Mission Analysis at EADS Astrium [email protected] [email protected] [email protected] [email protected] [email protected] [email protected] Page 1 Agenda 1. Introduction 2. Interplanetary Missions 3. Geostationary 4. Navigation satellites 5. Earth observation 6. Special utilities Page 2 1. Introduction Page 3 Mission Analysis at Astrium • Mission Analysis is a core engineering discipline – It is an integral part of an efficient spacecraft and mission design • Integrated Mission Analysis and Systems Engineering activities allow an efficient, iterative approach – This wholly integrated approach takes place within Astrium. Recent examples are Bepi-Colombo feasibility and Lisa- Pathfinder implementation • Mission Analysis is managed within Astrium’s Central Engineering group – The main engineering specialist skills group • Allows cross directorate and cross disciplinary activities – Full co-ordination of activities and research across France, Germany and UK • Close project collaboration with ESA/ESOC Page 4 Mission Analysis at Astrium (2) • Main applications of mission analysis are for – Earth Observation Navigation and Science – Communications • Many core ESA Science and EO missions supported directly in feasibility, feasibility and implementation phases. Examples are: Lisa-Pathfinder Feasibility and Implementation Bepi-Colombo Feasibility to date SOLO Pre-feasibility to date Venus Express Feasibility Rosetta Feasibility LISA Pre-feasibility to date SWARM Feasibility and Implementation Aeolus Implementation Page 5 Mission Analysis at Astrium (3) • Examples of future ESA missions and concepts direct support – Technology reference missions: · Mission designs for every planet of the solar system · Solar system escape missions · Detailed exploration of the Jovian system · Asteroid sample return missions – Asteroid exploration studies · Ishtar, Apies – Darwin + Xeus – Aurora (Exomars and MSR phase A1) – Lunar Lander/LEDA – Hyper Page 6 Flight Dynamics OPERATIONAL EXPERIENCE for GEO Commsats (Astrium in orbit deliveries) MISSION PLATFORM DATE LOCATION HISPASAT 1A E2000 September 1992 Arganda (Spain) HISPASAT 1B E2000 July 1993 Arganda (Spain) ORION F1 E2000 November 1994 Stevenage (UK) NILESAT 101 E2000 April 1998 Toulouse (France) ST-1 E2000+ August 1998 Stevenage (UK) NILESAT 102 E2000 August 2000 Toulouse (France) ASTRA 2B E2000+ September 2000 Toulouse (France) HELLAS SAT 2 E2000+ May 2003 Toulouse (France) AMAZONAS E3000 August 2004 Toulouse (France) ANIK F1R E3000 October 2005 Toulouse (France) ARABSAT 4A (1) E2000+ March 2006 Toulouse (France) Page 7 (1) Controlled re-entry following launcher failure Astrium FDS OPERATIONAL EXPERIENCE (In orbit deliveries or support with Astrosat platform MISSION Involvment PLATFORM DATE LOCATION Customer FORMOSAT-2 Astrosat 500 May 2004 Hsin-chu (Taïwan) Support Essaim (4 sat) Support Astrosat 100 Spring 2005 Toulouse (France) In orbit THEOS Astrosat 500 July 2007 Bangkok (Thailand) Delivery In orbit SPIRALE (2 sat) Delivery Astrosat 100 2008 Toulouse (France) Page 8 The Design Cycle: End to end mission Design •Preliminary mission design –Orbital Environment and/or science requirements –Transfer options and useful mass analysis •Detailed mission Design –Definition of operational orbit –Detailed transfer analyses/ trajectory and system optimisation •LEOP –Flight Dynamics development and application Page 9 2. Interplanetary Missions Page 10 Orbit and Environment Modelling • A general purpose tool has been developed • ORBITVIS • Planetary Orbit propagators • Mathematical models of environment effects – Atmosphere, Gravity • Planetary surface coverage • Lander contact analyses • Earth Ground station coverage • Manoeuvre design and simulation Page 11 ORBITVIS • Used both for preliminary and detailed mission design • High precision, high speed orbit propagators • Selectable environmental models (eg Atmospheres, Gravity models) • Multiple spacecraft simulation (formations or constellations) • MS Windows based, menu driven Page 12 ORBITVIS (2) • Examples (R to L): • Lisa-Pathfinder transfer from LEO to the Earth-Sun L1 Lagrange point: optimum apogee raising strategy • Satellite-Lander visibility simulation in Mars fixed frame • Aerobraking simulation at Mars from High elliptical Mars orbit to Low circular orbit • Venus Express Capture Page 13 Interplanetary Transfer orbit optimisation facility: ORBITOPTIM A set of transfer orbit optimisation tools interfaced with system optimisation tools Heliocentric transfers with special developments for Earth and Jupiter centred transfers Optional gravity assist manoeuvres (by patch conics or multiple gravity field modelling) Optional multiple low or high thrust arcs Target orbits about specified planets (From Mercury to Pluto + selected Asteroid/Comet targets) Optimises thrust vector steering profiles, locations and durations of thrust arcs. Can determine power constrained optimal thrust/Isp relationship Optimises approach and departure orbit parameters Can utilise 3 body gravity effects (Gravitational capture - Weak Stability Boundary effect) Page 14 ORBITOPTIM Features Direct Multiple shooting method Segmentation in space and time Facilitates initial solution estimation. High efficiency state transition matrix evaluation Variational calculus methods Detailed propulsion system modelling Electric propulsion models Solar sail models Can solve the optimal mass transport problem in conjunction with System optimisation tools Propulsion system design and operational parameters can be optimised Menus driven, Windows based system Page 15 ORBITOPTIM : Designing an optimal interplanetary transfer • Select preferred launch epoch and plan first Planet to Planet transfer • Select approximate epochs of subsequent fly-bys/Gravity Assists and estimate excess hyperbolic speed • Define propulsion system type for manoeuvres • Optionally make first estimates of any expected manoeuvres • Gravity assist design ‘wizzard’ assists in fly-by ephemeris estimation • Perform multi-starting point, forward/backward propagation from mission start, mission end and each fly-by • Now have a series of disconnected interplanetary trajectories • Select objective type, ie minimum fuel, minimum DeltaV (per mission phase) • Having generated an initial solution, perform trajectory optimisation. • NLP iterations remove trajectory discontinuities and optimise the objective Page 16 Examples of application of the Interplanetary Transfer orbit optimisation facility (1) • • L-E-V-M-(M3:2 Parametric studies on resonance)-M – Muliple GA, low thrust (M5:4 resonance)- transfers to Mercury M –(M1:1 resonance) • Include standard, Bepi- Colombo like transfers • L-E-V-(1:1 – (L-E-V-V-M-M GAs) resonance-)V-M- • Analysis of alternative (M3:2 resonance)- M – (M4:3 transfer types with different resonance) )-M – GA sequences (M1:1 resonance) • Explore DeltaV/time relationships • L-E-V-(1:1 resonance-)V-M- (M3:2 resonance)- M – (M5:4 resonance) )-M – (M1:1 resonance) (Langevin solution) Page 17 Examples of application of the Interplanetary Transfer orbit optimisation facility (2) • Bepi Colombo mission options with baseline mission variants • Sequence is – L-E-V-(1:1 resonant)-V-M- M • Number of revs V-M is related to on thrust selection and DeltaV minimisation • Optimal transport systems were derived Page 18 Examples of application of the Interplanetary Transfer orbit optimisation facility (3) • Capture at Mercury after low thrust transfer is a potential single point failure • An innovative strategy can be used employing a gravitational capture – Passage close to the Mercury- Sun Lagrange point allows a reduction in orbital energy 200000000 250 200 150000000 • Temporary capture is achieved 150 without manoeuvres (for 100 100000000 ) 50 g approximately one Mercury year) ) e m d ( ( sma a 50000000 0 e l True anom m g s • Total fuel mass is also reduced n -50 A 0 with such a strategy -100 -20 30 80 130 180 -150 -50000000 -200 Page 19 -100000000 -250 Time (days) Examples of application of the Interplanetary Transfer orbit optimisation facility (4) • Deriving SOLO Mission variants: • Nominal GA sequence with low thrust, fast transfer phase: – L-E-V-V-V-V-V-V • Increasing heliocentric inclination to 45 deg using additional SEP DeltaV • Intermediate Venus 3:2 resonant orbits used • Reducing mission duration by adopting Venus 2:1 resonant orbit Page 20 GENOPT: Evolutionary method for global optimisation of interplanetary missions • Generic technique based on genetic algorithms • Used for preliminary mission design activities, eg: • Optimisation of manoeuvre sequences in multi-gravity assist missions – Impulsive manoeuvre and low thrust cases • Optimisation of fuel consumption and launch strategy for the transfer of a constellation of spacecraft (after launch from a single launcher): – maximisation of the minimum spacecraft useful mass – Impulsive manoeuvre and low thrust cases • Optimisation of transfer propulsion system design: maximal transport problem Page 21 Geostationary Missions Page 22 Mission analysis for GEO • Mission analysis and design for: – Eurostar family – Military commsats • Launch and operations – Eurostar 2000 and 3000 series – Example of a SSTO. First commercial SSTO used for Orion 1 Page 23 Design Goals for the new FDS : QUARTZ++ • Design for SK and LEOP operations • Able to operate fleet composed of numerous satellites (including spacecraft of different types which all require specifics algorithms