Instrument Host Overview   ========================     The MErcury Surface, Space ENvironment, GEochemistry,     and Ranging (MESSENGER) spacecraft was launched from     the Cape Canaveral Air Station on 2004-08-03, on an     approximately 8 year mission to become the first     probe to orbit the planet Mercury. The initial mission     included one year orbit of Mercury which was subsequently     extended to a total of four years.      Most of the science data collected by the MESSENGER     mission will originate from instruments on the     spacecraft and be relayed via the telemetry system to     stations of the NASA Deep Space Network (DSN).  Radio     Science (RS) experiments (Doppler velocity and ranging     as well as radio occultations) require that DSN     hardware also participate in the data acquisition.     The following sections provide an overview first of     the spacecraft, then of the DSN ground system, and     the spacecraft clock reset and use of MET partitions.      For more information on the spacecraft see     [LEARYETAL2007]. For more information on the DSN see     [ASMAR&RENZETTI1993].     Instrument Host Overview - Spacecraft   =====================================     The MESSENGER spacecraft was built by the Johns     Hopkins Applied Physics Laboratory (JHUAPL) to     withstand the harsh environments encountered in     achieving and operating in a Mercury orbit.  It can be     divided into eight subsystems: structures and     mechanisms; propulsion; thermal; power; avionics;     software; guidance and control; radio frequency (RF)     telecommunications, and payload.      Structures and Mechanisms     -------------------------       The spacecraft's structures and mechanisms consist of       the composite core structure, the aluminum launch       vehicle adapter, and its deployables.        The core of the spacecraft is the integrated       structure/propulsion system. The structural load paths       are optimized by using lightweight titanium fuel tanks       designed specifically for the structural configuration       chosen. The tank struts transfer lateral loads to the       corners of the structure, allowing the use of       composite panels that are thin relative to their size.       The composite structure is designed to channel all       loads into a center column, which necessitates a       square-to-round launch vehicle adapter. This was a       machined aluminum forging, carefully tailored to       distribute evenly the structural loads from the       corners of the center column to the round vehicle       interface.        The solar panel design and development effort proved       to be challenging from a material engineering       prospective. Extensive, ply-by-ply structural analysis       was required for the solar array substrate due to the       large cantilever in the stowed configuration and the       use of high-conductivity, but relatively low-strength,       graphite-cyanate ester (GrCE) materials in the       sandwich face sheets.        The sunshade support structure is welded titanium       tubing construction. The tubing supports five       antennas, the solar monitoring sensor for the X-Ray       Spectrometer (XRS), four digital Sun sensors, and the       sunshade. The final shape of the sunshade projected       area was tailored to bring the center of solar       pressure as close to the measured center of mass of       the spacecraft as possible. By utilizing solar       pressure, tilting the spacecraft relative to the Sun       can unload the momentum wheels without expending fuel.        Three mechanical assemblies required deployment: the       two solar panels and the 3.6-m Magnetometer boom. The       solar array panels were released first and allowed to       settle; the arms were then released. The Magnetometer       (MAG) boom deployment used the same sequence. After       confirmation of full deployment, all six hinge-lines       are pinned in place to prevent hinge rotation during       high-thrust maneuvers.      Propulsion     ----------       The spacecraft's propulsion subsystem consists of its       state-of-the-art titanium fuel tanks, the thruster       modules, and the associated plumbing.        The subsystem is a pressurized bipropellant, dual-mode       system using hydrazine (N2H4) and nitrogen tetroxide       (N2O4) in the bipropellant mode and N2H4 in the       monopropellant mode. Three main propellant tanks, a       refillable auxiliary fuel tank, and a helium       pressurant tank provide propellant and pressurant       storage. The helium tank is a titanium-lined composite       over-wrapped leak-before-burst pressure vessel (COPV)       based on the flight-proven A2100 helium tank. A new       lightweight main propellant tank was developed and       qualified specifically for MESSENGER. The tank       configuration is an all-titanium, hazardous-leak-       before-burst design. A small 6Al-4V titanium auxiliary       tank is a hazardous-leak-before-burst design. It has       an internal diaphragm to allow positive expulsion of       propellant for use in small burns.        The propulsion subsystem includes a total of 17       thrusters of three types. The large velocity adjustment       (LVA) thruster is a flight-proven Leros-1b.  Four 22-N,       monopropellant LVA-TVC thrusters provide thrust vector       steering forces during main thrust burns and primary       propulsion for most of the smaller velocity adjustment       (delta-V) maneuvers.  Twelve monopropellant thrusters       provide 4.4 N of thrust for fine attitude control       burns, small delta-V burns, and momentum management.        The propulsion thermal system employs heaters to       maintain acceptable system temperatures. Heaters are       used during the cruise phase to maintain propellant       temperatures and in the operational phases to pre-heat       thrusters in preparation for operation.      Thermal     -------       The thermal subsystem consists of the spacecraft's       ceramic-cloth sunshade, its heaters, and its       radiators.        The primary element of the thermal design is the       ceramic-cloth sunshade, which protects the vehicle       from the intense solar environment encountered when       inside of Venus orbit. Creating a benign thermal       environment behind the sunshade allowed for the use of       essentially standard electronics, components, and       thermal blanketing materials. Non-standard thermal       designs were required for the solar arrays, sunshade,       digital Sun sensors, three of the seven instruments,       and the phased-array antennas. These components have       been designed to operate at Mercury perihelion       (Mercury closest to the Sun) and also during orbits       that cross over one of Mercury's hot poles.      Power     -----       The power subsystem consists of the spacecraft's solar       arrays, battery, and the controlling electronics.        The power subsystem utilizes a peak power tracker       architecture that isolates the battery and the power       bus from the variations of the solar array voltage and       current characteristics and optimizes the solar array       power output over the highly varied operating       conditions of the mission. The power system is       designed to support about 390 W of load power near       Earth and 640 W during Mercury orbit.        Because of the mission's large solar distance       variations, the requirements placed on the solar array       design are rather severe. The maximum solar array       voltage during normal operations is expected to vary       between 45 and 95 V, but this range does not include       the higher transient voltages expected on the cold       solar arrays at exits from eclipses. Triple junction       solar cells are used. The solar cell strings are       placed between Optical Solar Reflector (OSR) mirrors       with a cell-to-OSR ratio of 1:2 to reduce the panel       absorbance. Thermal control is performed by tilting       the panels away from normal incidence with increased       solar intensity. In case of an attitude control       anomaly near Mercury, the solar array temperature may       reach 270 degrees C. All material and processes used       in the solar panels are designed to survive the worst-       case predicted temperatures.      Avionics     --------       The avionics subsystem consists of the spacecraft's       processors, its solid-state data recorder, and the       data handling electronics.  These components are       packaged as an Integrated Electronics Module, or IEM.        The IEM implements command and data handling (C&DH),       guidance and control, and fault protection functions.       Its design utilizes five daughter cards: the Main       Processor (MP), the Fault Protection Processor (FPP),       the 8-Gbit Solid-State Recorder (SSR), the Interface       Card and the Converter Card.  These cards are tied       together by a backplane, and a chassis utilizing the       6U compact peripheral component interconnect (cPCI)       standard.  Three of the cards (MP, FPP and SSR)       implement fairly generic functions while the Interface       and Converter Card are much more mission-specific.        A primary driver of the IEM architecture was to       simplify spacecraft fault protection. The FPP       independently collects spacecraft health information       which are continuously evaluated by a rule-based       autonomy system. The FPP corrects faults by sending       commands to the MP and other components. The IEM       Interface board includes hardware limits to prevent a       failed FPP from continuously sending commands that       would disrupt the spacecraft operation.      Software     --------       The software subsystem consists of the spacecraft's       processor-supported code that performs command and       data handling (C&DH), and spacecraft guidance and       control (G&C). It consists of two applications, the MP       and the FPP, implemented as C code under the VxWorks       5.3.1 real-time operating system.        The MP software implements all C&DH and G&C       functionality in a single flight code application       running on the MP card.        The C&DH functionality includes uplink and downlink       management, command processing and dispatch, support       for stored and time-tagged commands, management of the       SSR and file system, science data collection, image       compression, telemetry generation, memory load and       dump functions, and support for transmission of files       from the SSR on the downlink using CCSDS File Delivery       Protocol (CFDP).  It also collects analog temperatures       and implements a peak power tracking algorithm to       optimize charging of the spacecraft battery via a       power subsystem interface.        The G&C functionality maintains spacecraft attitude,       manages spacecraft momentum, executes deep-space       propulsive maneuvers, controls the solar arrays for       optimized pointing to the Sun, manages spacecraft       thermal environment by ensuring the sunshade always       faces the Sun and enables a host of pointing options       and instrument pointing control in support of science       operations.        The MP also contains a boot mode which supports       rudimentary command processing and telemetry       generation for reporting health status and to support       uploads of code and parameters to EEPROM.        The FPP application runs on the FPP card and       implements an autonomy rule engine, which accepts       uploadable health and safety rules that can operate on       data collected from the various spacecraft subsystems       via several interfaces, including an interface to the       power subsystem. The action of each rule can dispatch       commands to the MP or to the power subsystem to       correct faults.        Fault correction can include actions such as switching       to redundant components, demotion to lower spacecraft       modes (Safe Hold or Earth Acquisition), or shedding       power loads. The FPP can swap the bus controller       functionality between MPs, power on and switch to the       redundant MP, select which of two stored applications       the MP loads and can reset the MP.        The spacecraft has two safing modes. During safing,       all time-tagged command execution is halted and the       spacecraft is taken to a pre-defined simple state.       Safe Hold is the first level of safing and assumes       knowledge of ephemeris time, orbit, and attitude.       Earth Acquisition, the lowest level of safing       responding to the most critical faults (e.g., battery       at low state of charge) and no knowledge of ephemeris       time, orbit, or attitude (with respect to the inertial       reference frame) is assumed. The spacecraft is put       into a slow rotation (one revolution every 3.5 hours)       allowing the antenna suite to sweep past the Earth       periodically regardless of location.      Guidance and Control     --------------------       The primary functions of the guidance and control       subsystem are to maintain spacecraft attitude and to       execute propulsive maneuvers for spacecraft trajectory       control. It consists of the spacecraft's attitude       sensors including star cameras and Sun sensors       integrated with controllers including reaction wheels.       The system enforces two attitude safety constraints:       the Sun Keep-In constraint that keeps the sunshade       pointed towards the Sun to protect the spacecraft bus       from extreme heat and the hot-pole keep-out constraint       that protects components on the top deck from re-       radiation of sunlight from the surface of Mercury.        The sensor suite consists of star trackers, an IMU,       and Sun sensors. The primary actuators for maintaining       attitude control are four reaction wheels, each of       which provides a maximum torque of 0.075 Nm and can       store up to 7.5 Nms of momentum. Thrusters in the       propulsion system are used for attitude control during       TCMs and momentum dumps and may also be used as a       backup system for attitude control in the event of       multiple wheel failures.        The G&C system also interfaces with actuators for       three other spacecraft components to position them       properly based on knowledge of the Sun, Earth, and       target planet directions relative to the spacecraft.       These are the two solar array drive assemblies; the       phased array antenna; and the pivot platform for the       Mercury Dual Imaging System (MDIS) instrument. In       addition, an interface to the Mercury Laser Altimeter       (MLA) instrument provides a range and 'slant angle'       used to set the instrument's internal configuration       parameters for surface observations.      Radio Frequency (RF) Telecommunications     ---------------------------------------       The RF telecommunications subsystem consists of       redundant General Dynamics small deep space       transponders, solid-state power amplifiers, phased-       array antennas, and medium- and low-gain antennas.       The phased-array antennas have no mechanical       components that could fail in the extreme thermal       environment of Mercury. They are designed to work at       the 350 degrees C temperatures to be encountered.  The       spacecraft is the first to utilize turbo coding for       downlink, resulting in an extra 0.9 dB margin,       corresponding to nearly a 25% increased in data       return.      Payload     -------       The MESSENGER payload consists of seven instruments:       the Mercury Dual Imaging System (MDIS), the Gamma-Ray       and Neutron Spectrometer (GRNS), the X-Ray       Spectrometer (XRS), the Magnetometer (MAG), the       Mercury Laser Altimeter (MLA), the Mercury Atmospheric       and Surface Composition Spectrometer (MASCS), and the       Energetic Particle and Plasma Spectrometer (EPPS).       They are described in the MESSENGER MISSION.CAT file       and in [SOLOMONETAL2007], as well as the individual       instrument catalog files in the MESSENGER PDS       archives.        Instrument design was constrained along several       dimensions.  The payload mass was limited 50 kg for       the seven instruments. The demanding thermal       requirements to stay warm enough during cruise and       eclipse periods, but cold enough on orbit, were       significant constraints. Although the spacecraft solar       arrays generate ample power during the orbital phase       of the mission, power is much lower during the early       cruise phase, restricting the size of instrument       heaters that could be used. Power is also limited to       the battery during eclipse.        The over-all spacecraft architecture specified       distributed power and data processing for the       instruments; each instrument had its own power supply       and microprocessor. Redundant Data Processing Units       (DPU)buffer all data interfaces between the payload       elements and the spacecraft. Common, flight-ready       power supply and processor boards, including basic       software functions, are used by all but one       instrument, allowing development of a common set of       ground support equipment hardware and software. This       system architecture allowed payload development and       testing to proceed separately from the rest of the       spacecraft.     Instrument Host Overview - DSN   ==============================     The Deep Space Network is a telecommunications     facility managed by the Jet Propulsion Laboratory of     the California Institute of Technology for the U.S.     National Aeronautics and Space Administration (NASA).      The primary function of the DSN is to provide two-way     communications between the Earth and spacecraft     exploring the solar system.  To carry out this     function it is equipped with high-power transmitters,     low-noise amplifiers and receivers, and appropriate     monitoring and control systems.      The DSN consists of three complexes situated at     approximately equally spaced longitudinal intervals     around the globe at Goldstone (near Barstow,     California), Robledo (near Madrid, Spain), and     Tidbinbilla (near Canberra, Australia).  Two of the     complexes are located in the northern hemisphere while     the third is in the southern hemisphere.      Each complex includes several antennas, defined by     their diameters, construction, or operational     characteristics: 70-m diameter, standard 34-m     diameter, high-efficiency 34-m diameter (HEF), and 34     m beam waveguide (BWG).      For more information see [ASMAR&RENZETTI1993].    Instrument Host Overview - Spacecraft Clock Reset and Use of Clock   Partitions   ==================================================================     A planned reset of the on-board clock of the MESSENGER spacecraft     occurred on January 8, 2013. This was commanded by Mission Operations and     was done because the integer seconds part of the on-board mission-     elapsed-time (MET) counter is not long enough to contain the larger MET     values that would occur due to the extended mission. The MESSENGER team     elected to command the clock reset and set MET to a small non-zero value     to prevent disruptions in on-board timekeeping and other effects (that     might have occurred if the clock were allowed to automatically rollover     to 0 in early 2013) and to ensure that the MET counter would accomodate     the remaining extended mission.      As a result of the spacecraft clock reset, a discontinuity was introduced     and MET values are no longer guaranteed to be unique throughout the     mission. This ambiguity is resolved in ground processing by the use of     SPICE 'clock partitions' (partition 1 for pre-reset METs and partition 2     for post-reset METs) in the Spacecraft Clock (SCLK) kernel (which     supports mapping MET to other time forms using SPICE routines as     described below) and with MET values stored in PDS products, labels, and     for some instruments, product file names.  For MET values in products or     labels,  a '1/' or '2/' preceding MET indicates the partition, as in:      SPACECRAFT_CLOCK_START_COUNT   = '1/265485874'     SPACECRAFT_CLOCK_STOP_COUNT    = '2/100005'      When using SPICE routines, clock partition numbers should be included     with MET input values. METs expressed without an explicit partition     number are associated with clock partition 1 by default. Use of clock     partition numbers in file names for some MDIS products is described in     the MDIS EDR and CDR/RDR SIS documents with those data sets.      Spacecraft Clock (SCLK) SPICE Kernel     -------------------------------------     The SCLK SPICE Kernel provides information that correlates mission     elapsed time (MET) as measured by the spacecraft's on-board clock with     Terrestrial Dynamical Time (TDT) as defined by the International     Astronomical Union (IAU). (TDT was later redefined by the IAU and     renamed 'Terrestrial Time' with the acronym 'TT.' However, that acronym     is ambiguous in the SPICE context.). A 'partition' is a segment of time     when the MET count increments continuously. A single clock partition can     continue for years. When a discontinuity in MET occurs, a new partition     is defined. MET discontinuities can occur for a number of reasons and can     result in either a jump forward in MET or a jump backwards.     Spacecraft clock jumps or discontinuities can be either the result of     anomalies or they can be deliberately commanded. When such a     discontinuity occurs, the previous correlation of MET to TDT is not valid     from that point forward. A new clock partition must be created in order     to correctly associate MET with TDT. The SCLK contains a list of all     partitions that have been defined and specifies the MET values at which     each ends. SPICE takes the partitions into account when computing the     encoded SCLK representations of MET that make up the first field in each     SCLK record 'triplet.' Because of this, the encoded SCLK values in the     kernel increment steadily regardless of partition changes.      Prior to the MESSENGER spacecraft clock reset, the MESSENGER SCLK kernel     defined a single partition (partition 1). A second post-reset partition     was introduced (partition 2) shortly before the reset. SCLK kernels from     that time forward include both partitions.      The MET partition change is largely transparent to users of SPICE and the     SCLK kernel (with the exception of users who are converting raw MET     counts), since the MET values in the kernel are provided in encoded SCLK     form.