Orbital control thrusters, orbital control propulsion systems, and orbital transfer vehicles

Abstract

The present invention provides an orbital control thruster, an orbital control propulsion system including the orbital control thruster, and an orbital transfer vehicle including the orbital control thruster. A thruster (100) for orbital maneuvering includes a divergent channel (111) having an axial axis of symmetry (S), a combustion chamber (110), a throat (112) interposed between the combustion chamber (110) and the divergent channel (111), an injection plate (115) facing the combustion chamber (110), a plurality of first injection passages (125) in fluid communication with the combustion chamber (110), each having an end (126) opening into the combustion chamber (110), and at least one second injection passage (135) in fluid communication with the combustion chamber (110), each having an end (136) opening into the combustion chamber (110). The ends (126) of the first injection passages (125) are located within an annular region (127) and extend along respective injection directions (d1). The end (136) of the at least one second injection passage (135) is located radially between the axis of symmetry (S) and the end (126) of the first injection passage (125). The injection direction d1 of each end 126 of the first injection flow passage 125 has an axial component directed toward the combustion chamber 110 and a radial component directed toward the axis of symmetry S.

Description

[Technical Field]

[0001] The present invention relates to a thruster for orbital control, a propulsion system for orbital control equipped with a thruster for orbital control, and an orbital transfer vehicle equipped with a thruster for orbital control. [Background technology]

[0002] An orbital transfer vehicle is a spacecraft capable of receiving, transporting, and releasing a payload. Such an orbital transfer vehicle can be transported into space on a launch vehicle and includes a dedicated propulsion system for performing orbital maneuvers, such as orbital changes, after separation from the launch vehicle.

[0003] The propulsion system of an orbital transfer vehicle includes thrusters for orbital control and propellant tanks that supply fuel and oxidizer (i.e., combustion promoter) to the thrusters.

[0004] Typically, an orbital maneuver thruster includes a combustion chamber in which propellant combustion occurs, an injection plate that injects the propellant into the combustion chamber, a throat that communicates with the combustion chamber and through which the burning mass accelerated in the combustion chamber reaches the sound barrier speed, a divergent channel (i.e., a supersonic nozzle) that extends from the throat opposite the combustion chamber and through which the burning mass accelerates beyond the sound barrier speed, and multiple cooling channels that cool the walls of the combustion chamber. The orbital maneuver thruster also includes a supply system that delivers propellant to the combustion chamber, a spark plug that ignites the combustion, and valves and sensors that control the combustion.

[0005] In a bipropellant orbital maneuver propulsion system, the propellants can be fed to the combustion chamber under pressure using a feed pump or by a pressurizer such as nitrogen or hydrogen that is used to maintain the fuel and oxidizer tanks at the appropriate pressurization conditions to allow the fuel and oxidizer to enter the combustion chamber at the appropriate flow rate and pressure.

[0006] Typically, orbital maneuver thrusters are turned on for a fixed period during the burn when each propellant charge is introduced into the combustion chamber and burned, accelerating the resulting burned mass and expelling it from the throat and divergent channel. The orbital maneuver thrusters generate a predetermined force (to generate a specific impulse) for a predetermined time required to perform a predetermined phase of the mission, for example, to modify the orbital parameters of the orbital maneuver vehicle. Summary of the Invention [Problem to be solved by the invention]

[0007] The present applicant has recognized that a small, lightweight, and low-cost inter-orbital transfer vehicle is needed for the transport and use of small payloads, such as picosatellites, cubesats, and other small satellites, preferably without motors. The present applicant has determined that to perform such a mission, a thruster capable of generating a thrust of 5-22 N continuously for at least 5 seconds, having a specific impulse greater than 285 seconds, and capable of multiple follow-on firings, with an overall length (in the thrust direction) of less than 250 mm and a total weight of less than 2 kg, would be required. The present applicant has recognized that a propulsion system for this type of inter-orbital transfer vehicle would desirably be highly versatile, lightweight, simple, inexpensive, and easy to operate.

[0008] The applicant has realized that the manufacture of a thruster for orbital control of the type summarized above can be carried out by additive manufacturing methods such as selective metal laser melting (laser powder bed fusion (L-PBF)), allowing the majority of the thruster to be produced at low cost as a monolithic block containing the combustion chamber, the supersonic nozzle, and at least some of the supply system. The applicant has further recognized that the injection plate, the cooling channels for the combustion chamber, and at least some of the supply channels that carry propellant to the injection plate can be integrated into the monolithic block. This manufacturing method requires that the monolithic block be made of a material suitable for additive manufacturing.

[0009] The applicant has discovered that handling propulsion system propellants during launch preparation typically requires extremely cumbersome safety procedures, and the use of the propellants themselves carries significant risks, resulting in increased deployment time and costs for the interorbital vehicle and its payload.

[0010] Applicant has realized that the use of safe, easy-to-handle propellants greatly simplifies pre-launch operations, reducing the time and cost involved.

[0011] The applicant has realized that the risks associated with the use of propellants can be further reduced and the safety procedures less complicated by using so-called "green propellants" which are less toxic and safer than conventional propellants.

[0012] The applicant has realised that if such a "green propellant" were self-pressurising, it could be stored in liquid form in each tank and evaporated when used, maintaining a constant pressure within the tank during consumption, eliminating the need for additional tanks for pressurising agent, thereby reducing the weight of the propulsion system.

[0013] Applicant has recognized that there are limited options for self-pressurizing green propellants. Specifically, Applicant has recognized that the only type of oxidizer that can be employed inherently possesses the above characteristics without the need for pre-decomposition in a decomposition system (e.g., electrolysis, pre-heating, etc.) or dissolving another component in it to make it self-pressurized is nitrous oxide (NO).

[0014] Applicant has determined that the use of nitrous oxide as an oxidizer and a suitable fuel (e.g., propylene) that is inherently green and self-pressurized for orbital maneuver thrust can result in flame front temperatures in excess of 2800°C.

[0015] Applicant has determined that even if an orbital maneuver thruster were additively manufactured using currently available materials that are best suited to withstand high temperatures, the flame front temperatures encountered with nitrous oxide would likely melt or degrade the combustion chamber, preventing the ignition period required to complete the mission phase. [Means for solving the problem]

[0016] Applicant has realized that by eliminating direct exposure of the combustion chamber walls to the flame front, the combustion chamber will have more time to deteriorate, allowing thrust to be ignited for the duration required to accomplish the mission phase.

[0017] Applicant has discovered that by introducing fuel and oxidizer into a combustion chamber in a stoichiometric ratio that is not useful for generating combustion and then into a space within the combustion chamber away from the combustion chamber wall (i.e., not directly touching the combustion chamber wall in any way), where they mix in a stoichiometric ratio that is useful for combustion, the wall of the combustion chamber is not directly impinged by the flame front.

[0018] Accordingly, in a first aspect, the present invention provides a combustion chamber comprising a divergent channel having an axial axis of symmetry, a combustion chamber laterally defined by a combustion chamber wall, a throat portion interposed between the combustion chamber and the divergent channel and fluidly connecting the combustion chamber with the divergent channel, an injection plate facing the combustion chamber, a plurality of first injection passages in fluid communication with the combustion chamber and configured to inject a first combustion component into the combustion chamber, each of the first injection passages having an end in the injection plate and opening into the combustion chamber, and a plurality of first injection passages in fluid communication with the combustion chamber and configured to inject a second combustion component into the combustion chamber, and at least one second injection flow path, the end of which is provided in the injection plate and opens into the combustion chamber, wherein the ends of the plurality of first injection flow paths are located approximately near an edge junction between the injection plate and the combustion chamber wall within an annular region whose center intersects with the axis of symmetry and extend along respective injection directions, the end of the at least one second injection flow path being located radially between the axis of symmetry and the ends of the plurality of first injection flow paths, and the injection direction of the end of each first injection flow path has an axial component directed toward the combustion chamber and a radial component directed toward the axis of symmetry.

[0019] In a second aspect, the present invention relates to a propulsion system for orbital maneuver, comprising: the thruster for orbital maneuver according to the first aspect; a first tank containing a first type of propellant and fluidly coupled to the plurality of first ejection passages; and a second tank containing a second type of propellant and fluidly coupled to the at least one second ejection passage, wherein the first combustion component and the second combustion component are self-pressurized.

[0020] In a third aspect, the present invention relates to an orbital transfer vehicle equipped with the thruster for orbital maneuver according to the first aspect.

[0021] The plurality of first injection passages enable the first combustion component to be injected into the combustion chamber along a substantially annular path.

[0022] The at least one second injection passage allows the second combustion component to be injected into the combustion chamber inside the annular passage of the first combustion component.

[0023] The applicant believes that this will result in the second combustion component not immediately mixing with the first combustion component, but rather remaining separate from the first combustion component, at least upon introduction, within the first combustion component.

[0024] The injection direction of the end of each first injection channel has an axial component directed towards the combustion chamber and a radial component directed towards the axis of symmetry, so that the jet of first combustion component entering the combustion chamber is directed towards the axis of symmetry, thereby causing the first combustion component injected into the combustion chamber to converge towards the axis of symmetry and towards the second combustion component.

[0025] The applicant believes that this will cause combustion to occur in the central region of the combustion chamber, and that the first combustion component, which has not yet mixed with the second combustion component, will flow through at least the initial part of the combustion chamber between the wall of the combustion chamber and the central region, and will adhere to the wall of the combustion chamber, interposing itself between the wall of the combustion chamber and the flame front, preventing the flame front from directly contacting the wall of the combustion chamber.

[0026] The divergent channels have an axis of symmetry about which expressions such as "axial," "axially," "radially," "radially," "radially inner," "radially outer," "circumferential," "circumferentially," "tangential," and "tangential" are used in this specification and the appended claims.

[0027] The terms "radial" and "radially inner / outer" refer to directions perpendicular to the axis of symmetry of the divergent channel, and the terms "axial" and "axially inner / outer" refer to directions parallel to the axis of symmetry.

[0028] The term "radially innermost" refers to a position closer to the axis of symmetry of the divergent channel, and the term "radially outermost" refers to a position further from the axis of symmetry.

[0029] The terms "circumferential" and "in" refer to a direction along the circumference of a circle lying on a plane perpendicular to the axis of symmetry of the divergent channel and whose center intersects the axis of symmetry of the divergent channel.

[0030] The terms "tangential" and "tangential to" are used to refer to a direction tangent to a circumference that lies in a plane perpendicular to the axis of symmetry of the divergent channel and whose center intersects the axis of symmetry of the divergent channel. This tangential direction is contained in the same plane as the circumference.

[0031] The expressions "first combustion component" and "second combustion component" in this specification and the appended claims should be understood to mean substances that, when properly mixed together, will undergo an exothermic oxidation-reduction reaction when activation energy is applied to the mixture.

[0032] The expression "injector plate" is understood to mean the part of the thruster facing the combustion chamber on the side opposite the throat and through which the first and second combustion components are injected into the combustion chamber.

[0033] In one or more of the aforementioned aspects, the present disclosure can be implemented according to one or more of the following embodiments, which may be combined with each other.

[0034] Preferably, the first combustion component is an oxidizer.

[0035] Preferably, the first combustion component comprises nitrous oxide (N2O).

[0036] Preferably, the first combustion component is nitrous oxide.

[0037] Preferably, the second combustion component is a fuel.

[0038] Preferably, the second combustion component comprises propylene (C3H6).

[0039] Preferably, the second combustion component is propylene.

[0040] Preferably, the end of the first ejection channel is straight.

[0041] Preferably, each end of the first jetting channels is an extension of a single channel.

[0042] Preferably, the end of the first injection channel is free of any branching.

[0043] Preferably, the end of the first injection channel has a constant cross section along the injection direction.

[0044] Preferably, the first ejection channel is straight.

[0045] Preferably, each of the first jetting channels is a single channel.

[0046] Preferably, the first injection channel has no branching portion.

[0047] Preferably, the first ejection flow path has a constant cross section along the ejection direction.

[0048] Preferably, the diameter of said end 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.

[0049] Preferably, the end of the at least one second ejection channel is straight.

[0050] Preferably, said end of said at least one second ejection channel is an extension of a single channel.

[0051] Preferably, the end of the at least one second injection channel is free of branches.

[0052] Preferably, said end of said at least one second injection channel has a constant cross section along the injection direction.

[0053] Preferably, the at least one second ejection channel is straight.

[0054] Preferably, the at least one second jetting channel is a single channel.

[0055] Preferably, the at least one second jetting channel is free of branches.

[0056] Preferably, said at least one second injection channel has a constant cross section along said injection direction.

[0057] Preferably, the diameter of the end of each second injection passage is 0.4 mm to 0.8 mm, more preferably 0.5 mm to 0.7 mm, even more preferably 0.55 mm to 0.65 mm, for example 0.6 mm. Preferably, the combustion chamber wall extends from the injection plate to the throat portion.

[0058] Preferably, the combustion chamber wall and the injection plate are joined to one another at an edge joint.

[0059] Preferably, at the edge joint, the injection plate and the combustion chamber wall form an edge angle.

[0060] Preferably, the edge angle is between 80° and 100°, more preferably between 85° and 95°, for example about 90°.

[0061] Preferably, the annular region is defined by an inner periphery and an outer periphery.

[0062] Preferably, the inner and outer circumferences are concentric.

[0063] Preferably, said periphery is located at said edge joint.

[0064] Preferably, the radius of the inner circumference of the annular region is 75% or more, more preferably 85% or more of the radius of the outer circumference.

[0065] Preferably, the first combustion component is injected near the combustion chamber wall.

[0066] Preferably, the end of the at least one second injection channel is located between the inner circumference of the annular region and the axis of symmetry.

[0067] The end of the at least one second injection passage is located outside the annular region defined between the inner and outer peripheries.

[0068] This causes the first combustion component to be injected away from the combustion chamber wall.

[0069] Preferably, the jet direction at the end of each first jet channel has no tangential component.

[0070] Preferably, said first combustion component is injected towards a point or area located along said axis of symmetry.

[0071] Preferably, the first combustion component is injected into the combustion chamber from the end of the first injection passage in a jet that diverges conically.

[0072] This helps to keep the second combustion component inside the cone of dispersion of the first combustion component.

[0073] Preferably, the ends of the at least one second injection channel extend along a respective injection direction, the injection direction having an axial component towards the combustion chamber.

[0074] In one embodiment, the injection direction of the end of the at least one second injection channel has no radial component.

[0075] In another embodiment, the injection direction of the end of the at least one second injection channel has a radial component directed towards the axis of symmetry.

[0076] In yet another embodiment, the injection direction of the end of the at least one second injection channel has a radial component directed away from the axis of symmetry, preferably having no tangential component.

[0077] This causes the second combustion component to be injected parallel to the axis of symmetry directly towards the central region of the combustion chamber where it is desired to combust.

[0078] Alternatively, the jet direction at the end of the at least one second jet channel has a component directed along a tangential direction.

[0079] This causes the second combustion component to be injected into the combustion chamber with a rotational motion about the axis of symmetry, which helps to keep combustion away from the injection plate and components attached to the injection plate.

[0080] Preferably, the projection of the injection direction of each of the ends of the plurality of first injection channels onto a plane containing the axis of symmetry forms an angle of 20° to 60°, preferably 30° to 50°, more preferably 35° to 45°, for example about 40°, with the axis of symmetry. This angle can be calculated as the arctangent of the ratio of the radial component to the axial component of the injection direction.

[0081] Preferably, the injection direction of the end of each of the plurality of first injection flow paths forms an angle of 30° to 70°, preferably 40° to 60°, more preferably 45° to 55°, for example approximately 50°, with respect to a plane perpendicular to the axis of symmetry.

[0082] Applicants have found that these angle values ​​are highly effective in keeping combustion away from the walls of the combustion chamber and contained within the central region of the combustion chamber.

[0083] Preferably, the number of said plurality of first injection channels is 7-28, preferably 10-22, even more preferably 13-16, for example 14.

[0084] Preferably, the projection of the injection direction of the end of the at least one second injection channel onto a plane containing the axis of symmetry forms an angle with the axis of symmetry of 0° to 20°, preferably 0° to 10°, more preferably 0° to 5°, which angle can be calculated as the arctangent of the ratio between the radial component (towards or away from the axis of symmetry) and the axial component of the second injection direction.

[0085] Preferably, the jetting direction of the end of the at least one second jetting channel forms an angle of 70° to 90°, preferably 80° to 90°, more preferably 85° to 90° with respect to a plane perpendicular to the axis of symmetry. If the angle is other than 90°, the jetting direction of the end of the at least one second jetting channel is incident on the axis of symmetry or diverges from the axis of symmetry.

[0086] Preferably, a plurality of second injection channels are provided.

[0087] Preferably, the number of the plurality of second injection channels is 2 to 8, preferably 3 to 6, for example 4.

[0088] Preferably, the second ejection channel has a smaller cross section than the first ejection channel.

[0089] Preferably, said second injection channels are distributed around a circumference about said axis of symmetry, preferably equidistant from one another.

[0090] Preferably, a spark plug is provided mounted on the injection plate at the axis of symmetry and configured to generate a spark within the combustion chamber.

[0091] Alternatively, a plurality of spark plugs may be mounted on the injection plate and equidistant about the axis of symmetry to generate sparks within the combustion chamber.

[0092] Preferably, a plurality of cooling passages are provided arranged around the periphery of the combustion chamber and in fluid communication with the first injection passage.

[0093] Preferably, the number of the plurality of cooling channels is 14 to 18, more preferably 15 to 17, for example 16.

[0094] The cooling passages also cool the combustion chamber walls and preheat the first combustion component.

[0095] Preferably, the plurality of cooling channels are integrated into the combustion chamber wall.

[0096] Preferably, the plurality of cooling passages are fluidly connectable to a first tank and configured to receive the first combustion component from the first tank prior to introducing the first combustion component into the combustion chamber.

[0097] Preferably, the cooling passage is configured to cause the first combustion component to flow inside the combustion chamber wall in a cooling direction from the throat portion toward the injection plate.

[0098] Preferably, a first distribution passage is provided arranged in the injection plate and configured to receive the first combustion component, the first injection passage extending from the first distribution passage to the combustion chamber.

[0099] Preferably, said first distribution channel is annular in shape and positioned around said axis of symmetry.

[0100] Preferably, a second distribution passage is provided arranged in the injection plate and configured to receive the second combustion component, and the at least one second injection passage extends from the second distribution passage to the combustion chamber.

[0101] Preferably, said second distribution channel is annular in shape and positioned around said axis of symmetry.

[0102] Preferably, the second distribution channel is radially inward of the first distribution channel.

[0103] In one embodiment, the injection plate is provided with a cylindrical slit formed around the axis of symmetry and facing the combustion chamber.

[0104] Preferably, a plurality of further injection channels are provided, each with an end facing the inside of the cylindrical slit and configured to inject the first injection component and / or the second injection component into the inside of the cylindrical slit with a rotational movement about the axis of symmetry.

[0105] Preferably, each end of the further injection channels extends along a respective injection direction tangent to the cylindrical slit.

[0106] Preferably, a block of material formed as a single piece is provided.

[0107] Preferably, the divergent channel, the combustion chamber, the throat portion, the injection plate, the plurality of first injection passages, and at least one second injection passage are formed in the block.

[0108] Preferably, at least a part of the thruster is formed by the block.

[0109] Preferably, the block is made by additive manufacturing.

[0110] Preferably, the block is made by selective metal laser melting.

[0111] Preferably, the blocks are made of Inconel® 718.

[0112] The features and advantages of the present disclosure will become apparent from the following detailed description of some embodiments of the disclosure, given by way of non-limiting example only, with reference to the accompanying drawings, in which: [Brief explanation of the drawings]

[0113] [Figure 1] 1 is a schematic diagram of a propulsion system for orbital control according to the present invention; [Figure 2] 1 is a perspective view of a thruster for orbital maneuvering according to the present invention, with some parts omitted for clarity; FIG. [Figure 3] 3 is a cross-sectional view of the thruster for orbital control of FIG. 2 taken along a first cutting plane. [Figure 4] 3 is a cross-sectional view of the thruster for orbital control of FIG. 2 taken along a second cross-sectional surface different from the first cross-sectional surface. [Figure 5] FIG. 5 is an enlarged view of the cross section of FIG. [Figure 6] 3 is a cross-sectional view of the thruster for orbital control of FIG. 2 taken along a third cut plane perpendicular to the first cut plane and the second cut plane. [Figure 7]3 is a cross-sectional view of the thruster for orbital control of FIG. 2 taken along a fourth cut plane different from the first cut plane and the second cut plane and perpendicular to the third cut plane. DETAILED DESCRIPTION OF THE INVENTION

[0114] FIG. 1 shows diagrammatically a propulsion system for orbital control, the object of the present invention, designated by the reference numeral 1.

[0115] The propulsion system 1 includes a first tank 10, a second tank 20, and a thruster 100 for orbital control fluidically coupled to the first tank 10 and the second tank 20.

[0116] The first tank 10 contains a first combustion component, specifically an oxidizer, which is a self-pressurizing substance contained in the first tank 10 in at least partially liquid form so as to maintain a pressure within the first tank 10 equal to its vapor pressure. In a preferred embodiment, the oxidizer is nitrous oxide (NO).

[0117] A first supply line 11 is fluidly connected to the first tank 10 for receiving the first combustion component from the first tank 10 at the vapor pressure of the first combustion component. The first supply line 11 extends from the first tank 10 to the thruster 100.

[0118] A pair of valves (not shown) are positioned along the first supply flow path 11 to regulate the flow rate of the first combustion component.

[0119] The second tank 20 contains a second combustion component, specifically a fuel, configured to react with the first combustion component to generate combustion. The second combustion component is a self-pressurizing substance contained in the second tank 20 in at least a partially liquid form so as to maintain a pressure within the second tank 20 equal to its vapor pressure. In a preferred embodiment, the fuel is propylene.

[0120] A second supply line 21 is fluidly connected to the second tank 20 for receiving the second type of propellant from the second tank 20 at the vapor pressure of the second type of propellant. The second supply line 21 extends from the first tank 20 to the thruster 100.

[0121] A pair of valves (not shown) are positioned along the second supply passage 21 to regulate the flow rate of the second combustion component.

[0122] The thruster 100 comprises a block of material 101 formed as a single piece by additive manufacturing. A preferred material for making the block 101 is a nickel alloy, which in the preferred embodiment is known by the trade name INCONEL® 718. The material contains: 50.00 to 55.00 weight percent nickel with cobalt; 17.00 to 21.00 weight percent chromium; 4.75 to 5.50 weight percent niobium with tantalum; 2.80 to 3.30 weight percent molybdenum; 0.65 to 1.15 weight percent titanium; 0.20 to 0.80 weight percent aluminum; 0 to 1.00 weight percent cobalt; 0 to 0.80 weight percent carbon; 0 to 0.35 weight percent manganese; 0 to 0.35 weight percent silicon; 0 to 0.015 weight percent phosphorus; 0 to 0.015 weight percent sulfur; 0 to 0.006 weight percent boron; 0 to 0.30 weight percent copper; and the balance iron.

[0123] The thruster 100 has a fixed portion 105 that can fixedly attach the thruster 100 to an orbital spacecraft. The fixed portion 105 is formed in the block 101.

[0124] The thruster 100 includes a combustion chamber 110, a divergent channel 111, and a throat 112 disposed between the combustion chamber 110 and the divergent channel 111 to provide fluid communication between the combustion chamber 110 and the divergent channel 111.

[0125] Combustion chamber 110 has a generally cylindrical first region 110a and a second region 110b adjacent to first region 110a that converges toward a throat 112.

[0126] The throat section 112 forms a sonic choke between the combustion chamber 110 and the divergent channel 111, and together with the second region 110b and the divergent channel 111, constitute a convergent-divergent supersonic nozzle.

[0127] Divergent channel 111 is preferably sized to operate in a vacuum and has an inlet at throat 112 and an outlet 113 opposite throat 112. Divergent channel 111 is configured to discharge the supersonic jet from combustion chamber 110 and directed away from stationary portion 105 through outlet 113. Divergent channel 111 has an axis of symmetry S.

[0128] Preferably, the combustion chamber 110 is also symmetrical about an axis of symmetry S. Preferably, the throat 112 is also symmetrical about an axis of symmetry S. Preferably, such symmetry is cylindrical.

[0129] The combustion chamber 110 , the divergent channel 111 and the throat portion 112 are formed in the block 101 .

[0130] The injection plate 115 faces the combustion chamber 110, specifically the first region 110a. The injection plate 115 is located axially between the fixed part 105 and the combustion chamber 110. The combustion chamber 110 is defined axially on the opposite side along the axis of symmetry S from the throat part 112 by the injection plate 115. The injection plate 115 has a circular shape, but is preferably flat and oriented perpendicular to the axis of symmetry S. The injection plate 115 is formed in the block 101.

[0131] The combustion chamber 110 is laterally defined by a combustion chamber wall 116. The combustion chamber wall 116 extends from the injection plate 115 to the throat portion 112. The combustion chamber wall 116 is symmetrical with respect to an axis of symmetry S. The combustion chamber wall 116 is formed in the block 101.

[0132] The combustion chamber wall 116 forms an edge joint 128 (shown in FIG. 5) with the injection plate. 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 is circular in shape, with its center located on the axis of symmetry S.

[0133] The divergent channel 111 is laterally bounded by a divergent channel wall 117. The divergent channel wall 117 extends from the combustion chamber wall 116 on the side opposite the combustion chamber 110. The divergent channel wall 117 is symmetrical about an axis of symmetry S. The divergent channel wall 117 is formed in the block 101.

[0134] The combustion chamber 110 is fluidly connectable to the first tank 10 .

[0135] The thruster 100 has a first connecting passage 118, shown in Figure 3, fluidly connectable to the first supply passage 11 and receiving the first combustion component from the first tank 10 from the first supply passage 11. The first connecting passage 118 is formed in the block 101.

[0136] The first connecting channel 118 has a first binding site 119 in the fixing part 105 that can be fluidly tightly connected to the first supply channel 11 .

[0137] An annular throat flow path 120 is arranged in the throat 112. The annular throat flow path 120 is formed in the block 101. The annular throat flow path 120 is arranged symmetrically around the throat 112 with respect to the rotation axis S.

[0138] A first connecting passage 118 extends from the first junction site 119 along the combustion chamber wall 116 to an annular throat passage 120 .

[0139] The combustion chamber wall 116 is formed with a plurality of cooling channels 121 that are fluidly connectable to the first type of tank 10 and that receive the first combustion component prior to introduction into the combustion chamber 110. The plurality of cooling channels 121 are formed in the block 101. The plurality of cooling channels 121 are embedded in and integrated into the combustion chamber wall 116. The first combustion component within the plurality of cooling channels 121 removes heat from the combustion chamber wall 116 and warms itself in preparation for insertion into the combustion chamber.

[0140] A first distribution flow path 122 is arranged around the injection plate 115. The first distribution flow path 122 is formed in the block 101. The first distribution flow path 122 has an annular shape and is symmetrical with respect to the axis of symmetry S.

[0141] The cooling passage 121 extends from the annular throat passage 120 to the first distribution passage 122 .

[0142] 4, 5 and 6 are configured to inject the first combustion component into the combustion chamber 110. The first injection passages 125 extend from the first distribution passages 122 into the combustion chamber 110.

[0143] The number of first injection channels 125 is 7 to 28, preferably 10 to 22, even more preferably 13 to 16, and preferably 14.

[0144] The first injection passage 125 is fluidly connectable to the first tank 10. When the first combustion component is injected into the combustion chamber 110, the first combustion component passes through the first joining portion 119, the first connecting passage 118, the annular throat passage 120, the cooling passage 121, the first distribution passage 122, and the first injection passage 125, preferably in this order.

[0145] Preferably, the first ejection channels 125 have equal lengths between them.

[0146] Each end 126 of the first injection passages 125 is provided in the injection plate 115 and opens into the combustion chamber 110 .

[0147] The ends 126 are located within an annular region 127 defined in the injection plate 115 about the axis of symmetry S. Each outlet of each end 126 faces the combustion chamber 110, entirely within the annular region 127.

[0148] The annular region 127 is bounded on the outside by an outer periphery CE defined by the combustion chamber wall 116. The outer periphery CE is defined at, or specifically at, an edge junction 128 between the injection plate 115 and the combustion chamber wall 116. That is, the radius of the outer periphery CE is approximately equal to the radius of the injection plate 115.

[0149] The annular region 127 is bounded internally by an inner circumference CI defined at the injection plate 115 between an edge junction 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 said radius.

[0150] The ends 126 are circumferentially distributed within the annular region 127. The ends 126 are angularly equidistant from each other about the axis of symmetry S. In the illustrated embodiment, the ends 126 are located approximately near an edge junction 128 between the injection plate 118 and the combustion chamber wall 116.

[0151] Each first injection passage 125 extends from the injection plate 115 to the combustion chamber 110 without any branches. At least the end 126 of each first injection passage 125 is straight. Preferably, each first injection passage 125 is straight throughout. At least the end 126 of each first injection passage 125 has a constant cross section. Preferably, the cross section of each first injection passage 125 is constant along its entire length.

[0152] Each end 126 has a respective injection direction d1 that is coincident with its main axis of extension and that generally determines the injection direction of the first combustion component within the combustion chamber 110. In the illustrated preferred embodiment, the injection direction d1 is coincident with the main axis of extension of each entire injection channel 125.

[0153] The injection direction d1 is based on the axis of symmetry S, and the axial component of the injection direction d1 is parallel to the axis of symmetry S and directed toward the combustion chamber 110 side.

[0154] Furthermore, the radial component of the injection direction d1 is directed toward the axis of symmetry S. That is, at least the end 126 of each injection flow path 125 approaches the combustion chamber 110 in a manner inclined toward the axis of symmetry S.

[0155] In the illustrated embodiment, the jet direction d1 of each end 126 has no tangential component, i.e., the tangential component of the jet direction d1 is zero.

[0156] The ejection directions d1 of the end 126 are incident on the axis of symmetry S, preferably at a common intersection point for all ejection directions d1.

[0157] Preferably, the injection direction d1 of each end 126 has an angle of 20° to 60°, even more preferably 30° to 50°, and even more preferably 35° to 45° relative to the axis of symmetry S. In the illustrated embodiment, this angle is 40°.

[0158] In an alternative embodiment not shown, the tangential component of the injection direction d1 of one or more ends 126 is non-zero, thereby defining a rotational movement of the first combustion component injected into the combustion chamber 110 about the axis of symmetry S.

[0159] The thruster 100 has a second connecting passage 130, shown in Figure 7, fluidly connectable to the second supply passage 21 and receiving the second combustion component from the second tank 20 from the second supply passage 21. The second connecting passage 130 is formed in the block 101.

[0160] The second connection channel 130 has a second binding site 131 in the fixing part 105 that can be fluidly tightly connected to the second supply channel 21 .

[0161] The injection plate 115 is provided with a second distribution channel 132. The second distribution channel 132 is formed in the block 101. The second distribution channel 132 is annular and symmetrical with respect to the axis of symmetry S. The second distribution channel 132 is located radially inward of the first distribution channel 122. Preferably, the second distribution channel 132 has a triangular cross section in a cut plane including the axis of symmetry S.

[0162] A second connecting channel 130 extends from the second binding site 131 to a second distribution channel 132 .

[0163] At least one second injection passage 135 is configured to inject the second combustion component into the combustion chamber 110. Preferably, there are multiple second injection passages 135. Preferably, the number of second injection passages 135 is between 2 and 8, and even more preferably between 3 and 6. In the embodiment shown in Figures 4, 5 and 6, there are four second injection passages 135.

[0164] A second injection passage 135 extends from the second distribution passage 132 to the combustion chamber 110 .

[0165] The second injection passage 135 is fluidly connectable to the second tank 20. When the second combustion component is injected into the combustion chamber 110, the second combustion component passes through the second coupling portion 131, the second connecting passage 130, the second distribution passage 132, and the second injection passage 135, preferably in this listed order.

[0166] Preferably, the second injection channels 135 have equal lengths between them.

[0167] Each end 136 of the second injection passages 135 is provided in the injection plate 115 and opens into the combustion chamber 110 .

[0168] The ends 136 of the second ejection passages 135 are located radially between the end 126 of the first ejection passage 125 and the axis of symmetry S. In other words, the end 136 of the second ejection passage 135 is located radially inward of the end 126 of the first ejection passage 125.

[0169] Preferably, the ends 136 of the second injection passages 135 are located radially inward of the annular region 127. That is, each outlet portion of each end 136 faces the combustion chamber 110 such that the entire end portion is located radially inward of the inner circumference CI of the annular region 127.

[0170] The ends 136 of the second ejection channels 135 are equidistant from the axis of symmetry S. The ends 136 of the second ejection channels 135 are distributed circumferentially around the axis of symmetry S. The ends 136 of the second ejection channels 135 are equidistant in the angular direction relative to the axis of symmetry S.

[0171] Each second injection passage 135 extends from the injection plate 115 to the combustion chamber 110 without any branches. At least the end 136 of each second injection passage 135 is straight. At least the end 136 of each second injection passage 135 has a constant cross section. Preferably, the cross section of each second injection passage 135 is constant along its entire length.

[0172] Each end 136 has a respective injection direction d2 that is coincident with its main axis of extension. The injection direction d2 generally determines the injection direction of the second combustion component within the combustion chamber 110. In the illustrated preferred embodiment, the injection direction d2 is coincident with the main axis of extension of each entire injection channel 135.

[0173] The injection direction d2 of each end 136 has an axial component that is parallel to the axis of symmetry S and points toward the combustion chamber 110. In a preferred embodiment, the injection direction d2 of each end 136 does not have a radial component. That is, the radial component of the injection direction d2 of each end 136 is zero. Preferably, the injection direction d2 of each end 136 does not have a tangential component. That is, the tangential component of the injection direction d2 of each end 136 is zero. That is, the injection direction d2 of each end 136 is parallel to the axis of symmetry S.

[0174] In an alternative embodiment (not shown), the radial component of the injection direction d2 of one or more end portions 136 is non-zero. The injection directions d2 of the end portions 136 may be incident on the axis of symmetry S, preferably at a common intersection point for all of the injection directions d2. Alternatively, the injection directions d2 of the end portions 136 may diverge from the axis of symmetry S. Preferably, the injection direction d2 of each end portion 136, whether converging or diverging with respect to the axis of symmetry S, has an angle of 0° to 20°, more preferably 0° to 10°, and even more preferably 0° to 5° with respect to the axis of symmetry S. This causes the second combustion component to be injected into the combustion chamber 110 with movement toward or away from the axis of symmetry S at an angle of 0° to 20°, more preferably 0° to 10°, and even more preferably 0° to 5° with respect to the axis of symmetry S. A 0° angle indicates that the injection direction d2 has no radial component.

[0175] In an alternative embodiment (not shown), the tangential component of the injection direction d2 at one or more ends 136 is non-zero, so that the second combustion component is injected into the combustion chamber 110 with a rotational motion about the axis of symmetry S.

[0176] 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 within the combustion chamber 110 to cause combustion between the first and second combustion components within the combustion chamber 110.

[0177] The spark plug is mounted in a first seat 140 formed on the block 101 at the axis of symmetry S. The first seat 140 is cylindrical.

[0178] Alternatively, in an embodiment not shown, a plurality of spark plugs may be mounted on the injection plate and equidistant about the axis of symmetry S.

[0179] The thruster 100 further includes a pressure sensor (not shown) attached to the injection plate 115 so as to face the combustion chamber 110. The pressure sensor is configured to measure the pressure in the combustion chamber 110 and control combustion.

[0180] The pressure sensor signal may be used to feedback regulate 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 .

[0181] The pressure sensor is attached to a second seat 141 formed in the block 101. The second seat 141 is formed radially between the end 136 of the second injection passage 135 and the axis of symmetry S. The second seat 141 is formed radially outward of the first seat 140.

[0182] In one embodiment not shown, a cylindrical slit is provided which extends around the axis of symmetry S facing the combustion chamber 110, in particular around the first seat 140 for the spark plug.

[0183] The cylindrical slit is provided with a plurality of additional injection channels (not shown) fluidically connected to the slit with their respective ends tangentially aligned. The additional injection channels are configured to inject the first and second combustion components into the cylindrical slit, imparting a circumferential motion thereto. This allows the first and second combustion components to partially or fully mix and exit the slit, forming a vortex around the axis of symmetry S within the combustion chamber 110. The vortex is introduced into the combustion chamber 110 at the spark plug. This facilitates ignition of the mixture of the first and second combustion components by the spark plug, while also moving the combustion away from the spark plug, thereby protecting the spark plug from excessively high temperatures.

[0184] The above-described propulsion system 1 with the thrusters 100 may be fixedly attached to an inter-orbital transfer vehicle (not shown) configured to travel in a given orbit and release a payload. The propulsion system 1 may be driven, for example, to transfer the inter-orbital transfer vehicle from a separation orbit, in which the inter-orbital transfer vehicle is separated from a rocket, to a deployment orbit, in which the inter-orbital transfer vehicle deploys a payload.

[0185] The propulsion system 1 may further be driven to transition the inter-orbital transfer vehicle from a first deployment orbit in which a first payload is deployed by the inter-orbital transfer vehicle to a second deployment orbit in which a second payload is deployed by the inter-orbital transfer vehicle.

[0186] The propulsion system 1 may also be powered to transfer the inter-orbital transfer vehicle from a deployment orbit, in which the vehicle deploys a payload, to a re-entry orbit, in which the vehicle is on a re-entry path into the atmosphere.

Claims

1. A thruster (100) for orbital control, A divergent channel (111) having an axial axis of symmetry (S), A combustion chamber (110) whose side is defined by the combustion chamber wall (116), A throat portion (112) is interposed between the combustion chamber (110) and the divergent channel (111), and the combustion chamber (110) is in fluid communication with the divergent channel (111). The injection plate (115) facing the combustion chamber (110), A plurality of first injection passages (125) are configured to be in fluid communication with the combustion chamber (110) and to inject a first combustion component into the combustion chamber (110), with each end (126) provided on the injection plate (115) and opening to the combustion chamber (110), At least one second injection passage (135) is in fluid communication with the combustion chamber (110) and is configured to inject a second combustion component into the combustion chamber (110), with its end (136) provided on the injection plate (115) and opening to the combustion chamber (110), Equipped with, The ends (126) of the plurality of first injection channels (125) are adjacent to the edge joint (128) between the injection plate (115) and the combustion chamber wall (116) within an annular region (127) whose center intersects the axis of symmetry (S), and extend along their respective injection directions (d1). The end (136) of at least one second injection channel (135) is located radially between the axis of symmetry (S) and the end (126) of the plurality of first injection channels (125), A thruster (100) for trajectory control, wherein the injection direction (d1) of each end (126) of the plurality of first injection passages (125) is such that the axial component is directed toward the combustion chamber (110) and the radial component is directed toward the axis of symmetry (S).

2. A thruster (100) for orbital control according to claim 1, wherein a block (101) of a material formed as a single unit has the divergent channel (111), combustion chamber (110), throat section (112), injection plate (115), a plurality of first injection channels (125), and at least one second injection channel (135) formed therein.

3. A thruster (100) for trajectory control according to claim 1 or 2, wherein the annular region (127) is defined by a concentric inner circumference (CI) and outer circumference (CE), the outer circumference (CE) is located at the edge joint (128), and the radius of the inner circumference (CI) of the annular region (127) is 75% or more of the radius of the outer circumference (CE).

4. A thruster (100) for orbital control according to Claim 3, wherein the radius of the inner circumference (CI) of the annular region (127) is 85% or more of the radius of the outer circumference (CE).

5. A thruster (100) for orbital control according to claim 1 or 2, wherein the annular region (127) is defined by a concentric inner circumference (CI) and outer circumference (CE), and the end (136) of the at least one second injection channel (135) is located between the inner circumference (CI) and the axis of symmetry (S).

6. A thruster (100) for orbital control according to claim 1 or 2, wherein the injection direction (d1) of each end (126) of the first injection channel (125) does not have a component oriented along the tangential direction.

7. A thruster (100) for orbital control according to claim 1 or 2, wherein the end (136) of the at least one second injection channel (135) extends along its respective injection direction (d2), and the projection of the injection direction (d2) of the end (136) of the at least one second injection channel (135) onto a plane including the axis of symmetry (S) forms an angle of 0° to 20° with respect to the axis of symmetry (S).

8. A thruster (100) for orbital control according to claim 1 or 2, wherein the projection of the injection direction (d1) of each end (126) of the plurality of first injection channels (125) onto a plane including the axis of symmetry (S) forms an angle of 20° to 60° with respect to the axis of symmetry (S).

9. A thruster (100) for orbital control according to claim 1 or 2, wherein the projection of the injection direction (d1) of each end (126) of the plurality of first injection channels (125) onto a plane including the axis of symmetry (S) forms an angle of 30° to 60° with respect to the axis of symmetry (S).

10. A thruster (100) for orbital control according to claim 1 or 2, wherein the projection of the injection direction (d1) of each end (126) of the plurality of first injection channels (125) onto a plane including the axis of symmetry (S) forms an angle of 35° to 60° with respect to the axis of symmetry (S).

11. A thruster (100) for orbital control according to claim 1 or 2, wherein there are multiple second injection channels (135).

12. A thruster (100) for orbital control according to claim 1 or 2, wherein the number of the plurality of first injection channels (125) is 7 to 28.

13. In the trajectory control thruster (100) according to claim 1 or 2, further, A spark plug is attached to the injection plate (115) along the axis of symmetry (S) and configured to generate a spark in the combustion chamber (110), A thruster (100) for orbital control, equipped with the following.

14. A thruster (100) for trajectory control according to claim 1 or 2, wherein the end (126) of the first injection channel (125) is straight without a branch and the cross-section is constant along the injection direction (d1).

15. A thruster (100) for orbital control according to claim 1 or 2, wherein the thrust direction (d1) of each end (126) of the plurality of first thrust channels (125) forms an angle of 30° to 60° with respect to a plane perpendicular to the axis of symmetry (S).

16. A thruster (100) for orbital control according to claim 1 or 2, wherein the thrust direction (d1) of each end (126) of the plurality of first thrust channels (125) forms an angle of 30° to 55° with respect to a plane perpendicular to the axis of symmetry (S).

17. A thruster (100) for orbital control according to claim 1 or 2, wherein the injection direction (d2) of the end (136) of the at least one second injection channel (135) does not have a component oriented along the radial direction.

18. A propulsion system (1) for orbital control, A thruster (100) for orbital control according to claim 1 or 2, A first tank (10) containing a first type of propellant and fluidly coupled to the plurality of first injection channels (125), A second tank (20) containing a second type of propellant and fluidly coupled to the at least one second injection channel (135), A propulsion system (1) for orbital control, comprising the first combustion component and the second combustion component being of the self-pressurizing type.

19. An inter-rail transport vehicle comprising a thruster (100) for track control according to claim 1 or 2.

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