An axially staged and circumferentially partitioned combustion chamber for an aircraft engine

Abstract

An axially graded and circumferentially partitioned combustion chamber of an aircraft engine comprises: an outer casing of a combustion chamber; an inner casing of a combustion chamber, which is arranged inside the outer casing of the combustion chamber; a flame tube, which is arranged between the outer casing of the combustion chamber and the inner casing of the combustion chamber, and has a head with a plurality of head mounting holes and a head air inlet hole distributed alternately along the circumferential direction; a plurality of air centrifugal atomizing fuel nozzles, which are arranged in each head mounting hole; a plurality of axially graded and circumferentially partitioned rectifying baffles, which are circumferentially arranged inside the flame tube, are located at the front end of the flame tube, face each blade of a turbine guide device, have circumferentially partitioned jet holes at the tail end, and have front and rear concave cavities distributed on both side walls, which are hollow inside; a plurality of front concave cavity fuel jet nozzles, which are connected to the side walls of each axially graded and circumferentially partitioned rectifying baffle, and are located at the front side of the front concave cavity; a plurality of rear concave cavity fuel jet nozzles, which are connected to the side walls of each axially graded and circumferentially partitioned rectifying baffle, and are located at the front side of the rear concave cavity.

Description

Technical Field

[0001] The present application belongs to the technical field of aero-engine combustion chamber design, and specifically relates to an aero-engine axially staged and circumferentially partitioned combustion chamber. Background Art

[0002] As the demand for aircraft engine thrust increases, the required turbine inlet temperature increases accordingly, that is, the combustion chamber outlet temperature needs to increase, and the combustion chamber temperature rise needs to increase.

[0003] In the current combustion chamber, high-pressure air from the compressor is used to carry out mixed combustion through the flame tube head intake, and the oil-gas ratio in the combustion area is controlled through the mixing holes on the flame tube side wall. The uniformity of the combustion chamber outlet temperature field is adjusted, and the flame tube side wall is cooled. The technical optimization and improvement are difficult. In the face of the demand for increasing the temperature rise of the combustion chamber, it is mainly achieved by increasing the flame tube head intake, which is used to reduce the uniformity of the combustion chamber outlet temperature field and the high-pressure air for cooling the flame tube side wall, which is easy to generate local high temperature, making the flame tube side wall susceptible to high-temperature ablation, and the outlet temperature field is uneven, the quality is reduced, and the turbine guide vane is easily subjected to high-temperature ablation, which increases the difficulty of cooling the turbine guide vane.

[0004] This application is proposed in view of the above-mentioned technical defects.

[0005] It should be noted that the disclosure of the above background technology content is only used to assist in understanding the inventive concept and technical solution of the present invention, and it does not necessarily belong to the prior art of this patent application. In the absence of clear evidence that the above content has been disclosed on the filing date of this application, the above background technology should not be used to evaluate the novelty and creativity of this application. Summary of the Invention

[0006] The purpose of the present application is to provide an axially staged and circumferentially partitioned combustion chamber for an aircraft engine to overcome or alleviate at least one of the existing technical deficiencies.

[0007] The technical solution of this application is:

[0008] An axially staged and circumferentially partitioned combustion chamber for an aero-engine, comprising:

[0009] combustion chamber casing;

[0010] The inner casing of the combustion chamber is arranged inside the outer casing of the combustion chamber;

[0011] The flame tube is arranged between the outer casing of the combustion chamber and the inner casing of the combustion chamber, and the head portion has a plurality of head mounting holes and head air inlet holes distributed alternately along the circumferential direction;

[0012] A plurality of air centrifugal atomizing fuel nozzles are arranged in the mounting holes of each head;

[0013] Multiple axially graded circumferentially partitioned rectifying baffles are circumferentially arranged inside the flame tube, at the front end of the flame tube, facing the blades of the turbine guide vane. The tail end has circumferentially partitioned jet holes, and the two side walls are provided with front and rear concave cavities, which are hollow inside and connected to the various head air inlet holes.

[0014] A plurality of front concave cavity fuel jet nozzles are connected to the side walls of each axially graded and circumferentially partitioned rectifying baffle and are located at the front side of the front concave cavity;

[0015] A plurality of rear concave fuel injection nozzles are connected to the side walls of each axially graded and circumferentially partitioned rectifying baffle and are located at the front side of the rear concave cavity;

[0016] in,

[0017] In the flame tube, an axial primary combustion zone is formed in front of the front concave cavity;

[0018] In the flame tube, an axial secondary combustion zone is formed between the front cavity and the rear cavity;

[0019] Inside the flame tube, an axial three-stage combustion zone is formed between the front concave cavity and the rear concave cavity;

[0020] Inside the flame tube, an axial four-stage combustion zone is formed behind the rear concave cavity.

[0021] According to at least one embodiment of the present application, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, there are three head mounting holes and their air centrifugal atomizing fuel nozzles located between two adjacent axially staged circumferentially partitioned rectifying baffles, which are distributed radially.

[0022] According to at least one embodiment of the present application, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, a through-connecting flame channel is provided on the side wall of each axially staged circumferentially partitioned rectifying baffle, connecting each axial primary combustion zone;

[0023] There are two head air inlet holes corresponding to each axially graded circumferentially partitioned rectifying baffle, located on both sides of the cross-flame channel;

[0024] The aero-engine axially staged circumferentially partitioned combustion chamber further comprises:

[0025] Multiple baffle inner support plates are laterally supported inside the front end of each axially graded circumferentially partitioned rectifying baffle and are located between the corresponding two head air inlet holes.

[0026] According to at least one embodiment of the present application, in the aforementioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, each axially staged circumferentially partitioned rectifying baffle has a plurality of circumferentially partitioned jet holes at its tail end, which are distributed radially;

[0027] or,

[0028] The circumferentially partitioned jet holes at the tail of each axially graded circumferentially partitioned rectifying baffle are strip-shaped and extend in the radial direction.

[0029] According to at least one embodiment of the present application, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, there are multiple front concave cavity fuel jet nozzles and rear concave cavity fuel jet nozzles corresponding to each axially staged circumferentially partitioned rectifying baffle, which are distributed in multiple points along the radial direction.

[0030] According to at least one embodiment of the present application, in the aforementioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, a plurality of partition film cooling holes are provided on the sidewall of each axially staged circumferentially partitioned rectifying partition, and the holes are distributed within the axial secondary combustion zone and the axial tertiary combustion zone;

[0031] The baffle film cooling holes located in the front concave cavity and the rear concave cavity on the side walls of each axially graded and circumferentially partitioned rectifying baffle are inclined along the return flow direction in the front concave cavity and the rear concave cavity.

[0032] According to at least one embodiment of the present application, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, there are multiple flame tube air film cooling holes on the side wall of the flame tube, which are distributed within the axial secondary combustion zone, axial tertiary combustion zone, and axial fourth combustion zone.

[0033] According to at least one embodiment of the present application, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, the turbine guide vane is formed at the tail of the flame tube, and the outer wall has a plurality of turbine guide vane film cooling holes distributed between the turbine guide vane blades.

[0034] According to at least one embodiment of the present application, in the aforementioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, the flame tube and its turbine guide vane are made of SiCf / SiC ceramic-based composite materials.

[0035] According to at least one embodiment of the present application, the above-mentioned axially staged and circumferentially partitioned combustion chamber of the aircraft engine further includes:

[0036] The diffuser is installed at the inlet between the outer casing of the combustion chamber and the inner casing of the combustion chamber. BRIEF DESCRIPTION OF THE DRAWINGS

[0037] Figure 1 Schematic diagram of an axially staged and circumferentially partitioned combustion chamber of an aero-engine provided in an embodiment of the present application;

[0038] Figure 2 This is a partial schematic diagram of an axially staged and circumferentially partitioned combustion chamber of an aero-engine provided in an embodiment of the present application;

[0039] Figure 3Schematic diagram of axially staged and circumferentially zoned combustion in an axially staged and circumferentially zoned combustion chamber of an aero-engine provided by an embodiment of the present application;

[0040] Figure 4 Schematic diagram of an axially staged and circumferentially partitioned combustion chamber of an aircraft engine provided by an embodiment of the present application, wherein the turbine guide vane is cooled by a jet of air behind an axially staged and circumferentially partitioned straightening baffle;

[0041] Figure 5-Figure 6 This is a schematic diagram of an axially staged and circumferentially partitioned combustion chamber of an aircraft engine provided by an embodiment of the present application, wherein the axially staged and circumferentially partitioned combustion chamber utilizes the air intake at the head of the flame tube to cool the axially staged and circumferentially partitioned rectifying baffles;

[0042] in:

[0043] 1-External combustion chamber casing; 2-Inner combustion chamber casing; 3-Diffuser; 4-Flame tube; 5-Air centrifugal atomizing fuel nozzle; 6-Axially graded and circumferentially partitioned rectifying baffle; 7-Turbine guide vane; 8-Front concave fuel jet nozzle; 9-Rear concave fuel jet nozzle; 10-Baffle inner support plate;

[0044] A-head mounting hole; B-head air inlet hole; C-circumferential partition jet hole; D-front concave cavity; E-rear concave cavity; F-cross-flame channel;

[0045] I-axial primary combustion zone; II-axial secondary combustion zone; III-axial tertiary combustion zone; IV-axial fourth combustion zone.

[0046] In order to better illustrate this embodiment, some parts of the drawings may be omitted, enlarged or reduced, and do not represent the size of the actual product. In addition, the drawings are only used for illustrative purposes and should not be understood as limitations on this patent. DETAILED DESCRIPTION

[0047] To make the technical solution and its advantages of the present application clearer, the technical solution of the present application will be described in further detail below in conjunction with the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of the present application and are only used to explain the present application, not to limit the present application. It should be noted that, for ease of description, only the parts related to the present application are shown in the accompanying drawings, and other related parts can refer to the general design. In the absence of conflict, the embodiments of the present application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0048] In addition, unless otherwise defined, the technical or scientific terms used in the description of this application should have the ordinary meanings understood by those of ordinary skill in the art to which this application belongs. The words "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," and "outer" used in the description of this application are only used to indicate relative directions or positional relationships, and do not imply that the device or component must have a specific orientation, be constructed, or operate in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. Therefore, they should not be understood as limitations on this application. The words "first," "second," "third," and similar terms used in the description of this application are used only for descriptive purposes to distinguish different components and should not be understood to indicate or imply relative importance. The words "one," "an," or "the" used in the description of this application should not be understood as absolute limitations on quantity, but should be understood as meaning the presence of at least one. The words "include" or "comprises" used in the description of this application mean that the element or object listed before the word includes the elements or objects listed after the word and their equivalents, but does not exclude other elements or objects.

[0049] In addition, it should be noted that, unless otherwise clearly stipulated and limited, the words "install", "connect", "connect" and similar terms used in the description of this application should be understood in a broad sense. For example, the connection can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, an indirect connection through an intermediate medium, or a connection between two components. Technical personnel in the field can understand their specific meanings in this application according to the specific circumstances.

[0050] The following is combined with Figures 1 to 6 This application is described in further detail.

[0051] An axially staged and circumferentially partitioned combustion chamber for an aero-engine, comprising:

[0052] Combustion chamber outer casing 1;

[0053] The inner combustion chamber casing 2 is arranged inside the outer combustion chamber casing 1;

[0054] The diffuser 3 is provided at the inlet portion between the combustion chamber casing 1 and the combustion chamber casing 2;

[0055] The flame tube 4 is arranged between the outer casing 1 of the combustion chamber and the inner casing 2 of the combustion chamber, and the head has a plurality of head mounting holes A and head air inlet holes B distributed alternately along the circumferential direction;

[0056] A plurality of air centrifugal atomizing fuel nozzles 5 are arranged in the mounting holes A of each head;

[0057] Multiple axially graded circumferentially partitioned rectifying baffles 6 are circumferentially arranged inside the flame tube 4, at the front end of the flame tube 4, facing the blades of the turbine guide vane 7, with circumferentially partitioned jet holes C at the tail end, and front and rear concave cavities D and E distributed on both side walls. The interior of the cavities is hollow and connected to the various head air inlet holes B;

[0058] A plurality of front cavity fuel jet nozzles 8 are connected to the side walls of the axially graded and circumferentially partitioned rectifying baffles 6 and are located at the front side of the front cavity D;

[0059] A plurality of rear cavity fuel jet nozzles 9 are connected to the side walls of the axially graded and circumferentially partitioned rectifying baffles 6 and are located at the front side of the rear cavity E;

[0060] in,

[0061] In the flame tube 4, an axial primary combustion zone I is formed in front of the front cavity D;

[0062] In the flame tube 4, an axial secondary combustion zone II is formed between the front cavity D and the rear cavity E;

[0063] Inside the flame tube 4, an axial three-stage combustion zone III is formed between the front cavity D and the rear cavity E;

[0064] Inside the flame tube 4, an axial four-stage combustion zone IV is formed behind the rear concave cavity E.

[0065] The above-mentioned embodiment discloses an axially staged circumferentially partitioned combustion chamber of an aircraft engine. When the combustion chamber is in operation, high-pressure air from the compressor can flow into the inlet of the outer wall 1 and the inner wall 2 of the combustion chamber through the diffuser 3, wherein part of the air flows into between the outer wall 1 of the combustion chamber and the outer wall of the flame tube 4, and part of the air flows into between the inner wall 2 and the inner wall of the flame tube 4. The two parts of the air flow flow axially backward, and can cool the inner and outer walls of the flame tube 3 along the way, and are then discharged into the turbine. The remaining part of the high-pressure air will enter the combustion zone between the axially staged circumferentially partitioned rectifying baffles 6 through each air centrifugal atomizing fuel nozzle 5, flow axially backward, and enter the axially staged circumferentially partitioned rectifying baffle 6 through each head air inlet hole B, and then be ejected backward through the circumferentially partitioned jet hole C, and can cool the axially staged circumferentially partitioned rectifying baffle 6 along the way.

[0066] The axially graded and circumferentially partitioned combustion chamber of the aircraft engine disclosed in the above embodiment can, when working, spray fuel into the combustion zone through the air centrifugal atomizing fuel nozzle 5. This part of the fuel will enter the axial primary combustion zone I, and can be fully mixed with the high-pressure air entering the combustion zone, and is in the reflow zone formed by the air centrifugal atomizing fuel nozzle 5, which is easy to achieve ignition and flame stabilization. In addition, fuel can be sprayed into the combustion zone through the front concave cavity fuel jet nozzle 8. This part of the fuel will vertically enter the axial secondary combustion zone II, and can be fully mixed with the airflow flowing backward in the combustion zone, which can be used. The recirculation zone formed in the front cavity D is used for low-speed stable combustion, and fuel can be sprayed into the combustion zone through the rear cavity fuel jet nozzle 9. This part of the fuel will vertically enter the axial three-stage combustion zone III and can be fully mixed with the airflow flowing backward in the combustion zone. The recirculation zone formed in the rear cavity E can be used for low-speed stable combustion. After the airflow flowing backward in the combustion zone flows out between the axial graded circumferentially partitioned rectifying baffles 6, it will enter the axial four-stage combustion zone IV for organized combustion, and then be discharged to the turbine guide 7. After being rectified by the turbine guide 7, it drives the turbine to do work.

[0067] As for the axially staged and circumferentially partitioned combustion chamber of the aircraft engine disclosed in the above embodiment, it can be understood by technicians in the field that its design constructs an axial first-stage combustion zone I, an axial second-stage combustion zone II, an axial third-stage combustion zone III, and an axial fourth-stage combustion zone IV in the flame tube 4, so that combustion is carried out in an axial stage in the flame tube 4, and the flame stabilization zone is designed to adopt centrifugal atomization + non-premixed combustion to meet the combustion stability requirements, and the core zone adopts direct atomization + concave cavity flame stabilization combustion to meet the high temperature rise requirements. By controlling the oil-gas ratio of each axial combustion zone, the uniformity of the combustion chamber outlet temperature field can be easily adjusted, and the heat load of each axial combustion zone can be controlled to avoid local high-temperature ablation of the flame tube 4 side wall, thereby reducing the demand for cooling air for the turbine guide 7 and the flame tube 4 side wall cooling, and can be well suitable for adjusting the uniformity of the combustion chamber outlet temperature field and reducing the high-pressure air for cooling the flame tube side wall.

[0068] The axially staged circumferentially partitioned combustion chamber of the aircraft engine disclosed in the above embodiment, when in operation, the airflow ejected backward from the circumferentially partitioned jet holes C at the tail end of each axially staged circumferentially partitioned straightening baffle 6 can aerodynamically isolate the axial four-stage combustion zone IV into multiple fan-shaped areas, each fan-shaped area facing the space between the blades of the turbine guide vane 7, so that the high-temperature combustion gas flowing out of the axial four-stage combustion zone IV can directly enter between the blades of the turbine guide vane 7, avoiding direct impact on the leading edge of the turbine guide vane 7 and ablating the leading edge of the turbine guide vane 7. In addition, the airflow ejected backward from the circumferentially partitioned jet holes C at the tail end of each axially staged circumferentially partitioned straightening baffle 6 faces the leading edge of the turbine guide vane 7, which can efficiently cool the leading edge of the turbine guide vane 7, and further flows backward to adhere to the wall surfaces on both sides of the turbine guide vane 7 to form a protective air film, which effectively thermally protects each blade of the turbine guide vane 7, thereby reducing the demand for cooling air for cooling the turbine guide vane 7.

[0069] In some optional embodiments, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, there are three head mounting holes A and their air centrifugal atomizing fuel nozzles 5 located between two adjacent axially staged circumferentially partitioned rectifying baffles 6, which are distributed radially, and the radial height occupied by the head mounting holes A is similar to the circumferential width.

[0070] In some optional embodiments, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, each axially staged circumferentially partitioned rectifying baffle 6 has a through flame connecting channel F on its side wall, connecting each axial primary combustion zone I, and can perform flame connecting in the circumferential direction for each axial primary combustion zone I, thereby ensuring the reliability of ignition and combustion;

[0071] There are two head air inlet holes B corresponding to each axially graded circumferentially partitioned rectifying baffle 6, located on both sides of the cross-flame channel F;

[0072] The aero-engine axially staged circumferentially partitioned combustion chamber further comprises:

[0073] A plurality of baffle inner support plates 10 are laterally supported within the front end of each axially graded and circumferentially partitioned rectifying baffle 6 and are located between the corresponding two head air inlet holes B.

[0074] In some optional embodiments, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, each axially staged circumferentially partitioned straightening baffle 6 has a plurality of circumferentially partitioned jet holes C at the tail end thereof, the number of which is greater than three, distributed radially, axially facing the leading edge of the turbine guide vane 7, and the distance between the circumferentially partitioned jet holes C and the leading edge of the turbine guide vane 7 is not less than 0.5 times the axial length of the turbine guide vane 7 and not more than 1 times the height of the axially staged circumferentially partitioned straightening baffle 6;

[0075] or,

[0076] The circumferentially partitioned jet holes C at the tail of each axially graded circumferentially partitioned rectifying baffle 6 are strip-shaped and extend in the radial direction.

[0077] In some optional embodiments, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, there are multiple front concave cavity fuel jet nozzles 8 and rear concave cavity fuel jet nozzles 9 corresponding to each axially staged circumferentially partitioned rectifying baffle 6, and the number is greater than three. They are distributed in multiple points along the radial direction, and the spacing is not less than 3 times the nozzle aperture and not greater than the axial length of the cavity.

[0078] In some optional embodiments, in the above-mentioned axially graded circumferentially partitioned combustion chamber of the aircraft engine, a plurality of partition film cooling holes are provided on the side walls of each axially graded circumferentially partitioned rectifying partition 6, which are distributed within the range of the axial secondary combustion zone II and the axial tertiary combustion zone III. The high-pressure air entering each axially graded circumferentially partitioned rectifying partition 6 can flow out through the partition film cooling holes, forming a protective air film on the wall surfaces on both sides of the axially graded circumferentially partitioned rectifying partition 6, thereby providing thermal protection for the axially graded circumferentially partitioned rectifying partition 6 and preventing ablation.

[0079] The baffle film cooling holes located in the front concave cavity D and the rear concave cavity E on the side walls of each axially graded circumferentially partitioned rectifying baffle 6 are inclined along the reflow direction in the front concave cavity D and the rear concave cavity E to avoid damaging the reflow zone in the front concave cavity D and the rear concave cavity E and affecting the stability of combustion.

[0080] In some optional embodiments, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, there are multiple flame tube air film cooling holes on the side wall of the flame tube 4, which are distributed in the axial secondary combustion zone II, the axial tertiary combustion zone III, and the axial fourth combustion zone IV. Part of the high-pressure air entering the outer casing 1 of the combustion chamber, the inner casing 2 of the combustion chamber and the flame tube 4 can enter the interior of the flame tube 4 through the flame tube air film cooling holes, forming an air film on the inner side of the flame tube 4, thereby performing thermal protection on the side wall of the flame tube 4 and avoiding ablation.

[0081] In some optional embodiments, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, the turbine guide vane 7 is formed at the tail of the flame tube 4, and the outer wall has a plurality of turbine guide vane air film cooling holes, which are distributed between the blades of the turbine guide vane 7. The high-pressure air entering between the outer casing 1 of the combustion chamber and the flame tube 4 can enter the interior of the turbine guide vane 7 through the turbine guide vane air film cooling holes after being discharged, forming an air film on the inner side of the outer wall of the turbine guide vane 7, thereby providing thermal protection for the outer wall of the turbine guide vane 7 and avoiding ablation.

[0082] In some optional embodiments, in the above-mentioned axially staged circumferentially partitioned combustion chamber of the aircraft engine, the flame tube 4 and its turbine guide vane 7 are made of SiCf / SiC ceramic-based composite materials, which have extremely high temperature resistance, and can reduce the need for cooling air in the cold zone of the flame tube 4 and its turbine guide vane 7. Under this scheme, there is no need to open air film cooling holes on the side wall of the flame tube 4 located within the axial first-level combustion zone I, which can greatly reduce the demand for cooling air.

[0083] The various embodiments in the specification are described in a progressive manner, and each embodiment focuses on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to in detail.

[0084] So far, the technical solution of the present application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the scope of protection of the present application is obviously not limited to these specific embodiments. Without departing from the principles of the present application, those skilled in the art can make equivalent changes or replacements to the relevant technical features, and the technical solutions after these changes or replacements will fall within the scope of protection of the present application.

Claims

1. An axially staged and circumferentially partitioned combustion chamber for an aircraft engine, characterized in that: include: Combustion chamber outer casing (1); The combustion chamber inner casing (2) is arranged in the combustion chamber outer casing (1); The flame tube (4) is arranged between the combustion chamber outer casing (1) and the combustion chamber inner casing (2), and the head portion has a plurality of head mounting holes (A) and head air inlet holes (B) distributed alternately along the circumferential direction; A plurality of air centrifugal atomizing fuel nozzles (5) are arranged in each head mounting hole (A); A plurality of axially graded circumferentially partitioned rectifying baffles (6) are circumferentially arranged inside the flame tube (4), at the front end of the flame tube (4), facing each blade of the turbine guide device (7), with a circumferentially partitioned jet hole (C) at the tail end, and a front concave cavity (D) and a rear concave cavity (E) distributed on both side walls, which are hollow inside and connected to each head air inlet hole (B); A plurality of front concave cavity fuel injection nozzles (8) are connected to the side walls of each axially graded and circumferentially partitioned rectifying baffle (6) and are located at the front side of the front concave cavity (D); A plurality of rear concave cavity fuel injection nozzles (9) are connected to the side walls of each axially graded and circumferentially partitioned rectifying baffle (6) and are located at the front side of the rear concave cavity (E); in, In the flame tube (4), an axial primary combustion zone (I) is formed in front of the front concave cavity (D); In the flame tube (4), an axial secondary combustion zone (II) is formed between the front concave cavity (D) and the rear concave cavity (E); In the flame tube (4), an axial three-stage combustion zone (III) is formed between the rear concave cavity (E) and the circumferential partitioned jet holes (C); In the flame tube (4), an axial four-stage combustion zone (IV) is formed behind the circumferentially partitioned jet holes (C).

2. The axially staged and circumferentially partitioned combustion chamber of an aircraft engine according to claim 1, characterized in that: There are three head mounting holes (A) and air centrifugal atomizing fuel nozzles (5) located between two adjacent axially graded circumferentially partitioned rectifying partitions (6) and distributed along the radial direction.

3. The axially staged and circumferentially partitioned combustion chamber of an aircraft engine according to claim 1, characterized in that: The side walls of each axially graded and circumferentially partitioned rectifying baffle (6) are provided with a through flame channel (F) which is connected to each axial primary combustion zone (I); There are two head air inlet holes (B) corresponding to each axially graded circumferentially partitioned rectifying baffle (6), which are located on both sides of the cross-flame channel (F); The aero-engine axially staged circumferentially partitioned combustion chamber further comprises: A plurality of baffle inner support plates (10) are laterally supported inside the front end of each axially graded circumferentially partitioned rectifying baffle (6) and are located between the corresponding two head air inlet holes (B).

4. The axially staged and circumferentially partitioned combustion chamber of an aircraft engine according to claim 1, characterized in that: Each axially graded circumferentially partitioned rectifying baffle (6) has a plurality of circumferentially partitioned jet holes (C) at the tail end thereof, which are distributed in the radial direction; or, The circumferentially partitioned jet holes (C) at the tail of each axially graded circumferentially partitioned rectifying baffle (6) are strip-shaped and extend in the radial direction.

5. The axially staged and circumferentially partitioned combustion chamber of an aircraft engine according to claim 1, characterized in that: There are a plurality of front concave cavity fuel jet nozzles (8) and rear concave cavity fuel jet nozzles (9) corresponding to each axially graded circumferentially partitioned rectifying baffle (6), which are distributed at multiple points along the radial direction.

6. The axially staged and circumferentially partitioned combustion chamber of an aircraft engine according to claim 1, characterized in that: A plurality of baffle film cooling holes are provided on the side walls of each axially graded and circumferentially partitioned rectifying baffle (6), and are distributed within the axial secondary combustion zone (II) and the axial tertiary combustion zone (III); The baffle air film cooling holes located in the front concave cavity (D) and the rear concave cavity (E) on the side walls of each axially graded and circumferentially partitioned rectifying baffle (6) are inclined along the return flow direction in the front concave cavity (D) and the rear concave cavity (E).

7. The axially staged and circumferentially partitioned combustion chamber of an aircraft engine according to claim 1, characterized in that: A plurality of flame tube air film cooling holes are provided on the side wall of the flame tube (4), which are distributed within the axial secondary combustion zone (II), the axial tertiary combustion zone (III), and the axial quaternary combustion zone (IV).

8. The axially staged and circumferentially partitioned combustion chamber of an aircraft engine according to claim 1, characterized in that: The turbine guide vane (7) is formed at the tail of the flame tube (4), and the outer wall thereof has a plurality of turbine guide vane air film cooling holes distributed between the blades of the turbine guide vane (7).

9. The axially staged and circumferentially partitioned combustion chamber of an aircraft engine according to claim 8, characterized in that: The flame tube (4) and the turbine guide vane (7) thereof are made of SiCf / SiC ceramic matrix composite material.

10. The axially staged and circumferentially partitioned combustion chamber of an aircraft engine according to claim 8, characterized in that: Also includes: The diffuser (3) is arranged at the inlet portion between the combustion chamber outer casing (1) and the combustion chamber inner casing (2).

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