PDS_VERSION_ID = PDS3 LABEL_REVISION_NOTE = "2006-11-21: A. Hulsbosch and O.Witasse, VSOC Team. 2006-12-01: Revised by Maud Barthelemy. 2006-12-14: 70 characters width update. 2008-04-28: remove non ascii char, MB 2009-02-10: editorial by JLV. 2011-01-07 NAIF:SEMENOV minor fixes to make ingestable into PDS catalog; " RECORD_TYPE = STREAM OBJECT = INSTRUMENT_HOST INSTRUMENT_HOST_ID = VEX OBJECT = INSTRUMENT_HOST_INFORMATION INSTRUMENT_HOST_NAME = "VENUS EXPRESS" INSTRUMENT_HOST_TYPE = "SPACECRAFT" INSTRUMENT_HOST_DESC = " This document largely includes sections of the following articles: The Venus Express Spacecraft System Design, by P. Sivac and T. Schirmann [SIVACAL2007]. Instrument Host Overview ======================== Venus Express is the second of the so-called Flexi-missions of ESA and with some differences imposed by the environment in orbit around Venus is, from a spacecraft design point of view almost an exact replica of the first Flexi-mission Mars Express. The satellite was launched on 9 November 2005 on a Soyuz rocket from Baikonur and was injected into its trajectory towards Venus using the Fregat upper stage. After a journey of about five months, Venus Express was inserted in orbit around the planet on 11 April 2006 and started its nominal scientific mission on 4 June 2006. The major objective of the Venus Express spacecraft design was to cope with the specific mission requirements of a spacecraft orbiting Venus (mainly associated with the much higher temperatures) whilst maximising the reuse of the Mars Express design and so minimising the development risks in order to guarantee the readiness of the spacecraft for the launch window. Mechanical design ================= The mechanical design was driven by the following considerations: - reuse of the Mars Express mechanical bus as far as possible; - the specific constraints of the Venus Express mission; - the need to minimise the spacecraft dry mass and optimise the location of the centre-of-mass. The reuse of the Mars Express mechanical bus (structure and propulsion system) minimised the development risks and helped to secure the programme very tight schedule. Drawing on the Mars Express qualification, the core structure design remained basically unchanged, which allowed qualification by similarity. The modifications to the secondary structure were strictly limited to accommodation of the new or modified units. Reusing the bus also meant that most of the Venus Express units have the same mechanical environment as on Mars Express. The main mechanical design modifications relate to the payloads, the additional High Gain Antenna and the constraints from the thermal design. As on Mars Express, the overall spacecraft mass was close scrutinised. The maximum mass of the spacecraft (including propellants) allowed by the launcher was agreed at 1270 kg. Through careful mass management, the propellant tanks could be filled to their maximum capacities. Close attention was paid to the centre-of-mass and the alignment of the main engine. During the main engine burn for Venus orbit insertion, the disturbing torques imposed on the spacecraft were directly linked to the offset between the direction of thrust and the centre-of-mass, which migrated as propellant was depleted. To improve the control torque margins, strict control of the centres-of-mass of individual units was maintained and balance masses were added for fine adjustment. Thermal control =============== The spacecraft thermal control subsystem is required to maintain the temperatures of spacecraft elements within their allowed temperature ranges during all mission phases. The spacecraft equipment fall into two categories namely: a) The units that are collectively controlled by the overall spacecraft thermal control b) The units that have their own thermal control such as individual heaters or radiators. The thermal control design of Venus Express is based on a robust and passive concept having a maximum commonality with Mars Express but having some specific design modifications in order to cope with the hot environment at Venus. The two main differences with Mars Express are: a) A harsher thermal environment. The solar constant is almost four times that of Mars. Furthermore, the Venus albedo makes a significant contribution to the overall flux on the spacecraft during the observation phase around the pericentre. b) Since the orbit of Venus is closer to the Sun than the one of the Earth, it is not possible to keep one wall of the spacecraft permanently in the shadow during communications with the Earth having a single antenna. For this reason, having analysed the allowable Sun aspect angles on the various faces of the spacecraft, it has been decided to include a second smaller high gain antenna on the top floor of the spacecraft. Most of the spacecraft units are collectively controlled inside thermal enclosures that are naturally created by the spacecraft mechanical configuration. The heat balance is controlled by radiators and heaters, allowing the unit temperatures to be maintained within their defined limits. The unit temperature control is achieved through the use and the selection of flight-tested materials used on numerous spacecraft. The key features of the thermal control are summarised as follows: a) Optical solar reflectors radiators which are electrically grounded to the aluminum sandwich panel face sheet and dissipate internal heat towards space b) Dedicated radiators provided for both VIRTIS and PFS which require low temperatures for operations c) Platform high dissipative units mounted directly behind the radiators in order to provide a good conductive path from unit to radiator panel d) Heat pipes implemented under the PCU and PDU units in order to evenly spread the high PCU thermal dissipation e) Thermal straps used to connect the Reaction Wheels and those payload units requiring a dedicated radiator (i.e PFS, SPICAV/SOIR,VIRTIS) Heaters are managed either under software control or by thermostats. Chemical propulsion subsystem ============================= The Chemical Propulsion Subsystem (CPS) is a helium-pressurised bipropellant system, using monomethyl hydrazine (MMH) as the fuel and nitrogen tetroxide (NTO) with 3% nitric oxide as the oxidiser. The main engine, essential for Venus orbit insertion, has a thrust of 414 N and a specific impulse of 317 s. Four pairs of 10 N thrusters (four primary, four redundant) provide trajectory corrections and attitude control and reaction wheel unloading. They are the same as those used on Mars Express. The CPS operated in a constant-pressure mode during main engine firings for the capture manoeuvre and the first part of the apocentre reduction manoeuvre, using a regulated helium supply. The CPS comprises two subsystems: the pressurant subsystem and the propellant feed subsystem. The helium pressurant subsystem, commonly referred to as the 'gas side', has two sections: high-pressure and low-pressure. The highpressure gas side comprises a 35.5-litre helium tank, normally-open and normally-closed pyrovalves, a high-range pressure transducer, a fill & drain valve, and a test port to support integration activities. This section has a maximum expected operating pressure (MEOP) of 276 bar. During all ground operations and launch, it was isolated from the pressure regulator by a pair of normally-closed pyrovalves. These are arranged in parallel for redundancy. The low-pressure gas side comprises a pressure regulator, non-return valves, a pair of low-flow latch valves, a low-range pressure transducer, normally-closed pyrovalves, and test ports and fill & vent valves. This section has an MEOP of 20 bar, controlled by the regulator that senses downstream pressure. The propellant feed subsystem, commonly referred to as the 'liquid side', supplies propellants to the main engine and thrusters. It comprises a pair of 267-litre propellant tanks, normally-open and normally-closed pyrovalves, propellant filters, low-range pressure transducers, main engine, reaction control thrusters, and test ports and fill & drain valves. This section is pressurised with helium from the low-pressure gas side, and has an MEOP of 20 bar. Power Subsystem =============== Electrical power is provided by two solar wings equipped with triple-junction GaAs cells. The array is oriented towards the Sun by two Solar Array Drive Mechanisms (SADM). During eclipses, power is provided by three lithium-ion batteries that recharge after the eclipse. Power management and regulation is performed by the Power Control Unit (PCU) that provides a regulated 28 V main bus. The PCU uses a Maximum Power Point Tracker (MPPT) in order to operate at the maximum power output of the solar array, which avoids the need to oversize the solar array to cope with both near-Earth and Venus orbit conditions. Battery management is performed using three Battery Charge and Discharge Regulators (BCDRs) under the control of a Main Error Amplifier (MEA) control loop. The resulting +28 V regulated bus is distributed to all spacecraft users by a Power Distribution Unit (PDU) featuring Latching Current Limiters (LCLs), which protect the bus from overcurrents at unit level. The PDU is also responsible for generating the commands for firing the pyrotechnics. The solar array consists of two identical low-weight deployable wings, each having two solar panels pointed towards the Sun by a one degree-of-freedom SADM. When stowed, each wing was clamped to the spacecraft side panel on four hold-down points and release mechanisms. For deployment (which was performed autonomously after launch as part of the separation sequence), redundant pyrotechnic bolt cutters released each wing individually. In order to meet the stringent requirements associated with the Venus radiation environment, the chosen solar cell technology was GaAs with 100 micrometer cover glass. The maximum array current is 18 A per wing. The total array power values are of the order of 820 W near Earth and 1400 W at Venus (end-of-life). Three batteries supply the spacecraft power if either the solar array is not illuminated by the Sun or if the power demand is higher than can be generated by the array. The energy is stored in three identical 24 Ah low-mass Li-ion batteries with a total capacity of around 500 Wh. Each has 16 parallel strings of six serial 1.5 Ah cells. The batteries are identical to those on Mars Express. Radio-frequency communications subsystem ======================================== The radio-frequency (RF) communication subsystem consists of a redundant set of dual-band transponders operating in both S-band and X-band for either the uplink or the downlink. The antennas on the spacecraft are: - two Low Gain Antennas (LGA), used primarily during the Launch and Early Operations Phase (LEOP), operating in S-band for omni-directional reception and hemispherical transmission; - the dual-band HGA1, operating in S-band and X-band for high-rate telemetry and telecommand; - the single-band offset HGA2, operating in X-band only, for high-rate telemetry and telecommand. All nominal operations are performed at X-band. Selection of which HGA to use depends on the mission phase and, particularly, the relative positions of the Earth, spacecraft and Venus. To maintain thermal control for instruments, solar illumination of the -X side of the spacecraft (opposite to HGA1) is minimised. With steady-state Earth communications, the spacecraft Z/X plane remains in the Sun-spacecraft-Earth plane. This means that: - no Sun impinges on the lateral sides (+/-Y sides); - the solar array can be pointed towards the Sun; - HGA1 or 2 can be pointed towards Earth; - the cold side of the spacecraft (-X) remains facing cold space. Before the Sun starts to impinge on the cold side, as the spacecraft and Venus orbit the Sun, the spacecraft is flipped to point the opposite antenna towards Earth. Ground stations =============== The ground station used throughout all mission phases are the ESA Cebreros 35m station, complemented by the ESA New Norcia 35m station, to support Venus Orbit Insertion and VeRa campaigns, and the Kourou 15m station during LEOP. Use of the NASA Deep Space Network (DSN) is envisaged for emergency cases or for special campaigns. Payload ======= The Venus Express payload includes seven instruments. Five of them are re-used from the Mars Express and Rosetta projects with small modifications: - ASPERA analyzer of space plasma and energetic neutral atom imager; - PFS a high-resolution IR Fourier spectrometer; - SPICAV/SOIR a versatile UV-IR spectrometer for solar/ stellar occultation and nadir observations; - VeRa a radio science experiment; - VIRTIS a sensitive visible and near infrared spectro-imager and high resolution spectrometer. These experiments are complemented by two small units specifically developed for the Venus Express mission. They are: - Venus Monitoring Camera (VMC) - magnetometer (MAG). Details about these instruments can be found in the INSTRUMENT.CAT files that are included in the experiment data sets. Payload operations ================== Payload scientific operations are conducted according to the observation scenarios or feasible combinations of the defined ten science cases whenever they are compatible with spacecraft and ground segment resources. Payload observation scenarios: I Observation scenario 1 Pericenter observation 2 Off-pericenter observation 3 Apocenter global spectral-imaging by VIRTIS 4 VeRa bi-static sounding 5 SPICAV stellar occultation 6 SPICAV/SOIR solar occultation 7 Limb observation 8 VeRa Earth radio occultation 9 VeRa solar superior and inferior conjunction 10 Venus Gravity Anomaly by VeRA Data downlink is performed during daily station passes. It is assumed that during the station pass science observations do not interfere with the Earth pointing. Under nominal conditions the station passes is such that an average of 8.5 hours continuous is available daily for stored data downlink. An average bandwidth of ca 7 kbps is allocated to the downlink of stored spacecraft telemetry, assuming a continuous generation rate of 2.5 kbps throughout the 24 hours for spacecraft housekeeping telemetry. Radio Science (VeRa) observations are conducted for Earth occultations with the spacecraft inertially fixed, for short periods when the line of sight between spacecraft HGA antenna and Earth ground station pass through the Venus atmosphere. In support of dual frequency measurements (X/S-band) above and beyond the Cebreros 35 m (X-band) occasional use is made of the New Norcia 35 m and DSN 34 and 70 m ground stations (S/X-band). As far as possible, the orbit phasing will be such that SC entry and exit of Earth occultation will occur during NNO passes. During the VeRa radio occultation the spacecraft is required to perform a slew manoeuvre to compensate for ray bending due to atmospheric refraction. Attitude and Orbit Control System ================================= The characteristics of the mission and the fact that the spacecraft has fixed High Gain Antennas and a single rigidly-mounted main engine means that there are demanding manoeuvering requirements. Attitude ascertainment is provided by two Star Trackers (STTs) which in turn are supported by redundant Inertial Measurement Units (IMUs), and a Sun Acquisition Sensor (SAS). Attitude control is established by a reaction wheel assembly supported by the attitude thruster system. Star Trackers ------------- The autonomous star tracker with its extensive star catalogue ensures that accurate attitude estimation can be achieved in almost any position. It is the main optical sensor of the AOCS and is used at the end of the attitude acquisition following each manoeuvre to acquire the 3-axis pointing required for almost all nominal operations. The star tracker includes a star pattern-recognition function and can perform attitude acquisition autonomously. Venus Express has two star trackers aligned with an angle of 30 degrees of arc between their optical axes for operational redundancy and to ensure that one can always see a recognizable region of the sky. Inertial Measurement Units -------------------------- Each of the two IMUs has a set of three gyros and three accelerometers aligned along three orthogonal axes. The AOCS can use either the three gyros of the same IMU or any combination of three gyros among the total of six. Only a full set of accelerometers from one single IMU is used; they are essential to improve the accuracy of manoeuvres performed with the main engine. The gyros are used during attitude acquisition phases for rate control, during observation and communication phases to ensure the required pointing performance and during the trajectory correction manoeuvres for control robustness and failure detection. The non-mechanical technology removes the possibility of mechanical failure during the mission. Sun Acquisition Sensor ---------------------- Two redundant SASs mounted on the spacecraft central body ensure pointing in the Sun Acquisition Mode at the time of first acquisition after launch or any subsequent reacquisition in the case of an FDIR activity following an onboard failure. The Mars Express sensors were modified, using different solar cells with ceramic backing in order to cope with the thermal environment at Venus. Reaction Wheel Assembly ----------------------- The reaction wheels are used for most of the attitude manoeuvres for specific activities such as Earth communication and scientific observations; they provide flexibility and reduce overall propellant consumption. The angular momentum of the wheels has to be carefully managed from the ground; regular off-loading keeps the wheel speeds within performance limits. The RSA comprises four individual wheels in a skewed configuration to manage the spacecraft momentum in all three axes. Although nominally all four wheels are used for operations, this configuration, which is identical to Mars Express, allows full performance with any three-wheel configuration. AOCS Modes ---------- The AOCS includes several modes for attitude acquisition/ reacquisition, the nominal scientific mission and orbit control. The attitude acquisition and reacquisition sequence has two basic modes: 1) the Sun Acquisition Mode, pointing the X-axis and the solar array towards the Sun in order to guarantee power. 2) the Safe/Hold Mode, which completes the acquisition phase and provides the final 3-axis pointing, with one High Gain Antenna pointing towards the Earth to guarantee telemetry and telecommanding. Trajectory correction manoeuvres are performed using two modes: 1) the Orbit Control Mode, for small corrections using the attitude thrusters; 2) the Main Engine Boost Mode, for major manoeuvres using the main engine; To ensure a smooth transition between the thruster-controlled modes and wheel-controlled modes there is also: 1) a Thruster Transition Mode Data Handling Subsystem ======================= The Data Management System (DMS) performs all the data handling functions for the spacecraft, including: 1) telecommand distribution throughout the spacecraft; 2) telemetry data collection from the spacecraft and data storage; 3) overall supervision and monitoring of the spacecraft and payload functions and health status; 4) timing functions, including the distributions of time and synchronization information. The DMS is based on a dual-processor architecture embedding standard communication links. The DMS includes four identical processor modules, located in two fully redundant Control & Data Management Units (CDMSs). One processor module is dedicated to the DMS software (in charge of the management of the platform subsystems), while the other is allocated to the AOCS software (in charge of acquisition and control of all platform sensors, actuators and Solar Array Driving Mechanism (SADM)). The data handling architecture is organized around the two CDMUs. They are in charge of controlling ground command reception and execution, onboard housekeeping and science data telemetry storage and formatting for transmission. Onboard data management, control-law processing and execution of onboard control procedures are also part of their function. Each CDMU features two MA3-1750 processor modules, each processing either DMS or AOCS software. A builtin failure operational reconfiguration module ensures system-level Failure Detection, Isolation and Recovery integrity and autonomously reconfigures the CDMU processor units as necessary. The CDMUs communicate directly with the Dual Band Transponder that is part of the RF Communications System. Three other units are part of the data handling subsystem: 1) the Remote Terminal Unit (RTU) is the interface between the DMS processor module and platform units (Power Control and Distributions Units) and payloads; 2) the AOCS Interface Unit (AIU) is the interface between the AOCS processor module and the sensors, actuators and solar array drive electronics; 3) the Solid State Mass Memory (SSMM) is a file-organized mass memory with 12 gigabit of storage that is used to store the housekeeping and science data collected by the CDMU; it also collects science data directly from the VIRTIS and VMC instruments. The use of SSMM is dictated by the same store-and-forward concept employed for Mars Express, meaning that every orbit is divided into two principal phases: an observation phase, where the instruments are pointed towards Venus, and a communication phase where a HGA is pointed towards the Earth. To support this, all scientific and housekeeping telemetry are stored in the SSMM. Failure Detection, Isolation and Recovery ========================================= As a deep space mission, Venus Express requires a high level of onboard autonomy because of the time needed for ground intervention in the event of an onboard anomaly or the extended periods when communications are not possible owing to the respective positions of the Sun, Venus and Earth. The automatic FDIR function is largely recurrent from Mars Express, with some adaptations. The FDIR function handles any anomalies onboard with a first goal of returning to the same spacecraft mode in order to preserve operational integrity. If this is not achievable, then the priority is to preserve spacecraft integrity and safety. In this case, the spacecraft autonomously enters safe mode, which ensures that power is available from the solar array, communication with Earth is available for diagnostics and recovery, and all non-essential loads have been switched off. In this way, spacecraft safety is ensured until operators on Earth can identify and correct the anomaly and proceed with the mission. " END_OBJECT = INSTRUMENT_HOST_INFORMATION OBJECT = INSTRUMENT_HOST_REFERENCE_INFO REFERENCE_KEY_ID = "SIVACAL2007" END_OBJECT = INSTRUMENT_HOST_REFERENCE_INFO END_OBJECT = INSTRUMENT_HOST END