| INSTRUMENT_HOST_DESC |
TABLE OF CONTENTS
----------------------------------
= Instrument Host Overview
- Spacecraft
= Spacecraft Coordinate System
= Mechanical Design
- Spacecraft structures
- Payload Interfaces
= Thermal Control
- Thermal Control Concept
- Thermal Control design
= Mechanisms
- Reaction Wheel
- Solar Arrays
- Beagle-2
- Marsis Antennas Mechanisms
= Attitude and Orbit Control System
- Star Tracker
- Inertial Measurement Units
- Sun Acquisition Sensors
- Reaction Wheel Assembly
- Propulsion Configuration
= AOCS Generic Functions
- Gyro-Stellar Estimation Function
- Reaction Wheel Off-Loading Function
- Reaction Wheel Management Function
- Thruster Modulation and Selection Function
= Propulsion Architecture Description
= RF Communications
- Overview
- Uplink
- Downlink
= Data Handling Architecture
= SSMM Software
= Instruments Software
= Ground Segment Overview
Instrument Host Overview
========================
Data obtained from the Mars Express instruments were send to ground
via the spacecraft on-board computer. As spacecraft to Earth
communication does typically exclude instrument operations, all data
are relayed from the instrument to the spacecraft mass memory, the
solid state mass memory (SSMM). The data was downlinked to Earth via
the telemetry subsystem using ESA's antenna in New Norcia, Australia,
and the Deep Space Network (DSN) antennas of NASA. The radio science
experiment required data from the New Norcia and the DSN ground
station hardware.
This catalogue file gives an overview of the spacecraft and the
ground stations used.
For more detailed information see the spacecraft user manual,
MEX-MMT-MA-1091.
Instrument Host Overview - Spacecraft
=====================================
The spacecraft baseline design was the combination of mission
customised configuration and mechanical / thermal architecture with a
Rosetta inherited avionics. The spacecraft design was driven by
mission requirements, science return and system concept. A
spacecraft articulation concept with body mounted instruments, fixed
High Gain Antenna and 1 degree of freedom steerable Solar Arrays was
baselined. The spacecraft design was based on a parallelipedic like
shape sizing about 1.7 m length, 1.7 m width and 1.4 m height. The
solar array was composed of two wings, providing a symmetrical
configuration favourable to aerobraking techniques and minimising
torques and forces applied on the arrays and the drive mechanisms
during the Mars insertion manoeuvres performed with the main engine.
So as to offer the adequate dry mass / propellant mass ratio and
large mounting surfaces and volumes for the Orbiter instruments
necessary for Mars Express, the traditional cone/cylinder central
structure has been found less efficient than a dedicated structural
concept with only a Launch Vehicle Adapter connected to a stiffened
'box' as now developed for light weight satellites. Within the
overall integrated design of the spacecraft, four main assemblies are
planned to simplify the development and integration process:
(1) the Propulsion Module with the core structure,
(2) the Y lateral walls, supporting the spacecraft avionics and the
solar arrays,
(3) the Y/+X shear wall and the lower and upper floors, supporting
the payload units. The +Zb face was nominally Nadir pointed
during science observation and Lander communication relay phases
around Mars, and supported Beagle 2 (released prior to Mars
capture) the Lander(s) relay antenna and ASPERA-3, and
(4) the X lateral walls supporting the High Gain Antenna (-X) and the
instruments radiators (-X).
Attitude and Orbit Control was achieved using a set of star sensors,
gyros, accelerometers and reaction wheels. A bi-propellant reaction
control system was used for orbit and attitude manoeuvres by either a
400 N main engine or banks of 10N thrusters. The Data Handling is
based on packet telemetry and telecommand. The Electrical Power
generation was performed by solar arrays, the power storage by a
Lithium-Ion battery. A standard 28 V regulated main bus is offered to
the payload instruments. The RF Communications function transmitted
X Band telemetry 8 hours per day via a High Gain Antenna at rates
between about 19 and 230 kbps depending of the Mars to Earth
Distance. A variable telecommand rate of 7.81 to 2000 bps was
foreseen during up to 8 hour per day.
Spacecraft Coordinate System
============================
The origin of the spacecraft Reference Frame, named Oa, was located
at the separation plane between the spacecraft and the adapter, at
the centre of the interface diameter of 937 mm.
- The Xa axis was contained in the spacecraft/launch vehicle
separation plane, and oriented toward the High Gain Antenna side
of the spacecraft.
- The Za axis was coincident with the launcher X1-axis. It
represents the SC line of sight toward Mars during science
operation, and the ejection direction for the Beagle 2 probe.
- The Ya axis was contained in the SC/LV separation plane, and
oriented so as to complete the right handed co-ordinate system. It
is therefore parallel to the solar array plane and positively
oriented opposite to Marsis antenna support wall.
The (Ob, Xb, Yb, Zb) Reference Frame is structure related, and is not
used at S/C or operations level.
Spacecraft Structure and Interface With Payload Units
=====================================================
The selected SC structure limits the number of complex elements to
the bare minimum. Indeed, the only cylindrical part of large
dimensions was the Launch Vehicle Adapter ring, the rest of the
structural items being principally flat, standard panels with
aluminium skins and aluminium honeycomb. The structure was composed
of:
1) a Core Structure, built up from :
- One Launch Vehicle Adapter ring machined from a solid aluminium
cylinder of approx. diameter 940 mm, 200 mm height with a
thickness of 3.5 mm. This LVA was the main load path transfer from
the Spacecraft body to the launch vehicle interface.
- Two Tank Beams supporting the lower tank bosses, and embedded in
the LVA ring,
- Two Upper Tank Floors, supporting the tanks upper bosses,
- One Lower Floor.
- Two X Shear Walls,
- One Shear Walls in the Y direction
2) an Outer Structure, built up from :
- a +Z Top Floor,
- two +Y and -Y Sidewalls,
- two +X and -X Lateral Closure Panels (split to allow separate
access into each quadrant),
- various dedicated equipment support panels (PFS, Omega and
pressurant tank)
- miscellaneous brackets (e.g. to support sensors, antennas,
propulsion items, instruments).
All these elements were made of Aluminium Alloy, either from forgings
(LVA ring, tank beams and main brackets) or from honeycomb sandwich
panels. The panels were made of honeycomb (generally type 1/8-5056-
0.001P) of 10 to 20 mm thickness, bonded to Aluminium facesheets of
thickness varying between 0.2 to 0.3 mm and up to 0.5mm additional
doubler for local reinforcements.
In general, the payload units were accommodated following their main
needs. The payloads needing a stringent thermal control and/or
pointing performances (HRSC, OMEGA, PFS, SPICAM) were gathered on, or
close to, the +X shear wall, inside the spacecraft and close to the
AOCS reference (namely the Inertial Measurement Package and the Star
Sensors). In order to meet the PFS scanner to PFS sensor co-alignment
without disturbance caused by dismounting, these PFS units were
installed on a stiff, removable mounting assembly which can be
integrated as a single unit on the spacecraft. To expedite
installation of the large Omega-SA, this instrument was installed via
edge-mounted inserts in the Y-shear wall and a dedicated Omega
support panel.
The payloads requiring a large field of view and not necessitating
stringent thermal control were located externally on the Top and
Bottom Floors (ASPERA) or lower edge of the +Y sidewall (MARSIS).
None of the payload units required isostatic mountings: they were
rigidly fixed to the spacecraft structure utilising standard space-
industry inserts and screws.
Some of the payload units were of significant mass and therefore
require the implementation of large, face-to-face inserts that are
bonded inside the sandwich panels at the time of panel moulding.
Beagle 2 was accommodated on the top floor of the S/C, in order to
minimise dynamic disturbance (centre of mass transfer in the (X, Y)
plane) and then maximise the reliability of the Mars orbit insertion
manoeuvre. The remaining Beagle 2 hardware after probe ejection is
constrained within 50 mm height and is thermally insulated to
minimise straylight and thermal distortion disturbances respectively.
Thermal Control
===============
The spacecraft thermal control was in charge of maintaining all
spacecraft equipment within their allowed temperature ranges during
all mission phases. The equipments fall into two categories:
- the collectively controlled units, for which the heat rejection
and heating capabilities (design and accommodation) are provided
by the spacecraft thermal control,
- the individually controlled units, self provided with their own
thermal control features (coatings selection, heaters,
insulators), for which the spacecraft thermal design controls the
thermal interfaces within the required ranges.
A passive thermal control design was implemented for the Mars Express
spacecraft; it was supplemented with an electrical heating system.
The heat rejection toward space was performed using radiators mainly
on the +/-Y panels for the platform internal units and the +X panel
for the payload equipments.
These sides of the spacecraft are the most favourable areas, being
most of the time protected from the direct sun inputs (always for the
+X side). The Mars planet flux are imposed by the spacecraft orbit
and attitude and mainly significant during the pericentre phase in
operation. The rest of the spacecraft is insulated with Multi Layer
Insulation blankets to minimise the heat exchange and the temperature
fluctuations.
The spacecraft external units (Platform and Payload units) were
thermally decoupled from the spacecraft and provided with their
individual radiator when needed. The electrical heater system allowed
to raise the temperature of the unit above their minimum allowed
limits, with temperature regulation functions provided either by
mechanical device or by the onboard software. Most of the spacecraft
units were collectively controlled inside defined thermal enclosures
in which the heat balances were controlled by proper sizing of heat
rejecting radiators and heating power implementation. It allowed to
maintain the unit temperatures to acceptable levels. The heat
transfer from the units to the radiators was performed by conduction
when unit baseplates were attached to the radiator honeycomb panels
and by radiation. In that case units and panels had a black finish to
maximise heat transfer inside the thermal enclosures.
For more demanding units like the HRSC and OMEGA cameras, and the PFS
spectrometer, featuring their own thermal control, special
precautions were taken by individual trimming of their conductive and
radiative isolation. The HRSC camera required a temperature control
within a narrow temperature range):
it was provided with a thermal strap connecting it to a dedicated
radiator tuned to limit the temperature excursion in operation within
a 10 degree C temperature range.
OMEGA and PFS are provided with dedicated radiators, implemented on
the +X side of the Spacecraft. Whatever the Sun / Earth / Mars /
Spacecraft geometry, the +X side of the Spacecraft was oriented away
from Sun over the complete Martian orbit, both during Nadir pointed
science phase and Earth pointed communication phase.
This allowed to provide the camera and the spectrometer with a
thermal interface at temperature lower than 175K and 190K
respectively during the Planet observation. The connection to the
radiators were performed by thermal straps, the radiators being
themselves decoupled from the rest of the spacecraft using thermal
blankets and insulating stand-offs.
Payload external units like MARSIS and MELACOM antennas, ASPERA-3
units, were individually controlled units. They required large field
of view and thus were directly affected by the external environment
and they had to withstand larger temperature ranges than the standard
units. They are as far as possible insulated from the spacecraft.
Their coatings were selected and trimmed to suit. The spacecraft
interface temperature had a very limited influence on their thermal
behaviour.
The propulsion equipments that were mounted internally were in
general isolated with MLI, and provided with their own thermal
control heaters: tanks, fluid lines, valves, pressure sensors. The
main engine and the thrusters had their thermal coupling with the
spacecraft tailored to meet their thermal requirement while
preserving the spacecraft thermal behaviour. They were provided with
individual electrical heaters sized to maintain these external units
within the acceptable temperature range accounting for wide change in
radiative environment.
The High Gain Antenna was using a passive thermal control: a
Kapton/Germanium sunshield was covering the whole antenna on its
front side, while a light weight MLI is used on the rear side of the
reflector.
Mechanisms
==========
The implementation of mechanisms into the spacecraft configuration
had been kept to the minimum. The mechanisms employed are those
associated with
- Reaction Wheel Assembly (RWA),
- Solar Array Drive Mechanism (SADM),
- Solar Array Hold-Down and Release Mechanism (HDRM),
- Solar Array Deployment System,
- Beagle-2 Spin-Up and Ejection Mechanism (SUEM) and the
- MARSIS antennas deployment mechanism.
REACTION WHEEL ASSEMBLY
The Attitude and Orbit Control System (AOCS) of the spacecraft
required implementation of four reaction wheels, used with a three
out of four redundancy. They were of ball bearing momentum / reaction
wheel type, for clock-wise and counter-clockwise operation, with the
wheel mass suspended by two angular contact ball bearings paired by
solid preloading. The main functions of the RWA was to ensure correct
orientation of the spacecraft in fine pointing modes, and to ensure
spacecraft manoeuvrability (e.g. at transition between Mars orbit
observation and communication phases), with minimum propellant
consumption (the only related consumption lied with wheel momentum
off-loading that had to be performed at regular intervals, typically
every 2 days.)
SOLAR ARRAY DRIVE MECHANISM
There were two SADM used on the spacecraft, one for each Solar Array
wing. The SADM were mounted on each side of the spacecraft, and were
independently controlled by the AOCS Processor Module. The main
functions of the SADM was to support the solar array wing throughout
the mission, to provide the electrical power and signal interfaces to
the spacecraft and to orient the solar array wing towards the Sun by
rotation about the Ys axis. The SADM was composed of a motor and
gearbox assembly, ensuring the orientation of the solar array by
rotation, a shaft and bearing assembly ensuring mechanical connection
and pointing accuracy, a twist capsule unit transferring electrical
power to the spacecraft. Those elements were mounted on a baseplate
which was attached to the spacecraft sidewall.
SOLAR ARRAY HOLD-DOWN AND RELEASE MECHANISM
Each wing of the solar array was attached on the spacecraft sidewall,
in launch configuration, by Hold Down and Release Mechanisms (HDRM).
Each HDRM consisted in a set of hold down bushings, attached to the
structure of each panel which were held together via a stainless
steel hold down pin of 3.5 mm diameter on a hold-down baseplate fixed
on the spacecraft sidewall. The HRDM also incorporated a pair of pyro
initiators, which were actuated after spacecraft separation from the
launcher under control of the Data Handling Processor Module. The
main functions of the HDRM was therefore to maintain safely stowed
each solar array wing and to ensure their release for proper solar
array power generation.
SOLAR ARRAY DEPLOYMENT SYSTEM
Each yoke and wing of the solar array was fitted with a deployment
mechanism that ensured proper deployment and latching of the solar
array after release of the HDRM. The deployment mechanism consisted
in a set of spring energy driven hinges mounted by pair between each
solar array panel, between the first panel and the yoke, and between
the yoke and the SADM. Each hingeline was then linked to the others
by a set of pulley and cables, that ensured a synchronised deployment
of the wing.
The torque margin of the Solar Array deployment system varied between
7.5 (at beginning of deployment) and 2.6 (at end of deployment).
BEAGLE 2 SPIN-UP AND EJECTION MECHANISM
Beagle 2 formed an integrated experiment, composed of a lander
(featuring investigation experiments) encapsulated in a Entry,
Descent and Landing System (EDLS). Those items were composing the
probe, which interfaced to the orbiter top floor through the Spin-Up
and Ejector Mechanism.
MARSIS ANTENNAS DEPLOYMENT MECHANISMS
The baseline configuration for MARSIS deployment mechanism had
departed from the Cassini (STEM) concept, i.e. a tubular antenna made
of 2 semi-circular formed strips made of Copper-Beryllium.
The selected design was the ASTRO one, consisting of a boom made of a
GFRP tube pierced with 2 diametrically opposed diamond shaped holes
at the selected distance to provide folding capability. The antenna
boom contained two wire elements forming the active radioelectrical
part of the antenna, and was folded at each hollowed hinge and held
flattened in specific containers. When release was initiated, the
container was opened through pyro devices, and the boom was self
deploying thanks to its intrinsic energy which had been stored during
the folding/flattening process necessary to meet the launch volume
constraints.
Attitude and Orbit Control System
=================================
AOCS BASIC CONCEPTS
Due to the selection of a fixed high Gain antenna (HGA), and to the
propulsion configuration including a Main Engine, the Mars Express
mission required a high level of attitude manoeuvrability for the
spacecraft. Attitude manoeuvres were performed:
- Between the observation phase and the Earth communication phase, or
to reach specific attitudes necessary for science observations (in
particular SPICAM).
- Before and after the Lander ejection, before and after each
trajectory correction manoeuvre, performed either with the Main
Engine or with the 10N thrusters.
- To optimise the Wheel Off-Loading, through the selection of an
adapted attitude for this operation.
All the attitude manoeuvres of operational phase were defined on
ground, using a polynomial description of the Quaternion to be
followed by the Spacecraft. The attitude estimation was based on Star
Tracker and gyros, ensuring the availability of the measurements in
almost any attitudes. Some constraints had however to be fulfilled,
the Star Tracker being unable to provide attitude data, when the sun
or the planet are close to or inside its Field of view. Reaction
wheels were used for almost all the attitude manoeuvres, providing a
great flexibility to the Spacecraft and reducing the fuel
consumption. The angular momentum of the wheels had however to be
managed carefully from ground.
STAR TRACKER (STR)
The Star Tracker (STR) was the main optical sensor of the AOCS, used
at the end of the attitude acquisition to acquire the final 3-axes
pointing, and during almost all the nominal operations of the
mission. A medium Field Of View (16.4 deg circular) and a sensitivity
to Magnitude 5.5 were used to provide a 3-axes attitude measurement
with at least 3 stars permanently present in the FOV.
The STR included a star pattern recognition function and can perform
autonomously the attitude acquisition. The Mars Express Star Tracker
was produced by Officine Galileo, and is similar to the Rosetta one,
except at S/W level. 2 Star Trackers were implemented on the minus Xa
face of the Spacecraft, with an angle of 30 degree between their
optical axes.
INERTIAL MEASUREMENT UNITS (IMU)
Two Inertial Measurement Units (IMU) were used by the AOCS, each IMU
including a set of 3 gyros and 3 accelerometers aligned along 3
orthogonal axes. The AOCS control used either the 3 gyros of the same
IMU (reference solution at the beginning of life) or any combination
of 3 gyros among the 6 provided by both IMUs. For the accelerometers,
only a full set of accelerometers of one single IMU was used, due to
the lower criticality of the accelerometer function, and to the
availability onboard of an alternative method for the delta V
measurement (pulse counting). The Gyros were useful during the
attitude acquisition phase for the rate control, during the
observation phase to ensure the required pointing performances and
during the trajectory corrections, for the control robustness and
failure detection. A non mechanical technology was selected to avoid
the mechanical sources of failure in flight. The Accelerometers were
essential during the main trajectory corrections such as the
insertion manoeuvre to improve the accuracy of the delta V. The IMU
of Mars Express is identical to the Rosetta unit. Only the number of
units and the onboard management of the configuration was different.
SUN ACQUISITION SENSORS (SAS)
Two redunded Sun Acquisition Sensors (SAS) were implemented on the
Spacecraft central body and are used for the pointing of the Sun
Acquisition Mode (SAM) during the attitude acquisition or
reacquisition in case of failure. The SAS are identical to Rosetta
units, but provided with customised baffles.
REACTION WHEEL ASSEMBLY (RWA)
The Reaction Wheel Assembly (RWA) included 4 Reaction Wheels (RW)
implemented on a skewed configuration. This configuration enabled to
perform most of the nominal operations of the mission with a 3 RWL
configuration among 4. During some critical phase during which the
transition to the SAM had to be avoided (before lander ejection and
before Mars Insertion Manoeuvre), a 4 wheels configuration was be
used, under ground request. The Reaction wheels provided the AOCS
control torques during all the phases of the mission except the
trajectory corrections, the attitude acquisition and back up modes.
PROPULSION CONFIGURATION
The Propulsion configuration included a Main Engine (414 N) which was
used to perform all the major trajectory changes, and 10 N thrusters
used for the attitude control and also to produce the thrust during
the small trajectory corrections. The 10 N thrusters configuration
was optimised to perform all the attitude control functions with only
4 redunded thrusters, each of them being implemented near a corner of
the -Z face of the spacecraft.
SOLAR ARRAY DRIVE MECHANISM
2 redunded Solar Array Drive Mechanisms (SADM) were implemented on
the Y+ and Y-walls of the spacecraft to control the orientation of
the Solar Arrays. The SADM was only used for large angle orientation
of the wings, the selected flight orientation during the observation
phase near pericentre requiring no SADM actuation, once the
observation attitude was reached. The SADM used a stepper motor, a
gear, and a twist capsule technology. The SADM motion is defined in
the range +/-180 deg (minus margins). The SADM is identical to the
Rosetta unit, except for the speed levels which are specific to Mars
Express.
AOCS HARDWARE ARCHITECTURE
AOCS unit Nb Technology / characteristics Heritage Supplier
------------ ---------------------------- ---------- --------
Star Tracker 2 CCD detector. 16.4deg Rosetta unit. Officine
circular FOV/ Magnitude 5.5 Galileo
Gyro/accelero 2 Ring Laser Gyros (RLG). Rosetta unit Honeywell
3 gyros/3 acceleros per
unit.
Sun Acquisition 2 Solar cells mounted on Rosetta/SOHO TPD-TNO
Sensor (SAS) a pyramid
Reaction Wheel 4 Ball bearing Momentum/ Telecom. Sat. Teldix
Reaction wheels. Unit
12 Nms/0.075 Nm
SADM 2 Stepper motor with gear. Rosetta unit Kongsberg
Twist capsule
AOCS Generic Functions
======================
The AOCS modes used generic functions for the guidance, the attitude
estimation and the actuators management. The role of the guidance was
to provide onboard the reference attitude to be followed at each time
of the mission by the attitude control. It concerned of course the
orientation of the Spacecraft but also the Solar Array position. The
analysis of the mission needs showed that 4 types of guidance are
necessary along Mars Express mission:
- Pointing of the High Gain Antenna (HGA) towards the Earth, and the
Solar Array cells towards the Sun. This kind of guidance was used
during the cruise phase and for communications during the
scientific mission phase, these two cases corresponding to the AOCS
Normal Mode, pointing on ephemerides (NM/ GSEP phase).
The information necessary to the guidance concerned the Spacecraft
to Earth and the Spacecraft to Sun directions. They were contained
in the ephemeris definition.
- This type of guidance was also used in a different way for the
Earth acquisition (SHM : Safe/Hold Mode), in order to perform the
autonomous orientation of the spacecraft towards the Earth. The
ephemeris data were then used to perform large angle slew
manoeuvres with thruster control.
- Attitude profiles : this type of guidance was necessary during the
observation phase for the Nadir pointing or to follow more specific
profiles. This function was ensured by an onboard profile
description based on Chebychev polynomial, the parameters being
uploaded from ground. This capability enabled also to ensure the
attitude slew manoeuvres.
- Fixed inertial pointing (fixed quaternion) : This type of guidance
was used for specific phases of the mission, during Orbit Control
Mode, Thruster Transition Mode or during the scientific mission
phase for SPICAM specific needs (in NM/FPIP and NM/WDP).
Three generic functions had been defined for this purpose at software
level :
- the Ground commanded guidance,
- the Onboard Ephemeris propagation,
- the Autonomous Attitude Guidance Function, this latter function
generating the guidance information necessary either for the fixed
Earth pointing or for the Earth acquisition in SHM.
GYRO-STELLAR ESTIMATION FUNCTION
The gyro-stellar estimation function was common to many AOCS modes :
It was initialised during the Sun Acquisition Mode (SAM) to prepare
the following Earth acquisition operation (SHM: Safe/Hold Mode). It
provided accurate attitude estimation during the Normal Mode of
course but also in the Orbit Control Mode (OCM) and Thruster
Transition Mode( TTM) for instance. The gyro-stellar estimator
processed gyro and star tracker (STR) measurements to provide an
accurate estimate of the spacecraft attitude. It was based on a
Kalman filter with constant covariance that allowed mixing
measurements at different rates (8 Hz for the gyros and 2 Hz for the
STR). The constant covariance reduces the computer load while
ensuring good performances. The estimated attitude was a quaternion
representing the spacecraft attitude in the J2000 inertial frame.
The gyro-stellar estimator also estimated the gyros drifts to limit
the attitude errors in case of STR measurement absence due, for
instance, to a temporarily STR occultation. A specific management of
the drift estimates was proposed for Mars Express, taking into
account the specific conditions of the scientific mission phase
(existence of rates due to varying profiles, and potential
occultation). The gyro-stellar estimator implemented a coherency test
between the gyro and STR measurements in order to detect failures
that could not be detected at equipment level.
REACTION-WHEEL OFF-LOADING FUNCTION
The wheel Off-Loading function enabled to manage the angular momentum
of the wheels to a target value, through thruster actuations. This
function was completely autonomous during the last phase of the Earth
acquisition sequence (SHM/EPP:Earth Pointing Phase). During the
nominal operations around Mars, it was preferable to command the
wheel Off-Loading from the ground, the date being optimised taking
into account the mission constraints. The Off-Loading function
managed simultaneously all the wheels. It included several sequences
of thruster pulses until angular momentum of each wheel was close to
the target value. This sequence was defined by a feed forward 3-axes
wheel torque command combined with a thruster pulse.
The sequence ended with a tranquillisation phase controlled by the
wheels, in order to damp the dynamic excitation generated by the
actuation of thrusters and wheels.
REACTION WHEEL MANAGEMENT FUNCTION
This function was active in all the modes controlled through wheel
torques (Normal Mode and Safe/Hold Mode at the end of the attitude
acquisition sequence), but also when the wheels were kept to a
constant speed through a specific control loop but not used in the
AOCS control, as in Orbit Control Mode, Thruster Transition Mode or
Braking Mode. Six states of the wheel configuration are possible with
this function depending on the control of the wheels in torques (t)
or in speed (s). For instance, the nominal operation in Normal Mode,
uses 3 wheels in torques (3t), but could sometime require a fourth
wheel if a hot redundancy is useful (4t). During trajectory
corrections the configuration included 3 wheels controlled in speed
(3s). Intermediate states are necessary between these basic
configurations in order to spin the wheels for instance (3t + 1s).
This function was also in charge of the generation of wheel torque
commands in wheel frame, and of the friction torque estimation
necessary for compensation and for the failure detection. It
interfaced also with the Wheel Off-Loading function.
THRUSTER MODULATOR AND SELECTION FUNCTION
The selected amplitude modulator and on-time summation algorithms
were re- used from Rosetta and adapted to match more efficiently the
Mars Express needs taking into account the specific thrusters
configuration.
The modulator had only one working phase where the four thrusters can
be used:
- to produce a force along the satellite Z axis direction
- to control the 3-axes satellite attitude (three torques are
commanded to the modulator).
The modulator working frequency was 8Hz. At each step, the modulation
type used (ON-modulation or OFF-modulation) was automatically
selected so as to maximise the available torque capacity for attitude
control. In the case the torque capacity was insufficient with
respect to the commanded control torque, priority is given to the
control and the commanded force ratio is automatically modified to
recover the required torque capacity. Moreover in order to limit the
actuation delay, the attitude control torque was always produced at
the beginning of the actuation period.
To limit the number of thrusters ON/OFF or to tune the control limit
cycle amplitude when using thrusters, the modulator output period had
to be changed to any period multiple of 125 ms.
Propulsion Architecture Description
===================================
A bi-propellant system based on a telecommunication spacecraft
heritage was adopted for the baseline. A set of isolation pyro valves
and latch valves had been added to ensure safe operations during
Launch and Cruise, and for a re- liable acquisition of the Mars orbit
for science mission.
At launch, the pressurant assembly (high and low pressure sections)
were all isolated from the propellant tanks by normally closed
pyrotechnic valves PVNC1 to PVNC6, by non return valves NRV1 to NRV4.
The propellant tanks are pressurised to 4 bar. Similarly, the
propellant was isolated from the Reaction Control Thrusters and Main
Engine assembly by normally closed pyrotechnic valves PVNC7 to PVNC14
and thruster/main engine Flow Control Valves (FCV).
Following separation, the pyro valves protecting the pressurant
assembly were fired to pressurise the system to its operating
pressure of 17 bar. Then the latch valves were closed, isolating the
non return valves from propellant. A pressure transducer (PT2)
located at the regulator outlet could monitor pressure build up at
the NRV location due to regulator leakage. When necessary the latch
valves were opened and the pressure relieved into the propellant
tanks. It was assumed that a pressure of up to 20.5 bar could be the
criterion to initiate an open/ close cycle of the latch valves by
telecommand.
The 20.5 bar pressure was an initial suggestion which needed to be
confirmed. It may affect component qualification issues because it
exceeds existing MEOP values for the components in the section. Short
duration opening times for the latch valves minimised propellant
vapour migration and it was essential for both oxidiser side and fuel
side latch valves to open simultaneously to limit vapour migration.
The system operates in this pressure regulated mode, using the 10 N
Reaction Control Thrusters only, during the period of the transfer
orbit to Mars.
A few days before Mars orbit insertion, the 400 N Main Engine was
primed and then calibrated by specific blank manoeuvres, combined
with re-targeting of the S/C after Beagle 2 probe ejection. This
ensured that the Main Engine could be used safely for the Mars orbit
insertion and acquisition of the operational orbit. Should a Main
Engine failure be detected at this stage, a back-up scheme, using the
Reaction Control Thrusters would have been implemented to reach at
least a degraded orbit around Mars.
After attaining the operational orbit, the pressurant and Main Engine
assemblies were re-isolated by firing all the normally open
pyrotechnic valves and closing the latch valves. The remainder of the
mission was per- formed in blow down mode, using only the 10 N
Reaction Control Thrusters. The number of Reaction Control Thruster
had been limited to 8 (4 nominal, 4 redundant), located at the bottom
(-Z) side of the spacecraft to provide thrust principally along Zb to
compensate for Main Engine thrust imbalance caused by Main Engine
alignment and Spacecraft Centre Of Mass (CoM) uncertainties. Adequate
tilting of the Reaction Control Thrusters is implemented so as to
provide the capability for torque around each main axis of the
spacecraft.
In order to maximise flexibility and adaptability to failure cases,
each Reaction Control Thruster was fitted with a Thruster Latch Valve
(TLV) upstream from the thruster Flow Control Valves, permitting
individual switch over from prime to redundant for each Reaction
Control Thruster. It had to be noted that this two-tank configuration
was compatible with a horizontal handling of the spacecraft as
required by Soyuz launch campaign, on the proviso that the tanks were
filled at least up to 62% of their maximum capacity. The
compatibility of this fill fraction wrt S/C global dynamic behaviour
was under investigation to avoid fluid/structural modes coupling.
RF Communications
=================
OVERVIEW
The communications with the Earth could be performed either in S-Band
or X-Band in accordance with ESA Standards. Two Low Gain Antennas
(LGA) allow omni-directional emission and reception in S-Band, while
a dual band 1.65 m High Gain Antenna (HGA) allows high rate TM
emission in S-Band and X-Band including TC reception in S-Band and X-
Band. Demodulation of the up-link signal was performed by the Dual
Band Transponder before routing the resulting bit flow to the Data
Handling. The stored TM within the SSMM is modulated in either SBand
or X-Band within the Dual Band Transponder, which also performed S-
Band signal amplification with 5 W. X-Band signal amplification is
performed using a 65 W Travelling Wave Tube Amplifier.
UPLINK
The communication from the ground station(s) to the spacecraft was
performed in S-Band or X-Band. Two Low Gain S- Band Antennas (LGA)
were accommodated, one on the upper Z-panel, aside of the High Gain
Antenna and the other one on the bottom of the spacecraft, thus
allowing a quasi omnidirectional coverage. The LGA was used mainly
during Launch and Early Operation Phase (LEOP), critical phases and
for emergency situations. A narrow-beam dual-band high-gain antenna
was used for all nominal mission operations for the uplink in X-Band,
like the Cruise Phase or when orbiting around Mars.
The RF uplink signal, which was modulated with packetised
telecommands as NRZ/PSK/PM data, was routed towards a diplexer,
performing frequency discrimination, and then to the Dual Band
Transponder input. The transponder performed carrier acquisition and
demodulation, and transmitted the extracted signal to the Data
Handling for further processing.
The frequencies for the uplinks are:
- 2114.676 MHZ (DSN 18) for S-Band,
- 7166.936 MHZ (DSN 18) for X-Band.
The following telecommand bit rates are handled by the Mars Express
Spacecraft as provided by the CDMU design: 7.8125 bps and 15.625 bps,
250 bps, 500 bps, 1000 bps and 2000 bps. These possible bit rates are
selectable by Memory Load Command (MLC). As a baseline, the lowest
bit rates was used in case of emergency via the Low Gain Antennas in
S-Band, while the highest ones were used operationally through the
High Gain Antenna in XBand.
DOWNLINK
A high data downlink capability was required, considering the large
data volume generated by the instruments. Nevertheless, downlink
capacity was limited by the large spacecraft to Earth distance. The
downlink of the telemetry data to the ground stations were performed
in either S or X-Band.
The frequencies for the downlink were:
- 2296.482 MHZ (DSN 18) for S-Band,
- 8420.432 MHZ (DSN 18) for X-Band.
Downlink was performed at a commandable, variable bit rate. The CDMU
design allowed to generate a telemetry flow at any bit rate
corresponding to a power of two multiplied by 32/n and lower than
262.144 bps, where n is equal to 2, 3, 5 or 7. The possible bit rates
were selected via Memory Load Command (MLC) and vary from 7.8 bps as
a minimum and can be up to 230 kbps. The bit rate to which reference
was made was the bit rate following Reed-Solomon encoding, but prior
to convolutional encoding, if any. Due to hardware limitations,
convolutional encoding was only performed for bit rates lower than
65536 bps. Above this value, only Reed-Solomon encoding was
performed.
As a baseline, the lowest bit rates were used in case of emergency
only using the Low Gain Antennas, whilst the highest ones were used
operationally through the High Gain Antenna in X Band. The variable
bit rate signal was transmitted to the Dual Band Transponder as SP-
L/PSK for bit rates lower than 65536 bps and as SP-L (no subcarrier)
for higher bit rates. This signal was phase- modulated in either S
Band or X Band by the Dual Band Transponder, and added to the MPTS
ranging signal if it had been detected on the uplink.
Data Handling Architecture
==========================
The Data Management System (DMS) was in charge of telecommand
distribution to the whole spacecraft, of telemetry data collection
from the spacecraft sub- systems and payloads and data formatting,
and of the overall supervision of spacecraft and payload functions
and health.
The DMS was based on a standard OBDH bus architecture enhanced by
high rate IEEE 1355 serial data link between the CDMU (Control and
Data Management Units) processors and the SSMM and STR. The OBDH bus
was the data route for platform and payloads data acquisition and
commands distribution via the RTU. The DMS included 4 identical
Processor Modules (PM, 1 to 4) located in the 2 CDMU. Two processor
modules were dedicated to the DMS (PM2 and PM3), and two to the
AOCS(PM1 and PM4). The PM selected for the DMS function acted as the
bus master. It was in charge of Platform subsystem management
(Communications, Power, Thermal, Payloads). The PM selected as the
AOCS computer was in charge of all sensors, actuators and Solar Array
Drive Electronics (SADE).
TC-decoder and Transfer Frame Generator (TFG) were included in each
CDMU. The Solid State Mass Memory (SSMM) was used for data storage
including 12 Gbits of memory at BOL. It was coupled to the two DMS
processors, the TFG, OMEGA, HRSC and MELACOM instruments. It stores
science and global housekeeping telemetry packets.
OVERVIEW
The Data Handling architecture was organised around the two CDMU.
They were in charge of controlling ground command reception and
execution, on-board housekeeping and science data telemetry storage
and formatting them for transmission. The on-board data management,
controlled processing and execution of on-board control procedures
belongs to their tasks as well. Each CDMU featured two MA3-1750
Processor Modules, each of them being able to process either Data
Management or AOCS software.
A built-in failure operational Reconfiguration Module (RM) ensured
system level FDIR and reconfigured the CDMU as necessary. Data
transfer with other Data Handling units were ensured using standard
links such as a redunded OBDH data bus or IEEE-1355 serial links. Two
Interface Units were performing inter- face adaptation between those
links and other spacecraft units. The AOCS Interface Unit (AIU) was
dedicated to AOCS equipment, while the RTU interfaces with the
remainders, including the Instruments. A file-organised 12 Gbits SSMM
was implemented to store the Housekeeping and the Science Data. It
also collected directly Science Data from the three high rate Payload
Instruments.
SSMM Software
=============
The Solid State Mass Memory (SSMM) consists of 2 processor systems:
- The Memory System Supervisor (MSS), dedicated to the communication
with the DMS MMS.
- The File and Packet Controller (FPC), dedicated to the file
management on the memory modules and to the data exchange with the
instruments and the TFG.
The SSMM software runs on the micro-processor based MSS and the
micro- controller located in the FPC. The main part of the SSMM-SW is
programmed in C language. Parts of the start-up function are
programmed in Assembler. The SSMM software consists in two parts:
- The Initialisation software covering the Init Mode and running in
the MSS. It was executed in MSS PROM after activation of the SSMM.
It performed the following main functions:
- initialisation of system controller and control interface
hardware, tables, data, etc.,
- load nominal software from EEPROM to RAM, (reduced) commands
handling, transition to Operational Mode.
- The Operational software covering the Operational Mode and Test
Mode. It did run in the MSS RAM and FPC RAM. It performed the
following main functions:
- execution and control of telecommands,
- configuration and test of the memory modules,
- control of data flow from instruments and to TFG to and from the
Memory Modules,
- failure handling, including management of failure log,
- failure recovery,
- creation of event report,
- housekeeping,
- TM packing for all required data, Watchdog control.
In case of fatal failure, the SW returns to the Init software to
allow for failure investigation.
Instruments Software
====================
Each instrument had its own autonomous SW, located in the instrument
electronic units. The command and control of the payloads was
performed by the dedicated Payload Management function of the DMS SW.
The physical interface of the DMS PM with the instruments is the
Remote Terminal Unit (RTU). Data exchange between the payloads and
the DMS software was performed by means of packetised TM/TC, both for
commands, housekeeping and science telemetry data.
- Commands from the Ground are routed by the DMS software to the
payloads through the RTU and the OBDH bus.
- Housekeeping data from all the instruments are transmitted from
the RTU to the DMS SW through the OBDH bus.
- Scientific data from low rate payloads (PFS, ASPERA, MARSIS,
SPICAM, VMC, OMEGA) are transmitted from the RTU to the DMS SW
through the OBDH bus.
- Scientific data from high rate payloads (OMEGA, MELACOM and HRSC)
are directly transferred to the SSMM through TM packets on the
IEEE-1355 link.
Ground Segment Overview
=======================
The Mars Express spacecraft will nominally be controlled from the ESA
New Norcia (Australia) station during the Routine Operations phase.
Shared operations with Rosetta provide a station availability of 8
hours a day (design assumption), though longer duration might be
achieved during Rosetta cruise phase. Additional Earth stations are
considered, such as ESA General Purpose Network Kourou 15m station
during LEOP and NASA DSN 34 m and 70 m stations in critical phases.
ESA GROUND SEGMENT
- ESA General Purpose Network Kourou station featuring 15 m antennas
with S-band uplink capability and S-band / X-band down-link
capability.
- ESA New Norcia station, featuring a 35 m S-band / X-band antenna
with S-band/X-band uplink and down-link capability.
DSN COMPATIBILITY
- NASA DSN stations featuring 34 m and 70 m antennas, with S-band and
X-band up-link and down-link capabilities, as described in DSN
Flight Project Interface Handbook (NASA/JPL 810.5).
Summary of ground stations nominal performances:
Kourou New Norcia DSN DSN
15m 35m 34 m 70m
S Band Uplink EIRP (2kW HPA) 81 87 98 117
Pointing Loss 0.05 0.1 0.1 0.1
Antenna Gain 48.5 55.0 55.2 61.7
X Band Uplink EIRP (2kW HPA) N/A 97 108.8 114.9
Pointing Loss N/A 0.1 0.3 0.3
Antenna Gain N/A 64.3 66.8 72.2
S Band Downlink G/T at 10 deg 29.85 37.5 40.5 46.9
Pointing Loss 0.03 0.1 0.1 0.1
Antenna Gain 49.2 56.0 56.9 62.3
X Band Downlink G/T at 10 deg 38 50.1 50.1 56.7
Pointing Loss 0.1 0.3 0.3 0.3
Antenna Gain 60.0 68.0 68.2 73.1
Acronyms
--------
AOCS Attitude and Orbit Control System
HDRM array hold-down and release mechanism
IMU INERTIAL MEASUREMENT UNITS
LV launch vehicle
MLI multi layer insulation
RWA reaction wheel assembly
SADM solar array drive mechanism
SAS sun acquisition sensors
SUEM Beagle2 spin-up and ejection mechanism
SC spacecraft
STR star tracker
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