Demonstrating a new technology for space debris removal using a bi-directional plasma thruster

Abstract

Space debris removal from Earth orbit by using a satellite is an emergent technological challenge for sustainable human activities in space. In order to de-orbit debris it is necessary to impart a force to decelerate it, resulting in its atmospheric re-entry. A satellite using an energetic plasma beam directed at the debris will need to eject plasma in the opposite direction in a controlled manner in order to maintain a constant distance between it and the debris during the deorbiting mission. By employing a magnetic nozzle plasma thruster having two open source exits, bi-directional plasma ejection can be achieved using a single electric propulsion device. Both the forces exerted on the thruster and the target plate simulating the debris are simultaneously measured in a laboratory space simulation chamber showing that a force decelerating the debris and a zero net force on the thruster can be successfully obtained. These two forces can be individually controlled by external electrical parameters, resulting in the ability to switch the acceleration and deceleration modes of the satellite and the debris removal mode using a single electric propulsion device.

Figure 1

(a) Concept for space debris removal by bi-directional momentum ejection from a satellite. (b) Schematic of a magnetic nozzle rf plasma thruster having two open source exits. (a) When plasmas carrying momentum fluxes F₁ and F₂ are expelled from two axially opposite satellite exits, the respective forces shown by the horizontal arrows F₁ (pointing to the left and providing the acceleration of the satellite with respect to the orbit velocity) and F₂ (providing the deceleration) are generated and used to adjust the satellite velocity relative to the debris. Continuously imparting momentum flux F₁ to the debris (horizontal arrow F₁ pointing to the right) will cause its deceleration, final re-entry into the Earth atmosphere and natural burn up. (b) The open exits magnetic nozzle rf plasma thruster forming the single electric propulsion device where control of the momentum flux imparted onto the debris is obtained via the control of the plasma momentum fluxes ejected at each open exit using variable external parameters (solenoids currents and propellant gas flow rates).

Figure 2

(a) Schematic diagram of the experimental setup, together with the calculated magnetic field lines for the (I_BL, I_BR) = (8 A, 8 A) Left/Right solenoidal current case. (b) Calculated magnetic field profiles on axis for various combinations of (I_BL, I_BR). Both the thruster (attached to the pendulum thrust balance) and separated insulating target acting as space debris are immersed in a space simulation chamber. The displacements of the thrust balance and the target plate are simultaneously measured and calibrated into forces (a positive value corresponds to a displacement and force pointing to the right).

Figure 3

(a) Photographs taken by a digital camera via a vacuum viewport on the chamber side wall for the solenoid currents of (I_BL, I_BR) = (8 A, 0 A), (0 A, 8 A), and (8 A, 8 A). (b) The raw (gray thin lines) and filtered (red bold lines) displacement signals of the thruster attached to the thrust balance. (c) The displacement signals of the target plate. The positive and negative displacement in b and c corresponds to the rightward and leftward directions, respectively. (a) Shows that the plasma exhausted from the left- and right-hand open source exits are changed by the magnetic field configuration. The displacement signals in b show that the deceleration (rightward, F₂ − F₁ > 0) and acceleration (leftward, F₂ − F₁ < 0) forces are exerted to the thruster (satellite) for the (I_BL, I_BR) = (8 A, 0 A) and (0 A, 8 A) cases, respectively (Fig. 1a). The results for the (I_BL, I_BR) = (8 A, 8 A) case demonstrates that zero thrust force is exerted to the thruster (b) while imparting the force to the target (c).

Figure 4

Simultaneously measured force to (a,b) the target and (c,d) the thruster as functions of the right- and left-hand solenoid currents (I_BR and I_BL). Either of these two solenoid currents is maintained at 8 A when surveying the other one. The momentum flux ejection to the left- and right-hand sides of the thruster can be controlled by the magnetic field configuration, yielding the space debris removal mode (zero thrust and the finite force to the target), the acceleration mode (leftward force to the thruster), and the deceleration mode (rightward force to the thruster).

Figure 5

(a) Ion saturation currents measured by the Langmuir probes located at z = −35.9 cm (filled square) and +35.9 cm (open circles) as functions of the solenoid currents, where either of the two solenoid currents is maintained at 8 A. (b) Axial profiles of the ion saturation current for (I_BL, I_BR) = (8 A, 0 A) (open triangles), (8 A, 8 A) (filled circles), and (0 A, 8 A) (open squares), where the lines are added as visual guides. The plasma densities in the plumes exhausted to the left- and right-hand sides of the single propulsion device can be controlled by the magnetic field configuration, resulting from the change in the density profile inside the source tube. When the densities exhausted on both sides are balanced, the zero net thrust is maintained while imparting the force to the target. The solenoids are axially centered at z = ±8.1 cm (Fig. 2).

Figure 6

Simultaneously measured forces to (a) the target and (b) the thruster as functions of the gas flow rates from the left- and right-hand gas inlets (C_ArL, C_ArR), where the total gas flow rate is maintained at C_ArL + C_ArR = 100 sccm and the solenoid currents are set as (I_BL, I_BR) = (8 A, 8 A). The control of the momentum flux ejection to the left- and right-hand sides can also be obtained by changing the gas flow rates (using the inlets positioned at z = ±5 cm), yielding the space debris removal mode and the two acceleration/deceleration modes of the thruster/satellite.

Figure 7

(a) Ion saturation currents measured by the Langmuir probes located at z = −35.9 cm (filled square) and +35.9 cm (open circles) as functions of the gas flow rates, where the total gas flow rate is maintained at C_ArL + C_ArR = 100 sccm and the solenoid currents are set as (I_BL, I_BR) = (8 A, 8 A). (b) Axial profiles of the ion saturation current for (C_ArL, C_ArR) = (100 sccm, 0 sccm) (open triangles), (50 sccm, 50 sccm) (filled circles), and (0 sccm, 100 sccm) (open squares), where the lines are added as visual guides. The densities exhausted from the source can be controlled by the gas flow rates by inducing changes in density profile inside the source tube. The maximum density appears very close to the gas inlets due to the high neutral density there. For equal gas flow rates from the two gas inlets, the plasma densities at the left- and right-hand sides are balanced; the zero net thrust and the force to the target is simultaneously obtained as shown in Fig. 6.

Figure 8

(a) Schematic diagram of the calibration basket and threads attached to the thrust balance. (b) Displacement signal of the laser sensor when putting a known mass to the basket at about every 20 sec (a gray line), together with a filtered signal (solid red line). (c) The relation between the applied horizontal force and the measured displacement (open circles), together with a fitted line (thin blue line) giving the calibration coefficient.

Figure 9

(a) Temporal evolution of the small solenoid current (I_cal: solid black line), the force (solid red line) measured by the load cell, and the displacement signal from the LED sensor (dashed blue line). (b) The relation between the horizontal force imparted to the target structure and the displacement (red dots), together with a fitted line (solid blue line) giving the calibration coefficient.