Instrument Host Information
= Spacecraft Overview
= Mission Requirements and Constraints
= Platform Definition
= Subsystem Accommodation
= Rosetta Spacecraft Frames
= Structure Design
  - Solar Array
  - Reaction Wheels
  - Propellant Tanks
  - Helium Tanks
  - Thrusters
  - High Gain Antenna
  - Gyros
= Mechanisms Design
  - Solar Array Drive Mechanism (SADM)
  - Solar Array Deployment Mechanisms
  - HGA Antenna Pointing Mechanism (APM)
  - Experiment Boom Mechanisms
  - Louvres
= Thermal Control Design
  - Thermal Control Concept
  - Thermal control design
  - General Heater Control Concept
  - Micrometeoroid and Cometary Dust Protection
= Propulsion Design
  - Operation
= Telecommunication Design
 - High Gain Antenna Major Assembly
 - High Gain Antenna Frame
 - Medium Gain Antenna
   - MGAS
   - MGAX
= Power Design
  - Power Conditioning Unit (PCU)
  - Payload Power Distribution Unit (PL-PDU)
  - Subsystems Power Distribution Unit (SS-PDU)
  - Batteries
  - Solar Array Generator
  - Mechanical Design of the Solar Panels
  - Rosetta Solar Array Frames
= Power Constraints in Deep Space
= Harness Design
= Avionics Design
  - Data Management Subsystem (DMS)
    - Solid State Mass Memory (SSMM)
  - Attitude and Orbit Control Measurement System (AOCMS)
  - Avionics external interface
= Avionics modes
  - Stand-By Mode
  - Sun Acquisition Mode
  - Safe/Hold Mode
  - Normal Mode
  - Thruster Transition Mode
  - Orbit Control Mode
  - Asteroid Fly-By Mode
  - Near Sun Hibernation Mode
  - Spin-up Mode
  - Sun Keeping Mode
= System Level Modes
  - Pre-launch Mode
  - Activation Mode
  - Active Cruise Mode
  - Deep Space Hibernation Mode
  - Near Sun Hibernation Mode
  - Asteroid Fly-by Mode
  - Near Comet Mode
  - Safe Mode
  - Survival Mode
= Ground Station Network
  - New Norcia
  - Cebreros
  - Kouru
= Acronyms

Spacecraft Overview

Please note: The ROSETTA spacecraft was originally designed for a
mission to the comet Wirtanen. Due to a delay of the launch a new
comet (Churyumov-Gerasimenko) had been selected. The compliance of
the design was checked and where necessary adapted for this new
mission. Therefore in the following all the details and
characteristics for this new mission are used (like min and max
distance to Sun).

The Rosetta design was based on a box-type central structure, 2.8 m x
2.1 m x 2.0 m, on which all subsystems and payload equipment were
mounted.  The two solar panels had a combined area of 64 m2 (32.7m
tip to tip), with each extending panel measuring 14 m in length.

The 'top' of the spacecraft accommodated the payload instruments, and
the 'base' of the spacecraft the subsystems. The spacecraft could be
physically separated into two main modules:

    * A Payload Support Module (PSM)
    * A Bus Support Module (BSM)

The Lander was attached to the rear face (-X), opposite the two-axes
steerable high-gain antenna (HGA). The two solar wings extended from
the side faces(+/-Y). The instrument panel pointed almost always
towards the comet, while the antennas and solar arrays pointed towards
the Sun and Earth (at such great distances the Earth is relatively
speaking in the same direction). The spacecraft attitude concept was
such that the side and back panels were shaded throughout all nominal
mission phases, offering a good location for radiators and louvres.
This was normally facing away from the comet, minimising the effects of
cometary dust.

The spacecraft was built around a vertical thrust tube, whose diameter
corresponded to the 1194 mm Ariane-5 interface. This tube contained two
large, equally sized, propellant tanks, the upper one containing
fuel, and the lower one containing the (heavier) oxidiser.  At launch
the total amount of stored propellant was roughly 1670 kg.

A coarse overview on the spacecraft main characteristics is
summarised hereafter:

Total launch mass requirement:  3065 kg
Propellant mass:                1718 kg
Overall size (xyz)
        Launch configuration:   225x256x318 cm
        SA deployed:            32.7 m tip-to-tip
power provided by SA:           440 W at max dist from sun (5.3 AU)
energy provided by 3 Batteries: 500 Wh
data management:                operation of s/c according to an on-
                                board master schedule and real-time
                                via ground-link

Mission Requirements and Constraints

In the following, the stringent mission requirements are summarised
and related to their consequences on the spacecraft system design.

The ambitious scientific goals of the ROSETTA mission required:

* a large number of complex scientific instruments, to be
accommodated on one side of the spacecraft, that would permanently face the
comet in the operational phase, . During cruise the instruments would serve
 for survival.
* one Surface Science Package (SSP), suitable for cruise survival and proper,
independent ejection from the orbiter (spacecraft). In addition, the orbiter
would provide the capability for SSP data relay to Earth.
* a complex spacecraft navigation at low altitude orbits around an
irregular celestial body with weak, asymmetric, rotating gravity
field, rendered by dust and gas jets.
These primary mission requirements were design driving for most of the
spacecraft layout and performance features, as:
* data rate (DMS, TTC)
* pointing accuracy (AOCMS, Structure)
* thermal layout
* closed loop target tracking (AOCMS, NAV Camera), derived
requirements from asteroid fly-by
* small-delta-v manoeuvre accuracy (RCS)

Other mission requirements, that related to the interplanetary cruise
phases rather than to the scientific objectives, drove mainly the
power supply, propulsion, autonomy, reliability and

For achieving the escape energy (C3=11.8 km^2/s^2) to the
interplanetary injection, an Ariane 5 Launch (delayed ignition) was
required, that constrained the maximum S/C wet mass and defined the
available S/C envelope in Launch configuration.

The total mission delta-v of more than 2100 m/s required a propulsion
system with over 1700 kg bi-propellant.

The environmental loads (radiation, micro meteoroids impacts) over
the mission duration of nearly 12 years was very demanding w.r.t.
shielding, reliability and life time of the S/C components.

The large S/C to Earth distance throughout most mission phases made
a communication link via an on-board high gain antenna (HGA)
mandatory. The spacecraft had to provide an autonomous HGA Earth-
pointing capability using star sensor attitude information and on-
board stored ephemeris table. TC link via spherical LGA coverage, and
TC/TM links via an MGA had to be possible as backup for a loss of the
HGA link.

The wide range of S/C to Sun distances (0.88 to 5.33 AU) drove the
thermal control and the size of the solar generator.

The long signal propagation time (TWTL up to 100 minutes), and the
extended hibernation phases (2.5 years the longest one), and the many
solar conjunctions/oppositions (the longest in active phases is 7
weeks) required a high degree of on-board autonomy, with corresponding
FDIR concepts.

Platform Definition

The ROSETTA platform was designed to fulfill the need to accommodate
the payload (including fixed, deployable and ejectable experiment
packages), high gain antenna, solar arrays and propellant mass in a
particular geometrical relationship (mass properties and spacecraft
viewing geometry) and with the specified modularity (Bus Support
Module and Payload Support Module incorporating Lander Interface
Panel). The thermal environment also drove the configuration such
that high dissipation units had to be mounted on the side walls with
thermal louvres providing trimming for changing external conditions
during the mission.

The design of the platform's electrical architecture was driven by the
need to meet specific power requirements at aphelion (the solar array
sizing case) and to incorporate maximum power point tracking.
Additional factors such as the uncertainty in the performance of the
Low Intensity Low Temperature solar cell technology had also
influenced the design.

The telecommunications design was driven by the need to be compatible
with ESA's 15m and 35m ground stations and the 34m and 70m DSN
stations. This had produced requirements for dual S/X band and
variable rate capability, together with an articulated High Gain
Antenna to maximise data transfer during the payload operations, and
a fixed Medium Gain Antenna to act as backup for the HGA in case of

Subsystem Accommodation

The majority of the subsystem equipments were accommodated together
within the BSM. The electronic units were located mostly on the Y
panels so that their thermal dissipations were closely coupled to the
louvred radiators on the sidewalls. So far as practical, functionally
related groups were located close together for harness, integration
and testability reasons. Where possible, equipments were positioned
towards the +X half of the S/C to counterbalance the mass of the
Lander on the opposite side.

Some subsystem equipments were deliberately located on the PSM. These
included the PDU and RTU for the payload, the NAVCAMS, two of the SAS
units and the +Z LGA. The PDU and RTU were located closer to the
payload instruments to reduce harness complexity and mass, and the
NAVCAMs and SASs and +Z LGA were located on the PSM for field of view
reasons. Other subsystem equipments had been located on the PSM
sidewalls as a result of BSM equipment/harness growth, or thermal
limitations. These comprised the STR electronics and SSMM as well as
the USO.

The RCS subsystem comprised tanks, thrusters and the associated
valves and pipework. The main tanks were accommodated within the
central tube while the helium pressurisation tanks were mounted on the
internal deck. Most of the valves and pipework were located on the +X
BSM, panel which became permanently attached to the BSM once RCS
assembly was completed. Sixteen of the twenty-four thrusters were
located at the four lower corners of the BSM. The remaining
thrusters were located in 4 groups near the top corners of the S/C.
They were installed as part of the BSM, but were attached to the PSM
after PSM/BSM mating.

The Star Trackers were mounted on the -X shearwalls. The STR B was
rotated by additional 10 degrees towards the -Z direction compared to
STR A to avoid the VIRTIS radiator rim to be seen in its field of
view. This location of the STRs was both thermally stable and
mechanically close to the -X PSM panel which accommodated the
instruments requiring high pointing accuracy. The reaction wheels were
located on the internal deck which provided them with a thermo-
elastically stable location.

A 2.2m diameter HGA was stowed face-outwards for launch against the
S/C +X face (so it would be partially usable even in the event of a
deployment failure). After deployment, the HGA could be rotated in two
axes around a pivot point on a tripod assembly some distance clear of
the lower corner of the S/C. This provided the HGA with greater than
hemispherical pointing range. The two MGAs were fixed mounted on the
S/C +X face, oriented in the +Xs/c direction, as this was the most
useful direction for a fixed MGA. The LGAs were located at the +Z and
-Z ends of the S/C but angled at 30 degs to the Z axis. This
accommodation provided spherical coverage with minimum need for

The solar array comprised two 5-panel wings folded against the
Spacecraft Y axis for launch. Because the arrays were sized to operate
at aphelion, the outwards facing outer panel could also generate useful
power before array deployment.

Two Sun Acquisition Sensors were located on the solar arrays and
another two on the S/C body. Their design and location of these also
allowed them to serve as fine Sun sensors.

Rosetta Spacecraft Frame

   Rosetta spacecraft frame was defined as follows:

      -  +Z axis was perpendicular to the launch vehicle interface
         plane and points toward the payload side;
      -  +X axis was perpendicular to the HGA mounting plane and
         points toward HGA;
      -  +Y axis completed the frame is right-handed.
      -  the origin of this frame was the launch vehicle interface

   These diagrams illustrate the ROS_SPACECRAFT frame:

   +X s/c side (HGA side) view:
                                   | toward comet

                              Science Deck
  .__  _______________.     |             |     .______________  ___.
  |  \ \               \    |             |    /               \ \  |
  |  / /                \   |  +Zsc       |   /                / /  |
  |  \ \                 `. |      ^      | .'                 \ \  |
  |  / /                 | o|      |      |o |                 / /  |
  |  \ \                 .' |      |      | `.                 \ \  |
  |  / /                /   |      |      |   \                / /  |
  .__\ \_______________/    |  +Xsc|      |    \_______________\ \__.
    -Y Solar Array          .______o-------> +Ysc   +Y Solar Array
                              .'       `.
                             /           \
                            .   `.   .'   .          +Xsc is out of
                            |     `o'     |             the page
                            .      |      .
                             \     |     /
                              `.       .'
                           HGA  ` --- '

   +Z s/c side (science deck side) view:
                                /     \  Lander
                               |       |
                            |             |
                            |             |
                            |  +Zsc       | +Ysc
  o==/ /==================o |      o------->o==================/ /==o
    -Y Solar Array          |      |      |        +Y Solar Array
                            |      |      |
                             `.   |   .'
                                .--V +Xsc
                         HGA  .'       `.
                                 `.|.'                 +Zsc is out
                                                      of the page

Structure Design
The ROSETTA platform structure consisted of two modules, the Bus
Support Module and the Payload Support Module (BSM and PSM). Mounted
to the BSM was the Lander Interface Panel (LIP), which could be handled
separately for the Lander integration.

The spacecraft structural design was based on a version with a central
cylinder accommodating the two propellant tanks. The general
dimensions were dictated on one hand by the need to accommodate the
two large tanks, to provide sufficient mounting area for the payload
and subsystems and the Lander, as well as being able to accommodate
two large solar arrays, and on the other hand by the requirement to
fit within the Ariane 5 fairing.

The spine of the structure was the central tube, to which the
honeycomb panels were mounted. The spacecraft box was closed by lateral
panels, which were connected to the central tube by load carrying
vertical shear webs and an internal deck.

The Bus Support Module (BSM) accommodated most of the platform and
avionic equipment.

The Payload Support Module (PSM) was accommodating all science
equipment. The PSM structure consisted of the PSM +z-panel, the PSM -x
panel, the PSM +y/-y panels and the Lander Interface Panel (LIP).

Most instrument sensors were located on a single face, the +Z panel,
with the exception of VIRTIS and OSIRIS mounted on the -X panel to
allow for the accommodation of their cold radiators, Alice mounted on
PSM -X and COSIMA mounted on the PSM -Y panel. The P/L electronics
were mounted on the +Y and -Y side of this module for heat radiation
via Louvers.

Special supports were provided by the structure for:

Solar Array
They provided stiff and accurately positioned points for the solar
array hold down points and for solar arrays drive mechanisms.

Reaction Wheels
The brackets provided stiff wheel support with alignment capability.
All 4 RW brackets were mounted together between the +X shear wall and
the central deck building one compact bracket unit which provided
high stiffness and stability.

Propellant Tanks
The two tanks were mounted via a circumferential ring of flanges to a
reinforced adapter ring on the tube with titanium screws.

Helium Tanks
The two helium tanks were mounted on the main deck of the BSM. They
were attached by an equatorial fixation in the middle of the tank
through internal deck holes.

Thrusters on the side of the spacecraft were mounted on lateral panel
extensions with aluminium machined brackets ensuring the angular
position of the thrusters. Thrusters underneath the spacecraft (-Z
pointing thrusters) were mounted on brackets on the corners of the
+/-Y panels.

High Gain Antenna
The HGA was stowed against the +X panel, in areas stiffened by the
+/-Y panels and the HGA support tripod. After launch, the HGA was
deployed and was connected to the S/C by the support tripod only. The
axis Antenna Pointing Mechanisms, fixed on the tripod, were located
close to the edge of the HGA.

A single bracket provided stiff gyro support and alignment capability
and orientated the 3 IMUs in the requested angular orientation. The
bracket was mounted on the -Y BSM panel for thermal dissipation

Mechanisms Design

The ROSETTA mechanisms comprised the following major equipments:
* Solar Array Drive Mechanism (SADM)
* Solar Array Deployment Mechanisms
* HGA Antenna Pointing Mechanism (APM)
* HGA Holddown & Release Mechanism (HRM)
* Experiment Booms & HRMs
* Louvres (mechanical elements)

Solar Array Drive Mechanism (SADM)
The SADM performed the positioning of the Solar Array w.r.t. the Sun
by rotation of the panels around the spacecraft Y-axis. There were two
identical SADMs on both sides of the spacecraft, which could be
individually controlled. The control authority rested with the AOCMS
subsystem, which always 'knew' the actual attitude and Sun direction
and was therefore in the position to determine the required
orientation of the solar panels. The positioning commands were routed
from the AOCMS I/F Unit via the SADE (SADM-Electronics) to the SADM.

The Solar Array rotation was limited to plus and minus 180 degrees to
the reference position. The array zero position was defined in the
section 'Power Design: Solar Array Generator' below.

The Solar Array Drive Mechanism baseline design comprised the
following major components:
* Housing structure from aluminium alloy
* Main bearing, pre-loaded angular contact roller bearing
* Drive unit consisting of a redundantly wound stepper motor, gear-
  reduction unit, anti-backlash pinion, and final stage gear ring
* Redundant position transducer and electronics, harness and
* Mechanical end-stop for +/-180 deg travel limit with redundant
  micro-switches (4 in all)
* Redundant electrical power and signal harnesses, and connectors
* Twist capsule unit, allowing +/-180 deg electrical circuit transfer
* Thermistor for temperature reading, with harness.

The SADM drive unit employed a 'pancake' configuration with one single
X-type ballbearing to provide high moment stiffness and strength
within a compact axial envelope. The central output shaft was of
hollow construction, providing sufficient space to accommodate the
power and signal transfer harness and a twist capsule allowing +/-180
degrees rotation of the harness. The drive unit contained a position
transducer and a drive train.

The Solar Arrays Drive Electronic was intended to manage two Solar
Array Drives that could be rotated so as to get the maximum energy from
the solar cell panels.

Solar Array Deployment Mechanisms
The baseline were 2 solar arrays, each with a full silicon 5-panel
wing, with panel sizes as used in the ARA MK3 5-panel qualification
wing (about 5.3 m2 per panel).

During launch the wings were stowed against the sidewalls of the
satellite. They were kept in this position by means of 6 hold-down
mechanisms per wing.

Approximately 3 hours after launch, the satellite was pointed towards
the Sun and the wings were deployed to their fully deployed position.
They were released for full deployment by 'cutting' Kevlar restraint
cables by means of thermal knives (actually degrading of the Kevlar
by heat).

The deployment system made use of spring driven hinges and was
equipped with a damper, that limited the deployment speed of the wing.
Thus, the deployment shocks on SADM hinge and inter-panel hinges were
kept relatively low.

The Rosetta wing was further equipped with:
* ESD protection on front and rear side,
* Solar Array sun acquisition sensor,
* Solar Array performance strings

HGA Antenna Pointing Mechanism (APM)
The APM was a two-axes mechanism which allowed motion of the HGA in
both azimuth and elevation. The control authority rested with the
AOCMS subsystem, which always 'knew' the actual attitude and Earth
direction and was therefore in the position to determine the required
orientation of the antenna. The positioning commands were routed from
the AOCMS I/F Unit via the APM-E (APM-Electronics) to the APMM. HGA
elevation rotation was physically limited to +30deg/ -165deg from the
reference position (after deployment). Before and during deployment
the range was -207deg and +30deg.

HGA azimuth rotation was physically limited to +80deg / -260deg from
the reference position.

The main functions of the APM were:

* Allow accurate and stable pointing of the antenna dish through
controlled rotation about azimuth and elevation axes.
* Minimise stresses on the waveguides by acting as load transfer path
between the HGA and the spacecraft.

It consisted of three main components:
* The motor drive units (APM-M) and RF Ancillary Equipment (Rotary
* The support structure (APM-SS).
* The electronic control of these units (APM-E).

The APM-M was mounted between the antenna dish and the APM-SS.

For thermal reasons the elements of the APM-M and APM-SS and the
Antenna HDRMs were covered with MLI.

Experiment Boom Mechanisms
Two deployable experiment booms supported a number of different
lightweight sensors from the plasma package which needed to be deployed
clear of the S/C body. These booms were deployed at beginning of the
mission after Launch.

Each boom consisted of a 76 mm dia CFRP tube. The lower boom was
approximately 1.3 m long and the upper boom 2m.

The boom deployment was performed by means of a motor driven unit. The
deployment mechanism consisted of:

* Hinge, Motor Gear Unit, Coupling system, Latching system and
  Position switches.

The Hold down and release mechanisms, one per boom, had the following
* Three Titanium blades to allow relative displacement in the boom
  centreline direction. This reduced the mechanical and thermo-
  elastic I/F forces.
* The separation device was the Hi-Shear low shock Separation Nut

The Rosetta Thermal Control Subsystem contained 14 louvers with 2
different set points which were located on the S/C Y walls in front of
white painted radiators. The louvers were designed, manufactured and
qualified by SENER.

The mechanisms of the 16 blade louver were the 8 temperature dependent
bi-metal springs (actuators), which supplied the fundamental function
of the louver. The actuators were driving the louver blades to its end
stops for the defined fully open / fully closed temperature set

Thermal Control Design

Thermal Control Concept

The thermal control design was driven on one side by the low heater
power availability together with the low solar intensity in the cold
case, and on the other side by the hot cases characterised by high
dissipation of the operational units and high external heat loads.

The thermal control concept mainly utilised conventional passive
components supported by active units like heaters and controlled
radiative areas, using well proven methods and classical elements.

This concept could be characterised as follows :

* Heat flows from and to the external environment were minimised using
  high performance Multi-Layer Insulation (MLI).
* Most unit heat was rejected through dedicated white paint radiator,
  actively controlled by louvers, located on very low Sun-illuminated
  +/-Y panels.
* High internal emissivity compartments reduced structural temperature
* Individually controlled instruments and appendages (booms, antennas
  ,...) were mounted thermally decoupled from the structure.
* High temperature MLI was used in the vicinity of thrusters.
* Optimised heaters, dedicated to operational, and hibernation modes,
  were monitored and controlled to judiciously compensate the heat
  deficit during cold environment phases.

Thermal control design
The thermal control subsystem (TCS) design was optimised for the
enveloping design cases of the end of life comet operations and the
aphelion hibernation. From the overall mission point of view the deep
space hibernation heater power request was the most critical thermal
design case. This heater power request was dependent on the radiator
sizing which needed to be performed for worst case end of mission
conditions. The very strong heater power limitation implied that to
a certain extent constraints in the operation and/or attitude needed to
be accepted for hot case.

The TCS used a combination of selected surface finishes, heaters,
multi-layer insulation (MLI) and louvres to control the units in the
allowable temperature ranges. The units were mostly mounted on the
main +/- Y panels of the spacecraft (and +Z for experiments), with
interface fillers to enhance the conductive link to the panel for the
collectively controlled units. The individually controlled
experiments were thermally decoupled from the structure.

Generated heat by the collectively controlled units was then rejected
via conduction into the panel and subsequent radiation from the
external surface of the panel to space. These surfaces were covered
with louvers over white painted radiators minimising any absorbed
heat inputs and heat losses in cold mission phases. The louvers were
selected as baseline being the best solution (investigated during
phase B) for flexibility, qualification status and reliability.

VIRTIS and OSIRIS cameras were located at the top of the -X (anti-sun
face) so that their radiator may have viewed deep space. The top floor was
extended over the top as a sunshield to prevent any direct solar
illumination of these instruments, while the sun angle on the -Z side
had to be limited to 80 degrees for the same reason.

Any external structural surface not required as a radiator, (or
experiment aperture) was covered with a high performance MLI blanket.
The bottom of the bus module, which was not enclosed with a structural
panel, was covered with a high performance MLI blanket used also as an
EMC screen. In the areas around thrusters, a high temperature version
of the MLI were implemented. All blankets were adequately grounded and

The bi-propellant propulsion subsystem needed to be maintained between
0 to +45 degrees throughout the mission. This was far warmer than some
units, particularly when the spacecraft was in deep space hibernation
mode. The tanks and RCS were therefore well isolated from the rest of
the spacecraft to allow their specific thermal control.

The antennae and experiment booms were passively thermally controlled
by the use of appropriate thermo-optical surface finishes and MLI.
The mechanism for the HGA had similar appropriate passive control but
also needed heaters to prevent the mechanism from freezing. It was
thermally decoupled from the rest of the spacecraft to allow its
dedicated thermal control.

The chosen solution for thermal control subsystem design used well
known and proven technologies and concepts.

General Heater Control Concept
The operation of the TCS shall have enabled to maintain all spacecraft
units within the required temperature range throughout the entire
mission coping with all possible spacecraft orientations and unit
mode operations.

The thermal heater concept used the following major control features:

* Thermistor controlled (software) heater circuits, which were used to
maintain platform, avionics and payload units within operating limits
when these units were operating.

* The S/W heater design included 3 control thermistors sited next to
each other and used the middle temperature reading to control the
heater switching. This method was used in order to maximise the
reliability of thermistor controlling temperature.

* Thermistors was also used to monitor the temperature at each
unit's temperature reference point (TRP) and at the System Interface
Temperature Points (STP).

* Thermostat controlled (hardware) heater circuits, which were used to
maintain platform, avionics and payload units within their non-
operating (or switch-on) limits when these units were non-operating.
These operated autonomously during satellite hibernation and Safe
modes to ensure thermal control.

* The hardware heater circuits was controlled by one thermostat
(cold guard) connected in redundant circuit. The prime circuits
without any thermostat was powered as long as the relevant LCL was
defined to be enabled. In the prime circuit a thermostat (hot guard)
was included to prevent from overheating. In the event of a failure in
the prime circuit the redundant circuit was automatically switched on
when the temperature fell because it was permanently enabled.

* The lower set points for the thermostats (cold guard) were at the
lower nonoperating limits of units. The hysteresis of the thermostats
was chosen to 35 degrees Celsius to limit the number of switching
cycles for the long Rosetta mission. The higher set points of the
prime thermostats (hot guard) was oriented to the upper operational
temperature limit, but will still have an appropriate margin to that

* Main and redundant heaters were in separate foil heaters. It was
necessary to define reserved unpainted areas on all units, which
would nominally have been black painted, specifically for the mounting of

All software and hardware heaters circuits comprised a simple
series connection of heaters with no parallel connections. The heater
concept assumed prime and redundant heater elements in different
mats. The heaters were mounted directly onto units as this
maximises the efficiency of the heating.

The sizing of the autonomous H/W heater circuits were based upon the
following criteria:

* Payload heaters shall have been designed to maintain non-operating
temperature limits at 5.33AU or switch-on limits at 3.25AU,
whichever gave the greater heater power requirement,

* Platform and Avionics units OFF in hibernation had heaters
designed to maintain non-operating temperature limits at 5.33AU
or switch-on limits at 4.5AU, whichever was the greater power

* Platform and Avionics units ON during hibernation had heaters
designed to maintain operating temperature limits at 5.33 AU.

The suppliers of individually controlled (I/C) units shall have
sized their S/W and H/W heaters by themselves and may have installed them
where they wished in order to control their unit temperatures.

Micrometeoroid and Cometary Dust Protection
The micrometeoroid protection used for Rosetta was composed of 2
layers of betacloth and a spacer. This protection was only applied to
the exposed +Z and -Z central tube areas of the propellant tanks as
the spacecraft honeycomb structure would form an effective shield

The first betacloth layer was underneath the outermost layer of the
S/C MLI acting as a bumper. To reach the agreed probability of no
micrometeoroid impacts in 998 out of 1000 strikes, a separation of 50mm
to the second betacloth layer (on top of the tank MLI) was needed. The
micrometeoroid protection was part of the overall MLI design.

The cometary dust had a velocity similar to that of Rosetta and
so hypervelocity impacts were not an issue. Of more concern was the
coating of the spacecraft surfaces by the cometary dust. Grounding of
the external surfaces prevented differential charging but the whole
spacecraft may have been charged to some potential.

Propulsion Design
The propulsion subsystem was based on a pressure fed bipropellant type
using MMH (MonoMethylHydrazine) and NTO (Nitrogen TetrOxide). It was
capable to operate in both regulated and in blow-down mode and
provided a delta v of more than 2100 m/s plus attitude control. It was
able to operate in three axis and in spin stabilised mode (about the
x-axis) provided that the spin rate does not exceed 1 rpm. The
subsystem provided a high degree of redundancy in order to cope with
the special requirements of the ROSETTA mission.

The materials used in the propulsion subsystem were proven to be
compatible with the propellants and their vapours the wetted area
being mainly made of titanium or suitable stainless steel alloys.

The components and most of the pipework were installed on the
spacecraft -X panel by means of supporting brackets made of material
with low thermal conductance.

The subsystem had 24 10 N thruster for attitude and orbit control.
They were located such that they could provide pure forces and pure
torques to the spacecraft. The 24 thrusters were grouped in pairs on
the brackets, one of each pair being the main and one the redundant
thruster. The subsystem allowed the operation of 8 thrusters

The subsystem was maintained within the temperature limits of the
components. The mixture ratio may have been adjusted by tank temperature
(i.e. pressure) manipulation in order to enhance thruster

The propulsion subsystem was operated in regulated mode as well
as in blow down mode. The pressurisation strategy must have taken into
account various constraints as the available propellant, the minimum
inlet pressures for the thrusters, the maximum allowable pressures in
the propellant tanks etc. Calculations had been performed to
demonstrate the capability of the subsystem to fulfil the mission
requirements in terms of delta-v provision under the various
constraints and also with respect to the requirement for additional
20% fuel.

Telecommunication Design

The Tracking, Telemetry and Command (TT & C) communications with the
Earth over the complete Rosetta mission was ensured by three antenna
concepts, operating at various stages throughout the overall
programme, combined with a number of electrical units performing
certain functions. The Telecommunication Subsystem was required to
interface with the ESA ground segment in normal operational mode and
with the NASA Deep Space Network during emergency mode.

The TT & C subsystem comprised a number of equipment's whose
descriptions appear below:

* Two Transponders interfacing with the S-Band RF Distribution Unit
(RFDU), with the High Power Amplifiers - in this case Travelling Wave
Tube Amplifiers (TWTA's) -, and with the Data Management System
(DMS). The Transponders modulated and transmitted the Telemetry stream
coming from both parts of the redundant Data Management System either
in S or X-Band or both simultaneously without any interference and
transponded the ranging signal in S and X-Band. The Transponders
provided hot redundancy for the receiving functions and cold
redundancy for transmitting functions. The receivers could receive
telecommands in S-Band or X-Band (selectable per command), but not
simultaneously in both frequency bands. The configuration was such
that both receivers could receive, demodulate and send the telecommand
signal to the DMS simultaneously. The transmitters were also able to
receive the telemetry stream from both parts of the redundant DMS.
Each transponder was capable of operating in a coherent or non-
coherent mode depending on the lock status of the receiver.

* An RF Distribution Unit (RFDU) providing an S-Band transmitted/received
switching function between the antennas and the two Transponder units
via two diplexers.

* Two TWTA's providing >28W of power at X-Band to the MGA or HGA via
the Waveguide Interface Unit (WIU). The input to the TWTA HPA's was
supplied by the Transponder X-Band modulators via a 3dB passive

* A Waveguide Interface Unit (WIU) comprising of diplexers, two
transfer switches and high power isolators so that it was possible to
switch between antennas without turning off the TWTA.

* The transmit frequency (and receiver rest frequency) could also be
derived from an external Ultra Stable Oscillator (USO) on request by
Telecommand which may have been used any time during the mission. This USO
had a superior performance compared to the Transponder internal
oscillator such that it is used for one-way ranging as part of the
Radio Science Investigations (RSI).

* Two Low Gain Antennas (LGA) providing a quasi omni directional
coverage for any attitude of the satellite which may have been used for:

      a)the near earth mission phase at S-Band for uplink telecommand
        and downlink telemetry.

      b)the telecommand Up Link at S-Band during emergency and
        nominal communications over large ranges up to 6.25 AU.

* A 2.2m High Gain Antenna (HGA) providing the primary communication
for Uplink at S/X-band and Downlink at S/X-Band.

* Two Medium Gain Antennas (MGA) providing emergency Up and Downlink
default communication after sun pointing mode of the S/C was reached.
The S-Band MGA was realised as a flat patch antenna whereas the X-
Band MGA was a offset-type 0.31m reflector antenna. The MGAs also
performed some mission communications functions at various phases
throughout their lifetime due to their much larger coverage area.

High Gain Antenna Major Assembly
The transmission of the high rate scientific data of the ROSETTA
spacecraft to earth was depending reliable operation of the High Gain
Antenna major assembly, which was therefore a critical element for
the mission success. The most important requirements for this
assembly were:
  * High reliability
  * conform to specified pointing requirements
  * minimize mechanical disturbances
  * comply to antenna gain requirements

The HGA Major Assembly comprised:
  * HRM Hold-down and Release Mechanism for the HGA dish during
    launch with three release points
  * Two axes APM Antenna Pointing Mechanism (HGAPM) mounted on
    a tripoid to offset the antenna from the +X panel
  * A Cassegrain (X-Band) quasiparaboloid highgain Antenna (HGA)
    with a dichoric subreflector and S-band primary feed
  * Antenna Pointing Mechanism Electronics (APME)
  * Waveguide (WG) and Rotary Joints (RJ) for the RF transmission

High Gain Antenna Frame

The Rosetta High Gain Antenna was attached to the +X side of the s/c
bus by a gimbal providing two degrees of freedom and it articulates
during flight to track Earth. Therefore, the Rosetta HGA frame,
ROS_HGA, was defined with its orientation given relative to the

The ROS_HGA frame was defined as follows:
   -  +Z axis was in the antenna boresight direction;
   -  +X axis pointed from the gimbal toward the antenna dish
      symmetry axis;
   -  +Y axis completed the right hand frame;
   -  the origin of the frame was located at the geometric center of
      the HGA dish outer rim circle.

The rotation from the spacecraft frame to the HGA frame could be
constructed using gimbal angles from telemetry by first rotating
by elevation angle about +Y axis, then rotating by azimuth about
+Z axis, and then rotating by +90 degrees about +Y axis to finally
align +Z axis with the HGA boresight.

   This diagram illustrates the ROS_HGA frame:

   +X s/c side (HGA side) view:
                                   | toward comet

                               Science Deck
  .__  _______________.     |             |     .______________  ___.
  |  \ \               \    |             |    /               \ \  |
  |  / /                \   |  +Zsc       |   /                / /  |
  |  \ \                 `. |      ^      | .'                 \ \  |
  |  / /                 | o|      |      |o |                 / /  |
  |  \ \                 .' |      |      | `.                 \ \  |
  |  / /                /   |      |      |   \                / /  |
  .__\ \_______________/    |  +Xsc|      |    \_______________\ \__.
    -Y Solar Array          .______o-------> +Ysc   +Y Solar Array
                              .'       `.
                             /           \
                            .   `.   .'   .           +Zhga and HGA
                            |     `o-------> +Yhga    boresight are
                            .      |      .           out of the page
                             \     |     /
                              `.   |   .'
                           HGA  ` -|- '
                                   V +Xhga

Medium Gain Antenna (MGA)
The MGA design had been split into two physically separated antennae
  * the MGAS operating in -S-Band frequencies,
  * the MGAX operating in -X-Band frequencies,

MGA S-band (MGAS)
- - - - - - - - -
The antenna design for the S-Band subsystem consisted of an array of
patch antenna elements providing a circularly symmetrical radiation
pattern. The maximum gain obtainable for this array surface area
(300mm x 300mm) ranged between 14.1 and 14.7 dBi in the receive and
transmit frequency bandwidths.

The MGAS assembly could be sub-divided into two parts, the RF active
part (radiators plus distribution network) and the support structure
(platform plus stand-offs).

The array elements were arranged in a hexagonal lattice to provide the
required symmetry to the antenna pattern. Six elements were used to
meet the required specification.

MGA X-band (MGAX)
- - - - - - - - -
The configuration of the X-band MGA (MGAX) was a single offset
parabolic reflector illuminated by a circular polarised conical horn.
Reflector dimensions were selected to reach a desired minimum gain and
to lead to a simple feeder design. This led to an aperture diameter
of about 310mm and a focal length of 186mm (F/D = 0.6). With these
values a large reflector subtended angle was obtained which ensured
small feeder dimensions and a compact antenna design.

The MGAX antenna assembly was composed of two sub-assemblies, a
reflector and a feeder, and of a platform which supported both these
sub-assemblies and provided the interface to the Rosetta spacecraft.
The total envelope of the antenna was length=600mm, width=320mm,

The thermal protection for the antenna consisted of:
* White paint on the radiant face (PYROLAC 120 FD + P128)
* Thermal blankets on the rear face of reflector, feeder, supports
  and platform.

Low Gain Antenna (LGA)
Two classical S-band Low Gain Antennae (LGA) of a conical quadrifilar
helix antenna type were implemented on the satellite in opposite
direction to achieve an omnidirectional coverage. One was located at
the +Z-panel in the near of the edge to the +X panel and thus was
orientated towards the comet during the comet mission phase. The
other one was mounted on the opposite face.

Ultra Stable Oscillator
An Ultra Stable Oscillator was implemented within the TTC subsystem
providing the required frequency stability (Allan Variance, 3s,
2.0e-13 at 38.2808642 MHz) for the RSI instrument. This USO would be
used by the TTC subsystem whenever needed and was available for RSI
measurements as well. Should the USO failed, each transponder would use
its own oscillator (TCX0), but with less stability and not harming
the performance.

Power Design
The Power Subsystem (PSS) conditions, regulated and distributed all
the electrical power required by the spacecraft throughout all phases
of the mission. Distribution involved the switching and protection of
power lines to all users, including the Avionics units and the
Payload instruments, and includes equipment power, thermal power and
keep-alive-lines. The PSS also switched, protected and distributed
power for the pyrotechnics and the thermal knives of the various
release mechanisms of the spacecraft.

Main power source for Rosetta was provided by the Solar Array
Subsystem from silicon solar cells mounted on 2 identical solar array
wings, which were deployed from the +Y and -Y faces of the spacecraft
and could be rotated to track the sun. The solar cells on the outer
panel of each wing were outward facing when in the launch (stowed)
configuration in order to provide power input to the PSS for loads
and battery recharge following separation from the launcher and prior
to array deployment.

Batteries provided power for launch and post-separation support until
the solar arrays were fully deployed and sun aligned, and thereafter
would support the main power bus as necessary to supply peak loads and
special situations during Safe Mode where the sun might not have been fully
oriented towards the sun. One special feature of the power supply was
the Maximum Power Point Tracker (MPPT), which would operate the solar
array in its maximum power point in case of power shortage. During
almost all time of the mission, except for short periods of peak
power demands, the PCU would operate in nominal mode, i.e. the PCU
took only the power required by the satellite from the solar array.
The delta power would remain in the solar array. Because of this
feature the actual performance of the array could only be assessed by
utilising 'performance strings' which operated some cells in short
circuit current mode and others in open circuit voltage mode. From
the data obtained from these cells the performance of the solar
generator could be determined.

Batteries were also the main power source for the pyrotechnics,
although pyrotechnic power was also available from the main bus as a
back-up in case there was no battery power.

The subsystem was designed in accordance to the ESA Power Standard

Power Conditioning Unit (PCU)
* Produced a fully regulated 28V single power bus from solar array
  and battery inputs.
* Main bus voltage control including triple redundant error
* Separate hot redundant array power regulators for each array wing.
* Separate hot redundant Maximum Power Point Trackers (MPPT) for
  each array wing
* Separate Battery Discharge Regulator (BDR) for each battery.
* Separate Battery Charge Regulator (BCR) for each battery.
* Array performance monitor.
* TM/TC interface.
* Some automatic functions to support power bus management.

Payload Power Distribution Unit (PL-PDU)
* Dedicated to payload power distribution.
* Fully redundant unit.
* Main bus power outlets were all switched and protected by Latching
  Current Limiters (LCL).
* LCLs had current measurement and input under-voltage protection.
* 7 LCL power rating classes covering 5.5W to 135W (nominal load
* Provision of Keep Alive Lines (KALs) for experiments.
* Pyrotechnic power protection and distribution, including firing
  current measurement and storage.
* Distributed power to the Thermal Control Subsystem hardware and
  software controlled heaters.
* Individual on/off switching for each software controlled heater
* TM/TC interface.

Subsystems Power Distribution Unit (SS-PDU)
* Dedicated to Platform and Avionics power distribution.
* Fully redundant unit.
* Fold-back Current Limiters (FCL) for non-switchable loads
  (Receivers and CDMUs).
* All other main bus power outlets were switched and protected by
  Latching Current Limiters (LCL).
* FCLs and LCLs had current measurement and FCLs had input under-
  voltage protection.
* LCL classes and power ratings as for PL-PDU.
* Pyrotechnic power protection and distribution, including firing
  current measurement and storage.
* Thermal Knives (TKs) power distribution (for Solar Array panels
* Distributes power to the Thermal Control Subsystem combined
  hardware -  software controlled heaters.
* Individual on/off switching for each software controlled heater
* TM/TC interface.

* 3 batteries each comprising 6 series and 11 parallel connected Li-
  Ion 1.5 Ah cells (corresponds to 16.5 Ah per battery).
* Power and monitoring connections to PCU.
* Power connections also to the PDUs for the pyrotechnics.
* Cells arrangement and wiring to minimise magnetic moment.
* 1 thermistors per battery for battery charge/discharge control.
* A combination of relay/heater mat in order to discharge the
  batteries for capacitance verification.

Solar Array Generator
The orbit of the S/C had an extremely wide variation of Spacecraft-
Earth-Sun angles and distances, hence it was mandatory to include an
electrical design based on LILT (Low Intensity Low Temperature) solar
cell technology.

The structural parts/units (deployment system, substrates, hold-down
& release system) were identical to the qualified ARA MK3 design of
Fokker Space.

The geometry and mechanical interface definition of the Rosetta
baseline Solar Array design was identical to the 5-panel qualification

The electrical architecture (cells, strings, sections & harness lay-
out) was uniquely designed for Rosetta. Electro static discharge (ESD)
protection design was qualified for the ARA MK3 type solar array.

The baseline were 2 solar arrays, each with a full silicon 5-panel
wing, with panel sizes as used in the ARA MK3 5-panel qualification
wing (about 5.3 m2 per panel).

   x---.---.---.---.---x  |       |  x---.---.---.---.---x
   |   |   |   |   |   |--|   x   |--|   |   |   |   |   |
   x---'---'---'---'---x  |       |  x---'---'---'---'---x

Mechanical Design of the Solar Panels
The basic skin design of the panels of the solar arrays consists of
two layers [0/90degres] M55J/950-1 CFRP prepreg (thickness per layer
0.06 mm) in closed lay-up. The panel substrate dimensions were 2.25 x
2.736 m2. The front side skin would use a 50^m Kapton foil to isolate
the solar cell network from the conductive CFRP layers. The Kapton
foil was co-cured with the CFRP layers.

The panel core consisted of Aluminium honeycomb with a core height of
22 mm. Local circular reinforcement plugs ('subassembly panels') were
used to provide the holddown areas with extra strength, stiffness and
fatigue resistance.

The hold-down and release system used a tie-down element (Kevlar
cable) under high preload which would be degraded by heat of the
thermal knife for release. The hold-down, SADM and yoke snubber
locations for Rosetta were fully identical to the ARA MK3
qualification hardware definition.

The stowed wing had a height of <239 mm at the wing tips (the gap
between inner panel and sidewall was increased from nominal 70 mm by
about 30mm by means of a dedicated bracket, the inter panel gap was 12
mm, and the panel substrate thickness was 22 mm).

The deployment mechanism concept relied on spring-driven hinges. The
spring characteristics were chosen such that the energy supply was
enough for the full range up to 5 maximum sized panels, while
maintaining the required deployment safety factors. In order to
reduce the shock loads on the SADM and inter-panel hinges, a damper
was introduced in the deployment system.

A stiff synchronisation system was applied to prevent a very non-
synchronous deployment, resulting in unpredictable high deployment
latch-up shocks at the interpanel hinges.

The V-yoke length was 1103 mm when measured from SADM hinge-line to
yoke/inner panel hinge-line. The yoke length used within the ARAFOM
5-panel QM wing programme was identical.

The arms of the V-shaped yoke consisted of M46J CFRP filament wound
with a circular cross section (inner diameter 43 mm; nominal wall
thickness 0.9 mm) with reinforcements at the ends of the yoke tubes.

Rosetta Solar Array Frames
The Rosetta solar arrays could be articulated (each having one degree
of freedom), the solar Array frames, ROS_SA+Y and ROS_SA-Y, were
defined with their orientation given relative to the ROS_SPACECRAFT

Both array frames were defined as follows :

      -  +Y axis was parallel to the longest side of the array,
         positively oriented from the end of the wing toward the

      -  +Z axis was normal to the solar array plane, the solar cells
         on the +Z side;

      -  +X axis was defined such that (X,Y,Z) is right handed;

      -  the origin of the frame was located at the geometric center
         of the gimbal.

The axis of rotation was parallel to the Y axis of the spacecraft and
solar array frames.

At zero (reference) position the array wing was aligned such that the
array surface was in the spacecraft Y-Z plane, with the face (cells)
aligned such that the array normal was parallel to the +X axis of the
spacecraft. This means that in stowed configuration (i.e. launch
configuration) the array position of the array on the +Y panel was -90
degrees and on the -Y panel +90 degrees.

This diagram illustrates the ROS_SA+Y and ROS_SA-Y frames:

+X s/c side (HGA side) view:
                                    | toward comet

                               Science Deck  +Xsa+y0
   .__  _______________.     |             ||    .______________  ___.
   |  \ \               \    |             ||   /               \ \  |
   |  / /                \   |  +Zsc       ||  /                / /  |
   |  \ \                 `. |      ^      ||.+Zsa+y0           \ \  |
   |  / /           +Zsa-y0 o-----> | <-----o  Zsa+y            / /  |
   |  \ \           +Zsa-y.'|+Ysa-y0|+Ysa+y0 `.                 \ \  |
   |  / /                /  ||+Ysa-y|+Ysa+y|   \                / /  |
   .__\ \_______________/   ||      |      |    \_______________\ \__.
     -Y Solar Array         |.______o-------> +Ysc   +Y Solar Array
                            v  +Xsc o__.
                     +Xsa-y0   .'       `.
                     +Xsa-y   /           \
                             .   `.   .'   . +Zsa+y0, +Zsa+y, +Zsa-y0,
                             |     `o'     | and +Zsa-y are out of
                             .      |      .       the page
                              \     |     /
                               `.       .'   Active solar cell is
                            HGA  ` --- '      facing the viewer

Power Constraints in Deep Space

In the phases with Sun distances above approximately 4.0 AU the
decreasing solar array power forced the use of economical strategies
for certain operations. Thereby the situation after the deep space
hibernation phase was much more severe. From radiation degradation
analysis it had been derived that after DSHM at 4.5 AU about 65 W
less solar array power would be available compared to 4.5 AU before
DSHM. This corresponded to about 13% of the power needed at that

In the deep space phases the general operational concept was the

  * minimise the overall power consumption by switching off all
  equipment not directly needed during the current operation

  * additionally, for certain operations with high extra power
  demand, perform a power sharing strategy by switching off some TCS
  heaters; as a consequence this puts a time limit on such operations

  * operate equipment like RWs and SSMM in reduced power mode

  * for autonomous operations, which were not directly under ground
  control, like in Safe Mode, the ground could set a Low Power Flag as
  invocation parameter in the call of the Safe Mode OBCP (which was
  loaded in the System Init Table) at the appropriate time in the
  mission, according to the current Sun distance. This flag would be
  checked by the OBCP; if the flag was set, the Safe Mode downlink
  would be performed in power sharing strategy and the SSMM was set
  into stand-by mode (memory modules remained powered, but memory
  controllers were switched off).

As a safety precaution the battery discharge alarm remained
enabled all the time. This would allow for nominal short (< 4 min)
peak power demands to be satisfied by the batteries, e.g. for RW
offloading, but would trigger a system alarm and transition to Safe
Mode in case of a creeping battery discharge due to a wrong power
configuration e.g. because of a missed command. If for such a case a
processor reconfiguration was not desired, it was possible to use the
monitoring of the MEA Voltage to trigger transition into Safe Mode
before the battery discharge alarm triggers (see Handling of On-board
Monitoring, [RO-DSS-TN-1155]).

Harness Design
The harness performed the electrical connection between all
electrical and electronic equipment in the ROSETTA spacecraft. It
provided distribution and separation of power supplies, signals,
scientific data lines, pyrotechnic firing pulses, and all connections
to the umbilical, safe/arm brackets/connectors and test connectors.

The harness consisted of the following subassemblies:
* Payload Support Module Harness
* Bus Support Module Harness
* Harness to the Lander I/F
Furthermore the harness / cables were divided into three harness EMC
classes: power, signal and data, and the pyro harness. Their routing
was physically separated. In addition to the appropriate twisting and
shielding techniques this minimised the probability of electrical
cross talking of critical lines.

The harness design followed a distributed single point grounding
scheme. Redundant functions had their own connectors and were routed
in separate bundles and in a different way as far as practical.

All connectors supplying power had female contacts.

To achieve a complete Faraday cage around the harness each of the
harnesses had its own overall shield made of aluminium tape with an
overlap of at least 50 % for harnesses within the spacecraft and a
double shield for harnesses outside the spacecraft. As fixation
points for the harness aluminium bases (Ty-bases) were bonded to the
structure with a two component conductive glue. The distance of the
Ty-bases was selected such that the harness withstands all specified
environmental conditions.

To avoid interruptions of the shield between the connector and the
overall shield, redundant connection wires were used between connector
case and harness overall shield. In case of pyro-lines and sensible
interfaces conductive connector boots were implemented.

To prevent contamination the harness were baked-out in a thermal
vacuum chamber prior to integration.

Avionics Design

The ROSETTA Avionics consisted of the Data Management Subsystem (DMS)
and the Attitude and Orbit Control and Measurement Subsystem (AOCMS)

Data Management Subsystem (DMS)
The data management subsystem was in charge of telecommand
distribution to other spacecraft subsystems and payload, of
telemetry data collection from spacecraft subsystems and payload and
formatting, and of 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 different Avionics
processors and the SSMM, STR and CAM. The OBDH bus was the data route
for data acquisition and commands distribution via the RTUs. Payload
Instruments were accessed via a dedicated Payload RTU. Subsystems were
accessed via a dedicated Subsystem RTU.

DMS included 4 identical Processor Modules (PM) located in 2 CDMUs.
Any of the processor modules could perform either the DMS or the AOCMS
processing. The PM selected for the DMS function acted as the bus
master. It was also in charge of Platform subsystem management (TTC,
Power, Thermal). The one selected as the AOCMS computer was in charge
of all sensors, actuators, HGA & SA drive electronics. TCdecoder and
Transfer Frame Generator (TFG) were included in each CDMU.

Telemetry could be downlinked via the TFG using the real time channel
(VC0) or in form of retrievals from the SSMM (VC1).

Solid State Mass Memory (SSMM)
- - - - - - - - - - - - - - - -
The Solid State Mass Memory (SSMM) was used like a 'Hard Disk Storage'
including 25 Gbit of memory. It contained a data compression module
which allowed lossy (for CAM image) and loss-less (for HK and science
data) compression of data to be stored. It was able of file management
capability. It stored CAM images, science and telemetry packets as
well as software data for the AOCMS and DMS computer.

It was coupled to:
* the 4 processors via an IEEE 1355 link,
* the TFGs of the 2 CDMUs via a serial link,
* VIRTIS, OSIRIS and the Navigation Camera via a high data rate
serial link (IEEE 1355)
* the High Power Command Module (HPCM) selecting the valid PM

The lossy compression method (WAVELET) was used for image data
compression of the NAVCAM or STR. The degree of compression could be
set by filter parameters from ground. The compression of OSIRIS and
VIRTIS image data could also be performed inside the SSMM. However
these two instruments did not request data compression from the system.

The SSMM SW run on a Digital Signal Processor. The SSMM SW was made

* The Init Mode Software
The Init mode software ensured the boot up of the SSMM and the
establishment of the communication with the DMS SW. It allowed the
loading of the operational SW from EEPROM to RAM, and its starting.

* The Operational Software
The operational SW managed the files located in the Memory Modules of
SSMM, and the Data Compression Function that performed Rice lossless
and Wavelet lossy data compression.

The functionality of the SSMM could be summarised with the three points
* Store on-board data in files. The on-board data could be both
scientific data and software images in files.
* Transmit the data stored in SSMM files to either an on-board User
or to the ground.
* Compress the stored files using both lossy and lossless compression

The Rosetta Solid State Mass Memory (SSMM) functionally consisted of
the following modules:
* 2 Memory Controllers (MC)
* 3 Memory Modules (MM)
* 2 Power Converters, which supplied power to the memory controller
and memory module boards.

The Memory Controllers were responsible for all data transfer to and
from the Mass Memory, compression of data in the mass memory and
basic housekeeping functions (collection of telemetry packets,
configuration of the SSMM etc.). The Memory Controllers worked in cold

The three Memory Modules were where the files are stored. The modules
could be turned on and off independently, giving the possibility to
increase and decrease the storage capacity of the SSMM. The Memory
Controllers accessed the Memory Modules via a memory module bus. Both
the Memory Controllers could access all three Memory Modules.

Attitude and Orbit Control Measurement System (AOCMS)

The AOCMS was in charge of attitude and orbit measurement and control
and was in charge with sensors and actuators for autonomous attitude
determination and control as well as pre-programmed manoeuvring.

The AOCMS used a decentralised architecture built around the AOCMS
Interface Unit (AIU) linked to all sensors / actuators and to the
Processor Modules included in the CDMUs:

* the AOCMS sensors: 2 Navigation Cameras (CAM) and 2 Star Trackers
(STR) having a common electronics unit, 4 Sun Acquisition Sensors
(SAS) and 3 Inertial Measurement Packages (3 IMP, each including 3
gyros + 3 acceleros),

* the AOCMS actuators: the Reaction Wheel Assembly (RWA), and
belonging to the Platform the Reaction Control System (RCS), the High
Gain Antenna Pointing Mechanism (HGAPM), and the 2 Solar Array Drive
Mechanisms (SADM).

AOCMS PM communication with AOCMS sensors (IMP, SAS, STR, CAM) and
actuators (RWA, RCS), and with pointing mechanism electronics
(SADE and HGAPE) was performed through the AIU. Functional AOCMS data
which needed to be put in the Telemetry and sent to the ground were
given packetised by the AOCMS processor and sent to the DMS processor
for further downlink to ground and storage in the SSMM.

The DMS PM permanently checked the AOCMS health by monitoring that the
AOCMS PM did not stop to communicate with DMS PM. This was done by
checking the correct reception of the so-called 'essential' AOCMS HK
packet every one second.

The AIU was the central data acquisition and distribution unit which
allowed access to the sensors and actuators with different type of
interfaces. It included RS 422, IEEE 1355 and MACS Bus interfaces as
well as analog and discrete digital interfaces for commanding and
data acquisition.

The AIU included furthermore a 12 bit A/D converter in order to
convert analog signals from the pressure transducers (temperature and
pressure) precise enough for the fuel level prediction on-board of
Rosetta late in the mission, when the fuel level was critical.

The major AOCMS components were the following:
 * AOCMS Interface Unit (AIU): it interfaced to all AOCMS sensors and

* The Sun Acquisition Sensors (SAS): they were internally redundant
and were used for Sun Acquisition and pointing. They provided full sky
coverage and ensured a permanent sensing of the Sun direction vector.

* The Inertial Measurement Packages (IMP): The IMP function provided
roll rate and velocity measurements along 3 orthogonal axes.

* 4 Reaction Wheels: they were arranged in the Reaction Wheel Assembly
(RWA) and the Reaction Control System (RCS), in a tetrahedral
configuration about the S/C Y-axis in order to enhance the torque and
momentum capacity about that axis for the asteroid fly-by.

* 2 Autonomous Star Trackers: they contained an Autonomous Star Pattern
Recognition function and provided autonomously to the AOCMS an
estimated attitude quaternion and stellar measurements data.

* 2 Navigation Cameras (A&B) were used in the AOCMS control loop
during the Asteroid Near Fly-by Phase. The navigation cameras could
also directly send image data to the SSMM through a high data rate

* Pointing mechanisms (through target pointing angles) and propulsion
thruster valves were commanded by the AOCMS through the AIU links.

Avionics external interface

The Avionics system had the following external interface to other
subsystems of the Rosetta spacecraft:

* Interface with the Ground through TTC Subsystem:
  Ground Telecommands (TC) were checked, decoded and executed
  internally or sent to other subsystems, Telemetry (TM) data
  generated on-board are collected, formatted (if needed) and sent to
  Ground through TTC S/S, either in real time or in play-back after
  storage in SSMM, on ground request.

* Interface with Platform and Payload:
  The Avionics provided the experiments and Platform equipment with a
  hardware command capability (power On/Off commands, heater On/Off

  The Avionics provided experiments with a time synchronisation
  capability, so that the Ground could later on correlate results
  coming from different experiments,

  The Avionics used for attitude and communication control purpose as
  well as for power generation Platform equipment: Reaction Control
  System (RCS), High Gain Antenna and Solar Array Pointing Mechanisms

  Housekeeping data and experiment science data were collected
  on-board to be sent to Ground in real time TM, or to be stored for
  play-back downlink,

  The Avionics S/W provided experiments and Platform with a
  processing capability, in form of application programs (AP) or
  On-board Control Procedures (OBCP), coded and implemented by the
  Avionics/OBCP contractor, but specified by the users to allow
  montoring/surveillance, thermal control, experiment or mechanism

Avionics modes

The Avionics modes derived from the AOCMS modes were the following:

Stand-By Mode
The SBM was used in Pre-launch and Launch Modes for general check
supervision. Only DMS functions were activated. It was possible to
command thrusters through AIU for RCS Priming.

Sun Acquisition Mode
This mode was used during Separation Sequence to perform rate
reduction (if necessary), Sun acquisition and Sun pointing. SAM was
also used as second level back-up mode to recover Sun pointing
attitude in case of an unsuccessful back-up to Sun Keeping Mode.

Safe/Hold Mode
The SHM followed the Sun Acquisition Mode / Sun Keeping Mode to
achieve a 3-axis attitude based on star trackers, gyros and reaction
wheels, with solar arrays pointing towards the Sun and Medium and
High Gain Antennae (i.e. S/C Xaxis) pointing towards the Earth and
the Y-axis normally pointing to the north of the ecliptic plane.

In some mission phases (i.e. defined by the minimum earth distance),
S/C X-axis pointing towards the Earth was forbidden because of thermal
constraints. Then, +X axis was pointed towards the Sun, and the High
Gain Antenna was pointed towards the Earth.

Normal Mode
The NM was used in Active Cruise and Near Comet phases for nominal
longterm operations, for comet observation and SSP delivery. Reaction
wheel off-loading was a function of the Normal Mode.

Thruster Transition Mode
The TTM was used for transition from Normal Mode to operational
thruster Modes, and vice-versa, for control tranquillisation.

Orbit Control Mode
The OCM was used in Active Cruise Mode for trajectory and orbit

Asteroid Fly-By Mode
The AFB mode was dedicated to asteroid observation.

Near Sun Hibernation Mode
The NSHM was a 3-axis controlled mode (with the attitude estimation
based on the use of STR only, and no gyro), with a dedicated thruster
control (i.e. single sided) to minimise the fuel consumption.

Spin-up Mode
The SpM was necessary to spin up the spacecraft at hibernation entry
(spin down at hibernation exit was achieved by Sun Keeping Mode). The
attitude control concept was a completely passive inertial spin during
the deep space hibernation phase.

There was no AOCMS Deep Space Hibernation Mode.

Sun Keeping Mode
The Sun Keeping Mode was used nominally at wake-up after Deep Space
hibernation, and as first level back-up mode to recover Sun pointing
attitude in case of a failure involving the Avionics and for which a
local reconfiguration on redundant units was not efficient. In case
the autonomous entry to Safe / Hold Mode was disabled or not
successful Earth Strobing Mode was established leading to a slow spin
motion around the Sun direction. Then the + X-axis was pointed towards
the expected earth direction (i.e. using the actual Sun/spacecraft/
Earth angle). The rotation along the Sun line was maintained therefore
the Earth crosses once per revolution the + X-axis which would allow
communication with the MGA.

System Level Modes

A basic configuration of the system level modes is given below:

Pre-launch      only DMS on, AOCMS PM on, external power supply

Launch Mode     Initially: DMS on, SSMM in standby with 1 MM,
                AOCMS PM on, separation sequence program running,
                power supply from batteries Finally: DMS on, AOCMS
                in Sun Acquisition Mode, TTC S-band downlink on,
                power supply from solar arrays, X-axis and solar
                arrays Sun pointing.

Activation      DMS on, AOCMS in Normal Mode, TTC S- or X-band
Mode            downlink via HGA (initially in S-band via LGA),
                3-axis stabilised, SA Sun pointing attitude

Active Cruise   DMS on, AOCMS in Normal Mode or Orbit Control
Mode            Mode, TTC S- or X-band downlink via HGA, 3-axis
                stabilised, SA Sun pointing attitude

Deep Space      CDMU on, AOCMS in SBM mode, inertial spin
Hibernation     stabilisation mode, wake-up timers on, thermostat
Mode            control of heaters

Near Sun        DMS on, AOCMS in NSHM, 3-axis active control mode
Hibernation     with 2 PMs, star tracker, thrusters, X-axis Sun or
Mode            Earth pointing

Asteroid        DMS on, TTC X-band downlink via HGA, SA Sun
Fly-by Mode     pointing, payload on, AOCMS in AFM mode: closed loop
                asteroid tracking with navigation camera, during Near
                Fly-by: HGA tracking stopped

Near Comet      DMS on, TTC X-band downlink via HGA, navigation
Mode            camera and payload on, AOCMS in Normal Mode: 3-axis
                stabilised, SA Sun pointing, instruments comet

Safe Mode       DMS on, AOCMS in Safe/Hold Mode; SA Sun pointing, X-
                axis Sun or Earth pointing, 3-axis stabilised using
                gyros, star tracker, RWs(if enabled by ground); TTC
                S-Band downlink via HGA; RXs on HGA/LGA; payload off

Survival Mode   DMS on, AOCMS in SKM submode 'MGA Strobing' (or in
                SKM if this submode is disabled), SA Sun pointing
                with offset from +X-axis = SSCE angle, fixed small
                residual rate around Sun vector; control by
                thrusters, Sun sensors, gyros; S-Band carrier
                downlink via MGA, RXs on MGA/LGA, load off

Ground Station Network

The Ground Station and Communications Network was performing telemetry,
telecommand and tracking operations within the S/X-band frequencies.
Telecommand was either in the S-band or X-band, and also telemetry was
switchable between S- and X-band, with the possibility to transmit
simultaneously in both frequency bands, only one of which was modulated
 (S-band downlink was primarily used during the near Earth mission
phases). The ground station used throughout all mission phases was the ESA
New Norcia (NNO 35m) deep space terminal (complemented by the ESA Kourou 15m
station during near-Earth mission phases and by the Cebreros and Malargue 35m
deep-space antennas during early comet phases up to Lander delivery). In
addition, the NASA Deep Space Network (DSN) 34m and/or 70m network was
envisaged for data downlink, back-up, and emergency cases.
The table below summarises the Ground Station Network usage.
 Ground Station  | Mission Phase Usage     | Frequency Utilisation |
 Kourou 15m      | Launch and LEOP         | Sband Uplink/Downlink |
                 |                         | Xband Uplink/Downlink |
 NNO 35m         | Launch, commissioning   | Sband Uplink/Downlink |
                 |                         | Xband Uplink/Downlink |
 Cebreros and/or | Comet approach, mapping | Xband Uplink/Downlink |
  Malargue 35m   |                         |                       |
 NASA/DSN        | Prime support for       | Sband Uplink/Downlink |
                 | critical phases and     | Xband Uplink/Downlink |
                 | Back up during inter-   |                       |
                 | planetary phases        |                       |

The information is extracted from the Rosetta Mission Implementation Plan -
[RO-ESC-PL-5100] and more details can be found in this document.

For more acronyms refer to Rosetta Project Glossary [RO-EST-LI-5012]

AFB     Asteroid Fly-By
AFM     Asteroid Fly-by Mode
AIU     AOCMS Interface Unit
AOCMS   Attitude and Orbit Control Measurement System
AOCS    Attitude and Orbit Control System
AP      Application Programs
APM     Antenna Pointing Mechanism
APME    APM Electronics
APM-M   APM Motor
APM-SS  APM Support Structure
ARA     Attitude Reference Assembly
AU      Astronomical Unit
BCR     Battery Charge Regulator
BDR     Battery Discharge Regulator
BSM     Bus Support Module
CAM     Navigation Camera
CAP     Comet Acquisition Point
CAT     Close Approach Trajectory
CDMU    Control and Data Management Unit
CFRP    Carbon Fibre Reinforced Plastic
CNES    Centre National d'Etudes Spatiales
COP     Close Observation Phase
DDOR    Delta Differential One-way Range
DLR     German Aerospace Center
DMS     Data Management Subsystem
DSHM    Deep Space Hibernation Mode
DSM     Deep Space Manouver
DSN     Deep Space Network
EEPROM  Electronically Erasable Programmable Read-Only Memory
EMC     Electromagnetic Compatibility
ESA     European Space Agency
ESD     Electro Static Discharge
ESOC    European Space Operations Center
ESTEC   European Space Research and Technology Center
EUV     Extreme UltraViolet
FAT     Far approach trajectory
FCL     Fold-back Current Limiters
FDIR    Failure Detection Isolation and Recovery
F/D     Focal Diameter
FOV     Field Of View
FUV     Far UltraViolet
GCMS    Gas Chromatography / Mass Spectrometry
GMP     Global Mapping Phase
HDRM    Hold-Down and Release Mechanism
HGA     High Gain Antenna
HGAPE   High Gain Antenna Pointing Electronics
HGAPM   High Gain Antenna Pointing Mechanism
HgCdTe  Mercury Cadmium Telluride
HIGH    High Activity Phase (Escort Phase)
HPA     High Power Amplifier
HPCM    High Power Command Module
HK      HouseKeeping
I/C     Individually Controlled
I/F     InterFace
IMP     Inertial Measurement Packages
IRAS    InfraRed Astronomical Satellite
IRFPA   InfraRed Focal Plane Array
IS      Infrared Spectrometer
HRM     HGA Holddown & Release Mechanism
H/W     Hard/Ware
KAL     Keep Alive Lines
LCC      Lander Control Center
LCL     Latching Current Limiters
LEOP    Launch and Early Orbit Phase
LGA     Low Gain Antenna
LILT    Low Intensity Low Temperature
LIP     Lander Interface Panel
LOW     Low Activity Phase (Escort Phase)
MACS    Modular Attitude Control System
MEA     Main Electronics Assembly
MC      Memory Controller
MGA     Medium Gain Antenna
MGAS    MGA S-band
MGAX    MGA X-band
MINC    Moderate Increase Phase (Escort Phase)
MLI     Multi Layer Insulation
MM      Memory Module
MMH     MonoMethylHydrazine
MPPT    Maximum Power Point Trackers
MS      Microscope
NM      Normal Mode
NNO     New Norcia ground station
NSHM    Near Sun Hibernation Mode
NTO     Nitrogen TetrOxide
OBCP    On-Board Control Procedures
OBDH    On-Board Data Handling
OCM     Orbit Control Mode
OIP     Orbit Insertion Point
PCU     Power Conditioning Unit
PDU     Power Distribution Unit
PI      Principal Investigator
P/L     PayLoad
PL-PDU  Payload Power Distribution Unit
PM      Processor Module
PSM     Payload Support Module
PSS     Power SubSystem
RAM     Random Access Memory
RCS     Reaction Control System
RF      Radio Frequency
RFDU    RF Distribution Unit
RJ      Rotary Joints
RMOC    Rosetta Mission Operations Center
RL      Rosetta Lander
RLGS    Rosetta Lander Ground Segment
RO      Rosetta Orbiter
RSI     Radio Science Investigations
RSOC    Rosetta Science Operations CenterRTU
RVM     Rendez-vous Manouver
RW      Reaction Wheel
RWA     Reaction Wheel Assembly
SA      Solar Array
SADE    Solar Array Drive Electronics
SADM    Solar Array Drive Mechanism
SAM     Sun Acquisition Mode
SAS     Sun Acquisition Sensors
SBM     Stand-By Mode
SHM     Safe/Hold Mode
SAS     Sun Acquisition Sensor
S/C     SpaceCraft
SI      Silicon
SINC    Sharp Increase Phase (Escort Phase)
STP     System Interface Temperature Points
SKM     Sun Keeping Mode
SONC    Science Operations and Navigation Center
SpM     Spin-up Mode
S/S     SubSystem
SSMM    Solid State Mass Memory
SSP     Surface Science Package
SS-PDU  Subsystems Power Distribution Unit
STR     Star TRacker
S/W     SoftWare
SWT     Science Working Team
TC      Telecommand
TC      Telecommunications
TCS     Thermal Control Subsystem
TFG     Transfer Frame Generator
TGM     Transition to global mapping
TK      Thermal Knives
TM      Telemetry
TRP     Temperature Reference Point
TTC    Tracking, Telemetry and Command
TTM     Thruster Transition Mode
TWTL    Two Way Travelling Lighttime
TWTA    Travelling Wave Tube Amplifiers
USO     Ultra Stable Oscillator
VC      Virtual Channel
WG      WaveGuide
WIU     Waveguide Interface Unit