ESAIL proof of concept mission JuhaPekka Luntama Pekka

ESAIL proof of concept mission Juha-Pekka Luntama Pekka Janhunen Petri Toivanen 1

Outline 1. 2. 3. 4. 5. 6. 7. Introduction Mission objectives Magnetosphere Mission elements Expected mission results Demo mission schedule Summary 2

Introduction • The physical background of the electric sail concept has been carefully studied and simulated • Sail manufacturing and deployment techniques are under development • Remaining problem: Electric sail can not be tested or demonstrated on the Earth surface => A concept demonstration mission is needed • to verify the analysis and the simulation results • to demonstrate the feasibility of the sail deployment and control • to test advanced concepts to improve electric sail efficiency 3

Mission objectives • Main objectives: – Successfully deploy and operate an electric sail in space – Measure the acceleration of the spacecraft in different solar wind conditions – Test enhancement of the sail efficiency by electron heating • Secondary objectives: – Many technical and scientific objectives considered: • Monitoring of the electric sail behaviour in the dynamic solar wind conditions • Spacecraft attitude control • Characteristics of the solar wind near the sail • Dust particle monitoring • … – The secondary objectives will be carefully assessed and selected based on the mission partners and main mission profile => focus in strictly on the main mission 4

Earth’s magnetosphere • Electric sail does not work (at least well) within the magnetosphere • Even outside the magnetosphere the solar wind is disturbed e. g. in the foreshock region Þ apogee of the test mission orbit has to be well outside the magnetosphere Þ the shortest distance to undisturbed solar wind is towards the sun 5

Elements of a proof of concept mission • Pre-phase A analysis – Payload – Spacecraft bus – Orbit – Launcher – Ground segment – Lifetime – Budget 6

Test mission payload • Main payload: Electric sail prototype – Sail: 8 X 1 km aluminium four-fold Hoytethers – Mass estimates: • Tethers: < 0. 1 kg (25 µm) • Reels: 4. 0 kg • Electron gun + radiator: 1. 5 kg (40 k. V & 1 k. W) • High-voltage power source: 2. 0 kg • tether direction sensor: 2. 0 kg • Spinup thrusters: 3. 0 kg • Accelerometer: 0. 5 kg 2 km • Ion and electron detector: 1. 5 kg • PCU: 0. 5 kg • Total: 15 kg 7

Spacecraft bus requirements • Essential requirements: – Spinner: spin rate 3 min per rotation – 200 W electric power – Spin control during sail deployment – Ground link from 46 Re (telemetry and telecommand) – Propulsion for reaching final orbit – Tether reels minimum of 30 cm radial distance from the spin axis – Cooling for the electron gun • Other requirements – Depend on the mission secondary objectives 8

Spacecraft requirements analysis • Spinner => symmetrical spacecraft, fixed solar panel • Very small payload => spacecraft mass impacts mostly perigee kick motor sizing • Electronics radiation hardened due to solar particles and Earth radiation belts • Spinup thrusters and tether reels benefit from the radial distance from the spacecraft rotation axis • Spacecraft spin axis points approximately to the sun direction during the main mission => spacecraft body can be used to shield the electron gun 9

Test mission spacecraft outline • Mission requirements can be fulfilled with a relatively simple, small weight spacecraft • Spacecraft body should have a relatively large diameter and a large sun pointing surface => spherical or octagonal cylinder with a diameter of 1 m • Payload constraints on the spacecraft body are modest => final design will depend on the launch vehicle and potential secondary payload instruments 10

Orbit selection criterias • Essential requirements: – Apogee well outside the magnetosphere – Mission life time minimum of 1 month – No passes through densely populated satellite orbit regions (our spacecraft has effective diameter of 2 km) • Important aspects: – No need for orbit maintenance – Simple spacecraft design => spin axis point to the sun – Minimize launch cost • Nice to have: – Option to perform other space science observations 11

Other orbit aspects • Extremely elliptical orbits unstable due to the Moon => either active orbit control or short mission lifetime • Final orbit not reachable without a perigee kick motor => Spacecraft design more complex => Up to 75% of launch mass fuel => Longer and more complex LEOP phase due to orbit manoeuvres • High initial orbit (e. g. GTO) => less fuel needed => higher launch costs • Satellite visibility => ground station antenna location 12

Orbit candidates Equatorial orbit Apogee radius: Perigee height: Inclination: Orbit period: Low/medium inclination orbit 47 Re 2800 km 0 0 - 45 7 days Deceleration zone Sun Acceleration zone Moon orbit Bow shock 13

Launcher options • Final orbit requires the use of a perigee kick motor => launch to either LEO or GTO • Demo mission spacecraft: – dry mass << 100 kg – fuel from LEO to final orbit: 75% of the launch mass => launch mass 200 – 400 kg • Piggy-back opportunities to be exploited => GTO orbit orientation potential limitation • Dedicated small launcher allows mission lifetime optimisation 14

Ground segment • Apogee height of 47 Re allows spacecraft control even from a high latitude station • No satellite link during the perigee pass => Single ground station, operations during “office hours” • One potential scenario: – Satellite ground station in Sodankylä, Finland – Mission control center at FMI premises – Mission operations by FMI staff – LEOP supported by launch provider – Data processing and analysis by mission partners 15

Mission lifetime • Main limiting factors: – Orbit stability – Apogee direction • Main mission objectives can be achieved during one month of experiments • Conservative mission plan: => a three month mission with the “prime time” during the second month • Next suitable observation period in 9 months => main mission objectives do not support extension of the mission life time beyond 3 months 16

Mission “prime time” definition Mission end Prime time Mission start Launch and LEOP 17

Mission budget estimate • Spacecraft bus: 2 M€ • E-sail payload: 1. 5 M€ • Launch: 1 M€ • Mission operations: 0. 5 M€ • Notes: – The budget outline has been estimated by assuming that all components can be procured based on competitive tenders. – Maximize the use of existing facilities – The spacecraft bus and the payload are produced and tested with reduced requirements policy 18

Expected mission results Main mission objectives • Successful deployment of E-sail tethers • Successful observation/direction sensing of tethers • Detected spacecraft acceleration: > 4 E-6 m/s 2 • Validation of E-sail theory: Dependence of acceleration on voltage and solar wind conditions • Electron heating test: Dependence of acceleration on A/C modulation of electron beam, for different frequencies Secondary objectives • E. g. monitoring of the dust particle hit rate and size distribution (effective detector area 1. 7 m 2, i. e. largest ever flown) 19

Demo mission schedule • One of the main schedule drivers is the development of the tether production line • Estimated payload delivery time after the tether production capability exists is 1 – 1. 5 years • Launch could take place within 6 months from the payload delivery • Nominal mission duration including LEOP is 4 months • Satellite will be deorbited at the end of the mission 20

Summary • Electric sail concept requires a test mission to: – Demonstrate deployment and operations of the sail in space – Measure the acceleration of the spacecraft in different solar wind conditions – Test enhancement of the sail efficiency by electron heating • Demonstration mission can be performed with a reasonably small, simple and inexpensive spacecraft <=> mission design driver is the need to fly outside the magnetosphere • Life time of the demonstration mission is only 4 months • E-sail demonstration can be combined with other space physics observations • Mission can be performed in 2 years from development of the tether manufacturing capability 21