Track maneuvering propeller, track maneuvering propelling system and track transportation aircraft

Abstract

A thruster (100) for orbital maneuver, comprising: a diffusion channel (111) having an axis of symmetry (S); a combustion chamber (110); a throat tube (112) interposed between the combustion chamber (110) and the diffusion channel (111); an injection plate (115) facing the combustion chamber (110); a plurality of first injection channels (125) in fluid communication with the combustion chamber (110), each first injection channel comprising an end portion (126) open in the combustion chamber (110); at least one second injection channel (135) in fluid communication with the combustion chamber (110), including an end portion (136) open in the combustion chamber (110); wherein an end portion (126) of the first injection channel (125) is located in the annular region (127) and extends along a respective injection direction (d1); wherein an end portion (136) of the at least one second injection channel (135) is located radially between the axis of symmetry (S) and an end portion (126) of the first injection channel (125); wherein the injection direction (d1) of the end portion (126) of each first injection channel (125) has an axial component directed towards the combustion chamber (110) and a radial component directed towards the axis of symmetry (S).

Description

Technical Field

[0001] The present invention relates to an orbital maneuvering propulsion device, an orbital maneuvering propulsion system including the orbital maneuvering propulsion device, and an orbital transport aircraft including the orbital maneuvering propulsion device. Background Art

[0002] An orbital transport vehicle is a vehicle that can store, transport and release payloads. This type of orbital transport vehicle can be carried on a space launch vehicle to perform space transportation missions and has a dedicated propulsion system to perform orbital maneuvers, such as orbit changes, after being released from the space launch vehicle.

[0003] The propulsion system of an orbital transport vehicle includes thrusters for orbital maneuvers and propellant tanks, which supply fuel and oxidizer (or oxidant) to the thrusters.

[0004] A thruster for orbital maneuvers generally includes: a combustion chamber in which the propellant is burned; an injection plate for injecting the propellant into the combustion chamber; a throat communicating with the combustion chamber through which the combustion material accelerated by the combustion chamber reaches the speed of sound under sound barrier conditions; a diffusion channel (or supersonic nozzle) extending from the throat to the opposite side of the combustion chamber through which the combustion material is accelerated to exceed the speed of sound; and a plurality of cooling channels for cooling the walls of the combustion chamber. A thruster for orbital maneuvers also includes a supply system for feeding the propellant into the combustion chamber, a spark plug for igniting the combustion, and valves and sensors for controlling the combustion.

[0005] In a propulsion system for bipropellant orbital maneuvers, the propellant can be delivered to the combustion chamber by a supply pump, or pressurized by a pressurizer (such as nitrogen or hydrogen) to maintain the fuel and oxidizer tanks at appropriate pressures so that the fuel and oxidizer can flow into the combustion chamber at appropriate flow rates and pressures.

[0006] Typically, thrusters for orbital maneuvers are turned on for a period of time during the combustion process, during which propellant is introduced into the combustion chamber for combustion, and the resulting combustion material is accelerated and discharged through the throat and diffusion channel. The thrusters for orbital maneuvers generate a predetermined force on the orbital transport vehicle for a predetermined time (so as to generate a specific pulse), which is necessary to complete a predetermined phase of the mission, such as to change the orbital parameters of the orbital transport vehicle.

[0007] Applicants have identified a need for a small size, low weight, low cost orbital transport vehicle for transporting and using small payloads, such as satellites (e.g., pico-satellites or cubesats) that are preferably non-motorized and miniaturized. Applicants have demonstrated that, in order to perform its mission, such an orbital transport vehicle requires a thruster with a total length (in the thrust direction) of less than 250 mm and a total weight of less than 2 kg, capable of generating a thrust force between 5 N and 22 N for at least 5 seconds, with a thrust specific impulse greater than 285 seconds, and capable of multiple consecutive firings. Applicants have recognized that a propulsion system for an orbital transport vehicle of this type should be versatile, lightweight, simple, inexpensive, and easy to operate.

[0008] The Applicant has noticed that a thruster for orbital maneuvers of the type described above can be produced using additive manufacturing techniques such as selective metal laser melting (laser powder bed fusion or L-PBF), which allows the production of most of the thruster, including the combustion chamber, the supersonic nozzle and at least part of the supply system, in a single block at low cost. The Applicant has also demonstrated that the cooling channels of the combustion chamber, the injection plate and at least part of the supply channels for conveying the propellant to the injection plate can all be integrated into the above-mentioned single block. This production technique requires that the single block is made of a material suitable for additive manufacturing.

[0009] Applicants have observed that handling the propellants of a propulsion system in preparation for launch generally requires particularly cumbersome safety procedures and involves non-negligible risks associated with the use of the propellants themselves, thereby increasing the deployment time and cost of orbital transportation vehicles and their payloads.

[0010] Applicants have recognized that the use of safe and easy-to-handle propellants would greatly simplify pre-launch operations, reducing the time and costs required.

[0011] The Applicant has noted that the use of so-called “green propellants”, which are less toxic and safer than conventional propellants, would allow further reduction of the risks associated with the use of propellants and would reduce the complexity of safety procedures.

[0012] Applicants have recognized that if such a "green propellant" is self-pressurizing, it can be stored in liquid form in corresponding tanks and evaporated during use to keep the pressure in the tank constant during consumption, avoiding the use of additional tanks for pressurizing agent, thereby reducing the weight of the propulsion system.

[0013] Applicants have found that the choices for both green propellants and self-pressurizing propellants are limited. In particular, Applicants have found that nitric oxide (N2O) is the only viable oxidizer that naturally possesses the above characteristics, does not require prior decomposition by a decomposition system (e.g., electrolysis or preheating), and does not require other elements to be dissolved within to become self-pressurizing.

[0014] Applicants have demonstrated that flame front temperatures above 2800°C can be achieved in thrusters for orbital manoeuvring using nitrous oxide as the oxidant in combination with a natural green, self-pressurising suitable fuel such as propylene.

[0015] Applicants have demonstrated that even when using currently available materials that are particularly suited to withstand high temperatures, and when manufacturing thrusters for orbital maneuvers by additive manufacturing techniques, the flame front temperature using nitrous oxide may not be able to reach the ignition time required to complete a certain phase of the mission due to melting or degradation of the combustion chamber.

[0016] Applicants have recognised that if direct exposure of the chamber walls to the flame front is avoided, the time required for chamber degradation will be increased and it is possible to ignite the thrusters in the time required to complete a certain phase of the mission.

[0017] The applicant has discovered that it is possible to introduce the fuel and the oxidant into the combustion chamber so that they come into contact with the combustion chamber walls in a stoichiometric ratio which is unfavorable for the development of combustion, and then to mix them in a stoichiometric ratio which is favorable for combustion after they have reached a volume of the combustion chamber remote from the combustion chamber walls (or in any case not in direct contact with the combustion chamber walls), in such a way that the walls of the combustion chamber will not be directly impacted by the flame front. Summary of the invention

[0018] Thus, the present invention in its first aspect relates to a thruster for orbital manoeuvring, comprising:

[0019] A diffusion channel having an axis of axial symmetry;

[0020] a combustion chamber whose sides are defined by combustion chamber walls;

[0021] a throat between the combustion chamber and the diffusion passage to connect the combustion chamber to the diffusion passage fluid;

[0022] an injection plate facing the combustion chamber;

[0023] a plurality of first injection channels in fluid communication with the combustion chamber, the plurality of first injection channels being configured to inject a first combustion component into the combustion chamber, wherein each of the plurality of first injection channels includes an end portion located at the injection plate and opening into the combustion chamber;

[0024] at least one second injection channel in fluid communication with the combustion chamber, the at least one second injection channel being configured to inject a second combustion component into the combustion chamber, and the at least one second injection channel including an end portion located at the injection plate and opening into the combustion chamber;

[0025] wherein the end portion of a first injection channel among the plurality of first injection channels is substantially adjacent to an edge joint between the injection plate and the combustion chamber wall, the end portion of the first injection channel is located within an annular region, the center of the annular region intersects the symmetry axis, and the end portion of the first injection channel extends along a respective injection direction;

[0026] wherein an end portion of at least one second injection channel is radially located between the symmetry axis and an end portion of a first injection channel among the plurality of first injection channels;

[0027] The injection direction of the end portion of each first injection channel has an axial component pointing to the combustion chamber and a radial component pointing to the axis of symmetry.

[0028] In its second aspect, the present invention relates to a propulsion system for orbital maneuvers comprising:

[0029] A thruster for orbital manoeuvring according to the first aspect;

[0030] a first tank containing a first type of propellant, the first tank being fluidly coupled to a first injection channel of the plurality of first injection channels;

[0031] a second tank containing a second type of propellant, the second tank being fluidly coupled to the at least one second injection passage;

[0032] Wherein, the first combustion component and the second combustion component are self-pressurizing.

[0033] In a third aspect thereof, the invention relates to an orbital transport vehicle comprising a thruster for orbital manoeuvring according to the first aspect.

[0034] A plurality of first injection passages allow the first combustion component to be injected into the combustion chamber along a substantially annular path.

[0035] At least one second injection channel allows a second combustion component to be injected into the combustion chamber within the annular path of the first combustion component.

[0036] Applicants believe that in this way, the second combustion component does not immediately mix with the first combustion component, but rather, at least upon its introduction, the second combustion component is located within the first combustion component and is separated from the first combustion component.

[0037] The injection direction of the end portion of each first injection channel has an axial component pointing to the combustion chamber and a radial component pointing to the axis of symmetry so as to orient the jet of the first combustion component entering the combustion chamber toward the axis of symmetry. In this way, the first combustion component injected into the combustion chamber converges toward the axis of symmetry and the second combustion component.

[0038] The applicant believes that in this way, combustion occurs in the central area of ​​the combustion chamber, and the first combustion component, which has not yet been mixed with the second combustion component, adheres to the wall of the combustion chamber itself, is located between the wall and the central area of ​​the combustion chamber, flows at least in the first part of the combustion chamber, and the first combustion component itself is between the combustion chamber and the flame front, preventing the flame front from directly contacting the wall of the combustion chamber.

[0039] The diffusion channel has an axis of symmetry. In the present specification and the appended claims, expressions such as "axial", "axially", "radial", "radially", "radially inside", "radially outside", "circumferential", "circumferentially", "tangentially", "tangentially" and the like are used with reference to this axis of symmetry.

[0040] The terms “radial” and “axial” and the expressions “radially inside / outside” and “axially inside / outside” are used with reference to the vertical direction and the direction parallel to the axis of symmetry of the diffusion channel, respectively.

[0041] The expressions “radially innermost” and “radially outermost” refer to positions close to and far from the axis of symmetry of the diffusion channel, respectively.

[0042] The terms "circumferential" and "circumferentially" refer to directions along a circumference that lies in a plane that is perpendicular to the axis of symmetry of the diffusion passage and has a center that intersects the axis of symmetry of the diffusion passage.

[0043] The terms "tangential" and "tangentially" are used to indicate a direction tangential to a circumference lying in a plane perpendicular to the axis of symmetry of the diffusion channel and having a center passed through the axis of symmetry of the diffusion channel. The tangential direction is contained in the same plane as the circumference.

[0044] In the present description and the appended claims, the expressions "first combustion component" and "second combustion component" are understood to mean respective substances which, if properly mixed together and if activation energy is provided to said mixture, will produce an exothermic redox reaction.

[0045] The expression "injection plate" is to be understood as that part of the impeller which faces the combustion chamber, opposite the throat, from which injection plate the first combustion component and the second combustion component are injected into the combustion chamber.

[0046] In at least one of the above-mentioned aspects, the present disclosure may be implemented according to one or more of the following embodiments, which may be combined with each other.

[0047] Preferably, the first combustion component is an oxidant.

[0048] Preferably, the first combustion component comprises nitrous oxide (N2O).

[0049] Preferably, the first combustion component is nitrous oxide.

[0050] Preferably, the second combustion component is a fuel.

[0051] Preferably, the second combustion component comprises propylene (C3H6).

[0052] Preferably, the second combustion component is propylene.

[0053] Preferably, said end portion of the first injection channel is straight.

[0054] Preferably, each end portion of the first injection channel is an extension of a single channel.

[0055] Preferably, said end portion of the first injection channel has no branches.

[0056] Preferably, the end portion of the first injection channel has a constant cross section along the injection direction.

[0057] Preferably, the first injection channel is straight.

[0058] Preferably, each first injection channel is a single channel.

[0059] Preferably, the first injection channel has no branches.

[0060] Preferably, the first injection channel has a constant cross section along the injection direction.

[0061] Preferably, the diameter of the end portion of each first injection channel is between 0.5 mm and 0.9 mm, more preferably between 0.6 mm and 0.8 mm, even more preferably between 0.65 mm and 0.75 mm, for example 0.7 mm.

[0062] Preferably, the end portion of at least one second injection channel is straight.

[0063] Preferably, the end portion of at least one second injection channel is an extension of a single channel.

[0064] Preferably, the end portion of at least one second injection channel has no branches.

[0065] Preferably, the end section of at least one second injection channel has a constant cross section along the injection direction.

[0066] Preferably, at least one second injection channel is straight.

[0067] Preferably, the at least one second injection channel is a single channel.

[0068] Preferably, at least one second injection channel has no branches.

[0069] Preferably, the at least one second injection channel has a constant cross section along the injection direction.

[0070] Preferably, the diameter of the end portion of each second injection channel is between 0.4 mm and 0.8 mm, more preferably between 0.5 mm and 0.7 mm, even more preferably between 0.55 mm and 0.65 mm, for example 0.6 mm. Preferably, the combustion chamber wall extends from the injection plate to the throat.

[0071] Preferably, the combustion chamber wall and the injection plate are connected to each other via an edge joint.

[0072] Preferably, at the edge joint, the injection plate and the combustion chamber wall form an edge angle.

[0073] Preferably, the edge angle is between 80° and 100°, more preferably between 85° and 95°, such as approximately 90°.

[0074] Preferably, the annular area is defined by an inner circumference and an outer circumference.

[0075] Preferably, the inner circumference and the outer circumference are concentric.

[0076] Preferably, the outer circumference is located at the edge joint.

[0077] Preferably, the radius of the inner circumference of the annular area is equal to or greater than 75% of the radius of the outer circumference, more preferably equal to or greater than 85% of the radius of the outer circumference.

[0078] Preferably, the first combustion component is injected close to the combustion chamber wall.

[0079] Preferably, the end portion of the at least one second injection channel is located between the inner circumference of the annular region and the axis of symmetry.

[0080] An end portion of the at least one second injection channel is located outside the annular region defined between the inner circumference and the outer circumference.

[0081] In this way, the first combustion component will be injected to a position away from the combustion chamber wall.

[0082] Preferably, the injection direction of the end portion of each first injection channel does not include any component along the tangential direction.

[0083] Preferably, the first combustion component is injected into points or areas arranged along the axis of symmetry.

[0084] Preferably, the first combustion component is injected into the combustion chamber from the end portion of the first injection passage to form a conically distributed jet.

[0085] In this way, the second combustion component tends to remain within the cone-shaped distribution of the first combustion component.

[0086] Preferably, the end section of at least one second injection channel extends along a respective injection direction having an axial component pointing towards the combustion chamber.

[0087] In one embodiment, the injection direction of the end portion of the at least one second injection channel does not include any component in the radial direction.

[0088] In various embodiments, the injection direction of the end portion of at least one second injection channel comprises a radial component pointing towards the axis of symmetry.

[0089] In another different embodiment, the injection direction of the end portion of at least one second injection channel includes a radial component pointing to the opposite side relative to the symmetry axis. Preferably, the injection direction of the end portion of at least one second injection channel does not include any component along the tangential direction.

[0090] In this way, the second combustion component is injected parallel to the axis of symmetry directly towards the central region of the combustion chamber where combustion is desired to take place.

[0091] Optionally, the injection direction of the end portion of at least one second injection channel includes a component along a tangential direction.

[0092] In this manner, the second combustion component is injected into the combustion chamber with a rotational motion about the axis of symmetry, which helps to keep combustion away from the injector plate and components mounted thereon.

[0093] Preferably, the angle formed by the projection of the injection direction of the end portion of each of the plurality of first injection channels on the plane containing the symmetry axis relative to the symmetry axis is between 20° and 60°, preferably between 30° and 50°, more preferably between 35° and 45°, for example about 40°. The angle can be calculated as the inverse tangent of the ratio of the radial component to the axial component of the injection direction.

[0094] Preferably, the injection direction of the end portion of each of the multiple first injection channels forms an angle with respect to a plane orthogonal to the axis of symmetry, which angle is between 30° and 70°, preferably between 40° and 60°, more preferably between 45° and 55°, for example approximately 50°.

[0095] Applicants have found that these angle values ​​are particularly effective in maintaining combustion in the central region of the combustion chamber and away from the combustion chamber walls.

[0096] Preferably, the number of first injection channels in the plurality of first injection channels is between 7 and 28, preferably between 10 and 22, even more preferably between 13 and 16, for example 14.

[0097] Preferably, the angle formed by the projection of the injection direction of the end portion of at least one second injection channel on the plane containing the axis of symmetry relative to the axis of symmetry is between 0° and 20°, preferably between 0° and 10°, more preferably between 0° and 5°. The angle can be calculated as the inverse tangent of the ratio of the radial component (pointing to the axis of symmetry or to the opposite side of the axis of symmetry) and the axial component of the second injection direction.

[0098] Preferably, the injection direction of the end portion of at least one second injection channel forms an angle of 70° to 90°, preferably 80° to 90°, more preferably 85° to 90° with respect to a plane orthogonal to the axis of symmetry. If the angle is not 90°, the injection direction of the end portion of at least one second injection channel intersects the axis of symmetry or diverges with respect to the axis of symmetry.

[0099] Preferably, a plurality of second injection channels are provided.

[0100] Preferably, the number of the second injection channels in the plurality of second injection channels is between 2 and 8, preferably between 3 and 6, for example 4.

[0101] Preferably, the second injection channel has a smaller cross section than the first injection channel.

[0102] Preferably, the second injection channels are distributed along a circumference centered on the axis of symmetry, preferably equidistant from each other.

[0103] Preferably, there is provided a spark plug which is mounted on the injection plate at the axis of symmetry and which is configured to generate a spark in the combustion chamber.

[0104] Optionally, a plurality of spark plugs may be provided mounted on the injection plate, equidistant relative to the axis of symmetry and configured to generate sparks in the combustion chamber.

[0105] Preferably, a plurality of cooling channels are provided which are arranged around the combustion chamber and are fluidically connected to the first injection channel.

[0106] Preferably, the number of cooling channels in the plurality of cooling channels is between 14 and 18, more preferably between 15 and 17, such as 16.

[0107] The cooling passages further cool the combustion chamber walls and preheat the first combustion components.

[0108] Preferably, a cooling channel of the plurality of cooling channels is integrated into the combustion chamber wall.

[0109] Preferably, a cooling passage of the plurality of cooling passages is fluidly coupled to the first tank, the cooling passage being configured to receive the first combustion component from the first tank prior to introducing the first combustion component into the combustion chamber.

[0110] Preferably, the cooling channel is configured to cause the first combustion component to flow in the combustion chamber wall in a cooling direction from the throat towards the injection plate.

[0111] Preferably, a first distribution channel is provided, which is arranged in the injection plate and is configured to receive the first combustion component, the first injection channel extending from the first distribution channel to the combustion chamber.

[0112] Preferably, the first distribution channel has an annular shape and is arranged around the axis of symmetry.

[0113] Preferably, a second distribution channel is provided, which is arranged in the injection plate and is configured to receive the second combustion component, the at least one second injection channel extending from the second distribution channel to the combustion chamber.

[0114] Preferably, the second distribution channel has an annular shape and is arranged around the axis of symmetry.

[0115] Preferably, the second distribution channel is radially inside the first distribution channel.

[0116] In one embodiment, a cylindrical slit injection plate is provided on the injection plate and is defined around a symmetry axis and faces the combustion chamber.

[0117] Preferably, a plurality of additional injection channels are provided, each of which has an end portion facing into the cylindrical slit and is configured to inject the first injection component and / or the second injection component into the cylindrical slit in a rotational motion around the axis of symmetry.

[0118] Preferably, each end portion of the further injection channel extends along a respective injection direction tangential to the cylindrical slit.

[0119] Preferably, there is provided a block of material which is produced in one piece.

[0120] Preferably, the diffusion channel, the combustion chamber, the throat, the injection plate, the plurality of first injection channels and the at least one second injection channel are defined in the block.

[0121] Preferably, at least a portion of said thruster is made from said block.

[0122] Preferably, the block is made by additive manufacturing.

[0123] Preferably, the block is made by selective metal laser melting.

[0124] Preferably, the block consists of Made in 718. BRIEF DESCRIPTION OF THE DRAWINGS

[0125] Features and advantages of the present disclosure will emerge from the following detailed description of some embodiments thereof, which are provided by way of non-limiting examples only and will be described with reference to the accompanying drawings, in which:

[0126] Figure 1 A schematic diagram showing a propulsion system for orbital maneuvering according to the present invention is shown;

[0127] Figure 2 shows a perspective view of a thruster for orbital maneuvering according to the present invention, with some parts removed for clarity;

[0128] Figure 3 Shows Figure 2 A cross-sectional view of a thruster for orbital maneuvering at a first cutting plane;

[0129] Figure 4 Shows Figure 2 A cross-sectional view of a thruster for orbital maneuvering at a second cutting plane different from the first cutting plane;

[0130] Figure 5 Shows Figure 4 Enlarged view of the cross section;

[0131] Figure 6 Shows Figure 2 A cross-sectional view of a thruster for orbital maneuvering on a third cutting plane orthogonal to the first cutting plane and the second cutting plane;

[0132] Figure 7 Shows Figure 2 A cross-sectional view of a thruster for orbital maneuvering on a fourth cutting plane that is different from the first cutting plane and the second cutting plane and is orthogonal to the third cutting plane. DETAILED DESCRIPTION

[0133] The propulsion system for orbital maneuvers of the present invention is schematically shown in Figure 1 In the figure, the propulsion system is represented by the reference numeral 1.

[0134] The propulsion system 1 includes a first tank 10 , a second tank 20 , and a thruster 100 for orbital maneuvering, and the thruster 100 for orbital maneuvering is fluidically coupled to the first tank 10 and the second tank 20 .

[0135] The first tank 10 contains a first combustion component, in particular an oxidant. The first combustion component is a self-pressurizing substance, which is at least partially contained in the first tank 10 in liquid form so as to maintain the pressure in the first tank 10 equal to its vapor pressure. In a preferred embodiment, the oxidant is nitrous oxide (N2O).

[0136] The first supply channel 11 is fluidly coupled to the first tank 10 for receiving the first combustion component from the first tank 10 at its vapor pressure. The first supply channel 11 extends from the first tank 10 to the propeller 100.

[0137] A pair of valves (not shown) are arranged along the first supply passage 11 to regulate the flow of the first combustion component.

[0138] The second tank 20 contains a second combustion component, which is configured to react with the first combustion component to produce combustion, and the second combustion component is particularly a fuel. The second combustion component is a self-pressurized substance, which is at least partially contained in the second tank 20 in the form of a liquid, so as to maintain the pressure in the second tank 20 equal to its vapor pressure. In a preferred embodiment, the fuel is propylene.

[0139] The second supply channel 21 is fluidly coupled to the second tank 20 for receiving the second type of propellant from the second tank 20 at its vapor pressure. The second supply channel 21 extends from the first tank 20 to the thruster 100.

[0140] A pair of valves (not shown) are arranged along the second supply passage 21 to regulate the flow of the second combustion component.

[0141] The thruster 100 comprises a block of material 101 made in one piece by additive manufacturing techniques. The preferred material of manufacture of the block 101 is a nickel alloy. In a preferred embodiment, this material is commercially known as 718. This material comprises (expressed in weight percent):

[0142] 50.00-55.00 nickel (with cobalt added);

[0143] 17.00-21.00 Cr;

[0144] 4.75-5.50 niobium (with tantalum added);

[0145] 2.80-3.30 molybdenum;

[0146] 0.65-1.15 titanium;

[0147] 0.20-0.80 aluminum;

[0148] 0-1.00 cobalt;

[0149] 0-0.80 carbon;

[0150] 0-0.35 manganese;

[0151] 0-0.35 silicon;

[0152] 0-0.015 phosphorus;

[0153] 0-0.015 sulfur;

[0154] 0-0.006 boron;

[0155] 0-0.30 copper;

[0156] The remainder is iron for balance.

[0157] The thruster 100 comprises a fixed portion 105 by means of which the thruster 100 can be further mounted in a fixed manner on the orbiter. The fixed portion 105 is defined in a block 101 .

[0158] The propeller 100 includes a combustion chamber 110 , a diffusion passage 111 , and a throat 112 . The throat 112 is arranged between the combustion chamber 110 and the diffusion passage 111 to connect the combustion chamber 110 and the diffusion passage 111 to each other in fluid communication.

[0159] The combustion chamber 110 includes a first region 110 a that is substantially cylindrical and a second region 110 b that is adjacent to the first region 110 a and converges toward a throat 112 .

[0160] The throat 112 forms a sonic choke between the combustion chamber 110 and the diffusion passage 111 , so as to form a supersonic convergent-divergent nozzle together with the second region 110 b and the diffusion passage 111 .

[0161] The diffuser channel 111 is preferably sized to operate in a vacuum and has an inlet at the throat 112 and an outlet 113 on the opposite side relative to the throat 112. The diffuser channel 111 is configured to discharge a supersonic jet from the combustion chamber 110 through the outlet 113 and directed to the opposite side of the fixed part 105. The diffuser channel 111 has an axis of symmetry S.

[0162] The combustion chamber 110 is also preferably symmetrical about the axis of symmetry S. The throat 112 is also preferably symmetrical about the axis of symmetry S. This symmetry is preferably cylindrical.

[0163] A combustion chamber 110 , a diffusion passage 111 and a throat 112 are defined in the block 101 .

[0164] The injection plate 115 faces the combustion chamber 110, in particular the first zone 110a. The injection plate 115 is placed axially between the fixed part 105 and the combustion chamber 110. The combustion chamber 110 is axially defined along the symmetry axis S, at the injection plate 115 located on the opposite side relative to the throat 112. The injection plate 115 has a circular shape, preferably flat, and the injection plate 115 is oriented orthogonal to the symmetry axis S. The injection plate 115 is defined in the block 101.

[0165] The combustion chamber 110 is delimited laterally by a combustion chamber wall 116. The combustion chamber wall 116 extends from the injection plate 115 to the throat 112. The combustion chamber wall 116 is symmetrical with respect to the symmetry axis S. The combustion chamber wall 116 is defined in the block 101.

[0166] The combustion chamber wall 116 forms an edge joint 128 with the injection plate (eg Figure 5 ). The edge angle of such edge joint 128 is between 80° and 100°, more preferably between 85° and 95°, for example about 90°. The edge joint 128 has a circular shape, the center of which is located on the symmetry axis S.

[0167] The diffuser channel 111 is laterally delimited by diffuser channel walls 117. The diffuser channel walls 117 extend from the combustion chamber wall 116 on the opposite side relative to the combustion chamber 110. The diffuser channel walls 117 are symmetrical relative to the symmetry axis S. The diffuser channel walls 117 are defined in the block 101.

[0168] The combustion chamber 110 may be fluidly connected to the first tank 10 .

[0169] The propeller 100 includes a first connecting channel 118 (such as Figure 3 As shown in FIG. 1 , the first connecting passage 118 may be fluidly connected to the first supply passage 11 for receiving the first combustion component from the first tank 10. The first connecting passage 118 is defined in the block 101.

[0170] The first connecting channel 118 includes a first coupling portion 119 at the fixing portion 105 and can be fluidically connected to the first supply channel 11 in a sealed manner.

[0171] An annular throat passage 120 is arranged at the throat 112. The annular throat passage 120 is defined in the block 101. The annular throat passage 120 is arranged around the throat 112 and symmetrically with respect to the axis S of rotation.

[0172] The first connecting passage 118 extends from the first coupling portion 119 along the combustion chamber wall 116 to the annular throat passage 120 .

[0173] A plurality of cooling channels 121 are formed in the combustion chamber wall 116, which are fluidically connected to the first type tank 10 for receiving the first combustion component before introduction into the combustion chamber 110. A plurality of cooling channels 121 are defined in the block 101. The plurality of cooling channels 121 are incorporated and embedded in the combustion chamber wall 116. Within the plurality of cooling channels 121, the first combustion component absorbs heat from the combustion chamber wall 116, heats itself, and prepares for entry into the combustion chamber.

[0174] The first distribution channel 122 is arranged around the injection plate 115. The first distribution channel 122 is defined in the block 101. The first distribution channel 122 has an annular shape and is symmetrical with respect to the axis S of symmetry.

[0175] The cooling passage 121 extends from the annular throat passage 120 to the first distribution passage 122 .

[0176] like Figure 4 , Figure 5 and Figure 6 As shown, the plurality of first injection channels 125 are configured to inject the first combustion component into the combustion chamber 110. The first injection channels 125 extend from the first distribution channel 122 to the combustion chamber 110.

[0177] The number of first injection channels 125 is between 7 and 28, preferably between 10 and 22, even more preferably between 13 and 16, preferably 14.

[0178] The first injection channel 125 may be fluidically connected to the first tank 10. In order to inject the first combustion component into the combustion chamber 110, the first combustion component preferably passes through the first connecting portion 119, the first connecting channel 118, the annular throat channel 120, the cooling channel 121, the first distribution channel 122 and the first injection channel 125 in sequence.

[0179] Preferably, the plurality of first injection channels 125 are of equal length to each other.

[0180] Each first injection channel 125 includes an end portion 126 located at the injection plate 115 and opening in the combustion chamber 110 .

[0181] The end portions 126 are located in an annular region 127 defined on the injection plate 115 and centered on the axis of symmetry S. Each end portion 126 has a respective outlet portion facing the combustion chamber 110 , which outlet portion is completely contained in the annular region 127 .

[0182] The outside of the annular area 127 is defined by an outer circumference CE defined at the combustion chamber wall 116. The outer circumference CE is defined at the edge joint 128 between the injection plate 115 and the combustion chamber wall 116, in particular at the joint thereof. In other words, the radius of the outer circumference CE is substantially equal to the radius of the injection plate 115.

[0183] The interior of the annular area 127 is delimited by an inner circumference CI defined on the injection plate 115 between the edge joint 128 and the axis of symmetry S. Preferably, the radius of the inner circumference CI is at least 75%, preferably at least 85% of the outer radius.

[0184] The end portions 126 are distributed circumferentially in an annular region 127. The end portions 126 are distributed at equal angles relative to the symmetry axis S. In the illustrated embodiment, the end portions 126 are placed substantially adjacent to an edge joint 128 between the injection plate 118 and the combustion chamber wall 116.

[0185] Each first injection channel 125 extends between the injection plate 115 and the combustion chamber 110 without branching. At least the end portion 126 of the first injection channel 125 is straight. Preferably, the first injection channel 125 is completely straight. At least the end portion 126 of the first injection channel 125 has a constant cross section. Preferably, the first injection channel 125 has a constant cross section along its entire length.

[0186] Each end portion 126 has an injection direction d1 coinciding with its main extension axis. The injection direction substantially determines the injection direction of the first combustion component in the combustion chamber 110. In the preferred embodiment shown, the injection direction d1 coincides with the main extension axis of the entire injection channel 125.

[0187] The injection direction d1 has an axial component parallel to the symmetry axis S relative to the symmetry axis S and points toward the combustion chamber 110 .

[0188] The injection direction d1 also has a radial component of the injection direction d1 , which is oriented toward the symmetry axis S. In other words, at least the end portion 126 of each injection channel 125 is inclined toward the symmetry axis S and close to the combustion chamber 110 .

[0189] In the illustrated embodiment, the injection direction d1 of each end portion 126 has no tangential component. In other words, the tangential component of the injection direction d1 is zero.

[0190] The injection directions d1 of the end portion 126 intersect the axis of symmetry S, preferably at a common intersection point of all injection directions d1 .

[0191] Preferably, the injection direction d1 of each end portion 126 is at an angle of between 20° and 60°, even more preferably between 30° and 50°, even more preferably between 35° and 45° relative to the axis of symmetry S. In the embodiment shown, this angle is 40°.

[0192] In an alternative embodiment not shown, the tangential component of the injection direction d1 of one or more end portions 126 is not zero. In this way, the movement of the first combustion component injected into the combustion chamber 110 in rotation about the axis of symmetry S is determined.

[0193] The propeller 100 includes a second connecting channel 130 (such as Figure 7 As shown in FIG. 1 , the second connecting passage 130 may be fluidly connected to the second supply passage 21 for receiving the second combustion component from the second tank 20. The second connecting passage 130 is defined in the block 101.

[0194] The second connection channel 130 includes a second coupling portion 131 at the fixing portion 105 , and can be fluidically connected to the second supply channel 21 in a sealed manner.

[0195] The second distribution channel 132 is arranged at the injection plate 115. The second distribution channel 132 is defined in the block 101. The second distribution channel 132 is annular and symmetrical with respect to the symmetry axis S. The second distribution channel 132 is radially located inside with respect to the first distribution channel 122. Preferably, the second distribution channel 132 has a triangular cross section in a cutting plane containing the symmetry axis S.

[0196] The second connecting channel 130 extends from the second coupling portion 131 to the second distributing channel 132 .

[0197] At least one second injection channel 135 is configured to inject the second combustion component into the combustion chamber 110. Preferably, a plurality of second injection channels 135 are provided. Preferably, the number of second injection channels 135 is between 2 and 8, and even more preferably between 3 and 6. Figure 4 , Figure 5 and Figure 6 In the illustrated embodiment, four second injection channels 135 are provided.

[0198] The second injection channel 135 extends from the second distribution channel 132 to the combustion chamber 110 .

[0199] The second injection channel 135 can be fluidically connected to the second tank 20. In order to inject the second combustion component into the combustion chamber 110, the second combustion component preferably passes through the second coupling portion 131, the second connecting channel 130, the second distribution channel 132 and the second injection channel 135 in sequence.

[0200] Preferably, the plurality of second injection channels 135 are of equal length to each other.

[0201] Each second injection channel 135 includes an end portion 136 located at the injection plate 115 and opening in the combustion chamber 110 .

[0202] The end portion 136 of the second injection channel 135 is radially located between the end portion 126 of the first injection channel 125 and the symmetry axis S. In other words, the end portion 136 of the second injection channel 135 is radially located inside relative to the end portion 126 of the first injection channel 125 .

[0203] Preferably, the end portions 136 of the second injection channel 135 are radially located inside relative to the annular region 127. In other words, each end portion 136 has an outlet portion facing the combustion chamber 110, which is completely arranged radially inside the inner circumference CI of the annular region 127.

[0204] The end portions 136 of the second injection channels 135 are equidistant from the symmetry axis S. The end portions 136 of the second injection channels 135 are circumferentially distributed around the symmetry axis S. The end portions 136 of the second injection channels 135 are distributed at equal angles relative to the symmetry axis S.

[0205] Each second injection channel 135 extends between the injection plate 115 and the combustion chamber 110 without branching. At least the end portion 136 of the second injection channel 135 is straight. At least the end portion 136 of the second injection channel 135 has a constant cross section. Preferably, the second injection channel 135 has a constant cross section along its entire length.

[0206] Each end portion 136 has an injection direction d2 coinciding with its main extension axis. The injection direction d2 substantially determines the injection direction of the second combustion component in the combustion chamber 110. In the preferred embodiment shown, the injection direction d2 coincides with the main extension axis of the entire injection channel 135.

[0207] The injection direction d2 of each end portion 136 has an axial component parallel to the symmetry axis S, pointing to the combustion chamber 110. In a preferred embodiment, the injection direction d2 of each end portion 136 has no radial component. In other words, the radial component of the injection direction d2 of each end portion 136 is zero. Preferably, the injection direction d2 of each end portion 136 has no tangential component. In other words, the tangential component of the injection direction d2 of each end portion 136 is zero. Therefore, the injection direction d2 of the end portion 136 is parallel to the symmetry axis S.

[0208] In an alternative embodiment not shown, the radial component of the injection direction d2 of one or more end portions 136 is not zero. The injection direction d2 of the end portion 136 may intersect the axis of symmetry S, preferably at the common intersection of all injection directions d2. Optionally, the injection direction d2 of the end portion 136 may diffuse relative to the axis of symmetry S. Preferably, the injection direction d2 of each end portion 136 has an angle of convergence or diffusion relative to the axis of symmetry S, which is between 0° and 20°, even more preferably between 0° and 10°, even more preferably between 0° and 5° relative to the axis of symmetry S. In this way, the second combustion component is injected into the combustion chamber 110 in a motion mode moving toward or away from the axis of symmetry S, and its angle relative to the axis of symmetry S is between 0° and 20°, even more preferably between 0° and 10°, even more preferably between 0° and 5°. An angle of 0° indicates that the injection direction d2 has no radial component.

[0209] In an alternative embodiment not shown, the tangential component of the injection direction d2 of one or more end portions 136 is not zero. In this way, the second combustion component is injected into the combustion chamber 110 in a rotational movement about the axis S of symmetry.

[0210] The thruster 100 further includes at least one spark plug (not shown) mounted on the injection plate 115 facing the combustion chamber 110. The spark plug is configured to generate a spark in the combustion chamber 110 to induce combustion between the first combustion component and the second combustion component in the combustion chamber 110.

[0211] The spark plug is mounted on a first seat 140, which is defined at the symmetry axis S on the block 101. The first seat 140 is cylindrical.

[0212] Alternatively, in an embodiment not shown, a plurality of spark plugs may be mounted on the injection plate, the spark plugs being equidistant relative to the axis S of symmetry.

[0213] The thruster 100 further includes a pressure sensor, not shown, mounted on the injection plate 115, facing the combustion chamber 110. The pressure sensor is configured to measure the pressure in the combustion chamber 110 in order to control the combustion.

[0214] The signal of the pressure sensor can be used to feedback adjust the flow rate of the first combustion component from the first tank 10 and the flow rate of the second combustion component from the second tank 20 .

[0215] The pressure sensor is mounted on a second seat 141 defined in the block 101 . The second seat 141 is defined radially between the end portion 136 of the second injection channel 135 and the axis of symmetry S. The second seat 141 is defined radially outside the first seat 140 .

[0216] In an embodiment not shown, a cylindrical slot is provided which extends around the axis of symmetry S, in particular around the first seat 140 for the spark plug and faces the combustion chamber 110 .

[0217] A plurality of additional injection channels (not shown) are arranged to be in fluid communication with the cylindrical slit so that their respective end portions are tangent to the slit itself. A plurality of additional injection channels are configured to inject the first combustion component and the second combustion component into the cylindrical slit so as to impart a circular motion thereto. Thus, the first combustion component and the second combustion component can escape from the slit, partially or completely mix, and form a vortex around the axis of symmetry S in the combustion chamber 110. The vortex is introduced into the combustion chamber 110 at the spark plug. This helps to ignite the mixture of the first combustion component and the second component by the spark plug, while keeping the combustion away from the spark plug to avoid subjecting it to excessive temperatures.

[0218] The above-mentioned propulsion system 1 including the thruster 100 can be fixedly mounted on an orbital transport vehicle (not shown) configured to move on a given orbit to release a payload. The propulsion system 1 can be activated, for example, to transfer the orbital transport vehicle from a release orbit, in which the orbital transport vehicle is released by a space vehicle, to a deployment orbit, in which the payload is deployed by the orbital transport vehicle.

[0219] The propulsion system 1 may also be activated to transfer the orbital transportation vehicle from a first deployment orbit, in which a first payload is deployed by the orbital transportation vehicle, to a second deployment orbit, in which a second payload is deployed by the orbital transportation vehicle.

[0220] The propulsion system 1 may still be activated to transfer the orbital transport vehicle from the deployment orbit in which the payload is deployed by the orbital transport vehicle to the re-entry orbit in which the orbital transport vehicle re-enters the atmosphere.

Claims

1. A thruster (100) for orbital maneuvers, comprising: A diffusion channel (111), wherein the diffusion channel (111) has an axial symmetry axis (S); A combustion chamber (110), the combustion chamber (110) being laterally delimited by a combustion chamber wall (116); a throat (112), the throat (112) being disposed between the combustion chamber (110) and the diffusion passage (111) so as to allow the combustion chamber (110) to be in fluid communication with the diffusion passage (111); an injection plate (115), the injection plate (115) facing the combustion chamber (110); a plurality of first injection channels (125) in fluid communication with the combustion chamber (110), the plurality of first injection channels (125) being configured to inject a first combustion component into the combustion chamber (110), wherein each first injection channel (125) of the plurality of first injection channels (125) comprises an end portion (126) located at the injection plate (115) and opening in the combustion chamber (110); at least one second injection channel (135), the at least one second injection channel (135) being in fluid communication with the combustion chamber (110), the at least one second injection channel (135) being configured to inject a second combustion component into the combustion chamber (110), and the at least one second injection channel (135) comprising an end portion (136) located at the injection plate (115) and opening in the combustion chamber (110); wherein an end portion (126) of a first injection channel (125) among the plurality of first injection channels (125) is substantially adjacent to an edge joint (128) between the injection plate (118) and the combustion chamber wall (116), the end portion (126) of the first injection channel (125) is located within an annular region (127), the center of the annular region (127) intersects the symmetry axis (S), and the end portion (126) of the first injection channel (125) extends along a respective injection direction (d1); wherein an end portion (136) of the at least one second injection channel (135) is radially located between the symmetry axis (S) and an end portion (126) of a first injection channel (125) among the plurality of first injection channels (125); The injection direction (d1) of the end portion (126) of each of the plurality of first injection channels (125) has an axial component pointing to the combustion chamber (110) and a radial component pointing to the axis of symmetry (S).

2. The thruster (100) for orbital maneuvers according to claim 1 comprises a block (101) made of a material that can be made into a single piece, and the diffusion channel (111), the combustion chamber (110), the throat (112), the injection plate (115), a plurality of first injection channels (125) and at least one second injection channel (135) are defined in the block (101).

3. The thruster (100) for orbital maneuvers according to claim 1 or 2, wherein: The annular region (127) is defined by concentric inner circumference (CI) and outer circumference (CE); wherein the outer circumference (CE) is located at the edge joint (128); And wherein, the radius of the inner circumference (CI) of the annular area (127) is equal to or greater than 75% of the radius of the outer circumference (CE), and more preferably equal to or greater than 85% of the radius of the outer circumference (CE).

4. The thruster (100) for orbital maneuvers according to any one of the preceding claims, wherein: The annular area (127) is delimited by a concentric inner circumference (CI) and an outer circumference (CE), and an end portion (136) of the at least one second injection channel (135) is located between the inner circumference (CI) and the axis of symmetry (S).

5. The thruster (100) for orbital maneuvers according to any one of the preceding claims, wherein: The injection direction (d1) of the end portion (126) of each first injection channel (125) does not include any component along the tangential direction.

6. The thruster (100) for orbital maneuvers according to any one of the preceding claims, wherein: The end portion (136) of the at least one second injection channel (135) extends along the respective injection direction (d2), and wherein the angle formed by the projection of the injection direction (d2) of the end portion (136) of the at least one second injection channel (135) on the plane containing the symmetry axis (S) relative to the symmetry axis (S) is between 0° and 20°, preferably between 0° and 10°, more preferably between 0° and 5°, for example 0°.

7. The thruster (100) for orbital maneuvers according to any one of the preceding claims, wherein: The angle formed by the projection of the injection direction (d1) of the end portion (126) of each of the multiple first injection channels (125) on the plane containing the symmetry axis (S) relative to the symmetry axis (S) is between 20° and 60°, preferably between 30° and 50°, more preferably between 35° and 45°, for example about 40°.

8. The thruster (100) for orbital maneuvers according to any one of claims 1 to 6, wherein: The angle formed by the projection of the injection direction (d1) of the end portion (126) of each of the plurality of first injection channels (125) on the plane containing the symmetry axis (S) relative to the symmetry axis (S) is between 30° and 60°.

9. The thruster (100) for orbital maneuvers according to any one of claims 1 to 6, wherein: The angle formed by the projection of the injection direction (d1) of the end portion (126) of each of the plurality of first injection channels (125) on the plane containing the symmetry axis (S) relative to the symmetry axis (S) is between 35° and 60°.

10. The thruster (100) for orbital maneuvers according to any of the preceding claims, comprising a plurality of second injection channels (135).

11. The thruster (100) for orbital maneuvers according to any one of the preceding claims, wherein: The number of the first injection channels (125) in the plurality of first injection channels (125) is between 7 and 28, preferably between 10 and 22, even more preferably between 13 and 16, preferably 14.

12. A thruster (100) for orbital maneuvers according to any one of the preceding claims, comprising a spark plug, which is mounted on the injection plate (115) at the axis of symmetry (S) and is configured to generate sparks in the combustion chamber (110).

13. The thruster (100) for orbital maneuvers according to any one of the preceding claims, wherein: The end portion (126) of the first injection channel (125) is straight, has no branches, and has a constant cross section along the injection direction (d1).

14. The thruster (100) for orbital maneuvers according to any one of the preceding claims, wherein: The angle formed by the injection direction (d1) of the end portion (126) of each of the plurality of first injection channels (125) relative to a plane orthogonal to the symmetry axis (S) is between 30° and 60°.

15. The thruster (100) for orbital maneuvers according to any one of claims 1 to 3, wherein: The angle formed by the injection direction (d1) of the end portion (126) of each of the plurality of first injection channels (125) relative to a plane orthogonal to the symmetry axis (S) is between 30° and 55°.

16. The thruster (100) for orbital maneuvers according to any one of the preceding claims, wherein: The injection direction (d2) of the end portion (136) of the at least one second injection channel (135) does not include any component in the radial direction.

17. A propulsion system for orbital maneuvers (1), comprising: A thruster (100) for orbital manoeuvring according to any one of the preceding claims; a first tank (10) containing a first type of propellant, the first tank (10) being fluidly coupled to a first injection channel (125) of the plurality of first injection channels (125); a second tank (20) containing a second type of propellant, the second tank (20) being fluidly coupled to the at least one second injection channel (135); Wherein, the first combustion component and the second combustion component are self-pressurizing.

18. Orbital transport vehicle comprising a thruster (100) for orbital manoeuvring according to any one of the preceding claims 1 to 16.

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