PDS_VERSION_ID = PDS3 RECORD_TYPE = STREAM 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" 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" END_OBJECT = INSTRUMENT_HOST_INFORMATION OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "PSS-02-10" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "RO-DSS-TN-1155" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "RO-ESC-PL-5000" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "RO-ESC-PL-5100" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "RO-EST-LI-5012" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO END_OBJECT = INSTRUMENT_HOST END