Hybrid thruster for space vehicle

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

2020368037   12 Oct 2026 injectors 71 and placed inside said external body, along the block of solid propellant 20. Figure 13 shows a hybrid propulsion unit 100’ according to another embodiment, wherein the combustion chamber 15 has an annular shape with two 5 slabs of solid propellant 20a and 20b attached respectively to an internal wall and an external wall of said chamber. The internal wall to which the first slab 20a of solid propellant is attached delimits the central reservoir 30 of liquid propellant. The liquid propellant 40 is self-pressurised by boiling, the heat necessary being provided by the heat-transfer fluid 512 that cools the walls of the nozzle 50 10 in accordance with a closed thermodynamic cycle. According to this embodiment, the propulsion unit 100’ comprises an injection dome 13 making it possible to provide gasified liquid propellant 41 coming from the top part of the reservoir 30 in the combustion chamber 15. Thus, a rich-mixture combustion (excess of solid propellant and lack of 15 liquid propellant) takes place in the combustion chamber 15 and vaporises the solid propellant without producing very hot combustion gases. The gases produced next pass through a bottom injection system 75 that provides the missing gasified liquid propellant for making the combustion complete. The mixing takes place in a combustion post-chamber 16 delimited by the walls of the nozzle 20 50. Furthermore, the cylindrical slabs of solid propellant 20a and 20b thermally protect the walls of the combustion chamber to which they are attached. According to the embodiment in figure 13, the combustion chamber 15 has upstream 151 and downstream 152 cavities without any solid propellant and 25 therefore includes additional thermal protections 155. The hybrid propulsion unit 100’ comprises a heat exchanger 80 enabling the liquid propellant 40 to boil and therefore to self-pressurise within the reservoir 30. The vaporised liquid propellant 41 supplies the combustion both in the combustion chamber 15 and in the combustion post-chamber 16. The supply to 30 the top 31a and bottom 31b injectors is controlled by a flow-rate regulator and a two-way valve 38. 2020368037   12 Oct 2026 The bottom injection system 75 makes it possible to convey the gases coming from the combustion chamber 15 to the combustion post-chamber 16, to support a series of bottom injectors 31b and to pass the cooling heat-transfer fluid 512 coming from the external walls to the heat exchanger 80, the latter being 5 located at the centre of the propulsion unit below the reservoir 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 the thermal energy by sensitive heat (heating) and optionally latent heat 10 (vaporisation). The heat-transfer fluid is cooled in the heat exchanger and reinjected into its storage zone. A pump 55 can serve to force the heat-transfer fluid to make its cooling cycle. Thus, the double slab of solid propellant multiplies by two the wetted surface area without making the geometry more complex (no multiple channels). 15 The arrangement with the reservoir of oxidant inside makes it possible to obtain high compactness and a length to diameter ratio compatible with a launch stage application. The slabs of solid propellant provide total thermal protection over the two walls (internal and external). The incomplete combustion in the combustion 20 chamber also reduces the thermal stresses at the end of operation when the slabs of solid propellant are no longer thick enough to provide sufficient thermal protection. The incomplete combustion also offers the possibility of obtaining a required or even optimum mixing ratio by virtue of the second injection of vaporised liquid 25 propellant into the combustion post-chamber. This architecture is not affected by the shift in ratio of the mixing due to the opening of the channel of the combustion chamber. Injecting vaporised liquid propellant avoids the combustion instabilities due to the liquid interactions (droplets) with the gases in the combustion chambers. 30 The injectors are also less complex than if they had to inject a fluid in the liquid state. 2020368037   12 Oct 2026 In the claims which follow and in the preceding description of the invention, except where the context requires otherwise due to express language or necessary implication, the word “comprise” or variations such as “comprises” or “comprising” is used in an inclusive sense, i.e. to specify the presence of the 5 stated features but not to preclude the presence or addition of further features in various embodiments of the invention.

Claims

1. Hybrid propulsion unit, in particular for a space vehicle or rocket, comprising an external body in which a solid propellant is stored, extending in a longitudinal direction X of the external body, a pressurised reservoir containing a liquid propellant, and a nozzle for ejecting the gases produced from the combustion of the propellants, wherein it includes a plurality of liquid- or gaseous-propellant injectors arranged axially between parts of the solid propellant, and in that the solid propellant comprises at least one hollow cylindrical block; wherein the liquid propellant is self-pressurised by boiling, the heat necessary being provided by a heat-transfer fluid that cools the walls of the nozzle in accordance with a closed thermodynamic cycle; wherein the hybrid propulsion unit comprises a heat exchanger enabling the liquid propellant to boil and therefore to self-pressurise within the reservoir; and wherein the injectors provide gasified liquid propellant.

2. Propulsion unit according to claim 1, wherein liquid- or gaseous-propellant injectors are arranged along an internal surface of the solid propellant.

3. Propulsion unit according to claim 2, wherein the pressurised reservoir is placed inside the external body, surrounded by the solid propellant, and wherein the injectors are disposed along and around said reservoir.

4. Propulsion unit according to claim 3, wherein the injectors are disposed uniformly in longitudinal rows, along a longitudinal axis of the pressurised reservoir, and in radial rows, with respect to said longitudinal axis, the radial rows being equidistant.

5. Propulsion unit according to any one of the preceding claims, wherein the external body and the pressurised reservoir are coaxial.

6. Propulsion unit according to any one of the preceding claims, wherein the liquid or gaseous propellant contained in the pressurised reservoir is sent into the injectors via a flow-rate control device and pipes.2020368037   12 Oct 20267. Propulsion unit according to any one of the preceding claims, wherein the nozzle is of the aerospike type and includes a central body secured to the pressurised reservoir and an annular body secured to the external body of said propulsion unit.

8. Propulsion unit according to claim 7, wherein the central body of the nozzle is connected to the pressurised reservoir by means of a swivel and / or slider connection and actuators, of the jack type, so as to make it possible to manoeuvre said nozzle by modifying the cross section of its neck and / or the inclination of said central body with respect to said longitudinal axis.

9. Propulsion unit according to any one of the preceding claims, wherein the pressurised reservoir is attached to the external body by means of connecting rods at its bottom end, said connecting rods being arranged radially with respect to a longitudinal axis of said reservoir.

10. Propulsion unit according to any one of the preceding claims, wherein the solid propellant is based on a polymer or a paraffin wax, and the liquid propellant is cryogenic liquid oxygen, hydrogen peroxide or nitrogen monoxide.

11. Aerospace vehicle of the launch vehicle type, wherein it includes a hybrid propulsion unit according to one of the preceding claims.

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