Hybrid propulsion systems for spacecraft

Abstract

The present invention relates to a hybrid propulsion system (100) for a spacecraft or rocket, comprising an external body (10) defining a combustion chamber (15) for containing a block of solid propellant (20), a pressurized tank (30) for containing a liquid propellant (40), and a nozzle (50) for ejecting combustion gases produced by the reaction of the two propellants. The pressurized tank is located inside the external body, surrounded by the block of solid propellant. The propulsion system further comprises a plurality of liquid propellant injectors (31) axially arranged between the portions of the solid propellant and for improving the effectiveness of combustion. The present invention also relates to a launch vehicle-type spacecraft equipped with such a hybrid propulsion system.

Description

[Technical field]

[0001] The present invention relates to the general field of aerospace hybrid propulsion, in particular the construction of hybrid rocket engines, and more specifically to hybrid propulsion systems (or thrusters) primarily intended for spacecraft and spacecraft, such as launch vehicles.

[0002] The present invention has direct application in the aerospace field. [Background technology]

[0003] So-called hybrid space propulsion is a technology that combines the advantages of solid and liquid propellants while reducing the effects of their drawbacks. The principles of this technology date back to the 1930s, but the first tests were carried out by the US Army in the 1950s. Patent document 1, for example, describes one of the early hybrid propulsion systems. Solid-liquid hybrid propulsion is characterized by the use of a solid fuel or propellant and a liquid oxidizer. Rarely, this configuration is reversed, with a liquid propellant and a solid oxidizer. The advantage of hybrid propulsion is its simplicity compared to the complex construction of liquid propulsion, but also because, unlike solid propulsion, the thrust can be adjusted by varying the supply of liquid oxidizer.

[0004] A considerable body of literature has extolled the potential of hybrid propulsion, but also highlighted the operational limitations of the technology, which currently hinder large-scale development.

[0005] In the 1990s, research into high-power hybrid rocket engines was carried out by Amroc, but was inconclusive.

[0006] In fact, SpaceShipOne was the first spacecraft to use hybrid propulsion (in 2004). Since then, hybrid propulsion has remained limited to a few specific applications, such as spaceships intended for passenger transport in suborbital flights, which see significant safety advantages (the propellants used are non-explosive) and potential cost savings.

[0007] Typically, in a hybrid propulsion system, a liquid oxidizer (LOX, N2O, H2O2, etc.) is injected into a combustion chamber containing a solid reductant (polymer, paraffin, etc.). The reaction between the two results in combustion, which creates the energy required for propulsion in a manner similar to that of solid or liquid propellant propulsion systems.

[0008] This principle allows for high thrust to be achieved, which can be adjusted by adjusting the flow rate of the injected oxidizer, and by choosing certain chemical species, hybrid propulsion can be made non-toxic and / or non-pyrotechnic in nature.

[0009] However, this propulsion method is limited by the performance of available materials, and in particular, the current consumption rate of the reductant is low, making it unsuitable for use in vehicles that require high thrust, such as launch vehicles.

[0010] The conventional structure of most known hybrid propulsion devices, as shown in Fig. 1a and Fig. 1b, firstly faces the problem of lack of compactness and volumetric efficiency due to the distance between the liquid propellant tank and the combustion chamber, and secondly faces the more critical problem of the phenomenon of swaying. Patent Document 2 describes a rocket engine structure in which the liquid propellant tank is entirely or partially arranged in the combustion chamber defined by the solid propellant, thereby addressing the above problems.

[0011] Despite this compact structure, the injection of the liquid propellant remains very localized within the combustion chamber, which makes it impossible to obtain an optimal mixture of the resulting gaseous species from the two propellants.

[0012] This mixing problem is exacerbated in larger combustion chambers, making existing hybrid propulsion systems unusable for large-scale space applications. This is because the scale of the turbulent structures in the gas flow does not scale linearly with the size of the combustion chamber. Therefore, in large combustion chambers, the combustion region is located farther away from the solid propellant burning surface than in small combustion chambers. This makes large combustion chambers less efficient than small combustion chambers.

[0013] Due to non-optimal mixing in current hybrid thrusters, the energy released by combustion remains below the maximum value corresponding to the optimal mixture, and some of the reagents exiting the thruster are unreacted and therefore exit without releasing energy, so the efficiency and therefore performance of the thrusters are clearly inferior to their theoretically achievable values.

[0014] Furthermore, for the reasons mentioned above, the heat released by the combustion is not fully transferred to the solid propellant. However, the heat transferred to the solid propellant determines the rate of consumption of the solid propellant, called the regression rate, represented by the solid arrow in Figure 2. This is because the more heat transferred to the solid propellant, the more quickly it is consumed. Since the thrust of the propulsion system depends in part on this regression rate, it is desirable to maximize the regression rate in high power applications.

[0015] For these many reasons, to the applicant's knowledge, there are no solutions available that would, inter alia, enable the use of a hybrid propulsion system large enough to propel a spacecraft until it reaches an extra-atmospheric orbit. [Prior art documents] [Patent documents]

[0016] [Patent Document 1] U.S. Pat. No. 3,274,771 [Patent Document 2] International Publication No. 2017 / 142590 Summary of the Invention [Problem to be solved by the invention]

[0017] The present invention aims to overcome the above-mentioned drawbacks and address the technical problems associated therewith. [Means for solving the problem]

[0018] To this end, the object of the invention is to provide a hybrid propulsion device, in particular for spacecraft or rockets, comprising an external body extending in a longitudinal direction X and accommodating therein a solid propellant, a pressurized tank containing a liquid or gaseous propellant and a nozzle for ejecting the gases resulting from the combustion of said two propellants, said propellant device being distinguished by the fact that it comprises a number of liquid or gaseous propellant injectors arranged axially between portions of solid propellant, said solid propellant consisting of at least one hollow cylindrical block.

[0019] According to one embodiment, the injector is disposed along an inner surface of the solid propellant.

[0020] According to a particularly advantageous embodiment, a pressurized tank is placed inside the external body and surrounded by said solid propellant, and the injectors are arranged along and around said pressurized tank.

[0021] In this manner, by locating the injector in close proximity to the solid propellant, a large amount of liquid propellant can be delivered throughout the combustion chamber formed by the channel between the inner surface of the solid propellant and the pressurized tank, thereby maintaining the combustion region close to the inner surface (burning surface) of the solid propellant regardless of the size (diameter and length) of the propellant device.

[0022] More specifically, the injectors are arranged in a longitudinal row along the longitudinal axis of the pressurized tank and in a longitudinal row relative to said longitudinal axis. Zhou Evenly spaced in columns of the direction, Zhou The rows of directions are equally spaced.

[0023] Advantageously, the outer body and the pressurized tank are coaxial.

[0024] According to one embodiment, liquid or gas propellant contained in a pressurized tank is delivered to the injector through a pipe connecting a valve-type flow control device and the injector, or through a double wall of the pressurized tank.

[0025] The pipes may be in contact with the walls of the pressurized tanks and may act as their cooling system, for example by conveying cryogenic liquid propellant therethrough.

[0026] According to an advantageous embodiment, the nozzle is of the aerospike type and includes a central body fixed to the pressurized tank and an annular body surrounding the central body and fixed to the external body of the propulsion device.

[0027] Preferably, the central body of the nozzle is connected to the pressurized tank by a swivel and / or slider connection and a jack-type actuator, making it possible to manipulate the nozzle by changing the cross section of its neck and / or the inclination of the central body relative to its longitudinal axis.

[0028] Advantageously, the pressurized tank is fixed at its lower end to the external body by a connecting rod, the connecting rod being arranged radially relative to the longitudinal axis of the pressurized tank.

[0029] According to an alternative embodiment, the pressurized tank is mounted so as to be movable within the external body by means of at least one jack at its lower end and an elastic connection at its upper end opposite said lower end, the jack or jacks being, for example, pneumatic jacks operated by the excess gas pressurizing the pressurized tank.

[0030] More specifically, each jack is arranged radially around the longitudinal axis X of the external body and connects the external body to the top of the central body of the nozzle, making it possible to manipulate said nozzle by modifying the cross section of its neck and / or the inclination of the central body relative to its longitudinal axis.

[0031] These jacks, which hold the pressurized tank and the nozzle central body in place, allow the geometry (eccentricity) of the nozzle neck to be changed, which allows the vectoring of the thrust and, therefore, the control of the thruster. Furthermore, these jacks allow the cross section of the nozzle neck to be changed to adapt the nozzle expansion while the thruster is in operation.

[0032] In terms of propellant chemistry, solid propellants may be based on polymers or paraffin wax, while liquid propellants may be cryogenic liquid oxygen (LOX), hydrogen peroxide (H2O2), or nitric oxide (N2O).

[0033] Another object of the present invention is to provide a spacecraft of launch vehicle type including a hybrid propulsion system as shown above.

[0034] The basic concept of the invention has been disclosed above in its most basic form, but other details and features will become clearer from a reading of the following description in conjunction with the accompanying drawings, which show, by way of non-limiting example, an embodiment of a hybrid propulsion device according to the principles of the invention. [Brief description of the drawings]

[0035] The drawings and elements in the drawings are not necessarily drawn to scale, and the same or similar elements are given the same reference numbers in all the drawings. [Figure 1a] 1 is a schematic diagram of a prior art hybrid propulsion system; [Figure 1b] FIG. 1b is a propulsion device according to FIG. 1a equipped with a turbopump for introducing liquid propellant. [Diagram 2]FIG. 1 is a schematic diagram showing the combustion chamber of a prior art hybrid propulsion device in operation. [Diagram 3] 1 is a schematic cross-sectional view of a hybrid propulsion device according to one embodiment of the present invention; [Figure 4] FIG. 2 is a partial cross-sectional perspective view of a propulsion device according to the present invention, showing the arrangement of an outer body, solid propellant, liquid propellant tank, and injector; [Figure 4a] FIG. 5 is a detail of FIG. 4 showing radial injection. [Diagram 5] FIG. 2 is a diagram showing an example of a configuration of a solid propellant block that can be used in a propulsion device according to the present invention. [Figure 6a] FIG. 2 is a cross-sectional view of a propulsion device according to one embodiment. [Figure 6b] 6 is a cross-sectional view of a propulsion device according to another embodiment using the solid propellant block of FIG. [Figure 7] FIG. 2 illustrates a partial exploded view of a propulsion device according to one embodiment. [Figure 8] FIG. 2 is a cross-sectional view of a propulsion device according to another embodiment, in which a liquid propellant tank is surrounded by a solid propellant second block. [Figure 9] FIG. 9 is a longitudinal cross-sectional detail of the propulsion device according to the embodiment of FIG. 8. [Figure 10] FIG. 2 is a partial longitudinal cross-sectional view of a propulsion device according to one embodiment, illustrating the connections between the outer body, the liquid propellant tank, and the nozzle. [Figure 11] FIG. 13 is an exploded view of a central body of a liquid propellant tank and nozzle of a propulsion device according to an embodiment using a deformable wall surrounding the central body. [Figure 12] FIG. 2 is a schematic diagram of a hybrid propulsion system according to one embodiment, in which the liquid propellant tanks are located outside the external body. [Figure 13] 2 is a schematic cross-sectional view of a hybrid propulsion device according to another embodiment of the present invention. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0036] In the embodiments described below, reference is made to a "hybrid" propulsion system intended primarily for spacecraft and rockets. This non-limiting example is provided for a better understanding of the invention and does not exclude the use of the propulsion system in tactical missiles, military aerial drones, and other adapted vehicles.

[0037] Elsewhere in this specification, the term "propulsion device" refers broadly to a space propulsion device, also known as a rocket propulsion device or rocket engine, and more precisely to a space propulsion system, and "hybrid propulsion device" refers to a hybrid space propulsion system of the solid and liquid propellant type.

[0038] With reference to Figures 1a, 1b and 2, it is important to recall that the prior art hybrid propulsion device, in simplified form, comprises a tank of liquid propellant LP pressurized by a pressurization system PS, and a combustion chamber CC storing solid propellant SP, the inner surface of which defines a combustion space and which terminates in a nozzle for ejecting the combustion gases to generate the thrust required for propulsion. As shown diagrammatically in Figure 2, the liquid propellant is forced into the combustion chamber through a valve that controls the flow rate and an injector that sprays the liquid propellant. If the pressure in the tank of liquid propellant does not exceed the pressure in the combustion chamber, a turbopump can be used to increase the pressure, as shown in Figure 1b. In the operating combustion chamber, shown diagrammatically in Figure 2, the droplets of injected liquid propellant are transformed into gaseous species LP, depending on the local temperature. * Similarly, from solid propellants, gaseous species SP * These gaseous species then mix in a combustion chamber and react highly exothermically to produce new species until one of the propellants is depleted. The combustion can be initiated by an external heat source such as a pyrotechnic igniter, electric arc, laser, or by spontaneous exothermic decomposition of a liquid propellant under the influence of a catalyst.

[0039] FIG. 3 shows a hybrid propulsion device 100 according to the invention, which mainly comprises an external body 10 for containing a solid propellant 20 defining a combustion chamber 15, a pressurized central tank 30 for containing a liquid propellant 40, the pressurized central tank being arranged inside the external body and extending along its longitudinal axis X, and a nozzle 50 for exhausting the combustion gases, connectively attached to the external body 10 by means of a jack 60.

[0040] According to the embodiment shown, the external body 10 has a generally cylindrical hollow shape with a circular base, and more precisely includes a lenticular cylindrical side wall 11 and a slightly curved top end wall 12 to increase mechanical strength against the pressure in the combustion chamber. Furthermore, for obvious aerodynamic and size reasons, the external body 10 has an elongated shape along the longitudinal axis X, which allows to reduce drag (in atmospheric flight) while allowing the storage of a large amount of solid propellant 20.

[0041] According to the embodiment shown, the solid propellant 20 is in the form of a block, which may be one piece (monoblock) or may be obtained by superposing several blocks, and has a shape and dimensions adapted to be housed in the external body 10. This is because the block of solid propellant 20 is of hollow form and has an outer surface that matches the shape of the external body 10 and an inner surface 21 of any shape, provided that it allows the arrangement of the other elements of the propulsion device 100, in particular the central tank 30, whose axial arrangement is essential in the context of the present invention, as emphasized below.

[0042] An inner surface 21 of the block of solid propellant 20 further defines a combustion chamber 15, the volume of which increases as the block of solid propellant is consumed during combustion.

[0043] The solid propellant 20 may be based on a polymeric material or a thermoplastic material such as fast-burning paraffin wax, for example, the solid propellant 20 is hydroxytelechelic polybutadiene (HTPB), a derivative of HTPB, or polyoxymethylene (POM).

[0044] The solid propellant 20 may also contain metallic additives such as aluminum, magnesium, lithium, or beryllium that increase the specific impulse of the thruster, i.e., the velocity of the ejected gases.

[0045] According to the embodiment shown, the central tank 30 is arranged coaxially inside the external body 10, has a cylindrical tubular shape with a circular base, is open at its upper end and is closed by a hemispherical bottom, and includes a number of injectors 31 arranged distributed along and around the central tank, a liquid propellant flow control device 32, a network of pipes 33 supplying the injectors 31, a thermal protector 34, and elastic connections 35 for attaching the central tank to the external body 10.

[0046] The central tank 30 needs to be pressurized since it is intended to contain a liquid propellant 40 or, according to an alternative embodiment, a gas propellant, for example, with liquid helium or, as will be shown below, with a liquid propellant, and for this purpose, the central tank 30 is provided with a suitable pressurization system, for example with a valve-type pressurization control device 36 and a pressurization channel 37, as shown diagrammatically in FIG.

[0047] With regard to the chemical nature of the liquid propellant 40, it may be cryogenic liquid oxygen (LOX), hydrogen peroxide (H2O2), nitric oxide (N2O), or any other suitable liquid propellant.

[0048] According to the embodiment shown, the injectors 31 are distributed radially around the central tank 30, or more precisely on its outer surface, in order to cover in the combustion space substantially the entire combustion surface of the block of solid propellant 20, which surface is also the inner surface 21. For this purpose, the injectors 31 are arranged in longitudinal rows, over most of the length of the central tank 30 and preferably equally spaced in the cross section of the central tank. The injectors 31 may also be inclined with respect to the cross section of the central tank 30.

[0049] It is important to note that the location of the injector 31 depends on the cross-sectional shape of the central tank 30, whether circular or polygonal, the shape of the inner surface of the block of solid propellant 20, and the specific parameters of the injector as previously mentioned.

[0050] This is because the injectors 31, which are preferably identical, are characterized by an atomized liquid jet 311 (or spray) that occupies a portion of the space bounded by a cone, which is not necessarily circular, and which is referred to as the "solid injection angle." Thus, by considering this solid injection angle, the optimal placement of the injectors 31 for a given configuration can be determined in terms of the configuration of the central tank 30, the inner surface 21 of the block of solid propellant 20, and the number of injectors to be incorporated, such that injection is uniform across the majority of the block of solid propellant 20.

[0051] Figure 4 shows the arrangement of an outer body 10, here a block of solid propellant 20 having a circular inner surface 21, a central tank 30 and injectors 31. The detailed view in Figure 4a shows the radial jet obtained by each injector, which is fed by a pipe 33.

[0052] Examples of possible peripheral configurations of the injector 31 are shown in Figures 6a and 6b, respectively, in a square arrangement when the solid propellant 20 is an annular block, and in a regular polygonal arrangement, here a pentagon, when the solid propellant is a star-shaped block having five arms.

[0053] It should be mentioned that the various morphologies of the internal surface 21 of the block of solid propellant 20 allow to increase the combustion surface area for the same volume of solid propellant, which favors the heat transfer and improves the quality of combustion. Thus, the combination of the enlarged internal surface of the solid propellant and the distribution of the injectors allows to significantly increase the efficiency of combustion compared to prior art solutions (which is not the case for prior art solutions with multi-channel blocks and upstream injectors), while ensuring a uniform distribution of the liquid propellant in the various cavities of the block of solid propellant.

[0054] Figure 5 shows a block of solid propellant 20 having a star-shaped inner surface 21 as in Figure 6b. The assembly of this block into a thruster is shown diagrammatically in an exploded view in Figure 7, which also shows the overall rotational symmetry of the thruster according to the described embodiment.

[0055] The injector 31 allows the injection of liquid propellant 40 into the combustion chamber 15 containing the solid propellant 20, the mixing of the two propellants resulting in high-heat combustion, providing the energy required for propulsion of the spacecraft equipped with the propulsion device 100. Combustion is maintained as long as the solid propellant 20 or the liquid propellant 40 is not completely consumed. In addition, the ignition, thrust regulation and extinguishing of the propulsion device 100 remain controllable by the amount of liquid propellant injected into the combustion chamber, as long as solid propellant remains in the combustion chamber.

[0056] The liquid propellant to be injected leaves the central tank 30 through a flow control device 32, for example a valve installed under said tank, and passes through a supply pipe 33 which conveys the pressurized liquid propellant to the injector 31. The pipe 33 has thin walls, for example preferably made of a copper alloy, and is arranged along the wall of the central tank 30 so as to constitute a convection cooling system by circulation of the liquid propellant (which may be cryogenic) for the central tank exposed to the extreme temperatures of the combustion chamber 15.

[0057] Additional thermal protection is provided by a thermal protector 34 which covers most of the outer surface of the central tank 30 exposed in the combustion chamber 15 in direct contact with the combustion gases. The thermal protector 34 may consist, for example, of a special cladding, tiles of heat-resistant material or a layer of solid propellant.

[0058] According to an alternative embodiment, not shown, the supply to the injector is realized through a double wall of a tank of liquid propellant.

[0059] The coaxial arrangement of the central tank 30 within the block of solid propellant 20 advantageously allows the injectors 31 to be close to and opposite said block so that the liquid propellant spray 311 effectively reaches the burning surface of the solid propellant, until the solid propellant is consumed. In addition, this arrangement favors the creation of a sustained local turbulence along the combustion chamber, improving gas mixing. The resulting gas mixing tends to bring the energy generated by combustion closer to the theoretical maximum (obtained if the reagents were completely reacted). These combustion gases are then ejected through the nozzles 50 located at the end of the central tank 30 and the external body 10.

[0060] According to the embodiment shown in FIG. 3, the nozzle 50 is of the long spike type, commonly referred to as an aerospike, known for its performance, particularly its effectiveness over a wide altitude range and its low fuel consumption at low altitudes, and includes a central body 51 surrounded by an annular body 52.

[0061] The central body 51 of the aerospike nozzle 50 can be attached to the extension of the central tank 30 by means of a joining member, for example, but is preferably manufactured integrally with the central body including the tank.

[0062] The central body 51 of the aerospike nozzle 50 has a conical shape, sometimes a slightly truncated hyperboloid of revolution, and defines the pipe through which the gas jet passes and therefore needs to be cooled during operation. For this purpose, the central body 51 is provided with a cooling system 511 provided by a cooling liquid 512 stored in a cavity bounded by said central body.

[0063] The cooling liquid 512 , which may be of the same nature as the liquid propellant 40 or another chemical, may serve, after vaporization, to supply the pressurization systems 36 and 37 of the central tank 30 .

[0064] According to an alternative embodiment not shown, the central body of the aerospike nozzle constitutes the bottom of a tank of liquid propellant, which acts as a heat transfer fluid for cooling said central body, so that part of the fluid is fed to the injector and injected into the combustion chamber, while the remaining part, vaporized, is used by the pressurization system.

[0065] The annular body 52 of the aerospike nozzle 50 is partly located at the lower end of the outer body 10 of the propulsion device and therefore has a shape adapted to be continuous with said outer body. The annular body 52 also defines the gas outlet and therefore needs to be cooled during operation. This is achieved, as in the case of the central body 51, through a cooling system 521 provided by a cooling liquid 522 stored in the internal cavity of the annular body 52.

[0066] Furthermore, the nozzle annulus 52 can be attached to the outer body 10 of the propulsion device by suitable joining means, but preferably can be manufactured integrally with said outer body for better aerodynamics and greater mechanical strength.

[0067] In the propulsion device 100 according to the illustrated embodiment, the aerospike nozzle 50 has a relative mobility with respect to the outer body 10 via a central body 51 of said nozzle, which is radially supported by at least one jack 60 allowing it to deviate substantially with respect to the longitudinal axis X of the propulsion device. The central body 51 may be connected to the outer body 10 by a number of jacks 60 uniformly distributed around its circumference.

[0068] For example, the central body 51 of the nozzle is connected to the outer body 10 by four jacks 60 arranged symmetrically so as to be orientable in different directions substantially inclined with respect to the longitudinal axis X.

[0069] The jack 60 is, for example, a pneumatic double-acting jack, itself powered by excess pressurized gas resulting from the evaporation of the cooling liquid in the nozzle 50 .

[0070] Since the central body 51 of the nozzle is fixed to the central tank 30, the latter must have a certain degree of mobility relative to the outer body 10 of the propulsion device in order to follow the movements of said central body. For this purpose, the central tank 30 is attached to the outer body 10 by an elastic connection 35 that acts like an accordion, allowing elastic deformation of the central tank 30 in response to the movements of the central body 51 of the aerospike nozzle 50.

[0071] The movement of the nozzle central body 51 makes it possible to vary the eccentricity of the cross section of the nozzle neck so as to direct the thrust during operation of the propulsion device 100 .

[0072] 10, the central tank 30 is fixed relative to the outer body 10 while being held centrally at its lower part by an adjustable length connecting rod 61 arranged radially between it and the outer body. The central body 51 of the aerospike nozzle 50 is connected to the central tank 30 by a swivel connection or preferably a slider connection combined with a swivel connection 62 and an actuator 63 such as a jack.

[0073] The connecting rod 61 makes it possible to adjust the centering of the central tank 30 during final assembly of the propulsion system to counter any bias remaining after manufacture or initial assembly.

[0074] A single swivel connection between the nozzle's central body 51 and the central tank 30 allows changing the orientation of the central body 51. On the other hand, a double slider swivel connection allows adjusting the cross section of the neck of the aerospike nozzle 50 in addition to the tilt of the central body 51.

[0075] Furthermore, the connection between the central body 51 and the central tank 30 can be made by two annular metal parts 65 spaced apart longitudinally from each other to allow their relative movement in response to the movement of the central body of the nozzle. Alternatively, referring to Figure 11, the connection can be made by a flexible wall 513 that deforms in an accordion-like manner under the effect of the movement of the central body 51 of the aerospike nozzle 50.

[0076] According to the embodiment of figures 8 and 9, the hybrid propulsion device further comprises a second block 25 of solid propellant arranged in layers around the central tank 30. An opening is provided in the second block 25 to receive the injector 31. This auxiliary block allows the generation of more gaseous species under the influence of the temperature of the combustion chamber, further improving the mixing of the propellants.

[0077] In view of the present invention, it is apparent that minor modifications may be applied to the hybrid propulsion system, and more specifically to the shape and arrangement of the liquid propellant tanks and injectors, without departing from the scope of the invention, the main object of which is claimed below.

[0078] For example, FIG. 12 shows an alternative construction according to the present invention, in which a central tank 30 is mounted on the outside of the outer body 10 and is connected to an injection system 70 equipped with multiple injectors 71 and mounted inside the outer body 10 along with a block of solid propellant 20.

[0079] 13 shows another embodiment of a hybrid propulsion device 100', in which the combustion chamber 15 is annular in shape with two solid propellants 20a and 20b attached to the inner and outer walls of the chamber, respectively. The inner wall on which the first solid propellant 20a is attached defines a central tank 30 of liquid propellant.

[0080] The liquid propellant 40 is self-pressurized by boiling, and the required heat is provided by a heat transfer fluid 512 which cools the walls of the nozzle 50 according to a closed thermodynamic cycle.

[0081] According to this embodiment, the propulsion device 100 ′ includes an injection dome 13 that allows vaporized liquid propellant 41 to be fed into the combustion chamber 15 from the top of the tank 30 .

[0082] In this way, a rich combustion mixture (excess solid propellant and deficit liquid propellant) occurs in the combustion chamber 15, vaporizing the solid propellant without producing very hot combustion gases. The gases then pass through the underside injection system 75, which provides the deficit vaporized liquid propellant to complete the combustion. The mixing takes place in the post-combustion chamber 16, defined by the wall of the nozzle 50.

[0083] Additionally, the blocks of cylindrical solid propellant 20a and 20b provide thermal protection to the walls of the combustion chamber in which they are mounted.

[0084] According to the embodiment of FIG. 13, the combustion chamber 15 includes an additional thermal protector 155 since it has upstream 151 and downstream 152 cavities that do not contain solid propellant.

[0085] The hybrid propulsion system 100 ′ includes a heat exchanger 80 that boils the liquid propellant 40 to enable self-pressurization within the tank 30 .

[0086] Vaporized liquid propellant 41 is used for combustion in both the combustion chamber 15 and the post-combustion chamber 16. Supply to the upper injector 31a and lower injector 31b is controlled by a flow regulator and two-way valve 38.

[0087] The lower injection system 75 routes gases from the combustion chamber 15 to the post-combustion chamber 16 and supports a series of lower injectors 31b, which pass cooling heat transfer fluid 512 from the outer wall through a heat exchanger 80. The heat exchanger is located in the center of the thruster, below the tank 30 of liquid propellant.

[0088] The aerospike nozzle 50 is cooled by the circulation of a heat transfer fluid within its double walls 51 and 52 (cooling system). The heat transfer fluid 512 carries thermal energy through sensible (heating) and possibly latent (evaporation). The heat transfer fluid is cooled in a heat exchanger and reinjected into its storage location. A pump 55 may function to force the heat transfer fluid through its cooling cycle.

[0089] Thus, the double block of solid propellant allows doubling the wetted surface area without making the geometry more complex (without multiple channels), while the internal oxidizer tank allows high compactness and a length-to-diameter ratio adapted to the launch stage.

[0090] The double block of solid propellant provides complete thermal protection for both walls (inner and outer) and also reduces thermal stresses at the end of the run when incomplete combustion in the combustion chamber causes the thickness of the solid propellant block to become too thick for sufficient thermal protection.

[0091] It is also possible that incomplete combustion can result in the necessary or optimal mixture ratio being achieved by re-injecting the vaporized liquid propellant into the post-combustion chamber. This structure is not susceptible to the variation in mixture ratio caused by the opening of channels in the combustion chamber.

[0092] Injecting vaporized liquid propellant avoids combustion instabilities caused by gas-liquid interaction (droplets) in the combustion chamber, and the injector is less complex than if a liquid fluid had to be injected. [Explanation of symbols]

[0093] 10 External Body 11 Side wall 12 Top wall 13 Jet Dome 15 Combustion chamber 16 Post-combustion chamber 20 Solid Propellants 20a Solid Propellant 20b Solid Propellant 21 Inside 25 2nd Block 30 Pressurized (center) tank 31 Injector 31a Upper injector 31b Lower injector 32 Flow Control Device 33 Pipe 34 Thermal protector 35 Elastic Connection 36 Pressure control device / pressure system 37 Pressurized Channel / Pressurized System 38 Two-way valve 40 Liquid Propellants 41 Vaporized Liquid Propellants 50 nozzles 51 Center body / cooling system 52 Annular body / cooling system 55 Pump 60 Jack 61 Connecting rod 62 Swivel Connection 63 Actuator 65 Circular metal parts 70 Injection System 71 Injector 75 Lower injection system 80 heat exchanger 100 Hybrid Propulsion Unit 100' Hybrid Propulsion Unit 151 Upstream 152 Downstream 155 Thermal protector 311 Atomizing Liquid Jet / Spray 511 Cooling System 512 Coolants / Heat Transfer Fluids 513 Flexible Wall 521 Cooling System 522 Coolant CC Combustion Chamber LP Liquid Propellant LP * Gaseous species PS Pressurization System SP Solid Propellant SP* Gaseous species X Longitudinal direction (axis)

Claims

1. an outer body (10) extending in a longitudinal direction X and accommodating a solid propellant (20) therein; a pressurized tank (30) containing a liquid or gas propellant (40); a nozzle (50) for ejecting gas generated by the combustion of the two propellants; A hybrid propulsion system (100, 100'), in particular for a spacecraft or rocket, comprising: the hybrid propulsion device (100, 100') includes at least one set of multiple liquid or gas propellant injectors (31) arranged along a longitudinal axis between the portions of the solid propellant; A hybrid propulsion system (100, 100'), wherein said solid propellant consists of at least one hollow cylindrical block.

2. 2. The hybrid propulsion device of claim 1, wherein the liquid or gas propellant injectors (31) are disposed along an inner surface (21) of the solid propellant (20).

3. The pressurized tank (30) is installed inside the outer body (10) and is surrounded by the solid propellant (20); 3. A hybrid propulsion system as claimed in claim 2, wherein the injectors (31) are arranged circumferentially along the pressurized tank.

4. 4. A hybrid propulsion device as claimed in claim 3, wherein the injectors (31) are uniformly arranged in longitudinal rows along a longitudinal axis of the pressurised tank (30) and in circumferential rows relative to said longitudinal axis, the circumferential rows being equally spaced.

5. A hybrid propulsion device according to any one of claims 1 to 4, wherein the outer body (10) and the pressurized tank (30) are coaxial.

6. 6. A hybrid propulsion device according to claim 1, wherein the liquid or gas propellant (40) contained in the pressurized tank (30) is delivered to the injector (31) via a flow control device (32) and a pipe (33).

7. 7. A hybrid propulsion device according to any one of claims 1 to 6, wherein the nozzle (50) is of the aerospike type and comprises a central body (51) fixed to the pressurized tank (30) and an annular body (52) fixed to the outer body (10) of the propulsion device.

8. 8. A hybrid propulsion device according to claim 7, characterized in that the central body (51) of the nozzle (50) is connected to the pressurized tank (30) by a swivel and / or slider connection (62) and a jack-type actuator (63) making it possible to manipulate the nozzle by changing the cross section of its neck and / or the inclination of the central body relative to the longitudinal axis.

9. 9. A hybrid propulsion device according to any one of claims 1 to 8, wherein the pressurized tank (30) is attached to the external body (10) by a connecting rod (61) at its lower end, the connecting rod being arranged radially with respect to the longitudinal axis of the pressurized tank.

10. 10. The hybrid propulsion system of claim 1, wherein the solid propellant (20) is based on a polymer or paraffin wax and the liquid propellant (40) is cryogenic liquid oxygen, hydrogen peroxide, or nitric oxide.

11. 11. A launch vehicle type aerospace vehicle comprising a hybrid propulsion system (100) according to any one of claims 1 to 10.

Images