In-orbit demonstration of an iodine electric propulsion system

Abstract

Propulsion is a critical subsystem of many spacecraft^1-4. For efficient propellant usage, electric propulsion systems based on the electrostatic acceleration of ions formed during electron impact ionization of a gas are particularly attractive^5,6. At present, xenon is used almost exclusively as an ionizable propellant for space propulsion^2-5. However, xenon is rare, it must be stored under high pressure and commercial production is expensive^7-9. Here we demonstrate a propulsion system that uses iodine propellant and we present in-orbit results of this new technology. Diatomic iodine is stored as a solid and sublimated at low temperatures. A plasma is then produced with a radio-frequency inductive antenna, and we show that the ionization efficiency is enhanced compared with xenon. Both atomic and molecular iodine ions are accelerated by high-voltage grids to generate thrust, and a highly collimated beam can be produced with substantial iodine dissociation. The propulsion system has been successfully operated in space onboard a small satellite with manoeuvres confirmed using satellite tracking data. We anticipate that these results will accelerate the adoption of alternative propellants within the space industry and demonstrate the potential of iodine for a wide range of space missions. For example, iodine enables substantial system miniaturization and simplification, which provides small satellites and satellite constellations with new capabilities for deployment, collision avoidance, end-of-life disposal and space exploration^10-14.

Fig. 1. Schematic of the NPT30-I2 iodine electric propulsion system.

Solid iodine (darker green region) is located in a storage tank upstream of the plasma source tube (blue region). Heating causes sublimation and a low-pressure gas (lighter green region) enters the source tube (green arrow). A plasma (purple region) is created by a RF antenna, and iodine ions (I⁺, I₂⁺ and I²⁺) are accelerated by a set of grids. A cathode emits electrons (e⁻) to neutralize the ion beam. Waste heat is conducted towards the iodine tank and structural frame (solid blue arrows) or radiated away (blue dashed arrows).

Fig. 2. Beam composition and ionization efficiency.

a, Example mass-to-charge ratio, m/z, spectrum obtained with the TOF diagnostic system. The labels indicate the ions I²⁺, I⁺ and I₂⁺. b, Relative current concentration of iodine species in the ion beam as a function of RF generator output power. c, Ion-beam current extracted from the plasma source as a function of RF power with iodine and xenon propellants. The black curve shows results of a numerical plasma discharge model (Methods). d, Propellant mass utilization efficiency as a function of total power for different iodine mass flow rates. The error bars represent estimates of measuring equipment precision and accuracy limitations. Source data

Fig. 3. Propulsion system performance.

a, Ion flux distribution functions (IFDF) in the plume for acceleration voltages of 900 V and 1,300 V. b, Direct thrust measurements from a thrust balance compared with indirect thrust measurements estimated from the ion-beam current, applied grid voltage, and extrapolated beam divergence and beam composition data. c, Measured ion-beam divergence half-angle with iodine and xenon. The normalized perveance, p/p_max, is a measure of the ion space charge (Methods). d, Thrust and specific impulse performance map of the propulsion system within the operating total power range, and for different iodine mass flow rates. The error bars represent 1 s.d. (b) or estimates of measuring equipment precision and accuracy limitations (c). Source data

Fig. 4. In-orbit manoeuvres performed by an iodine electric propulsion system.

a, Mean semi-major axis of the Beihangkongshi-1 satellite from the SSN and GPS data, and as predicted using numerical simulations and theory. The arrows indicate separate firings. b, Mean semi-major axis as a function of time during manoeuvre 1B. The green region indicates when the propulsion system is firing. c, Thrust and total power telemetry during manoeuvre 1B. d, Comparison between ion-beam current, b, electron neutralizer current, e, and current to the accel grid, a, during ground, g, and in-flight, f, operation for manoeuvre 1B. The GPS data have an accuracy of approximately 20 m. Source data

Extended Data Fig. 1. Propulsion system architecture and integration with satellite.

a, The propulsion system is a complete system that includes all necessary subsystems for operation. Power is supplied from the spacecraft (S/C) and used for flow control, plasma generation, ion acceleration, and beam neutralization. Solid iodine sublimates and enters the inductively coupled plasma source. An igniter initially strikes a plasma which is then maintained by an RF antenna wrapped around the outside of the source tube. Ions from the plasma are extracted and accelerated by the high-voltage grids, and the positive ion beam is neutralized by electrons thermionically emitted from the cathode filament. b, The propulsion system installed in the Beihangkongshi-1 satellite before launch. Photograph reproduced and adapted by the authors with permission from Spacety. © 2020 Spacety Co., Ltd. (Changsha).

Extended Data Fig. 2. Electrical system architecture and thruster electrical schematic.

a, The main control unit interfaces with the satellite onboard computer and implements global control and safety algorithms. The RF generator supplies power to the RF antenna via a matching network to match the impedance of the plasma and generator for efficient power transfer. The cathode supply controls and monitors the electron-emitting cathode filament, the flow control unit manages the propellant tank and flow path heaters, the grid control unit manages the applied voltage to the acceleration grids, and the ignition unit controls the igniter needed for initial gas breakdown in the source tube to produce a plasma. b, General electrical circuit showing the high-voltage grids and electron-emitting cathode. Ions (denoted X^z+) from the upstream plasma source are extracted and accelerated by the voltage applied across the screen and accel grids. A small ion current, I_a, flows to the accel grid due to charge-exchange collisions with any unionized propellant in the plume. To maintain charge balance in the source tube, an electron current equal to the sum of the ion beam and accel grid currents flows to the screen grid, I_s. A current equal to the ion beam current is then emitted from the filament. The accel grid is biased negatively with respect to the filament to prevent electron backstreaming into the plasma source.

Extended Data Fig. 3. Ion optics simulations and propulsion system performance.

a, Perveance (space-charge) and cross-over limits of the propulsion system grid set obtained from particle-in-cell (PIC) simulations. The black dash-dot line shows the perveance limit from the Child-Langmuir law (see Methods), while the green shaded region denotes the operating range of the propulsion system. b, Correction factor as a function of RF power applied to the indirect thrust measurements to account for ion beam divergence and the presence of multiple ion species. c, Thrust of the propulsion system as a function of total system power and iodine input mass flow rate. d, PIC simulation of a single set of grid apertures (black shaded regions) showing the steady-state spatial ion distribution. The simulation is 2D in cylindrical coordinates and the domain has been normalized by the axial and radial simulation dimensions. e, Specific impulse of the propulsion system as a function of total system power and iodine mass flow rate. Error bars represent estimates of the measuring equipment precision and accuracy limitations. Source data

Extended Data Fig. 4. Ground-flight data comparison and propulsion system operation timeline.

a, Comparison between the ion beam current, I_b, electron neutralizer current, I_e, and current to the accel grid, I_a, during ground and in-flight operation for manoeuvre 1B. b, Measured electronic subsystem temperatures during propulsion system operation on the ground, and in space for manoeuvre 1B. The figure presents data for the main control unit, or Motherboard (MB), the Radio-Frequency Generator (RFG), the Grid Supply Unit (GSU), the Cathode Supply Unit (CSU), and the Flow Control Unit (FCU). c, In-orbit telemetry data of the thrust and power as a function of time for manoeuvre 1B indicating the propellant heating, propulsion system operation and propellant cooling stages. Source data