Main combustion chamber of aero-engine

Abstract

The invention belongs to the technical field of configuration design of main combustion chambers of aero-engines, and particularly relates to a main combustion chamber of an aero-engine. The turning angle of airflow flowing among the last stage of the gas compressor, the main combustion chamber and the turbine guider can be reduced, the smoothness of the airflow flowing among components can be improved, the total pressure loss is reduced, the connection of the airflow flowing among the components is more reasonable, and the improvement of the overall performance of an aero-engine is facilitated.

Description

Technical Field

[0001] The present application belongs to the technical field of aircraft engine main combustion chamber configuration design, and specifically relates to an aircraft engine main combustion chamber. Background Art

[0002] The main combustion chamber is one of the core components of an aircraft engine. It is located after the compressor and before the turbine. The high-pressure gas generated by the compressor is burned to drive the turbine to do work.

[0003] At present, the main combustion chamber is set along the axial direction, and the design of the main combustion chamber is carried out under the boundary condition that the inlet and outlet airflow directions are both axial. The compressor outlet airflow has a certain circumferential rotation. The last-stage straightening blades are set at the compressor outlet to bend the compressor outlet airflow to the near axial direction. A turbine guide vane is set at the turbine inlet. The combustion chamber outlet airflow is greatly deflected in the turbine guide vane, and the deflection angle is on the order of 70 degrees, converting the potential energy of the high-temperature combustion gas into kinetic energy.

[0004] The compressor outlet and turbine guide vane are used to deflect the airflow. Based on the current main combustion chamber structure design, the airflow direction in the compressor, main combustion chamber, and turbine guide vane is as follows: Figure 1 As shown, the airflow undergoes two large deflections, resulting in poor flow between components and large total pressure loss.

[0005] This application is proposed in view of the above-mentioned technical defects. Summary of the invention

[0006] The purpose of the present application is to provide an aircraft engine main combustion chamber to overcome or alleviate at least one of the known technical deficiencies.

[0007] The technical solution of this application is:

[0008] A main combustion chamber of an aircraft engine, wherein the main combustion chamber is tilted along a circumferential direction as a whole;

[0009] The direction in which the main combustion chamber is tilted along the circumferential direction is the twisting tilt direction of the guide vanes in the turbine guide vane.

[0010] According to at least one embodiment of the present application, in the above-mentioned main combustion chamber of the aircraft engine, the angle of inclination of the main combustion chamber along the circumferential direction ranges from 20° to 50°.

[0011] According to at least one embodiment of the present application, in the above-mentioned main combustion chamber of the aircraft engine, the angle of inclination of the main combustion chamber along the circumferential direction is 25°.

[0012] According to at least one embodiment of the present application, in the above-mentioned main combustion chamber of the aircraft engine, the main combustion chamber is a full-annular flame tube.

[0013] According to at least one embodiment of the present application, in the above-mentioned main combustion chamber of the aircraft engine, the main combustion chamber is an annular tube-shaped flame tube.

[0014] This application has at least the following beneficial technical effects:

[0015] Provided is an aircraft engine main combustion chamber. Considering the distribution law of the airflow direction at the compressor last stage straightening blade inlet, the main combustion chamber interior and the turbine guide vane outlet, the main combustion chamber is designed to be tilted along the circumferential direction as a whole, which can reduce the turning angle of the airflow flowing between the compressor last stage, the main combustion chamber and the turbine guide vane, improve the smoothness of the airflow flowing between the components, reduce the total pressure loss, make the connection of the airflow between the components more reasonable, and be beneficial to the improvement of the overall performance of the aircraft engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] Figure 1 It is a schematic diagram of the flow direction of the airflow in the diffuser, main combustion chamber, and turbine guide vane based on the current main combustion chamber structure design;

[0017] Figure 2 It is a schematic diagram of the flame tube provided by the embodiment of the present application being tilted along the circumferential direction as a whole;

[0018] Figure 3 It is a schematic diagram of the flow direction of the airflow in the diffuser, main combustion chamber and turbine guide vane of the main combustion chamber of an aircraft engine provided in an embodiment of the present application.

[0019] In order to better illustrate the present 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 construed as limitations on the present application. DETAILED DESCRIPTION

[0020] In order to make the technical solution and advantages of the present application clearer, the technical solution of the present application will be described in further detail in detail and in detail with reference to the accompanying drawings. It can be understood that the specific embodiments described here are only partial embodiments of the present application, which are only used to explain the present application, not to limit the present application. It should be noted that, for the convenience 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.

[0021] In addition, unless otherwise defined, the technical terms or scientific terms used in the description of this application shall have the usual meanings understood by those of ordinary skill in the art to which this application belongs. The words used in the description of this application to indicate orientation are only used to indicate relative directions or positional relationships. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. The word "include" used in the description of this application indicates that the elements or objects appearing before the word include the elements or objects listed after the word and their equivalents, but does not exclude other elements or objects.

[0022] In addition, it should be noted that, unless otherwise clearly specified and limited, the words "installation", "connection" 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 or a detachable connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium. Technical personnel in the field can understand its specific meaning in this application according to the specific circumstances.

[0023] A main combustion chamber of an aircraft engine, wherein the main combustion chamber is tilted along a circumferential direction as a whole;

[0024] The direction in which the main combustion chamber is tilted along the circumferential direction is the twisting tilting direction of the guide vanes in the turbine guide vane. Figure 2 shown.

[0025] In the main combustion chamber of the aircraft engine disclosed in the above embodiment, the distribution law of the airflow direction at the inlet of the compressor last stage straightening blade, the inside of the main combustion chamber, and the outlet of the turbine guide vane is taken into consideration, and the main combustion chamber is designed to be tilted along the circumferential direction as a whole, so as to reduce the turning angle of the airflow flowing between the compressor last stage, the main combustion chamber, and the turbine guide vane. Figure 3 As shown, the smoothness of airflow between components can be improved, the total pressure loss can be reduced, and the connection of airflow between components can be made more reasonable, which is beneficial to the improvement of the overall performance of the aircraft engine.

[0026] In addition, in the main combustion chamber of the aircraft engine disclosed in the above-mentioned embodiment, the main combustion chamber is designed to be inclined along the circumferential direction as a whole, so as to reduce the turning angle of the airflow flowing between the compressor last stage, the main combustion chamber, and the turbine guide vane, which is beneficial to reducing the difficulty of aerodynamic design of the pre-diffuser of the main combustion chamber and the turbine guide vane of the compressor last stage. The inclination angle and chord length of the guide vanes in the compressor last stage, the main combustion chamber pre-diffuser, and the turbine guide vane can be reduced, which can correspondingly reduce the friction loss of the blade profile, reduce the possibility and severity of aerodynamic separation, shorten the overall length and mass of the aircraft engine, and reduce the number of guide vanes in the compressor last stage, the main combustion chamber pre-diffuser, and the turbine guide vane, thereby reducing the structural complexity and mass of the aircraft engine.

[0027] In some optional embodiments, in the main combustion chamber of the above-mentioned aircraft engine, the circumferential inclination angle of the main combustion chamber is in the range of 20° to 50°, and may be further preferably 25°. This can reduce the turning angle of the airflow between the compressor last stage, the main combustion chamber, and the turbine guide vane while ensuring the basic performance of the combustion chamber and reducing the difficulty of combustion chamber design.

[0028] In some optional embodiments, in the above-mentioned main combustion chamber of the aircraft engine, the main combustion chamber is a full annular main combustion chamber or an annular tube main combustion chamber.

[0029] 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 referenced to each other. In the absence of conflict, the embodiments in this application and the technical features in the embodiments can be combined with each other to obtain new embodiments.

[0030] 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 substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will fall within the scope of protection of the present application.

Claims

1. An aircraft engine main combustion chamber, characterized in that: The main combustion chamber is tilted in the circumferential direction as a whole; The direction in which the main combustion chamber is tilted along the circumferential direction is the twisting tilt direction of the guide vanes in the turbine guide vane.

2. The main combustion chamber of an aircraft engine according to claim 1, characterized in that: The main combustion chamber is inclined at an angle ranging from 20° to 50° along the circumferential direction.

3. The main combustion chamber of the aircraft engine according to claim 2, characterized in that: The main combustion chamber is inclined at an angle of 25° along the circumferential direction.

4. The main combustion chamber of an aircraft engine according to claim 1, characterized in that: The main combustion chamber is a full annular flame tube.

5. The main combustion chamber of the aircraft engine according to claim 1, characterized in that: The main combustion chamber is an annular flame tube.

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