Chandrayaan-1 Instrument Host Description Extracted from PDS3 Instrument Host Catalog File. Instrument Host Information =========================== Instrument Host Name: Chandrayaan-1 Orbiter Instrument Host Type: Spacecraft Instrument Host Overview ======================== 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): ------------------------ ^ | | +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 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