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