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NIAC Phase I Fellows Meeting Atlanta, Georgia March 7-8, 2006

Modular Spacecraft with Integrated Structural Electrodynamic Propulsion Nestor Voronka, Robert Hoyt, Brian Gilchrist, Keith Fuhrhop

TETHERS UNLIMITED, INC.

11807 N. Creek Pkwy S., Suite B-102 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 requirements • Propellant may be a source of contamination for optics and solar panels

Are there innovative alternatives?


NASA’s Vision of Exploration –

President’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

Isp

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) 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-andstructure 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 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| ≈ 0.3 Newtons

Proposed Integrated Structural Propulsion • Example: 100m conductor, 100 Ampere (!) in LEO – –

Thrust |iLxB| ≈ 0.3 Newtons Torque ≈ 750 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 selfassembling TinkertoyÂŽ 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% 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 Mission –

Small Missions for Advanced Research in Technology - Launched on 27 Sept 2003 • Arrived in lunar orbit 15 Nov 2004 • PPS-1350-G Hall Effect Ion Thruster (70 mNewton) –

Propellant mass fraction = 82.5 kg / 370 kg = 22.3 %

• 2nd time ion propulsion used for primary propulsion –

• •

1st 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


System Elements –

Nodes • Energy Storage • System Control

Booms • • •

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 ampacity 30% DOD for LEO 5 – 7 Year LEO life 5 – 10 whr/kg system SE

Li Expectations • • • • •

Flywheel Systems – Near Term

70 – 150 Cell whr/kg 20 – 60 A-hr ampacity 10 – 15% DOD for LEO 5 – 7 Year LEO life 10– 30 whr/kg system SE

• • • • • –

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 The Ultimate Spacecraft Battery

Auxiliary Bearings – touchdown and launch loads, stability, caging Housing – system and component integration, structural/dynamic response

Far Term Goals – Integrated Power & Attitude Systems • • • •

Magnetic Bearings – low losses, higher speeds, sensors, dynamic control

Motor/Generator – low losses, higher speeds, drive controls

– – – –

Composite Rotor – long life, safety without containment, light-weight hubs, design and cert. standards

75 whr/kg 92% efficiency 25 year LEO life -55-220°C

Energy Storage • 100 whr/kg • 30 year life

High System Specific Energy, Specific Power, Long Life High Round (Charge/Discharge) Trip Efficiency Multiple Functionality (Power and Torque) Long Storage Life Without Degradation

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 motorgenerator

Two counter-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


Electron Emitters –

Field Emissive Cathodes

• 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/cm2) 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


Electron Emitter Summary Device

Power Required Details

Thermionic Cathode+Gun

2.1 MW

18 emitters, Vf<1.25V for SCL

Field Emission Array

5.9 kW

10 emitters, Vf<0.4V for SCL

Hollow Cathode 1.25 – 10 kW

TO5 Header

A

1.8 cm

C

B

Consumable Required! (9-40 sccm Xe)


Electron Collection –

Passive 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 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/cm2 density, with 1mA-10mA total current 2015 – 10A/cm2 density, with >10A total current _

High Current Liquid Metal Ions (under development)

Low Current Gas Ions

Classical Field Ion Emission (a wetted needle)

+

+

Simple physics of field ionization and Taylor cones

Microfabricated Capillary Architecture Electric field and surface tension balance to form a “Taylor cone” at liquid surface

+

Accelerating Grid

+ Extracting Electrode

Liquid Metal Reservoir

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 direction • Competitive with tradition Electric Propulsion with added benefit 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 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


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