PDS_VERSION_ID            = PDS3                                              
LABEL_REVISION_NOTE       = "2004-09-23 KW: Initial draft.                    
                             2005-12-09 AC: Orbiter Information               
                                            Updated Added Inst_host           
                                            for lander References TBD         
                             2006-01-10 AC: Removed special                   
                                            characters                        
                             2006-02-15 PG: Added Inst_host for               
                                            lander                            
                             2007-01-26 MB: 70 char line length               
                             2007-08-14 MB: remove not ascii symbols          
                             2008-02-02 Maud Barthelemy                       
                             2008-04-11 JL Vazquez, SA                        
                             2008-05-09, MB                                   
                             2010-02-15, MB                                   
                             2011-06-07, MB, editorial                        
                             2012-06-06, M. Barthelemy after AST2             
                                         review;                              
                             2017-04-26, M. Barthelemy missing                
                                         reference and Lander updates.        
                             2017-09-27, Maud Barthelemy, Ground Station      
                                         Network updates                      
                             2017-11-17 Maud Barthelemy typo corrections      
                             2018-07-17 Maud Barthelemy typo corrections      
                             2018-08-30 DF: shall heve remained -> remained   
                                         "                                    
                                                                              
RECORD_TYPE               = STREAM                                            
                                                                              
OBJECT                    = INSTRUMENT_HOST                                   
 INSTRUMENT_HOST_ID       = RO                                                
                                                                              
 OBJECT                   = INSTRUMENT_HOST_INFORMATION                       
  INSTRUMENT_HOST_NAME    = "ROSETTA-ORBITER"                                 
  INSTRUMENT_HOST_TYPE    = SPACECRAFT                                        
  INSTRUMENT_HOST_DESC    = "                                                 
                                                                              
                                                                              
TABLE OF CONTENTS                                                             
----------------------------------                                            
= 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                                                                     
  - NASA DSN                                                                  
= 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                           
telecommunication.                                                            
                                                                              
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           
failure.                                                                      
                                                                              
                                                                              
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               
switching.                                                                    
                                                                              
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            
         point.                                                               
                                                                              
   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                                                                     
---------                                                                     
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.                                                 
                                                                              
Gyros                                                                         
-----                                                                         
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               
reasons.                                                                      
                                                                              
                                                                              
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                  
  connectors.                                                                 
* 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            
  Joint)                                                                      
* 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         
characteristics:                                                              
* 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             
  SN9422-M8                                                                   
                                                                              
                                                                              
Louvres                                                                       
--------                                                                      
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               
points.                                                                       
                                                                              
                                                                              
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        
  gradients.                                                                  
* 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        
vented.                                                                       
                                                                              
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          
limit.                                                                        
                                                                              
* 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     
heaters.                                                                      
                                                                              
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                 
requirement,                                                                  
                                                                              
* 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             
elsewhere.                                                                    
                                                                              
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                  
simultaneously.                                                               
                                                                              
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                     
performance.                                                                  
                                                                              
                                                                              
Operation                                                                     
----------                                                                    
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               
hybrid.                                                                       
                                                                              
* 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               
ROS_SPACECRAFT frame.                                                         
                                                                              
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            
                                .__o__.                                       
                              .'       `.                                     
                             /           \                                    
                            .   `.   .'   .           +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          
parts:                                                                        
  * 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,              
height=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            
PSS-02-10.                                                                    
                                                                              
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                   
  amplifiers                                                                  
* 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              
  capability).                                                                
* 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             
  circuit.                                                                    
* 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             
  release).                                                                   
* Distributes power to the Thermal Control Subsystem combined                 
  hardware -  software controlled heaters.                                    
* Individual on/off switching for each software controlled heater             
  circuit.                                                                    
* TM/TC interface.                                                            
                                                                              
Batteries                                                                     
----------                                                                    
* 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        
wing.                                                                         
                                                                              
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  |       |  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           
frame.                                                                        
                                                                              
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              
         gimbal;                                                              
                                                                              
      -  +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                          
                             ._____________.^+Xsa+y                           
   .__  _______________.     |             ||    .______________  ___.        
   |  \ \               \    |             ||   /               \ \  |        
   |  / /                \   |  +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              
distance.                                                                     
                                                                              
In the deep space phases the general operational concept was the              
following:                                                                    
                                                                              
  * 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)          
functions.                                                                    
                                                                              
                                                                              
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           
of:                                                                           
                                                                              
* 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       
below.                                                                        
* 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         
algorithms.                                                                   
                                                                              
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        
redundancy.                                                                   
                                                                              
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         
actuators                                                                     
                                                                              
* 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            
link.                                                                         
                                                                              
* 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           
  commands...),                                                               
                                                                              
  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         
  (HGAPM, SADM)                                                               
                                                                              
  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            
  management.                                                                 
                                                                              
                                                                              
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               
corrections.                                                                  
                                                                              
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               
Mode                                                                          
                                                                              
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                
                pointing;                                                     
                                                                              
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.              
                                                                              
Acronyms                                                                      
------------------------------                                                
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                                         
IMU     INERTIAL MEASUREMENT UNITS                                            
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                                              
                                                                              
                                                                              
"                                                                             
                                                                              
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/******************  LANDER PHILAE  ************************/                 
                                                                              
                                                                              
OBJECT                    = INSTRUMENT_HOST                                   
 INSTRUMENT_HOST_ID       = RL                                                
                                                                              
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  INSTRUMENT_HOST_NAME    = "ROSETTA-LANDER"                                  
  INSTRUMENT_HOST_TYPE    = SPACECRAFT                                        
  INSTRUMENT_HOST_DESC    = "                                                 
Lander overview                                                               
=============================================                                 
The Philae Lander was a box-type unit with the dimensions                     
of 850 x 850 x 640 mm3. On the comet, it rested on a tripod                   
called Landing Gear, with a diameter of 2.6 m and was supposed to be fixed to 
the comet's surface by harpoons.                                              
                                                                              
Philae was composed of three different parts, corresponding to its            
structural design:                                                            
                                                                              
1)    Internal compartment:                                                   
This compartment hosted almost all subsystems and most of the                 
experiment units. It provided a temperature controlled environment            
for all electronics and was built by the structural elements of an            
Instrument platform and so called Pi-plates. It was surrounded by             
Multilayer Insulation built of 2 tents to achieve the required                
insulation at a low power environment on the comet at 3 AU distance           
from Sun.                                                                     
                                                                              
2)    Solar Hood:                                                             
The solar hood was built around the internal compartment and its MLI          
tents, the shape followed the overall Lander shape. It hosted the solar       
arrays of the Lander composed by 6 different panels. In addition two          
absorber foils were mounted on the solar hood lid. These foils were           
built by thin copper foils with an external TINOX surface, high               
absorptivity and low emissivity, used to collect solar irradiation            
and transform it into heat radiated into the internal compartment.            
The solar hood also carried the camera system of the Lander, with one         
camera head on each panel, thus providing a 360 degrees panoramic             
view.                                                                         
                                                                              
3)    Baseplate / Balcony:                                                    
The baseplate was the central structural plate carrying the solar hood        
with the internal compartment underneath and providing at one end a           
special area called balcony. This area hosted all experiments or parts        
of them, especially the sensors, which required direct access to the          
comet environment and the comet surface.                                      
The baseplate was also the interface panel to the Landing Gear.               
In addition the baseplate hosted the Push plate, which was the                
interface to the Orbiter during the 10 years cruise from Launch to            
the Comet.                                                                    
                                                                              
The Lander mass was around 100 kg.                                            
                                                                              
In addition three units of the Lander system were mounted on the              
Orbiter, and remained there after Lander separation for the comet.            
These units provided the interfaces to the Orbiter: electrical and            
data (ESS) and mechanical (MSS). The third system was a TxRx system           
used to keep contact to the Lander during its operational phase on            
the comet.                                                                    
                                                                              
                                                                              
Lander Mission Requirements and Constraints                                   
=============================================                                 
The Lander was designed to fullfill the mission requirements given as:        
- survive the 10 years cruise phase with long hibernation phases under        
  autonomous thermal control powered by the Orbiter,                          
- land safely on the comet,                                                   
- provide a scientific phase after landing at 3 AU distance from Sun          
  with online data transmission,                                              
- provide a long term mission capability observing the comet on its           
  way from 3 AU to the Sun                                                    
                                                                              
                                                                              
Lander Platform Definition                                                    
=============================================                                 
The Lander platform was built by three major subsystems, required to          
operate the Lander throughout the mission:                                    
-    a Power subsystem (PSS) composed of a Battery system with a              
        Primary Battery and a Secondary Battery, the later refilled           
        by a Solar array generator, and the required electronics to           
        distribute and control the power flow inside the Lander,              
-    a Central Data Management System (CDMS), composed by two hot             
        redundant computers, controlling all activities on the                
        Lander, especially on the comet in an autonomous manner,              
-    a Thermal Control System, composed by a 2-tent                           
        MultiLayerInsulation supported by two absorber foils and an           
        electrical heater system. Additional independant heater               
        systems were used during the cruise phase, especially when the        
        Lander was in hibernation, and on the comet, when the Lander          
        run out of power and changed into a so called Wake-up mode,           
        to provide a thermal environment in the Internal compartment          
        as required to switch-on the Lander electronics.                      
                                                                              
                                                                              
Subsystem Definition                                                          
=============================================                                 
In addition to the already described platform units PSS, CDMS and TCS         
and the On-Orbiter units ESS, MSS and ESS-TxRx, a set of subsystems           
was installed on the Lander.                                                  
                                                                              
The Active Descent System ADS provided a 1-axis thruster system used          
at touch-down to support the landing and prevent a rebounding until           
the harpoons are shot.                                                        
An Anchoring system, built by two redundant harpoons, was used to fix         
the Lander to the comet's surface after landing and provide the               
required counter-force during drilling.                                       
A Flywheel providing a 1-axis momentum wheel used to stabilize the            
Lander's descent to the comet.                                                
The Landing gear provided the necessary interface between the Lander          
and the comet and supported Lander science operations by a rotation           
and tilting capability.                                                       
The structure subsystem provided the required structural elements to          
built up the Lander.                                                          
A TxRx system was installed to provide access to the Lander and enable        
data retrieval during its mission phase on the comet.                         
                                                                              
                                                                              
Lander Reference Frame                                                        
=============================================                                 
The Lander reference frame was defined as follows:                            
+Z-axis was perpendicular to the baseplate, generally pointing away           
from the comet towards space, during cruise parallel to the Orbiter           
+Z-axis, +X-axis was generally parallel to the comet surface, pointing        
opposite of the Lander's balcony, into the direction of Lander                
separation from the Orbiter, during cruise into Orbiter -X direction,         
+Y-axis completed the right-handed frame.                                     
                                                                              
The frame origin was located on the upper surface of the balcony              
(Z = 0), in the middle of the balcony (Y = 0), at the outer end               
(X = 0).                                                                      
                                                                              
                                                                              
                                                                              
Lander Operating Modes                                                        
=============================================                                 
The Lander was operated in the following modes:                               
                                                                              
Hibernation Mode:                                                             
This mode was defined as: Lander attached to the Orbiter, Orbiter LCL         
5A or 5B ON, Lander Hibernation heater ON (dissipation > 12W at 28V),         
no power on the Lander Primary Bus                                            
In this mode the Lander was non-operational but under thermal control         
with a hibernation temperature inside the internal compartment above          
minus 55 degC at the reference point.                                         
                                                                              
Wake-up Mode:                                                                 
This mode was applied on the comet, substituting the Hibernation Mode.        
The PSS wake-up thermostats were closed, because the temperature              
inside the internal compartment was below minus 53 degC. In this mode         
the Lander was non-operational, the Lander operational electronics were       
disconnected from the Primary Bus and the wake-up heaters were                
connected to the Primary Bus. In this mode NO thermal control was             
possible, since the wake-up heaters would only dissipate, if the              
Primary Bus was powered, which required Sun irradiation on the comet to       
operate the solar arrays. Without dissipation the compartment                 
temperature would drop until the comet environmental temperature. When        
the Lander was still attached to the Orbiter and powered from the             
Orbiter-LCL 15A/B, an additional heater set would also dissipate.             
                                                                              
Power Enough Mode:                                                            
This mode followed the Wake-up mode, the Lander Primary Bus was               
powered, but the voltage was still below 18.5V, which corresponded to a       
non-sufficient power situation. The available power was not lost,             
since special Power Enough loads were used to dissipate and heat the          
internal compartment.                                                         
                                                                              
Stand-by Mode:                                                                
The Lander was operational, since the Lander basic operational                
electronics (PCU, CDMS and one TCU) were connected to the Primary Bus         
and powered.                                                                  
In this mode thermal control would be performed from the dissipation          
of the activated units. If the temperature of the internal                    
compartment dropped below the TCU set-points, the respective TCU              
heaters would also dissipate.                                                 
                                                                              
Operational Modes:                                                            
These modes defined Lander operation of Experiments.                          
                                                                              
APXS:                                                                         
No activity during SDL and FSS.                                               
                                                                              
CIVA:                                                                         
CIVA-P mode Orbiter imaging : Imaging of the Orbiter after delivery           
with camera 1 & 6.                                                            
CIVA-P mode Agilkia Landing site : Imaging of the Landing site Agilkia        
just after touch-down Panorama with all 7 camera but only half of image       
received.                                                                     
CIVA-P mode Abydos location : Imaging of the Abydos after touch-down          
and move on the comet Panorama with all 7 cameras.                            
                                                                              
CONSERT:                                                                      
CONSERT Tuning mode: Instrument Switch ON and tuning of CONSERT               
Lander & Orbiter clocks.                                                      
CONSERT Sounding during descent (SDL) mode: CONSERT Lander emission and       
reception by CONSERT Orbiter. Active during the whole descent and stop during 
the touch down window (CONSERT remaining active during this window but not    
sounding (RF emission) for no interference with other instruments at time of  
Landing. Lander and Orbiter in visibility during all this period as CONSERT   
main objective was to monitor the Lander descent trajectory.                  
CONSERT Sounding after Landing mode: CONSERT Lander emission and reception by 
CONSERT Orbiter. Lander and Orbiter in occultation permitting the sounding of 
the comet structure.                                                          
                                                                              
COSAC:                                                                        
COSAC Taping station test and sniff mode at Agilkia: Evaluation of the COSAC  
taping station position and disengage of the possible taping station of an SD2
Carousel Oven (previous to any SD2 Carousel movement).                        
COSAC Abydos sniff mode: Analysis of the molecules present at the external    
entry of the COSAS Mass Spectrometer done in Agilkia just after landing and   
later on Abydos site.                                                         
                                                                              
MUPUS:                                                                        
MUPUS Anchor mode:                                                            
Measurement of the acceleration sensors ANC-M inside the harpoons during      
anchoring. Measurements of the temperature sensors ANC-T inside the harpoons. 
MUPUS MAPPER mode:                                                            
Calibration of the Thermal Mapper during descent.                             
Two sub-mode to MAPPER mode                                                   
  TM blackbody sub-mode:                                                      
  Calibration of the Thermal Mapper with blackbody in the TM field of view    
  (during descent).                                                           
  Infinite TEM sub-mode:                                                      
  Calibration of the Thermal Mapper with deep space in TM FOV (during descent)
MUPUS TEM mode:                                                               
Passive Thermal Measurement Mode.                                             
MUPUS LONGTERM (-> MAPPER) mode:                                              
Thermal Mapper longterm measurement after landing and during one comet        
rotation.                                                                     
MUPUS PENEL deployment mode:                                                  
Penetrator deployment.                                                        
MUPUS HAMMER mode:                                                            
Penetrator insertion to ground by hammering.                                  
                                                                              
PTOLEMY:                                                                      
PTOLEMY Sniff and CASE mode: Analyse of the molecules present at the external 
entry of PTOLEMY Mass Spectrometer.                                           
PTOLEMY Agilkia sniff mode and tapping station test: Analyse of the molecules 
present at the external entry of PTOLEMY Mass Spectrometer done in Agilkia    
just after landing. Evaluation of the PTOLEMY taping station position and     
disengage of the possible taping station of an SD2 Carousel Oven (previous to 
any SD2 Carousel movement).                                                   
PTOLEMY Abydos sniff mode: Analyse of the molecules present at the            
external entry of PTOLEMY Mass Spectrometer done in Abydos site.              
PTOLEMY Oven mass spectrum analysis:                                          
Analyse of the molecules present in the Oven by PTOLEMY Mass                  
Spectrometer done in Abydos site.                                             
                                                                              
ROLIS:                                                                        
ROLIS DIT mode: ROLIS imaging during descent Agilkia touch down               
location in the field of view.                                                
ROLIS DIS mode: ROLIS imaging during descent once landed.                     
ROLIS CUC mode: ROLIS imaging once on COMET in Abydos site during night period
with illumination by LED sources (blue, green dark, red, IR) Images of Abydos 
site.                                                                         
                                                                              
ROMAP:                                                                        
ROMAP Slow Mode: Magnetometer in Slow mode                                    
 (1Hz sampling 512 octet / mn).                                               
ROMAP Fast Mode: Magnetometer in Fast mode (916 octet / mn).                  
ROMAP Surface Mode SPM: Magnetometer and channeltron after Touch              
Down from Noon to Day/Night transition.                                       
                                                                              
SD2:                                                                          
SD2 Drill downward and upward for COSAC: Carousel movements to zero           
position Drill roto-translations downward and Sampling at 560 mm,             
translation upward to 0mm, sample in Oven#17(SD2 count, HTO) delivered        
to COSAC.                                                                     
SD2 Carousel movements: Carousel movement for PTOLEMY with rotation to        
0 arcmin.                                                                     
                                                                              
SESAME:                                                                       
SESAME CASSE mode:  Measurements executed to register the vibration           
environment generated by the Philae flywheel, intra-foot Soundings and        
inter-foot Soundings, touchdown impact, the cometary vibration background and 
any particles possibly dropping on the sole covers.                           
SESAME PP Passive Mode : Measurements conducted in order to determine         
the electromagnetic environment close to the orbiter.                         
SESAME PP Active Mode : PP calibrations, determine the permittivity           
of the comet surface material, monitor variations in the local                
plasma environment.                                                           
SESAME DIM mode: DIM  conducted to measure the particle environment.          
                                                                              
"                                                                             
                                                                              
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