PDS_VERSION_ID = PDS3
LABEL_REVISION_NOTE = "2010-02-16, CH1-ISRO-SAC-DP-TEAM"
RECORD_TYPE = STREAM
OBJECT = INSTRUMENT_HOST
INSTRUMENT_HOST_ID = "CH1ORB"
OBJECT = INSTRUMENT_HOST_INFORMATION
INSTRUMENT_HOST_NAME = "CHANDRAYAAN-1-ORBITER"
INSTRUMENT_HOST_TYPE = "SPACECRAFT"
INSTRUMENT_HOST_DESC = "
Instrument Host Overview
========================
Chandrayaan-1, the first Indian Mission to Moon was launched on
October 22, 2008 by PSLV-C11 at 00:52 UT from SDSC, SHAR. The lift
-off and dry mass break-up is as follows:
Lift-off mass : 1380 kg
Dry mass : 560 kg
Propellant mass : 818.2 kg
Pressurant mass : 2.84 kg
The Chandrayaan-1 spacecraft adopts a judicious choice of flight
proven as well as technology demonstration elements, while ensuring a
reliable lunar mission. The spacecraft is designed to meet the
mission specific needs such as solar array, payload pointing
requirements, data transmission, storage schemes and autonomous
operations required in different phases of the mission. Systems like
gyro, star sensors, and communication system are miniaturized.
Accommodation of eleven scientific instruments from various space
agencies and meeting their stringent technical requirements in small
satellite bus is a challenging task for spacecraft design.
The Chandrayaan-1 spacecraft design is adapted from flight proven
Indian Remote Sensing (IRS) Satellite bus. Chandrayaan-1 has a canted
solar array since the orbit around the Moon is inertially fixed,
resulting in large variation in solar incidence angle. A gimbaled
high gain antenna system is employed for downloading the payload data
to Deep Space Network established near Bangalore. The spacecraft is
cuboid in shape of approximately 1.5m side. It is a three-axis
stabilized spacecraft generating about 750W of peak power using the
solar array and will be supported by a Li-Ion battery for eclipse
operations. The spacecraft adopted bipropellant system to carry it
from the elliptical transfer orbits through lunar transfer orbit and
finally in attitude maintenance in lunar orbit. The TTC communication
is in S-band. The scientific payload data is stored in two solid
state recorders (SSR #1 & #2) and subsequently played back and
downlinked in X-band through 20 MHz bandwidth by a steerable antenna
pointing at DSN.
Spacecraft Structural Overview
==============================
The structure subsystem for CHANDRAYAN Spacecraft will provide
mechanical support for all satellite units and subsystems in a
configuration that meets the system requirements of thermal control,
mass properties, alignment, launch vehicle interface, assembly,
integration and test. The structure also provides interface with the
launch vehicle.
The structure is capable of sustaining all direct and cumulative load
combinations occurring during fabrication, testing, ground handling,
transportation, launch, orbit maneuvers and deployment. On-station,
the structure will maintain throughout the satellite?s mission
lifetime, the dimensional stability and alignment relationships
required to satisfy all mission requirements within specifications.
The structure for lunar mission is a cuboid of size 1.5m X 1.53m in
plan and 1.56m high.
The structure is designed with a central thrust bearing cylinder
extended above the cuboid to a height of 2.18m. The cylinder is made
of composite face skin/aluminium sandwich construction and has a
diameter of 916.6mm OD, 888mm ID and is 2061mm tall. The cylinder has
a bottom ring, which has provision for launch vehicle interface. Two
propellant tanks are housed inside the cylinder. Tanks are connected
to the cylinder at 18 discrete points using post-bonded inserts.
Extra stiffening layers are provided near interface ring, Ox and Fuel
tanks, intermediate stiffener, top deck and payload-top deck
interface to diffuse the joint interface stresses. Interface ring of
cylinder provides interface to lam engine support structure. Outer
skin of CFRP sandwich cylinder provides interface to the shear web
joining angles. There are four shear webs viz. which are connected to
cylinder by CFRP L-angles.
Two horizontal decks (Bottom deck and Top deck) and the four vertical
decks sun side (SS), anti sun side (ASS), moon view (MV) and anti
moon viewing (AMV) decks) and the payload deck (PD) are aluminium
sandwich panels. Payload top deck (PT) is of composite construction.
Majority of the payloads are accommodated on ASS, MV, PT and PD
decks. The SS panel supports the solar panel. The top deck carries
reaction wheel and star sensors. Bottom deck provides interface for
eight thrusters. The AMV panel provides interface for the Dual Gimbal
Antenna (DGA) mechanism support structure.
There are 4 main Shear Webs. The Sun Side (SS) shear panel is offset
from the center to transfer SS panel loads as well as provide support
stiffness to Solar Array. The Anti Sun Side (ASS) shear web provides
support to ASS panel. PD deck apart from accommodating payloads also
provides supports to the MV deck. Two AMV shear webs provide support
to pressurant tank and DGA support structure.
Various brackets are provided to mount the sensors and thrusters
keeping their requirement like Stiffness, FOV and non-interference
with other subsystems. The Reaction Wheel support bracket is
identical to that of IRS spacecraft with their location shifted from
Bottom deck to Top deck. A sandwich cylinder with the top closed with
a sandwich deck provides support to DGA.
Spacecraft Panels and Payload Interfaces
----------------------------------------
The co-ordinate system of the Chandrayaan-1 spacecraft is as follows:
o Origin is in the centre of the spacecraft to the launcher adapter.
o The Y-axis (Roll axis) is perpendicular to the launch interface
plane, directed positively through the spacecraft body
o The X-axis (Yaw axis) is perpendicular to the Y-axis and the solar-
array drive axis, directed negatively through the side of the
spacecraft containing the high gain antenna.
o The Z-axis (Pitch axis) completes the right-handed system.
Views of the spacecraft and the layout is provided below.
+X s/c side view (+Yaw):
------------------------ ^
|
| +Roll (+Y)
---------
| |
.___|_____|___.
/\ | |
/ \ | |
/ \ | |
/ o| +Ysc |
/ | ^ |
/ | | |
/ | | |
.______|______. +Xsc is out
| | | of the page
+Zsc <-------o____.
/ \
/_____\ Main Engine
The -X face (-Yaw face) of the box houses the high gain antenna,
mounted on a 2-axis orientation mechanism. The -Z face (-Pitch face)
is flat, containing just a thermal radiator.Solar panels are attached
to the +Z face (+Pitch face), canted at 30 deg. Two star sensors are
mounted on +Y(+Roll) deck with rotation of 63 deg about Yaixs towards
-Z axis of the spacecraft. Angle between two sensors is about 70 deg.
Eight numbers of 22N thrusters are mounted on -Y face (-Roll face) of
the spacecraft
TMC, LLRI, M3, SIR-2 are mounted on -Z face (-Pitch face) looking
towards to Moon view side (+X). HEX, HySI, C1XS are mounted on the
mounted on the payload panel of +X face. MiniSAR antenna is monuted
on +X face at an angle 32.8 deg frpm +Z axis. MIP is mounted on +Y
face (+Roll face).RADOM, SWIM, XSM, CENA are mounted on MIP deck of
+Y face (+Roll face).
Propulsion
==========
The propellant system consists of a unified bipropellant system for
orbit raising and attitude control. It consists of one 440N engine
and 8 number of 22N thrusters mounted on the negative roll face of
the lunarcraft. Two tanks, each with a capacity of 390 litre are used
for storing fuel and oxidizer. The attitude control thrusters provide
the attitude control capability during the various phases of the
mission like orbit raising using liquid motor, attitude maintenance
in LTT, lunar orbit maintenance and momentum dumping.
Thermal Control
===============
Thermal control system maintains the lunarcraft and it?s subsystem
within the operating temperature limits throughout the mission
phases. The large variations of lunar thermal heat flux with latitude
and longitude and the many constraints on vehicle attitude and orbit
combine to make the prediction of lunarcraft temperature a difficult
task. The influence of orbital variables on lunarcraft heating can be
appreciable, especially for lunarcraft orbiting close to a celestial
body. For typical moon orbits, the lunar heat striking a satellite
is considerable. However the large wavelength of lunar heat will have
different impact on the lunarcraft compared to the short wave length
solar heat. It should be noted that albedo of the moon is only about
1/5th the value compared to the earth albedo value. The absence of
atmosphere on the moon and the absence of convection currents do not
provide a uniform surface temperature of the moon compared to a
uniform temperature earth. These are to be addressed by suitable
mathematical modeling and simulation and suitable thermal control
will be adopted. The thermal control of scientific payloads requiring
special cooling requirements will be modeled and tested.A passive
thermal control system is proposed for the lunarcraft. MLI, OSR,
thermal coating, isolators, thermal shields etc. are used as thermal
elements. Both auto and manually controlled heaters are used to
maintain the lunarcraft above the minimum operating temperature level
in eclipse periods. To reduce the impact of the varying lunar surface
temperature conditions the lunarcraft time constant needs to be
increased. This is achieved by proper thermal isolation schemes.
Thermal design is based on the results of a thermal mathematical
model of the lunarcraft. The usual lumped parameter method is used to
build the thermal model. The lunar orbit conditions and the long
eclipses dictate the major thermal requirements during the lunar
phasing orbit.
Mechanisms
==========
The lunarcraft has the following mechanisms:
Solar array deployment mechanism - single wing with one panel
Dual Gimbal Antenna pointing mechanism (DGA)
Solar panel is canted by 30 deg
Solar Array Drive Mechanism
-----------------------------------
The solar array drive assembly (SADA) positions the solar array for
sun pointing and also provides power and signal transfer from solar
array to the spacecraft through sliprings. The drive electronics
provides power to the SADA motor windings with a provision for micro
stepping. SADA is capable of driving solar panel ar different orbital
rates.
Dual Gimbal drive Mechanism
----------------------------
The DGA drive electronics drives two brushless DC motors as per the
tracking profile generated through BMU in closed loop. DGA
electronics is RTX 2010 micro-controller based design with main
and redundant electronics housed in a single mechanical package.
Electronics has interface with DGA mechanism which contains resolvers
and motors. Resolvers give instantaneous antenna angular measurement.
Attitude and Orbit Control
==========================
The attitude and orbit control subsystem (AOCS) in lunarcraft uses
the body stabilized zero momentum system with reaction wheels to
provide a stable platform for the lunar mission payloads. Together
with the propulsion subsystem, AOCS provides the capability of 3-axis
attitude control with thrusters in the transfer orbit, momentum
dumping in the lunar orbit in addition to orbit rising and fine orbit
adjustment. Attitude and orbit control electronics (AOCE) integrated
in the bus management unit(BMU) receives the attitude data from the
star sensors, body rates using the data from the miniDTGs and
computes the necessary control torque commands and outputs to the
actuators. The various operational modes are
* Rate damp
* Sun pointing
* Inertial attitude control (IAC) with thrusters
* Gyrocalibration using star sensors
* Reorientation maneuver for orbit transfer
* Attitude control during liquid motor firing for LTT and LOI
* Midcourse correction in LTT and orbit adjusts after LOI
* Normal mode lunar pointing control with wheels
* Momentum dumping using 22N thrusters
* Seasonal maneuver for imaging
* Orbit maintenance
* Safe mode
* Suspended mode
AOCS Hardware Architecture
--------------------------
----------------------------------------
| EQUIPMENT | QUANTITY |
|---------------------------| ---------|
| BMU | 2 |
|---------------------------|----------|
| SENSORS | |
|---------------------------|----------|
| Coarse Analog Sun Sensors | 6 |
|---------------------------|----------|
| Star sensor | 2 |
|---------------------------|----------|
| Solar Panel Sun sensor | 1 |
|--------------------------------------|
| Gyroscope | 1 |
|--------------------------------------|
| Accelerometer | 1 |
|---------------------------|----------|
| ACTUATORS |
|---------------------------|----------|
| Reaction Wheel | 6 |
|---------------------------|----------|
| Wheel Drive Electronics | 2 |
|---------------------------|----------|
| Solar Array Drive | 1 |
|---------------------------------------
Tracking, Telemetry and RF Communications
=========================================
Communication system provides S-band uplink for telecommand and tone
ranging functions with near omni receive pattern onboard to carryout
these functions in all phases of the mission. S-band downlink
provides the house keeping telemetry, dwell data and retransmits the
ranging signals through an omni-link. X-band data downlink through a
steerable 0.7m parabolic antenna provides the payload data and any
other aux data stored in SSRs.RF system for TTC and Data transmission
is configured to provide link margins even with 18m ground antenna
system
Data Handling Overview
======================
The data rate of each of the 3 Stereo TMC chains are about 12.7Mbps,
i.e., a total of 38.1Mbps. For the HySI camera, the data rate is
about 3.1Mbps. Data handling system is required to suitably compress
the imaging data received from this camera, store the same in solid
state recorder before formatting and transmitting this data through 2
QPSK X-Band carriers from the lunar orbit to the earth. Similarly
data from the scientific payloads electronics received at a total
data rate of around 120kbps is to be formatted and stored before
transmitting this data through the same X-band carriers as the
imaging payload data. In view of the power , data rate and RF
visibility constraints, the imaging and other payload data cannot be
transmitted in real time. These data are stored in the solid-state
recorders while imaging and transmitted subsequently. However the
provision to play back some portion of the recorded SSR while other
portion being recorded is envisaged in Chandrayaan-1 imaging system
SSR. Considering the fact that generated power during the dawn/dusk
period would be approximately 50% and only the non imaging
scientific payloads will be ON, the solid state recorder has been
split into two parts to minimize the power consumption as follows: 32
Gb for imaging payload and 8Gb for other payloads, which will be
kept ON during non-imaging. Suitable error correcting codes are
required to be incorporated in the transmitting chain in order to
improve the link margin. In order to meet the mission requirements
of imaging and transmission durations, suitable data compression
techniques will be included in the transmitting chain prior to
formatting. However, provision is made to transmit raw data if
necessary. Solid-state recorders are to be designed to cater to the
mission requirements
Bus Management Unit
===================
The bus management unit (BMU) in lunarcraft consisting of MAR 31750
processor is a centralized electronic system with standard interfaces
to meet the various functional requirements of spacecraft bus. The
main functions of the lunarcraft to be taken care by the BMU are
Attitude and orbit Control, Command processing, House keeping
telemetry, Sensor data processing, Thermal management, Payload data
handling operations, dual gimbaled Data transmitting antenna pointing
,Fault detection and reconfiguration and Onboard mission management .
The salient features of BMU includes the following:
* MAR 31750 Processor based system
* 2kbps/1.0kbps/0.5k kbps (command selectable) House keeping
telemetry on 32kHz PSK sub carrier
* Dwell data on 128kHz PSK sub carrier
* Simultaneous normal and dwell TM data from same system is
available
* CCCSDS compatible TC system at 125bps PCM/PSK system (8Hz PSK sub
carrier)
* Object oriented software developed using UML
* High density connectors and surface mount packages for HMCs and
ASICs
* Double sided mounting and use of chips for passive components
* Use of solid state switches in place of relays for heater control
* Usage of high density CMOS PROMs and SRAMs
ACRONYM LIST
=============
AOCE Attitude and Orbit Control Electronics
AOCS Attitude and Orbit Control System
BDH Baseband Data Handling
BMU Bus Management Unit
BPSK Binary Phase Shift Keying
CASS Coarse Analog Sun Sensor
CCD Charge Coupled Device
CCSDS Consultative committee for Space Data Systems
CENA Chandrayaan-1 Energetic Neutral Analyzer
C1XS Chandrayaan-1 X-ray Spectrometer
DGA Dual Gimbal Antenna
DTG Dynamically Tuned Gyroscope
DSN Deep Space Network
H/W Hardware
HEX High Energy X-ray Spectrometer
HySI Hyper Spectral Imager
IAC Inertial Attitude Control
I/F Interface
LLRI Lunar Laser Ranging Instrument
LOI Lunar Orbit Insertion
LTT Lunar Transfer Trajectory
MIP Moon Impact Probe
MLI Multi Layer Insulation
M3 Moon Mineralogy Mapper
RADOM Radiation Dose Monitor
PM Phase Modulation
PSK Phase Shift Key
TMC Terrain Mapping Camera
SADA Solar Array Drive Assembly
SARA Sub - keV Atom Reflecting Analyzer
SIR-2 Short wave Infrared Radiometer
SPSS Solar Panel Sun Sensor
SSR Solid State Recorder
SWIM Solar Wind Monitor
TTC Telemetry, Tracking and Command
XSM X-ray Solar Monitor
"
END_OBJECT = INSTRUMENT_HOST_INFORMATION
OBJECT = INSTRUMENT_HOST_REFERENCE_INFO
REFERENCE_KEY_ID = "BHANDARI2005"
END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO
OBJECT = INSTRUMENT_HOST_REFERENCE_INFO
REFERENCE_KEY_ID = "GOSWAMI&ANNADUR2008"
END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO
END_OBJECT = INSTRUMENT_HOST
END