Airbus DS ESA Phase-0 L5 Spacecraft/Orbital Concept Overview. Emanuele Monchieri 6 th March 2017

Transcription

1 Airbus DS ESA Phase-0 L5 Spacecraft/Orbital Concept Overview Emanuele Monchieri 6 th March 2017

2 Airbus DS ESA Phase-0 L5 Spacecraft/Orbital Concept Overview Contents L5 Mission Outline Mission Concept Spacecraft Design Overview Design Drivers Payload Suite Platform Highlights Programmatics 2

3 L5 Mission Outline L5 Mission investigated as part of the P2-SWE-X Phase 0 Study Study activities span from observational requirements definition to mission design and preliminary cost assessment. An Operational Space Weather Monitoring Mission requires the capability to operate in severe SWE conditions L5 Mission provides a continuous stream of remote sensing and insitu data to support and significantly improve SWE forecasting capabilities Strong UK participation both on platform and on payloads sides 3

4 Mission Concept Launch Date: 2023, with 2024 as backup Transfer Strategy: multiple options available direct injection to escape orbit Injection into GTO followed by escape manoeuvre Injection into sub-lunar orbit, Lunar Gravity Assist, etc. Ariane 62 sequence for injecting into GTO Launchers: Ariane 62 Direct Transfer case: Insertion in orbit slightly larger than Earth orbit Stopping manoeuvre once at L5 (after 1.2 years) Operational Orbit: Sun-Earth L5 Mission Duration: 5-10 years Transfer Orbit for direct injection Sun Earth 3x15m Ground Stations supporting continuous P/L data download L5 4

5 Spacecraft Design Overview Design Drivers Objective To design a reliable operational platform capable to satisfy space weather requirements and operate in severe SWE conditions Design Drivers Re-use of existing technologies and processes to reduce costs and meet schedule requirements Robust design of platform and instrument. Heritage/lesson learned from other solar missions (SOHO, STEREO, Solar Orbiter) Space weather pointing and accommodation requirements from a large set of remote sensing and in-situ instruments Communication link established 24/7, 99% availability Large Propulsion System (large delta V manoeuvre for stopping at L5) Operational environments (launch, radiation, thermal, etc.) 5

6 Spacecraft Design Overview Payloads Suite Payloads Type Accommodation & Pointing Direction 1x Coronagraph Remote Sensing Inside the platform structure, 1x Magnetograph Remote Sensing looking at the Sun 1x EUV Imager 1x Heliospheric Imager Remote Sensing Remote Sensing Earth-facing side, looking at the Sun-Earth line 1x X-ray Flux Meter Remote Sensing On Sun-facing side 1x Magnetometer (2x sensors) In-Situ On 5m boom 1x Plasma Analyser In-Situ On Sun-facing side 2x Proton and Electron Detectors (HDRMs) In-Situ On two opposite sides of the platform, along the Parker and anti-parker lines 1x Radio Receivers In-Situ On the bottom side opposite to the Sun-facing side 60 deg EARTH 6

7 Spacecraft Design Overview Platform Highlights Three-axes stabilised and sun-pointing spacecraft with advanced AOCS capabilities to meet pointing requirements Deployable HGA Two main configurations supporting continuous high data rate from Sun-Earth L5: Fixed 1.6m High Gain Antenna Deployable 1.6m High Gain Antenna Re-use of Mars Express/Venus Express Platform Structure Large deployable solar panel(s) to generate ~kws of power Large Bi-propellant propulsion system, with >400N engine to perform large delta-v manoeuvres Long Deployable Magnetometer Rigid Boom (5m) for accurate magnetic field in-situ measurements Fixed HGA Maximised re-use of reliable electronics units designed for deep space missions (Solar Orbiter, Lisa PF, etc.) 7

8 VEX Platform Public Programmatics Mission ready to fly in 5.5 years, from B2 KO to Launch Campaign, no major show stoppers identified Cost-effective solution largely based on the experience matured in other solar and deep space missions and capable of meeting all observations requirements No pre-developments required, high TRL of selected equipment Highly reliable and robust platform design with focus on re-use as much existing hardware as possible to mitigate risks and optimise cost Although the mission is considered already viable with existing technology, further improvements (cost saving, mission performance, etc.) could derive from: Miniaturization/optimisation of instruments could lead to mass reduction, increased reliability/lifetime, improved performance Payload interface requirements (mechanical, electrical, data, thermal) frozen as early as possible in the programme Synergies with L1 Mission in terms of payloads (to improve data correlation) and launcher (dual launch option) SolO PFM RIU MEX HGA 8

9 Thank you! Any Question? Emanuele Monchieri 6 th March 2017