and placed into CCSDS-compliant
transfer frames. The transfer frames are directed to the DR, the
downlink, or both. Data recorded on the DR are read back, packed into
transfer frames and placed into the downlink on command. Recorder
playback data can be interleaved with real time data on the downlink.
The downlink data rate is selectable among 11 rates ranging from 8192
bps to 8 bps to match the communication link capability throughout the
mission.
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