EMITTER FOR A THERMIONIC DISPENSER CATHODE, METHOD OF MANUFACTURING IT AND THERMIONIC DEVICE

INCLUDING IT.

The invention relates to a method of manufacturing an emitter for a thermionic dispenser cathode, including: forming a porous emitter body with substantially interconnected pores, having an emission surface from which, upon application of heat, electrons are emitted; the porous emitter body being suitable for transporting through the substantially interconnected pores, a compound released upon application of heat, to the emission surface, which compound, when deposited on the emission surface, serves to lower an effective work function of the emitter; wherein the porous emitter body is formed by means of a process of deposition of material; and wherein at least a region of the porous emitter body is provided with a continuously varying porosity.

The invention also relates to an emitter for a thermionic dispenser cathode, including: a porous emitter body, having an emission surface from which, upon application of heat, electrons are emitted; wherein the porous emitter body includes substantially interconnected pores, and the emitter is arranged to supply, upon application of heat, a compound to the emission surface, which compound, when deposited on the emission surface, serves to lower an effective work function of the emitter; wherein the porous emitter body includes at least a region exhibiting an increase in porosity in a first direction.

The invention also relates to a thermionic device including a thermionic dispenser cathode provided with an emitter.

The invention also relates to a thermo-electric propulsion device, including: a hollow cathode, having an inlet for feeding propellant gas into the hollow cathode and provided with an orifice plate at an end opposite to the inlet; the orifice plate having a nozzle formed therein with a length L through the orifice plate, the orifice having a diameter D; and an anode, facing the orifice plate.

The invention also relates to a method of operating a thermo-electric propulsion device, including: a hollow cathode, having an inlet for feeding propellant gas into the hollow cathode and provided with an orifice plate at an end opposite to the inlet; and an anode, facing the orifice plate, wherein a propellant gas is fed to the inlet and a voltage difference is established between the hollow cathode and the anode.

The invention also relates to a spacecraft including: a reservoir of propellant; at least one thermo-electric propulsion device, the thermo-electric propulsion device; including a hollow cathode, having an inlet for feeding propellant into the hollow cathode and provided with an orifice plate at an end opposite to the inlet; and an anode, facing the orifice plate, wherein a propellant gas is fed to the inlet and a voltage difference is established between the hollow cathode and the anode; and

a control system for controlling the operation of the at least one thermo-electric propulsion device.

The invention also relates to a propulsion system for a spacecraft, including: a first propulsion apparatus; the first propulsion apparatus being one of a cold gas and a resistojet propulsion apparatus; an apparatus for storing a propellant; and an apparatus for feeding the propellant to the first propulsion apparatus.

Respective examples of a method of manufacturing an emitter for a thermionic dispenser cathode, an emitter for a thermionic dispenser cathode and a thermionic device as defined in the opening paragraphs are known. US-A1-2001/0019239 discloses a method for manufacturing an impregnated cathode having a cathode pellet in which the pore portion of a sintered body of porous metal is impregnated with electron emitting material. The method comprises the steps of press moulding metal raw material powder to form a porous substrate and sintering the porous substrate to form a sintered body of porous metal. The average porosity of the porous substrate is controlled by adjusting the pressure of press moulding, and the average porosity of the sintered body of porous metal after sintering is controlled by adjusting the sintering temperature. Consequently, the average porosity of the entire pellet can be controlled by the general process. The porosity distribution is controlled by adjusting the descending speed of the punch and the pressing time. It is therefore not necessary to use raw material powder having different particle sizes and to mould in multi-layers.

A problem of the known method is that the porosity can increase in only one direction, parallel to the direction of movement of the punch. Thus, the porosity in a particular region within the pellet is not independent of that in another region. This leads inevitably to a less than optimal porosity distribution.

Respective examples of a thermo-electric propulsion device, method of operating a thermo-electric propulsion device and spacecraft as defined in the opening paragraphs are known. Gessini, P. et al. , "The T6 Hollow Cathode as a Microthruster" , 41 ^st AIAA/ ASME/SAE/ ASEE Joint Propulsion Conference & Exhibit, 10-13 July 2005, Tucson, Arizona, describes a target based measurement system that has been developed to evaluate the thrust produced by a T6 ion engine hollow cathode. The hollow cathode under investigation was developed by Qinetiq for the T6 Kaufman-type ion thruster. The orifice diameter of 0.5 mm used to take the measurements reported in the paper is smaller than those used in the majority of previous investigations of the behaviour of this type of device. The hollow cathode was operated in an open diode configuration, without using a keeper electrode as in the majority of previous investigations in the literature. The results of the research described in the paper indicate a contribution to the thrust other than simple thermal heating through equilibrium with the hollow cathode wall, as in the case of a resistojet. It is speculated that this thrust contribution could be attributed to the presence of energetic neutrals, which could be produced by charge-exchange collisions with ions, provided that these were of higher than thermal energies.

A problem of the known test set-up is that it operates at around 300 W discharge power, which can generally only be provided on spacecraft with a high- voltage power supply.

An example of a propulsion system as defined in the opening paragraphs is known. A basic cold gas system for use in small satellites comprises a main storage tank for storing an inert gas at a pressure upwards of 40 bar, a plenum, maintained at a pressure of approximately 1 bar and at least two control valves. In use, one of the control valves remains open and the other is opened and closed in quick succession to maintain the pressure in the plenum. Such a system is suitable for basic attitude control functions, such as maintenance of LTAN (local time of ascending node) and launch injection error correction in small satellites in an Earth Observation constellation.

A disadvantage of the known propulsion system is that it is not very flexible, because it lacks the capability to provide the impulse needed to carry out manoeuvres requiring larger changes in velocity. This is because the exit velocity and specific impulse are relatively low. A typical spacecraft fitted exclusively with such a propulsion system in low earth orbit ( < 1500 km altitude) is therefore unable to meet the requirement that it be capable of de-orbiting within 25 years, for example.

It is an object of the invention to provide a method, emitter and thermionic device of the above-mentioned types that facilitate the attainment of improved lifetime and control of emission characteristics of the emitter, in particular the current density.

This object is achieved by the method of manufacturing an emitter for a thermionic dispenser cathode, wherein the porosity is continuously varied by controlling at least one parameter of the deposition process.

The porosity corresponds to the ratio of non-solid volume to the total volume of a region of emitter material. The porosity affects the ability of

the emitter material to conduct fluids. Because the porous emitter body is formed by means of a process of deposition of material and the porosity is continuously varied by controlling at least one parameter of the deposition process. It is possible to provide each region of the porous emitter body with the appropriate porosity as the porous emitter body is built up. An added advantage is that the process can be performed relatively rapidly, since at least one parameter of the deposition process is varied continuously. There is no need for a staged process to form the porous emitter body. Moreover, the continuously varying porosity avoids disruptions in the flow of the compound released upon application of heat, because there are no discontinuities between regions with different pore sizes, as there would be in a layered device. It is noted that the notion of continuous variation is used herein in the mathematical sense to denote an absence of discontinuities in the porosity profile of the emitter.

In an embodiment, the deposition process comprises thermal spraying of the material for forming the porous emitter body.

Examples of such thermal spraying processes include flame spraying and plasma spraying. It is relatively easy to vary parameters of such processes in order to vary the porosity of the porous body that is being formed.

Thermal spraying is also suited to the metals and ceramics that are the preferred materials for forming the porous emitter body. Furthermore, thermal spraying allows for continuous control of the material composition of the emitter body in a relatively easy way, for example to include emission enhancing compounds into the micro structure.

In an embodiment, the porous emitter body is provided with at least a region having a porosity increasing in a direction away from the emission surface and into the porous emitter body.

The smaller pore size nearer to the emission surface provides resistance to the flow of the compound towards the surface. This further contributes to an increase in lifetime. A further effect is that the porosity is larger in the interior of the porous emitter body, serving to increase the reservoir available for storing an impregnate, which, in certain embodiments, decomposes upon application of heat to release the compound responsible for lowering the effective work function. The lifetime of the emitter is thereby increased.

In an embodiment, the porous emitter body is provided with at least one region bordering on a further external surface of the porous emitter body and having a porosity increasing in a direction away from the further external surface and into the emitter body.

An effect is to limit the rate of desorption at the further external surface. Again, a relatively large reservoir of impregnate can be provided in the core of the porous emitter body. A further increase in useable lifetime is thereby achievable.

In an embodiment, the porous emitter body is provided with a region bordering on the emission surface and having a porosity increasing in a direction from an interior of the porous emitter body towards the emission surface.

An effect is that the pore size immediately adjacent the emission surface can be relatively large. This makes the surface roughness relatively large, without the need for further surface treatment. Moreover, the formation of undesired compounds near the surface and the movement of material under the conditions to which the emission surface is subjected are less likely to lead to blocked pores.

In an embodiment, the porous emitter body is provided with at least a region having a porosity increasing in a direction essentially parallel to the emission surface.

An effect is that account can be taken of gradients in external temperature, pressure and gas concentrations at the emission surface to provide essentially uniform emission.

In an embodiment, the process of deposition of material includes varying the composition of the deposited material.

In this way, emission-enhancing compounds can be included in the material near the emission surface, whereas compounds modifying the thermal conductivity can be included near other external surfaces.

According to another aspect, in the emitter for a thermionic dispenser cathode according to the invention the porous emitter body includes at least one region exhibiting an increase in porosity in a direction differing from the first direction.

The differing direction may be parallel to the first direction but in an opposite sense, or it may be at an angle thereto. Because porosity increases in more than one direction are present, optimisation of the porosity of different regions to increase the lifetime and emission performance of the emitter is made possible.

In an embodiment, the porous emitter body is provided with at least a region having a porosity increasing in a direction away from the emission surface and into the emitter body.

An effect is that a relatively large reservoir for storing an impregnate is available in the interior of the emitter, which is not depleted rapidly in use.

In an embodiment, the porous emitter body is provided with at least one region bordering on a further external surface of the porous emitter body and having a porosity increasing in a direction away from the further external surface and into the emitter body.

An effect is to make the rate of desorption at the further external surface relatively low.

In an embodiment, the porous emitter body is provided with a further region, bordering on the emission surface and having a porosity increasing in a direction towards the emission surface.

An effect is to increase the pore size at the emission surface. This serves to increase the surface roughness, and thus the effective emission surface area, as well as to prevent blockage of pores due to contaminants formed under external influences.

In an embodiment, the porous emitter body is provided with at least a region having a porosity increasing in a direction essentially parallel to the emission surface.

An effect is to make the emission characteristics relatively uniform across the emission surface in situations where the external conditions are nonuniform, e.g. due to the flow of a medium past the emission surface.

In an embodiment, the porous emitter body has a material composition varying towards at least one surface

Compounds to increase or decrease the thermal conductivity can be present in larger proportions towards the at least one surface.

According to another aspect, the thermionic device according to the invention includes a thermionic dispenser cathode provided with an emitter according to the invention and/or obtainable by means of a method according to the invention.

In an embodiment, the emitter is provided as a component attached to the thermionic dispenser cathode, e.g. as an insert for a hollow cathode.

An effect is that the cathode can be made of a refractory material, selected for mechanical strength and resistance to erosion. The material of the porous body is selected in combination with the composition of the impregnate according to its suitability for emitting electrons.

It is also an object of the invention to provide a thermo-electric propulsion device, method of operating such a device and spacecraft of the types mentioned in the opening paragraphs with relatively high suitability for use in a spacecraft with a low-voltage power system of limited capacity.

This object is achieved by the thermo-electric propulsion device according to the invention, wherein the nozzle has an aspect ratio L/D having a value within the range from 0.05 to 1.8.

It has been found that values of the aspect ration within this range allow the power to be varied by varying the discharge current over a relatively wide range, without requiring significant changes in discharge voltage.

■ The device can therefore be dimensioned to operate with a low discharge voltage, with the power being varied by varying the discharge current.

Preferably, the orifice diameter has a value within the range of 0.1 mm to 0.4 mm.

Together with the aspect ratio value, the orifice diameter defines the length of the nozzle formed in the orifice plate. The range of nozzle lengths thus defined allows operation of the propulsion device with a discharge voltage in a range about 28 V, the voltage commonly provided by current satellite bus systems.

In an embodiment, at least a section of the hollow cathode is internally lined by an emitter, the emitter having a lower effective work function than the hollow cathode.

An effect is to enable the thermionic generation of a sufficiently large flux of electrons to ionise the propellant gas at a lower temperature. The propulsion device is therefore able to operate at lower cathode temperatures.

In an embodiment of the thermo-electric propulsion device, the hollow cathode is internally lined by an emitter according to the invention and/or obtainable by means of a method of manufacturing an emitter according to the invention.

An effect is to provide a thermo-electric propulsion device with a relatively long working life and able to provide sufficient thrust throughout this working life.

In an embodiment, a heater is provided along at least a section of the hollow cathode for stimulating thermionic emission from a material exposed to the interior of the hollow cathode.

An effect is to enable control of the degree of ionisation of the propellant by controlling thermionic electron emission.

In an embodiment, the orifice is provided with a chamfer along a perimeter, preferably on a side facing the anode.

As a consequence, an aspect ratio in the preferred range can be attained without necessarily having to use a thin orifice plate. The orifice plate can be dimensioned to withstand the pressure in the hollow cathode, as well as to conduct a sufficient amount of heat away from the orifice. These features prolong the useable lifetime of the propulsion device.

In an embodiment, at least a section of the orifice plate defining the nozzle is made of a carbon material, preferably graphite.

An effect is that the nozzle can be machined to acquire the appropriate shape and dimensions relatively easily. The nozzle also has good thermal characteristics. Moreover, in particular graphite has a low sputter yield, meaning that the shape of the nozzle is retained relatively well during operation of the thermo-electric propulsion device. A high work function implies that there is little or no contribution to the discharge current.

In an embodiment, the nozzle is provided in the shape of a de Laval nozzle.

Near the throat of the nozzle, the gas velocity locally becomes transonic (Mach number = 1.0) . As the nozzle cross sectional area increases the

gas continues to expand and the gas flow increases to supersonic velocities where a sound wave will not propagate backwards through the gas as viewed in the rest frame of the nozzle. Thus, the exit velocity of the propellant gas is increased, providing a higher specific impulse.

According to another aspect, the method of operating a thermo-electric propulsion device according to the invention includes operating the thermo-electric device in a regime wherein gas flowing through the hollow cathode is substantially fully ionised within the cathode, at least upon reaching the orifice.

Due to the high degree of ionisation, good energy-equipartition - the process of transfer of thermal energy from electrons to ions - is achieved within the cathode. Heating of the plasma is substantially provided by Ohmic dissipation of the discharge current in the nozzle, rather than by conductive transfer of heat from the cathode. Because the resistance is highest at this point where the diameter of the plasma is smallest, the greatest amount of heating occurs in this area. The voltage difference required for ionisation in the relatively small orifice is relatively low, allowing operation at the low discharge voltage values that are compatible with small spacecraft without power electronics. In effect, a diffuse discharge is achieved, which requires a much lower discharge voltage. Thus, although the method uses an anode and a cathode in a manner similar to an arcjet, the fact that a hollow cathode is used allows operation at much lower discharge voltages and input power. An added advantage of the method is that, where the anode has a sufficiently large surface area, it is relatively easy to achieve operation in so-called spot mode, which is a stable mode of operation conducive to long useable lifetime of the thruster. Moreover, the emission of charged particles by the propulsion device is at a relatively low level under these conditions.

It is noted that it is known to use a hollow cathode without an anode to produce a stream of charged particles for neutralisation, e.g. in a Kaufman-type ion engine.

An embodiment includes applying heat to an emitter according to the invention and/or obtainable by means of a method of manufacturing an emitter according to the invention.

An effect is that there are sufficiently many thermionically generated electrons available for acceleration into the propellant gas to ionise it throughout the lifetime of the emitter, and that this lifetime is relatively long.

In an embodiment, an inert propellant gas is fed into the hollow cathode.

An effect is to minimise the risk of contamination of sensitive spacecraft equipment.

An embodiment of the method includes heating at least a section of the hollow cathode externally, so as to stimulate thermionic emission from a material exposed to an interior of the hollow cathode.

An effect is to allow a sufficient degree of ionisation of the propellant gas to be achieved in a controlled manner. Thermionically generated electrons are accelerated into the propellant gas, knocking out electrons to ionise the propellant gas. The acceleration occurs within a sheath established between plasma formed from the propellant gas, and the cathode wall.

An embodiment of the method includes adjusting a thrust produced by the thermo-electric propulsion device by varying the discharge current between the hollow cathode and the anode.

This embodiment involves the use of a thermo-electric propulsion device that is suitably dimensioned to ensure that the propellant is heated predominantly in an area of the hollow cathode interior adjacent the orifice. The voltage drop across the orifice is only weakly dependent on parameters other than the dimensions of the orifice, so that the input power is determined predominantly by the discharge current. This mode of operation can therefore be used in a spacecraft with a power system imposing constraints on the discharge voltage.

In an embodiment of the method, a mass flow of propellant fed into the hollow cathode and a discharge current between the hollow cathode and the anode are adjusted to satisfy the constraint: m -Ir m > ¹ AK 4 - π %

wherein: m is the mass flow [kg-s ¹ ], m is the atomic mass of the propellant [kg]

I _D is the discharge current [A] e is the electron charge [C] d _AK is the distance from the anode to the orifice in the orifice plate [m] , and

A _A is the surface area of the anode face exposed to the orifice plate [m ² ] .

As a consequence, the thermal flux of electrons to the anode, is at least equal to the discharge current. This ensures that the thermo-electric propulsion device is operated in a so-called "spot mode" as opposed to the so-called "plume mode" . The spot mode is a relatively stable mode of operation, in which the anode passively collects the discharge current and no potential drop is generated between the cathode and the anode. This is

conducive to a relatively long useable lifetime of the propulsion device. Moreover, efflux of charged particles is substantially reduced in this way. The spacecraft risk of exposing the spacecraft to ion bombardment is substantially reduced.

According to another aspect the control system in the spacecraft according to the invention is configured to allow the spacecraft to execute a method according to the invention.

The control system adjusts the mass flow, discharge current and variables determining thermionic emission within the hollow cathode to achieve substantially full ionisation of the propellant gas within the cathode, at least upon reaching the orifice.

It is a further object of the invention to provide a propulsion system for a spacecraft and spacecraft of the types defined in the opening paragraphs with a relatively high degree of flexibility of use at a relatively low cost.

This object is achieved by the propulsion system according to the invention, which further includes a second propulsion apparatus, provided with a connection to the propellant feeding apparatus and including a thermorelectric propulsion device for thermalising propellant fed to it to produce thrust.

The propulsion system offers a choice of propulsion mode, adapted to a specific situation. It is thus more flexible. The higher exhaust velocities attainable by the thermo-electric propulsion device enable greater total impulse to be provided, when needed. On the other hand, the lower power requirements of the first propulsion apparatus make it suitable for use as the default means of propulsion. Because the second propulsion apparatus is provided with a connection to the propellant feeding apparatus and

includes a thermo-electric propulsion device for thermalising the same propellant as is used in the first propulsion apparatus, a low-cost solution is provided, since the two propulsion apparatus share, the same architecture for storage and feeding of the propellant.

In an embodiment, the connection of the second propulsion apparatus to the propellant feeding apparatus is provided by the first propulsion apparatus.

Thus, the first and second propulsion apparatus are effectively connected in series to form an integrated propulsion system that can be operated in a "cold" mode, with the thermo-electric propulsion device forming the exhaust of the first propulsion apparatus, or in a mode in which the thermo-electric propulsion device is operational. In the latter mode, a higher total impulse is provided. An effect of this configuration is that the control valves needed to provide the propellant at the correct pressure are not duplicated.

In an embodiment, the thermo-electric propulsion device includes an anode and a gas-fed hollow cathode, configured, in use, to ionise internally at least part of the gas fed to it.

An effect is that a relatively low voltage is needed to provide an electrical discharge between the anode and the cathode, compared to, for instance, an arcjet. In combination with the use of a first propulsion apparatus of the cold gas or resistojet type, a relatively flexible propulsion system is formed without requiring relatively expensive and heavy power electronics on board the spacecraft.

In a variant of this embodiment, the hollow cathode is provided with an emitter insert having a low effective work function relative to the hollow cathode.

An effect of this is that less power is required to ionise the propellant gas in the thermo-electric propulsion device, because the emitter insert provides a sufficiently high flux of electrons at a relatively low temperature. The material of the cathode can be chose to meet other requirements, such as those relating to mechanical strength and durability.

In an embodiment, the hollow cathode is provided with an orifice plate positioned downstream of its connection to the propellant feeding apparatus.

An effect of this is to provide a sufficient level of plasma heating at lower voltages and input power. The channel in the orifice plate that forms the orifice operates as a nozzle to constrain the plasma within the hollow cathode. It also allows for resistive heating to take place at the nozzle, for which a relatively low voltage is required.

In an embodiment, the second propulsion apparatus is terminated by a de Laval nozzle at an end downstream of its connection to the propellant feeding apparatus.

This allows the cathode to function as a super-sonic nozzle for further acceleration of both the cold propellant (when the thermo-electric propulsion device is not operational) and the hot gas (when the thermoelectric propulsion device is in use) .

In an embodiment, the second propulsion apparatus includes a thermoelectric propulsion device provided with an emitter obtainable by means of a method of manufacturing an emitter according to the invention or an emitter according to the invention.

The invention will be explained in further detail with reference to the accompanying drawings, in which:

Fig. 1 is a very schematic cross-sectional view of a thermo-electric propulsion device;

Fig. 2 is a detailed schematic cross-sectional view of a downstream end of a hollow cathode in the thermo-electric propulsion device;

Fig. 3 is a diagram illustrating operating parameters for a first mode of operation of the thermo-electric propulsion device;

Fig. 4 is a diagram illustrating discharge parameters applicable to the first mode of operation;

Fig. 5 is a diagram illustrating operating parameters for a second mode of operation of the thermo-electric propulsion device;

Fig. 6 is a diagram illustrating discharge parameters applicable to the second mode of operation;

Fig. 7 is a diagram illustrating operating points of the thermoelectric propulsion device in a first configuration;

Fig. 8 is a diagram illustrating operating points of the thermoelectric propulsion device in a second configuration;

Fig. 9 is a diagram illustrating emission parameters for various internal diameters of an emitter insert for use in the thermo-electric propulsion device;

Fig. 10 is a very schematic view of a parallelepiped-shaped emitter for use in a thermionic device;

Fig. 11 is a very schematic view of a cylindrical emitter insert for a gas-fed hollow cathode as included in the thermo-electric propulsion device of Fig. 1; and

Fig. 12 is a block diagram of a satellite propulsion system.

A thermo-electric propulsion device, referred to herein as a thruster 1, includes a hollow cathode 2 and an anode 3. The hollow cathode 2 is made of a refractory metal, such as tungsten, thoriated tungsten, tantalum, molybdenum, rhenium, osmium, niobium, iridium, ruthenium, scandium or mixtures thereof, or alternatively a form of carbon, such as graphite, or a ceramic material. It has an elongated shape, in this example the shape of a circular cylinder. The hollow cathode 2 is similar in configuration to a typical neutraliser cathode.

As will be explained in more detail with reference to Fig. 12, a propellant gas is fed to the hollow cathode 2 through an inlet (not shown in detail) at an upstream end 4. The propellant used is an inert gas such as Xenon, Argon, Krypton, Neon, Nitrogen or Helium. This description will use Xenon as an example.

At a downstream end, the hollow cathode 2 is terminated by an orifice plate 5. The orifice plate 5 could be positioned slightly upstream of the

extreme end of the hollow cathode 2, but it should not be too far away from the anode 3. This would cause a potential drop between the anode 3 and the hollow cathode 2 to maintain a plasma bridge between the two.

Small spacecraft power systems operate at 28 V with no high power electronics and up to 100 W of available power. Thanks to an appropriate choice of dimensions and operating parameters, the thruster 1 can operate straight off these systems with no power conditioning. By contrast, power electronics are required for advanced propulsion systems, which all operate at kV voltages. Chemical propulsion systems come with immense handling costs and solid rocket alternatives may only be able to operate once.

The orifice plate 5 has a nozzle 6 formed therein with a length L through the orifice plate 5 (Fig. 2) in axial (i.e. longitudinal) direction. The orifice is defined by the nozzle 6 at its narrowest point in longitudinal direction. The diameter D of the orifice and the length L of the nozzle determine the aspect ratio AR = L/D .

It is observed that the aspect ratio AR can be set substantially independent of the thickness d of the orifice plate 5 by providing the orifice with a chamfered edge 7. The chamfer is such that the nozzle length may be about half the thickness of the orifice plate 5. In an example, the thickness of the orifice plate is of the order of 1 mm. Machining of the orifice plate 5 to provide the chamfered edge 7 is made relatively easy by using graphite as the material for at least a section of the orifice plate 5 surrounding the nozzle 6. Graphite also has low thermal conductivity, and low ' sputter yield to resist any change of orifice geometry. This is significant because high current densities will prevail in the nozzle to maximise Ohmic heating. The maximum tolerable tip temperature for

adequate thruster life will dictate the limiting current density and maximum acceptable power deposition into the orifice.

The hollow cathode 2 is internally lined along at least part of its length by an emitter insert 8. The emitter insert has a lower effective work function than the hollow cathode 2. The effective work function is the minimum amount of energy needed for an electron to leave a free surface, and varies from metal to metal.

The means by which the emitter insert 8 is provided with a low effective work function will be described more fully in relation to Figs. 10 and 11. Thermionic emission is stimulated by raising the temperature of the emitter insert 8, to which end a heater in the shape of a coil 9 is provided along at least a section of the hollow cathode 2 to provide resistive heating. Other types of heater can be used in place of the coil 9, as long as they enable a temperature of approximately 1500 K to be reached.

To operate the thruster 1, a trigger voltage applied between the cathode 2 and the anode 3 to initiate a discharge. The orifice plate 5 increases the pressure in the interior of the hollow cathode 2, generating a plasma within the hollow cathode 2. Self -heating is then maintained by the acceleration of ions through a sheath between the plasma and the cathode and recombination of these ions at an internal surface of the cathode 2 (or emitter insert 8) . Confinement of the plasma column to the interior of the hollow cathode 2 permits operation at low cathode fall voltages - the voltage difference between the cathode and the plasma at the nozzle 6 - while allowing high currents to be carried. The plasma column induces sheath-enhanced emission at the emitter surface due to the intense electric field ( ~ 10 ⁷ V/m) between the plasma column and the cathode over a distance of the order of the Debye length.

Thermionically emitted electrons liberated from the emitter insert 8 are accelerated through the sheath potential into the plasma, acquiring an energy level sufficient to excite an inert propellant like Xenon to a meta- stable state. Lower-energy electrons then contribute to ionisation of the atoms of propellant gas. The discharge current is drawn through the orifice towards the anode 3. This discharge current is responsible for Ohmic heating of the ionised propellant, providing useful thrust.

To achieve useable thrust and still be able to meet the constraints imposed by small spacecraft power systems, the aspect ratio AR is kept to a value within a range of 0.05 to 1.8, preferably 0.2 to 1.8, more preferably 0.25 to 1. Good results in this case means that a relatively wide range of discharge currents can be conducted through the cathode to the anode without significantly affecting the discharge voltage. The discharge voltage, defined as the voltage difference between the cathode and the anode can be kept to a value that can comfortably be delivered from a standard power bus in a commercial satellite.

Given the aspect ratio AR and orifice diameter D, the thruster 1 is controlled so as to operate in a regime wherein the propellant gas entering the hollow cathode 2 at the upstream end 4 is substantially fully ionised within the cathode 2, at least upon reaching the orifice. The model used to determine the geometric parameters and accompanying operating parameters will now be sketched in general terms .

The design process may be an iterative process. Given a certain propellant mass flow, the discharge current, power output, required input power, required aspect ratio AR and required emitter insert temperature are calculated. Then, a check is made to ensure that the required input power satisfies the constraints imposed by the target spacecraft's power system. If the constraints are not met, the process is repeated.

Figs. 3-5 illustrate the energy balance in the nozzle 6, with the point of stable operation being where the graphs of the power dissipated in the orifice and the power loss or output intersect.

Fig. 3 plots the power loss and dissipated power for an equivalent mass flow / = 0.1 A (the equivalent mass flow is found as I = — , m where m is the actual mass flow, m the propellant gas particle mass and e the electron charge) , a discharge current of 0.5 A and a nozzle length L of 1 mm. Fig. 4 assumes a fixed aspect ratio AR of 1.3 and the same value for the equivalent mass flow.

Fig. 5 plots the power loss and dissipated power for an equivalent mass flow I _eq = 0.15 A, a discharge current of 0.5 A and an orifice length L = I mm.

The power supplied to the thruster 1 is essentially equal to the sum of the power needed to heat the emitter insert 8 and the power needed to conduct the discharge current through the plasma at the discharge voltage V _d . The discharge voltage V _d is equal to the sum of the cathode fall voltage V _p , the Ohmic drop across the nozzle 6 and voltage drops across double sheaths which form at the entrance and exit of the nozzle:

K- = K _{p +} AK _{41 +} A^ ₊ AK ₄₂ (D

In this equation, the voltage drop across the nozzle 6 dominates. The voltage drop across the nozzle 6 is given by the following equation:

^δ ^ =H^ . ⁽²⁾ where I _d is the discharge current, and σ is the conductivity of the plasma. The conductivity is dependent on the electron temperature and

electron density within the orifice. Lowering the aspect ratio AR at a given diameter will decrease the voltage drop across the nozzle, but have essentially no effect on the electron density. Thus, the voltage drop across the nozzle can be kept low by keeping the aspect ratio at a low value.

The power dissipated in the orifice is given by the following relation:

PlOSS = Il - R (3)

Given the quadratic dependence of dissipative power on the discharge current, it is preferred to increase discharge current while reducing both orifice length and diameter to maintain low operating voltages whilst dissipating the same power into the flow. A further effect of such a design is to provide more control over plasma parameters by manipulating discharge current to effect almost quadratically the power deposition with little change in discharge voltages. Thus, the thrust is adjusted by varying the discharge current between the cathode and anode.

Fig. 6 shows the operating points for an aspect ratio AR = 1.3 and a nozzle length L = I mm. Fig. 7 is a similar plot for an aspect ratio AR = 0.25 and nozzle length L = 0.25 mm. The plots show that, at low aspect ratios and small orifice diameter values, the discharge is much less voltage-sensitive. Greater changes in discharge power can thus be effected at a limited value of the discharge voltage.

The required rate of mass flow can be calculated from an energy balance over the plasma within the nozzle 6. This energy balance is expressed as follows :

where I, is the ion fl<ux, I ^ψ _e is thhe e>lectron flux, ε _t is the degree of ionisation, and q _r is the radiative energy flux. I _eq is the equivalent mass flow.

Under these conditions and within the preferred range of geometric parameters, a choked flow regime is established in the nozzle 6, which is preferably a de Laval nozzle. In that case, the propellant gas is accelerated out of the hollow cathode 2 at supersonic velocities. It is necessary to produce a sufficiently large flux of electrons by thermionic emission in order to balance the energy loss due to convected thermionically generated electrons and energy gain through ion bombardment at the emitter insert 9. This energy balance is expressed as follows :

JJ.7,(ff, +V _f ~φ _eff )dA , (5) where the integral is over the emitter surface, φ _eff is the effective work function, V ₁ is the cathode fall voltage, ε _s is the degree of ionisation and J, is the ion current density. The thermionic current density follows from the Richards on-Schottky equation:

in which A _R is Richardson's constant, k _B is Boltzmann's constant and φ _eff the effective work function. Using equations (5) and (6), the cathode temperature T ₀ can be determined. Fig. 8 shows the required temperature for the emitter insert 8 in relation to its internal diameter

assuming a discharge current I ₀ of 0.5 A and an emitter length of 20 mm. Fig. 9 is a similar plot for a discharge current I _D = 0.6 A and emitter length of 10 mm.

It is known that a device like the thruster 1 can operate in two modes, namely plume mode and spot mode. When operating in spot mode. The spot mode is more stable, whereas the plume mode is characterised by strong discharge oscillations, noise and temperature variations. For this reason, the mass flow of propellant fed into the hollow cathode 2 and discharge current I _D are adjusted to satisfy the constraint:

Tn - I _n d A,, 3 _/ ,_ _* m > ^- ^■ ^^ ^{■ ■} 4 - π ^/l , (7) e A _A where d _AK is the distance from the anode to the cathode's orifice and A _A is the surface area of the anode face exposed to the orifice plate 5. With operating parameters adjusted to satisfy equation (7), it is ensured that the thermal flux of electrons to the anode 3 is at least equal to the discharge current.

A method of manufacturing an emitter with a low effective work function has been developed independently of the thruster 1, which method will now be explained in more detail.

Turning to Fig. 10, an emitter 10 has an emission surface 11 from which, in use, electrons are emitted. In one embodiment, the emitter 10 is attached to a cathode member (not shown) . In another embodiment, the emitter 10 is an integral part of a cathode. Upon application of heat and establishment of an electrical field between the cathode and an anode (not shown) , thermionic emission takes place. Such a dispenser cathode is used in a wide range of thermionic devices including, but not limited to, applications in RF and microwave engineering, klystrons, thyratrons (including those used for radar and broadcast, linear accelerators,

magnetrons, high power switching, synchrotrons, photo-cathodes, laser pulsing, lithotripsy, and trigger transformers) , travelling wave tubes, microwave tubes, lithography, electron-beam welders, X-ray sources, gas lasers, metal-vapour lasers, excimer lasers and free electron lasers, electron microscopy, high-voltage modulation, arc-lamps, and cathode lighting systems, projection and wide-screen displays, hollow cathodes and ion propulsion technologies.

The emitter 10 comprises a porous emitter body impregnated with a composition for supplying, upon application of heat, a compound to the emission surface 11, which compound serves to lower the effective work function of the emitter 10. The work function is the minimum amount of energy needed for an electron to leave the emission surface 11, and varies from metal to metal. Preferably, the porous emitter body is made of a transition metal or alloy, for example tungsten, thoriated tungsten, tantalum, molybdenum, rhenium, osmium, niobium, iridium, ruthenium, scandium or mixtures thereof. It is also possible to construct the emitter body from lanthanides and compounds of lanthanides and borides, such as lanthanum hexaboride and cerium hexaboride. Alternatively, alkaline metals, such as cesium, may be included. The method is also able to incorporate ceramic materials. Any combination of the listed compounds may be used.

The impregnate may be a BaO:CaO:Al ₂ O ₃ mixture or similar composition. However, the process is also applicable to a scandate-type mixture (barium and scandium) . Upon application of heat, the impregnate decomposes. In some embodiments, it is not necessary to provide an impregnate, because the material of the emitter body decomposes to release a compound suitable for lowering the work function of the emitter 10. In case of an impregnate comprising the BaO: CaO: Al ₂ O ₃ mixture, Ba and BaO vapour are transported through the metal matrix to

the surface to form a surface layer with a lower effective work function than the emitter body. Vapour diffusion has been shown to be dependent on the porosity, although it is also dependent on the gas pressure at the surface.

To produce the emitter body, a thermal spraying process is used. The porosity of the body is spatially varied in a continuous manner by adjusting at least one parameter of the spray process whilst depositing the metal that is to form the porous emitter body. This is done so as to provide the core of the emitter 1 with a relatively high porosity. Parameters of a flame spraying process that can be adjusted to vary the porosity are plasma power; plasma gas pressures and flow rates; powder injection details, such as particle size, composition, and carrier flow; torch-substance distance, traverse speed and position, and the roughness of the substrate on which the emitter body is deposited. Parameters of a plasma spraying process that can be adjusted to vary the porosity are similar. It is observed that the thermal spraying process nevertheless requires only a single metal grain size, thus simplifying the manufacture of the emitter 10.

In the illustrated embodiment, regions of the emitter 10 bordering on external surfaces 12-15 other than the emission surface 11 have a porosity increasing in directions from these external surfaces 12-15 towards the interior of the emitter 10. To prevent desorption at the external surfaces 12-15 other than the emission surface 11, the porosity may be reduced to essentially zero at the external surfaces 12-15.

Viewed in a direction z perpendicular to the emission surface 11, the porosity decreases in a core region. In a region bordering on the emission surface 11, the porosity increases in the direction z towards the emission surface 11. This has an effect on the surface roughness, which can thus

be made relatively high. A higher surface roughness implies a larger effective area from which electrons are emitted. A higher porosity in the immediate vicinity of the emission surface 11 also helps to avoid blockage of pores by Tungsten compounds that form during operation of the thermionic dispenser cathode provided with the emitter 10. This effect is particularly prominent in applications in which the emitter 10 is exposed to a plasma environment.

In the embodiment illustrated in Fig. 10, the porosity also varies continuously in a second direction x. Within a central region, away from the external surfaces 12-13, it increases in order to take account of any temperature and pressure gradients in the environment to which the emission surface 11 is exposed. A higher pressure impedes transport of gaseous products away from the emission surface 11 and encourages additional reactions at the emission surface 11, thereby affecting the lifetime of the emitter 10. By varying the porosity in the direction of the expected pressure gradient, uniform emission can still be achieved.

Turning to Fig. 11, a similar effect is provided by appropriate manufacture of the cylindrical emitter insert 8 for the gas-fed hollow cathode 2, which is additionally provided with the orifice plate 5 at its downstream end. In use, an inert gas, e.g. Xe, Kr, Ar, is fed through the hollow cathode 2. The heater coil 9 is provided to apply heat to the emitter insert 8. Electrons emitted from an emission surface 16 serve to maintain a plasma inside the hollow cathode 2.

Due to the presence of the orifice plate 5, the temperature inside the hollow cathode 2 increases quite considerably in a downstream direction corresponding to an axial direction ξ of the emitter insert 8. In order to achieve uniform emission of electrons from the emitter insert 8, at least a central region has a porosity increasing in the axial direction ξ.

Additionally, the porosity decreases in a radial direction r within a region bordering on the emission surface 16, to then increase towards the centre of the cylinder mantle. This increase is provided in order to provide a sufficiently large reservoir of impregnate within the emitter insert 8. The porosity again decreases, preferably to essentially zero, in radial direction within a region bordering on an outer surface 17 of the emitter insert 8.

The porosity may also decrease to zero within regions bordering on end surfaces 18, 19 in directions from the interior of the emitter insert 8 towards those end surfaces 18, 19. As in the other embodiments, this helps to prevent desorption of impregnate to improve the lifetime of the emitter insert.

The thruster 1 may be used in a propulsion system including a further propulsion apparatus, either a cold gas or a resistojet propulsion apparatus. A resistojet propulsion apparatus provides thrust by heating a typically non-reactive fluid. Heating is achieved by sending a current through a resistor. Like a cold gas propulsion apparatus, a resistojet propulsion apparatus provides a relatively limited amount of thrust, but is relatively efficient. To provide a more versatile type of propulsion system, it is proposed here to combine the thruster 1 with a cold gas and/or resistojet propulsion apparatus using the same propellant. The thruster is able to function as part of a larger control system for spacecraft, in particular those carrying electrostatic propulsion systems with a propellant common to the thruster 1. An example of a combined propulsion system is illustrated in Fig. 12. In this example, the thruster 1 is used in conjunction with a cold gas propulsion apparatus.

The propellant is stored in a propellant reservoir 20 at a pressure upwards of 40 bar. It is fed to a plenum 21 through a reduction valve 22. The

plenum 21 is maintained at a pressure of approximately 1 bar. In cold-gas mode, a control valve 23 is switched on and off in quick succession to expel the propellant out through the thruster 1, without the latter being energised.

When more thrust is required, the gold gas propulsion, apparatus is used to supply inert gas from the propellant reservoir 20 to the thruster 1. In this mode of operation, the plenum 21 is maintained at a slightly lower pressure of approximately 300 mbar. The control valve 23 is used to adjust the mass flow into the thruster 1.

In a variant (not shown), an electrically insulating expansion nozzle is provided between the orifice in the orifice plate 5 and the anode 3. Thus, an expansion nozzle (de Laval nozzle) is incorporated at the downstream end of the cathode 2, which allows the cathode to function as a supersonic nozzle in both the cold gas mode of operation and the mode of operation in which the thruster 1 is operational to provide more thrust.

A spacecraft, for instance a satellite in geo-stationary orbit, fitted with the propulsion system illustrated in Fig. 12 is capable of meeting more demanding mission scenarios, such as flying in formation, inspection, rendezvous, and servicing missions, requiring manoeuvres for drag compensation, constellation phasing, and proximity manoeuvring, using the thruster 1. Basic attitude control is carried out using the propulsion system in cold gas mode only. A synergic effect is provided due to the fact that the spacecraft can operate in either mode using a standard, low- cost power system commonly found in commercial satellites and the fact that the same, inert propellant is used in such a way as to prevent buildup of charge on the spacecraft.

The invention is not limited to the embodiments described above, which may be varied within the scope of the accompanying claims. For instance, a porous emitter body may undergo a surface treatment, e.g. to clean it, prior or subsequent to impregnation, in order to form the emitter.