ESAIL proof of concept mission
Juha-Pekka Luntama
Pekka Janhunen
Petri Toivanen
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Outline
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7.
Introduction
Mission objectives
Magnetosphere
Mission elements
Expected mission results
Demo mission schedule
Summary
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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
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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
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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
 apogee of the test mission
orbit has to be well outside
the magnetosphere
 the shortest distance to
undisturbed solar wind is
towards the sun
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Elements of a proof of concept mission
• Pre-phase A analysis
– Payload
– Spacecraft bus
– Orbit
– Launcher
– Ground segment
– Lifetime
– Budget
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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 kV & 1kW)
• 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
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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
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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
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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
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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
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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
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Orbit candidates
Equatorial orbit
Low/medium
inclination orbit
Apogee radius:
47 Re
47 Re
Perigee height:
2800 km
2800 km
Inclination:
0
0 - 45
Orbit period:
7 days
7 days
Deceleration zone
Sun
Acceleration zone
Moon orbit
Bow shock
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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
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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
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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
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Mission “prime time” definition
Mission end
Prime time
Mission start
Launch and LEOP
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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
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Expected mission results
Main mission objectives
• Successful deployment of E-sail tethers
• Successful observation/direction sensing of tethers
• Detected spacecraft acceleration: > 4E-6 m/s2
• 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 m2, i.e. largest ever flown)
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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
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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
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