Exomars Mission and Spacecraft Architecture
- Paper number
IAC-07-A3.3.02
- Author
Mr. Vincenzo Giorgio, Thales Alenia Space, Italy
- Coauthor
Mr. Giacinto Gianfiglio, European Space Agency (ESA)/ESTEC, The Netherlands
- Coauthor
Dr. Alessandro Gily, Alcatel Alenia Space Italia, Italy
- Coauthor
Mr. Carlo Cassi, Alcatel Alenia Space Italia, Italy
- Year
2007
- Abstract
FOREWORD The preliminary design of the Exomars Mission, Spacecraft and Modules is being performed in the frame of the ongoing phase B1. A System Requirements Review is currently being held between ESA and the Mission Prime Contractor (AAS-I) to review the top level requirements and the associate preliminary design. An Implementation Review will then follow to be conducted such that Participating States can decide the final mission configuration, the final Payload Complement and the launch date of the mission. This Irev is presently planned for May 2007. MISSION DESCRIPTION Exomars has a technology demonstration objective and a scientific objective. As part of the first objective the mission shall demonstrate feasibility of safe landing (Entry Descent Landing System) on the Mars surface, mobility (Rover) and access to sub-surface (Drill). The scientific objectives are: to search for signs of past and present life on to characterise the water/geochemical environment to study the surface environment and identify hazards for the future missions to investigate the planet’s deep interior. The space segment of the mission consists of: Orbiter, accommodating 30-kg, nadir-pointing instrument payload (in the “baseline” scenario this function is performed by a NASA satellite), to operate for a continuous period of at least two Martian years Descent Module, including the 20-kg Geophysics/Environment Science Payload (GEP) surface station Rover, a semi-autonomous surface mobile platform including the Pasteur scientific payload, to operate for a Nominal Mission of at least 180 sols on the Martian surface. This Nominal Mission shall be completed before the onset of the next global dust storm season. ExoMars shall be able to land and perform its scientific mission at any longitude, in a latitude band between 15ºS and 45ºN. The mission shall safely deliver the Rover within 25 km (15 km goal) of the target landing site; the Rover, in turn, shall be able to travel at least 25 km on the Martian surface. The Rover shall implement a two-way, high-speed communication link with the Orbiter, using every available Orbiter pass to transmit data and receive commands. This proximity link shall be compatible with communication services provided by NASA satellites. The Orbiter shall relay to Earth science and housekeeping data generated by the Rover, and relay back to the Rover Earth-generated commands and software updates. The performance of the links shall be such that the 1Gbit of data collected on the surface in an Experiment Cycle can be sent to the Orbiter and then transmitted to Earth. The DM shall transmit telemetry to the Orbiter during the descent using the UHF band. It shall also send tones directly to Earth by means of the X-band link. The mass budgets of the various launch options shall show a 5ExoMars will also validate the technology for safe Entry, Descent and Landing (EDL) of a large size spacecraft on Mars, the surface mobility and the access to subsurface. TRADE-OFFS As part of Phase B1, the Prime Contractor was asked to study a number of different scenarios for the ExoMars mission. The basic scenarios are defined as follows.
Baseline scenario The Baseline scenario features launch in 2013 of a Composite Spacecraft made up of Carrier and Descent Module, using a Soyuz 2-1B launch vehicle. In this scenario the Data Relay Satellite used to support the mission is the NASA Mars Reconnaissance Orbiter (MRO).
Option 1 scenario The option 1 scenario is the same as the baseline, except for the data relay function, which is performed by a European Mars Telecommunications Orbiter (MTO), launched separately with another Soyuz vehicle. It is under study the possibility to embark a scientific payload as well.
Option 2 scenario In the option 2 scenario, a composite of Orbiter and Descent Module is launched in 2013 by an Ariane-5 ECA vehicle. The Orbiter performs the carrier functions until delivery of the DM; thereafter it performs the data relay function from Mars orbit. This scenario shall also be available in the option of launch in 2015/2016.
The above basic scenarios allow numerous variants and options as concerns launch window, Earth escape strategy, Mars trajectory, Mars entry and landing, telecom orbit, and so on. This paper outlines the main selection factors. Then, it describes the proposed Soyuz and Ariane 5 mission profile and its rationale. SPACECRAFT DESCRIPTION The Soyuz Composite includes the Carrier Module (CM) and the Descent Module (DM) with the Rover.
The structural configuration selected after a dedicated trade-off is based on a 1194-mm central tube hosting a single 700-litre MON tank surrounded by four MMH tanks, the 400N main engine and 6+6 10N thrusters. This configuration is not compatible with passive spin stabilization during the cruise. Therefore the composite is 3-axis stabilized, with body-mounted solar array and sun-pointing attitude. At the end of the cruise, to release the DM, the composite is spun-up to 5 rpm.
A tightly integrated command and control architecture is implemented, minimizing mass while guaranteeing independency in the CM, DM and Rover developments. The solution bases on a shared computational capability (one computer, with proper redundancy, for CM and DM) and another dedicated computer, with its own redundancy, in the Rover. The RF subsystem has all transponders (1 main and 1 redundant for both UHF proximity link and X-band link to Earth) located in the Rover module, as this element is the final user of the subsystem. The Ariane-5 Composite includes the Orbiter Module (OM) and the Descent Module (DM) with the Rover.
The DM interfaces at six discrete separation points with the 1666 mm central tube of the Orbiter module. The same tube interfaces on the bottom side with the Ariane 5 adapter. The 400N main engines are accommodated below the floor panel, inside the interface ring. Eight 10N bipropellant RCS thrusters are mounted in four redundant clusters at the floor panel edges.
The composite is 3-axis stabilized, with Orbiter body-mounted solar array and sun-pointing attitude.
Different from the Soyuz composite, the Orbiter module is a fully fledged spacecraft. The command and control architecture is designed for working in both composite and stand-alone configuration with minimum changes. During the cruise, the Orbiter CDMU is the active one; the Rover and DM-CDMU will be switched on and interrogated through the 1553 bus from time to time for health checks and software uploads.
The Rover and DM RF units are the same as in the Soyuz configuration (X band transponder and UHF transceiver located in the Rover), while the Orbiter has its own X- band subsystem to support the Earth link in the cruise and the Mars operations, as well as its own UHF system to receive the data transmitted from the DM during the EDLP and the Rover during the Mars operations. Since the Orbiter data relay can be used in the descent and landing phase, the amount of data that can be returned to Earth in this phase is much greater than in the Soyuz configuration, and the 30-kpbs required data rate can be met.
The option 1 ETO (European Telecommunication Orbiter) is composed of two modules, the ETO-Propulsiom Module and the proper ETO Orbiter .
The ETO-PM is based on the Carrier Module (CM) used for the parallel Carrier+DM mission. It is a propulsion stage to carry the propellant for the escape and DSM burns, and it is separated after those burns have been performed. Thus, the structure can be simplified and the on-board avionics can be removed, leaving the commanding to the ETO orbiter. The main advantage of this approach is the improved dry mass budget available for the Orbiter, as the benefit of staging is greater than the penalty of duplicating the thruster equipment.
The proposed design can accommodate up to 2200kg of propellant, 1540kg of which in the Carrier. All tanks are off-the-shelf.
The configuration of the ETO-Orbiter is derived from a standard GEO platform. The structure is based on a shear web design with propellant tanks accommodated inside the web and the subsystem components mainly accommodated on the outer structural panels. The Unified propulsion system (MON-MMH) uses the same set of equipment as the Carrier, including a 400N engine, and tailored off-the-shelf tanks.
The MTO avionic architecture is very close to that of the Carrier/DM spacecraft composite, with the objective of using the same FDIR logic and possibly similar AOCS on-board software. A standard CDMU design is envisaged, tailored with 12 Gbit mass memory and TT C boards adapted to support X-band and UHF communications.
The AOCS sensor set comprises star trackers (two optical heads connected to a common electronics box), one internally redundant IMU unit, and six coarse sun sensors.
- Abstract document