Rocket Engine

Abstract

A rocket engine (fig.7,20) has nozzle 54 coupled to a combustion chamber (fig.5,52) via a throat (fig.4,56), the walls of the combustion chamber and nozzle are formed from a stack of sheets (fig.5,51), with channels in at least some of the sheets forming multiple propellant conduits 70 with cooling conduits (fig.4,72) for delivering propellant to the combustion chamber and providing cooling of the walls by propellant flow. The rocket engine may also have an electrolyser (fig.1,32) to generate hydrogen and oxygen gas from water, and the engine may be a bi-propellent engine using hydrogen and oxygen, where the cooling conduits only carry the hydrogen. There may be 10-30 sheets between 0.5-5.0mm thick formed from copper alloy diffusion bonded together. The engine may be used as a thruster of a spacecraft (fig.1,10) such as a satellite.

Description

The present disclosure relates to a rocket engine, for example for use as a thruster on board a satellite or other spacecraft. Some embodiments of the rocket engine use hydrogen and oxygen gas as propellants, for example produced by electrolysis of water on board the spacecraft. Introduction The design and construction of thrusters and rocket engines varies widely according to a range of constraints and requirements. For high specific impulse requirements but where low thrust to power ratios are acceptable, electrodynamic and electrostatic thrusters are frequently used, and these generally avoid storage of high energy density chemical propellants and storage of larger quantities of propellants at very high pressures. However, the very low thrust to power ratios which such electrical thrusters can achieve makes them impractical for many manoeuvre requirements. Stored chemical propellant systems are able to provide much higher thrust to power ratios, but with associated difficulties of propellant storage, toxicity and other issues. Use of electrolysis of water stored on board the spacecraft to provide hydrogen and oxygen propellants is an attractive alternative, but design of compact, reliable and cost effective rocket engine structures to use such propellants remains a challenge. It would be desirable to address limitations of the related prior art. Summary of the invention The invention relates to a rocket engine, which may also be referred to as a thruster, which is built using a laminated or bonded sheet construction, and which is arranged to use chemical propellant flow to cool parts of the engine for example at least a combustion chamber and / or a nozzle of the engine. The invention may particularly apply to such a rocket engine using gaseous hydrogen and oxygen propellants, where flow of hydrogen is used for the cooling, but various other propellants and propellant combinations may instead be used. More generally, the inventors have found that the laminated or bonded sheet construction techniques described below can be advantageous in engineering high efficiency cooling conduits in both the described and in various other space propulsion applications. The laminated or bonded sheet construction may at the same time be used for other parts of a vehicle, such as a space vehicle, incorporating the rocket or thruster, for example with the same stack of sheets providing an entire vehicle chassis, body or other framework integrally formed with the rocket engine or thruster. In some aspects the invention provides a rocket engine comprising: a stack of sheets or layers or wafers bonded together to define at least a combustion chamber, a nozzle coupled to the combustion chamber via a throat to receive combustion products from the combustion chamber, and a plurality of propellant conduits arranged to deliver one or more propellants to the combustion chamber, the propellant conduits being formed by channels defined in at least some of the sheets, the stack of sheets defining walls of at least the combustion chamber and nozzle, wherein the propellant conduits comprise cooling conduits located within the walls so as to provide cooling of the walls, and optionally of other parts of the rocket engine, through flow of propellant through the cooling conduits. More particularly, the walls may comprise end walls of the stack (corresponding to the major planar faces of the sheets) and side walls disposed between, and connecting, the end walls. The side and end walls then also form side and end walls of elements within the stack such as the combustion chamber and the nozzle. The propellant conduits comprise one or more feed conduits, each feed conduit being arranged to carry flow of propellant rearwards past the combustion chamber and then past at least a portion of the nozzle for delivery to the cooling conduits. These feed conduits may in particular be disposed in one or both of the side walls, in which case for each such side wall, the one or more feed conduits may be provided by a single feed conduit, which optionally extends through a majority, i.e. more than half, of the sheets. For either or each side wall, the cooling conduits may comprise one or more side wall cooling conduits. These side wall cooling conduits are preferably disposed inwardly within the side wall from the one or more feed conduits, and are arranged to carry a flow of propellant received through the one or more feed conduits forwards through the side wall past the at least a portion of the nozzle and then past at least a portion of the combustion chamber for delivery into the combustion chamber, for example past at least 50% or past at least 80% of the length of each of either or both of these structures. For either or each side wall, the one or more side wall cooling conduits may extend in parallel or adjacent to the one or more feed conduits along a majority of the length of the combustion chamber, and / or along a majority of the length of the nozzle, for example past at least 50% or past at least 80% of the length of each of these chambers. For either or each side wall, the one or more side wall cooling conduits may comprise a plurality of stacked side wall cooling conduits which are distributed through the stack of sheets, for example with each of the stacked side wall cooling conduits being defined by a different pair or group of the sheets. These stacked side wall cooling conduits may be said to form an array of such side wall cooling conduits, typically a linear array forming a straight line in the depth direction through the stack although other configurations are of course possible. For either or each side wall, each of the plurality of stacked side wall cooling conduits may be defined by channels formed in just one or in just two of the sheets, and in particular formed using surface but not through channels in just one or two of the sheets. If deeper cooling conduits are required each cooling conduit may additionally comprise one or more through channels and may be defined by two, three or more adjacent sheets. For each or either side wall, there may be more stacked side wall cooling conduits than the number of adjacent feed conduits. For example, there may be just one or just two feed conduits, and six or more side wall cooling conduits, for example twelve side wall cooling conduits. To allow for a different number of feed conduits and side wall cooling conduits in a particular side wall, for each side wall, at the end of the side wall adjacent to the nozzle and furthest from the combustion chamber, one or more manifold sections may be provided to couple flow of propellant from the one or more feed conduits into the one or more side wall cooling conduits. Each such manifold may be being positioned inwardly from the one or more feed conduits within the side wall. For each of one or more of the side wall cooling conduits passing the combustion chamber, a side wall film cooling port may be arranged to divert a portion of the propellant flow from the side wall cooling conduit into the combustion chamber, optionally via one or more cooling port manifolds, for example into a lower portion of the combustion chamber than main injection ports, and proximal to the throat so as to provide film cooling within the lower portion of the combustion chamber proximally to the throat. The cooling conduits may also or instead comprise one or more nozzle end wall cooling conduits arranged to direct a flow of propellant forwards through the end wall past at least a portion of the nozzle, and / or one or more combustion chamber end wall cooling conduits which in are arranged to direct a flow of propellant past at least a portion of the combustion chamber. An end wall convergence may be provided for channelling the propellant flow from the nozzle end wall cooling conduits to the one or more combustion chamber end wall cooling conduits. A narrowest cross section of the flow through the end wall convergence may be less than 20%, or less than 5%, of the widest cross section of the flow through either or both of the nozzle end wall cooling conduits to the one or more combustion chamber end wall cooling conduits. The flow of propellant for directing forwards via the one or more nozzle end wall cooling conduits, and / or the end wall convergence, and / or the one or more combustion chamber end wall conduits, may at least partly be received from the one or more feed conduits located within each side wall, for example via the manifold sections mentioned above. The one or more nozzle end wall cooling conduits, the end wall convergence and the one or more combustion chamber end wall cooling conduits may be defined by two adjacent sheets, one or both of the adjacent sheets comprising surface channels defining the conduits. The one or more nozzle end wall cooling conduits may for example comprise first and second arrays of nozzle end wall cooling conduits. These first and second arrays may be arranged to respectively direct the flow of propellant so as to converge from the two sides of the nozzle, for example from each manifold section, to a central plenum or central common channel arranged to carry the flow of propellant to the end wall convergence. Similarly, the one or more combustion chamber end wall cooling conduits may comprise an array of combustion chamber end wall cooling conduits, the array directing the flow of propellant to diverge from the end wall convergence. Each of these arrays may for example comprise from about three to twelve adjacent and / or parallel conduits. One or more end wall film cooling ports may also be provided, arranged to divert a portion of the propellant flow from the end wall convergence and / or from one or more of the combustion chamber end wall cooling conduits, optionally via one or more suitable manifolds, into a lower portion of the combustion chamber proximal to the throat, so as to provide film cooling within the lower portion of the combustion chamber. The rocket engine or thruster may for example be a bi-propellant engine arranged to use two propellants. Typically in this case, all of the cooling conduits will be arranged to carry the same one of the two propellants, although if two separate sets of cooling conduits were used each set could use a different one of the propellants. In some arrangements, the propellants are gaseous hydrogen and gaseous oxygen, and the cooling conduits would typically then be arranged to carry the gaseous hydrogen, although they could instead be arranged to carry the gaseous oxygen. The rocket engine, or a spacecraft comprising the engine, may also comprise an electrolyser arranged to generate the gaseous hydrogen and gaseous oxygen from water, and optionally one or more tanks to store the gaseous hydrogen and gaseous oxygen for subsequent delivery to the combustion chamber via the propellant conduits, either storing the two propellants separately in different tanks, or mixed together in one or more common tanks. Aspects of the invention also provide a spacecraft or other vehicle comprising the rocket engine or thruster arranged to provide motive force. The stack of sheets or layers of wafers which are laminated or bonded may be a stack of metal or metallic sheets, and these may be diffusion bonded together. A variety of metals may be used to form the sheets, but in some arrangements some or all of the sheets, or at least the sheets used to define the combustion chamber, nozzle and propellant conduits, may be formed of copper or of one or more copper alloys. The number of sheets, layers or wafers used to form the stack may typically be from about 10 to 30, but fewer or more may be used. A suitable thickness for each sheet, or the average thickness of the sheets within the stack (if some are of different thicknesses to each other), may for example be from about 0.2 to 5.0 mm or from about 0.1 to 10.0 mm. Each sheet may have a thickness which is less than about 5%, or less than about 2%, of its longest dimension or along the major axis of the engine. In some embodiments, all of the sheets have the same or substantially the same thickness. However in some other embodiments some of the sheets have different thicknesses. The use of one or more sheets within the stack which have a smaller thickness may for example assist in the construction of cooling conduits with smaller feature sizes or cross sectional areas. Aspects of the invention also provide methods corresponding to the above apparatus, for example including methods of manufacturing and methods of operating the described rocket engine or thruster. Such methods may for example include a method of fabricating a rocket engine comprising forming each sheet, layer or wafer, and bonding these together to form the described stack, including for example the described combustion chamber, nozzle, propellant conduits and other features. Such methods may also include a method of operating the described rocket engine or thruster, for example by supplying one or more propellants to the propellant conduits such that the one or more propellants pass to the combustion chamber for combustion, and thereby cooling at least either or both of the combustion chamber and the nozzle by at least one of the one or more propellants flowing through the cooling conduits. Brief summary of the drawings Embodiments of the invention will now be described, by way of example only, with reference to the drawings of which: Figure 1 shows a spacecraft within which a rocket engine or thruster according to the present invention is implemented; Figure 2 shows how the rocket engine of figure 1 may be implemented using a bonded or laminated stack of sheets or wafers; Figures 3A to 3D show in cross section some ways in which surface and through channels in the sheets of figure 2 can be used and combined to form propellant conduits with the stack of sheets; Figure 4 is a schematic showing some of the functional elements of the rocket engine of figures 1 and 3; Figure 5 illustrates in perspective view a single sheet of the stack of figure 2; Figure 6 provides an exploded view of the stack of sheets of figure 2 with the visible top internal sheet being similar to that of figure 5; Figure 7also provides an exploded view of the stack of sheets of figure 2, but with the visible top internal sheet providing arrays of cooling conduits within one of the end walls of the stack; and Figures 8B and 8C show the cross sections views indicated in figure 8A of the stack of sheets shown in figures 6 and 7. Detailed description of the embodiments Referring first to figure 1 there is shown schematically a spacecraft 10 comprising a rocket engine 20 according to the present invention. The spacecraft 10 may be a satellite in, or intended for, Earth orbit, a space launch vehicle, a lunar, interplanetary or planetary probe, or any of a variety of other types of spacecraft. A rocket engine 20 according to the present invention may also be used in other vehicles such as aircraft, water borne craft, or land vehicles of various types and sizes, and in various other application areas. When used to provide impulse to an orbiting satellite or in similar applications, the rocket engine 20 may frequently be referred to as a thruster. As illustrated in figure 1, the rocket engine 20 may advantageously be comprised in a vehicle, such as spacecraft 10, in which electrolysis is used to form propellants for use in the rocket engine 20, typically where the propellants are hydrogen and oxygen formed from the electrolysis of water, and for example where such electrolysis is carried out using electrical power generated using solar cells or thermoelectric generation, and / or electrical power stored in batteries. To this end, in figure 1 the spacecraft 10 comprises a solar array 30 arranged to deliver electrical power to an electrolysis unit 32. The electrolysis unit uses the electrical power to produce hydrogen and oxygen gas from water stored in a water store 34, and stores these in gaseous form in hydrogen and oxygen tanks 36, 38. Gas pressures of the order of 1x107 Pa (100 bar) or more can readily be achieved by electrolysis without additional mechanical compression, and so the tanks may be used to store the hydrogen and oxygen at pressures of this order (for example around 4x106 Pa) for subsequent delivery via control valves 40 to the rocket engine 20 to generate thrust through controlled combustion. In some arrangements, the hydrogen and oxygen tanks may be omitted and the hydrogen and oxygen fed directly from the electrolysis unit 32 to the rocket engine. In some practical working embodiments the gas pressures in the tanks 38 may be around 2.4x106 Pa, with the working pressure in a combustion chamber of the rocket engine being about 0.8x106 Pa. Such rocket engines using gaseous hydrogen and oxygen propellants sourced using electrolysis of an on board water supply are able to achieve much higher thrust to power ratios than electrostatic, electrodynamic, or other mono / bipropellant chemical thrusters typically used on satellites, although the specific impulse achievable is likely to be lower. However, the achievable specific impulse is likely to be higher than that typically attained using cold gas or electrothermal thrusters. The propellants produced by electrolysis of water are stoichiometric 2:1 of H2:O2, and it may therefore be desirable to feed and combust these propellants in the rocket engine in approximately this ratio, although it may be desirable to combust a slightly hydrogen rich mixture, partly to help reduce combustion chamber and nozzle temperatures. Although figure 1 illustrates rocket engine 20 being provided with hydrogen and oxygen gaseous propellants created locally using electrolysis of water, the described rocket engine may be used with hydrogen and oxygen gas propellants provided in other ways for example stored in the tanks 36, 38 before launch, and using hydrogen and oxygen propellants stored in liquid form. However the rocket engine 20 may use a variety of other propellants and forms of propellants, including different combinations of bipropellants, or monopropellants. For example, instead of using gaseous hydrogen for the fuel propellant, methane, ethane or hydrazine could be used. Instead of using gaseous oxygen for the oxidant propellant, nitrous oxide or dinitrogen tetroxide could be used. Figure 2 shows in external perspective view an example of a rocket engine 20 according to the invention which may be used in the spacecraft 10 of figure 1 or in other vehicles and applications as noted above. The rocket engine comprises a stack 50 of sheets 51 bonded together. Prior to stacking and bonding, the individual sheets 51 are cut or otherwise shaped so as to subsequently define within the stack 50 at least a combustion chamber 52 (not visible in this figure), a nozzle 54 coupled to the combustion chamber via throat 56 (again not visible) so as to receive combustion products from the combustion chamber, and a plurality of propellant conduits 70 (again not visible in this figure). The propellant conduits are formed or defined using channels 58 created in the sheets 51. Figures 3A to 3C show some different ways in which such channels 58 may be used or combined to form the propellant conduits 70, each figure showing a number of sheets 51 of stack 50 in cross section. In figure 3A a single surface channel in one sheet, closed by the planar surface of an adjacent sheet forms the propellant conduit. In figure 3B opposing surface channels 58 in adjacent sheets combining to form the propellant conduit. In figure 3C a through channel in a single sheet is closed by planar surfaces of adjacent sheets to form the propellant conduit. In figure 3D a deeper propellant channel is formed by providing two adjacent through channels and one surface channel, although of course various other combinations of channels may be used. These figures show characteristic cusping in the channel cross sections which is typical where single or double sided etching has been used to form the channels, but whether such cusping will be present and the extent of any such cusping will depend on various details of the techniques used to form the channels. Referring back to figure 2, the propellant conduits 70 serve to deliver the propellant or propellants from one or more propellant inlets 59, where the propellants are received under a suitable working pressure, to the combustion chamber, and optionally to other working parts of the rocket engine as needed. Importantly however, some of the propellant conduits also or instead serve as cooling conduits 72 located within various parts of the rocket engine to provide a cooling function so as to maintain a safe working temperature, and others serve as feed conduits arranged to feed propellant to or from the cooling conduits . For example, these cooling conduits 72 may be located within walls of the combustion chamber and nozzle which are defined by the stack of sheets. These walls may include end walls 62 which are typically oriented parallel to the planes of the sheets 51, and side walls 60 which are typically oriented transversely to the planes of the sheets 51 and / or extend between opposing end walls. Because of the stacked formation of the rocket engine 20, the cross sectional shape of the interior of the combustion chamber and / or of the nozzle may conveniently be rectangular, or at least approximately so, although other cross sectional shapes may be used. Although in figure 2 the external side walls 60 of the stack 50 are essentially flat and perpendicular to the planes of the sheets, these side walls could of course take other shapes and forms for example by suitable adjustment of the perimeters of the sheets forming the stack to form a different side wall profile. Similarly, although in figure 2 the external end walls 62 of the stack are essentially flat and parallel to the planes of the sheets, each being defined in figure 2 by the major face of one of the endmost sheets in the stack, these end walls 62 could take other forms for example by building a different profile using suitable shaped sheets. The sheets 51 may be suitably shaped before stacking so as to form or define the channels 58, propellant conduits 70, and other structures mentioned above within the stack 50 when stacked and bonded, using one or more of a variety of techniques, one of which is photochemical etching, for example using a ferric chloride solution. Other suitable techniques for forming the channels include other wet or dry etching techniques, electrochemical processes, laser cutting, stamping, and milling. Each of the stacked sheets may be formed of a metal or metallic material, for example of copper or of a copper alloy such as a suitable formulation of copper chromium zirconium. The sheets may be bonded together to form the stack 50 using a variety of suitable techniques, one of which is diffusion bonding which can be achieved by simultaneously applying heat along with pressure to press the sheets of the stack together. Such diffusion bonding may be carried out in a vacuum furnace using jigs, coated with a ceramic such as boron nitride, to apply the required pressure to achieve a strong bond. Prior to such diffusion bonding the sheets may be solvent cleaned and then oxide stripped using acid. Practical operating temperatures of the hottest parts of the rocket engine 20 such as the interior walls of the combustion chamber 52, throat 56 and nozzle 54, may be limited in practice by the material or materials used for the sheets, but typical working temperatures of 400 - 800 C suitable for copper alloy construction may be appropriate and easily achievable when suitable cooling arrangements such as those described below using propellant cooling conduits are implemented within the stack. The rocket engine depicted in figure 2 may be constructed to be of a wide range of sizes, but in some examples the total length (the longest depicted dimension along the main axis of the engine) may be around 50 to 500 mm, and the total width around 20 to 200 mm. The total thickness of the stack of sheets may typically be around 5 to 50 mm. A typical weight for the depicted stack of sheets may be around 0.1 to 10 kg. Typically, each sheet may have a thickness (between its two major faces) of around 0.2 to 5.0 mm, although thicknesses outside this range may be used if required, and this range may be taken as an average of the sheet thicknesses in a particular stack. Typically also, all of the sheets forming the stack 50 may be of the same or of approximately the same thickness, but this need not be the case. The number of sheets forming the stack may typically be in the range of about 5 to 50, without around 10 to 30 being convenient for many designs to provide sufficient complexity and control of form and structure. Figure 4 shows in a more schematic view some of the structures and components of the rocket engine 20 which are defined by the bonded stack 50 of sheets 51. For convenience of description, the left side of the figure will be described as the front of the engine, and the right side where the nozzle terminates will be described as the rear, with the terms forward and rearward being used accordingly. Thickness or depth refers to the dimension extending though the multiple layers of the stacked sheets, i.e. perpendicular to the plane of the figure. The stack 50 of sheets 51 defines within it the combustion chamber 52 and the nozzle 54, the nozzle being coupled to the rear of the combustion chamber via the narrow throat 56 so as to receive combustion products from the combustion chamber and to allow these to expand in further rearwards motion as part of the process of generating thrust. The stack of sheets may also define an ignition chamber 57, typically positioned forward of the combustion chamber and serving to better enable ignition of the combustion process in the combustion chamber, for example using a sparking device located in the ignition chamber but not shown in figure 4. In other arrangements a catalyst may be used to support or ensure ignition, for example a packed bed catalyst within the ignition chamber. Although for convenience in this schematic view the spacing between the combustion chamber and nozzle appears quite large, in practice this spacing will typically be defined by the very short extent of the throat 56 connecting the two. The throat may typically have a width of the order of 1.0 mm if the width of the combustion chamber is of the order of 20 mm. More generally, the cross sectional area of the throat at the narrowest point may typically be around 2% to 15% of the widest cross sectional area of the combustion chamber. The nozzle is depicted as being closed on all four sides transverse to the rearward flow of exhaust gases, while expanding in the direction of flow, but in some embodiments other more open nozzle forms may be used, such as an aerospike design in which an expansion region extends along an open plug or spike profile. The one or more propellant inlets 59 are typically positioned forward of the combustion chamber, and forward of the ignition chamber if provided, and accept the propellant(s) under pressure to be delivered to the combustion chamber at one or more injection ports 68 for subsequent combustion. If the rocket engine 20 is designed to use a monopropellant then there may be just a single propellant inlet 59, and optionally also a single injection port 68. In figure 4 the propellant inlets 59 are depicted as being located on an end wall 62 of the stack, but may instead be located one on either end wall, on a front side wall (i.e. at the left end of figure 4), or on either or both of the side walls 60, or suitably located in other ways. The stack 50 of sheets defines within it a plurality of propellant conduits 70 as discussed above, which are arranged to deliver the propellants to the combustion chamber at the injection ports 68. Optionally some propellant may also be delivered via the ignition chamber if provided. Some portion of the propellant may also or instead be delivered to other parts of the engine for various purposes for example directly to the interior of the nozzle, or to be released from the engine altogether for example at the end of the nozzle. Such releases may provide functions such as film cooling. The propellant conduits are provided using channels 58 defined in at least some of the sheets, for example as already illustrated in figures 3A to 3D. As noted above, some of these channels may be provided as surface channels in just one side of a particular sheet, and some may be provided as through channels connecting both major faces of a sheet. Any one conduit section may be defined by a combination of two or more sheets, and be defined by different combinations of surface channel, through channel and unchanneled planar sheet surface. Some lengths of the propellant conduits 70 may be very shallow, being defined by just two adjacent sheets, so potentially having a depth of rather less than half, and up to two sheet thicknesses. Other lengths of propellant conduit may be considerably deeper being defined by and formed through several or even most of the sheets, so could have a depth close to that of the entire stack of sheets. Because the various pathways of these propellant conduits 70 are potentially quite complex, they are not illustrated in any detail in this schematic figure. However, as already noted above some of the propellant conduits 70 also act as cooling conduits 72 so that propellant entering the stack via the propellant inlets 56 can be used to cool various parts of the rocket engine formed by the stack 50. If the engine is a bi-propellant engine then typically just one of the propellants will be used for such cooling purposes. If the propellants are hydrogen and oxygen then hydrogen would usually be used. However, both of the propellant types in a bi-propellant form of the engine could be used for these cooling purposes if required. The inventors have found that cooling using gaseous hydrogen, for example following production using electrolysis from water, can be very effective in providing the required cooling of the walls of the combustion chamber and nozzle. To achieve the desired cooling, cooling conduits 72 (shown as dashed lines in figure 4) are typically provided within walls of the rocket engine or stack 50, and more particularly within walls of at least the combustion chamber 52 and nozzle 54. Although the cooling conduits of figure 4 are illustrated as being within the side walls 60 of the combustion chamber and nozzle, where each side wall is oriented largely transversely to the planes of the sheets and / or extending between the end walls, other cooling conduits 72 not shown in this figure may be provided within the end walls 62 themselves which lie largely within the plane of the sheets as seen in figure 2. Cooling conduits 72 may extend through the side and end walls of the combustion chamber and nozzle along substantially the entire length of each, or along a portion of the length of both or each, for example extending along at least 50%, or along at least 80%, of the length of the nozzle in a frontrear direction, and / or similarly extending along at least 50%, or along at least 80%, of the length of the combustion chamber in a front-rear direction. The stack 50 may also comprise a plurality of feed conduits 71 (shown as dotted lines in figure 4) arranged to deliver propellant received from the propellant inlets to the cooling conduits 72. In some embodiments, the feed conduits 71 channel the propellant in a rearwards direction towards the ends of the nozzle where it is passed into cooling conduits which then pass the propellant forward again past the nozzle and combustion chamber to the injection ports 58, and a manifold section 78 may be provided to couple flow of propellant at the end of each side of the nozzle into the cooling conduits 72. In various embodiments as discussed in more detail below, the feed conduits 71 are formed within the side walls of the stack, and in particular within the side walls of the combustion chamber and nozzle. Some of the cooling conduits 72 may then also be formed within the same side walls, but inwardly of the feed conduits, that is closer to the interior walls of the combustion chamber and nozzle than the feed conduits, and further from the external walls of the stack. Other cooling conduits may be formed within the end walls. Notably, the cooling conduits 72 located within the side walls 60, for example as illustrated in more detail in figure 5, may extend alongside, adjacent to, or in parallel to the one or more feed conduits also located in the side walls, for example extending in this way along a majority of the length of the combustion chamber and / or along a majority of the length of the nozzle. In some arrangements, the cooling conduits 72 located in the side walls may follow meandering or oscillatory paths, so as to increase their overall length and internal surface area, and therefore their ability to provide cooling to the side walls. If so constructed, these meandering side wall cooling conduits may still be described as extending substantially in parallel to the respective adjacent feed conduit(s). Cooling conduits located in the end walls, for example as shown in more detail in figure 7, may also be arranged to follow meandering or oscillatory paths. In either or both of the side wall and end wall cooling conduits, such meanders might extend the overall path length of a conduit (for example as measured by a central axis of the conduit) by around 5% to 25%, and therefore increase the wall area of the conduit by about the same amount. If film cooling is also implemented as mentioned above, this may most beneficially be used around the region of the throat 56. Film cooling ports (not shown in figure 4) for reducing internal wall temperatures, especially at the throat, could be constructed using small feature size propellant conduits running within the side walls 60 and / or end walls 62 into the combustion chamber, either from the feed conduits 71 or cooling conduits 72 themselves, or from one or more manifolds provided for that purpose which are themselves fed from the feed or cooling conduits. Figure 5 shows, using a perspective engineering view of a single sheet 51 forming part of stack 50, how feed conduits 71 and cooling conduits 72 may be implemented in the side walls 60. This particular sheet 51 may be located roughly midway within the stack 50 between the end walls 62. In particular, the figure shows in more detail how both feed conduits 71 and cooling conduits 72 may be provided in the side walls 60 of one or both of the combustion chamber 52 and the nozzle 54. Although in the design depicted in figure 5 no ignition chamber is provided, if such an ignition chamber was provided then the feed and cooling conduits may also be provided in side walls of the ignition chamber if desired, although the amount of heat produced by the ignition chamber would likely be relatively low in comparison to the combustion chamber. The propellant conduits 70 shown in figure 4 include one or more feed conduits 71 on each side of the engine, each feed conduit being arranged to carry flow of propellant rearwards through a respective side wall past at least a portion of the combustion chamber and then past at least a portion of the nozzle, for example extending along at least 50%, or along at least 80%, of the length of the combustion chamber in a front-rear direction, and / or similarly extending along at least 50% or at least 80% of the length of the nozzle in a frontrear direction. Each feed conduit 71 is supplied with propellant, for example with hydrogen in the water electrolysis arrangement described above, from the associated propellant inlet 59. In the arrangement of figure 4 each feed conduit runs close to and parallel to the outer face of sidewall 60 past the combustion chamber and nearly to the end of the nozzle. In the arrangement depicted in figure 5, a single feed conduit 71 is provided on each side of the engine, with each such conduit extending through most of the available depth of the stack of sheets on its respective side, for example extending through at least 50% or at least 80% of the depth of the stack. Providing just a single feed conduit 71 within each side wall can be advantageous in reducing resistance to propellant flow, especially if the feed conduit supplies propellant to both side wall and end wall cooling conduits. Each feed conduit could be formed using a structure similar to that shown in figure 3D, but typically with the through channels extending through a larger number of the sheets for example through between 5 and 20 sheets. However, if required, for example to improve thermal conduction across the sidewall or to improve mechanical strength, more than one such feed conduit 71 may be provided on each side of the engine, with such multiple feed conduits 71 being distributed through some or all of the depth of the stack of sheets. Alternatively or additionally, one or more of the stacked sheets could be used to provide one or more bridges across a said feed conduit 71 to help address such thermal conduction or mechanical strength issues. The propellant conduits in figure 5 also include one or more side wall cooling conduits 76 on each side of the engine, each side wall cooling conduit 76 being arranged to carry flow of propellant forwards through a respective side wall past at least a portion of the nozzle and then past at least a portion of the combustion chamber, for example extending along at least 50% or at least 80% of the length of the nozzle in a rear-front direction, and / or similarly extending along at least 50% or at least 80% of the length of the combustion chamber in a rear-front direction. Each side wall cooling conduit 76 is supplied with propellant, for example with hydrogen in the water electrolysis arrangement described above, from the one or more feed conduit(s) in the same sidewall, with the feed conduit(s) passing the propellant on to the side wall cooling conduits 76 of the same engine side in the vicinity of the end of the nozzle. In the arrangement of figure 5, the side wall cooling conduit(s) 76 run parallel and / or adjacent to the feed conduit(s) in the same side wall, at least along a majority such as at least 80% of the length of the combustion chamber, and / or along a majority such as at least 80% of the length of the nozzle, or more preferably both. Although not shown in figure 5, as noted above some or all of the side wall cooling conduits may follow an oscillatory or meandering path so as to increase the path length of each such cooling conduit along the side wall and increase the heat transfer from the side wall into the propellant passing along such a side wall cooling conduit. Bold arrows in figure 5 are provided to show the direction of propellant flow rearwards along the outer side wall cooling conduits, and then forwards along the inner side wall cooling conduits. In the arrangement depicted in figure 5, multiple side wall cooling conduits 76 are provided on each side of the engine, with each such conduit extending through a limited depth of the stack, for example extending through less than one or less than two sheets 51 of the stack, in such a manner that the multiple side wall cooling conduits are mutually distributed or stacked through the stack of sheets, for example such that each side wall cooling conduit in a particular side wall is defined by a different pair or group of adjacent sheets. Of course, not all sheets of the stack need be used to define any such side wall cooling conduit. For example, in the arrangement of figures 8B and 8C several adjacent sheets making up the end walls are not used to define any such side wall cooling conduit. The total number of side wall cooling conduits 76 stacked in this way through each side wall of the engine could for example be from 5 to 20. Providing multiple, stacked, side wall cooling conduits in this way may be advantageous in providing a higher rate of heat absorption in this interior portion of each side wall. Each one of these stacked, side wall cooling conduits could be provided in a manner similar to that shown in figure 3A or 3B, so that each conduit is defined by channels within just one sheet or two adjacent sheets. In some arrangements, each sheet defining part of the interior of the side walls of the combustion chamber and / or nozzle comprises a channel on each side of the engine which defines a separate one of the side wall cooling conduits. Although above it is suggested that, for each side wall, a single feed conduit carries propellant in a rearwards direction, and multiple stacked side wall cooling conduits disposed inwardly from the feed conduit then carry the propellant forwards again, more generally the number of feed conduits in each side wall may be more than one, but may be less than, for example less than half the number, of the corresponding stacked side wall cooling conduits carrying the propellant forwards again. The smaller number of outer feed conduits serves to promote propellant flow, while the larger number of inner side wall cooling conduits serves to promote heat exchange into the propellant from the combustion chamber and nozzle. In order to provide an effective coupling of propellant flow from the one or more outer side wall cooling conduits to a larger number of stacked inner side wall cooling conduits in each side wall, at least one manifold section 78 is provided within each side wall of the nozzle to make this coupling. Each manifold section 78 may be disposed within the sidewall inwardly of the one or more feed conduits so that the side wall cooling conduits then effectively begin at the forward end of the manifold sections 78. For example, if each side wall comprises a feed cooling conduit, the manifold section may then comprise a single length of propellant conduit feeding propellant flow into all of the stacked inner side wall cooling conduits. Because the amount of heat absorbed from the exhaust gases near the rear end of the nozzle is likely to be less than further up the nozzle, the use of these manifold structures can promote improved fluid flow of the propellant in this area while maintaining an adequate level of cooling of the nozzle structure. These manifold sections 78 may also provide an important function in delivering flow of propellant to end wall cooling conduits with reduced resistance to that flow as discussed below in connection with figures 7 and 8. For each side wall, the side wall cooling conduit(s) may terminate in an injection manifold 80 arranged to receive the propellant flow from all of the inner side wall cooling conduits of the corresponding side wall. Each injection manifold 80 then passes the propellant flow to the combustion chamber through one or more injection ports 68. The number of injection ports could be the same as the number of stacked side wall cooling conduits, and if so the injection manifold 80 might be dispensed with altogether, but by first combining the propellant flow from each side wall into a single corresponding injection manifold 80 the number and design of the injection ports 68 can be better optimized. Figure 5 also shows how each of some or all of the inner side wall cooling conduits may be provided with a corresponding film cooling port 82 which short cuts a small portion of the propellent flow from the side wall cooling conduit into a lower, that is rearward, portion of the combustion chamber, to provide enhanced film cooling of the wall of this portion of the combustion chamber downstream of the film cooling port as it enters the high temperature throat region. Figure 6 shows in exploded view the stack 50 of sheets of figure 2, so that an internal sheet similar to that of figure 5 is visible. The visible internal sheet 51 seen in figure 6 is higher up in the stack than that of figure 5, so that the propellant port 56 providing the propellant flow to the cooling conduits 72 is no longer visible. In the interests of clarity, not all of the features labelled in figure 5 are labelled in figure 6, but their correspondence will be readily apparent to the reader. Figure 7 shows, in exploded view, how cooling conduits 72 may also or instead be implemented in the end walls 62 of the stack 50 to provide enhanced cooling of the end walls through flow of propellant, received via the feed conduits, through the end walls, and especially enhanced cooling of the end walls of the combustion chamber 52 and the nozzle 54. Typically, the arrangement shown will be repeated at the other end wall 62 not visible in this figure. To this end, the cooling conduits 72 schematically shown in figure 4 may comprise one or more nozzle end wall cooling conduits 92 arranged to direct a flow of propellant forwards through the end wall 62 past at least a portion of the nozzle 54, for delivery via an end wall convergence 100 which is proximal to the throat 56 between the combustion chamber and the nozzle, to one or more combustion chamber end wall cooling conduits 102. These in turn are arranged to direct the flow of propellant forwards through the end wall 62 past at least a portion of the combustion chamber 52, for subsequent delivery into the combustion chamber for example via injection manifolds 80 which have already been discussed in connection with figures 5 and 6. Although not shown in figure 7, one or more end wall film cooling ports may be provided and arranged to divert a portion of the propellant flow from the end wall convergence 100 and / or from one or more of the one or more combustion chamber end wall cooling conduits, into a lower portion of the combustion chamber proximal to the throat so as to provide film cooling within the lower portion of the combustion chamber. Optionally this may be achieved by diverting the flow via one or more suitable end wall film cooling manifolds. These end wall film cooling ports may be provided in addition to the film cooling ports already described above and illustrated in figure 5. For example, the nozzle end wall cooling conduits 92 may be arranged to direct the flow of propellant forwards through the end wall 62 past or over at 80% of the length of the nozzle 54 in a direction along the rear-front axis of the engine, and the combustion chamber end wall cooling conduits 102 may be arranged to direct the flow of propellant forwards through the end wall 62 past or over at least 80% of the length of the combustion chamber 52 in a direction along the rear-front axis of the engine. If the side walls 60 of the engine are used to direct propellant rearwards through feed conduits 71 as already discussed above, then the manifold sections 78 coupling the feed conduits to the side wall cooling conduits 76 may also be used to feed propellant into the end wall cooling conduits. However, in other arrangements the propellant for flowing through the end wall cooling conduits may be delivered in other ways, for example through feed conduits which do not also feed the side wall cooling conduits The nozzle and combustion chamber end wall cooling conduits for each end wall 62 may typically be provided as a single layer of such conduits defined between two sheets forming part of each such end wall. However, if required two or more such layers of conduits could be defined in each end wall. In figure 7 the end wall cooling conduits for the illustrated end wall 62 are defined using surface channels formed in a single sheet of the stack which forms part of the end wall 62, similar to the arrangement of figure 3A, or using confronting surface channels in two adjacent sheets similar to the arrangement of figure 3B if greater conduit depth is required. The end wall convergence 100 between the nozzle and combustion chamber end wall cooling conduits for each end wall 62 may be provided in the same way, or may be of increased depth using one or more through channels in additional sheets if required. As illustrated in figure 7, the nozzle end wall cooling conduits 92 may comprise first and second arrays 94, 96 of nozzle end wall cooling conduits, the first and second arrays causing the flow of propellant to converge from the two sides of the nozzle, for example from the manifold sections 68, to a central plenum 98 arranged to carry the flow of propellant on to the end wall convergence 100. Also as shown in figure 7, each array of nozzle end wall cooling ducts may comprise a first region in which the propellant is directed more transversely towards the central axis of the engine, and then more longitudinally towards the end wall convergence 100 with the second regions delivering the propellant in a longitudinally distributed manner into the plenum 98. Flow of the propellant from the nozzle end wall cooling conduits 92 via the plenum 98 to the end wall convergence 100 will typically involve a considerable narrowing of the cross sectional area of the flow, and therefore significantly increased velocity and therefore substantially increased cooling effect in the corresponding vicinity of the throat 56. For example, the cross sectional area of the end wall convergence at its narrowest point may be less than 20% or less than 5% of the cross sectional area of the combined nozzle end wall cooling conduits 92 at their widest point. Following emergence from the end wall convergence 100, the flow of propellant diverges into an array of the combustion chamber cooling conduits 102, and this divergence may be guided by some divergence of the cooling conduits in the array which otherwise may be essentially linear as depicted in the figure. After passing through the array of combustion chamber cooling conduits 102 the propellant is delivered to the one or more injection manifolds 80, where it is combined with propellant delivered via the side wall cooling conduits if these are also used in the design. Bold arrows along one side of the end wall of figure 7 show the general direction of flow of propellant from the manifold sections 78 to the injection manifolds 80. Figure 8A shows a plan view outline of the stack of sheets already illustrated in the previous figures, marking cross sections B and C which are then shown in the corresponding cross section, directional views of figures 8B and 8C. These show the respective views in the directions of the arrows B and C following the making of the cross sectional cut. In figure 8C the central open area is that of the nozzle 54, with regions 120 representing the expanding side walls 60 of the nozzle as the structure recedes into the page- In this particular arrangement, the nozzle 54 is defined by an aperture in twelve of the sheets which is rectangular in this cross sectional view. In each such sheet, and on each side of the nozzle 54, a surface channel is provided to define a corresponding inner side wall cooling conduit 76, so that each side wall contains twelve stacked inner side wall cooling conduits. In a location outwardly from these individual surface channels, each sheet is also provided with a through channel, and in each side wall these line up to provide a single feed conduit 71. Each end wall 62 is several sheets thick in order to provide adequate strength, and surface channels in the innermost sheet of each end wall provide the nozzle end wall cooling conduits 92 in first and second arrays 94, 96 directing propellant flow into the plenum 98 for injection into the end wall convergence 100 (not shown). In the centre of figure 8B the inside of the combustion chamber 52 is seen, with regions 122 representing the expanding side walls 60 of the stack towards the front of the engine. At the back of the combustion chamber the injection ports 68 are visible. Note that there are only six injection ports on each side in this particular arrangement, compared with the twelve side wall cooling conduits 76, but in other arrangements significantly more injection ports may be used. Each side wall also contains the single outer feed conduit 71 already seen in section in figure 8C which conducts propellant from one of the propellant inlets towards the end of the nozzle for return to the injection manifolds and injection ports 68 via the side wall cooling conduits 76 as well as via the nozzle end wall cooling conduits 92 seen in figure 8C and the combustion chamber end wall cooling conduits 102 seen in this figure 8B. Similar to the nozzle end wall cooling conduits 92, the combustion chamber end wall cooling conduits 102 seen in this figure are defined by surface channels formed in the sheets which define the inside end walls of the combustion chamber in the manner shown in figure 3A, although deeper conduits could be formed using confronting surface channels in two sheets in the manner shown in figure 3B. In some particular embodiments, where each sheet has a thickness of around 0.5 mm, the feed conduits in the side walls which in figures 8B and 8C span twelve sheets may be around 6mm in depth, and the inner side wall cooling conduits each of which is formed using a surface channel in just one sheet may be around 0.25 mm in depth. Similarly, the end wall cooling conduits may be around 0.25 mm in depth. Although particular embodiments have been described, a number of alternatives and variations will be apparent to the skilled person without departing from the invention for example as set out in the claims.

Claims

1. A rocket engine comprising:a stack of sheets bonded together to define at least a combustion chamber, a nozzle coupled to the combustion chamber via a throat to receive combustion products from the combustion chamber, and a plurality of propellant conduits arranged to deliver one or more propellants to the combustion chamber, the propellant conduits being formed by channels defined in at least some of the sheets,the stack of sheets defining walls of the combustion chamber and nozzle, the propellant conduits comprising cooling conduits located within the walls so as to provide cooling of the walls through flow of propellant through the cooling conduits.

2. The rocket engine of claim 1 wherein:the walls comprise a side wall on each side of both the combustion chamber and nozzle, each side wall being oriented transverse to the sheets; andwithin each side wall, the propellant conduits comprise one or more feed conduits, each feed conduit being arranged to carry flow of propellant rearwards through the side wall past the combustion chamber and then past at least a portion of the nozzle for delivery to the cooling conduits.

3. The rocket engine of claim 2 wherein, for each side wall, the one or more feed conduits are provided by a single feed conduit which extends through a majority of the sheets.

4. The rocket engine of claim 2 or 3, wherein for each side wall, the cooling conduits comprise one or more side wall cooling conduits, the side wall cooling conduits being positioned inwardly within the side wall from the one or more feed conduits, and being arranged to carry a flow of propellant received through the one or more feed conduits forwards through the side wall past the at least a portion of the nozzle and then past at least a portion of the combustion chamber for delivery into the combustion chamber.

5. The rocket engine of claim 4 wherein, for each side wall, the one or more side wall cooling conduits extend in parallel to the one or more feed conduits along a majority of the length of the combustion chamber and along a majority of the length of the nozzle.

6. The rocket engine of claim 4 or 5 wherein, for each side wall, the one or more side wall cooling conduits comprise a plurality of stacked side wall cooling conduits which are distributed through the stack of sheets.

7. The rocket engine of claim 6 wherein, for each side wall, each of the plurality of stacked side wall cooling conduits is defined by channels formed in just one or in just two of the sheets.

8. The rocket engine of claim 6 or 7 wherein, for each side wall, there are more stacked side wall cooling conduits than the number of feed conduits.

9. The rocket engine of claim 8 wherein, for each side wall, at the end of the side wall adjacent to the nozzle and furthest from the combustion chamber, one or more manifold sections couple flow of propellant from the one or more feed conduits into the one or more side wall cooling conduits, the manifold being positioned inwardly within the side wall from the one or more feed conduits.

10. The rocket engine of any of claims 4 to 9 further comprising, for each of one or more of the side wall cooling conduits passing the combustion chamber, a side wall film cooling port arranged to divert a portion of the propellant flow from the side wall cooling conduit into a lower portion of the combustion chamber proximal to the throat so as to provide film cooling within the lower portion of the combustion chamber.

11. The rocket engine of any of claims 4 to 10 wherein at least some of the side wall cooling conduits following a meandering path so to increase the path length of these side wall cooling conduits within the side wall.

12. The rocket engine of any preceding claim wherein:the walls comprise at least one end wall of the combustion chamber and nozzle, the end wall being oriented parallel to the sheets; andthe cooling conduits comprise one or more nozzle end wall cooling conduits arranged to direct a flow of propellant forwards through the end wall past at least a portion of the nozzle, for delivery via an end wall convergence which is proximal to the throat between the combustion chamber and the nozzle, to one or more combustion chamber end wall cooling conduits which in turn are arranged to direct the flow of propellant forwardsthrough the end wall past at least a portion of the combustion chamber, for subsequent delivery into the combustion chamber.

13. The rocket engine of claim 12 when dependent on any of claims 2 to 11, wherein the flow of propellant for directing forwards via the one or more nozzle end wall cooling conduits, the end wall convergence, and the one or more combustion chamber end wall conduits, is at least partly received from the one or more feed conduits within each side wall.

14. The rocket engine of claim 12 or 13 wherein the one or more nozzle end wall cooling conduits, the end wall convergence and the one or more combustion chamber end wall cooling conduits are defined by two adjacent ones of the sheets, one or both of the adjacent sheets comprising surface channels defining the conduits.

15. The rocket engine of any of claims 12 to 14 wherein the one or more nozzle end wall cooling conduits comprise first and second arrays of nozzle end wall cooling conduits, the first and second arrays respectively directing the flow of propellant to converge from the two sides of the nozzle to a central plenum arranged to carry the flow of propellant to the end wall convergence.

16. The rocket engine of any of claim 12 to 15 wherein the one or more combustion chamber end wall cooling conduits comprise an array of combustion chamber end wall cooling conduits, the array directing the flow of propellant to diverge from the end wall convergence.

17. The rocket engine of any of claims 12 to 16 further comprising one or more end wall film cooling ports arranged to divert a portion of the propellant flow from the end wall convergence and / or from one or more of the combustion chamber end wall cooling conduits into a lower portion of the combustion chamber proximal to the throat so as to provide film cooling within the lower portion of the combustion chamber.

18. The rocket engine of any preceding claim wherein the rocket engine is a bipropellant engine arranged to use two propellants, and all of the cooling conduits are arranged to carry the same one of the two propellants.

19. The rocket engine of claim 18 wherein the propellants are gaseous hydrogen and gaseous oxygen, and the cooling conduits are arranged to carry gaseous hydrogen.

20. The rocket engine of any preceding claim further comprising an electrolyser arranged to generate the gaseous hydrogen and gaseous oxygen from water, and optionally one or more tanks to store the gaseous hydrogen and gaseous oxygen for subsequent delivery to the combustion chamber via the propellant conduits.

21. The rocket engine of any preceding claim wherein the stack of sheets bonded together is a stack of metal sheets diffusion bonded together.

22. The rocket engine of any preceding claim wherein each of the stack of sheets is formed of copper or a copper alloy.

23. The rocket engine of any preceding claim wherein the number of sheets forming the stack is from 10 to 30, and / or an average thickness of the sheets within the stack is from 0.5 to 5.0 mm.

24. A method of fabricating a rocket engine according to any preceding claim, comprising:forming each sheet;bonding the sheets together to form the stack defining the combustion chamber, nozzle, throat, and propellant conduits.

25. A method of operating a rocket engine according to any of claims 1 to 23 comprising:supplying one or more propellants to the propellant conduits such that the one or more propellants pass to the combustion chamber for combustion; andcooling at least the combustion chamber and the nozzle using at least one of the one or more propellants flowing through the cooling conduits.

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