An aircraft engine afterburner

Abstract

The present application belongs to the technical field of design of afterburner combustion chambers for aircraft engines, and specifically relates to an afterburner combustion chamber for aircraft engines, comprising: an annular outer wall, the front end of which is connected to the rear end of an outer casing of a turbine; a merging ring, arranged inside the annular outer wall, the front end of which is connected to the rear end of an inner casing of a turbine, forming an outer duct between the annular outer wall, and the inner diameter of the front section gradually expands, and the inner diameter of the rear section gradually expands; an inner cone, arranged inside the merging ring, the front end of which is connected to the rear end of an inner ring of a last-stage stator of a turbine, forming an inner duct between the merging ring, and the outer diameter of the front section gradually expands, exceeding the tip height of a turbine blade, and the outer diameter of the rear section gradually shrinks to form a cone; the flow area of ​​the flow path formed between the inner cone and the merging ring gradually expands, and is a pressure-expanding flow path.

Description

Technical Field

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

[0002] The turbine is a high-temperature component in an aircraft engine, and its blades are the main source of infrared radiation behind the aircraft engine.

[0003] The afterburner is located behind the turbine and can effectively increase the thrust of the aircraft engine without changing the main engine. It is the main component visible from the rear of the aircraft engine.

[0004] The current afterburner cannot effectively shield the turbine blades and cannot cool itself efficiently, especially the inner cone, which limits the improvement of the infrared detection performance of aircraft engines.

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

[0006] 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 this application, 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

[0007] The purpose of the present application is to provide an aircraft engine afterburner to overcome or alleviate at least one of the existing technical deficiencies.

[0008] The technical solution of this application is:

[0009] An aircraft engine afterburner chamber, comprising:

[0010] An annular outer wall, with a front end connected to the rear end of the turbine outer casing;

[0011] The confluence ring is arranged inside the annular outer wall, with its front end connected to the rear end of the turbine inner casing, forming an outer duct between it and the annular outer wall, and the inner diameter of the front section gradually expands, and the inner diameter of the rear section gradually expands;

[0012] The inner cone is arranged in the merging ring, with its front end connected to the rear end of the inner ring of the turbine's final stator, forming an inner channel between the merging ring and the front section. The outer diameter of the front section gradually expands to exceed the tip height of the turbine blade, and the outer diameter of the rear section gradually shrinks to form a cone shape.

[0013] The flow area of ​​the flow path between the inner cone converging rings gradually expands, forming a pressure-expanding flow path.

[0014] According to at least one embodiment of the present application, in the above-mentioned aircraft engine afterburner, the side wall of the rear section of the inner cone has multiple rows of film cooling holes distributed along the circumferential direction;

[0015] The aero-engine afterburner chamber further comprises:

[0016] The air-cooled support plate is circumferentially supported between the merging ring and the front section of the inner cone, located in the inner duct. It is hollow inside and is connected to the outer duct and the interior of the inner cone through grooves on the merging ring and the inner cone.

[0017] According to at least one embodiment of the present application, in the above-mentioned aircraft engine afterburner, the rear end of the merging ring has a plurality of notches distributed along the circumferential direction;

[0018] The aero-engine afterburner chamber further comprises:

[0019] Multiple radial stabilizers, which are hollow structures with a V-shaped cross-section, are stuck in each notch. The upper section extends into the outer duct, and the side wall has multiple air-cooling air inlet holes. The lower section is in the inner duct, and the side wall has multiple air-cooling air outlet holes and fuel ejection holes.

[0020] A plurality of fuel injection rods are arranged through the annular outer wall and extend into each radial stabilizer. The side wall is connected to a plurality of fuel injection nozzles, which extend from each fuel injection hole.

[0021] According to at least one embodiment of the present application, in the above-mentioned aircraft engine afterburner, the radial stabilizers are distributed at long and short intervals in the circumferential direction.

[0022] According to at least one embodiment of the present application, the above-mentioned aircraft engine afterburner further includes:

[0023] A plurality of flame transfer grooves are arc-shaped and circumferentially connected between the radial stabilizers.

[0024] According to at least one embodiment of the present application, the above-mentioned aircraft engine afterburner further includes:

[0025] The fuel main pipe is annular in shape, sleeved on the outer circumference of the annular outer wall, and connected to each injection rod.

[0026] According to at least one embodiment of the present application, the above-mentioned aircraft engine afterburner further includes:

[0027] A plurality of brackets are supported circumferentially between the annular outer wall and the confluence ring.

[0028] According to at least one embodiment of the present application, in the above-mentioned aircraft engine afterburner, each bracket is located at the position where the confluence ring has the maximum diameter.

[0029] According to at least one embodiment of the present application, the above-mentioned aircraft engine afterburner further includes:

[0030] The anti-vibration and heat-insulating screen is annular and is arranged inside the rear end of the annular outer wall. It has a plurality of cooling circulation holes and anti-vibration holes distributed along the circumferential direction.

[0031] This application has at least the following beneficial technical effects:

[0032] Provided is an aircraft engine afterburner chamber, wherein an inner cone is improved in design. The outer diameter of the front section of the inner cone is gradually enlarged to exceed the tip height of the turbine blades. In this way, the turbine blades can be completely shielded in the rear direction of the aircraft engine without adding additional components, thereby reducing the infrared radiation of the turbine blades in the rear direction of the aircraft engine and ensuring the infrared detection performance of the aircraft engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] Figure 1 Schematic diagram of an aircraft engine afterburner provided in an embodiment of the present application;

[0034] Figure 2 This is a rear view of the afterburner chamber of an aircraft engine provided by an embodiment of the present application;

[0035] Figure 3 is a schematic diagram of a confluence ring provided in an embodiment of the present application;

[0036] Figure 4 is a schematic diagram of an inner cone provided in an embodiment of the present application;

[0037] Figure 5 Schematic diagram of an air-cooled support plate provided in an embodiment of the present application;

[0038] Figure 6 Schematic diagram of key dimension control of an aircraft engine afterburner chamber according to an embodiment of the present application;

[0039] Figure 7 This is a schematic diagram of the air cooling flow path of the afterburner of an aircraft engine provided by an embodiment of the present application;

[0040] in:

[0041] 1-annular outer wall; 2-turbine outer casing; 3-merging ring; 4-turbine inner casing; 5-inner cone; 6-turbine last stage stator inner ring; 7-turbine blades; 8-air cooling support plate; 9-radial stabilizer; 10-injection rod; 11-bracket; 12-anti-vibration and heat insulation screen.

[0042] 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

[0043] 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.

[0044] 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.

[0045] 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.

[0046] The following is combined with Figures 1 to 7 This application is described in further detail.

[0047] An aircraft engine afterburner, such as Figure 1-Figure 2 As shown, including:

[0048] The annular outer wall 1 has a front end connected to the rear end of the turbine outer casing 2;

[0049] The confluence ring 3 is set in the annular outer wall 1, and its front end is connected to the rear end of the turbine inner casing 4. It forms an outer duct with the annular outer wall 1, and the inner diameter of the front section gradually expands, and the inner diameter of the rear section gradually expands, such as Figure 3 As shown, the inner diameter of the front segment can be expanded rapidly at a faster speed, and the inner diameter of the back segment can be expanded slowly at a lower speed;

[0050] The inner cone 5 is arranged in the merging ring 3, and its front end is connected to the rear end of the inner ring 6 of the final stage stator of the turbine, and an inner channel is formed between the merging ring 3. The outer diameter of the front section gradually expands, exceeding the tip height of the turbine blade 7, and the outer diameter of the rear section gradually shrinks to form a cone shape, such as Figure 4 As shown, the outer diameter of the front section can be expanded at a relatively fast speed, and the specific expansion speed can be slightly smaller than the expansion speed of the inner diameter of the front section of the merging ring 3, and the outer diameter of the rear section can be contracted at a relatively fast speed, and the contraction speed can be significantly larger than the expansion speed of the inner diameter of the rear section of the merging ring 3;

[0051] The flow area of ​​the flow path formed between the inner cone 5 and the merging ring 3 gradually expands, forming a pressure-expanding flow path.

[0052] As for the afterburner combustion chamber of the aircraft engine disclosed in the above embodiment, it can be understood by those skilled in the art that the inner cone 5 is improved in design, and the outer diameter of the front section of the inner cone 5 is gradually enlarged to exceed the tip height of the turbine blade 7. In this way, the turbine blade 7 can be completely blocked in the rear direction of the aircraft engine without adding additional components, thereby reducing the infrared radiation of the turbine blade 7 in the rear direction of the aircraft engine and ensuring the infrared detection performance of the aircraft engine.

[0053] As for the afterburner combustion chamber of the aviation engine disclosed in the above embodiment, it can be understood by those skilled in the art that, based on the design of the outer diameter of the front section of the inner cone 5 gradually expanding and the outer diameter of the rear section gradually shrinking, the inner diameter of the front section of the merging ring 3 is adaptively designed to gradually expand and the inner diameter of the rear section gradually shrinks, thereby constructing a diffusion flow path and gradually expanding the flow area of ​​the inner channel. This allows the inner airflow entering the afterburner to undergo initial diffusion and deceleration between the merging ring 3 and the front section of the inner cone 5, complete diffusion and deceleration between the merging ring 3 and the rear section of the inner cone 5, reach the trailing edge of the merging ring 3, mix with the outer airflow, and be able to decelerate to a range where combustion can be stably organized. This can be achieved by designing the profiles of the merging ring 3 and the inner cone 5 to ensure that no aerodynamic separation occurs.

[0054] For the afterburner chamber of the aero-engine disclosed in the above embodiment, the dimensions of its key parts can be designed as follows, referring to Figure 6 :

[0055] b1 ≥ r1;

[0056] π(a22 -a1 2 )<π(b2 2 -b1 2 );

[0057] in,

[0058] r1 is the distance from the tip of turbine blade 7 to the center axis of the aircraft engine;

[0059] b1 is the distance from the maximum outer diameter of the inner cone 5 to the center axis of the aircraft engine;

[0060] b2 is the distance from the maximum inner diameter of the merging ring 3 to the center axis of the aircraft engine;

[0061] a1 is the distance from the outer diameter of the front end of the inner cone 5 to the center axis of the aircraft engine;

[0062] a2 is the distance from the inner diameter of the front end of the merging ring 3 to the central axis of the aircraft engine.

[0063] In some optional embodiments, in the above-mentioned aircraft engine afterburner, the inner cone 5 is subjected to radar modification design to enhance its own radar performance.

[0064] In some optional embodiments, in the above-mentioned aircraft engine afterburner, the rear sidewall of the inner cone 5 has multiple rows of film cooling holes distributed along the circumferential direction;

[0065] The aero-engine afterburner chamber further comprises:

[0066] Multiple air-cooled support plates 8 are supported circumferentially between the merging ring 3 and the front section of the inner cone 5, and are located in the inner duct. The inner duct is hollow and is connected to the outer duct and the inner cone 5 through the grooves on the merging ring 3 and the inner cone 5. Figure 5 shown.

[0067] For the afterburner combustion chamber of the aircraft engine disclosed in the above embodiment, it can be understood by those skilled in the art that part of the by-pass airflow entering the afterburner combustion chamber can flow into the inner cone 5 through the air-cooling support plate 8 and flow out through the air film cooling holes on the side wall of the rear section of the inner cone 5, thereby cooling the air-cooling support plate 8 and the inner cone 5 along the way. The by-pass air can be used to efficiently cool the inner cone 5, reduce the infrared radiation of the inner cone 5, and ensure the infrared detection performance of the aircraft engine. Figure 7 shown.

[0068] In some optional embodiments, in the above-mentioned aircraft engine afterburner, the rear end of the merging ring 3 has a plurality of notches distributed along the circumferential direction;

[0069] The aero-engine afterburner chamber further comprises:

[0070] Multiple radial stabilizers 9, which are hollow structures with a V-shaped cross section, are stuck in each notch. The upper section extends into the outer duct, and the side wall has multiple air-cooling air inlet holes. The lower section is located in the inner duct, and the side wall has multiple air-cooling air outlet holes and fuel ejection holes.

[0071] A plurality of fuel injection rods 10 are provided through the annular outer wall 1 and extend into each radial stabilizer 9. The side wall is connected to a plurality of fuel injection nozzles extending from each fuel injection hole.

[0072] For the aircraft engine afterburner disclosed in the above embodiment, technical personnel in the field can understand that the internal airflow entering the afterburner can be mixed with the external airflow after flowing out at the rear end of the merging ring 3, and at the same time, a recirculation zone can be formed at the trailing edge of each radial stabilizer 9. The fuel sprayed by each injection rod 10 through each injection nozzle can be quickly atomized and enter the corresponding recirculation zone to form a local oil-rich zone, which can be easily ignited and organized for combustion, thereby ensuring the reliability of the ignition of the afterburner and the stability of its combustion.

[0073] For the afterburner combustion chamber of the aircraft engine disclosed in the above embodiment, it can be understood by those skilled in the art that the outer airflow entering the combustion chamber can enter the radial stabilizer 9 through the air-cooling air inlet hole and be discharged through the air-cooling air outlet hole, thereby cooling the radial stabilizer 9 and the fuel injection rod 10 along the way, thereby protecting the radial stabilizer 9 and the fuel injection rod 10 from high temperature damage. Figure 7 shown.

[0074] In some optional embodiments, in the above-mentioned aircraft engine afterburner, the radial stabilizers 9 are distributed at long and short intervals in the circumferential direction.

[0075] In some optional embodiments, the above-mentioned aircraft engine afterburner further includes:

[0076] A plurality of flame transfer grooves are arc-shaped and circumferentially connected between the radial stabilizers 9 so as to be able to transfer flame in the circumferential direction, thereby ensuring the reliability of the ignition of the afterburner and the stability of its combustion.

[0077] In some optional embodiments, the above-mentioned aircraft engine afterburner further includes:

[0078] The fuel main pipe is annular and is sleeved on the outer circumference of the annular outer wall 1 to connect the various injection rods 10. It can be connected to the oil source through pipelines to supply oil to the various injection rods 10.

[0079] In some optional embodiments, the above-mentioned aircraft engine afterburner further includes:

[0080] A plurality of brackets 11 are supported between the annular outer wall 1 and the merging ring 3 along the circumferential direction.

[0081] In some optional embodiments, in the above-mentioned aircraft engine afterburner, each bracket 11 is located at the position where the diameter of the merging ring 3 is the largest, so as to reliably support the merging ring 3.

[0082] In some optional embodiments, the above-mentioned aircraft engine afterburner further includes:

[0083] The anti-vibration and heat-insulating screen 12 is annular and is arranged inside the rear end of the annular outer wall 1. It has a plurality of cooling circulation holes and anti-vibration holes distributed along the circumferential direction.

[0084] For the aircraft engine afterburner combustion chamber disclosed in the above embodiment, technical personnel in the field can understand that the external airflow can partially flow into the anti-vibration and heat insulation screen 12 and the annular outer wall 1, and then flow out through the cooling circulation holes and anti-vibration holes on the anti-vibration and heat insulation screen 12, thereby cooling and insulating the rear end of the annular outer wall 1, and suppressing oscillating combustion in the afterburner combustion chamber as a whole.

[0085] 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.

[0086] 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 aircraft engine afterburner, characterized in that: include: An annular outer wall (1), the front end of which is connected to the rear end of the turbine outer casing (2); The confluence ring (3) is arranged in the annular outer wall (1), the front end of which is connected to the rear end of the turbine inner casing (4), and an outer duct is formed between the confluence ring and the annular outer wall (1), and the inner diameter of the front section gradually expands, and the inner diameter of the rear section gradually expands; The inner cone (5) is arranged in the merging ring (3), and its front end is connected to the rear end of the inner ring (6) of the final stage stator of the turbine, and an inner channel is formed between the inner cone and the merging ring (3). The outer diameter of the front section gradually expands to exceed the tip height of the turbine blade (7), and the outer diameter of the rear section gradually shrinks to form a cone shape. The flow area of ​​the flow path formed between the inner cone (5) and the merging ring (3) gradually expands to form a pressure-expanding flow path.

2. The aircraft engine afterburner according to claim 1, characterized in that: The rear side wall of the inner cone (5) is provided with multiple rows of air film cooling holes distributed along the circumferential direction; The aero-engine afterburner chamber further comprises: A plurality of air-cooling support plates (8) are supported circumferentially between the merging ring (3) and the front section of the inner cone (5), are located in the inner duct, are hollow inside, and are connected to the outer duct and the interior of the inner cone (5) through grooves on the merging ring (3) and the inner cone (5).

3. The aircraft engine afterburner according to claim 1, characterized in that: The rear end of the merging ring (3) has a plurality of notches distributed along the circumferential direction; The aero-engine afterburner chamber further comprises: A plurality of radial stabilizers (9) are hollow structures with a V-shaped cross section, which are inserted into each notch, with the upper section extending into the outer duct and the side wall having a plurality of air-cooling air inlet holes, and the lower section being in the inner duct and the side wall having a plurality of air-cooling air outlet holes and fuel ejection holes; A plurality of fuel injection rods (10) are arranged through the annular outer wall (1) and extend into each radial stabilizer (9); the side wall is connected to a plurality of fuel injection nozzles, which extend from each fuel injection hole.

4. The aircraft engine afterburner according to claim 3, characterized in that: The radial stabilizers (9) are distributed at long and short intervals in the circumferential direction.

5. The aircraft engine afterburner according to claim 3, characterized in that: Also includes: A plurality of flame transfer grooves are arc-shaped and connected between the radial stabilizers (9) along the circumferential direction.

6. The aircraft engine afterburner according to claim 3, characterized in that: Also includes: The fuel main pipe is annular, sleeved on the outer periphery of the annular outer wall (1), and connected to each fuel injection rod (10).

7. The aircraft engine afterburner according to claim 1, characterized in that: Also includes: A plurality of brackets (11) are supported circumferentially between the annular outer wall (1) and the merging ring (3).

8. The aircraft engine afterburner according to claim 1, characterized in that: Each bracket (11) is located at the position where the diameter of the merging ring (3) is the largest.

9. The aircraft engine afterburner according to claim 1, characterized in that: Also includes: The anti-vibration heat-insulating screen (12) is annular and is arranged inside the rear end of the annular outer wall (1). It has a plurality of cooling circulation holes and anti-vibration holes distributed along the circumferential direction.

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