ROCKET ENGINE WITH GROUND-BASED IGNITION

Abstract

The present disclosure relates to a rocket engine obtaining safer and better controlled ground-based ignition, the rocket engine comprising an axisymmetric propulsion chamber (12), including a throat (12c) at which the diameter of the propulsion chamber (12) is a minimum, an injection head (11) configured to inject at least one liquid propellant into the propulsion chamber (12), and a destructible tubular guide (40), applied coaxially in the propulsion chamber (12) so as to channel said propellant downstream of the throat (12c) of the propulsion chamber (12).

Claims

1. A rocket engine, comprising an axisymmetric propulsion chamber, including a throat at which the diameter of the propulsion chamber is a minimum, an injection head configured to inject at least one liquid propellant into the propulsion chamber, and a destructible tubular guide, applied coaxially in the propulsion chamber so as to channel said propellant downstream of the throat of the propulsion chamber, at least one portion of the guide being configured to resist at least 2 seconds at a temperature of 1700 C.

2. The rocket engine according to claim 1, wherein the guide is axisymmetric.

3. The rocket engine according to claim 1, wherein the guide is attached in a sealed fashion to the throat of the propulsion chamber.

4. The rocket engine according to claim 1, comprising a divergent nozzle connected to the downstream end of the propulsion chamber, wherein the guide extends into the divergent nozzle over at most 20% of the length of the divergent nozzle.

5. The rocket engine according to claim 1, wherein the guide is configured to tear when a pressure greater than 5 bar is exerted on its interior face.

6. The rocket motor according to claim 1, wherein the guide extends from the throat of the propulsion chamber.

7. The rocket motor according to claim 1, wherein the propulsion chamber is devoid of an internal ignition device.

8. An assembly comprising a rocket engine according to claim 1 and a launch pad, the rocket engine being positioned on the launch pad, wherein the launch pad comprises at least one ignition torch configured to project a flame toward the interior space of the guide of the rocket engine.

9. The assembly according to claim 8, wherein the rocket engine comprises a divergent nozzle connected to the downstream end of the propulsion chamber, and wherein said ignition torch points toward a zone situated along the axis of the rocket engine and not situated lower than 20% of the length of the divergent nozzle of the rocket engine starting from the upstream end of the divergent nozzle.

10. The assembly according to claim 8, wherein said ignition torch does not penetrate into the interior of the rocket engine.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0039] The appended drawings are schematic and aim primarily to illustrate the principles of the invention.

[0040] FIG. 1 is a section plan of a rocket engine according to the invention.

[0041] FIG. 2 is a section plan of a rocket engine according to the prior art.

DETAILED DESCRIPTION OF EXEMPLARY EMBODIMENT(S)

[0042] In order to make the invention more concrete, an example of a rocket engine is described in detail hereafter, with reference to the appended drawings. It is recalled that the invention is not limited to this example.

[0043] FIG. 1 shows, in section along a vertical plane passing through its main axis A, a rocket motor 1 according to the invention. It includes, from upstream to downstream, an injection head 11, a propulsion chamber 12 and a divergent nozzle. This assembly is mounted on a launcher 2 by means of a gimbal hinge 21, mounted at its upper end, and lateral rams 22.

[0044] The propulsion chamber 12, symmetrical about the axis A of the engine 1, is connected by its upstream end to the injection head 11. The diameter of the propulsion chamber then decreases downstream until it reaches a minimum, forming a throat 12c, before increasing again until the downstream end of the propulsion chamber. This downstream end is connected to the divergent nozzle 13 of the engine 1.

[0045] During the normal operation of the engine 1, in steady-state operation, the injection head 1 injects a mixture of two liquid propellants into the upstream portion 12m of the propulsion chamber 12; the combustion of the propellants then occurs in this portion 12m of the propulsion chamber 12, forming a combustion chamber, and generates a considerable quantity of combustion gases ejected downstream at high speed; passing the throat 12c of the propulsion chamber allows the combustion gases to be accelerated while its downstream portion 12v and the divergent nozzle 13 ensure the expansion of the combustion gases prior to the ejection at the downstream end of the divergent nozzle 13, thus ensuring opposite thrust directed upward and allowing the launcher to be propelled.

[0046] During preparation for launch, the launcher 2 is placed on the launching pad 3 of a firing point, this launching pad 3 being equipped with ignition torches 31 which do not penetrate inside the rocket engine 1.

[0047] In addition, a tubular guide 40, cylindrical with a circular base in this case, the diameter of which is substantially equal to that of the throat 12c of the propulsion chamber 12, is mounted in the rocket engine 1 coaxially with the main axis A. This tubular guide 40 has two portions connected to one another by sealed means: an upstream portion 41, made of plastic, and a downstream portion 42, made of a thermally insulating material comprising silica fibers.

[0048] The upstream end of the tubular guide 40 is attached to the throat 12c of the propulsion chamber 12, for example by means of adhesive tape. It thus extends into the downstream portion 12v of the propulsion chamber 12 from the throat 12c and penetrated in part into the divergent nozzle 13. In this example, the downstream end of the tubular guide 40 is situated at approximately 15% of the length of the divergent nozzle 13, starting from the upstream end of the divergent nozzle 13.

[0049] The ignition torches 31 are then oriented so as to point toward the downstream portion 42 of the tubular guide 40.

[0050] To ignite the engine 1, the two liquid propellants are injected by means of the injection head 11; the latters then flow into the upstream portion 12m of the propulsion chamber 12, then into the tubular guide 40 without being dispersed laterally. The ignition torches 31 are then lit and each projects a flame in the direction of the interior volume of the downstream portion 42 of the tubular guide 40, i.e. in the direction of the main axis A of the engine 1 and at the interface 15 between the propulsion chamber 12 and the divergent nozzle 13. This step can last for one to two seconds before the propellants fire: the thermally insulating material of the downstream portion 42 of the tubular guide 40 then allows the latter to resist until the effective ignition of the engine.

[0051] Thus, thanks to the tubular guide 40, the ignition of the propellants occurs in a controlled ignition zone 50 situated in the interior of the tubular guide 40, rather precisely on the axis of symmetry A of the rocket engine 1, which causes symmetrical, and therefore relatively weak, lateral loads 51.

[0052] During detonation, the tubular guide 40 tears and is ejected out of the rocket engine 1 by the detonation blast. Once ignition occurs, the heat and the speed of the combustion gases allow the destruction and/or ejection of possible residues of the tubular guide 40 and of its means of attachment.

[0053] The embodiments or exemplary embodiments described in the present disclosure are given by way of illustration and are not limiting, a person skilled in the art being able to easily, upon seeing this disclosure, modify these embodiments or exemplary embodiments, or conceive others, while remaining within the scope of the invention.

[0054] Moreover, the different features of these embodiments or exemplary embodiments can be used alone or be combined together. When they are combined, these features can be combined as well as described above or differently, the invention not being limited to the specific combinations described in the present disclosure. In particular, unless otherwise specified, a features described in relation with an embodiment or exemplary embodiment can be applied analogously to another embodiment or exemplary embodiment.