A dual-element nozzle for nacelle bleed air cooling of an aircraft engine

Abstract

Provided is a two-dimensional nozzle for bleed air cooling in an aircraft engine cabin. High-temperature gas in the aircraft engine can be discharged through a channel formed by a circular square section, a convergent section, an expansion section, a front section side wall, and a rear section side wall. At the same time, when the pressure in the two cooling bleed air cavities is lower than the pressure in the aircraft engine cabin, the low-temperature gas in the aircraft engine cabin can be used to cool the internal and external components along the way in addition to the external air flow, thereby increasing the cooling effect on the internal and external components, effectively suppressing the infrared radiation characteristics of the internal and external components, ensuring the infrared stealth performance of the aircraft engine, and avoiding the problems of internal components being ablated and external components being deformed and warped. Moreover, when the pressure in the two cooling bleed air cavities is higher than the pressure in the aircraft engine cabin, the high-temperature gas in the aircraft engine can be prevented from flowing reversely into the aircraft engine cabin along the way and damaging the components in the aircraft engine cabin.

Description

Technical Field

[0001] The present application belongs to the technical field of aircraft engine dual nozzle design, and specifically relates to an aircraft engine cabin bleed air cooling dual nozzle. Background Art

[0002] The dual-element nozzle of an aero-engine is installed at the end of the aero-engine. It uses the contraction and expansion movement of internal components such as the convergent section and the divergent section for vector adjustment. It is also designed with external components such as external adjustment plates and side wall covers to protect the internal components from direct damage by external forces, to rectify the exterior, and to reduce rear body resistance.

[0003] The internal components of the dual nozzle of an aircraft engine are in direct contact with the high-temperature airflow, subjected to extremely high temperatures, easily ablated, and have obvious infrared radiation characteristics. To this end, at present, the internal components are mostly cooled by external airflow to avoid internal components from being ablated and suppress infrared radiation characteristics. However, the demand for external airflow is difficult to match the state of the aircraft engine in many cases, and it does not have the ability to actively adjust and cannot achieve the expected results. In addition, the internal components of the dual nozzle of an aircraft engine will be irradiated by the high-temperature airflow, subjected to high temperatures, and prone to deformation and warping problems. The infrared radiation characteristics are obvious, which is not conducive to the infrared stealth performance of the aircraft engine.

[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 the present application. In the absence of clear evidence that the above content has been disclosed on the filing date of the present application, the above background technology should not be used to evaluate the novelty and creativity of the present application. Summary of the Invention

[0006] The purpose of the present application is to provide an aircraft engine nacelle bleed air cooling dual nozzle to overcome or alleviate at least one of the existing technical defects.

[0007] The technical solution of this application is:

[0008] An aircraft engine nacelle bleed air cooling dual nozzle, comprising:

[0009] The circular rotating square section 1 has a front end connected to the rear end of the aircraft engine, and the outer wall of the rear end has a connecting edge;

[0010] The front ends of the two convergent sections 2 are hinged to the connecting edges of the two opposite side walls at the rear end of the circular rotating square section 1. The connecting edges of the two opposite side walls at the rear end of the circular rotating square section 1 are provided with a cabin air bleed hole A; the cabin air bleed hole A is connected to the aircraft engine cabin;

[0011] Two expansion sections 3, the front ends of which are hinged to the rear ends of the two convergence sections 2;

[0012] Two outer adjustment plates 4 are located outside the two convergent sections 2 and the expansion section 3, with their front ends hinged to the connecting edges of the two opposite side walls at the rear end of the circular rotating square section 1. The rear end side walls are provided with a plurality of cooling exhaust holes B;

[0013] The front ends of the two front side walls 5 are connected to the connecting edges of the other two opposite side walls at the rear end of the circular rotating square section 1, and form a cooling air bleed cavity between the two convergent sections 2, the expansion section 3, and the outer adjustment plate 4. A plurality of cooling communication holes C are distributed on the cooling air bleed cavity. The cooling air bleed cavity connects to the cabin air bleed holes A, the cooling exhaust holes B, and the cooling communication holes C.

[0014] The front ends of the two rear side walls 6 are connected to the rear ends of the two front side walls 5, and a rear exhaust gap D is formed between the rear ends of the two front side walls 5;

[0015] Two side wall covers 7 are located outside the two front side walls 5 and the rear side walls 6, with their front ends correspondingly connected to the connecting edges of the two opposite side walls at the rear end of the circular rotating square section 1, and connected to the two front side walls 5 and the rear side walls 6 along the edges, forming a cooling exhaust cavity between the two front side walls 5 and the rear side walls 6; the cooling exhaust cavity is connected to each cooling communication hole C and the rear exhaust slit D;

[0016] Two valves 8 are located in the two cooling air bleed cavities and connected to the corresponding two connecting edges;

[0017] When the pressure in the two cooling air bleed cavities is greater than the pressure in the aircraft engine cabin, the two valves 8 are pressed against the connecting edges of the two opposite side walls at the rear end of the circular rotating square section 1 under the action of gravity, thereby blocking the two cabin air bleed holes A;

[0018] When the pressure in the two cooling air bleed cavities is lower than the pressure in the aircraft engine cabin, the two valves 8 overcome gravity and tilt upward under the pressure difference between the aircraft engine cabin and the cooling air bleed cavities, thereby releasing the blockage of the two cabin air bleed holes A.

[0019] According to at least one embodiment of the present application, in the above-mentioned aircraft engine nacelle bleed air cooling dual nozzle, the rear ends of the two expansion sections 3 and the outer adjustment plate 4 are in a pointed cone shape;

[0020] The cooling exhaust holes B are distributed along the rear edges of the two expansion sections 3 and the outer adjustment plate 4;

[0021] The two front side walls 5, the rear side wall 6, and the side wall cover 7 are V-shaped as a whole;

[0022] The rear exhaust slit D is in a V-shape that expands rearward.

[0023] According to at least one embodiment of the present application, in the above-mentioned aircraft engine cabin bleed air cooling dual nozzle, the flow area of the two cabin bleed air holes A is 1 to 2 times the flow area of each cooling exhaust hole B and cooling connecting hole C;

[0024] The flow area of each cooling connecting hole C is 1 to 2 times the flow area of the two rear exhaust slits D.

[0025] According to at least one embodiment of the present application, in the aforementioned aircraft engine nacelle bleed air cooling two-dimensional nozzle, the two outer adjustment blades 4 converge inwardly.

[0026] According to at least one embodiment of the present application, in the above-mentioned aircraft engine nacelle bleed air cooling two-dimensional nozzle, the two valves 8 have hinge holes;

[0027] Also includes:

[0028] Two double-ear brackets 9 are located in the two cooling air ducting cavities and connected to the corresponding two connecting edges;

[0029] Two bolts 10, the screw rods of which pass through the ear holes and hinge holes on the two double-ear brackets 9;

[0030] Two nuts 11 are screwed onto the two bolts 10 to restrict the two valves 8 between the ears of the double-ear bracket 9.

[0031] According to at least one embodiment of the present application, in the above-mentioned aircraft engine cabin bleed air cooling dual nozzle, the single-side gap between the two valves 8 and the ears of the double-ear bracket 9 is controlled to be 0.2-0.5 mm;

[0032] The single-side gap between the heads of the two bolts 10, the nuts 11 and the ears of the double-ear bracket 9 is controlled to be 0.2-0.3 mm.

[0033] This application has at least the following beneficial technical effects:

[0034] Provided is a dual-element nozzle for bleed air cooling in an aircraft engine cabin. High-temperature gas in the aircraft engine can be discharged through a channel formed by a circular square section, a convergent section, an expansion section, a front section side wall, and a rear section side wall. At the same time, when the pressure in the two cooling bleed air cavities is lower than the pressure in the aircraft engine cabin, low-temperature gas in the aircraft engine cabin can be used, in addition to the external air flow, to cool internal and external components such as the convergent section, the expansion section, an outer adjustment plate, the front section side wall, the rear section side wall, and the side wall outer cover along the way, thereby increasing the cooling effect on the internal and external components, effectively suppressing the infrared radiation characteristics of the internal and external components, ensuring the infrared stealth performance of the aircraft engine, and avoiding the problems of internal components being ablated and external components being deformed or warped. The cooling by using the low-temperature gas in the aircraft engine cabin does not require attention to the matching problem with the aircraft engine state, and when the pressure in the two cooling bleed air cavities is higher than the pressure in the aircraft engine cabin, it can prevent the high-temperature gas in the aircraft engine from flowing reversely into the aircraft engine cabin along the way and damaging the components in the aircraft engine cabin. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Figure 1 This is an outline diagram of a dual nozzle for bleed air cooling in an aircraft engine nacelle provided in an embodiment of the present application;

[0036] Figure 2 This is a cross-sectional view of an aircraft engine cabin bleed air cooling dual nozzle provided by an embodiment of the present application, which utilizes low-temperature gas in the cabin for cooling;

[0037] Figure 3 This is a side view of an aircraft engine cabin bleed air cooling dual nozzle provided by an embodiment of the present application, which utilizes low-temperature gas in the cabin to cool the nozzle without the side wall cover;

[0038] Figure 4 This is a schematic diagram of a dual nozzle for bleed air cooling of an aircraft engine cabin provided by an embodiment of the present application, which prevents high-temperature gas from flowing in the reverse direction into the cabin;

[0039] Figure 5 This is a schematic diagram of the valve structure provided in an embodiment of the present application;

[0040] Figure 6 This is a partial cross-sectional view of a dual nozzle for bleed air cooling in an aircraft engine nacelle provided in an embodiment of the present application;

[0041] in:

[0042] 1-Circular rotating square section; 2-Converging section; 3-Expanding section; 4-External adjustment plate; 5-Front section side wall; 6-Rear section side wall; 7-Side wall cover; 8-Valve; 9-Double ear bracket; 10-Bolt; 11-Nut;

[0043] A-cabin air bleed hole;

[0044] B-cooling exhaust hole;

[0045] C- cooling connecting hole.

[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 limiting this application. 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 aircraft engine nacelle bleed air cooling dual nozzle, comprising:

[0052] The circular-to-square section has a front end connected to the rear end of the aircraft engine, and an outer wall of the rear end has a connecting edge;

[0053] The two convergent sections have front ends hinged to the connecting edges of two opposite side walls at the rear end of the circular rotating square section. The two opposite side walls at the rear end of the circular rotating square section are provided with cabin air bleed holes. The cabin air bleed holes are connected to the aircraft engine cabin.

[0054] Two expansion sections, the front end of which is hinged to the rear end of the two convergent sections;

[0055] Two outer adjustment plates are located outside the two convergent sections and the expansion section, with their front ends hinged to the connecting edges of the two opposite side walls at the rear end of the circular-square section, and multiple cooling exhaust holes are distributed on the rear end side walls;

[0056] The front ends of the two front section side walls are connected to the connecting edges of the other two opposite side walls at the rear end of the circular turning square section, and form a cooling air bleed cavity between the two convergent sections, the expansion section, and the outer adjustment plate, on which multiple cooling communication holes are distributed; the cooling air bleed cavity is connected to each cabin air bleed hole, cooling exhaust hole, and cooling communication hole;

[0057] The two rear side walls have front ends connected to the rear ends of the two front side walls, forming a rearward exhaust gap between the two rear ends of the front side walls;

[0058] The two side wall covers are located outside the two front side walls and the rear side walls, with the front ends correspondingly connected to the connecting edges of the two opposite side walls at the rear end of the circular rotating square section, and connected to the two front side walls and the rear side walls along the edges, forming a cooling exhaust cavity between the two front side walls and the rear side walls; the cooling exhaust cavity is connected to each cooling communication hole and the rear exhaust slit; the middle of the cooling exhaust cavity is separated by a partition, forming two independent front and rear cavities corresponding to the front side wall 5 and the rear side wall 6;

[0059] Two valves are located in the two cooling air bleed cavities and connected to the corresponding two connecting edges;

[0060] When the pressure in the two cooling air bleed cavities is greater than the pressure in the aircraft engine cabin, the two valves, under the action of gravity, abut against the connecting edges of the two opposite side walls at the rear end of the circular rotating square section, thereby blocking the two cabin air bleed holes.

[0061] When the pressure in the two cooling air bleed cavities is lower than the pressure in the aircraft engine cabin, the two valves overcome gravity and tilt upward under the action of the pressure difference between the aircraft engine cabin and the cooling air bleed cavities, thereby releasing the blockage of the two cabin air bleed holes.

[0062] As for the dual nozzle for cooling the aircraft engine cabin bleed air disclosed in the above embodiment, it can be understood by those skilled in the art that the high-temperature gas in the aircraft engine can be discharged through the channel formed by the rotating square section 1, the convergent section 2, the expansion section 3, the front section side wall 5, and the rear section side wall 6. At the same time, when the pressure in the two cooling bleed air cavities is lower than the pressure in the aircraft engine cabin, the cabin bleed air hole A can be used to introduce the low-temperature gas in the aircraft engine cabin into the cooling bleed air cavity, and part of the low-temperature gas can be discharged through the cooling exhaust hole B. This part of the low-temperature gas can cool the convergent section 2, the expansion section 3, and the outer adjustment plate 4 along the way, and the other part of the low-temperature gas can enter the cooling exhaust cavity through the cooling connecting hole C, and then be discharged through the rear exhaust gap D. This part of the low-temperature gas can cool the front section side wall 5 and the side wall outer cover 7 along the way, and the low-temperature gas discharged from the rear exhaust gap D can form a cooling air film on the inner side of the rear section side wall 6 to cool the rear section side wall 6 along the way.

[0063] As for the aircraft engine cabin bleed air cooling dual nozzle disclosed in the above embodiment, it can be understood by those skilled in the art that, in addition to the external airflow, it can utilize the low-temperature gas in the aircraft engine cabin to cool the internal and external components such as the convergent section 2, the expansion section 3, the outer adjustment plate 4, the front section side wall 5, the rear section side wall 6, and the side wall outer cover 7 along the way, thereby increasing the cooling effect on the internal and external components, effectively suppressing the infrared radiation characteristics of the internal and external components, ensuring the infrared stealth performance of the aircraft engine, and avoiding the internal components from being ablated and the external components from being deformed or warped. The cooling using the low-temperature gas in the aircraft engine cabin does not require attention to the matching problem with the aircraft engine state.

[0064] As for the dual nozzle for cooling the aircraft engine cabin bleed air disclosed in the above embodiment, those skilled in the art will understand that, in its design, when the pressure in the two cooling bleed air cavities is greater than the pressure in the aircraft engine cabin, the two valves, under the action of gravity, abut against the connecting edges of the two opposite side walls at the rear end of the rotating square section, thereby blocking the two cabin bleed air holes. At this time, the high-temperature gas in the aircraft engine can be prevented from flowing back into the aircraft engine cabin and damaging the components inside the aircraft engine cabin.

[0065] The above embodiment discloses a dual nozzle for bleed air cooling of an aircraft engine nacelle, wherein the connection relationship among the circular-to-square section 1, the convergent section 2, the divergent section 3, and the outer adjustment plate 4 and its specific structural design can be specifically referred to CN113107706, and no further detailed description is given here.

[0066] In some optional embodiments, in the above-mentioned aircraft engine cabin bleed air cooling dual nozzle, the two expansion sections 3 and the rear end of the outer adjustment plate 4 are in a pointed cone shape, and the cooling exhaust holes B are distributed along the rear end edges of the two expansion sections 3 and the outer adjustment plate 4. The two front side walls 5, the rear side wall 6, and the side wall outer cover 7 are V-shaped as a whole, and the rear exhaust gap D is a V-shape that expands backward, so as to ensure the aerodynamic performance of the aircraft engine dual nozzle, and to achieve better cooling effect, and to have better radar stealth performance.

[0067] In some optional embodiments, in the above-mentioned aircraft engine cabin bleed air cooling dual nozzle, the flow area of the two cooling bleed air holes A is 1 to 2 times the flow area of each cooling exhaust hole B and the cooling connecting hole C, and the flow area of each cooling connecting hole C is 1 to 2 times the flow area of the two rear exhaust slits D, so as to ensure the flow capacity of the introduced external cooling air and the cooling effect on the internal and external components.

[0068] In some optional embodiments, in the above-mentioned aircraft engine cabin bleed air cooling dual nozzle, the two outer adjustment plates converge inward, so that the cooling air discharged through each cooling exhaust hole B converges inward, can cover and mix with the high-temperature gas discharged from the aircraft engine dual nozzle, so as to reduce noise, reduce exhaust temperature, and enhance stealth effect.

[0069] In some optional embodiments, in the above-mentioned aircraft engine nacelle bleed air cooling dual nozzle, two valves are provided with hinge holes;

[0070] Also includes:

[0071] Two double-ear brackets are located in the two cooling air ducting cavities and connected to the corresponding two connecting edges;

[0072] Two bolts, the screw rods of which pass through the ear holes and hinge holes on the two double-ear brackets;

[0073] Two nuts are screwed onto the two bolts to restrict the two valves between the ear pieces of the double-ear bracket, thereby realizing the hinge connection of the two valves on the corresponding two connecting edges.

[0074] In some optional embodiments, in the above-mentioned aircraft engine cabin bleed air cooling dual nozzle, the single-sided gap between the two valves and the double-ear bracket ear is controlled at 0.2-0.5 mm, and the single-sided gap between the two bolt heads, nuts and the double-ear bracket ear is controlled at 0.2-0.3 mm, so that there is a small gap between the valves, double-ear bracket, bolts and nuts, so that the valves can move flexibly.

[0075] In the dual nozzle for bleed air cooling of the aircraft engine cabin disclosed in the above embodiment, the cooling air flows out from the holes on the nozzle cover and the side wall of the nozzle, and can be mixed with the high-temperature mainstream, especially discharged from the holes on the nozzle cover. While being discharged to the high-temperature mainstream, it also draws in air from the outside, which can enhance heat exchange with the high-temperature mainstream and reduce the length of the high-temperature core area of the exhaust mainstream behind the nozzle, thereby effectively reducing the infrared radiation intensity of the jet.

[0076] 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.

[0077] 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 nacelle bleed air cooling dual nozzle, characterized in that: include: The circular rotating square section (1) has a front end connected to the rear end of the aircraft engine, and an outer wall of the rear end has a connecting edge; The two convergent sections (2) have front ends hinged to the connecting edges of two opposite side walls at the rear end of the circular rotating square section (1), and the connecting edges of the two opposite side walls at the rear end of the circular rotating square section (1) are provided with cabin air bleed holes (A); the cabin air bleed holes (A) are connected to the aircraft engine cabin; Two expansion sections (3), the front ends of which are hinged to the rear ends of the two convergence sections (2); Two outer adjustment plates (4) are located outside the two convergent sections (2) and the expansion section (3), and their front ends are hinged to the connecting edges of the two opposite side walls at the rear end of the circular rotating square section (1), and a plurality of cooling exhaust holes (B) are distributed on the rear end side wall; The front ends of the two front section side walls (5) are connected to the connecting edges of the other two opposite side walls at the rear end of the circular rotating square section (1), and a cooling air bleed cavity is formed between the two convergent sections (2), the expansion section (3), and the outer adjustment plate (4), and a plurality of cooling connecting holes (C) are distributed on the cooling air bleed cavity; the cooling air bleed cavity is connected to each cabin air bleed hole (A), the cooling exhaust hole (B), and the cooling connecting hole (C); The two rear side walls (6) have front ends connected to the rear ends of the two front side walls (5) and form a rearward exhaust slit (D) between the two rear ends of the front side walls (5); Two side wall covers (7) are located outside the two front side walls (5) and the rear side wall (6), with the front ends correspondingly connected to the connecting edges of the two opposite side walls at the rear end of the circular rotating square section (1), and connected to the two front side walls (5) and the rear side wall (6) along the edges, forming a cooling exhaust cavity between the two front side walls (5) and the rear side wall (6); the cooling exhaust cavity is connected to each cooling connecting hole (C) and the rear exhaust slit (D); Two air valves (8) are located in the two cooling air bleed cavities and connected to the corresponding two connecting edges; When the pressure in the two cooling air bleed cavities is greater than the pressure in the aircraft engine cabin, the two valves (8) are pressed against the connecting edges of the two opposite side walls of the rear end of the circular rotating square section (1) under the action of gravity, thereby blocking the two cabin air bleed holes (A); When the pressure in the two cooling air bleed cavities is lower than the pressure in the aircraft engine cabin, the two valves (8) overcome gravity and tilt upwards under the pressure difference between the aircraft engine cabin and the cooling air bleed cavities, thereby releasing the blockage of the two cabin air bleed holes (A).

2. The aircraft engine nacelle bleed air cooling dual nozzle according to claim 1, characterized in that: The rear ends of the two expansion sections (3) and the outer adjustment piece (4) are in a pointed cone shape; The cooling exhaust holes (B) are distributed along the two expansion sections (3) and the rear end edges of the outer adjustment plate (4); The two front side walls (5), the rear side wall (6), and the side wall outer cover (7) are V-shaped as a whole; The rear exhaust slit (D) is in a V-shape that expands rearward.

3. The aircraft engine nacelle bleed air cooling dual nozzle according to claim 1, characterized in that: The flow area of the two cabin air bleed holes (A) is 1 to 2 times the flow area of each cooling exhaust hole (B) and cooling connecting hole (C); The flow area of each cooling connecting hole (C) is 1 to 2 times the flow area of the two rear exhaust slits (D).

4. The aircraft engine nacelle bleed air cooling dual nozzle according to claim 1, characterized in that: The two outer adjustment pieces (4) converge inwards.

5. The aircraft engine nacelle bleed air cooling dual nozzle according to claim 1, characterized in that: The two valves (8) are provided with hinge holes; Also includes: Two double-ear brackets (9) are located in the two cooling air ducting cavities and connected to the corresponding two connecting edges; Two bolts (10), the screw rods of which pass through the ear holes and hinge holes on the two double-ear brackets (9); Two nuts (11) are screwed onto the two bolts (10) to restrict the two valves (8) between the ears of the double-ear bracket (9).

6. The aircraft engine nacelle bleed air cooling dual nozzle according to claim 1, characterized in that: The single-side gap between the two valves (8) and the ear pieces of the double-ear bracket (9) is controlled to be 0.2-0.5 mm; The single-side gap between the heads of the two bolts (10), the nuts (11) and the ear pieces of the double-ear bracket (9) is controlled to be 0.2-0.3 mm.

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