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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.
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