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 " 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 Segment - 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 (Churyumow-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 is 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 are mounted. The two solar panels have 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 accommodates the payload instruments, and the 'base' of the spacecraft the subsystems. The spacecraft can be physically separated into two main modules: * A Payload Support Module (PSM) * A Bus Support Module (BSM) The Lander is attached to the rear face (-X), opposite the two-axes steerable high-gain antenna (HGA). The two solar wings extend from the side faces(+/-Y). The instrument panel points almost always towards the comet, while the antennas and solar arrays point towards the Sun and Earth (at such great distances the Earth is relatively speaking in the same direction). The spacecraft attitude concept is such that the side and back panels are shaded throughout all nominal mission phases, offering a good location for radiators and louvres. This will normally be facing away from the comet, minimising the effects of cometary dust. The spacecraft is built around a vertical thrust tube, whose diameter corresponds to the 1194 mm Ariane-5 interface. This tube contains 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 require: * a large number of complex scientific instruments, to be accommodated on one side of the spacecraft, that shall, in the operational phase, permanently face the comet. During cruise the instruments shall be served for survival. * one Surface Science Package (SSP), to be accommodated, suitable for cruise survival and proper, independent ejection from the orbiter (spacecraft). In addition, the orbiter shall 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 are 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 relate to the interplanetary cruise phases rather than to the scientific objectives, drive 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) is required, that constrains the maximum S/C wet mass and defines the available S/C envelope in Launch configuration. The total mission delta-v of more than 2100 m/s requires a propulsion system with over 1700 kg bi-propellant. The environmental loads (radiation, micro meteoroids impacts) over the mission duration of nearly 12 years is 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 makes a communication link via an on-board high gain antenna (HGA) mandatory. The spacecraft must 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 shall 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) drive 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) require a high degree of on-board autonomy, with corresponding FDIR concepts. Platform Definition ===================================================================== The ROSETTA platform is 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 drives the configuration such that high dissipation units must 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 is 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 have also influenced the design. The telecommunications design is driven by the need to be compatible with ESA's 15m and 32m ground stations and the 34m and 70m DSN stations. This has 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 are accommodated together within the BSM. The electronic units are located mostly on the Y panels so that their thermal dissipations are closely coupled to the louvred radiators on the sidewalls. So far as practical, functionally related groups are located close together for harness, integration and testability reasons. Where possible, equipments are positioned towards the +X half of the S/C to counterbalance the mass of the Lander on the opposite side. Some subsystem equipments are deliberately located on the PSM. These include the PDU and RTU for the payload, the NAVCAMS, two of the SAS units and the +Z LGA. The PDU and RTU are located closer to the payload instruments to reduce harness complexity and mass, and the NAVCAMs and SASs and +Z LGA are located on the PSM for field of view reasons. Other subsystem equipments have been located on the PSM sidewalls as a result of BSM equipment/harness growth, or thermal limitations. These comprise the STR electronics and SSMM as well as the USO. The RCS subsystem comprises tanks, thrusters and the associated valves and pipework. The main tanks are accommodated within the central tube while the helium pressurisation tanks are mounted on the internal deck. Most of the valves and pipework are located on the +X BSM, panel which becomes permanently attached to the BSM once RCS assembly is completed. Sixteen of the twenty-four thrusters are located at the four lower corners of the BSM. The remaining thrusters are located in 4 groups near the top corners of the S/C. They are installed as part of the BSM, but are attached to the PSM after PSM/BSM mating. The Star Trackers are mounted on the -X shearwalls. The STR B is 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 is both thermally stable and mechanically close to the -X PSM panel which accommodates the instruments requiring high pointing accuracy. The reaction wheels are located on the internal deck which provides them with a thermo- elastically stable location. A 2.2m diameter HGA is 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 can 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 provides the HGA with greater than hemispherical pointing range. The two MGAs are fixed mounted on the S/C +X face, oriented in the +Xs/c direction, as this is the most useful direction for a fixed MGA. The LGAs are located at the +Z and -Z ends of the S/C but angled at 30 degs to the Z axis. This accommodation provides spherical coverage with minimum need for switching. The solar array comprises two 5-panel wings folded against the Spacecraft Y axis for launch. Because the arrays are sized to operate at aphelion, the outwards facing outer panel can also generate useful power before array deployment. Two Sun Acquisition Sensors are located on the solar arrays and another two on the S/C body. Their design and location of these also allow them to serve as fine Sun sensors. Rosetta Spacecraft Frame ===================================================================== Rosetta spacecraft frame is defined as follows: - +Z axis is perpendicular to the launch vehicle interface plane and points toward the payload side; - +X axis is perpendicular to the HGA mounting plane and points toward HGA; - +Y axis completes the frame is right-handed. - the origin of this frame is 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 consists of two modules, the Bus Support Module and the Payload Support Module (BSM and PSM). Mounted to the BSM is the Lander Interface Panel (LIP), which can be handled separately for the Lander integration. The spacecraft structural design is based on a version with a central cylinder accommodating the two propellant tanks. The general dimensions are 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 is the central tube, to which the honeycomb panels are mounted. The spacecraft box is closed by lateral panels, which are connected to the central tube by load carrying vertical shear webs and an internal deck. The Bus Support Module (BSM) accommodates most of the platform and avionic equipment. The Payload Support Module (PSM) is accommodating all science equipment. The PSM structure consists of the PSM +z-panel, the PSM -x panel, the PSM +y/-y panels and the Lander Interface Panel (LIP). Most instrument sensors are 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 are mounted on the +Y and -Y side of this module for heat radiation via Louvers. Special supports are provided by the structure for: Solar Array ----------- They provide stiff and accurately positioned points for the solar array hold down points and for solar arrays drive mechanisms. Reaction Wheels --------------- The brackets provide stiff wheel support with alignment capability. All 4 RW brackets are mounted together between the +X shear wall and the central deck building one compact bracket unit which provides high stiffness and stability. Propellant Tanks ---------------- The two tanks are mounted via a circumferential ring of flanges to a reinforced adapter ring on the tube with titanium screws. Helium Tanks ------------ The two helium tanks are mounted on the main deck of the BSM. They are attached by an equatorial fixation in the middle of the tank through internal deck holes. Thrusters --------- Thrusters on the side of the spacecraft are mounted on lateral panel extensions with aluminium machined brackets ensuring the angular position of the thrusters. Thrusters underneath the spacecraft (-Z pointing thrusters) are mounted on brackets on the corners of the +/-Y panels. High Gain Antenna ----------------- The HGA is stowed against the +X panel, in areas stiffened by the +/-Y panels and the HGA support tripod. After launch, the HGA is deployed and is connected to the S/C by the support tripod only. The axis Antenna Pointing Mechanisms, fixed on the tripod, are located close to the edge of the HGA. Gyros ----- A single bracket provides stiff gyro support and alignment capability and orientates the 3 IMUs in the requested angular orientation. The bracket is mounted on the -Y BSM panel for thermal dissipation reasons. Mechanisms Design ===================================================================== The ROSETTA mechanisms comprise 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 performs the positioning of the Solar Array w.r.t. the Sun by rotation of the panels around the spacecraft Y-axis. There are two identical SADMs on both sides of the spacecraft, which can be individually controlled. The control authority rests with the AOCMS subsystem, which always 'knows' the actual attitude and Sun direction and is therefore in the position to determine the required orientation of the solar panels. The positioning commands are routed from the AOCMS I/F Unit via the SADE (SADM-Electronics) to the SADM. The Solar Array rotation is limited to plus and minus 180 degrees to the reference position. The array zero position is defined in the section 'Power Design: Solar Array Generator' below. The Solar Array Drive Mechanism baseline design comprises 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 employs 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 is 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 contains a position transducer and a drive train. The Solar Arrays Drive Electronic is intended to manage two Solar Array Drives that can be rotated so as to get the maximum energy from the solar cell panels. Solar Array Deployment Mechanisms ---------------------------------- The baseline are 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 are stowed against the sidewalls of the satellite. They are kept in this position by means of 6 hold-down mechanisms per wing. Approximately 3 hours after launch, the satellite is pointed towards the Sun and the wings are deployed to their fully deployed position. They are released for full deployment by 'cutting' Kevlar restraint cables by means of thermal knives (actually degrading of the Kevlar by heat). The deployment system makes use of spring driven hinges and is equipped with a damper, that limits the deployment speed of the wing. Thus, the deployment shocks on SADM hinge and inter-panel hinges are kept relatively low. The Rosetta wing is 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 is a two-axes mechanism which allows motion of the HGA in both azimuth and elevation. The control authority rests with the AOCMS subsystem, which always 'knows' the actual attitude and Earth direction and is therefore in the position to determine the required orientation of the antenna. The positioning commands are routed from the AOCMS I/F Unit via the APM-E (APM-Electronics) to the APMM. HGA elevation rotation is physically limited to +30deg/ -165deg from the reference position (after deployment). Before and during deployment the range is -207deg and +30deg. HGA azimuth rotation is physically limited to +80deg / -260deg from the reference position. The main functions of the APM are: * 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 consists 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 is 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 are covered with MLI. Experiment Boom Mechanisms --------------------------- Two deployable experiment booms support a number of different lightweight sensors from the plasma package which need to be deployed clear of the S/C body. These booms are deployed at beginning of the mission after Launch. Each boom consists of a 76 mm dia CFRP tube. The lower boom is approximately 1.3 m long and the upper boom 2m. The boom deployment is performed by means of a motor driven unit. The deployment mechanism consists of: * Hinge, Motor Gear Unit, Coupling system, Latching system and Position switches. The Hold down and release mechanisms, one per boom, has the following characteristics: * Three Titanium blades to allow relative displacement in the boom centreline direction. This reduces the mechanical and thermo- elastic I/F forces. * The separation device is the Hi-Shear low shock Separation Nut SN9422-M8 Louvres -------- The Rosetta Thermal Control Subsystem contains 14 louvers with 2 different set points which are located on the S/C Y walls in front of white painted radiators. The louvers are designed, manufactured and qualified by SENER. The mechanisms of the 16 blade louver are the 8 temperature dependent bi-metal springs (actuators), which supply the fundamental function of the louver. The actuators are 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 is 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 utilises conventional passive components supported by active units like heaters and controlled radiative areas, using well proven methods and classical elements. This concept can be characterised as follows : * Heat flows from and to the external environment are minimised using high performance Multi-Layer Insulation (MLI). * Most unit heat is rejected through dedicated white paint radiator, actively controlled by louvers, located on very low Sun-illuminated +/-Y panels. * High internal emissivity compartments reduce structural temperature gradients. * Individually controlled instruments and appendages (booms, antennas ,...) are mounted thermally decoupled from the structure. * High temperature MLI is used in the vicinity of thrusters. * Optimised heaters, dedicated to operational, and hibernation modes, are monitored and controlled to judiciously compensate the heat deficit during cold environment phases. Thermal control design ----------------------- The thermal control subsystem (TCS) design is 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 is the most critical thermal design case. This heater power request is dependent on the radiator sizing which need to be performed for worst case end of mission conditions. The very strong heater power limitation implies that to a certain extent constraints in the operation and/or attitude need to be accepted for hot case. The TCS uses a combination of selected surface finishes, heaters, multi-layer insulation (MLI) and louvres to control the units in the allowable temperature ranges. The units are 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 are thermally decoupled from the structure. Generated heat by the collectively controlled units is then rejected via conduction into the panel and subsequent radiation from the external surface of the panel to space. These surfaces are covered with louvers over white painted radiators minimising any absorbed heat inputs and heat losses in cold mission phases. The louvers are selected as baseline being the best solution (investigated during phase B) for flexibility, qualification status and reliability. VIRTIS and OSIRIS cameras are located at the top of the -X (anti-sun face) so that their radiator may view deep space. The top floor is extended over the top as a sunshield to prevent any direct solar illumination of these instruments, while the sun angle on the -Z side has to be limited to 80 degrees for the same reason. Any external structural surface not required as a radiator, (or experiment aperture) is covered with a high performance MLI blanket. The bottom of the bus module, which is not enclosed with a structural panel, is 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 are implemented. All blankets are adequately grounded and vented. The bi-propellant propulsion subsystem needs to be maintained between 0 to +45 degrees throughout the mission. This is far warmer than some units, particularly when the spacecraft is in deep space hibernation mode. The tanks and RCS are therefore well isolated from the rest of the spacecraft to allow their specific thermal control. The antennae and experiment booms are passively thermally controlled by the use of appropriate thermo-optical surface finishes and MLI. The mechanism for the HGA has similar appropriate passive control but also needs heaters to prevent the mechanism from freezing. It is thermally decoupled from the rest of the spacecraft to allow its dedicated thermal control. The chosen solution for thermal control subsystem design uses well known and proven technologies and concepts. General Heater Control Concept ------------------------------- The operation of the TCS shall enable 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 uses the following major control features: * Thermistor controlled (software) heater circuits, which are used to maintain platform, avionics and payload units within operating limits when these units are operating. * The S/W heater design includes 3 control thermistors sited next to each other and uses the middle temperature reading to control the heater switching. This method is used in order to maximise the reliability of thermistor controlling temperature. * Thermistors will be 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 are used to maintain platform, avionics and payload units within their non- operating (or switch-on) limits when these units are non-operating. These operate autonomously during satellite hibernation and Safe modes to ensure thermal control. * The hardware heater circuits will be controlled by one thermostat (cold guard) connected in redundant circuit. The prime circuits without any thermostat will be powered as long as the relevant LCL is defined to be enabled. In the prime circuit a thermostat (hot guard) is included to prevent from overheating. In the event of a failure in the prime circuit the redundant circuit is automatically switched on when the temperature falls because it is permanently enabled. * The lower set points for the thermostats (cold guard) are at the lower nonoperating limits of units. The hysteresis of the thermostats is 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) is oriented to the upper operational temperature limit, but will still have an appropriate margin to that limit. * Main and redundant heaters will be in separate foil heaters. It is necessary to define reserved unpainted areas on all units, which would nominally be black painted, specifically for the mounting of heaters. All software and hardware heaters circuits will comprise a simple series connection of heaters with no parallel connections. The heater concept assumes prime and redundant heater elements in different mats. The heaters will be mounted directly onto units as this maximises the efficiency of the heating. The sizing of the autonomous H/W heater circuits are based upon the following criteria: * Payload heaters shall be designed to maintain non-operating temperature limits at 5.33AU or switch-on limits at 3.25AU, whichever gives the greater heater power requirement, * Platform and Avionics units OFF in hibernation have heaters designed to maintain non-operating temperature limits at 5.33AU or switch-on limits at 4.5AU, whichever is the greater power requirement, * Platform and Avionics units ON during hibernation have heaters designed to maintain operating temperature limits at 5.33 AU. The suppliers of individually controlled (I/C) units shall size their S/W and H/W heaters by themselves and may install them where they wish in order to control their unit temperatures. Micrometeoroid and Cometary Dust Protection -------------------------------------------- The micrometeoroid protection used for Rosetta is composed of 2 layers of betacloth and a spacer. This protection is only applied to the exposed +Z and -Z central tube areas of the propellant tanks as the spacecraft honeycomb structure will form an effective shield elsewhere. The first betacloth layer is underneath the outermost layer of the S/C MLI acting as a bumper. To reach the agreed probability of no micrometeroid impacts in 998 out of 1000 strikes, a separation of 50mm to the second betacloth layer (on top of the tank MLI) is needed. The micrometeoroid protection is part of the overall MLI design. The cometary dust will have a velocity similar to that of Rosetta and so hypervelocity impacts are not an issue. Of more concern is the coating of the spacecraft surfaces by the cometary dust. Grounding of the external surfaces prevents differential charging but the whole spacecraft may be charged to some potential. Propulsion Design ===================================================================== The propulsion subsystem is based on a pressure fed bipropellant type using MMH (MonoMethylHydrazine) and NTO (Nitrogen TetrOxide). It is capable to operate in both regulated and in blow-down mode and provides a delta v of more than 2100 m/s plus attitude control. It is 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 provides a high degree of redundancy in order to cope with the special requirements of the ROSETTA mission. The materials used in the propulsion subsystem are 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 are installed on the spacecraft -X panel by means of supporting brackets made of material with low thermal conductance. The subsystem has 24 10 N thruster for attitude and orbit control. They are located such that they can provide pure forces and pure torques to the spacecraft. The 24 thrusters are grouped in pairs on the brackets, one of each pair being the main and one the redundant thruster. The subsystem allows the operation of 8 thrusters simultaneously. The subsystem will be maintained within the temperature limits of the components. The mixture ratio may be adjusted by tank temperature (i.e. pressure) manipulation in order to enhance thruster performance. Operation ---------- The propulsion subsystem will be operated in regulated mode as well as in blow down mode. The pressurisation strategy must take 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 have 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 is 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 is 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 comprises 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 modulate and transmit 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 transpond the ranging signal in S and X-Band. The Transponders provide hot redundancy for the receiving functions and cold redundancy for transmitting functions. The receivers can receive telecommands in S-Band or X-Band (selectable per command), but not simultaneously in both frequency bands. The configuration is such that both receivers can receive, demodulate and send the telecommand signal to the DMS simultaneously. The transmitters are also able to receive the telemetry stream from both parts of the redundant DMS. Each transponder is 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 transmit/receive 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 is 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 is possible to switch between antennas without turning off the TWTA. * The transmit frequency (and receiver rest frequency) can also be derived from an external Ultra Stable Oscillator (USO) on request by Telecommand which may be used any time during the mission. This USO has 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 be 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 is reached. The S-Band MGA is realised as a flat patch antenna whereas the X- Band MGA is a offset-type 0.31m reflector antenna. The MGAs also perform 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 is depending reliable operation of the High Gain Antenna major assembly, which is therefore a critical element for the mission success. The most important requirements for this assembly are: * High reliability * conform to specified pointing requirements * minimize mechanical disturbances * comply to antenna gain requirements The HGA Major Assembly comprises: * 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 is 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, is defined with its orientation given relative to the ROS_SPACECRAFT frame. The ROS_HGA frame is defined as follows: - +Z axis is in the antenna boresight direction; - +X axis points from the gimbal toward the antenna dish symmetry axis; - +Y axis completes the right hand frame; - the origin of the frame is located at the geometric center of the HGA dish outer rim circle. The rotation from the spacecraft frame to the HGA frame can 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 has 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 consists 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) ranges between 14.1 and 14.7 dBi in the receive and transmit frequency bandwidths. The MGAS assembly can 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 are arranged in a hexagonal lattice to provide the required symmetry to the antenna pattern. Six elements are used to meet the required specification. MGA X-band (MGAX) - - - - - - - - - The configuration of the X-band MGA (MGAX) is a single offset parabolic reflector illuminated by a circular polarised conical horn. Reflector dimensions are selected to reach a desired minimum gain and to lead to a simple feeder design. This leads 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 is obtained which ensures small feeder dimensions and a compact antenna design. The MGAX antenna assembly is composed of two sub-assemblies, a reflector and a feeder, and of a platform which supports both these sub-assemblies and provides the interface to the Rosetta spacecraft. The total envelope of the antenna is length=600mm, width=320mm, height=320mm. The thermal protection for the antenna consists 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 are implemented on the satellite in opposite direction to achieve an omnidirectional coverage. One is located at the +Z-panel in the near of the edge to the +X panel and thus is orientated towards the comet during the comet mission phase. The other one is mounted on the opposite face. Ultra Stable Oscillator ------------------------ An Ultra Stable Oscillator is 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 will be used by the TTC subsystem whenever needed and is available for RSI measurements as well. Should the USO fail, each transponder will use it's own oscillator (TCX0), but with less stability and not harming the performance. Power Design ===================================================================== The Power Subsystem (PSS) conditions, regulates and distributes all the electrical power required by the spacecraft throughout all phases of the mission. Distribution involves 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 switches, protects and distributes power for the pyrotechnics and the thermal knives of the various release mechanisms of the spacecraft. Main power source for Rosetta is provided by the Solar Array Subsystem from silicon solar cells mounted on 2 identical solar array wings, which are deployed from the +Y and -Y faces of the spacecraft and can be rotated to track the sun. The solar cells on the outer panel of each wing are 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 provide power for launch and post-separation support until the solar arrays are fully deployed and sun aligned, and thereafter will support the main power bus as necessary to supply peak loads and special situations during Safe Mode where the sun might not be fully oriented towards the sun. One special feature of the power supply is the Maximum Power Point Tracker (MPPT), which will 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 will operate in nominal mode, i.e. the PCU takes only the power required by the satellite from the solar array. The delta power will remain in the solar array. Because of this feature the actual performance of the array can only be assessed by utilising 'performance strings' which operate 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 can be determined. Batteries are also the main power source for the pyrotechnics, although pyrotechnic power is also available from the main bus as a back-up in case there is no battery power. The subsystem is designed in accordance to the ESA Power Standard PSS-02-10. Power Conditioning Unit (PCU) ----------------------------- * Produces 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 are all switched and protected by Latching Current Limiters (LCL). * LCLs have 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. * Distributes 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 are switched and protected by Latching Current Limiters (LCL). * FCLs and LCLs have current measurement and FCLs have 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 has an extremely wide variation of Spacecraft- Earth-Sun angles and distances, hence it is 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) are identical to the qualified ARA MK3 design of Fokker Space. The geometry and mechanical interface definition of the Rosetta baseline Solar Array design is identical to the 5-panel qualification wing. The electrical architecture (cells, strings, sections & harness lay- out) is uniquely designed for Rosetta. Electro static discharge (ESD) protection design is qualified for the ARA MK3 type solar array. The baseline are 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 are 2.25 x 2.736 m2. The front side skin will use a 50^m Kapton foil to isolate the solar cell network from the conductive CFRP layers. The Kapton foil is co-cured with the CFRP layers. The panel core consists of Aluminium honeycomb with a core height of 22 mm. Local circular reinforcement plugs ('subassembly panels') are used to provide the holddown areas with extra strength, stiffness and fatigue resistance. The hold-down and release system uses a tie-down element (Kevlar cable) under high preload which will be degraded by heat of the thermal knife for release. The hold-down, SADM and yoke snubber locations for Rosetta are fully identical to the ARA MK3 qualification hardware definition. The stowed wing has a height of <239 mm at the wing tips (the gap between inner panel and sidewall is increased from nominal 70 mm by about 30mm by means of a dedicated bracket, the inter panel gap is 12 mm, and the panel substrate thickness is 22 mm). The deployment mechanism concept relies on spring-driven hinges. The spring characteristics are chosen such that the energy supply is 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 is introduced in the deployment system. A stiff synchronisation system is applied to prevent a very non- synchronous deployment, resulting in unpredictable high deployment latch-up shocks at the interpanel hinges. The V-yoke length is 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 is identical. The arms of the V-shaped yoke consist 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 can be articulated (each having one degree of freedom), the solar Array frames, ROS_SA+Y and ROS_SA-Y, are defined with their orientation given relative to the ROS_SPACECRAFT frame. Both array frames are defined as follows : - +Y axis is parallel to the longest side of the array, positively oriented from the end of the wing toward the gimbal; - +Z axis is normal to the solar array plane, the solar cells on the +Z side; - +X axis is defined such that (X,Y,Z) is right handed; - the origin of the frame is located at the geometric center of the gimbal. The axis of rotation is parallel to the Y axis of the spacecraft and solar array frames. At zero (reference) position the array wing is aligned such that the array surface is in the spacecraft Y-Z plane, with the face (cells) aligned such that the array normal is 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 is -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 forces the use of economical strategies for certain operations. Thereby the situation after the deep space hibernation phase is much more severe. From radiation degradation analysis it has been derived that after DSHM at 4.5 AU about 65 W less solar array power will be available compared to 4.5 AU before DSHM. This corresponds to about 13% of the power needed at that distance. In the deep space phases the general operational concept is 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 are not directly under ground control, like in Safe Mode, the ground can set a Low Power Flag as invocation parameter in the call of the Safe Mode OBCP (which is loaded in the System Init Table) at the appropriate time in the mission, according to the current Sun distance. This flag will be checked by the OBCP; if the flag is set, the Safe Mode downlink will be performed in power sharing strategy and the SSMM is set into stand-by mode (memory modules remain powered, but memory controllers are switched off). As a safety precaution the battery discharge alarm shall remain enabled all the time. This will allow for nominal short (< 4 min) peak power demands to be satisfied by the batteries, e.g. for RW offloading, but will 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 is not desired, it is 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 performs the electrical connection between all electrical and electronic equipment in the ROSETTA spacecraft. It provides 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 consists of the following subassemblies: * Payload Support Module Harness * Bus Support Module Harness * Harness to the Lander I/F Furthermore the harness / cables are divided into three harness EMC classes: power, signal and data, and the pyro harness. Their routing is physically separated. In addition to the appropriate twisting and shielding techniques this minimises the probability of electrical cross talking of critical lines. The harness design follows a distributed single point grounding scheme. Redundant functions have their own connectors and are routed in separate bundles and in a different way as far as practical. All connectors supplying power have female contacts. To achieve a complete Faraday cage around the harness each of the harnesses has 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) are bonded to the structure with a two component conductive glue. The distance of the Ty-bases is 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 are used between connector case and harness overall shield. In case of pyro-lines and sensible interfaces conductive connector boots are implemented. To prevent contamination the harness was baked-out in a thermal vacuum chamber prior to integration. Avionics Design ===================================================================== The ROSETTA Avionics consists 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 is 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 is 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 is the data route for data acquisition and commands distribution via the RTUs. Payload Instruments are accessed via a dedicated Payload RTU. Subsystems are accessed via a dedicated Subsystem RTU. DMS includes 4 identical Processor Modules (PM) located in 2 CDMUs. Any of the processor modules can perform either the DMS or the AOCMS processing. The PM selected for the DMS function acts as the bus master. It is also in charge of Platform subsystem management (TTC, Power, Thermal). The one selected as the AOCMS computer is in charge of all sensors, actuators, HGA & SA drive electronics. TCdecoder and Transfer Frame Generator (TFG) are included in each CDMU. Telemetry can 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) is used like a 'Hard Disk Storage' including 25 Gbit of memory. It contains a data compression module which allows lossy (for CAM image) and loss-less (for HK and science data) compression of data to be stored. It is able of file management capability. It stores CAM images, science and telemetry packets as well as software data for the AOCMS and DMS computer. It is 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) will be used for image data compression of the NAVCAM or STR. The degree of compression can be set by filter parameters from ground. The compression of OSIRIS and VIRTIS image data could also be performed inside the SSMM. Present baseline however is that these two instruments do not request data compression from the system. The SSMM SW runs on a Digital Signal Processor. The SSMM SW is made of: * The Init Mode Software The Init mode software ensures the boot up of the SSMM and the establishment of the communication with the DMS SW. It allows the loading of the operational SW from EEPROM to RAM, and its starting. * The Operational Software The operational SW manages the files located in the Memory Modules of SSMM, and the Data Compression Function that performs Rice lossless and Wavelet lossy data compression. The functionality of the SSMM can be summarised with the three points below. * Store on-board data in files. The on-board data can 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 consists of the following modules: * 2 Memory Controllers (MC) * 3 Memory Modules (MM) * 2 Power Converters, which supplies power to the memory controller and memory module boards. The Memory Controllers are 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 work in cold redundancy. The three Memory Modules are where the files are stored. The modules can be turned on and off independently, giving the possibility to increase and decrease the storage capacity of the SSMM. The Memory Controllers access the Memory Modules via a memory module bus. Both the Memory Controllers can access all three Memory Modules. Attitude and Orbit Control Measurement System (AOCMS) ----------------------------------------------------- The AOCMS is in charge of attitude and orbit measurement and control and is in charge with sensors and actuators for autonomous attitude determination and control as well as pre-programmed manoeuvring. The AOCMS uses 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) is performed through the AIU. Functional AOCMS data which need to be put in the Telemetry and sent to the ground are given packetised by the AOCMS processor and sent to the DMS processor for futher downlink to ground and storage in the SSMM. The DMS PM permanently checks the AOCMS health by monitoring that the AOCMS PM does not stop to communicate with DMS PM. This is done by checking the correct reception of the so-called 'essential' AOCMS HK packet every one second. The AIU is the central data acquisition and distribution unit which allows access to the sensors and actuators with different type of interfaces. It includes RS 422, IEEE 1355 and MACS Bus interfaces as well as analog and discrete digital interfaces for commanding and data acquisition. The AIU includes 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 is critical. The major AOCMS components are the following: * AOCMS Interface Unit (AIU): it interfaces to all AOCMS sensors and actuators * The Sun Acquisition Sensors (SAS): they are internally redundant and are used for Sun Acquisition and pointing. They provide full sky coverage and ensure a permanent sensing of the Sun direction vector. * The Inertial Measurement Packages (IMP): The IMP function provides roll rate and velocity measurements along 3 orthogonal axes. * 4 Reaction Wheels: they are 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 contain an Autonomous Star Pattern Recognition function and provide autonomously to the AOCMS an estimated attitude quaternion and stellar measurements data. * 2 Navigation Cameras (A&B) are used in the AOCMS control loop during the Asteroid Near Fly-by Phase. The navigation cameras can also directly send image data to the SSMM through a high data rate link. * Pointing mechanisms (through target pointing angles) and propulsion thruster valves are commanded by the AOCMS through the AIU links. Avionics external interface ---------------------------------------------- The Avionics system has the following external interface to other subsystems of the Rosetta spacecraft: * Interface with the Ground through TTC Subsystem: Ground Telecommands (TC) are 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 provides the experiments and Platform equipment with a hardware command capability (power On/Off commands, heater On/Off commands...), The Avionics provides experiments with a time synchronisation capability, so that the Ground can later on correlate results coming from different experiments, The Avionics uses 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 are collected on-board to be sent to Ground in real time TM, or to be stored for play-back downlink, The Avionics S/W provides 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 are the following: Stand-By Mode -------------- The SBM is used in Pre-launch and Launch Modes for general check supervision. Only DMS functions are activated. It is possible to command thrusters through AIU for RCS Priming. Sun Acquisition Mode --------------------- This mode is used during Separation Sequence to perform rate reduction (if necessary), Sun acquisition and Sun pointing. SAM is 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 follows 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 noth of the ecliptic plane. In some mission phases (i.e. defined by the minimum earth distance), S/C X-axis pointing towards the Earth is forbidden because of thermal constraints. Then, +X axis is pointed towards the Sun, and the High Gain Antenna is pointed towards the Earth. Normal Mode ------------ The NM is used in Active Cruise and Near Comet phases for nominal longterm operations, for comet observation and SSP delivery. Reaction wheel off-loading is a function of the Normal Mode. Thruster Transition Mode ------------------------- The TTM is used for transition from Normal Mode to operational thruster Modes, and vice-versa, for control tranquillisation. Orbit Control Mode ------------------ The OCM is used in Active Cruise Mode for trajectory and orbit corrections. Asteroid Fly-By Mode -------------------- The AFB mode is dedicated to asteroid observation. Near Sun Hibernation Mode ------------------------- The NSHM is 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 is necessary to spin up the spacecraft at hibernation entry (spin down at hibernation exit is achieved by Sun Keeping Mode). The attitude control concept is a completely passive inertial spin during the deep space hibernation phase. There is no AOCMS Deep Space Hibernation Mode. Sun Keeping Mode ---------------- The Sun Keeping Mode is 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 is not efficient. In case the autonomous entry to Safe / Hold Mode is disabled or not successful Earth Strobing Mode is established leading to a slow spin motion around the Sun direction. Then the + X-axis is pointed towards the expected earth direction (i.e. using the actual Sun/spacecraft/ Earth angle). The rotation along the Sun line is maintained therefore the Earth crosses once per revolution the + X-axis which will allow communication with the MGA. System Level Modes ===================================================================== A basic conficuration 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 Segment ===================================================================== Ground Station and Communications Network performing telemetry, telecommand and tracking operations within the S/X-band frequencies. Telecommand will always be in the S-band, whilst telemetry will be switchable between S- and X-band, with the possibility to transmit simultaneously in both frequency bands, only one of which will be modulated (S-band down-link is primarily used during the near Earth mission phases). The ground station used throughout all mission phases will be the ESA Perth 32m deep-space terminal (complemented by the ESA Kourou 15m station during near-Earth mission phases). In addition, the NASA Deep Space Network (DSN) 34m and/or 70m network is envisaged for back-up and emergency cases. New Norcia Dur. Start-Date End-Date -------------------------------------------------- NNO Daily 129d 26/02/04 03/07/04 NNO Weekly 64d 04/07/04 05/09/04 NNO Daily 56d 06/09/04 31/10/04 NNO Weekly 61d 01/11/04 31/12/04 NNO Weekly 30d 01/01/05 30/01/05 NNO Daily 116d 31/01/05 26/05/05 NNO Daily 52d 27/05/05 17/07/05 NNO Weekly 63d 18/07/05 18/09/05 NNO Daily 7d 19/09/05 25/09/05 NNO Weekly 21d 26/09/05 16/10/05 NNO Daily 7d 17/10/05 23/10/05 NNO Weekly 28d 24/10/05 20/11/05 NNO Monthly 41d 21/11/05 31/12/05 NNO Monthly 50d 01/01/06 19/02/06 NNO Daily 16d 20/02/06 07/03/06 NNO Weekly 13d 08/03/06 20/03/06 NNO Daily 48d 21/03/06 07/05/06 NNO Weekly 14d 08/05/06 21/05/06 NNO Daily 3d 22/05/06 24/05/06 NNO Weekly 28d 25/05/06 21/06/06 NNO Monthly 32d 22/06/06 23/07/06 NNO Weekly 35d 24/07/06 27/08/06 NNO Daily 63d 28/08/06 29/10/06 NNO Weekly 28d 30/10/06 26/11/06 NNO Daily 28d 27/11/06 24/12/06 NNO Weekly 7d 25/12/06 31/12/06 NNO Weekly 31d 01/01/07 31/01/07 NNO Daily 122d 01/02/07 02/06/07 NNO Weekly 28d 03/06/07 30/06/07 NNO Monthly 71d 01/07/07 09/09/07 NNO Weekly 21d 10/09/07 30/09/07 NNO Daily 74d 01/10/07 13/12/07 NNO Weekly 18d 14/12/07 31/12/07 NNO Weekly 10d 01/01/08 10/01/08 NNO Daily 7d 11/01/08 17/01/08 NNO Weekly 28d 18/01/08 14/02/08 NNO Monthly 136d 15/02/08 29/06/08 NNO Daily 129d 30/06/08 05/11/08 NNO Weekly 56d 06/11/08 31/12/08 NNO Weekly 21d 01/01/09 21/01/09 NNO Weekly 28d 22/01/09 18/02/09 NNO Daily 65d 19/02/09 24/04/09 NNO Weekly 28d 25/04/09 22/05/09 NNO Monthly 105d 23/05/09 04/09/09 NNO Weekly 28d 05/09/09 02/10/09 NNO Daily 79d 01/10/09 18/12/09 NNO Weekly 13d 19/12/09 31/12/09 NNO Daily 63d 01/01/10 04/03/10 NNO Monthly 62d 05/03/10 05/05/10 NNO Daily 144d 06/05/10 26/09/10 NNO Weekly 42d 27/09/10 07/11/10 NNO Daily 54d 08/11/10 31/12/10 NNO Daily 102d 01/01/11 12/04/11 NNO Weekly 37d 13/04/11 19/05/11 NNO Daily 55d 20/05/11 13/07/11 NNO Daily 343d 23/01/14 31/12/14 NNO Daily 365d 01/01/15 31/12/15 Cebreros --------------------- Support of the ESA Cebreros ground station is scheduled for 90 days between the 18-Aug-2014 and the 15-Nov-2014 to support comet characterization and Lander delivery. Kourou Dur. Start-Date End-Date --------------------------------------------------- Kourou 1 14d 26/02/2004 11/03/2004 Kourou 2 30d 04/02/2005 05/03/2005 Kourou 3 30d 22/10/2007 20/11/2007 Kourou 4 30d 22/10/2009 20/11/2009 The support around the Earth swing-by is limited to a few passes close to the swing-by and a few weekly passes prior to the swing-by to verify the compatibility between the ground station and the spacecraft. NASA DSN Dur. Start-Date End-Date -------------------------------------------------- DSN1 14d 26/02/04 10/03/04 DSN2 93d 26/02/04 29/05/04 DSN3 7d 03/06/04 09/06/04 DSN4 42d 06/09/04 17/10/04 DSN5 30d 17/02/05 18/03/05 DDOR Check 14d 07/08/06 20/08/06 DSN6 38d 01/09/06 08/10/06 DDOR1 14d 09/10/06 22/10/06 DSN7 155d 23/10/06 26/03/07 DSN8 30d 31/10/07 29/11/07 DSN9 40d 08/08/08 16/09/08 DSN10 30d 28/10/09 26/11/09 DSN11 40d 12/06/10 21/07/10 DSN12 115d 10/11/10 04/03/11 DSN13 30d 05/03/11 03/04/11 DSN14 153d 23/01/14 24/06/14 DSN15 34d 23/07/14 25/08/14 DSN16 28d 25/10/14 21/11/14 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 Controlller 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 Sience 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 /********** REFERENCES **************/ 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 = "PSS-02-10" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO END_OBJECT = INSTRUMENT_HOST /****************** LANDER PHILAE ************************/ OBJECT = INSTRUMENT_HOST INSTRUMENT_HOST_ID = RL OBJECT = INSTRUMENT_HOST_INFORMATION INSTRUMENT_HOST_NAME = "ROSETTA-LANDER" INSTRUMENT_HOST_TYPE = SPACECRAFT INSTRUMENT_HOST_DESC = " Lander overview ============================================= The Philae Lander is a box-type unit with the dimensions of 850 x 850 x 640 mm3. On the comet, it will rest on a tripod called Landing Gear, with a diameter of 2.6 m and will be fixed to the comet's surface by harpoons. Philae is composed of three different parts, corresponding to its structural design: 1) Internal compartment: This compartment hosts almost all subsystems and most of the experiment units. It provides a temperature controlled environment for all electronics and is built by the structural elements of an Instrument platform and so called Pi-plates. It is 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 is built around the internal compartment and its MLI tents, the shape follows the overall Lander shape. It hosts the solar arrays of the Lander composed by 6 different panels. In addition two absorber foils are mounted on the solar hood lid. These foils are 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 carries 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 is 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 hosts all experiments or parts of them, especially the sensors, which require direct access to the comet environment and the comet surface. The baseplate is also the interface panel to the Landing Gear. In addition the baseplate hosts the Push plate, which is the interface to the Orbiter during the 10 years cruise from Launch to the Comet. The Lander mass is around 100 kg. In addition three units of the Lander system are mounted on the Orbiter, and will remain there after Lander separation for the comet. These units provide the interfaces to the Orbiter: electrical and data (ESS) and mechanical (MSS). The third system is a TxRx system used to keep contact to the Lander during its operational phase on the comet. Lander Mission Requirements and Constraints ============================================= The Lander is 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 is 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 are used during the cruise phase, especially when the Lander is in hibernation, and on the comet, when the Lander runs out of power and changes 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 Definiton ============================================= 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 is installed on the Lander. The Active Descent System ADS provides 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, is 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 provides the necessary interface between the Lander and the comet and supports Lander science operations by a rotation and tiliting capability. The structure subsystem provides the required structural elements to built up the Lander. A TxRx system is installed to provide access to the Lander and enable data retrievel during its mission phase on the comet. Lander Reference Frame ============================================= The Lander reference frame is defined as follows: +Z-axis is perdendicular to the baseplate, generally pointing away from the comet towards space, during cruise parallel to the Orbiter +Z-axis, +X-axis is 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 completes the right-handed frame. The frame origin is 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 is operated in the following modes: Hibernation Mode: This mode is 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 is 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 is applied on the comet, substituting the Hibernation Mode. The PSS wake-up thermostats are closed, because the temperature inside the internal compartment is below minus 53 degC. In this mode the Lander is non-operational, the Lander operational electronics are disconnected from the Primary Bus and the wake-up heaters are connected to the Primary Bus. In this mode NO thermal control is possible, since the wake-up heaters will only dissipate, if the Primary Bus is powered, which requires Sun irradiation on the comet to operate the solar arrays. Without dissipation the compartment temperature will drop until the comet environmental temperature. When the Lander is still attached to the Orbiter and powered from the Orbiter-LCL 15A/B, an additional heater set will also dissipate. Power Enough Mode: This mode follows the Wake-up mode, the Lander Primary Bus is powered, but the voltage is still below 18.5V, which correspond to a non-sufficient power situation. The available power is not lost, since special Power Enough loads are used to dissipate and heat the internal compartment. Stand-by Mode: The Lander is operational, since the Lander basic operational electronics (PCU, CDMS and one TCU) are connected to the Primary Bus and powered. In this mode thermal control will be performed from the dissipation of the activated units. If the temperature of the internal compartment drops below the TCU set-points, the respcetive TCU heaters will also dissipate. Operational Modes: These modes define Lander operation of Experiments. ###########TO BE COMPLETED BY SONC ############ " END_OBJECT = INSTRUMENT_HOST_INFORMATION OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "BIBRINGETAL2007B" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO END_OBJECT = INSTRUMENT_HOST END