Colloid Micro-Newton Thruster Development for the ST7-DRS and LISA Missions

Ziemer, John; Gamero-Castaño, Manuel; Hruby, Vlad; Spence, Doug; Demmons, Nathaniel; McCormick, Ryan D.; Roy, Tom

doi:10.2514/6.2005-4265

Cited by 27 publications

(17 citation statements)

References 10 publications

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“…The Darwin infrared space interferometer and the Laser Interferometer Space Antenna (LISA) [1] are examples of missions that require technology capable of producing 0.1 to 1000 N of thrust with low thrust noise and high thrust precision. Two ideal candidates that could meet the thrust requirements are the colloid micronewton thruster (CMNT) and the field emission electric propulsion (FEEP) thruster [1][2][3]. CMNT and FEEP thrusters are unique in that they do not require a cathode for propellant ionization.…”

mentioning

confidence: 99%

Regenerable Field Emission Cathode for Spacecraft Neutralization

Makela

Washeleski

King

2009

Journal of Propulsion and Power

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This research investigates the discharge characteristics of a field emission cathode for use in electric propulsion that has the ability to be regenerated when the emitter tip becomes damaged. Emitter tip regeneration is achieved by taking advantage of Taylor cone formation from an operating liquid-metal ion source. Tip formation is accomplished by solidifying, or quenching, the ion-emitting cone to preserve the sharp protrusion so that it can then be used for electron emission. Electron emission I-V curves were taken after tips were formed by quenching the liquid-metal ion source at ion discharge currents ranging from 1 to 25 A. Fowler-Nordheim modeling was then used to estimate the emitter tip radii of each quenched liquid-metal ion source. Results of the Fowler-Nordheim modeling were promising, showing the ability to regenerate tips and to control the features of the resulting tips by varying the ion current during the quench process. The set of experiments that are reported demonstrated the regeneration process of emitter tip radii ranging from approximately 30-45 nm from a tip quenched at 2 A down to tip radii of 15-22 nm when quenched at 25 A.

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confidence: 99%

Regenerable Field Emission Cathode for Spacecraft Neutralization

Makela

Washeleski

King

2009

Journal of Propulsion and Power

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“…However, any miniaturization only achieves its full potential if its merit is not outweighed by the mass and volume required by auxiliary propulsion components in the propulsion system. While advances have been made in the miniaturization of auxiliary components such as the Busek Microvalve used for the colloid thruster system on LISA Pathfinder [91], the propellant stored itself constitutes a natural barrier towards miniaturization. Any meaningful miniaturization of a thruster technology therefore needs to consider the entire propulsion mass, highlighting the importance of maintaining, or increasing, the propulsion system efficiency with advancing miniaturization.…”

Section: Systems Considerationmentioning

confidence: 99%

Space Propulsion Technology for Small Spacecraft

2018

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Abstract-As small satellites become more popular and capable, strategies to provide in-space propulsion increase in importance. Applications range from orbital changes and maintenance, attitude control and desaturation of reaction wheels to drag compensation and de-orbit at spacecraft end-of-life. Space propulsion can be enabled by chemical or electric means, each having different performance and scalability properties. The purpose of this review is to describe the working principles of space propulsion technologies proposed so far for small spacecraft. Given the size, mass, power and operational constraints of small satellites, not all types of propulsion can be used and very few have seen actual implementation in space. Emphasis is given in those strategies that have the potential of miniaturization to be used in all classes of vehicles, down to the popular 1-liter, 1 kg CubeSats and smaller.

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“…Charged droplets are generated by applying a strong electrostatic field and extracted to generate thrust. Colloid propulsion technology has advanced significantly during the last decade, through the development of the thrusters for NASA's ST-7 [23] and the Laser Interferometry Space Antenna missions [24]. The technology leveraged from the ST-7 design is now available for use in nanosatellites [25].…”

Section: Propulsion Subsystemmentioning

confidence: 99%

Feasibility for Orbital Life Extension of a CubeSat in the Lower Thermosphere

Blandino

Martinez-Baquero

Demetriou

et al. 2016

Journal of Spacecraft and Rockets

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Orbital flight of CubeSats at altitudes between 150 and 250 km has the potential to enable a new class of scientific, commercial, and defense-related missions. A study is presented to demonstrate the feasibility of extending the orbital lifetime of a CubeSat in a 210 km orbit. Propulsion consists of an electrospray thruster operating at a 2 W, 0.175 mN thrust, and an specific impulse (Isp) of 500 s. The mission consists of two phases. In phase 1, the CubeSat is deployed from a 414 km orbit and uses the thruster to deorbit to the target altitude of 210 km. In phase 2, the propulsion system is used to extend the mission lifetime until propellant is fully expended. A control algorithm based on maintaining a target orbital energy is presented that uses an extended Kalman filter to generate estimates of the orbital dynamic state, which are periodically updated by Global Positioning System measurements. For phase 1, the spacecraft requires 25.21 days to descend from 414 to 210 km, corresponding to a ΔV 96.25 m∕s and a propellant consumption of 77.8 g. Phase 2 lasts 57.83 days, corresponding to a ΔV 119.15 m∕s, during which the remaining 94.2 g of propellant are consumed. Nomenclature a = semimajor axis, km a d = acceleration due to atmospheric drag, km∕s 2 a ns = acceleration due to aspherical geopotential, km∕s 2 a T = control acceleration vector, km∕s 2 C = controller gain C d = drag coefficient E e = estimated energy per unit mass, km 2 ∕s 2 E SA = energy generated by solar arrays, J E t = target energy per unit mass, km 2 ∕s 2 F = Jacobian of nonlinear model G = state matrix of process noise H k = Jacobian of measurement model i = orbital inclination, deg K k = Kalman gain mt = spacecraft mass, kg m p;0 = initial propellant mass, kg _ mt = mass flow rate, mg∕s P EP = power available for electric propulsion system, W P k = covariance matrix of residual noise P S∕C = power available for spacecraft bus and payload, W Q = covariance of process noise R E = equatorial radius of Earth, km r a = actual position vector, km r e = estimated position vector, km r t = target position vector, km S = reference area, 0.01 m 2 T d = time spent in sunlight during one orbital period, s T e = time spent in eclipse during one orbital period, s T T = commanded thrust in the alongtrack direction, mN v a = actual velocity of the spacecraft, km∕s v e = estimated velocity of the spacecraft, km∕s v rel = velocity of the spacecraft relative to the atmosphere, km∕s v t = target velocity of the spacecraft, km∕s W = work done by atmospheric drag, J X d = power transfer efficiency: solar array to loads X e = power transfer efficiency: solar array to batteries x a = actual state, 6 × 1 x e = estimated state, 6 × 1 y k = Global Positioning System state measurement, 6 × 1 ΔE = energy per unit mass error, km 2 ∕s 2 Δt = integration time step, 0.1 s ΔV = change in velocity, m∕s δ = declination (latitude), rad δr = error between estimated and actual positions, m δ _ r = error between estimated and actual velocities, cm∕s ε = eccentricity ε GPS;k = Glob...

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Colloid Micro-Newton Thruster Development for the ST7-DRS and LISA Missions

Cited by 27 publications

References 10 publications

Regenerable Field Emission Cathode for Spacecraft Neutralization

Regenerable Field Emission Cathode for Spacecraft Neutralization

Space Propulsion Technology for Small Spacecraft

Feasibility for Orbital Life Extension of a CubeSat in the Lower Thermosphere

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