Method for igniting a mixture of an oxidizer and a fuel in a combustion chamber of a space launch vehicle

Abstract

The invention relates to a method for initiating or igniting a mixture of an oxidizer and a fuel in a combustion chamber of a space launcher, characterized in that it comprises implementing a step of injecting decomposed hydrogen peroxide (HTP) via a catalyst so as to obtain a predetermined ignition temperature. Figure to be published with the abstract: Fig. 1

Description

Title of the invention: Method for igniting a mixture of an oxidizer and a fuel in a combustion chamber of a space launcher

[0001] The present invention relates to a method for initiating the combustion of an oxidizer, in particular liquid oxygen called LOX, and a fuel, in particular RP1, taking place in a combustion chamber of a space launcher.

[0002] In the space sector, an ignition system ignites the mixture generated by the injection system. For example, ignition systems for oxidizer and fuel mixtures in a combustion chamber of a space launcher are known to include a GOX / CH4 flare. However, this type of flare is connected to two tanks: a tank of gaseous oxygen (GOX) and a tank of methane (CH4), resulting in an additional vessel with added mass and bulk.

[0003] The invention aims to remedy this problem by proposing a new embodiment. SUBJECT OF THE INVENTION

[0004] To this end, and according to a first aspect, the invention proposes a method for initiating combustion or igniting a mixture of an oxidizer and a fuel in a combustion chamber of a space launcher characterized in that it includes the implementation of a step of injecting hydrogen peroxide, called HTP, decomposed via a catalyst, so as to obtain a predetermined ignition temperature, called the combustion initiation temperature.

[0005] Thus, it is possible to obtain combustion initiation efficiently. The decomposition of HTP through a catalytic bed provides the energy necessary to reach the ignition conditions of an LOX and RP1 mixture. In the case of RP1 as fuel, it is necessary to reach a temperature of 577 Kelvin to achieve ignition.

[0006] Preferably, the fuel used by the engine is RP1 and the oxidizer used is liquid oxygen, also known as LOX.

[0007] RP1 (rocket petroleum 1) refers to the fuel known to those skilled in the art in the space sector: a form of kerosene specially refined for use as a storable liquid fuel for space launch vehicles. In particular, this fuel is a mixture of several alkanes and aromatics, for example including methyl biphenyl (with %M = 17.4%), n-heptylcyclopentane (with %M = 45.4%), n-tridecane (with %M = 37.2%).

[0008] Preferably, the process provides for a jet of HTP emitted into the combustion chamber at a temperature of at least 1000 kelvins, preferably at least 1400 kelvins. In one embodiment, the temperature is approximately 1473 kelvins. This jet makes it possible to obtain a droplet temperature of RP1 of at least 500 kelvins, preferably at least 580 kelvins.

[0009] The injection can have the following parameters: flow rate between 25 and 40 g / s, and a pressure in the torch chamber between 15 and 30 bar. According to a particular embodiment, the injection parameters can be: flow rate 33 g / s, duration 3 seconds, pressure in the torch chamber of 9.63 bar.

[0010] Preferably, the catalyst can be Manganese Oxide (MnO2) pellets.

[0011] It may be provided that only HTP is used.

[0012] According to one embodiment, at least one additive is added to improve the torch flame temperature, in particular to make the decomposition of HTP more energetic. In particular, the additive is an alcohol, for example ethanol.

[0013] The decomposition of HTP is an external ignition source so that the torch provides the energy necessary to ignite a LOx / RPl mixture. The decomposition of HTP alone may be sufficient to ignite the propellant mixture in the engine's combustion chamber. The addition of additional RPl is not planned.

[0014] According to a second aspect, the invention also relates to an ignition system for igniting a mixture of an oxidizer and a fuel in a combustion chamber of a space launcher comprising or being connected to:

[0015] - a single reservoir containing hydrogen peroxide, known as HTP,

[0016] - a catalyst that decomposes hydrogen peroxide, known as HTP,

[0017] - a torch chamber having a hydrogen peroxide injection orifice decomposed.

[0018] Preferably, the injection orifice may have an orifice between 5 and 9 millimeters.

[0019] Preferably, the torch chamber has an internal diameter of between 20 and 30 millimeters, preferably between 25 and 29 millimeters, a height of between 40 and 70 millimeters, preferably between 46 and 50 millimeters, an angular orientation of said system with respect to the fixing plane of the combustion chamber of between 80 and 100 degrees, preferably the system being located at a distance of at least 70 millimeters from the orifice of the injection system, in particular of the "pintle" or needle type.

[0020] For example, the so-called "pintle" injector can be a propellant injector for a dual-propellant or dual-fuel rocket engine. It ensures proper flow and fuel mixing when the fuels are forcefully injected under high pressure. Pressure in the combustion chamber is maintained so that an efficient and controlled combustion process can occur. In particular, the needle injector can be a type of coaxial injector. It consists of two concentric tubes and a central protrusion. Fuel A (usually the oxidizer) flows through an outer tube, exiting as a cylindrical jet, while fuel B (usually the fuel) flows into an inner tube and strikes a central needle-shaped protrusion (similar in shape to a mushroom valve like those found on four-stroke engines), spraying into a wide cone or flat sheet that cuts off the cylindrical jet of fuel A.

[0021] Preferably, the ignition system includes a device for controlling the flow rate of HTP injection into the catalyst, for example of the injection plate type.

[0022] The ignition system is arranged to implement the process

[0023] According to a third aspect, the invention also relates to a space launcher engine comprising an ignition system including one or more of the features of the second aspect. Preferably, the engine includes a "pintle" type injection system known to those skilled in the art. BRIEF DESCRIPTION OF THE FIGURES

[0024] Other features and advantages of the invention will become apparent from the detailed description of the invention which follows with reference to the accompanying figures and in which:

[0025] [Fig.1] Fig.1 represents a cross-sectional view of a combustion chamber of a space launcher engine, comprising at least part of a "pintle" type injection system and at least part of an ignition system;

[0026] [Fig.2] Fig.2 represents an arrangement of a torch connected to a chamber of combustion according to a particular embodiment.

[0027] For greater clarity, identical or similar elements of the different embodiments are identified by identical reference signs throughout the figures. DETAILED DESCRIPTION OF THE INVENTION

[0028] With reference to Figures 1 and 2, an embodiment of an ignition system or torch is described. Figure 1 shows the arrangement of an ignition system 1 and an injection system 2 whose orifices open into the combustion chamber 3.

[0029] The ignition system comprises:

[0030] - a single reservoir containing hydrogen peroxide, also known as HTP,

[0031] - a catalyst that decomposes hydrogen peroxide, known as HTP,

[0032] - a torch chamber having a hydrogen peroxide injection orifice decomposed and

[0033] - a device for controlling the HTP injection flow rate into the catalyst, by example of the injection plate type.

[0034] Figure 2 shows a particular embodiment in terms of dimensions and positioning relative to the combustion chamber and the pintle-type injection system. For example, the ignition system has an orifice diameter of 7 mm, an inner torch diameter of 27.6 mm, and a torch chamber height of 48.6 mm. Preferably, the orifice of the ignition system is located at a distance of 70 mm. The angular orientation of said system with respect to the mounting plane of the combustion chamber is approximately 95 degrees.

Claims

Demands

1. A method for initiating or igniting a mixture of an oxidizer and a fuel in a combustion chamber of a space launcher characterized in that it comprises implementing a step of injecting decomposed hydrogen peroxide (HTP) via a catalyst so as to obtain a predetermined ignition temperature.

2. The method according to claim 1, wherein the fuel used is RP1 and the oxidant used is liquid oxygen, also known as LOX.

3. A method according to claim 1 or 2, wherein a jet of HTP is emitted into the combustion chamber at a temperature of at least 1000 kelvins, preferably at least 1400 kelvins.

4. A method according to any one of claims 1 to 3, wherein only HTP is used.

5. A process according to any one of claims 1 to 3 wherein at least one additive, preferably alcohol, for example ethanol, is added.

6. A method according to any one of the preceding claims, wherein the injection has the following parameters: flow rate between 25 and 40 g / s, and a pressure in the torch chamber between 15 and 30 bar.

7. Ignition system for igniting a mixture of an oxidizer and a fuel in a combustion chamber of a space launcher comprising or being connected to: - a single tank containing hydrogen peroxide called HTP, and - a catalyst decomposing the hydrogen peroxide called HTP, - a torch chamber having an injection orifice for the decomposed hydrogen peroxide.

8. System according to the preceding claim comprising an orifice diameter is between 5 and 9 millimeters.

9. A system according to the preceding claim comprising an inner diameter of the torch chamber of between 20 and 30 millimeters, preferably between 25 and 29 millimeters, a height of between 40 and 70 millimeters, preferably between 46 and 50 millimeters, an angular orientation of said system with respect to the mounting plane of the combustion chamber of between 80 and 100 degrees, of 6. Preferably the system located at a distance of at least 70 millimeters from the orifice of the injection system.

10. Space launcher engine comprising an ignition system according to any one of claims 7 to 9 and an injection system of the needle injector type.