ESA Science & Technology - Engineering

Planetary Orbiter

The Planetary Orbiter

The remaining side is made up of a large radiator, the size of which (1.5 m²) is driven by the internal power dissipation (200 W). The nadir pointing attitude law, plus a 180^o rotation every half Mercury year, keeps the radiator away from the Sun.

A deployable shield is used to protect against the planetary radiation. The shield is deployed at 120^o from the radiator to avoid solar back reflection into it, and it is large enough to block infrared radiation at any permitted view factor to the planet.

The optical instruments are recessed into the spacecraft and view the planet through dichroic mirrors. Multi-layer Insulation (MLI) sheets are used for the non-optical instruments. High-temperature MLI insulates the interior everywhere, except for the radiator and the instrument apertures.

The main externally mounted element is a deployable, two-axis articulated, 1.5 m High Gain Antenna (HGA), mounted on a short boom on the Zenith side. A suitable latch mechanism restrains the antenna at launch. Two data relay antennas (UHF dipole arrays) are mounted to the Zenith side (for communications with the Magnetospheric Satellite) and the Nadir side (for communications with the Surface Element). Three star sensors view through the radiator side, and are mounted to a stable bench-like structure, in common with the cameras.

Magnetospheric Orbiter

The Magnetospheric Orbiter

The magnetometers and search coils are mounted on a rigid boom. The electric antenna consists of two 30 metre long wires that mounted at right angles to the rigid booms. The wire booms are deployed by centrifugal force, but the rigid booms will be deployed in a controlled manner because of the large inertia ratio with respect to the satellite.

All experiments share a common electronics subsystem. Mass and volume are minimised by use of miniaturised advanced electronics. An integrated approach is also applied to the data handling and attitude control functions, which share a common processor. The mass memory (250 Mbit) is integrated into the central avionics unit. The AOCS includes sun and horizon sensors, and nitrogen thrusters for actuators. The data relay communications are performed by two UHF antennas mounted to the top and bottom sides, with TM/TC functions integrated in the common electronics.

Mercury Surface Element

The Mercury Surface Element (MSE) performs measurements on the surface of Mercury for at least one week. The lander is aimed at a latitude of 85^o in the terminator region where the mean surface temperature lies in the range -50 to +65^oC. Following the release of MMO, a burn of the 4000 N CPM thruster inserts MSE on a low-periherm (10 km) orbit. At the end of a 75-minute autonomous descent, a final braking manoeuvre is performed from an altitude of about 10 km until zero velocity is reached at a height of approximately 120 m above the surface. This manoeuvre is controlled with gyros/accelerometers and an optical range/range-rate sensor.

MSE (top) seperating from CPM

The MSE design is driven by severe mass constraints, and by the risk inherent in a landing in a largely uncharted and possibly shadowed area; 40 % of the terrain is estimated to be in shadow at a latitude of 85^o. The lander concept therefore relies on a primary battery with high energy density (1.7 kWh capacity). Operations can therefore be carried out even in a completely shadowed area for about one week. An optional small solar generator (about 0.5 kg) could nevertheless feed the load, or part of it, directly without a secondary battery, thus allowing extended surface operations.

The craft is fully insulated to cope with the low-temperature environment in a shadowed area. Should the landing occur in sunlight, a jettisonable cover would be expelled to enable a topside radiator to dump waste heat from MSE. The radiator is slightly recessed inside the MSE structure and is protected from the Sun for elevations of up to 20^o.

Data transmission between the modules

MSE's communications, data handling and power supply subsystems will consume about 1.4 kWh, leaving 300 Wh of primary energy for the payload. The scientific data are stored in a mass memory; either MPO or MMO can be used as a relay at each overhead pass. The MSE to MMO UHF link provides for a mean usable data rate of 8.7 kbit/s, corresponding to a total of 75 Mb for 7 days of operation (18 contact periods of 480 s). Nearly twice this total data volume (138 Mb) is provided by the MSE to MPO link (more frequent contacts due to the lower orbit). In addition, by appropriately phasing the MSE descent with the MPO orbit, it is possible to ensure the visibility of MPO from MSE for at least a portion of the descent trajectory, so that the low rate transmission of vital parameters can be maintained until landing.

Solar Electric Propulsion Module

The SEPM

SEPM is a simple rectangular prism. The main structural element is a central thrust cone which houses the Xe propellant tanks and transmits the loads to the launcher interface. Two deployable wings, consisting of several panels equipped with GaAs cells, provide the power required during the cruise; they are mounted on drive mechanisms to enable their orientation to be modified during the cruise.

Electric propulsion is optimal for slow cruise manoeuvres, but unsuitable for quick insertion needs. Moreover, the large cruise arrays can hardly withstand the near-Mercury thermal environment without major developments. Solar arrays derived from a standard design developed for communication satellites (with a maximum operational temperature of 150^oC), and optimised for extreme thermal conditions, have been adopted. The available electrical power and the temperature of the array increase as the spacecraft approaches the Sun. Once the temperature has risen to 150^oC, the array is progressively tilted away from the Sun direction, by up to 65^o. In this way, the power remains approximately constant, from 0.6 to 0.32 AU, at a level twice that available at 1 AU.

Main characteristics of the Solar Electric Propulsion Module

Chemical Propulsion Module

The CPM-MPO composite after jettison of SEPM

The Chemical Propulsion Module (CPM) hosts the bi-propellant propulsion system employed for attitude control during the cruise, Mercury orbit insertion, and MSE de-orbiting and descent. The attitude control functions are performed with a redundant set of eight 20 N thrusters, while the other manoeuvres are achieved with a 4000 N engine. Alternative engines with lower thrust (down to 1500 N) may also be used, depending on their availability.

CPM also serves as a structural interface between SEPM and the scientific modules, and carries attitude and orbit control system (AOCS) sensors and the CLAM-D camera head (for the MMO-MSE mission).

The CPM-MSE composite after seperation from MMO and interface cone

The CPM structure consists essentially of a main platform, which supports the elements of the propulsion system, and of a short interface cone, which caps the SEPM thrust cone and provides a mounting interface for MMO and MPO. The propellant tanks protrude under the platform into SEPM to minimise the total height of the system. The lower side of CPM is insulated to provide thermal protection after the jettisoning of SEPM.

MPO is mounted on CPM in such a way that it maintains the same attitude with respect to the Sun and planet as during the operational phase. In the MMO-MSE mission, MMO remains momentarily attached to the interface cone after separating from the CPM-MSE composite. The interface cone is jettisoned afterwards. MSE is supported by brackets which are welded to the propulsion tanks.

Main characteristics of the Chemical Propulsion Module