Cooling structure for combustion chamber of air turbine rocket engine and application

Abstract

The invention discloses a cooling structure for a combustion chamber of an air turbine rocket engine and application, the cooling structure comprises a first barrel arranged in the combustion chamber and a second barrel arranged in the first barrel, a first annular cavity is formed between the inner wall of the combustion chamber and the outer wall of the first barrel, and a second annular cavity is formed between the inner wall of the combustion chamber and the outer wall of the second barrel. A second annular cavity is formed between the inner wall of the first barrel and the outer wall of the second barrel, a plurality of gas impact holes are formed in the first barrel, and the second barrel is made of a porous medium with a microporous structure; and the microporous structures of the first annular cavity, the gas impact holes, the second annular cavity and the second barrel jointly form a cooling gas channel. The invention aims to solve the problems that the cooling effect of a combustion chamber of an air turbine rocket engine is poor, and cooling gas cannot be efficiently utilized.

Description

Technical Field

[0001] The invention belongs to the technical field of combined engines, and in particular relates to a cooling structure for a combustion chamber of an air turbine rocket engine and its application. Background Art

[0002] As a type of combined engine, the air turbine rocket engine has great development potential. The air turbine rocket engine uses a gas generator to generate fuel-rich gas to drive the turbine to work and drive the compressor to work. The air captured by the inlet is compressed by the compressor and mixed with the gas to enter the combustion chamber for combustion. The generated high-temperature gas expands through the nozzle to generate thrust. However, the temperature of the combustion chamber of the air turbine rocket engine is as high as 2500K or more. When there is no thermal protection, the near-wall temperature will exceed 2000K, which exceeds the allowable temperature of most high-temperature materials. Therefore, the research and development of the cooling structure of the combustion chamber of the air turbine rocket engine is of great significance to the normal operation of the air turbine rocket engine.

[0003] The existing cooling structure of air turbine rocket engines is mostly discrete hole film cooling. Specifically, discrete hole film cooling is to arrange a flame tube inside the combustion chamber, and there are several film holes along the radial direction of the flame tube on the flame tube. A part of the air compressed by the compressor enters the cavity formed by the inner wall of the combustion chamber and the outer wall of the flame tube as cooling gas. This part of the cooling gas will be ejected from each film hole to the combustion area in the middle of the flame tube. After that, this part of the cooling gas will adhere to the inner surface of the flame tube with the flow of the intermediate high-temperature combustion gas, forming an air film that blocks the heat from being transferred to the combustion chamber wall. However, this air film cooling method is subject to the distribution and density of the air film holes. In fact, the air film holes of this cooling method cannot be dense enough, and the thickness of the air film formed is relatively uneven, resulting in poor cooling effect and the cooling gas cannot be efficiently used. Summary of the invention

[0004] In view of the problems existing in the prior art, the present invention provides a cooling structure and application for an air turbine rocket engine combustion chamber, which aims to solve the problem that the cooling effect of the air turbine rocket engine combustion chamber is poor and the cooling gas cannot be efficiently utilized.

[0005] In order to solve the above technical problems, the present invention is implemented by the following technical solutions: According to a first aspect of the present invention, there is provided a cooling structure for a combustion chamber of an air turbine rocket engine, comprising a first cylinder arranged in the combustion chamber, and a second cylinder arranged in the first cylinder, a first annular cavity is formed between an inner wall of the combustion chamber and an outer wall of the first cylinder, a second annular cavity is formed between an inner wall of the first cylinder and an outer wall of the second cylinder, a plurality of gas impact holes are formed in the first cylinder, and the second cylinder is made of a porous medium with a microporous structure; the first annular cavity, the gas impact holes, the second annular cavity and the microporous structure of the second cylinder together form a cooling gas channel.

[0006] In a possible implementation manner of the first aspect, a diameter of the gas impact hole is 1 mm to 10 mm.

[0007] In a possible implementation manner of the first aspect, the distance between the inner wall of the combustion chamber and the outer wall of the first cylinder is 5 to 10 times the diameter of the gas impact hole.

[0008] In a possible implementation manner of the first aspect, a distance between an inner wall of the first cylinder and an outer wall of the second cylinder is 5 to 10 times the diameter of the gas impact hole.

[0009] In a possible implementation manner of the first aspect, the thickness of the first cylinder and the thickness of the second cylinder are both 1.5 to 3 times the diameter of the gas impact hole.

[0010] In a possible implementation manner of the first aspect, the gas impact hole is a cylindrical hole opened along the radial direction of the first cylinder.

[0011] In a possible implementation manner of the first aspect, along the flow direction of the cooling gas, the density of the plurality of gas impact holes opened on the first cylinder gradually increases.

[0012] In a possible implementation manner of the first aspect, central axes of the combustion chamber, the first cylinder, and the second cylinder coincide with each other.

[0013] In a possible implementation manner of the first aspect, the porous medium having a microporous structure is a silicon carbide composite material.

[0014] According to a second aspect of the present invention, there is provided an application of a cooling structure for a combustion chamber of an air turbine rocket engine, for cooling the combustion chamber of the air turbine rocket engine.

[0015] Compared with the prior art, the present invention has at least the following beneficial effects: The present invention provides a cooling structure for an air turbine rocket engine combustion chamber, which forms a more efficient and complex cooling gas channel through the gas impact holes on the first cylinder and the second cylinder made of a porous medium with a microporous structure. The cooling gas not only exchanges heat with the first cylinder in the first annular cavity, but also forms a strong impact airflow through the gas impact holes, and performs efficient convection heat exchange with the second cylinder in the second annular cavity. At the same time, the cooling gas can penetrate into the microscopic pores of the second cylinder and continuously exchange heat with the pore walls, thereby greatly reducing the temperature of the second cylinder, and the cooling gas passes through the inner wall of the second cylinder through the microscopic pores, forming a low-temperature gas film on the inner wall of the second cylinder to reduce the transfer of heat to the wall. This multi-level collaborative cooling cooling mechanism significantly improves the cooling efficiency and effectively reduces the temperature near the wall of the combustion chamber. In other words, the cooling structure of the present invention realizes uniform distribution and efficient utilization of cooling gas in the combustion chamber through continuous gas impact holes and porous media with a microporous structure. The cooling gas can fully carry out convective heat exchange with the walls of each cylinder, the walls of the gas impact holes and the walls of the microscopic pores, thereby reducing the waste of cooling gas, improving the cooling efficiency and optimizing the utilization of the cooling gas.

[0016] In order to make the above-mentioned objects, features and advantages of the present invention more obvious and easy to understand, preferred embodiments are given below and described in detail with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] In order to more clearly illustrate the technical solutions in the specific implementation modes of the present invention, the drawings required for use in the description of the specific implementation modes will be briefly introduced below. Obviously, the drawings described below are some implementation modes of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0018] Figure 1 This is an axonometric diagram of a cooling structure for an air turbine rocket engine combustion chamber according to an embodiment of the present invention; Figure 2 It is a front cross-sectional view of a cooling structure for a combustion chamber of an air turbine rocket engine according to an embodiment of the present invention; Figure 3 This is a schematic diagram of cooling gas flow in a cooling structure for a combustion chamber of an air turbine rocket engine according to an embodiment of the present invention; Figure 4 This is a schematic diagram of convective heat transfer and air film formation of various parts in a cooling structure for a combustion chamber of an air turbine rocket engine according to an embodiment of the present invention; Figure 5 A side view of a cooling structure for a combustion chamber of an air turbine rocket engine according to an embodiment of the present invention; Figure 6 This is a partial isometric view of a cooling structure for an air turbine rocket engine combustion chamber according to an embodiment of the present invention; Figure 7 This is a microscopic schematic diagram of a braided structure of a silicon carbide composite material used in a cooling structure of an air turbine rocket engine combustion chamber according to an embodiment of the present invention.

[0019] In the figure: 1-combustion chamber; 2-first cylinder; 20-gas impact hole; 3-second cylinder; 4-first annular cavity; 5-second annular cavity. DETAILED DESCRIPTION

[0020] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution of the present invention will be clearly and completely described below in conjunction with the accompanying drawings. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0021] Combination Figures 1 to 7 As shown, an embodiment of the present invention provides a cooling structure for a combustion chamber of an air turbine rocket engine, comprising a first cylinder 2 disposed in a combustion chamber 1, and a second cylinder 3 disposed in the first cylinder 2, a first annular cavity 4 is formed between the inner wall of the combustion chamber 1 and the outer wall of the first cylinder 2, a second annular cavity 5 is formed between the inner wall of the first cylinder 2 and the outer wall of the second cylinder 3, a plurality of gas impact holes 20 are provided on the first cylinder 2, and the second cylinder 3 is made of a porous medium with a microporous structure. The first annular cavity 4, the gas impact holes 20, the second annular cavity 5 and the microporous structure of the second cylinder 3 together form a cooling gas channel.

[0022] Specifically, the cooling structure of the present invention includes a first barrel 2 arranged inside the combustion chamber 1. The material of the first barrel 2 has good heat resistance and mechanical strength to withstand the high temperature and high pressure environment in the combustion chamber. A first annular cavity 4 is formed between its outer wall and the inner wall of the combustion chamber 1. This cavity is used to introduce and distribute cooling gas. Inside the first barrel 2, a second barrel 3 is provided. The second barrel 3 is made of a porous medium material with a microporous structure. The selection of this porous medium with a microporous structure is to achieve uniform distribution and efficient heat exchange of cooling gas on a microscopic scale. A second annular cavity 5 is formed between the outer wall of the second barrel 3 and the inner wall of the first barrel 2, which serves as a space for further distribution and impact heat exchange of cooling gas. On the wall of the first barrel 2, there are a number of gas impact holes 20. These gas impact holes 20 are for introducing the cooling gas in the first annular cavity 4 into the second annular cavity 5 to form a strong impact jet and enhance the convective heat exchange effect.

[0023] It should be noted that the number, size and distribution of the gas impact holes 20 are optimized and designed according to specific cooling requirements and combustion chamber dimensions. For example, the gas impact holes 20 may be cylindrical or irregular in shape.

[0024] Combination Figure 4 As shown, the working principle of the cooling structure for the combustion chamber of the air turbine rocket engine is specifically as follows: a part of the air output from the compressor is mixed with the fuel gas and enters the main combustion area of ​​the combustion chamber, while the other part is guided to the first annular cavity 4 as cooling gas. In the first annular cavity 4, the cooling gas flows along the flow direction and performs preliminary heat exchange with the outer wall of the first cylinder 2. Through the gas impact hole 20 on the first cylinder 2, part of the cooling gas is ejected into the second annular cavity 5 to form a strong impact jet. This impact jet not only enhances the convective heat exchange with the outer wall of the second cylinder 3, but also promotes the mixing and uniform distribution of the cooling gas in the second annular cavity 5. The cooling gas entering the second annular cavity 5 further penetrates into the microporous structure of the second cylinder 3. In the microscopic pores, the cooling gas continuously exchanges heat with the pore wall, effectively reducing the temperature of the second cylinder 3. At the same time, part of the cooling gas penetrates the inner wall of the second cylinder 3 through the microporous structure to form a layer of low-temperature gas film, which effectively blocks the heat transfer from the high-temperature fuel gas in the combustion area to the combustion chamber wall. The efficient cooling structure can ensure the stable operation of the air turbine rocket engine in a high temperature and high pressure environment for a long time, avoiding performance degradation or failure due to overheating.

[0025] Through the above process, the cooling gas not only performs effective convection heat exchange in the first annular cavity 4 and the second annular cavity 5, but also realizes microscopic heat exchange in the microporous structure of the second cylinder 3. In addition, the formation of a low-temperature gas film further enhances the cooling effect and reduces the transfer of heat to the combustion chamber wall. This comprehensive cooling mechanism significantly improves the cooling efficiency of the air turbine rocket engine combustion chamber and optimizes the use of cooling gas, effectively solving the problem of poor cooling effect and inefficient use of cooling gas in the prior art.

[0026] It should be understood that the combustion chamber 1 includes a columnar section and a convergent section, so the first cylinder 2 and the second cylinder 3 also adaptively have a columnar section and a convergent section. In addition, the first cylinder 2 can be fixed by connecting the two ends to other main structures of the air turbine rocket engine, or by connecting the first cylinder 2 to the combustion chamber 1 through connecting ribs. Similarly, the second cylinder 3 can be fixed by connecting the two ends to other main structures of the air turbine rocket engine, or by connecting the first cylinder 2 to the second cylinder 3 through connecting ribs.

[0027] In one achievable manner, the aperture of the gas impact hole 20 is 1 mm to 10 mm. Specifically, an aperture less than 1 mm causes the flow rate of the cooling gas to be too high when passing through, forming turbulence, increasing the difficulty of flow control, and also causing erosion to the hole wall due to excessive local pressure. On the contrary, if the aperture is too large (more than 10 mm), the amount of cooling gas released from a single hole is too much, making it difficult to adjust the cooling effect, and even interfering with the mainstream airflow in the combustion chamber, affecting combustion stability. The aperture range of 1 mm to 10 mm not only meets the processing technology, such as laser drilling, electric spark drilling, etc., but can ensure the smoothness of the hole wall, reduce processing costs and time, and at the same time ensure the strength and durability of the structure.

[0028] Preferably, in order to maximize cooling efficiency, the holes are evenly distributed to ensure that all areas of the inner wall of the combustion chamber can be effectively cooled.

[0029] In one achievable manner, the spacing between the inner wall of the combustion chamber 1 and the outer wall of the first barrel 2 is 5 to 10 times the aperture of the gas impact hole 20. Specifically, controlling the spacing to 5 to 10 times the aperture of the gas impact hole 20 not only ensures the compactness of the cooling structure, but also avoids the reduction of the effective volume of the combustion chamber due to excessive spacing. The cooling gas enters the first annular cavity 4 through the gas impact hole 20, and forms a certain flow rate and pressure distribution therein to achieve impact cooling of the inner wall of the combustion chamber. The size of the spacing directly affects the flow state of the cooling gas in the first annular cavity 4. Too small a spacing causes the cooling gas flow to be blocked, the flow rate to increase, the flow resistance to increase, and even turbulence to affect the cooling effect. Too large a spacing causes the cooling gas to be unevenly distributed in the first annular cavity 4, and some areas are insufficiently cooled. Therefore, controlling the spacing to 5 to 10 times the aperture of the gas impact hole 20 can not only ensure that the cooling gas has enough flow space to form a stable flow state, but also ensure that the cooling gas can evenly and effectively impact the inner wall of the combustion chamber.

[0030] In one achievable manner, the spacing between the inner wall of the first barrel 2 and the outer wall of the second barrel 3 is 5 to 10 times the aperture of the gas impact hole 20. Specifically, the spacing between the inner wall of the first barrel 2 and the outer wall of the second barrel 3 is controlled to be 5 to 10 times the aperture of the gas impact hole 20, which not only ensures the compactness of the combustion chamber, but also avoids the reduction of the effective volume of the combustion chamber caused by the excessive spacing. After the cooling gas enters the first annular cavity 4 through the gas impact hole 20, part of the gas continues to flow to the second annular cavity 5 to further realize the cooling of the outer peripheral structure of the combustion chamber. Therefore, the spacing design between the first barrel 2 and the second barrel 3 directly affects the flow state of the cooling gas in the second annular cavity 5. Too small a spacing causes the cooling gas flow to be blocked, the flow velocity to increase, the flow resistance to increase, and even turbulence to affect the cooling effect. Too large a spacing causes the cooling gas to be unevenly distributed in the second annular cavity 5, and some areas are insufficiently cooled. Controlling the spacing to 5 to 10 times the diameter of the gas impact hole 20 can not only ensure that the cooling gas has enough flow space to form a stable flow state, but also ensure that the cooling gas can evenly and effectively cover the wall surface in the second annular cavity 5.

[0031] In one achievable manner, the thickness of the first cylinder 2 and the thickness of the second cylinder 3 are both 1.5 to 3 times the diameter of the gas impact hole 20. Specifically, controlling the cylinder thickness within the range of 1.5 to 3 times the diameter of the gas impact hole 20 can ensure that the cylinder has sufficient strength and rigidity, while avoiding weight increase and space waste caused by excessive thickening.

[0032] In one possible implementation, the gas impact hole 20 is a cylindrical hole opened radially along the first cylinder 2. Specifically, this design allows the cooling gas to be vertically sprayed onto the outer wall of the second cylinder 3 after passing through the gas impact hole 20, and then dispersed into the second annular cavity 5, achieving uniform distribution of the cooling gas in the second annular cavity 5, thereby improving the cooling effect. When the cooling gas is vertically sprayed onto the outer wall of the second cylinder 3, an effective impact cooling effect can be formed, quickly taking away the heat on the inner wall of the combustion chamber.

[0033] In one feasible method, along the flow direction of the cooling gas, the density of the plurality of gas impact holes 20 opened on the first cylinder 2 gradually increases. According to the heat load distribution in the combustion chamber, the density of the gas impact holes is reasonably adjusted to ensure that more cooling gas is allocated to the high heat load area to achieve uniform dispersion and effective heat extraction.

[0034] Specifically, during the combustion process of an air turbine rocket engine, the heat load distribution in the combustion chamber is usually uneven. Therefore, by adjusting the density of the gas impact holes according to the heat load distribution, the cooling gas can be used more effectively to achieve focused cooling of the high heat load area. Gradually increasing the density of the gas impact holes along the flow direction of the cooling gas can ensure that more cooling gas passes through the high heat load area, thereby taking away the heat more effectively, which can not only improve the cooling efficiency, but also reduce the waste of cooling gas, making the entire cooling structure more efficient. By adjusting the density of the gas impact holes, the heat can also be evenly dispersed. Arranging more dense holes in the high heat load area can make the cooling gas cover the area more evenly and avoid the occurrence of local overheating. At the same time, the holes in the low heat load area are relatively sparsely arranged to save cooling gas and reduce unnecessary cooling burden.

[0035] In one practicable manner, the central axes of the combustion chamber 1, the first barrel 2, and the second barrel 3 coincide, ensuring the coaxiality between the components of the cooling structure, which is conducive to the uniform distribution and effective flow of the cooling gas, thereby improving the cooling efficiency. Specifically, the cooling gas can flow more evenly along the central axis direction, avoiding the occurrence of biased flow or eddy flow of the cooling gas during the flow process, thereby improving the utilization efficiency of the cooling gas.

[0036] In one possible implementation, Figure 7 As shown, the porous medium with a microporous structure is a silicon carbide composite material.

[0037] Exemplarily, when preparing a silicon carbide composite material, first, the silicon carbide fibers are woven into a preform according to a predetermined weaving method. The shape and size of the preform are designed according to the shape of the combustion chamber and the cooling requirements. Finally, the silicon carbide matrix material is filled into the pores of the preform. In the preparation process of the silicon carbide composite material, the porosity of the porous medium can be controlled by adjusting parameters such as the amount of matrix material introduced. The filled preform is sintered at a high temperature to form a good bond between the silicon carbide fibers and the matrix material.

[0038] In one feasible embodiment, the first cylinder 2 is made of a high-temperature alloy, which has excellent high-temperature performance, oxidation resistance and corrosion resistance, and can maintain high strength and good mechanical properties in a high-temperature environment. The first cylinder 2 is made of a high-temperature alloy to cope with the extreme high-temperature environment in the combustion chamber.

[0039] In one embodiment, the present invention provides an application of a cooling structure for a combustion chamber of an air turbine rocket engine, which is used for cooling the combustion chamber of an air turbine rocket engine. That is, the cooling structure is applied to cooling the combustion chamber of an air turbine rocket engine.

[0040] In the description of the present invention, it is necessary to understand that the terms "upper", "lower", "bottom", "inside", "outside", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present invention and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on the present invention.

[0041] In addition, the terms "first" and "second" are used for descriptive purposes only and should not be understood as indicating or implying relative importance or implicitly indicating the number of the indicated technical features. Therefore, the features defined as "first" and "second" may explicitly or implicitly include at least one of the features. In the description of the present invention, the meaning of "plurality" is at least two, such as two, three, etc., unless otherwise clearly and specifically defined.

[0042] In the present invention, unless otherwise clearly specified and limited, the terms "connected", "connection" and the like should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection, an electrical connection, or communication with each other; it can be a direct connection, or an indirect connection through an intermediate medium, it can be the internal connection of two elements or the interaction relationship between two elements, unless otherwise clearly defined. For ordinary technicians in this field, the specific meanings of the above terms in the present invention can be understood according to specific circumstances.

[0043] In the present invention, unless otherwise clearly specified and limited, a first feature being "above" or "below" a second feature may mean that the first and second features are in direct contact, or the first and second features are in indirect contact through an intermediate medium. Moreover, a first feature being "above", "above" or "above" a second feature may mean that the first feature is directly above or obliquely above the second feature, or simply means that the first feature is higher in level than the second feature. A first feature being "below", "below" or "below" a second feature may mean that the first feature is directly below or obliquely below the second feature, or simply means that the first feature is lower in level than the second feature.

[0044] In the present invention, the terms "one embodiment", "some embodiments", "examples", "specific examples", or "some examples" etc. mean that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present invention. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the described specific features, structures, materials or characteristics may be combined in any one or more embodiments or examples in a suitable manner. In addition, those skilled in the art may combine and combine the different embodiments or examples described in this specification and the features of the different embodiments or examples, without contradiction.

[0045] Finally, it should be noted that the above-described embodiments are only specific implementations of the present invention, which are used to illustrate the technical solutions of the present invention, rather than to limit them. The protection scope of the present invention is not limited thereto. Although the present invention is described in detail with reference to the above-described embodiments, ordinary technicians in the field should understand that any technician familiar with the technical field can still modify the technical solutions recorded in the above-described embodiments within the technical scope disclosed by the present invention, or can easily think of changes, or make equivalent replacements for some of the technical features therein; and these modifications, changes or replacements do not make the essence of the corresponding technical solutions deviate from the spirit and scope of the technical solutions of the embodiments of the present invention, and should be included in the protection scope of the present invention. Therefore, the protection scope of the present invention shall be based on the protection scope of the claims.

Claims

1. A cooling structure for an air turbine rocket engine combustion chamber, characterized in that: The invention comprises a first cylinder (2) arranged in a combustion chamber (1), and a second cylinder (3) arranged in the first cylinder (2); a first annular cavity (4) is formed between the inner wall of the combustion chamber (1) and the outer wall of the first cylinder (2); a second annular cavity (5) is formed between the inner wall of the first cylinder (2) and the outer wall of the second cylinder (3); a plurality of gas impact holes (20) are formed on the first cylinder (2); and the second cylinder (3) is made of a porous medium having a microporous structure; the first annular cavity (4), the gas impact holes (20), the second annular cavity (5) and the microporous structure of the second cylinder (3) together form a cooling gas channel.

2. A cooling structure for an air turbine rocket engine combustion chamber according to claim 1, characterized in that: The diameter of the gas impact hole (20) is 1 mm to 10 mm.

3. A cooling structure for an air turbine rocket engine combustion chamber according to claim 2, characterized in that: The distance between the inner wall of the combustion chamber (1) and the outer wall of the first cylinder (2) is 5 to 10 times the diameter of the gas impact hole (20).

4. A cooling structure for an air turbine rocket engine combustion chamber according to claim 2, characterized in that: The distance between the inner wall of the first cylinder (2) and the outer wall of the second cylinder (3) is 5 to 10 times the diameter of the gas impact hole (20).

5. A cooling structure for an air turbine rocket engine combustion chamber according to claim 2, characterized in that: The thickness of the first cylinder (2) and the thickness of the second cylinder (3) are both 1.5 to 3 times the diameter of the gas impact hole (20).

6. A cooling structure for an air turbine rocket engine combustion chamber according to claim 1, characterized in that: The gas impact hole (20) is a cylindrical hole opened along the radial direction of the first cylinder (2).

7. A cooling structure for an air turbine rocket engine combustion chamber according to claim 1, characterized in that: Along the flow direction of the cooling gas, the density of the plurality of gas impact holes (20) opened on the first cylinder (2) gradually increases.

8. The cooling structure for the combustion chamber of an air turbine rocket engine according to claim 1, characterized in that: The central axes of the combustion chamber (1), the first cylinder (2) and the second cylinder (3) coincide with each other.

9. A cooling structure for an air turbine rocket engine combustion chamber according to claim 1, characterized in that: The porous medium with a microporous structure is a silicon carbide composite material.

10. Application of a cooling structure for a combustion chamber of an air turbine rocket engine according to any one of claims 1 to 9, characterized in that: Used for cooling the combustion chamber of air turbine rocket engines.

Images