Modular Spacecraft with Integrated Structural Electrodynamic Propulsion

Transcription

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

2 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?

3 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

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

5 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

6 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

7 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

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

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

10 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!

11 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.

12 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 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

13 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

14 Energy Storage Technologies Battery Systems NiH cell whr/kg A-hr A ampacity 30% DOD for LEO 5 7 Year LEO life 5 10 whr/kg system SE Li Expectations Cell whr/kg A-hr A ampacity 10 15% DOD for LEO 5 7 Year LEO life whr/kg system SE Flywheel Systems Near Term whr/kg >4 kw hrs capacity 90% DOD for LEO 15 Year LEO life whr/kg system SE Far Term whr/kg Unlimited thru paralleling 90% DOD for LEO > 15 Year LEO life whr/kg system SE Courtesy NASA GRC P&PO

15 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 C Energy Storage 100 whr/kg 30 year life Pulse Power 2,000 W/kg Courtesy NASA GRC P&PO

16 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)!

17 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

18 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, 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

19 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 20 sccm Xenon 61.6 kg Hydrogen 0.47 kg High current -> > High temperature -> > lifetime limit

20 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 kw TO5 Header A 1.8 cm C B Consumable Required! (9-40 sccm Xe)

21 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 sccm to collect 100A of electrons 6.6 kg of Hydrogen for 1 year

22 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 1440 sccm Xenon 4400 kg (!) Hydrogen 33 kg 1440 sccm to emit 100A of ions OPTION: Combo plan ion thruster (without neutralizer) as contactor/thruster

23 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: ma/cm 2 density, with 1mA-10mA 10mA total current A/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

24 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)

25 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