Nestor Voronka - NASA's Institute for Advanced Concepts

NIAC Phase I Fellows Meeting
Atlanta, Georgia
March 7-8, 7
2006
Modular Spacecraft with
Integrated Structural
Electrodynamic Propulsion
Nestor Voronka, Robert Hoyt,
Brian Gilchrist, Keith Fuhrhop
TETHERS
UNLIMITED,
INC. NC.
11807 N. Creek Pkwy S., Suite B-102B
Bothell, WA 98011
(425) 486-0100 Fax: (425) 482-9670
voronka@tethers.com

Motivation
– Traditional propulsion uses propellant as reaction
mass
– Advantages (of reaction mass propulsion)
• Can move spacecraft center of mass, readily and relatively
quickly
• Multiple thrusters offer independent and complete control of
spacecraft (6DOF)
– Disadvantages
• Propellant is a finite and mission limiting resource
• Propellant mass requirements increases exponentially with
mission ∆V V requirements
• Propellant may be a source of contamination for optics and
solar panels
– Are there innovative alternatives

NASA’s s Vision of Exploration
– President’s s Vision Mandates NASA to “implement a
sustainable and affordable human and robotic program to
explore the solar system and beyond”
– Current architectures require very large total masses
to be launched from Earth
– Propellant mass fractions for In-situ resource
utilization (ISRU) and mining based architectures are
significant and costly
There exists a critical need for highly efficient low-cost
propulsion to assure access to space & in-space propulsion

Space Propulsion Landscape
10,000 sec
2,000 sec
I sp
Courtesy Gallimore, A., UMich

Electrodynamic Space Tether Propulsion
– In-space propulsion system
– PROS:
• Converts electrical energy into
thrust/orbital energy
• Little or no consumables (propellant) are
required
– CONS:
• Long (1-100km) 100km) flexible structures
exhibit complex dynamics, especially in
higher current/thrust cases
• Gravity gradient tethers have constrained
thrust vector
• Relies on ambient plasma to close
current loop

Proposed Solution
– Multifunctional propulsion-and
and-
structure system that utilizes
Lorentz forces generated by
current carrying booms to
generate thrust with little or
no propellant expenditure
• Utilizes same principles as
electrodynamic tether propulsion
– Utilize relatively short (≈100(
meter), rigid booms with
integrated conductors capable
of carrying large currents, that
have plasma contactors at the
ends

Performance of Proposed Approach
– Current flowing in a moving wire through space
interacts with the ambient magnetic field
• Earth’s s Magnetic Field in LEO ≈ 30,000 nT
• Interplanetary Magnetic Field ≈ 5 nT
– Lorentz Force: F = iL x B
– Space Tether Electrodynamic Propulsion
• Example: 10km conductor, 1Ampere in LEO
– Thrust |iLxB|
iLxB| ≈ 0.3 Newtons
– Proposed Integrated Structural Propulsion
• Example: 100m conductor, 100 Ampere (!) in LEO
– Thrust |iLxB|
iLxB| ≈ 0.3 Newtons
– Torque ≈ 750 N· N m

‘Structural’ ED Propulsion
– By connecting six booms to a spacecraft along
orthogonal axes, full 6DOF of motion can be
controlled (translational and rotational)

Modular Spacecraft
– By making booms and spacecraft modules
modular and interconnectable, , we create self-
assembling Tinkertoy ® like components for
space structures and systems

Optimal Path Planning
• Chemical Systems near-impulsive
– Hohmann and Bi-elliptical transfers
• Low-thrust trajectory planning (e.g.
electric propulsion)
– Near continuous low level thrust
– Additional constraints for optimization
problem
• Available Power (eclipse periods)
• Tethers and Structural
Electrodynamic Propulsion
– Additional constraints due to ambient
magnetic field
• Thrust Vector direction limited
• Thrust dependent on magnetic field
strength!

Low-Thrust Trajectory Optimization
– EP Orbit Raising from GTO to GEO
• Optimizing both thrust magnitude & angle
• Variable thrust can increase payload mass
fraction up to 3%, and be 5-10% 5
more fuel
efficient
– Secondary Effects to consider
• J2 effects, solar eclipsing, solar cell
degradation due to radiation
Kimbrel, M.S., “Optimization of EP Orbit Raising”, MIT, 2002.

ESA’s SMART-1 1 Mission
– Small Missions for Advanced Research in
Technology - Launched on 27 Sept 2003
• Arrived in lunar orbit 15 Nov 2004
• PPS-1350
1350-G G Hall Effect Ion Thruster (70 mNewton)
– Propellant mass fraction = 82.5 kg / 370 kg = 22.3 %
• 2 nd time ion propulsion used for primary propulsion
– 1 st
st was NASA Deep Space 1 launched Oct 1998
• Utilized near-constant thrust
• Trajectory optimization
– Propellant consumption
– Radiation Belt Transit Time
– Available power (limited thrust duration during eclipse)
• Thruster 1190W max out of available 1850W BOL

– Nodes
• Energy Storage
• System Control
– Booms
System Elements
• Structural Propulsion Booms
• Plasma Contactors
• Docking Mechanisms and
Sensors
– Key Elements
• Energy Source (Solar)
• Energy Storage
• Electron and Ion Sources

Energy Storage Technologies
Battery Systems
– NiH2
• 35 – 55 cell whr/kg
• 20 – 300 A-hr A
ampacity
• 30% DOD for LEO
• 5 – 7 Year LEO life
• 5 – 10 whr/kg system SE
– Li Expectations
• 70 – 150 Cell whr/kg
• 20 – 60 A-hr A
ampacity
• 10 – 15% DOD for LEO
• 5 – 7 Year LEO life
• 10– 30 whr/kg system SE
Flywheel Systems
– Near Term
• 25 – 40 whr/kg
• >4 kW hrs capacity
• 90% DOD for LEO
• 15 Year LEO life
• 10 – 20 whr/kg system SE
– Far Term
• 50 – 75 whr/kg
• Unlimited thru paralleling
• 90% DOD for LEO
• > 15 Year LEO life
• 40 – 75 whr/kg system SE
Courtesy NASA GRC P&PO

Flywheel Technology Challenges and Goals
Auxiliary Bearings –
touchdown and launch loads,
stability, caging
Magnetic Bearings – low
losses, higher speeds,
sensors, dynamic control
The Ultimate Spacecraft Battery
Motor/Generator – low losses,
higher speeds, drive controls
Housing – system and
component integration,
structural/dynamic
response
Composite Rotor – long life,
safety without containment,
light-weight hubs, design and
cert. standards
– High System Specific Energy, Specific Power, Long Life
– High Round (Charge/Discharge) Trip Efficiency
– Multiple Functionality (Power and Torque)
– Long Storage Life Without Degradation
Far Term Goals
– Integrated
Power &
Attitude Systems
• 75 whr/kg
• 92% efficiency
• 25 year LEO life
• -55-220°C
– Energy Storage
• 100 whr/kg
• 30 year life
– Pulse Power
• 2,000 W/kg
Courtesy NASA GRC P&PO

Flywheel Benefits
– Life is virtually independent of Depth of Discharge
– Performs equally well with low- and high-power loads
– State of charge easily determined by measuring flywheels
rotational velocity
– Demonstrated net (charge/discharge) efficiencies up to 93.7%
• Eddy-current and hysteresis losses in magnetic bearings and motor-
generator
– Two counter-rotating rotating flywheels produce no net torque (OR
can be used for attitude control)
!

Integrated Structural ED Boom
– Requirements
• Rigidity based on Application
• Conductive Element(s)
– Boom (Tether) Optimization
• Goal: Maximize Efficiency of Power
to Orbital Energy Conversion
– There is no optimal tether length, nor
optimal current level for a desired
thrust force
– Resistive Losses in boom (tether)
should be minimized

Integrated Structural ED Boom Construction
– Tensegrity (tensile integrity)
Structures
• “an assemblage of tension and compression components arranged in a
discontinuous compression system..” R.B. Fuller Patent, 1962.
– Tubular Booms (e.g. Stem)
– Rigidized Inflatables
• Foam Rigidized
• Mechanically Rigidized
• UV Cured Thermoset Composites
• Thermally Cured Thermoset Composites
• Work Hardened Aluminum Laminates
– On-orbit Construction
strength
and
conductive
elements
UV
dissolving
film

– Field Emissive Cathodes
Electron Emitters
• Microfabricated Emitter tips rely on sharp emitter
tips, and close non-intercepting electrodes to
generate high field required to enable electrons to
quantum tunnel out of the material into space
• High current densities (5000A/cm 2 ) have been
demonstrated
• Development undergoing to increase total current
output and reduce environmental constraints
– Hollow Cathodes
• Electric discharge ionizes neutral gas
• Technology well developed – neutralizers for EP
• 100A HCs have been tested (9-40sccm
Xe flow)
– Annual fuel requirement for 100A @ 20 sccm
• Xenon – 61.6 kg
• Hydrogen – 0.47 kg
• High current -> > High temperature -> > lifetime limit

Device
Thermionic
Cathode+Gun
Field Emission
Array
Electron Emitter Summary
Power Required
Details
2.1 MW 18 emitters, V f

– Passive Electron Collection
Electron Collection
• Space Tethers typically utilize large
collection areas
– Solid or grid spheres, bare tethers
• To collect 100A, 46.6kV needed
(4.7 MW) for a 1 meter sphere (!)
– Hollow Cathode
• 6.2 kW @ 280 sccm to collect
100A of electrons
– 6.6 kg of Hydrogen for 1 year

Hollow Cathode Ion Source
– Hollow cathode Ion Emission
• VERY inefficient as compared to electron emission
(ionization efficiency is 1:1)
• Ion emission requires ≈ 14 sccm /Ampere of emission
– Annual fuel requirement for 100A @ 1440 sccm
• Xenon – 4400 kg (!)
• Hydrogen – 33 kg
• 4.7kW @ 1440 sccm to emit 100A of ions
– OPTION: Combo plan – ion thruster
(without neutralizer) as contactor/thruster

Liquid Metal Ion Source
– Micro Ion Source Technology – Liquid Metal Ion
• Scalable system, including a passive material supply (no valves)
• Goal: Wide range of ion currents from addressable large area arrays
ays
• Goal: Optimized Power (> 80%) and Mass (≈100%)(
efficiencies
• Power efficiencies on the order of 300 Watts/Ampere expected
• Controllable current over 7 orders of magnitude
• Development Objectives:
– 2006 – 100 mA/cm 2 density, with 1mA-10mA 10mA total current
– 2015 – 10A/cm 2 density, with >10A total current
High Current Liquid Metal Ions
(under development)
Low Current Gas Ions
Classical
Field
Ion Emission
(a wetted
needle)
+ +
_
Microfabricated
Capillary
Architecture
Electric field and surface tension
balance to form a “Taylor cone” at
liquid surface
+ +
Liquid Metal
Reservoir
Accelerating Grid
Extracting Electrode
Simple physics of
field ionization and
Taylor cones
No energy loss, only ionization energy
Less contamination, can only produce ions
Increased reliability from lower voltage
operation, reduced arcing

Applications
– Self-Assembling Modular Spacecraft (SAMS)
– Self-Assembling Structure for Refueling Station
– Self-Assembling Space Tug
– Self-Assembling Structure for Large Mirror
or Antenna Arrays
– Formation Flying Space Systems
• Terrestrial Planet Finder (TPF)

Summary
– Proposed Concept IS feasible
• Almost propellantless – required consumable for ion source
• Almost full 6DOF control – no thrust in B-field B
direction
• Competitive with tradition Electric Propulsion with added benefit t of
structural elements
– Technology Challenges
• High Current Plasma Contactors
– Devices exist – robust units with higher efficiencies needed
• Plasma Contactor Space Charge Limiting
– High current densities may be environmentally limited
• Collision proof coordinated control laws for formation flight, and a
self-assembly
– Additional constraints imposed on low-thrust control laws
– Potential Applications
• Space Tug and Commodity Depot
• Structure for Beamed Power Solar Array/Antenna Fields
• Structure for Space Habitats with Integral Drag Makeup