CN211819717U - Rocket engine combustion chamber and rocket engine - Google Patents

Abstract

The utility model provides a rocket engine combustion chamber and rocket engine. The rocket engine combustion chamber includes: the main body is provided with a primary combustion chamber and a secondary combustion chamber, the secondary combustion chamber is arranged at the outlet end of the primary combustion chamber, a flow guide structure is arranged at the junction of the primary combustion chamber and the secondary combustion chamber, and the flow guide structure is used for guiding a propellant at the inner wall surface of the primary combustion chamber to the central position of the main body; an injector disposed at an inlet end of the primary combustion chamber; and the spray pipe is arranged at the outlet end of the secondary combustion chamber. The utility model discloses a rocket engine combustion chamber can avoid fuel or oxidant in the coolant to be heated under the heat of center high temperature gas and decompose and partial combustion produces chemical reaction to the one-level combustion chamber inner wall face, avoids the combustion chamber wall phenomenon such as oxidation, ablation to appear.

Description

Rocket engine combustion chamber and rocket engine

Technical Field

The utility model relates to a rocket driving system technical field particularly, relates to a rocket engine combustion chamber and rocket engine.

Background

The traditional liquid rocket engine mainly comprises an electromagnetic valve and a thrust chamber, wherein an injector sprays fuel and oxidant at a high speed, and the fuel and the oxidant are mixed and combusted in the thrust chamber. There are currently a number of proven fuel and oxidant combinations, each having specific properties. They also release different amounts of energy during combustion and also different thermodynamic reactions. All propellant combinations are exothermic, i.e., the fuel and oxidizer will combine to release a large amount of heat, depending on the nature of their reaction. The efficiency of the propellant reaction is particularly important, i.e. the efficiency of the reaction of the two propellant components depends to a large extent on the sufficiency of mixing of the two components. Insufficient mixing can result in wasted engine performance in a limited space.

Injectors for rocket engines are important components for mixing fuel and oxidant and for combustion. A typical injector may consist of several to several thousand orifices through which fuel and oxidant are injected into the combustion chamber. The fuel and the oxidant are jetted from the jet orifice to form a spray fan, and are mixed and combusted at a certain position in the combustion chamber body.

The combustion of the propellant produces a gas temperature that generally exceeds the melting temperature of most known materials. If cooling technology is not used, the combustion chamber wall can burn through and melt, and due to metal connection between the combustion chamber and the injector, heat is transferred to the injector through the combustion chamber wall, so that the injector is overheated, and air resistance of the injector can be caused, and the engine can explode seriously.

At present, two methods are mainly used for cooling the wall of the combustion chamber of the mainstream rocket engine. The first is regenerative cooling, in which a coolant is circulated through cooling channels in the combustion chamber walls, and the coolant may be one or two propellants which act as a coolant and remove heat. The utility of this approach is limited in its use, it does not provide adequate cooling for small engines because their propellant flow is too small, and for other reasons, this approach may not be suitable for use with large high pressure engines because cooling needs to flow at a rate in the cooling channels, thus resulting in injection inlet pressures that are too high and causing significant disadvantages to the front end supply system.

Another method is to use liquid film cooling, where the holes of the propellant injectors are ejected, where the central area of the injector ejects a central combustion stream and the side areas of the injector eject cooling streams that cool the walls of the combustion chamber. In the central region, the propellant is well mixed and combusted through the central orifice of the injector. The boundary coolant is sprayed to a certain angle to the wall surface of the combustion chamber through the peripheral spray holes of the injector. The propellant forms a liquid film on the wall of the combustion chamber to separate the wall of the combustion chamber from the central high-temperature gas. The fuel or oxidant ejected from the injector absorbs heat by evaporation of the droplets, thereby lowering the temperature of the wall of the combustion chamber. This approach can be used for most rocket engines, but results in significant engine performance loss due to the coolant not participating in combustion.

The liquid film is exposed to high temperature combustion gas, and the coolant is evaporated and decomposed by the transfer of heat. The decomposition products react with the unreacted propellant in the central region to form various unknown reactants (due to the widely varying concentrations of fuel and oxidizer) that may chemically react with typical chamber wall materials such as copper, nickel, platinum, iridium, gold, rhenium, and copper, which can corrode and oxidize the combustion chamber walls and cause their failure.

The cooling of the liquid film may also result in localized oxidant or fuel concentrations in the combustion chamber. These differences in concentration can result in the fuel and oxidant producing unknown chemical components that, when in contact with the combustion chamber walls, can cause oxidation, ablation, etc. of the combustion chamber walls. Meanwhile, the injector hole is influenced, so that the phenomena of expansion, cracks and the like of the injector hole occur.

SUMMERY OF THE UTILITY MODEL

The utility model discloses a main aim at provides a rocket engine combustion chamber and rocket engine to solve the problem that oxidation, ablation appear easily in the rocket engine combustion chamber among the prior art.

In order to achieve the above object, according to an aspect of the present invention, there is provided a rocket engine combustion chamber, including: the main body is provided with a primary combustion chamber and a secondary combustion chamber, the secondary combustion chamber is arranged at the outlet end of the primary combustion chamber, a flow guide structure is arranged at the junction of the primary combustion chamber and the secondary combustion chamber, and the flow guide structure is used for guiding a propellant at the inner wall surface of the primary combustion chamber to the central position of the main body; an injector disposed at an inlet end of the primary combustion chamber; and the spray pipe is arranged at the outlet end of the secondary combustion chamber.

Further, the inner wall surface of the outlet end of the first-stage combustion chamber is provided with the flow guide structure.

Further, the flow guide structure is an annular baffle, and the central axis of the annular baffle is consistent with the central axis of the main body.

Furthermore, the primary combustion chamber and the secondary combustion chamber are connected through a connecting flange or in a welding mode after being arranged in a split mode.

Further, the injector is connected with the primary combustion chamber through a connecting flange or in a welding mode.

Furthermore, the injector is provided with a plurality of first injection holes and a plurality of second injection holes, the injection direction of the first injection holes is consistent with the axis direction of the main body, the second injection holes surround the periphery of the second injection holes, and the injection direction of the second injection holes faces the inner wall surface of the primary combustion chamber.

Further, the nozzle and the secondary combustion chamber are integrally formed.

Further, the height of the flow guide structure higher than the inner wall surface of the primary combustion chamber is 0.76-2.6 mm.

According to another aspect of the present invention, there is provided a rocket engine comprising a combustion chamber, said combustion chamber being as defined above.

Use the technical scheme of the utility model, the utility model provides a coolant inwards deflects under the effect of water conservancy diversion structure and mixes it with high temperature center gas torrent, makes all chemical components of coolant and high temperature center gas contact, can make all propellant burn basically completely, then enters into the second grade combustion chamber, avoids discharging the propellant of complete combustion, and then furthest improves the efficiency of combustion process.

Because of the inward deflection of the coolant, the inner wall of the secondary combustion chamber is mainly contacted with a completely combusted propellant, and high-temperature fuel gas cannot chemically react with many materials such as molybdenum, copper, nickel, gold, platinum, iridium, rhenium and the like. Therefore, the inner wall of the secondary combustion chamber can be made of high-thermal-conductivity materials, and the thin-wall structure provides the required high heat transfer rate, so that the corrosion rate of the wall of the combustion chamber is compensated without increasing the thickness of the wall of the combustion chamber, and the production cost of the combustion chamber of the rocket engine can be effectively reduced.

In addition to the above-described objects, features and advantages, the present invention has other objects, features and advantages. The present invention will be described in further detail with reference to the drawings.

Drawings

The accompanying drawings, which form a part of the specification, are included to provide a further understanding of the invention, and are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and together with the description serve to explain the invention without unduly limiting the scope of the invention. In the drawings:

figure 1 shows schematically a cross-sectional view of a rocket engine combustion chamber according to the invention.

Wherein the figures include the following reference numerals:

10. a main body; 11. a primary combustion chamber; 12. a secondary combustion chamber; 13. a flow guide structure; 20. an injector; 21. a first injection hole; 22. a second injection hole; 30. and (4) a spray pipe.

Detailed Description

It should be noted that, in the present invention, the embodiments and features of the embodiments may be combined with each other without conflict. The present invention will be described in detail below with reference to the accompanying drawings in conjunction with embodiments.

In order to make the technical solution of the present invention better understood, the technical solution of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments in the present invention, all other embodiments obtained by a person skilled in the art without creative efforts shall belong to the protection scope of the present invention.

It should be noted that the terms "first," "second," and the like in the description and claims of the present invention and in the drawings described above are used for distinguishing between similar elements and not necessarily for describing a particular sequential or chronological order. It is to be understood that the terms so used are interchangeable under appropriate circumstances for describing embodiments of the invention herein. Furthermore, the terms "comprises," "comprising," and "having," and any variations thereof, are intended to cover a non-exclusive inclusion, such that a process, method, system, article, or apparatus that comprises a list of steps or elements is not necessarily limited to those steps or elements expressly listed, but may include other steps or elements not expressly listed or inherent to such process, method, article, or apparatus.

It is noted that the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of example embodiments according to the present application. As used herein, the singular forms "a", "an" and "the" are intended to include the plural forms as well, and it should be understood that when the terms "comprises" and/or "comprising" are used in this specification, they specify the presence of stated features, steps, operations, devices, components, and/or combinations thereof, unless the context clearly indicates otherwise.

Referring to fig. 1, according to an embodiment of the present invention, there is provided a rocket engine combustion chamber, which in this embodiment includes a main body 10, an injector 20, and a nozzle 30.

The main body 10 comprises a primary combustion chamber 11 and a secondary combustion chamber 12, the secondary combustion chamber 12 is arranged at the outlet end of the primary combustion chamber 11, a flow guide structure 13 is arranged at the junction of the primary combustion chamber 11 and the secondary combustion chamber 12, and the flow guide structure 13 is used for guiding the propellant on the inner wall surface of the main body 10 to the central position of the main body 10; the injector 20 is arranged at the inlet end of the primary combustion chamber 11; a nozzle 30 is provided at the outlet end of the secondary combustion chamber 12.

In operation, propellant is injected into the primary combustion chamber 11 of the main body 10 through the injector 20, high temperature central gas is formed in the primary combustion chamber 11, and the combustion position of the propellant in the main body 10 is related to the design of the injector 20.

Specifically, the injector 20 in the present embodiment is provided with a plurality of first injection holes 21 and a plurality of second injection holes 22, the injection direction of the plurality of first injection holes 21 coincides with the axial direction of the primary combustion chamber 11, the second injection holes 22 are provided around the outer periphery of the second injection holes 22, and the injection direction of the second injection holes 22 is provided toward the inner wall surface of the primary combustion chamber 11. The injection direction herein is directed toward the inner wall surface of the primary combustion chamber 11, which means that the propellant injected from the second injection holes 22 is directly injected toward the inner wall surface of the primary combustion chamber 11.

The propellant entering from the second injection orifice 22 is a coolant, i.e. a fuel or an oxidizer. The propellant entering from the first injection holes 21 is a high temperature fuel. The propellant entering from the second injection hole 22 is continuously transmitted forward along the inner wall of the primary combustion chamber 11, a cooling boundary layer is formed near the inner wall surface of the primary combustion chamber 11, the fuel or oxidant in the coolant is prevented from being heated and decomposed under the heat of central high-temperature fuel gas and being partially combusted to generate chemical reaction on the inner wall surface of the primary combustion chamber 11, and the phenomena of oxidation, ablation and the like of the wall surface of the combustion chamber are avoided. This is because the coolant boundary layer is thermally decomposed and partially burned first near the central gas side, and the coolant thermal decomposition and combustion components temporarily do not affect the wall surface of the combustion chamber because the coolant moves in the axial direction.

The distance between the injector 20 and the inner surface of the flow guiding structure 13 is selected such that the coolant is radially deflected into the central combustion zone before it is completely decomposed or burnt, and is mixed and burnt in the central high temperature zone with the coolant boundary layer zone.

The utility model provides a coolant inwards deflects under the effect of water conservancy diversion structure 13 and mixes its and high temperature center gas torrent, makes all chemical composition of coolant and high temperature center gas contact, can make all propellant burn basically completely, then enters into second grade combustion chamber 12, avoids discharging the propellant of complete combustion not, and then furthest improves the efficiency of combustion process.

Due to the inward deflection of the coolant, the inner walls of the secondary combustion chamber 12 are primarily in contact with the fully combusted propellant, and the hot gases do not chemically react with many materials such as molybdenum, copper, nickel, gold, platinum, iridium, rhenium, etc. Therefore, the inner wall of the secondary combustion chamber 12 can be made of high-thermal-conductivity materials, and the thin-wall structure provides the required high heat transfer rate, so that the corrosion rate of the wall of the combustion chamber is compensated without increasing the thickness of the wall of the combustion chamber, and the production cost of the combustion chamber of the rocket engine can be effectively reduced.

It can be seen that, the utility model discloses a rocket engine combustion chamber during operation, in one-level combustion chamber 11, coolant parcel center gas to the restriction is from center high temperature gas to the inner wall transmission heat of one-level combustion chamber 11.

In a preferred embodiment of the present invention, the flow guiding structure 13 is disposed at the outlet end of the first-stage combustion chamber 11, so as to guide the coolant on the inner wall surface of the first-stage combustion chamber 11 to the center of the first-stage combustion chamber 11 for combustion. Of course, in other embodiments of the present invention, the diversion structure 13 can be disposed at the inlet end of the secondary combustion chamber 12, and other deformation manners under the concept of the present invention are all within the protection scope of the present invention.

Preferably, the flow guiding structure 13 in the present embodiment is an annular baffle plate having a central axis coinciding with a central axis of the main body 10, so as to guide the coolant at the inner wall surface of the primary combustion chamber 11 to a central position of the primary combustion chamber 11.

In other embodiments of the present invention, the ring-shaped baffle can be set as a round protrusion, if it is in other deformation modes under the concept of the present invention, all the deformation modes are within the protection scope of the present invention. When the flow guide structure 13 is provided as a circle of bulges, the flow guide effect is not as strong as that of the annular baffle plate, and partial coolant flows into the secondary combustion chamber 12 between every two adjacent bulges.

Referring to fig. 1 again, the first-stage combustion chamber 11 and the second-stage combustion chamber 12 in this embodiment are separated from each other, and can be connected by a connecting flange or by welding, and other deformation modes under the concept of the present invention are all within the protection scope of the present invention.

Similarly, the injector 20 and the primary combustion chamber 11 in this embodiment may be connected by a connecting flange or by welding.

The nozzle 30 and the secondary combustion chamber 12 are integrally formed, so that the structure is simple and the stability is strong.

Preferably, in the present embodiment, the height of the flow guiding structure 13 above the inner wall surface of the primary combustion chamber 11 is 0.76mm to 2.6 mm.

As described above, the present invention is divided into the first-stage combustion chamber 11 and the second-stage combustion chamber 12, which are coaxially disposed, in the main body 10 of the rocket engine combustion chamber. The inner wall of the primary combustion chamber 11 is in contact with the coolant phase only. Instead, the inner walls of the secondary combustion chamber 12 are only in contact with substantially completely combusted gas. Thus, the rocket engine cooling efficiency is greatly improved, the combustion chamber is protected from chemical corrosion by decomposition and combustion propellants, the combustion efficiency is significantly improved compared to similar cooling, and the respective combustion chamber walls can be made of different materials, reducing engine cost and weight.

Specifically introduce the utility model discloses a rocket engine combustion chamber's scheme as follows:

a. the primary combustion chamber 11 and the secondary combustion chamber 12 are axially connected;

b. the injector 20 injects fuel and oxidant in proportion to a first injection hole 21 in the central region of the primary combustion chamber 11, and simultaneously, a second injection hole 22 on the periphery of the injector 20 injects a certain amount of coolant (fuel or oxidant) to the inner wall surface of the primary combustion chamber 11;

c. the injected fuel and oxidant are proportioned, so that the fuel and oxidant are fully mixed and combusted, and the stable combustion of the central area is ensured. The inner wall surface of the first-stage combustion chamber 11 forms low-temperature flow, and the coolant is ensured to be not reacted or slightly reacted with the central combustion area as far as possible;

d. the distance control between the injector 20 and the flow guiding structure 13 needs to be determined by comprehensive consideration, such as the proportion of the coolant to the whole propellant, the design parameters (such as pressure, thrust, size and the like) of the engine and the physical parameters of the coolant; the maximum distance is to ensure that the coolant is not completely decomposed and combusted as an end point, and the design is relatively difficult because chemical combustion is complex to calculate; to improve design efficiency, a simplified calculation approach may also be employed. The temperature rise of the coolant is informed to be calculated to determine the distance, namely the decomposition temperature of the coolant is used as a terminal point, and the distance is further determined, and the design of the mode is conservative; or calculated using CFD fluid-combustion simulation software. Regardless of the method, the final distance is determined by a hot test;

e. the height of the flow-guiding structure 13 is also an important design parameter, and the height and shape are required to ensure that the coolant is guided into the central combustion zone before entering the secondary combustion chamber 12, and the height of the baffle ring is generally selected to be between 0.76mm and 2.6mm according to simulation calculation and experimental data.

f. The length of the primary combustion chamber 11 needs to be determined by the material of the injector 20 and the vaporization temperature of the propellant in the injector 20, so that the situation that the primary combustion chamber 11 is too short, the high temperature of the inner wall surface of the secondary combustion chamber 12 is transmitted to the panel of the injector 20, and the air resistance and ablation phenomena of the injector occur is prevented.

According to another aspect of the present invention, there is provided a rocket engine comprising a combustion chamber, the combustion chamber being as defined above.

From the above description, it can be seen that the above-mentioned embodiments of the present invention achieve the following technical effects:

the utility model discloses a rocket engine's combustor can restrain the chemical reaction between various chemical components and the one-level combustion chamber wall. Thus, the structural integrity of the combustion chamber wall is preserved. The utility model discloses a rocket engine combustion chamber has still improved the efficiency of propellant burning. The edge region coolant is deflected to the central combustion region from the combustion chamber wall radially inwards through the flow guide structure, so that the edge region coolant can be completely combusted with the central high-temperature fuel gas secondarily.

Due to the flow guide structure, the proportion of the coolant in the primary combustion chamber to the total propellant can be increased, and the concentration of the oxygen-fuel ratio in the central combustion area needs to be ensured to be stable in combustion. The utility model discloses a this can increase cooling efficiency, simultaneously when the coolant before not taking place to decompose or decompose on a small scale, the coolant mixes, the burning through water conservancy diversion structure and central high temperature gas. The height of the flow guide structure and the distance of the injector need to be matched with the design parameters of the engine and the combination of the propellant for detailed calculation, and some experimental parameters are given in the following experiments and only used as references.

The coolant separates the combustion chamber walls from the central combustion gases before the flow directing structure. Thus, combustor wall temperature and erosion are minimized. There are greater possibilities in the choice of materials between the injector and the baffle ring. Materials with lower thermal conductivity are generally adopted, so that the heat flow transmitted by high-temperature fuel gas behind the flow guide structure to the combustion chamber wall and the injector in front of the flow guide structure is reduced.

And the coolant of the combustor part at the downstream of the flow guide structure is fully combusted with the central high-temperature fuel gas. The decomposed components of the coolant are mixed and combusted in the center by high-temperature fuel gas. Because the wall of the combustion chamber at the downstream of the flow guide structure is in contact with high-temperature fuel gas, high-temperature and high-thermal conductivity materials can be selected, and the materials can be possibly corroded by decomposition components of the coolant and the like but can not be influenced by completely combusted components. The combustion chamber at the downstream of the flow guide structure mainly adopts a radiation cooling thin-wall structure which can increase the radiation transfer rate.

As demonstrated above, the present invention effectively cools the combustor wall temperature before the injector and flow guide structure and further inhibits heat flux to the injector by selecting a material with a low thermal conductivity. Thus, heat transfer through the interface between the combustion chamber wall and the injector may be limited and controlled to prevent the injector from experiencing vapor lock. The utility model discloses further keep apart one-level combustion chamber wall and coolant decomposition component or combustion component, central high temperature gas and one-level combustion chamber wall to prevent that coolant decomposition component or combustion component from oxidizing or corroding combustion chamber wall and one-level combustion chamber high temperature. The cooling agent of the present invention may be realized by an injector, and only one of the propellant components, for example, fuel or oxidizer, is used as the cooling agent. However, the invention also makes it possible to use different propellants as coolant which only have to be introduced into the coolant at different locations. Only the different propellants are required to be ensured not to be mixed with each other.

In the axial interface region before the flow guiding structure, there may be partial combustion of some coolant or decomposition of the coolant. However, the overall combustor wall erosion is significantly reduced as compared to the prior art coolant film technology because the baffle structure can divert the coolant to the central combustion zone for combustion before the coolant decomposes and the combustion components affect the combustor wall. Furthermore, the utility model has proved advantageous in improving the combustion efficiency due to the foregoing discussion.

Unless specifically stated otherwise, the relative arrangement of the components and steps, the numerical expressions, and numerical values set forth in these embodiments do not limit the scope of the present invention. Meanwhile, it should be understood that the sizes of the respective portions shown in the drawings are not drawn in an actual proportional relationship for the convenience of description. Techniques, methods, and apparatus known to those of ordinary skill in the relevant art may not be discussed in detail but are intended to be part of the specification where appropriate. In all examples shown and discussed herein, any particular value should be construed as merely illustrative, and not limiting. Thus, other examples of the exemplary embodiments may have different values. It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, further discussion thereof is not required in subsequent figures.

Spatially relative terms, such as "above … …," "above … …," "above … …," "above," and the like, may be used herein for ease of description to describe one device or feature's spatial relationship to another device or feature as illustrated in the figures. It will be understood that the spatially relative terms are intended to encompass different orientations of the device in use or operation in addition to the orientation depicted in the figures. For example, if a device in the figures is turned over, devices described as "above" or "on" other devices or configurations would then be oriented "below" or "under" the other devices or configurations. Thus, the exemplary term "above … …" can include both an orientation of "above … …" and "below … …". The device may be otherwise variously oriented (rotated 90 degrees or at other orientations) and the spatially relative descriptors used herein interpreted accordingly.

In the description of the present invention, it should be understood that the orientation or positional relationship indicated by the orientation words such as "front, back, up, down, left, right", "horizontal, vertical, horizontal" and "top, bottom" etc. are usually based on the orientation or positional relationship shown in the drawings, and are only for convenience of description and simplification of description, and in the case of not making a contrary explanation, these orientation words do not indicate and imply that the device or element referred to must have a specific orientation or be constructed and operated in a specific orientation, and therefore, should not be interpreted as limiting the scope of the present invention; the terms "inner and outer" refer to the inner and outer relative to the profile of the respective component itself.

The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims (9)

1. A rocket engine combustion chamber, comprising:

the main body (10) is provided with a primary combustion chamber (11) and a secondary combustion chamber (12), the secondary combustion chamber (12) is arranged at the outlet end of the primary combustion chamber (11), a flow guide structure (13) is arranged at the junction of the primary combustion chamber (11) and the secondary combustion chamber (12), and the flow guide structure (13) is used for guiding the propellant at the inner wall surface of the primary combustion chamber (11) to the central position of the main body (10);

an injector (20), said injector (20) being arranged at an inlet end of said primary combustion chamber (11);

a nozzle (30), the nozzle (30) being arranged at an outlet end of the secondary combustion chamber (12).

2. A rocket engine combustion chamber according to claim 1, characterized in that said flow guiding structure (13) is provided on the inner wall surface of the outlet end of said primary combustion chamber (11).

3. A rocket engine combustion chamber according to claim 1, characterized in that said flow guiding structure (13) is an annular baffle, the central axis of which coincides with the central axis of said body (10).

4. A rocket engine combustion chamber according to claim 1, characterized in that said primary combustion chamber (11) and said secondary combustion chamber (12) are separately arranged and then connected by means of connecting flanges or by means of welding.

5. A rocket engine combustion chamber according to claim 1, characterized in that said injector (20) is connected to said primary combustion chamber (11) by means of a connecting flange or by means of welding.

6. A rocket engine combustion chamber according to claim 1, wherein a plurality of first injection holes (21) and a plurality of second injection holes (22) are provided on the injector (20), the injection direction of the plurality of first injection holes (21) is consistent with the axial direction of the main body (10), the second injection holes (22) are arranged around the periphery of the second injection holes (22), and the injection direction of the second injection holes (22) is arranged towards the inner wall surface of the primary combustion chamber (11).

7. A rocket engine combustion chamber as recited in claim 1, wherein said nozzle (30) is provided integrally with said secondary combustion chamber (12).

8. A rocket engine combustion chamber as claimed in any one of claims 1 to 7, wherein the height of the flow guiding structure (13) above the inner wall surface of the primary combustion chamber (11) is 0.76mm-2.6 mm.

9. A rocket engine comprising a combustion chamber, characterized in that said combustion chamber is a combustion chamber according to any one of claims 1 to 8.

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