Instrument Host Information
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       
    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).       
    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.                              
    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     
  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                                      
         Coarse Analog Sun Sensors     6                                      
         Star sensor                   2                                      
         Solar Panel Sun sensor        1                                      
         Gyroscope                     1                                      
         Accelerometer                 1                                      
         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
  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    
    * 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      
    * 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