Abstract:
A plasma micro-thruster, including: an elongate and substantially non-conductive tube having a first end to receive a supply of propellant gas, and an open second end to act as an exhaust; first, second, and third electrodes extending circumferentially around the tube and being mutually spaced along a longitudinal axis of the tube, the third electrode being longitudinally interposed between the first and second electrodes; wherein the tube and the first, second and third electrodes are configured to generate a plasma from propellant gas flowing though the tube from the first end of the tube when the third electrode receives radio frequency power and the first and second electrodes are electrically grounded relative to the third electrode, such that the expansion of the plasma from the open end of the tube generates a corresponding thrust.

Description:
TECHNICAL FIELD 
       [0001]    The present invention relates to micro-thrusters for use in space applications, where thrust (force) is achieved through the generation of a plasma plume. 
       BACKGROUND 
       [0002]    Micro-thrusters find use in space applications where thrusts of the order of milli Newton are used to manoeuvre spacecraft. Such manoeuvring may be, for example, to direct a spacecraft into a desired orbit, to maintain the spacecraft&#39;s position within a desired orbit, or to remove the spacecraft from one orbit to another (e.g., parking in a so-called ‘graveyard’ orbit, or atmospheric re-entry). One matter of concern in the design of thrusters for spacecraft is to minimise weight. 
         [0003]    It is desired to provide a plasma micro-thruster that alleviates one or more difficulties of the prior art, or that at least provides a useful alternative. 
       SUMMARY 
       [0004]    In accordance with the present invention, there is provided a plasma micro-thruster, including:
       an elongate and substantially non-conductive tube having a first end to receive a supply of propellant gas, and an open second end to act as an exhaust;   first, second, and third electrodes extending circumferentially around the tube and being mutually spaced along a longitudinal axis of the tube, the third electrode being longitudinally interposed between the first and second electrodes;   wherein the tube and the first, second and third electrodes are configured to generate a plasma from propellant gas flowing though the tube from the first end of the tube when the third electrode receives radio frequency power and the first and second electrodes are electrically grounded relative to the third electrode, such that the expansion of the plasma from the open end of the tube generates a corresponding thrust.       
 
         [0008]    The present invention also provides a plasma micro-thruster, including:
       a tube having a length greater than its width, receiving at one end a supply of propellant gas, and having the other end open as an exhaust;   a first and a second conductive electrodes in a spaced-apart arrangement surrounding the tube, each electrodes being connected to zero relative potential; and   a third conductive electrode interposed between the first and second electrodes and surrounding the tube and adapted to be supplied with radio frequency power; and   wherein a plasma is ignited within the tube with the flow of propellant gas into said tube and the application of radio frequency power to said third electrode.       
 
         [0013]    The tube of the micro-thruster is preferably composed of a ceramic material. In a preferred form the micro-thruster includes a plenum chamber configured to supply a positive pressure of the propellant gas to the corresponding end of the tube. Advantageously, a gas flow rate controller is disposed between the plenum chamber and the corresponding end of the tube. The micro-thruster preferably includes a radio frequency power supply connected to the third electrode. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0014]    Some embodiments of the invention are hereinafter described, by way of example only, with reference to the accompanying drawings, wherein: 
           [0015]      FIG. 1  is a schematic side view of a micro-thruster in accordance with some embodiments of the present invention; 
           [0016]      FIG. 2  is a schematic side view of a micro-thruster in accordance with some embodiments of the present invention and in an experimental arrangement to measure parameters of the plasma generated by the micro-thruster, including a camera and a Langmuir probe; 
           [0017]      FIG. 3  is a graph of the measured intensity of the 488 nm Ar II line as a function of radial distance from the central axis of the plasma plume, for upstream Argon gas pressures of 0.54 Torr, 1.6 Torr. 2.3 Torr and 3.1 Torr, respectively, and 40 W RF power; 
           [0018]      FIGS. 4 and 5  are camera images of plasma plumes generated by the micro-thruster of  FIG. 2  for an Argon gas pressure of 1.6 Torr and RF powers of 40 W and 6 W, respectively; 
           [0019]      FIG. 6  is a graph of (i) normalized ion current measured by the Langmuir probe biased at −27 V and located at z=15 mm (solid circles), and (ii) normalized RF current I nos   2  (open squares), both as a function of RF power; the normalization being to the corresponding values for the maximum RF power of 30 W; and 
           [0020]      FIG. 7  is a graph of the ion saturation current as a function of position along the longitudinal axis of the micro-thruster, as measured by the Langmuir probe biased at −27 V for 9.5 W RF power (V rf =250 V) and a plenum pressure of 1.5 Torr. The solid vertical arrow  502  and the dotted vertical arrow  504  indicate the Langmuir probe&#39;s respective positions for the measurement of the full characteristic (to determine the electron temperature) and the measurements of  FIG. 6 . The solid horizontal line  506  indicates the position of the RF electrode. 
       
    
    
     DETAILED DESCRIPTION 
       [0021]    As shown in  FIG. 1 , a micro-thruster  10  includes an elongate tube  12  composed of a substantially rigid and substantially electrically non-conducting material. In the described embodiments, the tube  12  is composed of alumina, but it will be apparent that other materials with the described properties can be used in other embodiments, including other ceramic materials. The relative dimensions of the tube  10  are typically such that it is considerably longer than its outer diameter; for example. in some embodiments the aspect -ratio is about a factor of ten. Two mutually spaced and electrically conductive outer electrodes  14 ,  16  surround the tube  12 , and are maintained at a zero relative potential. In the described embodiments, the outer electrodes  14 ,  16  are in the form of generally cylindrical metal bands that extend circumferentially to around the tube  12  and whose height (i e., dimension along the longitudinal axis of the tube  12 ) is approximately equal to the outer diameter of the tube  12 , and the outer electrodes  14 ,  16  are mutually spaced along the longitudinal axis of the tube  12  by a distance of about 3 outer diameters (between the nearest edges of the electrodes  14 ,  16 ). A third or central electrode or metal band  18 , also surrounding the tube  12 , is situated centrally between the first and second bands  14 ,  16 , and in use is connected to a radio frequency source or generator  20 . The micro-thruster  10  can be encased in a non-conducting and vacuum-tight support structure (not shown). 
         [0022]    One end of the tube  12  is connected to a gas plenum chamber  22  that, in use, contains a propellant gas under positive pressure. The propellant gas is introduced into the tube  12  in a controlled manner by a suitable mechanism (e.g., a mass flow controller)  24 , that allows the flow rate of gas into the tube  12  to be controlled as desired. The resulting flow of gas  26  escaping from the open (exhaust) end of the tube  12  in itself generates thrust due to Newton&#39;s third law of motion. 
         [0023]    The application of radio frequency power with a frequency from below 100 kHz to above 1 GHz to the central electrode  18  causes an avalanche breakdown of the gas passing through the tube  12  to establish a plasma plume  28 . The plasma plume  28  projects outwards from the exhaust end of the tube  12  and increases the overall thrust over that generated by the gas stream  26  alone due to ion acceleration (possibly to supersonic velocities) caused by the plasma expansion. 
         [0024]    When used to control the movement of a spacecraft, the micro-thruster  10  is mounted to the spacecraft so that the open (exhaust) end of the tube  12  is directed away from the spacecraft into space, and, where a single micro-thruster  10  is used, in a direction opposite to the desired direction of the spacecraft&#39;s movement. In order to control the direction of thrust relative to the spacecraft, the micro-thruster  10  can be mounted to the spacecraft via an adjustable support or mount that allows the spatial orientation of the micro-thruster  10  relative to the spacecraft to he remotely and correspondingly adjusted and controlled, for example by mechanical means (e.g., using gimbals), and/or by electrical means (e.g., using magnetic or electric fields). Additionally or alternatively, a plurality of micro-thrusters  10  can be mounted orthogonally to allow for 3-axis control of the spacecraft. 
         [0025]    The micro-thrusters  10  described herein are compact and efficient in converting electrical energy to thrust, and therefore can be much lighter than prior art thrusters. As the described micro-thrusters  10  use non-metallic materials (e.g., ceramics) in contact with the plasma  28 , this avoids another of the difficulties suffered by prior art thrusters, namely metallic particles generated by sputtering endangering the spacecraft&#39;s solar panels. 
         [0026]    In one embodiment, the ceramic tube  12  has an outside diameter of 3 mm and an inside diameter of 1.5 mm, and a length of about 2 cm. The propellant gas used is argon, having a flow rate of about 10 to 1000 seem, more preferably about 100 sccm. The pressure in the plenum chamber  22  is about 7 Torr, and the pressure downstream of the tube  12  in the gas exhaust  26  is about 1 Torr. For about 10 watts generated by the radio frequency generator  20  at a frequency of 13.56 MHz, a plasma  28  was ignited, and observed to extend many centimeters downstream in a cone-shaped plume  28  with a half angle of less than 5 degrees. 
         [0027]    In a further embodiment, illustrated schematically in  FIG. 2 , a micro-thruster  10  has cylindrical ceramic tube  12  that is 2 cm long with inner and outer diameters of 4.2 mm and 5.3 mm, respectively. The central electrode  18  is in the form of a 6 mm high copper ring (A rf ˜1 cm 2 ) and the two outer electrodes  14 ,  16  are 3 mm high grounded copper rings  14 ,  16  placed upstream and downstream of the central electrode  18  and separated from it (edge-to-edge) by 3 mm. A vertical z axis with z=0 cm defined as the location of the upstream (gas inlet) end of the tube  12 , so that z=20 mm corresponds to the open (exhaust) end of the tube  12  and hence the start of the geometric expansion of the plasma plume  28 . 
         [0028]    The lower open (exhaust) end of the tube  12  projects into a 72 cm long, relatively large (5 cm) diameter glass tube  202  contiguously attached to a 30 cm long, 16 cm diameter aluminum vacuum chamber (not shown) equipped with a primary pump and a Baratron gauge. Argon gas is introduced upstream of the micro-discharge into a small cavity or plenum chamber  22  (1.2 cm wide and 4 cm in diameter) equipped with a Convectron gauge. The system was pumped down to a base pressure of ˜3×10 −3  Torr, and gas flows ranging from a few tens to hundreds of sccm resulted in an operating pressure range of 0.3-7 Torr as measured in the plenum chamber  22  and about 2.2 times lower as measured in the aluminium vacuum chamber. 
         [0029]    RF power from about 5 to about 40 W was coupled to the plasma using a π impedance matching network  204  equipped with a Rogowski coil to measure the RF current and a×1/1000 HV Tektronics probe to measure the RF voltage. A Bird power meter was inserted between the RF generator  20  and the impedance matching box  204  to measure both the forward and reflected power and deduce the RF power P rf  dissipated in the discharge. At any time, either a digital camera (Casio Exilim EX-F1) or an axially movable Langmuir probe (LP) with a 1 mm in diameter nickel tip was mounted on a back port/window  206  of the plenum chamber  22  to measure either the radial profile or the axial (longitudinal) profile of the plasma density. Although an RF filter was used in the LP data acquisition system, the small plasma cavity size did not allow for the LP to be fully RF compensated. Previous experiments with and without RF compensation in a larger scale device operating at lower gas pressure (a few mTorr) have shown that the error bar for T e  is of the order of ±0.5 eV for the electron bulk. 
         [0030]    The resulting capacitive radiofrequency (13.56 MHz) micro-discharge was about 2 cm long and 4.2 mm in diameter. Images of the discharge cross section were taken using a 488 nm filter of 10 nm bandwidth inserted between the plenum viewing port  206  and the digital camera lens. Although the focus was manually set about halfway into the cylindrical discharge, the measurement was integrated over the whole discharge volume. The results of the Ar II line intensity across the horizontal diameter as a function of radial distance are shown in  FIG. 3  for an RF power of 40 W and four upstream pressures of 0.54 Torr, 1.6 Torr, 2.3 Torr and 3.1 Torr, respectively. The 487.986 nm Ar II line corresponds to the 4p 2 D o -4s 2 P transition and the light intensity is n e   2  in the coronal model, assuming a two-step ionization where n e  is the electron density. Above 3 Torr. the discharge exhibits an annulus of maximum intensity located about mid-radius. and expands as a collimated beam over a few cm with striations. presumably resulting from shock waves from the gas flow appearing above 5 Torr. The mode of interest is the low pressure mode (less than ˜3 Torr) where the density peaks on the central axis with a broader plasma plume extending over about 1 cm. 
         [0031]    Images of the discharge cross section and of the discharge expansion were taken (without the Ar II filter) and are shown in  FIGS. 4 and 5  for a pressure of 1.6 Torr and RF powers of 40 W and 6 W, respectively. Although the radial sheath edge position cannot be spatially resolved, the density ratio between centre (r=0 mm) and edge (r=2 mm) in the coronal model is estimated to be about 4 at 1.5 Torr ( FIG. 3 ). Measurements of the peak breakdown voltage V break  using the HV probe provide a Paschen curve with a minimum of V break =230 V around 1.5 Torr. Once ignited, the plasma can be sustained for peak electrode voltages lower than V break  and RF powers of a few watts only. 
         [0032]      FIG. 6  shows both the ion saturation current I sat  measured with the LP biased at −27 V and positioned at z=15 mm, and I rf   2  (where I rf   2  is the mean square value of the current measured with the Rogowski probe) versus increasing RF power from 5 to 30 W. The linear variation of I rf   2  with RF power demonstrates that the impedance of the discharge is constant. The linear variation of I sat  with RF power suggests acceleration of secondary electrons across the RF sheath as the dominant electron heating process rather than RF sheath heating. A LP characteristic taken from −100 V to 80 V was measured at 19.7 W (for a peak RF voltage V rf =380 V), 1.5 Torr with the probe located at z=4 mm (near the upstream edge of the discharge), giving a plasma potential of 15 V and a bulk electron temperature of 3±0.5 eV. The density estimated using this electron temperature of 3 eV and Sheridan&#39;s sheath expansion model for a probe bias of −80 V is about 2.8×10 11  cm −3  at z=4 mm. Using a particle balance for a cylindrical argon discharge of length 20 mm and radius 2.1 mm and a single Maxwellian distribution for electrons yields a calculated electron temperature of about 2 eV for a gas temperature of 300 K. 
         [0033]    The I sat  axial profile obtained with the probe biased at about &#39;27 V is shown in  FIG. 7  for 9.5 W RF power (V rf =250 V) and a plenum chamber pressure of 1.5 Torr. When the probe was inserted into the discharge by more than 8 mm. the upstream pressure gradually increased by 0.1 Torr every 2 mm to reach 2.3 Torr at z=20 mm as a result of flow constriction. From  FIG. 3 , this would give a value underestimated by at least 25%. The flow constriction could also be the source of the density dip around z=5 mm, where the uncertainty on I sat  could be as high as 50%.  FIG. 7  shows that towards the upstream side of the tube  12  (z=6-10 mm), the ion current (and hence the plasma density) increases exponentially by an order of magnitude to peak at z=10 mm which corresponds to the centre of the RF electrode (z˜9 mm)  20 . From this maximum value, the ion current decays exponentially towards the exhaust opening of the tube  12 . This asymmetry in the axial profile is likely a result of the gas flow and geometric expansion. Since the ion current has been measured to increase linearly with power ( FIG. 6 ), scaling factors for RF power and axial position can be applied to the full characteristic taken at z=4 mm for 19.7 W to deduce a peak plasma density of 1.8×10 12  cm −3  at z=10 mm (the ‘centre’ of the discharge) for a power of 9.5 W. 
         [0034]    These measurements allow the development of a global model of the discharge where the plasma parameters can be derived from a power balance assuming a single Maxwellian for the electrons (T e =3 eV): 
         [0000]    
       
         
           
             
               
                 
                   
                       
                   
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         [0000]    where P rf  is the RF power, q is the electron charge. A plasma ˜2.9 cm 2  is the plasma wall loss area (ceramic surface area and two ends), n sh  is the plasma density at the radial sheath edge. 
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         [0000]    is the Bohm velocity (M is the ion mass), E c (T e ) is the collisional energy loss per electron-ion pair in argon. 
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         [0000]    corresponds to the voltage divider formed by the ceramic and the plasma sheath in between the RF electrode and the plasma bulk (the capacitance of the ceramic of thickness d=0.6 mm and dielectric constant ˜10×ε 0  is 
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         [0000]    and V rf  is the peak voltage applied on the RF electrode. The coefficient of 0.83 in equation (1) results from the asymmetry of the discharge (A plasma ˜3×A rf ). 
         [0035]    Since the sheath capacitance, hence β, is also a function of n sh , an iterative procedure is applied to determine both β and n sh . The sheath capacitance is written as 
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         [0000]    where s is the collisionless sheath thickness (K i ˜0.82 for RF Child law). For P rf =9.5 W (V rf =250 V which is larger than V break ). β is 0.26 (most of the RF voltage is dropped across the ceramic and V sheath ˜65 V), C sheath =4.2 pF˜2.9×C ceramic , n sh  is 6.1×10 11  cm −3 , and n axis  would be about 4× larger at 2.4×10 12  cm −3  as deduced from the radial profile of  FIG. 3 . This value is probably overestimated since the plume loss area is not taken into account which minimizes A plasma  (equation (1)). Since this value is of the same order as the measured density of 1.8×10 12  cm −3  for 9.5 W at z=10 mm, important parameters can be derived from the model. The mean free path for ion-neutral collisions (elastic and charge exchange) at 1.5 Torr is 45 μm. The sheath thickness from equation (2) is about 160 μm, giving an average number of 3.5 ion-neutral collisions in the sheath (the Debye length is 16 μm). No self-bias was measured on the blocking capacitor in the impedance matching box  204  due to the presence of the ceramic. The plasma potential in the region of the RF electrode  18  will be of the order of 22 V on axis (the value of 15 V measured at z=4 mm and an extra 
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         [0000]    and about 20 V at the radial sheath edge 
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         [0000]    which indicates that the inner wall of the ceramic tube  12  will develop a negative bias of ˜−36 V, since 0.83 βV rf ˜56 V at 9.5 W. 
         [0036]    At 1.5 Torr, the gas flow of about 100 seem corresponds to 3 mg s −1  or to 4.5×10 19  argon atoms per second. If this were being expelled from a nozzle at the sound speed (Mach 1) of v g =300 m s −1 , the corresponding thrust would be 
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         [0000]    If 10 W (10 J of kinetic energy) are effectively transferred into heating the gas, then 
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         [0000]    (M 1  is the total mass ejected per second). However, considering all degrees of freedom, i.e. 3×(½) then 
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         [0000]    along the z-axis which would correspond to a gas temperature of 
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         [0000]    (k is the Boltzmann constant). This value can be increased by increasing the RF power and the gas flow can be reduced by reducing the discharge diameter or introducing pressure gradients by modifying the cavity geometry (e.g. with a nozzle). Using the particle balance discussed above but for a gas temperature of 1430 K yields a calculated electron temperature of 2.5 eV compared with 2 eV obtained with 300 K (the gas temperature which would yield the measured electron temperature of 3 eV is 3200 K). 
         [0037]    Many modifications will be apparent to those skilled in the art without departing from the scope of the present invention.