EP4045786A1

EP4045786A1 - Hybrid thruster for space vehicle

Info

Publication number: EP4045786A1
Application number: EP20803224.3A
Authority: EP
Inventors: Alexandre MANGEOT; Sylvain BATAILLARD; Vincent ROCHER; Alexis AZOULAI
Original assignee: Hybrid Propulsion For Space
Current assignee: Hybrid Propulsion For Space
Priority date: 2019-10-17
Filing date: 2020-10-16
Publication date: 2022-08-24
Also published as: AU2020368037A1; FR3102219B1; KR20220078710A; IL292237A; FR3102219A1; CN114599871A; WO2021074553A1; CA3154256A1; US20220381201A1; JP2022553637A

Abstract

The present invention relates to a hybrid thruster (100), for a spacecraft or vehicle, comprising an outer body (10), defining a combustion chamber (15) intended for receiving a block of solid propellant (20), a pressurised tank (30) intended for receiving a liquid propellant (40), and a nozzle (50) for ejecting the combustion gases produced by the reaction of the propellants, the pressurised tank being placed inside the outer body, surrounded by the block of solid propellant, the thruster further comprising a plurality of liquid propellant injectors (31) arranged axially between parts of the solid propellant in order to improve combustion efficiency. The invention also relates to a space vehicle, of the booster type, equipped with such a hybrid thruster.

Description

Hybrid thruster for space vehicle

TECHNICAL AREA

The present invention belongs to the general field of aerospace hybrid propulsion, in particular of the architecture of hybrid rocket engines, and more particularly relates to a hybrid propulsion (or propellant) system intended mainly for vehicles and spacecraft such as launchers.

The present invention finds direct application in the field of astronautics.

STATE OF THE ART So-called hybrid space propulsion consists in using both solid propellants and liquid propellants, in order to combine their advantages while reducing the effects of their individual disadvantages. The principle of this technology dates back to the 1930s but the first tests were carried out by the US military in the 1950s. The patent document US3274771A describes, for example, one of the first hybrid propulsion systems. Hybrid solid-liquid propulsion is characterized by the use of a fuel, or fuel, solid and a liquid oxidizer. In rare cases, this configuration is reversed with a liquid fuel and a solid oxidizer. The advantage of hybrid propulsion stems from its simplicity compared to the complex architectures of liquid propulsion, but also from the fact that the thrust can be modulated by varying the supply of liquid oxidizer, unlike solid propulsion.

A fairly abundant literature extols the potential of hybrid propulsion, but also highlights the operational limits of this technology which, to date, hinder its development on a larger scale. In the 1990s inconclusive research was carried out by the Amroc company on high power hybrid rocket motors. Indeed, SpaceShipOne was the first space plane (2004) to use hybrid propulsion. Since then, hybrid propulsion has remained confined to a few niche applications such as the SpaceShip which, with a view to transporting passengers in suborbital flights, sees a considerable safety advantage (the propellants used are not explosive) and a possibility of cost reduction.

Typically, in a hybrid propellant, a liquid oxidant (LOX, N2O, H202, etc.) is injected into a combustion chamber containing a solid reducing agent (polymers, paraffins, etc.). The reaction of these two species generates combustion which produces the energy necessary for propulsion, in a manner comparable to the operation of a solid propellant or liquid propellant propellant.

According to this principle, it is possible to obtain strong thrusts, moreover, can be modulated by adjusting the flow rate of the injected oxidant. In addition and by means of the choice of certain chemical species, the hybrid propulsion can have a non-toxic and / or non-pyrotechnic character.

However, this propulsion has drawbacks linked to the limited performance of the materials available, in particular the low consumption speeds of current reducers, making its use inappropriate in vehicles requiring a large thrust such as launchers.

The usual architecture of the majority of known hybrid thrusters, shown schematically in Figures 1a and 1b, comes up against, on the one hand, a lack of compactness and volumetric efficiency due to the distance between the liquid propellant tank and the combustion chamber, and on the other hand, to more critical sloshing phenomena. Patent document WO2017142590 addresses these issues by describing a rocket engine architecture in which the liquid propellant reservoir is placed, entirely or partially, in the combustion chamber defined by the solid propellant.

Despite this compact architecture, the injection of liquid propellant remains very localized in the combustion chamber and does not allow an optimal mixture of the gaseous chemical species from the two propellants to be obtained.

This mixing problem is amplified when the combustion chamber is large, and the existing hybrid propulsion systems cannot therefore be used in large-scale space applications. This is due to ensures that the turbulent structures that form in the gas flow do not change in scale with the size of the combustion chamber. As a result, the combustion zone is located further from the solid propellant combustion surface in a large combustion chamber than in a small one. So a large combustion chamber is less efficient than a small combustion chamber.

Due to a non-optimal mixture in current hybrid thrusters, the energy released by combustion remains below a maximum value corresponding to an optimal mixture, with some of the reactants ejected from the thruster without having reacted and therefore without having released d 'energy. As a result, the efficiency and hence the performance of the thruster are significantly lower than their theoretically achievable values.

In addition, the heat given off by combustion is less well transferred to the solid propellant for the aforementioned reasons. Now, this heat transferred to the solid propellant determines the rate of consumption of said propellant, called the rate of regression, represented by solid arrows in FIG. 2. In fact, the more heat transferred to the solid propellant, the greater the latter. is consumed quickly. The thrust of the thruster being in part dependent on this rate of regression, maximization of the latter is desirable for high power applications.

For these multiple reasons, no solution, to the knowledge of the applicant, makes it possible to use a hybrid propulsion system of sufficient size to, among other things, propel a space vehicle until it reaches an extraatmospheric orbit.

PRESENTATION OF THE INVENTION

The present invention aims to overcome the drawbacks set out above and to respond to the technical problems relating thereto.

To this end, the present invention relates to a hybrid propellant, in particular for a vehicle or spacecraft, comprising an outer body in which is stored a solid propellant, extending in a longitudinal direction X of the outer body, a pressurized tank containing a liquid or gaseous propellant, and an ejection nozzle for the gases produced by the combustion of propellants. This propellant is remarkable in that it comprises a plurality of liquid or gaseous propellant injectors arranged axially between parts of the solid propellant, and in that said solid propellant comprises at least one hollow cylindrical block.

According to one embodiment, the injectors are arranged along an internal surface of the solid propellant.

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

Thus, this proximity of the injectors with the solid propellant makes it possible to supply a large quantity of liquid propellant all along the combustion chamber formed by the channel between the internal surface of the solid propellant and the pressurized tank, so maintaining the combustion zone close to the internal (combustion) surface of the solid propellant regardless of the size (diameter and length) of the propellant.

More particularly, the injectors are arranged uniformly in longitudinal rows, along a longitudinal axis of the pressurized tank, and in radial rows, with respect to said longitudinal axis, the radial rows being equidistant.

Advantageously, the outer body and the pressurized tank are coaxial.

According to one embodiment, the liquid or gaseous propellant contained in the pressurized tank is sent to the injectors via a flow control device, of the valve type, and pipes connecting the injectors or via a double wall of the pressurized tank.

The pipes can be in contact with the walls of the pressurized tank and serve them as a cooling system by conveying a cryogenic liquid propellant, for example.

According to an advantageous embodiment, the nozzle is of the aerospike type and comprises a central body secured to the pressurized tank and an annular body, surrounding the central body, secured to the outer body of said thruster.

Advantageously, the central body of the nozzle is connected to the pressurized tank by means of a ball and / or rail connection and actuators, of the type jacks, so as to enable said nozzle to be maneuvered by modifying the section of its coi and / or the inclination of said body centered with respect to said longitudinal axis

Advantageously, the pressurized tank is fixed to the outer body by means of connecting rods at its lower end, said connecting rods being arranged radially with respect to a longitudinal axis of said tank.

According to an alternative embodiment, the pressurized tank is movably mounted in the outer body by means of at least one jack, at its lower end, and elastic connections, at its upper end opposite said lower end. The cylinder or cylinders are for example pneumatic cylinders operating thanks to the excess pressurization gas from the pressurized tank.

More particularly still, each jack is arranged radially around the longitudinal axis X of the outer body and connects said outer body to an upper part of the central body of the nozzle, so as to allow said nozzle to be maneuvered by modifying the section of its neck. and / or the inclination of said central body relative to said longitudinal axis.

These jacks for maintaining the pressurized tank and the central nozzle body in position make it possible to modify the shape (eccentricity) of the nozzle throat. This modification makes it possible to orient the thrust and therefore to control the thruster. The cylinders also make it possible to modify the section of the nozzle throat to adapt the rate of expansion of the nozzle during operation of the thruster.

With regard to the chemical nature of propellants, the solid propellant can be based on a polymer or a paraffin wax, and the liquid propellant can be cryogenic liquid oxygen (LOX), peroxide of 'hydrogen (H202) or nitrous oxide (N2O).

The invention also relates to a space vehicle, of the launcher type, comprising a hybrid thruster as has been presented.

The fundamental concepts of the invention having just been explained above in their most elementary form, other details and characteristics will emerge more clearly on reading the description which follows and with reference to the appended drawings, giving by way of nonlimiting example an embodiment of a hybrid propellant in accordance with the principles of the invention. BRIEF DESCRIPTION OF THE FIGURES

The different figures and the elements of the same figure are not necessarily represented on the same scale. In all of the figures, identical or equivalent elements bear the same reference numeral.

It is thus illustrated in:

- Figure 1a: a schematic view of a hybrid propellant of the prior art;

- Figure 1b: the propellant of Figure 1a equipped with a liquid propellant inlet turbopump;

- Figure 2: a schematic view of the combustion chamber of the prior art hybrid thruster in operation;

- Figure 3: a schematic sectional view of a hybrid thruster according to one embodiment of the invention;

- Figure 4: a partial sectional and perspective view of a propellant according to the invention, showing the arrangement of the outer body, the solid propellant, the liquid propellant tank and the injectors;

- Figure 4a: a detail of Figure 4 showing a radial injection;

- Figure 5: an example of the shape of a solid propellant block that can be used in a propellant according to the invention;

- Figure 6a: a cross section of a thruster according to one embodiment;

- Figure 6b: a cross section of a thruster according to another embodiment, with the solid propellant block of Figure 5;

- Figure 7: a partial exploded view of a thruster according to one embodiment;

- Figure 8: a cross section of a propellant according to another embodiment, in which the liquid propellant tank is surrounded by a block of secondary solid propellant;

- Figure 9: a detail in longitudinal section of a thruster according to the embodiment of Figure 8; FIG. 10: a partial view in longitudinal section of a thruster according to one embodiment, showing the connections between the outer body, the liquid propellant tank and the nozzle;

Figure 11: an exploded view of the liquid propellant tank and of the central body of the nozzle of a propellant according to one embodiment, with a deformable wall surmounting said central body;

Figure 12: a schematic view of a hybrid thruster according to an embodiment in which the liquid propellant tank is outside the outer body;

Figure 13: a schematic sectional view of a hybrid thruster according to another embodiment of the invention.

DETAILED DESCRIPTION OF EMBODIMENTS

In the embodiment described below, reference is made to a “hybrid” propulsion system intended mainly for vehicles and spacecraft. This non-limiting example is given for a better understanding of the invention and does not exclude the use of the propulsion system in tactical missiles, military aerial drones or any other suitable vehicle.

In the remainder of the description, the term “thruster” designates by extension a space thruster, also called a rocket thruster or rocket engine, and more exactly a space propulsion system, and the expression “hybrid thruster” designates a hybrid space propulsion system. with solid propellant and liquid propellant.

It is important to remember that the hybrid propellants of the prior art, with reference to FIGS. 1a, 1b and 2, comprise, in a simplified manner, a tank of liquid propellant LP pressurized by a pressurization system PS and a combustion chamber CC storing a solid SP propellant, the internal surface of which delimits a combustion volume, and ending in a combustion gas ejection nozzle generating the thrust required for propulsion. The liquid propellant is pushed into the combustion chamber passing through a flow control valve and an injector which sprays the liquid propellant as a spray as shown schematically in FIG. 2. When the pressure in the liquid propellant tank is not does not exceed the pressure in the combustion chamber, a turbopump can be used for further pressure as shown in Figure 1b. In the combustion chamber in operation shown schematically in FIG. 2, the injected liquid propellant droplets produce, in response to the local temperature, gaseous chemical species LP ^* . Likewise, solid propellant produces gaseous chemical species SP ^* . These gaseous chemical species then mix in the combustion chamber and operate in a very exothermic manner to produce new chemical species until one of the propellants is exhausted. This combustion can be initiated by an external heat source such as a pyrotechnic igniter, an electric arc, a laser, etc., or by spontaneous exothermic decomposition of the liquid propellant under the effect of a catalyst.

FIG. 3 represents a hybrid propellant 100 according to the invention, mainly comprising an outer body 10, intended to receive a solid propellant 20 which defines a combustion chamber 15, a pressurized central tank 30 intended to receive a liquid propellant 40, said tank being placed inside the outer body and extending along a longitudinal axis X thereof, and a combustion gas outlet nozzle 50, mounted articulated on the outer body 10 by means of jacks 60.

The outer body 10, according to the illustrated embodiment, has the shape of a generally cylindrical cavity with a circular base, and more precisely comprises a cylindrical side wall 11 and a slightly curved upper end wall 12, of lenticular shape for better resistance. mechanical to the pressures in the combustion chamber. In addition and for obvious reasons of aerodynamic performance and size, the outer body 10 has an elongated shape along the longitudinal axis X, thus reducing the drag (in atmospheric flight) while allowing the storage of a large quantity of 'solid propellant 20.

The solid propellant 20, according to the illustrated embodiment, is in the form of a block, which can be in one piece (monoblock) or obtained by the superposition of several blocks, having a shape and dimensions suitable for be stored in the outer body 10. In fact, the solid propellant block 20 is hollow in shape and has an outer surface the shape of which matches that of the outer body 10 and an inner surface 21 of any shape, such as underlined below, on condition that the other elements of the thruster 100 can be arranged, in particular the central reservoir 30, the axial arrangement of which is essential in the context of the present invention.

The internal surface 21 of the solid propellant block 20 also makes it possible to delimit the combustion chamber 15, the volume of which increases as the said block is consumed during combustion.

The solid propellant 20 may be based on a polymerized material or on a thermoplastic material such as a fast burning paraffin wax. For example, solid propellant 20 is hydroxytelechelic polybutadiene (PBHT), a derivative of PBHT, or polyoxymethylene (POM).

The solid propellant 20 can also include metallic additives such as aluminum, magnesium, lithium or beryllium which make it possible to increase the specific impulse of the propellant, in other words the speed of the ejected gases.

The central reservoir 30, according to the illustrated embodiment, is arranged coaxially inside the outer body 10, has a cylindrical tubular shape with a circular base, open at its upper end and closed by a hemispherical bottom, and comprises a plurality of injectors 31 distributed along and around said tank, a device for controlling the flow of liquid propellant 32, a network of pipes 33 supplying the injectors 31, thermal protections 34 and elastic connections 35 by which the central tank is attached to the outer body 10.

The central tank 30 is intended to contain the liquid propellant 40, or according to an alternative embodiment a gaseous propellant, and therefore needs to be pressurized. To this end, the central tank 30 is provided with a suitable pressurization system, for example with liquid helium or using liquid propellant as indicated below, comprising for example a pressurization control device 36, of the valve type, as well as a pressurization channel 37 shown diagrammatically in FIG. 3.

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

The injectors 31, according to the illustrated embodiment, are distributed radially along and around the central reservoir 30, more precisely on a external surface of said reservoir, so as to allow the most efficient possible injection into the combustion volume, reaching almost the entire combustion surface of the solid propellant block 20 (which is also its internal surface 21). To this end, the injectors 31 are arranged in longitudinal rows, over a major part of the length of the central reservoir 30, and along cross sections of said reservoir, preferably equidistant. The injectors 31 can also be inclined relative to the transverse planes of the central reservoir 30.

It is important to note that the location of the injectors 31 depends on the shape of the cross section of the central tank 30, circular or polygonal, on the shape of the internal surface of the solid propellant block 20 and on the intrinsic parameters of said injectors.

In fact, the injectors 31, which are preferably identical, are characterized by the sprayed liquid jet 311 (or spray) which occupies a portion of the space delimited by a cone which is not necessarily circular, the latter will be designated by "solid injection angle. ". Thus, taking into account this solid injection angle makes it possible to determine the optimal arrangement of the injectors 31 for a given configuration in terms of shapes of the central reservoir 30, of the internal surface 21 of the solid propellant block 20 and of the number injectors which will be integrated, so as to obtain a uniform injection over a major part of the solid propellant block 20.

FIG. 4 represents the arrangement of the external body 10, of the solid propellant block 20, here with a circular internal surface 21, of the central reservoir 30 and of the injectors 31. The detail of FIG. 4a illustrates the radial injection obtained by each injector, the injectors being supplied by the pipes 33.

Examples of possible peripheral arrangements of the injectors 31 are given in FIGS. 6a and 6b, with respectively a square arrangement in the case of an annular solid propellant block and a regular polygonal arrangement, here in pentagon, in the case of an annular solid propellant block. a solid star propellant block, here with five branches.

This is reminiscent of the various shapes of the internal surface 21 of the solid propellant block 20 make it possible to increase the combustion surface for the same volume of solid propellant, this combustion surface favoring the heat transfer and improving the quality of combustion. Thus, the combination of an enlarged solid propellant internal surface and the distribution of the injectors makes it possible to significantly increase the efficiency of the combustion compared to the solutions of the prior art, while ensuring a homogeneous distribution of the fuel. liquid propellant in the various cavities of the solid propellant block (which is not the case in the previous solutions having multichannel blocks and an upstream injector).

Figure 5 shows a solid propellant block 20 having a star-shaped internal surface 21 such as that of Figure 6b. The integration of this block in the thruster is shown schematically by the exploded view of Figure 7 which also reflects the overall rotational symmetry of the thruster according to the embodiment described.

The injectors 31 then make it possible to inject liquid propellant 40 into the combustion chamber 15 containing the solid propellant 20, the mixture of the two propellants producing a very exothermic combustion which provides the energy necessary for the propulsion of a machine. space equipped with the propellant 100. Combustion is maintained as long as no propellant, solid 20 or liquid 40, is completely consumed. In addition, the ignition, the thrust modulation and the extinction of the propellant 100 remain controllable by the amount of liquid propellant injected into the combustion chamber, provided that there is still solid propellant in said chamber.

The injected liquid propellant comes from the central tank 30 passing through the flow control device 32, which is for example a valve placed under said tank, and through the supply pipes 33 which allow the liquid propellant to be conveyed under pressure towards the injectors 31. The pipes 33 are for example thin-walled, preferably copper alloy, lining the walls of the central tank 30 to thus constitute a convective cooling system, thanks to the circulation of the liquid propellant (which can be cryogenic), for said tank subjected to the extreme temperatures of the combustion chamber 15.

Additional thermal protection is provided by thermal protections 34 which cover a major part of the external surface of the central tank 30, said surface bathing in the combustion chamber 15 while being in direct contact with the combustion gases. Thermal protections 34 consist for example of a specific coating, tiles made of heat-resistant material, or a layer of solid propellant.

According to an alternative embodiment not shown, the injectors are supplied by a double wall of the liquid propellant tank.

The coaxial arrangement of the central reservoir 30 inside the solid propellant block 20 advantageously allows the multiple injectors 31 to be opposite and close to said block so that the liquid propellant sprays 311 effectively reach the combustion surface of the solid propellant, until the solid propellant is used up. In addition, this arrangement promotes the formation of local turbulence maintained along the combustion chamber for improved gas mixing. The resulting mixture of gases is such that the energy produced by combustion tends to approach a theoretical maximum (obtained during a total reaction of the reactants). These combustion gases are then ejected by the nozzle 50 located at the end of the central tank 30 and of the outer body 10.

The nozzle 50, according to the embodiment illustrated in FIG. 3, is of the long point type, commonly called an aerospike, known for its performance, in particular its efficiency in a wide range of altitudes and its low fuel consumption at low altitude, and comprises a central body 51 surrounded by an annular body 52.

The central body 51 of this aerospike nozzle 50 can be attached as an extension of the central reservoir 30 by means of a junction member, for example, or preferably made in one piece with a central body also containing said reservoir.

The central body 51 of the aerospike nozzle 50 has a conical shape, sometimes slightly of a truncated hyperboloid of revolution, defining a ramp which channels the gas jet and which, therefore, must be cooled in operation. To this end, the central body 51 is provided with a cooling system 511 supplied by a cooling liquid 512 stored in a cavity delimited by said central body.

The cooling liquid 512, which can be of the same nature as the liquid propellant 40 or another chemical substance, can also be used, after vaporization, to supply the pressurization system 36 and 37 of the central tank 30. According to an alternative embodiment not shown, the central body of the aerospike nozzle constitutes the bottom of the liquid propellant reservoir so that the liquid propellant serves as a coolant to cool said body. Thus, part of the fluid feeds the injectors to be injected into the combustion chamber and the other part, vaporized, is used by the pressurization system.

The annular body 52 of the aerospike nozzle 50 is for its part located at the level of the lower end of the outer body 10 of the thruster and thus has a shape adapted in continuity with said outer body. The annular body 52 also defines gas ejection ramps and therefore needs to be cooled in operation. This is achieved, as in the case of the central body 51, via a cooling system 521 supplied with a cooling liquid 522 stored in internal cavities of the annular body 52.

In addition, the annular body 52 of the nozzle can be attached to the outer body 10 of the thruster by suitable connecting means, or preferably made in one piece with said outer body for better aerodynamics and more mechanical strength.

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

For example, the central body 51 of the nozzle is connected to the outer body 10 by four cylinders 60 arranged symmetrically so as to be able to point in different directions substantially inclined relative to the longitudinal axis X.

The cylinders 60 are for example double-acting pneumatic cylinders supplied with the surplus pressurization gas, itself resulting from the vaporizations of the cooling liquids from the nozzle 50.

The central body 51 of the nozzle being integral with the central reservoir 30, the latter must have a certain degree of mobility with respect to the outer body 10 of the thruster in order to follow the movements of said central body. As a result, the central reservoir 30 is fixed to the outer body 10 by means of elastic links 35 which work in accordion fashion to allow elastic deformations of the reservoir. central 30 in response to the movements of the central body 51 of the aerospike nozzle 50.

The movements of the central body 51 of the nozzle make it possible to modify the eccentricity and the section of the neck of the nozzle so as to orient the thrust during the operation of the thruster 100.

Preferably and according to the embodiment of FIG. 10, the central reservoir 30 is fixed relative to the outer body 10 by being centered and maintained at the level of its lower part by connecting rods 61 adjustable in length arranged radially between said reservoir and said body. exterior; and the central body 51 of the aerospike nozzle 50 is connected to the central reservoir 30 by a ball joint, or preferably a slide connection combined with a ball joint 62, as well as by actuators 63 such as jacks.

The connecting rods 61 are used to adjust the centering of the central tank 30 during the final assembly of the thruster, to make up for any residual off-center after manufacture or initial assembly.

A single ball joint between the central body 51 of the nozzle and the central reservoir 30 allows the orientation of said body to be changed. On the other hand, a double slide-ball joint connection makes it possible, in addition to the inclination of the central body 51, to adjust the section of the neck of the aerospike nozzle 50.

In addition, the junction between the central body 51 and the central reservoir 30 can be made by two annular metal parts 65 spaced apart longitudinally from one another to allow their relative movement in response to the movement of the central nozzle body. Alternatively and with reference to Figure 11, this junction can be made by means of a flexible wall 513 which deforms in an accordion shape under the effect of the movements of the central body 51 of the aerospike nozzle 50.

According to the embodiment of FIGS. 8 and 9, the hybrid propellant further comprises a block of secondary solid propellant 25 arranged in a layer around the central tank 30. Openings receiving the injectors 31 are made in the block 25. This additional block allows the production of more gaseous chemical species under the effect of the temperature of the combustion chamber, further improving the mixing of propellants. In view of the invention, it clearly appears that minor modifications can be applied to the hybrid propellant, and more particularly to the geometry and the arrangement of the liquid propellant tank and of the injectors, without however departing from the scope of the invention. invention, the main object of which is claimed below.

For example, FIG. 12 represents an alternative architecture according to the invention in which the central reservoir 30 is placed outside the outer body 10, being connected to an injection system 70 provided with a plurality of injectors 71 and placed inside said outer body, along the solid propellant block 20.

FIG. 13 represents a hybrid propellant 100 'according to another embodiment, in which the combustion chamber 15 has an annular shape with two blocks of solid propellant 20a and 20b respectively fixed on an internal wall and an external wall of said chamber. . The internal wall on which is fixed the first block 20a of solid propellant delimits the central reservoir 30 of liquid propellant.

The liquid propellant 40 is self-pressurized by boiling, the necessary heat being provided by the coolant 512 which allows the walls of the nozzle 50 to be cooled according to a closed thermodynamic cycle.

According to this embodiment, the propellant 100 ’comprises an injection dome 13 making it possible to supply gasified liquid propellant 41 from the upper part of the tank 30 into the combustion chamber 15.

Thus, combustion in a rich mixture (excess of solid propellant and lack of liquid propellant) occurs in the combustion chamber 15 and makes it possible to vaporize the solid propellant without producing very hot combustion gases. The gases produced then pass through a lower injection system 75 which supplies the missing gasified liquid propellant to complete combustion. The mixing takes place in a post-combustion chamber 16 delimited by the walls of the nozzle 50.

In addition, the cylindrical blocks of solid propellant 20a and 20b thermally protect the walls of the combustion chamber to which they are fixed. According to the embodiment of FIG. 13, the combustion chamber 15 has upstream 151 and downstream 152 cavities without solid propellant and therefore comprises additional thermal protections 155.

The hybrid propellant 100 ’comprises a heat exchanger 80 allowing the boiling of the liquid propellant 40 and therefore its self-pressurization within the tank 30.

The vaporized liquid propellant 41 makes it possible to feed the combustion both in the combustion chamber 15 and the post-combustion chamber 16. The feed to the upper 31a and lower 31b injectors is controlled by a flow regulator and a double-way valve 38.

The lower injection system 75 makes it possible to pass the gases from the combustion chamber 15 to the post-combustion chamber 16, to support a series of lower injectors 31b and to pass the coolant heat transfer fluid 512 from the outer walls towards the heat exchanger 80, the latter being located in the center of the propellant below the tank 30 of liquid propellant.

The aerospike nozzle 50 is cooled by the circulation of the heat transfer fluid in its double walls 51 and 52 (cooling systems). The heat transfer fluid 512 transports thermal energy by sensible heat (heating) and possibly latent (vaporization). The heat transfer fluid is cooled in the heat exchanger and reinjected into its storage area. A pump 55 can be used to force the coolant to complete its cooling cycle.

Thus, the double block of solid propellant makes it possible to double the wetted surface without complicating the geometry (no multiple channels). The arrangement with the oxidant reservoir inside achieves high compactness and a length-to-diameter ratio compatible with a launcher stage application.

The solid propellant blocks provide total thermal protection on both walls (internal and external). Incomplete combustion in the combustion chamber also makes it possible to reduce thermal stresses at the end of operation when the solid propellant blocks are no longer thick enough to provide sufficient thermal protection. Incomplete combustion also offers the possibility of obtaining a desired or even optimal mixing ratio thanks to the second injection of liquid propellant vaporized into the post-combustion chamber. This architecture is not affected by the shift in mixing ratio due to the opening of the combustion chamber channel.

The injection of vaporized liquid propellant avoids combustion instabilities due to liquid interactions (droplets) with the gases in the combustion chambers. The injectors are also less complex than if they had to inject a fluid in the liquid state.

Claims

1. Hybrid thruster (100, 100 '), in particular for vehicle or spacecraft, comprising an outer body (10) in which is stored a solid propellant (20), extending in a longitudinal direction X of the outer body, a tank pressurized (30) containing a liquid or gaseous propellant (40), and a nozzle (50) for ejecting the gases produced from the combustion of the propellants, characterized in that it comprises a plurality of propellant injectors (31) liquid or gas arranged axially between parts of the solid propellant, and in that the solid propellant comprises at least one hollow cylindrical block,

2. Propellant according to claim 1, wherein injectors (31) of liquid or gaseous propellant are arranged along an internal surface (21) of the solid propellant (20).

3. Propellant according to claim 2, wherein the pressurized tank (30) is placed inside the outer body (10), surrounded by the solid propellant (20), and in which the injectors (31) are arranged on the long and around said reservoir.

4. Thruster according to claim 3, wherein the injectors (31) are arranged uniformly in longitudinal rows, along a longitudinal axis of the pressurized tank (30), and in radial rows, with respect to said longitudinal axis, the radial rows being equidistant.

5. Thruster according to any one of the preceding claims, in which the outer body (10) and the pressurized tank (30) are coaxial.

6. Thruster according to any one of the preceding claims, in which the liquid or gaseous propellant (40) contained in the pressurized tank (30) is sent to the injectors (31) via a flow control device (32) and pipes (33).

7. Thruster according to any one of the preceding claims, wherein the nozzle (50) is of the aerospike type and comprises a central body (51) integral with the pressurized tank (30) and an annular body (52) integral with the outer body ( 10) of said propellant.

8. Thruster according to claim 7, wherein the central body (51) of the nozzle (50) is connected to the pressurized tank (30) by means of a connection (82) ball and / or slide and actuators (83 ), of the jacks type, so as to allow said nozzle to be maneuvered by modifying the section of its neck and / or the inclination of said central body with respect to said longitudinal axis.

9. Thruster according to any one of the preceding claims, wherein the pressurized tank (30) is fixed to the outer body (10) by means of connecting rods (81) at its lower end, said connecting rods being arranged radially with respect to a longitudinal axis of said reservoir.

10. A thruster according to any preceding claim, wherein the ergo! solid (20) is based on a polymer or dre of paraffin, and liquid propellant (40) is cryogenic liquid oxygen, hydrogen peroxide or nitrous oxide.

11. A launcher type aerospace vehicle characterized in that it comprises a hybrid propellant (100) according to one of the preceding claims.