A thruster

Abstract

The invention is directed to a thruster for generating thrust, the thruster comprising: a body having an inlet and an outlet, the inlet for introducing a propellant stream into the body and the outlet providing a nozzle for exiting a heated propellant stream from an interior of the body; a heat transfer element located within the interior of the body and configured to interact with the propellant stream; and a heater positioned to transfer heat to the heat transfer element, wherein the heat transfer element defines a plurality of curved channels connecting the inlet to the outlet to generate turbulent mixing of the propellant stream and to transfer heat energy from the heat transfer element to the propellant stream, to thereby generate thrust as the heated propellant stream is forced through the nozzle of the outlet.

Description

A THRUSTERTECHNICAL FIELD

[0001] The present invention relates to thrusters for spacecraft, aircraft, and satellites. More particularly, the invention is directed to thrusters where thrust is achieved by electro- thermally heating a propellant. The invention is also directed to methods for manufacturing thrusters, and systems for propelling satellites, spacecraft, and aircraft.BACKGROUND

[0002] Resistojets are a type of in-space propulsion thruster that are used for many satellites. Resistojets can provide attitude control and pointing for larger spacecraft and can be used for primary propulsion and orbital manoeuvring of smaller spacecraft.

[0003] Resistojets typically function by heating a fluid propellant before expelling the propellant through a nozzle to generate thrust. Their performance is characterised by their thermal efficiency, mass, and size.

[0004] Specifically, electrothermal propulsion electrically heats the propellant (using a resistive heating element to increase the temperature of the fluid propellant) and then expands the heated propellant though the nozzle. The velocity of the propellant exhausted (and thus the performance of the Resistojet) is primarily a function of temperature of the expelled propellant.

[0005] An example of a conventional resistojet thruster is described in US 7,665,292 B2 which relates to a thruster having electro-thermal thrust augmentation which utilises a catalytic bed to decompose a propellant before heating. The catalytic bed is limited by catalytic attrition, caused by mechanical breakdown from thermal and pressure cycling as well as chemical contaminants in the propellant. This results in a reduction of the thruster's operative life.

[0006] Another example of a conventional resistojet thruster is described in US2023 / 0130545 A1 which relates to an omnivorous solar thermal thruster, and which utilises solar energy as a power source. Solar energy is highly inefficient suffering from heat losses to the environment and is therefore not an efficient heat source. Furthermore, there are operational limitations from solar energy, as no electrothermal augmentation is achieved when the thruster is not in direct sunlight.

[0007] It would be advantageous to provide a thruster with improved thermal efficiency, for example providing an increased propellant outlet temperature to improve efficiency of thrust generation. It would be beneficial to provide a thruster having heat transfer enhancement. It would be advantageous if a thruster could be miniaturised while maintaining high thermal efficiency.

[0008] The present invention was conceived with the above shortcomings in mind.SUMMARY OF THE INVENTION

[0009] The thruster described herein utilises a heat exchange surface with complex geometry enabled through the application of additive manufacturing, to efficiently heat a propellant, allowing the thruster to be miniaturised, to decrease overall dimensions and mass thereof. The thruster utilises at least one heat transfer element configured as mathematical lattice structures to achieve high thermal efficiency for heating propellant, for the propulsion of spacecraft and the like.

[0010] The thruster has significant benefits that result at least in part from the application of a heat transfer element comprising a lattice structure. The lattice structure is characterised by a complex geometry and continuous, smooth internal surfaces.

[0011] The term lattice structure is a generally understood term to refer to a form of cellular structure and is distinct from other cellular structures such as honeycombs and foams. A lattice structure as described herein is understood to relate to a porous, three-dimensional structure formed from a pattern of tessellated, repeating cells with varying topological geometries. While honeycombs, foams and lattices all belong to the family of cellular structures, honeycombs and stochastic foams are a subcategory within cellular structures, and lattices are their own subcategory. For example, honeycomb structures are two- dimensional cellular structures which contain an array of polygon unit cells (not just hexagonal prisms as the name suggests) that have the same shape and size. Stochastic foams comprise cell walls that are randomly orientated in space with high porosity and are commonly found in nature.

[0012] The lattice structure provided in embodiments of the present invention offers improved heat exchange having an increased surface area to volume ratio compared to known heat transfer elements. The complex geometry can induce turbulent mixing to further enhance heat transfer to the propellant. The continuous, smooth, internal surfaces may reduce drops in pressure, preventing the formation of dead zones in the propellantstream. These features result in a thruster with a high thermal efficiency and an improved performance. The thruster will require reduced fuel to produce the same amount of thrust, extending the duration at which any satellite using the thruster can stay in orbit, or enabling manoeuvres to be conducted that were not previously possible.

[0013] In some embodiments, the heat transfer element has minimalist dimensions, which allow for a highly miniaturised thruster. In addition to reducing mass (which can also increase performance) the reduced dimensions make the thruster well suited for providing propulsion to CubeSats and similar small satellites which have inherently strict size and mass constraints. The thruster and heat transfer element comprising the lattice structure can be fabricated with additive manufacturing techniques as a single piece, there are advantages provided in a more cost-effective design and a more reliable, consistent end product.

[0014] Nitrous oxide is a non-toxic, safe and easily transportable type of propellant. In some embodiment, the thruster enables self-sustaining decomposition of nitrous oxide (N2O). The decomposition of N2O is an exothermic reaction, which when sustained during use of thruster enables reducing the amount of energy required to produce thrust as long as propellant stream flow is continued.

[0015] In a first aspect, the invention provides a thruster for generating thrust, the thruster comprising: a body having an inlet and an outlet, the inlet for introducing a propellant stream into the body and the outlet providing a nozzle for exiting a heated propellant stream from an interior of the body; a heat transfer element located within the interior of the body and configured to interact with the propellant stream; and a heater positioned to transfer heat to the heat transfer element, wherein the heat transfer element defines a plurality of curved channels connecting the inlet to the outlet to generate turbulent mixing of the propellant stream and to transfer heat energy from the heat transfer element to the propellant stream, to thereby generate thrust as the heated propellant stream is forced through the nozzle of the outlet.

[0016] In some embodiments, the heat transfer element defines a cavity for receiving the heater therein. The cavity may be a bore that is centrally located in the heat transfer element. The bore may be a blind bore.

[0017] In some embodiment, the cavity is coaxially aligned with the nozzle.

[0018] In some embodiments, the heater is in direct contact with the heat transfer element. In such embodiments, the heat may be transferred by conduction between the heater and the heat transfer element.

[0019] In some other embodiments, the heater is positioned within the heat transfer element with a gap separating the heater from the heat transfer element. In these embodiments, heat may be radiatively transferred between the heater and the heat transfer element, and the propellant stream may be radiatively heated during use of the thruster.

[0020] In some embodiments, the heater is an electrical heater that can be powered by a battery or an alternative power source. The heater may be a resistive heater.

[0021] The heater may be configured to have a maximum operating temperature exceeding 1000°C. For example, the heater may be configured to reliably operate at a temperature of approximately 1100°C. The heater may be configured to heat the heat transfer element to a temperature sufficient for attaining a temperature of the propellant stream up to approximately 800°C. The heater may be configured to heat the heat transfer element while consuming a minimised amount of power. For example, the heater may be configured to heat the heat transfer element to a temperature sufficient for attaining a temperature of the propellant stream up to approximately 800°C while consuming less than 250 Watts of power.

[0022] Where the propellant stream comprises nitrous oxide, the heater may be configured to heat the heat transfer element to a temperature sufficient to initiate nitrous oxide (N2O) decomposition during use of the thruster.

[0023] The decomposition of N2O is an exothermic reaction and as such by heating the heat transfer element to a temperature sufficient to initiate the N2O decomposition, the decomposition produces enough energy and heat to be self-sustaining and reduce the energy demand to heat the heater and maintain the reaction. In this embodiment, the thruster can function thermally efficiently to produce thrust continuously with a reduced need for energy input as long as the flow of the N2O propellant stream is continuous.

[0024] In some embodiments, the heater is a resistive heater comprising ceramic material. The heater may be a ceramic heater. For these embodiments, it may be advantageous to position the heater within the heat transfer element with a gap separating the heater from the heat transfer element, wherein heat is radiatively transferred between the heater and the heat transfer element. Ceramic material is highly brittle and inserting the heater comprising the ceramic material into the centrally located cavity during manufacture of the thruster maybe challenging due to the tight dimensions and resulting tolerance in the cavity to maximise conductive heat transfer from the heater to the heat transfer element during use of the thruster. A heater comprising ceramic material in contact with the heat transfer element may shatter or break during use of the thruster as the heater undergoes temperature cycling due to thermal expansion inside the bore. A heater positioned with a gap separating it from the heat transfer element such that there is no direct contact between the heater and heat transfer element is advantageous as the heater does not have to be inserted directly into the bore in the centre of the thruster and does not have to be limited by the tight tolerance of the bore.

[0025] In some embodiments, the thruster comprises a resistive heater that can reach very high temperatures while minimising the electrical energy required for heating. The ceramic material may in use reach temperatures above 1000°C. The heater comprising the ceramic material may be compact in size.

[0026] In some embodiments, the thruster described herein utilises a resistive heater that enables minimising thermal radiation losses.

[0027] The plurality of curved channels may be configured to split the propellant stream into a plurality of tortuous flows that traverse the heat transfer element. In some embodiments, the curved channels of the plurality of curved channels may be configured to intersect one another via a plurality of holes throughout the heat transfer element providing a plurality of intersection points which induce turbulent mixing between the plurality of tortuous flows of the propellant stream as the propellant stream traverses the heat transfer element. In some embodiment, the plurality of intersection points increases the divergence of the plurality of tortuous flows between the inlet and the outlet, to increase heat transfer therebetween.

[0028] The configuration of the heat transfer element facilitates efficient heat transfer from the heater to the heat transfer element, and from the heat transfer element to the propellant stream. The plurality of curved channels may be configured to induce turbulent mixing of the propellant stream as the propellant stream traverses the heat transfer element thereby achieving efficient mixing of the plurality of tortuous flows of the propellant stream to increase the propellant stream's bulk thermal energy. The plurality of tortuous flows can also provide advantages because they may disrupt a near wall viscous layer in the turbulent flows of the propellant stream to promote turbulent mixing which enhances heat transfer from the heat transfer element to the propellant stream.

[0029] The heat transfer element may comprise a lattice structure.

[0030] The lattice structure may define a complex geometry. The lattice structure may define the plurality of curved channels traversing the lattice structure. In one embodiment, the lattice structure and plurality of curved channels have continuous, smooth internal surfaces. The plurality of curved channels traversing the lattice structure may induce strong turbulent mixing splitting the propellant stream into the plurality of tortuous flows to enhance heat transfer thereto. The continuous, smooth surfaces may lower pressure drops and prevent the formation of dead zones in the propellant flow. The features of surface area to volume ratio, complex geometry, and continuous smooth surface result in a thruster with an increased thermal efficiency and an improved performance when compared to conventional thrusters. The thruster uses less fuel to produce thrust, extending the duration at which objects such as satellites can stay in orbit. In addition, the reduction in fuel can allow the thruster to conduct manoeuvres that are not currently possible, using traditional thrusters.

[0031] In one embodiment, the lattice structure may comprise the plurality of curved channels, each channel of the plurality of curved channels traversing the lattice structure being separated from an adjacent curved channel by a channel wall. A thickness of the channel walls may be constant. In some embodiments, the thickness of the channel walls may vary. In some embodiments, the thickness of the channel walls decreases with increasing distance from a location of the heater within the thruster. Where the heater is positioned within the cavity defined by the heat transfer element, the thickness of the channel walls may decrease with increasing distance from the heater. In some embodiments, the thickness of the channel walls may range from 0.05mm to 5mm. In some embodiments, the thickness of the channel walls may vary from 0.1 mm to 1.5 mm, with the walls thinning towards an exterior of the thruster. In some embodiments, the walls taper from a thickness of 1.0 mm - 0.8 mm in proximity to the heater and reduce to 0.6mm - 0.4mm in proximity to an external periphery of the heat transfer element, adjacent the exterior of the thruster. This is advantageous in that the conductive surface area of each wall is increased in close proximity to the heater to increase transfer of thermal energy more effectively from the heater into the lattice structure of the heat transfer element. Where a central core of the heat transfer element is in direct contact with the heater, there is advantage to increasing the thickness of the walls in contact / proximity to the heater because it increases the surfacematerial (area) available to be heated and thereby increases the transfer of heat energy into the heat transfer element.

[0032] The lattice structure may comprise an array of a plurality of unit cells, the unit cells of the plurality of unit cells being tessellated to form the lattice structure. This is advantageous in that the lattice structure offers an increased surface area to volume ratio compared to conventional heat transfer elements. A volume of each cell of the array of tessellated cells may range from 1 mm3to 8000mm3. The dimensions of each cell may range from 1x1x1 mm to 20x20x20 mm. In one embodiment, the dimensions of each cell of the array of tessellated cells may be 6x6x6 mm, providing a volume of 216 mm3. In another embodiment, the dimensions of the cells of the array of tessellated cells can be scaled within the lattice structure. That is, the dimensions of the cells can vary in size from one end of the thruster to an opposing end of the thruster. In yet another embodiment, the distribution of the cells of the array of tessellated cells can vary within the lattice structure. In yet another embodiment, the distribution of the cells of the array of tessellated cells can be evenly distributed within the lattice structure. The cells can be varied in size and varied such that they are evenly distributed. There can be embodiments with a graduation in size from one end of the heat transfer element to the other, referred to as scaling or warping of the unit cell size distribution.

[0033] The array of unit cells may provide varying topological geometries.

[0034] In some embodiments, the lattice structure comprises an array of unit cells that are tessellated to form a Triply Periodic Minimal Surface (TPMS) lattice structure. In other words, the lattice structure may be a TPMS lattice structure.

[0035] In some embodiment, each unit cell of the array has a gyroid topological geometry. The lattice structure may be a Gyroid Triply Periodic Minimal Surface lattice structure.

[0036] In another embodiment, each unit cell of the array has a diamond topological geometry. The lattice structure may be a Diamond Triply Periodic Minimal Surface lattice structure. This is advantageous in that the Diamond TPMS lattice structure has an improved thermal performance to that of a Gyroid TPMS lattice structure. Numerical simulations indicate that the Diamond TPMS lattice structure can enhance the heat transfer efficiency of the lattice structure, resulting in a 300°C propellant stream outlet temperature, as compared to a temperature of 297°C for the propellant stream as simulated under the same conditions from a Gyroid TPMS lattice structure.

[0037] In embodiments wherein the heat transfer element comprises a lattice structure, the lattice structure may define the cavity for receiving the heater therein. The cavity may be a bore. The bore may be a blind bore in the lattice structure. The cavity may be coaxially aligned with the nozzle of the outlet. The cavity may be a bore and may be centrally located in the lattice structure. The central location of the heater within the lattice structure can reduce radiation loss to the environment as the thermal energy conducted away from the heater in a radial direction is primarily transferred to the propellant stream, thereby reducing the temperature of the exterior surface of the thruster. In addition, the centrally located heat source, and a low power operation of the heater, in combination, can reduce a need for radiation shielding.

[0038] In some embodiments, radiation shielding may be omitted entirely.

[0039] In some other embodiments, radiation shielding may be incorporated to further reduce thermal energy loss from the exterior of the thruster. For example, radiation shielding may sheath the body to reflect radiation back toward the lattice structure of the heat transfer element.

[0040] In some embodiments, the heat transfer element may include a reinforcing member. The lattice structure of the heat transfer element may comprise the reinforcing member. The reinforcing member may be formed from a densified portion of the heat transfer element. The reinforcing member may be formed integrally with the heat transfer element and advantageously provides structural support for the heater within the heat transfer element. The reinforcing member may extend from the cavity defined within the heat transfer element and terminate at a free end of the lattice structure. The reinforcing member may extend from the inlet and terminate at a point adjacent a free end of the lattice structure. The reinforcing member may extend longitudinally within the heat transfer element. The reinforcing member may have a variable thickness. The thickness of the reinforcing member may taper towards the free end of the lattice structure. The reinforcing member may be formed as a separate insertable component. The reinforcing member may taper towards the free end of the lattice structure. The reinforcing member may taper to a conical tip.

[0041] In some embodiments, the inlet may be coaxially aligned with a longitudinal axis of the body. In some embodiments, the inlet may be arranged tangentially to the longitudinal axis of the body, such that the propellant is introduced into the body at an inclination angle to a direction of propellant flow through the thruster (along the longitudinal axis). The inletmay be positioned at an inclination angle of approximately 90 degrees to the longitudinal axis of the body. The tangential inlet arrangement can facilitate heater placement in the centre of the heat transfer element, which in turn can increase the efficiency of the thruster by decreasing the amount of thermal radiation energy that is lost to the surrounding environment. The tangential inlet arrangement is designed to promote uniform flow distribution of the propellant stream into the heat transfer element, for example the lattice structure of the heat transfer element. The tangential inlet further causes the propellant stream to swirl around an internal surface of the thruster in a circular motion, referred to as centrifuging. This aids effective vaporization of the propellant stream on entering the thruster body. This effect is exemplified in low gravity conditions.

[0042] In some embodiments, the body of the thruster provides a first end and a second end. The inlet may be positioned towards the first end of the body. In some embodiments, the inlet may be positioned towards the second end of the body.

[0043] In some embodiments, the inlet may be positioned adjacent to the nozzle.

[0044] In some embodiments, the inlet is positioned distal to the nozzle.

[0045] In some embodiments, the thruster may comprise a plenum chamber within the body. The plenum chamber may be configured to guide the propellant stream in a circulating motion within the heat transfer element. In some embodiments, the plenum chamber may be positioned between the heat transfer element and the inlet.

[0046] The plenum chamber may be annular. The plenum chamber may have an internal radius. The internal radius of the plenum chamber guides the propellant stream into the heat transfer element and can assist in reducing thermal hotspots. When the propellant stream enters the thruster, it impinges on the plenum chamber radius and the sloped wall about the radius guides the propellant stream into the heat transfer element efficiently. The internal radius of the plenum chamber may be between 2mm and 5mm. In some embodiments, the plenum chamber radius may be about 3mm.

[0047] In some embodiment, the lattice structure of the heat transfer element may be annular or cylindrical. The lattice structure may have an internal radius.

[0048] An annular or cylindrical lattice structure and an annular plenum chamber are advantageous in that the plenum chamber guides the propellant stream in a circulating motion around the cavity or bore, and within the heat transfer element, with a funnel-likeeffect, reducing the occurrence of thermal hotspots within the lattice structure of the heat transfer element.

[0049] The body may be manufactured from a range of different conductive and non- conductive materials, for example steel, aluminium, nickel, copper, and alloys thereof. In some embodiments, the heat transfer element may be manufactured from the same material as that of the body. In some embodiments, the heat transfer element may be manufactured from a different material to that of the body.

[0050] The body and heat transfer element may be manufactured from GRCop-42 alloy. GRCop-42 has a high thermal conductivity (approximately 280 to approximately 325 W / mK), strong oxidation resistance and high service temperature (also known as an operating temperature) making it highly desirable for use in a heat transfer element.

[0051] The body and heat transfer element may be manufactured using additive manufacturing. Additive manufacturing may be used to manufacture the heat transfer element and / or the body as separate components for later assembly. The body and heat transfer element may be manufactured using three-dimensional (3D) printing techniques. The body and the heat transfer member may be manufactured together as a unitary component, thereby obviating the need for the formation of internal joints within or between the various parts of the thruster.

[0052] In some embodiments, the body and the heat transfer element may be manufactured contemporaneously as a monolithic piece. Additive manufacturing may be used to manufacture various embodiments of the body and heat transfer element in combination, as a monolithic piece. The manufacturing process involves the layer-by-layer deposition of a material by computer control, to form a three-dimensional object. Integrally forming / manufacturing the body and the heat transfer element provides advantage in that no joints or angled edges are formed in the thruster. These joint conditions and edges can cause areas of increased mechanical stress and strain and by eliminating these concentrations the thruster is made more robust and less susceptible to damage.

[0053] Additive manufacturing may provide advantages for production and assembly (in comparison to manufacturing traditional thrusters) because it combines the manufacture of multiple parts into a single operation. Instead of creating individual parts and assembling them at a later point, additive manufacturing combines the manufacturing actions and eliminates the need for assembly processes. This can improve the reliability, tolerance, andconsistency of the finished thruster, while reducing costs, labour, and production, time compared with conventional thruster manufacture.

[0054] The thermal conductivity K of a material is a measure of the material's ability to conduct heat (thermal energy)Q d = K A 8T

[0055] In the above equation: Q is the amount of heat transferred; d is the distance between two isothermal planes; K is the thermal conductivity of the material; A is the surface area; and 8T is the difference in temperature.

[0056] Furthermore, radiation heat transfer is typically a primary mode of thermal loss for thrusters operating at high temperatures. A rate of radiation heat loss is governed by Stefan-Boltzmann Law:Q = CT S A (Ts4- Ta4)

[0057] In the above equation, ct is the Stefan-Boltzmann constant (equal to 5.67 x 10’8J / sm2K4), e is the material emissivity, A is the surface area, Tsis the surface temperature and Tais the ambient temperature.

[0058] The rate of heat transfer via radiation being proportional to the fourth square of the surface temperature, thermal design to minimise radiation losses is critical for thrusters operating at high temperatures.

[0059] Material emissivity is a relative measure of an ability of an object to emit infrared radiation compared to a blackbody, which is a perfect emitter of radiation and has an emissivity of 1. Dark, rough and oxidised surfaces are much more effective at emitting and absorbing radiation and thus have high emissivity values. Light, clean and polished materials will have a very low emissivity as they tend to reflect infrared radiation. Thus, the surface emissivity can be controlled through the selection of materials and post processing to either absorb the radiation emitted from the heater or reflect radiation that is emitted from the thruster body.

[0060] A common method to reduce radiation losses for high temperature thrusters is using a radiation shield. Radiation shields are usually thermally isolated cylinders that surround the thruster and are constructed from highly polished aluminium or steel to reflect the thermalradiation emitted from the thruster back onto itself. Highly polished steels have a surface emissivity of 0.075 and can thus reflect more than 90% of incoming infrared radiation.

[0061] In embodiments for which the propellant stream is nitrous oxide and the heater is a resistive heater comprising a ceramic material, numerical simulations indicate that radiation shielding of the thruster body can assist attaining a temperature of the heat transfer element sufficient for initiating the thermal decomposition of nitrous oxide during use of the thruster. A temperature suitable for initiating thermal decomposition of nitrous oxide typically is between approximately 700°C and approximately 800°C, or approximately 800°C. A resistive heater comprising a ceramic material used in the thruster can typically reach a temperature close to or approaching a maximum reliable operating temperature of approximately 1000°C to 1100°C.

[0062] In some embodiments, the thruster comprises a radiation shield surrounding the body. The radiation shield may comprise stainless steel.

[0063] In some embodiments, the thruster further comprises a plurality of cooling channels, each cooling channel configured to provide a passage for the propellant stream from the inlet to the heat transfer element.

[0064] The cooling channels may comprise GRCop-42 alloy.

[0065] The thruster may comprise at least two cooling channels, at least five cooling channels, or at least ten cooling channels.

[0066] In some embodiment, each cooling channel of the plurality of cooling channels is embedded and extends longitudinally in an outer wall of the thruster. The cooling channels of the plurality of cooling channels may be arranged spaced apart and circumferentially in the outer wall of the thruster.

[0067] The cooling channels may take the form of a series of small fluid channels embedded in the outer wall of the thruster. The cooling channels are commonly used in rocket engines and spacecraft thrusters for cooling the thruster wall to prevent material failure.

[0068] In use, a cold propellant stream is fed through the cooling channels before being injected into the heat transfer element. The flow of the cold propellant stream through the cooling channels results in a cooling of the thruster outer wall as the propellant stream absorbs the heat from the solid thruster outer wall. The thruster outer wall temperature decreases while the propellant stream temperature increases. In addition to preventing material failure, the reduction in temperature of the outer wall improves the efficiency of thethruster as lower outer wall temperatures emit less heat, and more heat is retained within the thruster. The propellant is pre-heated in the cooling channels before being injected into the heat transfer element, and higher propellant temperatures can thus be achieved, which improves specific impulse (propulsion efficiency).

[0069] The cooling channels are termed "regenerative" because they use the propellant stream itself to cause the cooling of the thruster outer wall, which simultaneously results in an increase of the propellant stream temperature as the propellant stream absorbs the heat from the solid thruster outer wall. Cooling channels contribute further to minimising thermal radiation losses.

[0070] Once the propellant stream passes through the cooling channels, it may be uniformly distributed into the plenum chamber via a ring of injector orifices, each injector orifice having a same cross-sectional area as a respective cooling channel. This configuration assists enforcing uniform flow distribution into the heat transfer element and subsequently a uniform radial temperature distribution, preventing the formation of hot spots.

[0071] In embodiments where the propellant stream is nitrous oxide and the heater is a resistive heater comprising a ceramic material, the use of GRCop-42 alloy for the body, heat transfer element and regenerative cooling channels may be particularly advantageous. The heat transfer element requires a high thermal conductivity to reduce warm-up time in use of the thruster while simultaneously having a high melting temperature and strong oxidation resistance to prevent degradation under the environment with hot oxygen reactants under which nitrous oxide decomposes.

[0072] In some embodiments, the thruster may be a resistojet. In some embodiments, the thruster may be a miniaturised thruster in comparison to conventional thrusters, wherein a length and a width of the thruster is greatly reduced as compared to a length and a width of conventional thruster, without significant reduction in performance (predominantly generation of thrust).

[0073] In a second aspect, the invention provides a method of generating thrust comprising the steps of: supplying a propellant stream to an interior of a body; propelling the propellant stream through a plurality of curved channels in an interior of the body; heating the propellant stream as it traverses the plurality of curved channels and splitting the propellant stream into a plurality of tortuous flows; and expelling the heated propellant stream from the interior of the body through an outlet nozzle to generate thrust.

[0074] The method may be implemented using the thruster provided in accordance with the first aspect.

[0075] Embodiments herein also relate to a satellite comprising a thruster as herein described.

[0076] Embodiments herein also relate to a spacecraft comprising a thruster as herein described.

[0077] In some further aspect, the technology defined and described in embodiments of the present invention may be used for a device other than a thruster.

[0078] In a third aspect, the invention provides a device for decomposing nitrous oxide, the device comprising: a body having an inlet for introducing a stream of nitrous oxide into the body; a heat transfer element located within the body and configured to interact with the stream of nitrous oxide; and a heater positioned to transfer heat to the heat transfer element, wherein the heat transfer element defines a plurality of curved channels connecting the inlet to an outlet, the plurality of curved channels arranged to generate turbulent mixing of the stream of nitrous oxide, and to transfer heat energy from the heat transfer element to the stream of nitrous oxide, and wherein the heater is configured to generate and transfer heat to the heat transfer element to cause decomposition of the nitrous oxide.

[0079] In an embodiment, the heater is a resistive heater comprising ceramic material.

[0080] In an embodiment, the heater is positioned within the heat transfer element with a gap separating the heater from the heat transfer element, the heat transfer element receiving heat energy from the heater by radiation.

[0081] In an embodiment, the plurality of curved channels is configured to split the stream of nitrous oxide into a plurality of tortuous flows that traverse the heat transfer element.

[0082] In an embodiment, the heat transfer element comprises a lattice structure.

[0083] The lattice structure may comprise the plurality of curved channels, each curved channel of the plurality of curved channels being separated from an adjacent curved channel by a channel wall.

[0084] The lattice structure may comprise an array of a plurality of unit cells that are tessellated to form a Triply Periodic Minimal Surface (TPMS) lattice structure. Each unit cell of the array has a diamond topological geometry and the lattice structure is a Diamond Triply Periodic Minimal Surface lattice structure.

[0085] In an embodiment, the body and heat transfer element are manufactured from GRCop-42 alloy.

[0086] In an embodiment, the body and heat transfer element are manufactured as a monolithic piece using additive manufacturing.

[0087] In some embodiment, the body and heat transfer element are manufactured using 3D printing. The thruster may be manufactured using 3D printing using a suitable material.

[0088] In some embodiment, the cooling channels are 3D printed directly into the thruster outer wall. This can assist reducing the complexity and cost of a manufacturing of the thruster. 3D printed regenerative cooling channels can also have improved thermal performance, as the rough surface finish on the interior of the cooling channel that results from the printing process increases the heat transfer rate from the thruster outer wall to the cold propellant stream during use.

[0089] Embodiments of the device present significant advantages in that they not only enable initiating decomposition of the nitrous oxide but also enable sustaining the reaction of decomposition of the nitrous oxide in a stable manner. These advantages result at least in part from the use of materials for the device that enable reaching and sustaining an operating temperature of the stream of nitrous oxide that causes the reaction of decomposition of the nitrous oxide, and from the application of a structure of the heat transfer element that has a complex geometry with continuous and smooth internal surfaces,

[0090] In a fourth aspect, there is provided a method to decompose nitrous oxide, the method comprising: supplying a stream of nitrous oxide to an interior of a body transferring heat to a heat transfer element located within the body, the heat transfer element comprising a plurality of curved channels to generate turbulent mixing of the stream of nitrous oxide; and heating the stream of nitrous oxide to a temperature that causes the decomposition of the nitrous oxide by passing the stream of nitrous oxide through the plurality of curved channels.

[0091] The method may be conducted using the device provided in accordance with the third aspect.BRIEF DESCRIPTION OF THE DRAWINGS

[0092] Embodiments of the invention will now be described with particular reference to the accompanying drawings. However, it is to be understood that the features in and describedwith reference to the accompanying drawing are illustrated by way of example, and not by way of limitation, of which:

[0093] Figure 1 A is a cross-sectional view of a thruster according to one embodiment of the present invention, the section taken along line A— A (a longitudinal axis of the thruster), illustrating a body, a heat transfer element, a cavity which is a bore, and a nozzle of the thruster;

[0094] Figure 1 B is a cross-sectional view of the thruster of Figure 1 A with an indication of the flow of the propellant stream;

[0095] Figure 2A is a perspective cross-sectional view along line A— A of the thruster of Figure 1, illustrating a plenum chamber and a plenum chamber exit;

[0096] Figure 2B is a cross-sectional view of the thruster of Figure 2A with an indication of the flow of the propellant stream;

[0097] Figure 3 is a perspective cross-sectional view along line B — B of the thruster of Figure 1 B, illustrating the cavity which is a bore, a plurality of curved channels and channel walls of the heat transfer element;

[0098] Figure 4 is a perspective view of the heat transfer element of Figures 1a and 1 b, illustrating a plurality of curved channels and a plurality of holes through the lattice structure of the heat transfer element, and illustrating a core of the heat transfer element;

[0099] Figure 5 is a cross-sectional view along C— C of the heat transfer element of Figure 4, illustrating a core portion and an outer portion of the heat transfer element;

[0100] Figure 6 is a perspective view of the heat transfer element, illustrating a plurality of curved channels and a plurality of holes through the lattice structure of the heat transfer element, and illustrating a core of the heat transfer element, the heat transfer element having a diamond Triply Periodic Minimal Surface (TPMS) lattice structure;

[0101] Figure 7 is a cross-sectional view cut through an embodiment of a thruster according to the invention along a central axis X, illustrating an inlet and a nozzle positioned in proximity to one another at one end of the thruster;

[0102] Figure 8 is a cross-sectional view cut through an embodiment of a thruster according to the invention along a central axis X, illustrating a heat transfer element having a central bore, a core of the heat transfer element and a reinforcing member tapering to a conical tip tapering towards a free end of the heat transfer element;

[0103] Figure 9A is a computational heat map simulation of the temperature of the body of the thruster and of the temperature of the propellant stream wherein the heat transfer element is configured as a Gyroid TPMS lattice;

[0104] Figure 9B is a computational heat map simulation of the temperature of the body of the thruster and of the temperature of the propellant stream wherein the heat transfer element is configured as a Diamond TPMS lattice;

[0105] Figure 10 is a thermal simulation of a propellant stream flowing through the heat transfer element in the thruster of Figure 1, illustrating a plurality of flows of the propellant stream that interleave and intersect with one another and a gradual increase in temperature as the propellant stream advances from the inlet to the outlet of the thruster when the heat transfer element is configured as a Diamond TPMS lattice;

[0106] Figure 11 A is a perspective view of a thruster according to one embodiment, the thruster cutaway along a longitudinal axis X of the thruster) to illustrate the internal components, including a radiation shield, a body, cooling channels, a heat transfer element, a bore, a heater, an inlet and a nozzle of the thruster;

[0107] Figure 11 B is a perspective view of the thruster of Figure 11A with an indication of the flow of the propellant stream;

[0108] Figure 12 is a perspective view of a ceramic heater suitable for use in the thruster;

[0109] Figure 13 is a perspective view of a thruster, cut across the longitudinal axis X-X as shown in Figure 11 A, illustrating a central cavity, a plurality of curved channels and channel walls of a heat transfer element and a plurality of cooling channels embedded in an outer wall of the thruster;

[0110] Figure 14 is a cross-sectional view cut along the longitudinal axis X of the thruster of Figure 11 A, illustrating the central cavity, an inlet and an outlet, a diffuser, the plurality of curved channels and channel walls of the heat transfer element and the plurality of cooling channels;

[0111] Figure 15 is a schematic of a concentric cylinder arrangement used for a thermal model of a thruster, in accordance with an embodiment;

[0112] Figure 16 is a schematic of a thermal network used for the thermal model of Figure 15;

[0113] Figure 17 is graph illustrating variation in a thruster solid body temperature at steady state with and without a radiation shield;

[0114] Figure 18 is a block diagram of a propulsion system using the thruster of Figure 1, illustrating a propellant tank, a filter, an isolation valve, a thruster valve, and a controller;

[0115] Figure 19 is a perspective view of the thruster of Figure 1 or Figure 11A, illustrating the thruster connected to a valve and a connection cable of the propulsion system of Figure 18;

[0116] Figure 20A is a heat map from a thermal simulation of a body manufactured from steel, illustrating thermal hotspots and thermal stress across the heat transfer element;

[0117] Figure 20B is a heat map from a thermal simulation of a body manufactured from GRCop-42 alloy, illustrating thermal hotspots and thermal stress across the heat transfer element;

[0118] Figure 21 A is a three-dimensional representation of a unit cell having a gyroid topological geometry, which can be used in one embodiment to form the array of unit cells tessellated to form the lattice structure;

[0119] Figure 21 B is three-dimensional representation of a unit cell having a diamond topological geometry, which can be used in one embodiment to form the array of unit cells tessellated to form the lattice structure;

[0120] Figure 22 is a flow chart of an embodiment of a method of generating thrust;

[0121] Figure 23 is a schematic representation of an embodiment of a device for decomposing nitrous oxide; and

[0122] Figure 24 is a flow chart of an embodiment of a method of decomposing nitrous oxide.

[0123] Embodiments will now be described more fully hereinafter with reference to the accompanying drawings, in which various embodiments, although not the only embodiments, of the invention are shown. The invention may be embodied in many different forms and should not be construed as being limited to the embodiments described below.DETAILED DESCRIPTION OF EMBODIMENTS

[0124] While embodiments of the invention are described herein in relation to a resistojet thruster for use in a satellite, it is contemplated that the thruster as described herein can also have valuable application when used in spacecraft and aircrafts.

[0125] With particular reference to Figures 1A, 1B, 2A and 2B, there is illustrated a thruster 1 for generating thrust T, the thruster 1 comprising: a body 2 having an inlet 11 and an outlet 3a, the inlet 11 introducing a propellant stream S into the body 2 and the outlet 3a providing a nozzle 33 for exiting a heated propellant stream H from an interior 19 of the body 2; a heat transfer element 20 located within the interior 19 of the body 2 configured to interact with the propellant stream S; and a heater 30 positioned to transfer heat to the heat transfer element 20. The heat transfer element 20 defines a plurality of curved channels 25 connecting the inlet 11 to the outlet 3a, to generate turbulent mixing of the propellant stream S and to transfer heat energy from the heat transfer element 20 to the propellant stream S, as the propellant stream S traverses the heat transfer element 20, to thereby generate thrust T as the heated propellant stream H is forced through the nozzle 33 of the outlet 3a.

[0126] The nozzle 33 increases a velocity of the heated propellant stream H as the stream exits the body 2 via the outlet 3a. The outlet 3a defines the nozzle 33 and then transitions into and expands to define a diffuser 34 through which the heated propellant H is ejected to generate thrust T. In some embodiments, the outlet 3a can have a radius of between approximately 3mm and approximately 20mm. In some variants, the radius of the outlet 3a may be about 7mm and may have a surface finish ranging from approximately 20 microinches to approximately 35 microinches. It will further be understood that the outlet 3a may have any other dimensions as considered suitable by a person skilled in the art in the context of the present disclosure.

[0127] As shown in Figures 1 A and 2A the body 2 is cylindrical in shape and has an exterior surface 2a and an interior surface 2b, the interior surface 2b defining the interior 19 of the body 2. The body 2 extends from a first end 5 to a second end 8 and has a central longitudinal axis X. The body 2 has a length LB measured from the first end 5 to the second end 8, and an internal radius R. The thruster 1 has a length LT, which equals the length LB of the body 2 plus the length of the diffuser 34. In some embodiments, the body 2 can have a length ranging from approximately 25mm to approximately 200mm and the internal radius R of the body 2 can have a range of approximately 10mm - 65mm. It will be understood that the body 2 and its internal radius may have any other dimensions as would be considered suitable by a skilled person in the context of the present disclosure.

[0128] The heat transfer element 20 of thruster 1 defines a cavity 27 for receiving the heater 30 therein. The cavity 27 is illustrated in Figure 2A as a bore. More specifically, the cavity 27 is a blind bore. The cavity 27 is coaxially aligned with the nozzle 33. The cavity 27 is centrally located in the heat transfer element 20 and the heater 30 positioned therein is centrally located within the heat transfer element 20. In this embodiment, the heater 30 is illustrated to be in direct contact with the heat transfer element 20. This arrangement provides an even distribution of heat (thermal energy) into the heat transfer element 20 but may be varied for different applications. In other embodiments, the heater 30 can be located within the heat transfer element 20 but not to be in direct contact therewith, such that a gap 7 is formed between the heater 30 and the heat transfer element 20 - this will be described in further detail in relation to Figure 11 A.

[0129] Although not illustrated, it will be understood that in some other embodiments, depending on the application of the thruster 1 and the components in proximity thereto, the heater may not be centrally located. An uneven heat distribution through the body may be configured by placing the heater 30 in closer proximity to a first portion of the body 2 and distal to a second portion of the body 2.

[0130] In some embodiments, the heat transfer element 20 is cylindrical in shape and the cavity 27 therein is cylindrical. The heater 30 fitted therein is also cylindrical in shape.

[0131] In thruster 1, the heat transfer element 20 comprises a lattice structure 22. The lattice structure 22 defines the cavity 27, which in the illustrated embodiment is a blind bore centrally located within the lattice structure 22 for receiving the heater 30 therein.

[0132] In thruster 1, the first end 5 comprises a circular mounting flange 4, having at least one mounting aperture 6 for mounting the thruster 1 to a propellant source for use. The first end 5 of the body 2 defines a plenum chamber 17 for receiving the propellant stream S and forcing the propellant stream S into the heat transfer element 20 and lattice structure 22 of the heat transfer element 20. The plenum chamber 17 is annular. The plenum chamber 17 at least partially surrounds the cavity 27. The plenum chamber 17 has an internal radius r and smooth surface walls for swirling the propellant stream S around the heat transfer element 20.

[0133] An opening 21 in the body 2 provides access to the cavity 27 from the exterior of the body 2. In some embodiments, the cavity 27 has a radius ranging from approximately 2mmto approximately 20mm. In one example, the cavity 27 may have a radius of approximately 3mm. The cavity 27 may have a length ranging from approximately 10mm to approximately 50mm. In one example, the cavity 27 may have a length of approximately 24mm. It will be understood that the cavity may have any other dimensions as would be considered suitable by a skilled person in the context of the present disclosure.

[0134] The cavity 27 may be formed by drilling or reaming the heat transfer element 20. In some embodiments, the lattice structure 22 can be configured such that the central cavity 27 is geometrically created within the heat transfer element 20.

[0135] In some embodiments, the heater 30 is frictional fitted inside the cavity 27, wherein a shape of the heater 30 corresponds to the internal shape of the cavity 27.

[0136] The central location of the heater 30 within the heat transfer element 20 reduces radiation loss to the environment as the thermal energy conducted away in a radial direction continues to be transferred to the propellant stream S, thereby reducing the temperature at the exterior surface 2a of the body 2. In such embodiments, it is possible to omit radiation shielding from within or surrounding the thruster 1, in other embodiments the radiation shielding material can be reduced.

[0137] The heat transfer element 20 can additionally comprise a reinforcing member 31 therein providing structural support for the heater 30. The reinforcing member 31 can extend from the cavity 27 to terminate at a free end 23. A thickness of the reinforcing member 31 can taper along a length thereof. The reinforcing member 31 can provide a conical tip or a cone shape.

[0138] With reference to Figure 2A, the inlet 11 is positioned at the first end 5 of the body 2, and the nozzle 33 is formed towards the second end 8 of the body 2. The inlet 11 is thus positioned distal to the nozzle 33. The inlet 11 is arranged tangentially to the body 2. The inlet 11 opens into the annular plenum chamber 17 within the body 2. The tangential location of the inlet 11 is advantageous because it allows the heater 30 to be placed in the centre of the heat transfer element 20, which increases the efficiency of the thruster 1 by decreasing the amount of thermal radiation energy that is lost to the surrounding environment. The tangential arrangement of the inlet 11 is designed to allow uniform flow distribution of the propellant into the heat transfer element 20, and to cause the propellant stream S to swirl around a surface boundary 16 of the plenum chamber 17 in a circularmotion, referred to as centrifuging, which aids in effective vaporisation of the propellant stream S after entering the thruster 1, particularly in low gravity.

[0139] The heat transfer element 20 extends from the plenum chamber 17 towards the second end 8 of the body 2, guiding the propellant stream S through the body 2. Adjacent the outlet 3a of the body 2, the heat transfer element 20 terminates to define the free end 23 of the lattice structure 22. An air gap 24 is formed between the free end 23 of the lattice structure 22 and the outlet of the body 3a. The air gap 24 is peripherally bounded by the nozzle 33. In some embodiments, an internal radius R of the body 2 is smaller than the internal radius r of the plenum chamber 17 (see Figure 1 B). The plurality of curved channels 25 is configured to split the propellant stream S into a plurality of tortuous flows 32 that traverse the heat transfer element 20. The lattice structure 22 is formed by the plurality of curved channels 25. The curved channels 25 have continuous, smooth internal surfaces.Each curved channel 25 is separated from an adjacent curved channel 25 by a channel wall 26. The plurality of curved channels 25 and channel walls 26 interleave throughout the heat transfer element 20. In some embodiments, the free end 23 of the lattice structure 22 may comprise a rotor shaped termination of the plurality of channel walls 26 and curved channels 25.

[0140] Figure 10 further provides an illustration of a plurality of flows 32 of the propellant stream S that interleave and intersect with one another while weaving through the lattice structure 22 of the heat transfer element 20 after the propellant stream S enters the body 2 via the inlet 11 and flows towards the outlet 3a. As the propellant stream S traverses the lattice structure 22 of the heat transfer element 20, the propellant stream S is heated. Variations in temperature of the propellant stream S and heated propellant stream H are shown in Figure 10. With additional reference to Figures 1 to 5, in some embodiment, from the first end 5 of the body 2 at the inlet 11 to the second end 8 of the body 2 at the outlet 3a, a temperature of the propellant stream S, H can vary from approximately 300K to approximately 600K at a steady state of the thruster 1. It will be understood that when not in the steady state, the thruster 1 can operate such that the propellant stream S, H can reach higher temperatures wherein, for example, the propellant stream S, H may have a temperature varying from approximately 300K to approximately 800K.

[0141] With reference to Figures 1 to 5, the curved channels 25 form a series of holes through the heat transfer element 20. The series of holes comprises holes 29 arranged inand around a central core 12 of the heat transfer element 20 (shown in Figure 5). The series of holes further comprises peripheral holes 28 arranged in and around an outer peripheral portion 13 of the heat transfer element 20 (shown in Figure 4). The curved channels 25 and the channel walls 26 are illustrated in Figures 2 to 5. The plurality of holes 28, 29 together defines a plurality of intersection points between the curved channels 25 whereby the propellant stream S can be repeatedly split to form the plurality of tortuous flows 32 that intersect with one another while weaving through the lattice structure 22. The plurality of intersection points induces turbulent mixing between the plurality of tortuous flows 32 of the propellant stream S as the propellant stream S traverses the heat transfer element 20. The plurality of intersection points increases the divergence of the plurality of tortuous flows 32 between the inlet 11 and the outlet 3a, to increase heat transfer therebetween.

[0142] The central core 12 supports a weight of the heater 30. It can provide direct contact between the heater 30 and the heat transfer element 20. It will however be understood that in some embodiment described herein, the central core 12 can be configured to enable the gap 7 to be formed between the heater 30 and heat transfer element 20 such that the heater and heat transfer element do not have direct contact. The cavity 27 is a bore that is centrally located within the central core 12 of the heat transfer element 20. The cavity 27 can be a blind bore. The cavity 27 is located within the heat transfer element 20 such that the cavity 27 is at least partially surrounded by the plenum chamber 17. The cavity 27 extends from the first end 5 of the body 2 towards the free end 23 of the lattice structure 22 proximate the air gap 24. The cavity 27, the nozzle 33, and the central longitudinal axis X are coaxially aligned. The internal radius "r" of the plenum chamber 17 guides the flow of propellant stream S around the cavity 27 and can prevent thermal hotspots.

[0143] The flow of the propellant stream S is indicated in Figures 1 B and 2B for clarity. Propellant stream S is introduced via the inlet 11 at the first end 5 of the body 2. The propellant stream S then flows into the annular plenum chamber 17 from the inlet 11 and is propelled out into the heat transfer element 20 via a chamber exit 17a. The inlet 11 and the plenum chamber exit 17a are substantially perpendicular to each other to start the circulation of the propellant stream S about the plenum chamber 17 and thus about the central core 12 and heater 30 therein. The internal radius r of the annular plenum chamber 17 in some embodiments can be similar to the internal radius R of the body 2. As the propellant stream S enters the inlet 11 and subsequently the plenum chamber 17, angularvelocity is imparted to the propellant stream S (from the offset angle a of the inlet in relation to the longitudinal axis X of the body 2). The propellant stream S starts circulating about the plenum chamber 17 and about the central core 12 of the heat transfer element 20 and heater 30 therein. The swirling propellant stream S is channelled around the core 12 of the heat transfer element 20, which offers an expanse of heated surface area for heat energy to be absorbed by the propellant stream S. As more propellant is fed into the inlet 11, the swirling, turbulent motion of the propellant stream S drives and forces the propellant stream S into the curved channels 25 of the heat transfer element 20, and begins to weave its way through the network of plurality of tortuous flows 32, and through holes 28, 29 of the heat transfer element 20, all the while gathering thermal energy from the thermally charged channel walls 26 of the heat transfer element 20. The propellant stream S flows across the heat transfer element 20 to be heated. The resulting heated propellant stream H emerges at the free end 23 of the lattice structure 22 and subsequently passes out of the body 2 while being forced into the nozzle 33, squeezed through the outlet 3a, and out of the thruster 1 diffused by the diffuser 34 as a super-heated propellant stream H to generate thrust T. The body 2 sharply tapers, in the region comprising the air gap 24 and nozzle 33, from radius R to a narrow aperture at the outlet 3a.

[0144] Figure 3 is a perspective view of the thruster 1 cut transversely across longitudinal axis X through a thickness of the thruster 1. In this view, the channel walls 26 are displayed as a plurality of arcuate walls 26 of varying thickness, that radiate outwardly from the cavity 27 in which the heater 30 is housed (not shown here). In this cross-section, the curved channels 25 present a plurality of fin-shaped openings, having a narrower end near the central core 12 and opening into a wider opposing end as the opening extends through the outer peripheral portion 13 of the heat transfer element 20. Figure 3 also illustrates the holes 28 within the outer peripheral portion 13 of the heat transfer element 20.

[0145] The first end 5 of the body 2 is at least partially surrounded by the flange 4, providing here three mounting apertures 6. The flange 4 is shown in Figure 3 integrally formed with the body 2. It will however be understood that the flange 4 can be manufactured as a separate component and joined to the body 2 as a sub-process. The inlet 11 positioned tangentially to the body 2 is also illustrated in Figure 3, inclined at an angle a to the longitudinal axis X and centrally offset thereto.

[0146] The inlet 11 is defined by a connector 14 fluidly connecting a propellant stream S to the interior 19 of the body 2. The connector 14 protrudes from the exterior surface 2a of the body 2 and extends away therefrom. The connector 14 facilitates coupling of the thruster 1 with a propellant source for feeding the propellant stream S to the thruster 1. In some examples, the inlet 11 can have a radius ranging from approximately 1 mm to approximately 10mm and a surface finish ranging from approximately 20 microinches to approximately 35 microinches. It will be understood that the inlet 11 may be any other dimensions as may be considered suitable by a skilled person in the context of the present disclosure.

[0147] With further reference to Figures 4 to 6, the heat transfer element 20 is cylindrical in overall shape and comprises the central core 12 and the outer peripheral portion 13 circumferentially surrounding the core 12. In Figures 4, 5 and 6, the body 2 has been removed to better show the lattice structure 22 of the heat transfer element 20. The outer peripheral portion 13 defines the plurality of curved channels 25 and channel walls 26 therethrough, with each of the channels 25 having a plurality of spaced apart holes 28 therealong. The curved channels 25 extend from the plenum chamber 17 and terminate at the free end 23 of the lattice structure 22 to define the air gap 24. The plurality of channels 25 are separated from each other by the channel walls 26. The plurality of curved channels 25 divide the propellant stream S into the plurality of torturous flows 32 when flowing from the inlet 11 to the nozzle 33. As shown in Figures 3 and 4, the channel walls 26 are equidistantly spaced about the central core 12 and radiate outwardly from the core 12 towards the interior surface 2b of the body 2.

[0148] As shown in Figure 5, the central core 12 of the heat transfer element 20 comprises a plurality of equidistantly spaced holes 29 therethrough. On passing through these holes 29, the propellant stream S is continuously split thereby forming the plurality of flows 32, each of which can have direct access to the heater 30. The holes 28 of the plurality of curved channels 25 are fluidly connected with each other and are also fluidly connected to the plurality of holes 29 of the central core 12. The fluid connection between the holes 28 and holes 29 enables the plurality of flows 32 of the propellant stream S to interact with each other thereby causing constant mixing of the plurality of flows 32 within the propellant stream S.

[0149] In thruster 1, the heater 30 is electrically powered via cable 35. In some embodiments, the heater 30 can be an electrical heater, which can be powered by a battery or an alternative power source. The heater 30 may be a resistive heater. In some embodiments, the heater 30 transfers heat to the heat transfer element 20 via the cavity 27 by conduction and then heat is transferred to the propellant stream S by convection. In some embodiments, the heater 30 comprises sensors such as thermocouple sensors to measure the temperature of the heater 30. The heater 30 can be selected to supply between 5 Watts to 5 kW in power or can be configured to provide a variable output.

[0150] In some embodiments, the holes 29 of the central core 12 of the heat transfer element 20 are not in fluid communication with the interior of the cavity 27. The indirect heating of the propellant stream S by the heater 30 prevents direct contact between the heater 30 and propellant stream S. This arrangement can be set-up to help prevent or at least reduce oxidation that can reduce the lifetime of the heater 30.

[0151] In other embodiments, the holes 29 are in fluid communication with the interior of the cavity 27, which can increase the rate of thermal energy transferred to the propellant S.

[0152] The complex geometry of the heat transfer element 20 is defined by each of the plurality of curved channels 25 and the plurality of holes 28, 29 that form the lattice structure 22. The complex geometry induces turbulent mixing of the propellant stream S as the flows 32 travel through the heat transfer element 20, which also enables convective heat transfer of thermal energy to the heat transfer element 20. Furthermore, the plurality of channels 25 increase a surface area to volume ratio of the thruster 1, as compared to traditional heat exchangers, where propellant typically traverses through only one or two channels therethrough (these channels may spiral but do not interact or intersect one another and as such do not cause any additional mixing of the flows). The increased surface area to volume ratio can improve the convective heat transfer to the propellant S resulting in a hotter propellant stream H when using a heat exchange element 20 of reduced dimensions.

[0153] The continuous smooth internal surfaces of each of the plurality of curbed channels 25 can reduce the magnitude of any pressure drops (differentials) in the propellant stream S and can also prevent (or at least reduce) the formation of dead zones in the flows 32 of the propellant stream S.

[0154] A large reduction in pressure in the propellant stream S is referred to as a pressure drop. Pressure drops are detrimental because they can cause instability in the flows 32 ofT1 the propellant stream S for example a stuttering or series of fluctuations. Dead zones are areas within the heat exchanger 20 that remain stagnant or trapped and where flows 32 are notably lesser compared to the remainder of the heat exchanger 20, these dead zones can often be seen around voids or corners in traditional heat exchangers. Dead zones can lead to significant fouling as well as reduced heat transfer efficiency and are therefore detrimental to the overall heat transfer efficiency. The term "fouling" refers to damage to heat transfer surfaces or build-up of unwanted material. This effect is of lesser relevance to spacecraft heat exchangers and thrusters but can have greater significance for terrestrial heat exchangers that operate for extended durations, particularly in dirty / dusty environments.

[0155] The increased surface area to volume ratio of the heat transfer element 20 has been demonstrated to maintain a steady performance while facilitating a reduction in the overall length LB and a reduction in the internal radius R of the body 2, thereby enabling minimalist dimensions and miniaturisation of the thruster 1.

[0156] In some embodiments, the features of: (i) increased surface area to volume ratio of the heat transfer element 20; (ii) the complex geometry of the lattice structure 22, which induces turbulent mixing of the propellant stream S; (iii) the continuous, smooth surfaces along each of the plurality of curved channels 25; and (iv) the miniaturisation of the thruster 1, either taken alone or in combinations thereof, provide a thruster with an increased thermal efficiency and an improved performance when compared to conventional thrusters.

[0157] In some embodiments, the features of: (i) a central location of the heater 30 within the heat transfer element 20; (ii) a low power operation of the heater 30; and (iii) reduced temperatures at the exterior surface 2a of the body 2, either taken alone and in combinations thereof, provide opportunities for further mass reduction of the thruster 1. These mass reductions can be realised by reducing the need for radiation shielding. In some embodiments it can be possible to omit radiation shielding from within or surrounding the thruster 1. Alternatively, in some embodiments it may be possible to at least reduce the radiation shielding material.

[0158] The body 2 of the thruster 1 can be integrally formed with the heat transfer element 20. In some embodiments, the body 2 and heat transfer element 20 can be manufactured contemporaneously as a monolithic piece.

[0159] In some embodiment, the body 2 can also be integrally formed with the flange 4 and the connector 14. In some embodiments, the body 2 can also be integrally formed with thenozzle 33 and diffuser 34. In some embodiments, the thruster 1 is integrally formed as a single, monolithic structure.

[0160] The thruster 1 can be manufactured using additive manufacturing. The thruster 1 can be manufactured using 3D printing from a suitable material.

[0161] In some embodiments, the lattice structure 22 has a warped structure, mathematically controlled such that a thickness t of the channel walls 26 decreases with increasing distance from the cavity 27 in which the heater 30 is located (shown in Figure 5). A thickness tz of the channel walls 26 in proximity to the bore 27 is largest, and a thickness ti is smallest adjacent to the exterior surface 2a of the body 2. The channel wall 26 thickness t can range from approximately 0.05 mm to approximately 5 mm. In one embodiment, the thickness t of the channel wall 26 decreases from around 0.8 mm adjacent the cavity 27 to approximately 0.6 mm adjacent the exterior surface 2a of the body 2. This increases the conductive surface area of the channel walls 26 near the cavity 27, and heater 30 located therein, to transfer thermal energy more effectively from the heater 30 throughout the heat transfer element 20.

[0162] In one embodiment, the lattice structure 22 comprises an array formed from a plurality of unit cells (471, 473) that are tessellated to form a Triply Periodic Minimal Surface (TPMS) lattice structure. This TPMS form of lattice provides a labyrinthine network of curved channels 25 through the heat transfer element 20. TMPS lattice structure can be generated in multiple formats, two such formats being Diamond and Gyroid.

[0163] With reference to Figure 21A a unit cell 471 is illustrated that can be tessellated to form a lattice structure that is a Gyroid TPMS lattice structure, each unit cell 471 of the array having a gyroid topological geometry. With reference to Figure 21 B there is illustrated a unit cell 473 that can be tessellated to form a lattice structure that is a Diamond TPMS lattice structure, each unit cell 473 of the array having a diamond topological geometry.

[0164] Diamond TPMS lattice structure can be mathematically defined with the following equation: sin (x) sin (y) sin (z) + sin(x)cos(y)cos(z) + cos(x)sin(y)cos(z) + cos(x)cos(y)sin(z) = 0 Figure 6 illustrates a diamond TPMS lattice structure for a heat transfer element 20 as described above.

[0165] Gyroid TPMS lattice structure can be mathematically defined with the following equation:cos(x)sin(y) + cos(y)sin(z) + cos (z) sin (x) = 0

[0166] In one example, TPMS diamond lattice structure cells 471, 473 were selected to range in dimension from approximately 1 x 1 x 1 mm to approximately 20 x 20 x 20 mm. In one embodiment, the cell dimensions approximately 6 x 6 x 6 mm were used. The cells 471, 473 were then repeated and / or tessellated to define the volume of the lattice structure 22.

[0167] Computational Fluid Dynamic (CFD) simulations comparing performance of a diamond TPMS lattice structure and a gyroid TPMS lattice structure (confined within the same body 2) show that the diamond TPMS lattice structure has an improved thermal performance. The computational Fluid Dynamic (CFD) simulations consisted in undertaking a heat transfer analysis to compare the convective heat transfer performance of the gyroid and the diamond TPMS lattice structures. The CFD simulations demonstrated that the diamond TPMS lattice structure had a 3% increase in the heat transfer rate from the heater transfer element 20 to the propellant stream S compared to the gyroid TPMS lattice structure. In other words, the simulations indicated that the diamond TPMS lattice structure enhanced the heat transfer efficiency of the lattice structure 22.

[0168] Example heat maps of some of these computational simulations and heat transfer analysis are shown in Figures 9A and 9B. At the inlet 11, the propellant stream S enters and is introduced into the plenum chamber 17. The coldest temperature of about 301 K (approximately 28°C) is in the plenum chamber 17, which represents a coldest temperature of the propellant stream S and the body 2 of the thruster 1 in this region of the thruster 1. The propellant stream S and body 2 temperatures can then be seen to have variations as the propellant stream S circulates and traverses the lattice structure 22. The propellant stream S and body 2 temperatures combined are the hottest within regions 22A of the lattice structure 22 that surround the heater (not shown but which would be positioned centrally in this embodiment). The propellant stream S and body 2 temperatures combined are also the hottest within the region 22B surrounding the outlet 3a. The propellant stream S begins to swirl around the plenum chamber 17 as it is forced into the lattice structure 22 of the heat transfer element 20. With additional reference to Figure 10, as the propellant stream S interacts with the lattice structure 22, the propellant stream S is split repeatedly into the plurality of tortuous flows 32. Each flow 32 is forced and drawn through the holes 28 and 29 while continuously intersecting and weaving through adjacent flows 32 and being repeatedlysplit, all the while drawing thermal energy from the surface 20a of the lattice structure 22 and increasing in temperature. The temperature scale is higher in the embodiment wherein the lattice structure 22 comprises a gyroid TPMS lattice structure compared to the diamond TPMS lattice structure. This is because when using a gyroid TPMS lattice structure, heat transfer from the heat transfer element 20 to the propellant stream S is not as effective as when using a diamond TPMS lattice structure and while the propellant stream S reaches lower temperatures, the body 2 gets hotter in the embodiment using the gyroid TPMS lattice structure than in the embodiment using a diamond TPMS lattice structure. The simulations illustrate a steady state temperature of the propellant stream S and body 2 of the thruster 1. Much higher outlet temperatures of up to about three-fold can be possible however are not steady state.

[0169]

[0170] The thruster 1 has two operative phases: phase (1) is a warm-up phase with no propellant in the system; and phase (2) is a productive or operative phase with propellant flowing over the heated heat transfer element 20 to generate thrust T.

[0171] The thruster 1 is typically run for about 10 minutes with no propellant (warm up phase); this gets the internal temperatures in the thruster into a range between 500°C and 600°C. Once up to temperature, the propellant stream S is introduced (productive phase), which will instantly cool the internal temperatures of the thruster. After about 200 seconds (3 to 4 minutes), the thruster 1 will reach a "steady state" wherein the internal temperature has stabilised for the given flow rate of the propellant S and the heat energy inputted by the heater 30. In practice, this steady state is seldom reached and as such the steady state output temperatures are more often used for comparative purposes.

[0172] Higher performance (i.e. higher output temperatures) can be possible compared to steady state results. The productive phase of the thruster 1 is approximately 30 seconds. This short burst of running propellant through the thruster 1 does not cool the heat transfer element 20 sufficiently to reach steady state, and results in significantly hotter outlet temperatures (i.e. temperatures at the outlet 3a). Experiments have shown that outlet temperatures can be increased three-fold when the thruster is heated for approximately 10 minutes before allowing propellant to enter.

[0173] Figure 10 is a thermal simulation of a propellant stream S in the thruster 1 of Figure 1, illustrating a rapid increase in temperature from about 300K to about 600K as the propellantstream S is split into the plurality of flows 32 and the propellant stream S advances from the inlet 11 to the outlet 3a of the thruster. The simulation model was configured to use the heat transfer element 20 configured as a Diamond TPMS lattice structure 22, as shown in Figure 6. The complexity of the plurality of flows 32 is illustrated in Figure 10, as the tangential inlet 11 introduces the propellant stream at an angle a of about 90 degrees to the longitudinal axis X. The plenum chamber 17 immediately imparts angular velocity to the propellant stream S, which begins to circulate about the core 13 and on being forced into the lattice structure 22 of the heat transfer element 20 begins to split and weave through the holes 28 and 29 of the lattice 22. In this manner, the plurality of flows 32 is intricately interwoven, and mixed while being heated creating a turbulent flow and an even thermal gradient across the lattice structure 22.

[0174] Another embodiment of a thruster 101, configured for testing, is illustrated in Figure 7. The thruster 101 provides a body 102, an outlet 103a, a first end 105, a second end 108, a plenum chamber 117, a heat transfer element 120 that defines a cavity 127 for receiving a heater (not shown) therein. The cavity 127 can be a bore. The thruster 101 has an inlet (not shown) for introducing a propellant stream S into the body 102. The thruster 1 further defines a nozzle 133 at the outlet 103a and an air gap 124 defined between the outlet 103a and a free end 123 of the heat transfer element 120.

[0175] The thruster 101 further comprises a measurement port 115 arranged tangentially to the body 102. The measurement port 115 is positioned at the second end 108 of the body 102 proximate to nozzle 133. The measurement port 115 is configured to house a thermocouple (for measuring pressure and temperature). Once inserted, the thermocouple prevents any ingress or egress of propellant from the measurement port 115. The measurement port 115 is used predominantly for data acquisition and receives data from the propellant immediately as it exits the heat transfer element 120, and immediately before the heated propellant stream H is pushed through the outlet 103a. The thruster 101 further comprises a diffuser 134. The thruster 101 and all its components and parts can have characteristics identical or otherwise similar to the ones herein described in relation to thruster 1. The heat transfer element 120 may have a lattice structure and have characteristics identical or similar to the ones described in relation to thruster 1 and heat transfer element 20.

[0176] Figure 8 illustrates another embodiment of a thruster 201, which provides a body 202, a plenum chamber 217, and a heat transfer element 220 having a reinforcing member 231 extending from a cavity 227. The cavity 227 is defined within the heat transfer element 220 for receiving a heater (not shown) therein. The reinforcing member 231 terminates proximal to a free end 223 of the heat transfer element 220. A thickness of the reinforcing member 231 tapers along a length thereof and terminates with a conical tip. The reinforcing member 231 has a conical shape that tapers in a direction towards a nozzle 233 of the body 202. The body 202, plenum chamber 217, heat transfer element 220 and reinforcing member 231 are coaxially aligned with the longitudinal axis X of the thruster 201. The thruster 201 has an inlet (not shown) for introducing a propellant stream S into the body 202. The thruster 201 also has an outlet 203a for exiting the heat propellant stream H and a diffuser 234. The thruster 201 also comprises an air gap 224 defined between the outlet 203a and a free end 223 of the heat transfer element 120. The thruster 201 and all its components and parts may have characteristics identical or otherwise similar to the ones herein described in relation to thruster 1. The heat transfer element 220 may have a lattice structure and have characteristics identical or similar to the ones described in relation to thruster 1 and heat transfer element 20.

[0177] Figures 11A and 11 B illustrate another embodiment of the present invention. The flow of the propellant stream S and heated propellant H for this embodiment is indicated in the enlarged view of Figure 11B for clarity. With references to Figure 11A and 11 B, there is provided a thruster 301 for generating thrust T. The thruster 301 comprises a body 302 having an inlet 311 and an outlet 303a, the inlet 311 introducing a propellant stream S into the body 302 and the outlet 303a providing a nozzle 333 for exiting a heated propellant stream H from an interior of the body 302. The inlet 311 is arranged tangentially to the body 302. The outlet 303a defines the nozzle 333 and then transitions into and expands to define a diffuser 334 through which the heated propellant H is ejected to generate thrust T.

[0178] With further reference to Figures 13 and 14, the thruster 301 comprises a heat transfer element 320 located within the interior 319 of the body 302 and configured to interact with the propellant stream S. The thruster 301 further comprises a heater 330 positioned to transfer heat to the heat transfer element 320. The heater 330 is positioned within the heat transfer element 320 with a gap 307 separating the heater 330 from the heat transfer element 320. In this embodiment, heat is radiatively transferred between the heater330 and the heat transfer element 320, and the propellant stream S is radiatively heated during use of the thruster. The heater 330 does not contact the heat transfer element 320. The heater 330 is centrally located within the heat transfer element 320. The heater 330 is connected to a power source (not shown) via connection cable 335. The heat transfer element 320 comprises a lattice structure 322, as described herein in relation to thrusters 1, 101 and 201. The lattice structure 322 and heat transfer element 320 define a cavity 327 for receiving the heater 330 therein. The cavity 327 is centrally located in the heat transfer element 320. The cavity 327 is a blind bore.

[0179] The heater 330 is a resistive heater. The heater 330 comprises a material that has a maximum operating temperature above that of traditional resistive heaters, for example that of traditional metal sheathed cartridge heaters. The heater 330 can have a maximum operating temperature of approximately or exceeding 1000°C to approximately 1100°C. In some embodiment, the heater 330 can heat the heat transfer element 320 to a temperature sufficient to attain a temperature of the propellant stream S of up to approximately 800°C, when introduced into the body 302 via the inlet 311. Where the propellant stream S comprises nitrous oxide, the heater 330 can heat the heat transfer element 320 to a temperature sufficient to initiate the decomposition of the nitrous oxide. To achieve these temperatures, the heater 330 can comprise a ceramic material.

[0180] In the present specification, the term "ceramic heater" will be used to refer to a heater that comprises a ceramic material, and may be entirely made of ceramic material, or may partially be made of ceramic material, wherein the heater comprises enough ceramic material to enable heating of the heat transfer element by radiation to the desired temperature in accordance with embodiments of the present invention.

[0181] The ceramic heater 330 can be manufactured from known methods and materials and can be manufactured from a ceramic film with an embedded platinum conductor path that enables reliable operation at 1100°C. For example, a ceramic material manufactured by The Rauschert Group company in Germany may be used. The use of a ceramic material for the heater 330 enables operation of the heater 330 wherein the heater 330 can reach a temperature of at least approximately 1000°C, whereby the heat transfer element 320 is heated by radiation and sufficiently heated such that the propellant stream S can attain a temperature of up to approximately 800°C, suitable for nitrous oxide decomposition, andwhile consuming a substantially minimised amount of energy. For example, while consuming less than 250 Watts of power.

[0182] The decomposition of N2O is an exothermic reaction and as such by heating the heat transfer element to a temperature sufficient to initiate the N2O decomposition, the decomposition produces enough energy and heat to be self-sustaining thereby reducing the energy demand to heat the heater 330 and maintain the reaction. In this embodiment, the thruster 301 can function thermally efficiently to produce thrust T continuously with a reduced need for energy input as long as the flow of the N2O propellant stream S is continuous.

[0183] Ceramic materials are highly brittle and inserting a ceramic heater 330 into a tight tolerance bore can cause ceramic heaters to shatter or break when they undergo temperature cycling due to thermal expansion inside the metal bore. Positioning the ceramic heater 330 within the heat transfer element 320 with the gap 307 separating the ceramic heater 330 from the heat transfer element 320 can assist in overcoming manufacturing difficulties, structural difficulties, and usage difficulties including providing space for insertion of the heater 330 within the heat transfer element 320 and reducing the risk of shatter and breakage of the heater associated with thermal expansion when submitted to temperature cycling.

[0184] Figure 12 shows a schematic representation of a ceramic heater 330 as the one manufactured by The Rauschert Group company in Germany and that may be used for applications in accordance with embodiments of the present invention wherein the heater is positioned to radiatively heat the heat transfer element, the heater being positioned within the heat transfer element with a gap in between the heater and the heat transfer element. The ceramic heater 330 comprises a first part 330A for encapsulating cables, electronics and other electric elements, and a second part 330B that is the heated element. Heat distribution within the heated element 330B is schematically indicated with a central part 330C that can reach highest temperatures.

[0185] Returning to Figure 11A, the body 302 extends from a first end 305 to a second end 308 and has a central longitudinal axis X. The first end 305 of the body 302 defines an annular plenum chamber 317 for receiving the propellant stream S and forcing the propellant stream S into the heat transfer element 320. In this embodiment, the inlet 311 is located at the second end 308 adjacent the nozzle 333. This is advantageous because theregion near the nozzle 333 is likely to attain the highest temperatures within the thruster 301. The inlet 311 leads to an inlet chamber 362 surrounding the nozzle 333. The location of the inlet 311 at the second end 308 further enables a maximum heat rejection from a wall 364 of the inlet chamber 362 to the cold propellant stream S on introduction through the inlet 311.

[0186] The body 302 further comprises a chamber temperature measurement port 315 that is arranged tangentially to the body 302.

[0187] The body 302 comprises a plurality of cooling channels 352 configured to provide passage for the propellant stream S from the inlet 311 and plenum chamber 317 to the heat transfer element 320.

[0188] For example, the thruster 301 can comprise at least two cooling channels 352. In the embodiment shown in Figure 13, the thruster 301 comprises ten cooling channels 352 equidistantly spaced around the thruster 301. In other embodiments, the thruster 301 can comprise more than ten cooling channels 352.

[0189] The cooling channels 352 are cavities embedded within an outer wall 354 of the thruster 301 and configured to reduce the temperature of the thruster outer wall 354. This provides benefits, including but not limited to reducing the likelihood of material failure during use, and reducing thermal radiation losses.

[0190] In the embodiment illustrated in Figures 11 A, 11 B, 13, and 14, the cooling channels 352 are arranged spaced apart, circumferentially in the outer wall 354 of the thruster 301.

[0191] The cold propellant stream S is introduced into the cooling channels 352 before being brough into contact with the heat transfer element 320, such that the flow of the cold propellant stream S through the cooling channels 352 results in a cooling of the outer wall 354 as the propellant stream S absorbs heat (thermal energy) from the outer wall 354. The temperature of the outer wall 354 is thus decreased while the temperature of the propellant stream S is increased. These properties of the cooling channels 352 make the cooling channels 352 regenerative.

[0192] In addition to preventing material failure, the reduction in temperature of the outer wall 354 improves the efficiency of the thruster 301 as lower outer wall temperatures emit less heat, thereby retaining more heat (thermal energy) within the thruster 301. In this manner, the propellant stream S is pre-heated in the cooling channels 352 before beingforced into the heat transfer element 320. Higher propellant temperatures H can thus be achieved, which improves specific impulse (propulsion efficiency).

[0193] The cooling channels 352 are termed "regenerative" because they use the propellant streams S to cause the cooling of the thruster outer wall 354, which simultaneously results in an increase of the propellant stream temperature as the propellant stream absorbs the heat from the solid thruster outer wall 354.

[0194] Once the propellant stream S passes through the cooling channels 352, it is uniformly distributed into the plenum chamber 317 via a ring of injector orifices (not shown), each injector orifice having a cross-sectional area substantially equal to a cross-sectional area of a respective cooling channel 352. This assists in enforcing uniform flow distribution of propellant S into the heat transfer element 322 and subsequently a uniform radial temperature distribution, preventing the formation of hot spots.

[0195] Once the propellant stream S flows around the inlet chamber 362, the propellant stream S enters the cooling channels 352 in the outer wall 354, before being redirected 180 degrees to travel through the heat transfer element 320. The propellant stream S is then heated in the lattice structure of the heat transfer element 320 before being expelled through the nozzle 333 to generate thrust T.

[0196] The heat transfer element 320 extends from a region adjacent the first end 305 towards the second end 308 of the body 302 to a region adjacent the outlet 303a, guiding the propellant stream S through the body 302. Adjacent the outlet 303a of the body 302, the heat transfer element 320 terminates to define a free end 323. An air gap 324 is formed between the free end 323 and the outlet 303a. The air gap 324 is peripherally bounded by the nozzle 333.

[0197] In this embodiment, the body 302 has a substantially cylindrical shape and conically tapers in the region comprising the air gap 324 and nozzle 333 to a narrow outlet 303a.

[0198] In some embodiment, the gap 307 between an outer surface of the heater 330 and an outer surface of the heat transfer element 320 may have a length within a range between approximately 0.5mm and approximately 3mm. It will be understood that the length of the gap 307 is not limited to this range however is limited by structural limitations to ensure that performance of the heater 330 and thruster 301 as a whole is not affected. Structural limitations include the dimensions of the thruster, and dimensions of all its elements including the body, the heater and the heat exchange element.

[0199] In the embodiment illustrated in Figure 11 A, the heater 330 is secured along the thruster 301 longitudinal axis X via a mounting plate (not shown). The heater 330 is encased by the cavity 327 but does not contact the inner surface of the cavity 327 to ensure that only radiation heat transfer occurs with the heat transfer element 320.

[0200] The thruster 301 can further comprise a radiation shield 356. The radiation shield 356 surrounds the body 302 of the thruster 301 as illustrated in Figure 11 A. The radiation shield 356 can be separated from the outer surface of the thruster by a shield gap G. In one embodiment, the gap G separating the radiation shield 356 and outer surface of the thruster 301 is equal to the air gap 307 between the inner surface of the cavity 327 and the outer surface of the heater 330. In some embodiment, the radiation shield 356 is provided in the form of a thermally isolated cylinder. The radiation shield 356 enables reduction of radiation losses by enabling reflection of the thermal radiation emitted from the thruster 302 back onto itself.

[0201] Ceramic spacers 358 can be used between the thruster 301, the radiation shield 356 and mounting bolts 360 to minimise thermal losses via conductive heat transfer.

[0202] With reference to Figures 13 and 14, cross-sectional views of the thruster 301 are illustrated to comprise the heat exchange element 320 in the form of a lattice structure 322 as described for thrusters 1, 101, 201 and comprising the plurality of cooling channels 352.

[0203] Once the propellant stream S passes through the cooling channels 352, it is uniformly distributed into the plenum chamber 317 via a ring of injector orifices (not shown) with the same cross-sectional area as the cooling channels 352. This is intended to enforce uniform flow distribution into the lattice structure and subsequently a uniform radial temperature distribution, preventing the formation of hot spots.

[0204] The heat exchange element 320 has a lattice structure 322 similar to those described in relation to thrusters 1, 101, 201. In one embodiment, the lattice structure of the heat exchange element 320 has a diamond TPMS structure defined by a plurality of curved channels 325, channel walls 326, and holes 328, 329. In Figure 13, the cooling channels 352 are arranged circumferentially around the outer surface of the heat transfer element 320 and have a rectangular cross-section embedded within the outer wall 354 of the thruster body 302.

[0205] The cross-sectional area of the regenerative cooling channels 352 can be reduced in cross-section to further increase a velocity of the propellant stream S. This can subsequentlyincrease the convective heat transfer from the outer wall 354 to the propellant S. In one embodiment, the cooling channels 352 can have a rectangular cross-section with dimensions for an area within a range from approximately 0.5mm2to approximately 5 mm2. The dimensions of the thruster 301, lattice the curved channels 325 of the lattice structure 322, the bore 327, and other components of the thruster 302 will generally be determined based on manufacturing limitations and intended use of the thruster 302.

[0206] It will be understood that cross-sectional shapes of the cooling channels 352 other than rectangular may be contemplated. For example, the cooling channels 352 may have a circular cross-sectional shape, or an oval or any other rounded cross-sectional shape. In another example, the cooling channels 352 may have a square cross-sectional shape. In addition, the cross-sectional dimensions of the cooling channels provided herein are merely exemplary of one embodiment of the invention and may be varied within the scope of the present invention to enable an improved thruster having one or more of a reduced size, weight, energy consumption, more efficient thrust production, increased thermal efficiency and an improved performance when compared to conventional thrusters.

[0207] Material emissivity is an important element considered in designing the thruster 1, 101, 201, 301 in accordance with embodiments of the present invention. Emissivity is a relative measure of an objects ability to emit infrared radiation compared to a blackbody, which is a perfect emitter of radiation and thus has an emissivity of 1. Dark, rough and oxidised surfaces are much more effective at emitting and absorbing radiation and thus have high emissivity values. Light, clean and polished materials in contrast have a very low emissivity as they tend to reflect infrared radiation. Thus, surface emissivity can be controlled through a selection of materials and post processing to either absorb the radiation emitted from the heater or reflect radiation that is emitted from the body.

[0208] In some embodiment, the radiation shield 356 comprises an aluminium material. In some embodiments, the radiation shield 356 comprises a steel material. In some embodiment, the radiation shield 356 is constructed from a highly polished aluminium material. In some other embodiments, the radiation shield 356 is constructed from a highly polished steel material. In one example, the radiation shield 356 comprises a highly polished steel material having a surface emissivity of about 0.075. This type of radiation shield can reflect more than 90% of incoming infrared radiation.

[0209] The thruster 1, 101, 201, 301 can reduce fuel requirements for producing an equivalent amount of thrust (when compared to know thrusters), thereby extending the duration for which an object, for example, a satellite, can stay in orbit. The fuel saved as a result of this reduction in fuel requirements can be used to undertake manoeuvres that are not currently possible with a traditional thruster, or to extend range.

[0210] The compact dimensions of the heat transfer element 20, 120, 220, 320 allow for a highly miniaturised thruster. In addition, the miniaturised thruster facilitates a reduction in mass of the thruster, thereby further increasing the performance of the thruster. This makes the thruster particularly well suited for providing propulsion for objects such as CubeSats and small satellites which have inherently strict size and mass constraints.

[0211] The modes of heat transfer for a thruster in a vacuum environment is typically via conduction and radiation. Conduction is governed by Fourier's Law as seen in the following equation:In this equation, k is the thermal conductivity of the conducting material, A is the area available for heat transfer, L is the length of the heat conducting path, and AT is the temperature difference between the heat source and interface. Conduction losses only occur where a solid material provides a path for thermal energy to be transferred away from the heater and heat exchanger to the thruster interface.

[0212] For high temperature thrusters as described in embodiments of the present invention, ceramic thermal spacers such as ceramic spacers 358 are typically utilised to minimise conduction losses. Ceramic thermal spacers such as ceramic spacers 358 can withstand very high operating temperatures, have a low a thermal conductivity (typically between approximately 0.1 - 2 W / mK), and are much longer in length than a washer to minimise conduction losses. Assuming three ceramic standoffs with a thermal conductivity of approximately 1 W / mK suitable for an M3 bolt are used at the thruster interface (an inner and outer diameter of approximately 3.2 and approximately 7 mm respectively), with the ceramic spacer being approximately 30 mm long and exhibiting a thermal conductivity of approximately 1 W / mK, the thermal losses via conduction can be estimated for the worstcase scenario from the above Fourier's Law equation. For a temperature difference of WOOK, the conduction losses are approximately 1 Watt per spacer, or 3 Watts in total, whichaccounts for approximately 1.5 % in thermal energy loss from a 205-Watt ceramic heater. It is considered that losses via conduction are consequently negligible with the use of ceramic spacers and have not been considered for the analytical thermal analysis as described herein.

[0213] Radiation heat transfer is by far the primary mode of thermal loss for thrusters that operate at high temperature. The rate of radiation heat loss is governed by the Stefan- Boltzmann Law in accordance with the following equation:Q = a s A (Ts4- Ta4)

[0214] In the above equation, o is the Stefan-Boltzmann constant (equal to 5.67 x 10’8J / sm2K4), e is the material emissivity, A is the surface area, Tsis the surface temperature and Tais the ambient temperature.

[0215] The rate of heat transfer via radiation is proportional to the fourth square of the surface temperature, and thus thermal design to minimise radiation losses becomes critical for thrusters that operate at high temperature. This is particularly applicable to the thrusters as described in accordance with the embodiments of the present invention, which aim to attain very high thruster temperatures from a resistive heating element, where a primary design objective is to reduce (and where possible minimise) the electrical energy required to power the heater and heat the propellant stream, and thus subsequently minimise thermal radiation losses.

[0216] Conduction and convection thermal losses are in some embodiment considered negligible, and the inventor has carried out an analytical steady state thermal model for a ceramic heater 330 and thruster 301 as described in embodiments of the present invention, based on known radiation heat transfer principles. The model can be used to estimate an attainable temperature of the thruster 301 as a function of the temperature of the heater 330 and the impact of including a radiation shield. The purpose of the model has been to determine in one particular embodiment whether the thruster 301 can attain temperatures of approximately or exceeding 800°C to initiate thermal decomposition of nitrous oxide (before the propellant is admitted). The model simplifies the overall geometry of the thruster 301 and its components, with the heater 330, thruster 301 and radiation shield 356 being concentric cylinders comprised of diffuse and grey surfaces. Figure 15 illustrates the thermal model analogy. A space between the cylinders is assumed to be evacuated, since the ambient environment is considered to be vacuum. The analysis is one dimensional, the bodies are made of thin conductive materials, and internal conduction resistance isneglected. Two cases were considered for the analytical model - one without a radiation shield 356 and one with a radiation shield 356.

[0217] The thermal network analogy as illustrated in Figure 15 is used to describe net radiation heat transfer between the inner and outer surfaces of each cylinder. For such a problem, the heat transfer is treated similar to an electric current, the temperature difference is analogous to the voltage, and a thermal resistance is introduced. The following mathematical equation is followed to solve for the net radiative heat transfer rate:

[0218] A thermal radiation network as illustrated in Figure 16 can then be constructed with thermal resistances calculated according to well-known radiation theory between two body enclosures following the equation as follows with reference to Figure 16:Where:The variable F is the shape factor - a non dimensionless variable used to quantify the percentage of radiation that leaves one cylinder and is incident upon another, which is entirely dependent on the body geometric configuration. For the problem of concentric cylinders, the shape factors are all assumed to be equal to unity, since the cylinders are fully enclosed within each other. T1 is the heater 330 temperature, T2 is the thruster 301 solid body 302 temperature, T3 is the radiation shield 356 temperature and T4 is the ambient temperature.

[0219] To determine the net radiative heat transfer Qnet from the heater 330 to the environment and the thruster 301 solid body 302 temperature T2 once the net heat transfer rate is determined, the following equations are used:

[0220] The inventor has then made thermal design considerations. To maximise the solid body 302 temperature (T2) of the thruster 301, the thermal resistance (R12) between the heater 330 and thruster 301 must be minimised. This can be achieved by maximising the surface emissivity (si) of the heater and the surface emissivity (82) of the internal bore of the thruster 301 which surrounds the heater 330. For this reason, the ceramic heater 330 was coated in a high emissivity black paint which has an average emissivity of 0.9. Furthermore, the cavity 327 will be heated in atmosphere before testing to oxidise the surface of a GRCop- 42 material used in this example for the body 302 and cavity 327, which is assumed to have a surface emissivity of 0. 76 equal to that of oxidised copper.

[0221] To further increase the body 302 surface temperature, the net heat transfer from the thruster 301 to the environment needs to be minimised, which can be done by maximising the thermal resistance R23 between the thruster 301 and radiation shield 356 and the thermal resistance R34 between the radiation shield 356 and the ambient environment. This can be achieved by minimising the surface emissivity (83) of the external thruster 301 surface, and the surface emissivity (84 and 85) of respectively, the internal surface of the radiation shield 356 and external surface of the radiation shield 356. The emissivity of the external surface of the thruster 301 cannot be controlled easily and is assumed to be equal to oxidised copper (0. 76), the same as the internal surface. The radiation shield 356 can be highly polished steel with a surface emissivity of 0.1 in order to reflect as much radiation as possible back onto the thruster 301 while also reducing the heat transfer to the ambient.

[0222] A python software code was developed to investigate how the body 302 temperature would vary as a function of the heater 330 temperature. The primary purpose was to undertake the analysis with and without a radiation shield 356 (by emitting the thermal resistance R23 between the thruster 301 and radiation shield 356), to determine the necessity of a radiation shield 356 and ensure that body 302 temperatures above 800 °C could be achieved at a heater 330 temperature that is below its maximum operating temperature of approximately 1000°C to 1100°C. The steady-state body 302 temperature as a function ofthe heater 330 temperature is illustrated in Figure 17. The results from the analytical thermal model highlight the importance of a radiation shield 356 to reduce the rate of heat transfer to the environment. The results show that the minimum desired thruster 301 solid body 302 temperature of approximately 800°C using the heater 330 cannot be achieved without a radiation shield 356. At the maximum operating temperature of the heater 330 (approximately 1000°C to 1100 °C), the expected steady state thruster 301 temperature is approximately 676 °C, which is below the temperatures required for thermal decomposition of nitrous oxide. With the addition of a radiation shield 356, it can be seen that the minimum desired thruster 301 solid body 302 temperature can be attained at a heater 330 temperature of 847 °C, which is below the maximum operating temperature of the heater 330. At the maximum heater 330 operating temperature, a steady state solid body 302 temperature of 1043 °C was calculated. Subsequently, the analytical thermal model indicates radiative heating using a ceramic heater 330 can enable increasing the thruster 301 temperature to above 800°C with inclusion of a radiation shield 356.

[0223] A size of the thruster 1, 101, 201, 301 is to some extent dictated by the size of the heater 30, 130, 230, 330. For example, with reference to the embodiment described in relation to thruster 301, the ceramic heater 330 can have a length of around 90mm to 100mm and a diameter of around 10mm to 15mm. It is further envisaged that the ceramic heater 330 can have a length within a range between approximately 10mm and approximately 90mm up to approximately 100mm, and a diameter within a range between approximately 3mm and approximately 15mm. In some embodiments with the radiation shield 356, the thruster 301 can have an overall diameter of around 30mm to 35mm and a length within a range that accommodates for the length of the ceramic heater 330.

[0224] For a given heater power, it is known that minimising the mass flow rate through the thruster 301 results in higher temperatures of the propellant H exiting the heat exchange element 320 according to the following equation:Q = mc,,&T

[0225] An estimation of the propellant stream S mass flow rate that can be sustained at steady state can then be acquired by rearranging the latter equation and including a thermal efficiency parameter (r|t) as in the following equation:This then considers the thermal losses of the system, as not all of the energy used in the resistive heater 330 is transferred to the propellant stream S.

[0226] As such, r|twas assumed to be equal to around 0.65, implying that 35% of the heat generated by the resistive heater 330 is not transferred to the thruster 301 and is lost to the surroundings. Using the thermal properties of nitrous oxide and imposing an 800°C change in temperature across the heat transfer element 320, wherein the propellant stream S is nitrous oxide, a nitrous oxide mass flow rate of 0.19 g / s should be maintainable according to the above latter equation. Therefore, to attain the desired 800°C increase of the propellant stream S nitrous oxide across the heat transfer element 320 to initiate nitrous oxide decomposition, the mass flow rate should be below 0.19 g / s.

[0227] Low propellant operating pressures within the thruster 301 and body 302 are also desirable for nitrous oxide decomposition. According to Le Chatelier's principle, when the gaseous nitrous oxide pressure is reduced, the reaction equilibrium, shifts toward the reactants to produce more gaseous molecules and increase the pressure to restore equilibrium. The nozzle 333 throat diameter may be reduced to reduce the mass flow rate as much as possible for a given propellant operating pressure. At the intended operational conditions for testing (0.05 < m < 0.1 g / s) the thrust produced can be on the order of tens of millinewtons according to isentropic gas dynamic theory.

[0228] In any of the embodiments described herein, the body 2, 102, 202, 302 with the heat transfer element 20, 120, 220, 320 can be manufactured as a single monolithic piece using additive manufacturing. This is advantageous in that a lack of edges / joints within the thruster 1, 101, 201, 301 reduces, if not eliminates, areas where mechanical stresses can accumulate. In addition, when integrally formed, the geometry of the heat transfer element 20, 120, 220, 320 and transitions from the plenum chamber 17, 117, 217, 317 to the heat transfer element 20, 120, 220, 320, and the transition from the heat transfer element 20, 120, 220, 320 to the nozzle 33, 133, 233, 333 and to the diffuser 34, 134, 234, 334 can be very smooth and therefore present minimal impedance to the ongoing flow of the heated propellant stream H. The body 2, 102, 202, 302 can be manufactured from any one of steel, aluminium, nickel, and their alloys. In some embodiment, the body 2, 102, 202, 302 and heat transfer element 20, 120, 220, 320 can be manufactured using GrCop-42- a copper alloy.

[0229] GRCop-42 is a copper / chromium / niobium alloy. The alloy was developed to additively manufacture parts in need of high-strength dispersion and high conductivity. Itretains strength at high temperature, due to the use of chromium and niobium in the alloy. GRCop-42 also has excellent creep resistance, and a low cycle fatigue life. The aforementioned properties are particularly valuable for rocket engine components such as fuel injector faces and combustion chamber linings with regenerative cooling.

[0230] GRCop-42 resistance to oxidation at high temperatures is a result of chromium in the alloy. The chromium initially oxidises on the surface of the metal, forming chromium oxide (C^Cb) which acts as a protective layer.

[0231] The data sheet in Table 1 below specifies expected mechanical properties and characteristics of the GRCop-42 alloy when manufactured. All data is based on parts built using Velo3D standard 50 pm layer thickness parameters, using Praxair TruForm™ CU42- N30, a Velo3D-approved powder. Parts built from GRCop-42 on a Sapphire System can be heat treated like those manufactured by other methods.

[0232] Mechanical Properties at Room Temperature - GRCop-42:3. Mechanical and test samples printed in vertical orientation.4HIP conditions: 1750 ± 25F, 15 ± 0.5 ksi; 3 hours (+ 15 / -0 min) in inert environment.Table 1

[0233] Materials that have high thermal conductivity and high maximum service temperatures are preferrable. GrCop-42 has a high thermal conductivity (approximately 280W / mK to approximately 325 W / mK), strong oxidation resistance, and a high service temperature making it highly desirable for the heat transfer element 20, 120, 220, 320.

[0234] Utilising copper was found to significantly reduce thermal hotspots and thermal stress on the heat transfer element 20 through transient thermal simulations. Example results from these thermal simulations are shown in the heat maps represented in Figures 20A and 20B.

[0235] Figure 20A is a heat map from a thermal simulation of a body 2, 102, 202 manufactured from steel, illustrating thermal hotspots and thermal stress across the heat transfer element 20, 120, 220 after the heater 30, 130, 230 has generated heat for 70 seconds. The areas of direct contact with the heater 30, 130, 230 can be seen to have reached a temperature of approximately 620°C. This heating effect is localised, and the extremities of the lattice structure 22 appear to remain significantly cooler (approximately 238°C to 390°C) even after the heater 30, 130, 230 has generated heat for 70 seconds. In other words, the steel body 2, 102, 202 would require approximately 70 seconds of warm-up time with no propellant stream S flowing to reach a desired approximate 300°C average thruster 1, 101, 201 temperature. After 70 seconds, the steel body 2, 102, 202 varied in temperature from approximately 162°C to approximately 620°C, resulting in large thermal stress. The bold arrow labelled "T" indicates the direction of thrust produced by the thruster 1, 101, 201.

[0236] Figure 20B is a heat map from a thermal simulation of a thruster 1, 101, 201 body 2, 102, 202 manufactured from GRCop-42 alloy, illustrating thermal hotspots and thermal stress across the heat transfer element 20, 120, 220 after 48 seconds of the heater 30, 130, 230 generating heat. As shown in Figure 20B, the areas of transfer of heat through the lattice structure 22 are similar to those of Figure 20A, however, the highest temperature is about 317K (as contrasted to 620K) with the thermal energy appearing to be more evenly distributed through the lattice 22. This gives a smaller temperature range of 298K - 317K across the lattice structure 22 after 48 seconds of heat. In other words, a GrCop-42 body 2, 102, 202 was shown in the simulations to take approximately 48 seconds of warm-up time, and the copper alloy body 2, 102, 202 was only seen to vary from approximately 290°C to approximately 320°C due to the high thermal conductivity, resulting in very small thermal stress. The bold arrow labelled "T" indicates the direction of thrust produced by the thruster 1, 101, 201.

[0237] In one simulated example, the thruster 1, 101, 201 has the following operating parameters shown in Table 2 below.Table 2

[0238] The thruster 1, 101, 201 can support a wide variety of operating conditions and can be used to produce thrust from approximately 20mN to approximately 10N, which is achieved by varying the propellant operating pressure and throat diameter (i.e., body outlet 3a).

[0239] The simulations carried for thrusters 1, 101, 201 have been considered for the thruster 301. In some embodiment, the thruster 1, 101, 201, 301 is additively manufactured from GRCop-42 as a monolithic part with an internal lattice structure 22 that provides compact heat exchange ability due to an inherently high volume to surface area ratio.

[0240] All thrusters 1, 101, 201, 301 can be configured to operate with nitrous oxide as a propellant stream S. Thrusters 1, 101, 201, 301 can also be configured for other propellant streams such as hydrazine monopropellant, ammonia, hydrazine, butane, nitrogen, or water.

[0241] The use of the heat transfer element 20, 120, 220, 320 has been shown to result in a high thermal efficiency and result in significant mass and volume reductions for the thruster 1, 101, 201, 301. In one example, the use of the heat transfer element 20, 120, 220, 320 can result in a reduction of the mass of the thruster 1, 101, 201, 301 to less than 200 grams, providing advantages over known thrusters.

[0242] The thrusters 1, 101, 201, 301 produce more than double the thrust of a UoSAT-12 thruster whilst having a length that is 25% of the length of the UoSAT-12 thruster and a width that is 20% the width of the UoSAT-12 thruster, therefore having a total length, size,and volume reduction of 75%. A thruster 1, 101, 201, 301 in accordance with described embodiments of the present invention provide a thruster with reduced mass and volume and make a thruster highly suitable for a small satellite platform.

[0243] The thruster 1, 101, 201, 301 can be used as part of a propulsion system 50 for an object such as, for example, a spacecraft or a satellite. With reference to Figures 18 and 19, the thruster 1 - and similarly the thruster 101, 201, 301 - can be integrated with the propulsion system 50 such that the propellant stream S flows from a propellant tank 37 to a filter 38, from the filter 38 to an isolation valve 39, from the isolation valve 39 to a thruster valve 40, and from the thruster valve 40 to a pressure reducing orifice 18 (not shown in Figure 19) and subsequently to the inlet 11, 311 of the thruster 1, 101, 201, 301. The pressure reducing orifice 18 reduces pressure in the propellant stream S from the tank 37 to the thruster 1. It is anticipated that a fixed orifice will be used having a diameter of between approximately 0.1 mm and approximately 0.8mm. However, a variable orifice can also be incorporated into the propulsion system 50 to allow the operating pressure of the propellant stream S entering the thruster 1, 101, 201, 301 to be varied. The pressure reducing orifice 18 can be incorporated into the thruster 1, 101, 201, 301 or thruster body 2, 102, 202, 302 or the connector 14 immediately adjacent the inlet 11, 311.

[0244] The propellant tank 37 supplies the propellant stream S to the thruster 1, 101, 201, 301. The filter 38 filters out any contaminants from the propellant stream S prior to the stream entering the interior 19 of the body 2, 102, 202, 302. In this manner, the filtration protects the downstream holes 28, 29 and channels 25 from contamination and helps reduce the risk of any contaminant particles in the propellant stream S from blocking the nozzle 33, 133, 233, 333. The filter 38 can be in the form of a mesh.

[0245] The isolation valve 39 is used to stop the flow of propellant stream S into the thruster 1, 101, 201, 301 during maintenance procedures or for safety purposes. The thruster valve 40 is used to meter the flow of propellant stream S into the thruster 1, 101, 201, 301.

[0246] A controller 36 such as a computer can be connected to the valves 39, 40 to control the opening and closing of the valves. The controller 36 can also be connected to the heater 30, 130, 230, 330 and any heater sensors.

[0247] With reference to Figure 19, the thruster 1 is secured into the propulsion system 50 via the flange 4 and the connector 14, wherein a power source can be connected to theheater 30 via connection cable 35. The propellant tank 37 can be connected to the inlet 11 via the thruster valve 40.

[0248] The propulsion system 50 can further comprise a second controller (not shown) for command and on-board data handling associated with the propulsion system 50 and for attitude control systems for attitude determination and control. Attitude control includes the ability to rotate the spacecraft or satellite to re-orient sensors or dump momentum. Attitude determination can be made by any one or more of: magnetometers; sun sensors; horizon sensors; star cameras; and GPS receivers. Alternatively, the controller 36 can also be used for command and on-board data handling associated with the propulsion system 50 and for attitude control systems for attitude determination and control, in addition to controlling the opening and closing of the valves 39, 40.

[0249] In use, the body 2 receives the propellant stream S from the propellant tank 37 via the valve 40, which feeds the propellant stream through the tangential inlet 11 and into the plenum chamber 17. The propellant stream S spirals around the heater 30 in the plenum chamber 17 before uniformly entering the heat transfer element 20. The heater 30 uses electrical energy supplied by the power source via the connection cable 35 to heat the heat transfer element 20, increasing the surface temperature of the heat transfer element 20. The propellant stream then flows through the intricate and complex geometry of the lattice structure 22 of the heat transfer element 20 where the propellant stream is effectively mixed and heated H before it is expelled through the nozzle 33 to produce thrust T.

[0250] Figure 22 provides a flow chart of a method 470 of generating thrust T comprising the step 472 of supplying a propellant stream to an interior of a body. The method 470 further comprises the step 474 of traversing the propellant stream S through a plurality of curved channels in the interior of the body. The method 470 further comprises the step 476 of heating the propellant stream S as it traverses the plurality of curved channels. The method 470 then comprises the step 478 of expelling the heated propellant stream H from the interior of the body to generate thrust T. The body and plurality of curved channels can have similar or the same characteristics as those described herein for the body 2, 102, 202, 302 and curved channels 25, 325. The curved channels can be part of a lattice structure of a heat transfer element as described in relation to heat transfer element 20, 120, 220, 320, using lattice structure 22, 122, 222, 322. The propellant stream S can be heated by transfer of energy from the heat transfer element, which can be heated with a heater configured asdescribed in relation to heater 30, 130, 230, 330. The method 470 can be implemented using the thruster described herein including embodiments of the thruster 1, 101, 201, 301.

[0251] It is envisaged to use the technologies described herein in a device other than a thruster, and in applications other than, for example, satellite applications. For example, it is envisaged to have a device for decomposing nitrous oxide that uses the technology described herein. It is also envisaged to have a method for decomposing nitrous oxide that uses the technology described herein.

[0252] With reference to Figure 23, there is described a device 501 for decomposing nitrous oxide. The device 501 comprises a body 502 that may be similar and have the same characteristics as the bodies 2, 102, 202, 302 described herein. The body 502 has an inlet 511 for introducing a stream of nitrous oxide N into the body 502. The device 501 further comprises a heat transfer element 520 located within the interior of the body 502 and configured to interact with the stream N of nitrous oxide. The device 501 further comprises a heater (not shown) positioned to transfer heat to the heat transfer element 520. The heat transfer element 520 defines a plurality of curved channels 525 connecting the inlet 511 to an outlet 503a. The plurality of curved channels 525 is arranged to generate turbulent mixing of the stream of nitrous oxide N, and to transfer heat energy from the heat transfer element 520 to the stream of nitrous oxide N. The heater is configured to generate and transfer heat to the heat transfer element to cause the decomposition of the nitrous oxide. The device 501 provides similar features as described herein in relation to the thruster 1, 101, 201, 301. The inlet 511 provides similar characteristics to those described herein in relation to inlet 11. In a specific embodiment, the heater of the device 501 is a resistive heater comprising ceramic material as described herein in relation to the heater 330.

[0253] The heat transfer element 520 of device 501 defines a cavity 527 for receiving the heater therein. The cavity 527 is illustrated in Figure 23 as a bore and can be a blind bore. The cavity 527 is centrally located in the heat transfer element 520 and a heater is centrally located within the heat transfer element 520. The heater can be positioned within the heat transfer element 520 with a gap separating the heater from the heat transfer element 520. The gap can be assimilated to gap 307 between heater 330 and heat transfer element 320. The heat transfer element 520 thereby receives heat energy from the heater by radiation.

[0254] The body 502 extends from a first end 505 to a second end 508. The first end 505 defines a plenum chamber 517 for receiving the stream of nitrous oxide N and forcing thestream N into the heat transfer element 520. The plenum chamber 517 can be annular and at least partially surrounds the cavity 527.

[0255] The plurality of curved channels 525 of the heat transfer element 520 is configured to split the stream of nitrous oxide N into a plurality of tortuous flows (not shown) that traverse the heat transfer element 520 in a similar manner to heat transfer element 20, 120, 220, 320 described herein, and the curved channels 25, 335 and tortuous flows 32 described herein.

[0256] The heat transfer element 520 preferably comprises a lattice structure 522 as has been described in relation to the thruster 1, 101, 201, 301, i.e. the lattice structure 22, 322. The lattice structure 522 of the device 501 comprises the plurality of curved channels 525, with each curved channel 525 of the plurality of curved channels 525 being separated from an adjacent curved channel 525 by a channel wall 526. The plurality of curved channels 525 form a series of holes 528, 529 through the heat transfer element 520. The plurality of holes 528, 529 together define a plurality of intersection points between the curved channels 525 whereby the stream of nitrous oxide N can be repeatedly split to form the plurality of tortuous flows 32 that intersect with one another while weaving through the lattice structure 522.

[0257] The lattice structure 522 of device 501 further comprises an array formed from a plurality of the unit cells 471, 473 that are tessellated to form a Triply Periodic Minimal Surface (TPMS) lattice structure. In a specific embodiment, each unit cell of the array has a diamond topological geometry (unit cell 473) and the lattice structure is a Diamond Triply Periodic Minimal Surface lattice structure, as has been described herein in relation to the lattice structure 22, 322. The body 502 and heat transfer element 520 of the device 501 can be manufactured from GRCop-42 alloy, as has been described herein in relation to the body 2, 102, 202, 302 and heat transfer element 20, 120, 220, 320. The body 502 and heat transfer element 520 of the device 501 can also be manufactured as a monolithic, single piece using additive manufacturing.

[0258] As previously described, a heater comprising a ceramic material enables operation of the heater wherein the heater can reach a temperature of at least approximately 1000°C. The heat transfer element 520 being positioned with a gap 307 separating it from the heater, it enables heating by radiation whereby sufficient heat energy can be transferred to heat the stream of nitrous oxide N to a maximum operating temperature, which is up to approximately 800°C, suitable for nitrous oxide decomposition. As the reaction ofdecomposition of nitrous oxide is an exothermic reaction, the device 501 presents significant advantages in that the components and characteristics enable initiating decomposition of the nitrous oxide and subsequently sustaining the reaction of decomposition of the nitrous oxide in a stable manner. The application of a lattice structure 522 for the heat transfer element 520 provides a complex geometry with continuous and smooth internal surfaces. This complex geometry combined with the use of the GRCop-42 material for the body 502 and heat transfer element 520 and ceramic material for the heater, and the thermal properties of the nitrous oxide, enable achieving a stable and maintainable reaction of decomposition of the nitrous oxide. The device 501 is characterised by high thermal conductivity, oxidation resistance, and a reduction of warm-up energy consumption for achieving appropriate conditions causing initiation of the reaction of decomposition.

[0259] A method 570 of decomposing nitrous oxide is illustrated in Figure 24, comprising the step 572 of supplying a stream of nitrous oxide to an interior of a body. The body can have similar characteristics to the body 2, 102, 202, 302, 502 described herein in relation to the thruster 1, 101, 201, 301, 501. The method 570 further comprises the step 574 of transferring heat to a heat transfer element located within the body, the heat transfer element comprising a plurality of curved channels to generate turbulent mixing of the stream of nitrous oxide N. The heat transfer element and curved channels provide similar characteristics to the heat transfer element 20, 120, 220, 320, 520 and curved channels 25, 325, 525 described herein in relation to the thruster 1, 101, 201, 301, 501. The heat can be provided by a ceramic heater like that described herein as heater 330 of thruster 301. The heater can be a resistive heater comprising ceramic material that is positioned within the heat transfer element with a gap separating the heater from the heat transfer element. The method 570 further comprises the step 576 of heating the stream of nitrous oxide N to a temperature that causes the decomposition of the stream of nitrous oxide N by passing the stream of nitrous oxide through the plurality of curved channels 523.

[0260] The method 570 can be implemented, for example, using the device 501 disclosed herein, having similar features and characteristics as the thruster 1, 101, 201, and 301 described herein. The method 570 provides a stable and sustainable decomposition of nitrous oxide while consuming a minimised amount of energy.

[0261] The method and device proposed herein to decompose nitrous oxide produce a self- sustaining reaction of decomposition of nitrous oxide due to the exothermic aspect of thereaction. Oxygen (O2) can be produced continuously from this decomposition as long as the inflow of nitrous oxide is continuous.

[0262] Carrying oxygen in space or under water typically poses a danger because the oxygen needs to be compressed (increasing the risk of explosion). It also requires significant volume for storage. While oxygen can be derived from electrolysis of water, producing oxygen via electrolysis is energy intensive. In contrast, the device 501 and method as set our herein provides a means to generate oxygen efficiently.

[0263] It will be appreciated by persons skilled in the art that numerous variations and modifications may be made to the above-described embodiments, without departing from the scope of the following claims. The present embodiments are, therefore, to be considered in all respects as illustrative of the scope of protection, and not restrictive.

[0264] Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this invention belongs. Although any methods and materials similar or equivalent to those described herein can also be used in the practice or testing of the present invention, a limited number of the exemplary methods and materials are described herein.

[0265] It is to be understood that, if any prior art publication is referred to herein, such reference does not constitute an admission that the publication forms a part of the common general knowledge in the art, in Australia or any other country.

[0266] In the claims which follow and in the preceding description of the invention, except where the context requires otherwise due to express language or necessary implication, the word "comprise" or variations such as "comprises" or "comprising" is used in an inclusive sense, i.e., to specify the presence of the stated features but not to preclude the presence or addition of further features in various embodiments of the invention.

[0267] LEGEND

Claims

CLAIMSThe claims defining the invention are as follows:

1. A thruster for generating thrust, the thruster comprising: a body having an inlet and an outlet, the inlet for introducing a propellant stream into the body and the outlet providing a nozzle for exiting a heated propellant stream from an interior of the body; a heat transfer element located within the interior of the body and configured to interact with the propellant stream; and a heater positioned to transfer heat to the heat transfer element, wherein the heat transfer element defines a plurality of curved channels connecting the inlet to the outlet to generate turbulent mixing of the propellant stream and to transfer heat energy from the heat transfer element to the propellant stream, to thereby generate thrust as the heated propellant stream is forced through the nozzle of the outlet.

2. The thruster of claim 1, wherein the heat transfer element defines a cavity for receiving the heater therein.

3. The thruster of claim 2, wherein the cavity is a bore that is centrally located in the heat transfer element.

4. The thruster of claim 2 or claim 3, wherein the cavity is coaxially aligned with the nozzle.

5. The thruster of any one of claims 1 to 4, wherein the heater is in direct contact with the heat transfer element.

6. The thruster of any one of claims 1 to 4, wherein the heater is positioned within the heat transfer element with a gap separating the heater from the heat transfer element.

7. The thruster of any one of claims 1 to 6, wherein the heater is a resistive heater.

8. The thruster of claim 7, wherein the heater comprises a ceramic material.

9. The thruster of any one of claims 1 to 8, wherein the plurality of curved channels is configured to split the propellant stream into a plurality of tortuous flows that traverse the heat transfer element.

10. The thruster of claim 9, wherein the curved channels of the plurality of curved channels are configured to intersect one another via a plurality of holes throughout the heat transfer element providing a plurality of intersection points which induce turbulent mixing between the plurality of tortuous flows of the propellant stream as the propellant stream traverses the heat transfer element.

11. The thruster of claim 10, wherein the plurality of intersection points increases the divergence of the plurality of tortuous flows between the inlet and the outlet, to increase heat transfer therebetween.

12. The thruster of any one of claims 1 to 11, wherein the heat transfer element comprises a lattice structure.

13. The thruster of claim 12, wherein the lattice structure comprises the plurality of curved channels, each curved channel of the plurality of curved channels being separated from an adjacent curved channel by a channel wall.

14. The thruster of claim 13, wherein a thickness of the channel wall decreases with increasing distance from the heater.

15. The thruster of any one of claims 12 to 14, wherein the lattice structure comprises an array of a plurality of unit cells that are tessellated to form a Triply Periodic Minimal Surface (TPMS) lattice structure.

16. The thruster of 15, wherein each unit cell of the array has a gyroid topological geometry and the lattice structure is a Gyroid Triply Periodic Minimal Surface lattice structure.

17. The thruster of claim 15, wherein each unit cell of the array has a diamond topological geometry and the lattice structure is a Diamond Triply Periodic Minimal Surface lattice structure.

18. The thruster of any one of claims 1 to 17, wherein the heat transfer element comprises a reinforcing member.

19. The thruster of claim 18, wherein the reinforcing member is formed from a densified portion of the heat transfer element and is formed integrally with the heat transfer element.

20. The thruster of claim 18 or claim 19, wherein the reinforcing member extends longitudinally within the heat transfer element and tapers towards a free end of the heat transfer element.

21. The thruster of any one of claims 1 to 20, wherein the inlet is arranged tangentially to the body.

22. The thruster of claim 21, wherein the inlet is positioned adjacent to the nozzle.

23. The thruster of any one of claims 1 to 22, comprising a plenum chamber within the body, the plenum chamber being configured to guide the propellant stream in a circulating motion into the heat transfer element.

24. The thruster of any one of claims 1 to 23, further comprising a plurality of cooling channels configured to provide a passage for the propellant stream from the inlet to the heat transfer element.

25. The thruster of claim 24, wherein each cooling channel of the plurality of cooling channels is embedded and extends longitudinally in an outer wall of the thruster.

26. The thruster of any one of claims 1 to 25, comprising a radiation shield surrounding the body.

27. The thruster of any one of claims 1 to 26, wherein the body and heat transfer element are manufactured from GRCop-42 alloy.

28. The thruster of any one of claims 1 to 27, wherein the body and heat transfer element are manufactured using additive manufacturing.

29. The thruster of any one of claims 1 to 28, wherein the body and heat transfer element are manufactured contemporaneously as a single piece.

30. A satellite comprising a thruster according to any one of claims 1 to 29.

31. A spacecraft comprising a thruster according to any one of claims 1 to 29.

32. A method of generating thrust comprising the steps of: supplying a propellant stream to an interior of a body; traversing the propellant stream through a plurality of curved channels in the interior of the body; heating the propellant stream as it traverses the plurality of curved channels; and expelling the heated propellant stream from the interior of the body to generate thrust.

33. The method of claim 33, conducted using the thruster of any one of claims 1 to 29.

34. A device for decomposing nitrous oxide, the device comprising: a body having an inlet for introducing a stream of nitrous oxide into the body;a heat transfer element located within the body and configured to interact with the stream of nitrous oxide; and a heater positioned to transfer heat to the heat transfer element, wherein the heat transfer element defines a plurality of curved channels connecting the inlet to an outlet, the plurality of curved channels arranged to generate turbulent mixing of the stream of nitrous oxide, and to transfer heat energy from the heat transfer element to the stream of nitrous oxide, and wherein the heater is configured to generate and transfer heat to the heat transfer element to cause decomposition of the nitrous oxide.

35. The device of claim 34, wherein the heater is a resistive heater comprising a ceramic material.

36. The device of claim 34 or claim 35, wherein the heater is positioned within the heat transfer element with a gap separating the heater from the heat transfer element, the heat transfer element receiving heat energy from the heater by radiation.

37. The device of any one of claims 34 to 36, wherein the plurality of curved channels is configured to split the stream of nitrous oxide into a plurality of tortuous flows that traverse the heat transfer element.

38. The device of any one of claims 34 to 37, wherein the heat transfer element comprises a lattice structure.

39. The device of claim 38, wherein the lattice structure comprises the plurality of curved channels, each curved channel of the plurality of curved channels being separated from an adjacent curved channel by a channel wall.

40. The device of claim 38 or claim 39, wherein the lattice structure comprises an array of a plurality of unit cells that are tessellated to form a Triply Periodic Minimal Surface lattice structure.

41. The device of claim 40, wherein each unit cell of the array has a diamond topological geometry and the lattice structure is a Diamond Triply Periodic Minimal Surface lattice structure.

42. The device of any one of claims 34 to 41, wherein the body and heat transfer element are manufactured from GRCop-42 alloy.

43. The device of any one of claims 34 to 42, wherein the body and heat transfer element are manufactured as a monolithic piece using additive manufacturing.

44. A method of decomposing nitrous oxide comprising the steps of: supplying a stream of nitrous oxide to an interior of a body; transferring heat to a heat transfer element located within the body, the heat transfer element comprising a plurality of curved channels to generate turbulent mixing of the stream of nitrous oxide; and heating the stream of nitrous oxide to a temperature that causes the decomposition of the nitrous oxide by passing the stream of nitrous oxide through the plurality of curved channels.

45. The method of claim 44, conducted using the device of any one of claims 34 to 43.

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