ESAIL proof of concept mission

1
ESAIL proof of concept mission

• Juha-Pekka Luntama
• Pekka Janhunen
• Petri Toivanen

2
Outline

1. Introduction
2. Mission objectives
3. Magnetosphere
4. Mission elements
5. Expected mission results
6. Demo mission schedule
7. 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
• gt 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
• gt focus in strictly on the main mission

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Earths 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 lt 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
• Ion and electron detector 1.5 kg
• PCU 0.5 kg
• Total 15 kg

2 km
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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 gt symmetrical spacecraft, fixed solar
  panel
• Very small payload gt 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
• gt 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
• gt spherical or octagonal cylinder with a
  diameter of ?1 m
• Payload constraints on the spacecraft body are
  modest
• gt 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 gt 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
• gt either active orbit control or short mission
  lifetime
• Final orbit not reachable without a perigee kick
  motor
• gt Spacecraft design more complex
• gt Up to 75 of launch mass fuel
• gt Longer and more complex LEOP phase due to
  orbit manoeuvres
• High initial orbit (e.g. GTO)
• gt less fuel needed
• gt higher launch costs
• Satellite visibility gt 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
Bow shock
Moon orbit
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Launcher options

• Final orbit requires the use of a perigee kick
  motor
• gt launch to either LEO or GTO
• Demo mission spacecraft
• dry mass ltlt 100 kg
• fuel from LEO to final orbit 75 of the launch
  mass
• gt launch mass 200 400 kg
• Piggy-back opportunities to be exploited
• gt 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
• gt 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
• gt a three month mission with the prime time
  during the second month
• Next suitable observation period in 9 months
• gt 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 gt 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
• ltgt 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