An impact-divergence cooling combustor liner of an aeroengine

Abstract

The present application relates to an impact-divergent cooling flame tube wall of an aero-engine, comprising: an outer annular impact wall having a plurality of impingement cooling holes thereon, a plurality of pairs of return channels inside, and a plurality of arc-shaped flow-blocking protrusions on the inner side; the inlet ends and outlet ends of each pair of return channels extend to the inner side of the outer annular impact wall, and the outlet ends are relatively inward and turn back inward; each arc-shaped flow-blocking protrusion is located between each pair of return channels; a plurality of inner-layer fan-shaped divergent walls having film cooling holes thereon, which are connected to the outer annular impact wall by bolts to form an inner annular divergent wall, and an annular impingement cavity is formed between the inner annular divergent wall and the outer annular impact wall; there are floating gaps between the two side edges of each adjacent inner-layer fan-shaped divergent wall; each floating gap is located between each pair of return channels; a trapezoidal flow guiding protrusion is provided on the inner side of the inner annular divergent wall within the span range of the inlet ends and outlet ends of each return channel; and a plurality of groups of support ribs are supported in each return channel.

Description

Technical Field

[0001] This application belongs to the technical field of the design of the barrel wall of an aero-engine combustion liner, and particularly relates to a barrel wall of an aero-engine impingement-diffusion cooling combustion liner. Background Art

[0002] In order to pursue higher thrust, it is necessary to increase the temperature before the turbine of the aero-engine, which causes a significant increase in the temperature inside the combustion liner in the combustion chamber. The barrel wall of the combustion liner is subjected to a high temperature load and is prone to ablation. Therefore, an impingement-diffusion cooling technology is designed for the barrel wall of the combustion liner to efficiently cool the barrel wall of the combustion liner.

[0003] Currently, for the barrel wall of the combustion liner using the impingement-diffusion cooling technology, the design includes an outer annular impingement wall 1 and a plurality of inner sector diffusion walls 2 arranged inside the outer annular impingement wall 1. Among them, the outer annular impingement wall 1 is provided with a plurality of impingement cooling holes; each inner sector diffusion wall 2 is provided with film cooling holes and is connected to the outer annular impingement wall 1 by bolts to form an inner annular diffusion wall, and an annular impingement cavity is formed between the inner annular diffusion wall and the outer annular impingement wall 1. As Figure 1 shown, the cooling air can enter the annular impingement cavity through each impingement cooling hole for impingement cooling, and then is discharged through each film cooling hole to form a film on the inner side of the inner annular diffusion wall for film cooling, so as to achieve efficient cooling of the barrel wall of the combustion liner and protect the barrel wall of the combustion liner from ablation.

[0004] For the currently used barrel wall of the combustion liner with the impingement-diffusion cooling technology, in order to facilitate assembly and compensate for the thermal deformation amount, a floating connection is adopted between the two side edges of each adjacent inner sector diffusion wall 2 inside the outer annular impingement wall 1, and there is a floating gap. As Figure 2 shown, the cooling air entering the annular impingement cavity can leak through the floating gaps everywhere, resulting in a reduction in the utilization rate of the cooling air. In the case where the temperature conditions inside the combustion liner are becoming more and more severe and the available cooling air volume is continuously decreasing, it is difficult to ensure the cooling effect on the barrel wall of the combustion liner.

[0005] In view of the existence of the above technical defects, this application is proposed.

[0006] It should be noted that the disclosure of the above background art 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. Without clear evidence indicating that the above content was publicly available on the filing date of this application, the above background art should not be used to evaluate the novelty and inventiveness of this application. Summary of the Invention

[0007] The purpose of this application is to provide a barrel wall of an aero-engine impingement-diffusion cooling combustion liner to overcome or mitigate at least one aspect of the known technical defects.

[0008] The technical solution of the present application is as follows:

[0009] An impact-divergent cooling flame tube wall of an aeroengine, comprising:

[0010] An outer annular impact wall, which has a plurality of impact cooling holes thereon, has a plurality of pairs of return channels inside, and has a plurality of arc-shaped flow-blocking protrusions on the inner side; the inlet ends and outlet ends of each pair of return channels extend to the inner side of the outer annular impact wall, and the outlet ends are relatively inward and turn back inward; each arc-shaped flow-blocking protrusion is located between each pair of return channels;

[0011] A plurality of inner-layer fan-shaped divergent walls, which have film cooling holes thereon, are connected to the outer annular impact wall by bolts to form an inner annular divergent wall, and form an annular impact cavity with the outer annular impact wall; there is a floating gap between the two side edges of each adjacent inner-layer fan-shaped divergent wall; each floating gap is located between each pair of return channels and is opposite to the position of each arc-shaped flow-blocking protrusion; the inner side of the inner annular divergent wall is within the span range of the inlet ends and outlet ends of each return channel and has a trapezoidal flow guiding protrusion;

[0012] A plurality of groups of support ribs, which are supported in each return channel.

[0013] According to at least one embodiment of the present application, in the above-mentioned impact-divergent cooling flame tube wall of the aeroengine, the two side edges of each adjacent inner-layer fan-shaped divergent wall are in a stepped fit to form a floating gap.

[0014] According to at least one embodiment of the present application, in the above-mentioned impact-divergent cooling flame tube wall of the aeroengine, the inward turning angle of the outlet end of each return channel is greater than 45°.

[0015] According to at least one embodiment of the present application, in the above-mentioned impact-divergent cooling flame tube wall of the aeroengine, the height of each return channel is not greater than 0.5 mm, and the maximum thickness from the inner side of the outer annular impact wall is not greater than 1 mm.

[0016] According to at least one embodiment of the present application, in the above-mentioned impact-divergent cooling flame tube wall of the aeroengine, the height of each trapezoidal flow guiding protrusion is not greater than 1 mm, and the distance from the outer annular impact wall is not greater than 0.5 mm.

[0017] According to at least one embodiment of the present application, in the above-mentioned impact-divergent cooling flame tube wall of the aeroengine, each trapezoidal flow guiding protrusion has a plurality of return channels, which are supported by a plurality of support ribs inside, and have the same configuration as the return channels on the outer annular impact wall and are arranged in parallel. Description of the Drawings

[0018] Figure 1 is a schematic diagram of the current flame tube wall using the impact-divergent cooling technology;

[0019] Figure 2 It is a schematic diagram of the floating gap between adjacent inner-layer sector-shaped divergent walls of the flame tube wall currently adopting the impingement divergent cooling technology;

[0020] Figure 3 It is a schematic diagram of an embodiment of the flame tube wall of an aero-engine impingement divergent cooling provided by an embodiment of the present application;

[0021] Figure 4 It is a schematic diagram of another embodiment of the flame tube wall of an aero-engine impingement divergent cooling provided by an embodiment of the present application;

[0022] Wherein:

[0023] 1 - Outer-layer annular impingement wall; 2 - Inner-layer sector-shaped divergent wall; 3 - Support rib.

[0024] To better illustrate this embodiment, some components in the drawings are omitted, enlarged or reduced, which do not represent the dimensions of the actual product. In addition, the drawings are only for illustrative purposes and cannot be construed as a limitation of this patent. Detailed implementation manners

[0025] To make the technical solutions and their advantages of the present application clearer, the technical solutions of the present application will be further described clearly and completely in conjunction with the drawings. It can be understood that the specific embodiments described herein are only part of the embodiments of the present application, which are only used to explain the present application rather than 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 drawings, and other related parts can refer to the general design. Without conflict, the embodiments and technical features in the embodiments of the present application can be combined with each other to obtain new embodiments.

[0026] In addition, unless otherwise defined, the technical terms 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 pertains. The words indicating directions such as "upper", "lower", "left", "right", "center", "vertical", "horizontal", "inner", "outer", etc. used in the description of this application are only used to indicate relative directions or positional relationships, rather than implying that the device or component must have a specific orientation, be constructed and operated in a specific orientation. When the absolute position of the object being described changes, its relative positional relationship may also change accordingly. Therefore, it should not be construed as a limitation to this application. The terms "first", "second", "third", and similar terms used in the description of this application are only for descriptive purposes to distinguish different components and should not be construed as indicating or implying relative importance. The similar terms such as "a", "an", or "the" used in the description of this application should not be construed as an absolute limitation on the quantity but should be understood as meaning at least one. The terms such as "including" or "comprising" used in the description of this application are intended to mean that the elements or objects appearing before this word cover the elements or objects listed after this word and their equivalents, without excluding other elements or objects.

[0027] In addition, it should be noted that, unless otherwise clearly specified and limited, the similar terms such as "installed", "connected", "joined", etc. 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 directly connected or indirectly connected through an intermediate medium, and can also be the communication inside two components. Those skilled in the art can understand its specific meaning in this application according to the specific situation.

[0028] The following will further elaborate on this application in conjunction with the attached Figures 1 to 4 drawings.

[0029] An impact-divergent cooling flame tube wall of an aeroengine, comprising:

[0030] An outer annular impact wall 1, which has a plurality of impact cooling holes thereon, has a plurality of pairs of return channels inside, and has a plurality of arc-shaped flow-blocking protrusions on the inner side; the inlet ends and outlet ends of each pair of return channels extend to the inner side of the outer annular impact wall 1, and the outlet ends are relatively inward and turn back inward; each arc-shaped flow-blocking protrusion is located between each pair of return channels;

[0031] Multiple inner-layer fan-shaped divergent walls 2, which are provided with film cooling holes and are connected to the outer-layer annular impact wall 1 by bolts, form an inner-layer annular divergent wall, and an annular impact cavity is formed between the inner-layer annular divergent wall and the outer-layer annular impact wall 1; there are floating gaps between the two side edges of adjacent inner-layer fan-shaped divergent walls 2; each floating gap is located between each pair of return channels and is opposite to the position of each arc-shaped flow-blocking protrusion; within the span range of the inlet end and the outlet end of each return channel on the inner side of the inner-layer annular divergent wall, there are trapezoidal guiding protrusions.

[0032] Multiple groups of support ribs 3 are supported within each return channel.

[0033] For the wall of the aero-engine impingement-divergent cooling flame tube disclosed in the above embodiments, those skilled in the art can understand that the cooling air can enter the annular impact cavity through each impingement cooling hole for impingement cooling, and then be discharged through each film cooling hole to form a film on the inner side of the inner-layer annular divergent wall for film cooling. Part of the cooling air entering the annular impact cavity can flow into each return channel and flow back and forth in the return channel, so as to prevent the cooling air from flowing towards the floating gap, reduce the possibility of leakage from the floating gap, and ensure the cooling effect on the wall of the flame tube.

[0034] For the wall of the aero-engine impingement-divergent cooling flame tube disclosed in the above embodiments, those skilled in the art can understand that the design of having arc-shaped flow-blocking protrusions between each pair of return channels on the outer-layer annular impact wall 1 and having trapezoidal guiding protrusions within the span range of the inlet end and the outlet end of each return channel on the inner side of the inner-layer annular divergent wall can bend the flow path of the cooling air in the annular impact cavity towards the floating gap, and can narrow the flow path and increase the flow resistance, so as to further reduce the possibility of leakage of the cooling air in the annular impact cavity from the floating gap and ensure the cooling effect on the wall of the flame tube. In addition, the trapezoidal guiding protrusions can guide part of the air flow in the annular impact cavity into the return channels, ensuring that there is enough cooling air flowing out through the return channels and preventing the cooling air from flowing towards the floating gap.

[0035] In some alternative embodiments, in the above-mentioned wall of the aero-engine impingement-divergent cooling flame tube, the two side edges of adjacent inner-layer fan-shaped divergent walls 2 are in a stepped fit to form a floating gap.

[0036] In some alternative embodiments, in the above-mentioned wall of the aero-engine impingement-divergent cooling flame tube, the inward folding angle of the outlet end of each return channel is greater than 45°, ensuring that the air flow flowing back and forth from the return channel can effectively prevent the cooling air from flowing towards the floating gap, reduce the possibility of leakage from the floating gap, and ensure the cooling effect on the wall of the flame tube.

[0037] In some alternative embodiments, in the above-mentioned barrel wall of the impact-divergent cooling combustor of an aeroengine, the height of each reflux channel is not greater than 0.5 mm, and the maximum thickness from the inner side of the outer annular impact wall 1 is not greater than 1 mm.

[0038] In some alternative embodiments, in the above-mentioned barrel wall of the impact-divergent cooling combustor of an aeroengine, the height of each trapezoidal flow guiding protrusion is not greater than 1 mm, and the distance from the outer annular impact wall 1 is not greater than 0.5 mm.

[0039] In some alternative embodiments, in the above-mentioned barrel wall of the impact-divergent cooling combustor of an aeroengine, each trapezoidal flow guiding protrusion is provided with a plurality of reflux channels, which are supported by a plurality of support ribs 3 inside, and have the same configuration as the reflux channels on the outer annular impact wall 1, and are arranged in parallel, as Figure 4 shown, so as to enhance the resistance to prevent the cooling air flow from flowing into the floating gap, effectively reduce the possibility of leakage from the floating gap, and ensure the cooling effect on the barrel wall of the combustor.

[0040] The various embodiments in the specification are described in a progressive manner. The key point of each embodiment is to illustrate the differences from other embodiments. For the same or similar parts among the various embodiments, reference can be made to each other.

[0041] So far, the technical solutions of the present application have been described in combination with the preferred embodiments shown in the drawings. Those skilled in the art should understand that the protection scope of the present application is obviously not limited to these specific embodiments. Without departing from the principle 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 protection scope of the present application.

Claims

1. An impact-divergence cooling flame tube wall of an aeroengine, characterized in that, Comprising: An outer annular impact wall (1) having a plurality of impingement cooling holes thereon, having a plurality of pairs of return channels inside, and having a plurality of arc-shaped flow-blocking protrusions on the inner side; the inlet ends and outlet ends of each pair of return channels extend to the inner side of the outer annular impact wall (1), and the outlet ends are relatively inward and turn back inward; each arc-shaped flow-blocking protrusion is located between each pair of return channels; A plurality of inner-layer fan-shaped divergence walls (2) having film cooling holes thereon, connected to the outer annular impact wall (1) by bolts to form an inner annular divergence wall, and forming an annular impingement cavity with the outer annular impact wall (1); there is a floating gap between the two side edges of each adjacent inner-layer fan-shaped divergence wall (2); each floating gap is located between each pair of return channels and is opposite to the position of each arc-shaped flow-blocking protrusion; the inner side of the inner annular divergence wall is within the span range of the inlet ends and outlet ends of each return channel and has a trapezoidal flow-guiding protrusion; A plurality of groups of support ribs (3) supported inside each return channel.

2. The wall of the impingement-divergence cooling combustion chamber of an aero-engine according to claim 1, wherein The two side edges of each adjacent inner-layer fan-shaped divergence wall (2) are in a stepped fit to form a floating gap.

3. The wall of the impingement-divergence cooling combustion chamber of an aero-engine according to claim 1, wherein The inward turning angle of the outlet end of each return channel is greater than 45°.

4. The wall of the impingement-divergence cooling combustion chamber of an aero-engine according to claim 1, wherein The height of each return channel is not greater than 0.5 mm, and the maximum thickness from the inner side of the outer annular impact wall (1) is not greater than 1 mm.

5. The wall of the impingement-divergence cooling combustion chamber of an aero-engine according to claim 1, wherein The height of each trapezoidal flow-guiding protrusion is not greater than 1 mm, and the distance from the outer annular impact wall (1) is not greater than 0.5 mm.

6. The wall of the impingement-divergence cooling combustion chamber of an aero-engine according to claim 1, wherein Each trapezoidal flow-guiding protrusion has a plurality of return channels, which are supported by a plurality of support ribs (3) inside, and have the same configuration as the return channels on the outer annular impact wall (1) and are arranged side by side.

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