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Instrument Host Overview
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Chandrayaan-1, the first Indian Mission to Moon, was launched on 22
October 2008 at 00:52 UT on-board an upgraded Polar Satellite Launch
Vehicle (PSLV-C11) from the Satish Dhawan Space Center (SDSC) in
Sriharikota located along the southeast coast of India. The lift-off and
dry mass break-up were as follows:
Lift-off mass : 1380 kg
Dry mass : 560 kg
Propellant mass : 818.2 kg
Pressurant mass : 2.84 kg
The Chandrayaan-1 orbiter spacecraft adopted a judicious choice of flight
proven as well as technology demonstration elements, while ensuring a
reliable lunar mission. The spacecraft was designed to meet the mission
specific needs such as solar array power, payload pointing requirements,
data transmission, storage schemes, and autonomous operations required in
different phases of the mission. Systems such as gyroscopes, star
sensors, and communications were miniaturized. Accommodation of eleven
scientific instruments from various space agencies and meeting their
stringent technical requirements in a small satellite bus was a
challenging task for spacecraft design.
The Chandrayaan-1 orbiter spacecraft design was adapted from the flight
proven Indian Remote Sensing (IRS) Satellite bus. Chandrayaan-1 had a
canted solar array because the orbit around the Moon was inertially fixed,
resulting in large variation in solar incidence angle. A gimballed high
gain antenna system was employed for downloading the payload data to a
Deep Space Network (DSN) established near Bangalore. The spacecraft was
cuboid in shape, measuring approximately 1.5 meters per side. It was a
three-axis stabilized spacecraft which generated about 750 W of peak power
using the solar array and was supported by a lithium ion battery for
eclipse operations. The spacecraft used a bi-propellant system to carry
it from the elliptical transfer orbits through lunar transfer orbit and to
maintain attitude during lunar orbit. The Telemetry, Tracking and Command
(TTC) communication was in the S-band. The scientific payload data were
stored in two solid state recorders (SSR #1 and SSR #2) and subsequently
played back and downlinked in X-band with a 20-MHz bandwidth by a
steerable antenna pointing at the DSN.
Spacecraft Structural Overview
==============================
The structure subsystem for the Chandrayaan-1 orbiter spacecraft provided
mechanical support for all satellite units and subsystems in a
configuration that met the system requirements of thermal control, mass
properties, alignment, launch vehicle interface, assembly, integration,
and test. The structure also provided an interface with the launch
vehicle.
The structure was 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 maintained the dimensional stability and alignment relationships
required to satisfy all mission requirements within specifications
throughout the lifetime of the satellite.
The Chandrayaan-1 orbiter spacecraft had the shape of a cuboid with a
length of 1.5 meters, a width of 1.53 meters, and a height of 1.56 meters.
The structure was designed with a central thrust bearing cylinder extended
above the cuboid to a height of 2.18 meters. The cylinder was made of
composite face skin/aluminium sandwich construction with an outer diameter
of 916.6 mm, an inner diameter of 888 mm, and a height of 2061 mm. The
cylinder had a bottom ring with a provision for an interface to the launch
vehicle. Two propellant tanks were housed inside the cylinder. The tanks
were connected to the cylinder at 18 discrete points using post-bonded
inserts. Extra stiffening layers were provided near the interface ring,
oxygen and fuel tanks, intermediate stiffener, top deck, and payload-top
deck interface to diffuse the joint interface stresses. The interface
ring of the cylinder provided interface to lam engine support structure.
The outer skin of a carbon fiber reinforced polymer (CFRP) sandwich
cylinder provided interface to the shear web joining angles. There were
four shear webs which were 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), two anti
moon viewing (AMV) decks, and the payload deck (PD) were aluminium
sandwich panels. The payload top deck (PT) was a composite construction.
The majority of the payloads were accommodated on the ASS, MV, PT, and PD
decks. The SS panel supported the solar panel. The top deck carried a
reaction wheel and the star sensors. The bottom deck provided an
interface for the eight thrusters. The AMV panel provided an interface for
the Dual Gimbal Antenna (DGA) mechanism support structure.
There were four main shear webs. The sun side (SS) shear panel was 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 provided
support to ASS panel. The PD deck apart from accommodating payloads also
provided support to the MV deck. The two AMV shear webs provided support
to pressurant tank and DGA support structure.
Various brackets were used to mount the sensors and thrusters maintaining
their requirements for stiffness, field of view, and non-interference with
other subsystems. The reaction wheel support bracket was identical to
that of ISRO spacecraft but with the location shifted from the bottom deck
to the top deck. A sandwich cylinder with the top closed with a sandwich
deck provided support to DGA.
Spacecraft Panels and Payload Interfaces
----------------------------------------
The coordinate system of the Chandrayaan-1 orbiter spacecraft is
defined as follows:
* Origin is in the centre of the spacecraft to the launcher adapter.
* The Y-axis (Roll axis) is perpendicular to the launch interface
plane, directed positively through the spacecraft body
* 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.
* The Z-axis (Pitch axis) completes the right-handed system.
A view of the spacecraft and the layout is provided here.
+X s/c side view (+Yaw):
------------------------ ^
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| +Roll (+Y)
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/ \ | |
/ \ | |
/ o| +Ysc |
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.______|______. +Xsc is out
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+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 only 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 degrees about Y axis towards the -Z
axis of the spacecraft. The angle between the two sensors is about 70
degrees. Eight, 22-Newton thrusters are mounted on the -Y face (-Roll
face) of the spacecraft.
The TMC, LLRI, M3, and SIR-2 instruments are mounted on the -Z face (-
Pitch face) looking towards to Moon view side (+X). HEX, HySI, and C1XS
are mounted on the payload panel of the +X face. Mini-SAR antenna is
mounted on the +X face at an angle 32.8 degree from the +Z axis. The
MIP instrument is mounted on +Y face (+Roll face). RADOM, SWIM, XSM,
and CENA are mounted on the MIP deck of the +Y face (+Roll face).
Propulsion
==========
A unified bi-propellant system was employed for orbit raising and attitude
control. It consisted of one 440-Newton engine and eight, 22-Newton
thrusters mounted on the negative roll face of the lunar craft. Two tanks,
each with a capacity of 390 litres, were used for storing fuel and
oxidizer. The attitude control thrusters provided the attitude control
capability during the various phases of the mission such as orbit raising
using liquid motor, attitude maintenance in LTT, lunar orbit maintenance,
and momentum dumping.
Thermal Control
===============
The thermal control system maintained the temperature of the Chandrayaan-1
orbiter spacecraft and it subsystems within the operating limits
throughout the mission. The large variations of lunar thermal heat flux
with latitude and longitude and the many constraints on vehicle attitude
and orbit combined to make the prediction of the lunar craft temperature a
difficult task. The influence of orbital variables on the heating of the
lunar craft could be appreciable, especially when the spacecraft was
orbiting close to a celestial body. For typical moon orbits, the lunar
heat striking a satellite was considerable. However, the large wavelength
of lunar heat had a different impact on the lunar craft compared to solar
heat at shorter wavelengths. It should be noted that the albedo of the
Moon is only about 1/5th the value compared to Earth. The absence of
atmosphere at the Moon and the absence of convection currents do not
provide a uniform lunar surface temperature compared to Earth. These were
addressed by suitable mathematical modeling and simulation, and a suitable
thermal control was adopted. The thermal control of scientific payloads
requiring special cooling requirements were modeled and tested.
A passive thermal control system was designed for the lunar craft. Multi
layer insulation, optical solar reflectors, thermal coating, isolators,
thermal shields, etc., were used as thermal elements. Both automatically
and manually controlled heaters were used to maintain the lunar craft
above the minimum operating temperature level in eclipse periods. To
reduce the impact of the varying lunar surface temperature conditions, the
lunar craft time constant needed to be increased. This was achieved by
proper thermal isolation schemes. Thermal design was based on the results
of a thermal mathematical model of the lunar craft. The usual lumped
parameter method was used to build the thermal model. The lunar orbit
conditions and the long eclipses dictated the major thermal
requirements during the lunar phasing orbit.
Mechanisms
==========
The Chandrayaan-1 orbiter spacecraft had the following mechanisms:
* A solar array deployment mechanism - single wing with one panel
* A Dual Gimbal Antenna (DGA) pointing mechanism
* A solar panel, canted by 30 degrees
Solar Array Drive Mechanism
---------------------------
The solar array drive assembly (SADA) positioned the solar array for sun
pointing and also provided power and signal transfer from solar array to
the spacecraft through slip rings. The drive electronics provided power
to the SADA motor windings with a provision for micro stepping. SADA
was capable of driving solar panel at different orbital rates.
Dual Gimbal drive Mechanism
----------------------------
The DGA drive electronics operated two brushless DC motors as per the
tracking profile generated through the bus management unit in closed
loop. DGA electronics was an RTX 2010 micro-controller based design with
main and redundant electronics housed in a single mechanical package.
The electronics interfaced with the DGA mechanism which contains
resolvers and motors. Resolvers gave instantaneous antenna angular
measurement.
Attitude and Orbit Control
==========================
The attitude and orbit control subsystem (AOCS) in the Chandrayaan-1
orbiter spacecraft used 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 provided 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), received the attitude data
from the star sensors and body rates using the data from the mini DTGs 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
* Gyro calibration 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 22-N thrusters
* Seasonal maneuver for imaging
* Orbit maintenance
* Safe mode
* Suspended mode
The AOCS hardware architecture included these components:
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
==========================================
The communication system provided an S-band uplink for telecommand and
tone ranging functions with a near-omni receive pattern on-board to
carryout these functions for all phases of the mission. An S-band
downlink provided the housekeeping telemetry and dwell data and
retransmits the ranging signals through an omni-link.
An X-band data downlink through a steerable 0.7-meter parabolic antenna
provided the payload data and any other auxiliary data stored in the
solid-state recorders (SSRs).
The radio frequency (RF) system for TTC and Data transmission was
configured to provide link margins even with a 18-meter ground antenna
system.
Data Handling Overview
======================
The data rate of each of the 3 Stereo TMC chains were about 12.7 Mbps,
i.e., a total of 38.1 Mbps. For the HySI camera, the data rate was about
3.1 Mbps. Data handling system was required to suitably compress the
imaging data received from this camera, store the same in a solid-state
recorder (SSR) before formatting and transmitting the data through 2 QPSK
X-Band carriers from the lunar orbit to Earth. Similarly data from the
scientific payloads electronics received at a total data rate of around
120 kbps was formatted and stored before transmitting the 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 could be transmitted in real time. These data were stored in the
solid-state recorders while imaging then subsequently transmitted.
However the provision to play back some portion of the recorded SSR while
other portions are being recorded was 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 would be on, the SSR was split into two parts to
minimize the power consumption as follows: 32 Gb for imaging payload and
8 Gb for other payloads, which was kept on during non-imaging. Suitable
error correcting codes were incorporated into the transmitting chain to
improve the link margin. To meet the mission requirements of imaging and
transmission durations, suitable data compression techniques were included
in the transmitting chain prior to formatting. However, a provision was
made to transmit raw data if necessary. The solid-state recorders were
designed to cater to the mission requirements.
Bus Management Unit
===================
The bus management unit (BMU) in the Chandrayaan-1 orbiter consisting of a
MAR 31750 processor was a centralized electronic system with standard
interfaces to meet the various functional requirements of the spacecraft
bus. The main functions of the lunar craft to be taken care by the BMU
were attitude and orbit control, command processing, housekeeping
telemetry, sensor data processing, thermal management, payload data
handling operations, dual-gimballed data transmitting antenna pointing,
fault detection, and reconfiguration and on-board mission management. The
salient features of BMU included:
* MAR 31750 Processor based system
* 2kbps/1.0kbps/0.5k kbps (command selectable) housekeeping telemetry
on 32 kHz PSK sub carrier
* Dwell data on 128 kHz PSK sub carrier
* Simultaneous normal and dwell telemetry data from the same system is
available
* CCSDS-compatible telecommand system at 125 bps PCM/PSK system
(8 Hz PSK sub carrier)
* Object-oriented software developed using Unified Modeling Language (UML)
* High density connectors and surface mount packages for hybrid
micro circuits (HMCs) and application-specific integrated circuits
(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 complementary metal oxide semiconductor (CMOS)
PROMs (programmable read-only memories) and SRAMs (static random
access memories).
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
CIXS Chandrayaan-1 Imaging 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
ISRO Indian Space Research Organization
I/F Interface
LLRI Lunar Laser Ranging Instrument
LOI Lunar Orbit Insertion
LTT Lunar Transfer Trajectory
M3 Moon Mineralogy Mapper
MIP Moon Impact Probe
MLI Multi Layer Insulation
Mini-SAR Miniaturized Synthetic Aperture Radar
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 Solar X-ray Monitor
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