Aero-engine combustion chamber fuel cooling structure

Abstract

The application belongs to the technical field of aero-engine combustion chamber design, and particularly relates to a fuel cooling structure of an aero-engine combustion chamber, which comprises a combustion chamber outer wall, a combustion chamber inner wall, a flame tube, a fuel passage, an internal partition plate, and a cavity partition plate.

Description

Technical Field

[0001] This application belongs to the field of aero-engine combustor design technology, specifically relating to a fuel cooling structure for an aero-engine combustor. Background Technology

[0002] Aero engines use hydrocarbons as fuel, which easily produces a large amount of greenhouse gases, causing serious environmental impact. For this reason, aero engines are designed to use hydrogen as fuel. However, the high temperature generated after hydrogen combustion increases the emission of pollutants such as nitrogen oxides. Currently, in order to reduce the emission of pollutants from hydrogen-fueled aero engines, most designs adopt multi-point injection combustion chambers to have a more uniform combustion temperature field, thereby reducing pollutant emissions.

[0003] In existing multi-point injection combustion chambers, most of the compressed air from the compressor enters the flame tube directly through the air inlet at the head of the flame tube to participate in combustion. This reduces the amount of air used to cool the flame tube, making the sidewalls of the flame tube prone to ablation. If additional air is drawn from the compressor to cool the flame tube, it will lead to a reduction in compressor efficiency.

[0004] This application is made in view of the aforementioned technical deficiencies.

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

[0006] The purpose of this application is to provide a fuel cooling structure for an air engine combustion chamber to overcome or mitigate at least one of the known technical defects.

[0007] The technical solution of this application is:

[0008] A fuel cooling structure for an aircraft engine combustion chamber includes:

[0009] The combustion chamber has an outer wall fuel inlet.

[0010] The inner wall of the combustion chamber is installed inside the outer wall of the combustion chamber;

[0011] The flame tube is set between the outer wall of the combustion chamber and the inner wall of the combustion chamber. It has multiple air inlets formed at its head and fuel flow holes formed at the upper end of its outer wall. Its side walls are hollow. Each air inlet has multiple fuel injection holes distributed circumferentially on its side wall.

[0012] The fuel passage connects the outer wall of the combustion chamber and the outer wall of the flame tube, and connects the fuel inlet and fuel flow hole on the outer wall.

[0013] An internal baffle is installed inside the side wall of the flame tube, dividing the inside of the flame tube into an outer cavity and an inner cavity; the inner cavity and the outer cavity are connected at the ends of the outer and inner walls of the flame tube; the inner cavity is connected to each fuel injection hole;

[0014] A partition plate is installed inside the inner cavity, located at the head of the flame tube, dividing the inner cavity into two parts circumferentially.

[0015] According to at least one embodiment of this application, in the above-described fuel cooling structure for an aero-engine combustion chamber,

[0016] Each air intake is divided into multiple rows radially and multiple columns circumferentially;

[0017] Multiple crisscrossing circumferential and radial flow channels are formed between the internal baffle and the inner wall of the flame tube head.

[0018] Each circumferential airflow channel and each row of air intake holes are arranged alternately in the radial direction;

[0019] Each radial guide channel and each row of air inlets are distributed alternately in the circumferential direction and are connected to the fuel injection holes on the side wall of the air inlets.

[0020] According to at least one embodiment of this application, the above-described fuel cooling structure for an aircraft engine combustion chamber further includes:

[0021] Multiple flow guide baffles are arranged circumferentially in the inner cavity and intersect with the cavity partition baffle at the head of the flame tube, thus cutting off each circumferential flow guide channel radially.

[0022] According to at least one embodiment of this application, the above-described fuel cooling structure for an aircraft engine combustion chamber further includes:

[0023] The diffuser is connected at the inlet between the outer wall of the combustion chamber and the inner wall of the combustion chamber. Attached Figure Description

[0024] Figure 1 This is a schematic diagram of the fuel cooling structure of the aircraft engine combustion chamber provided in the embodiments of this application;

[0025] Figure 2 yes Figure 1 A partial sectional view;

[0026] Figure 3 yes Figure 2 A partial sectional view;

[0027] in:

[0028] 1-Outer wall of combustion chamber; 2-Inner wall of combustion chamber; 3-Flame tube; 4-Fuel passage; 5-Internal baffle; 6-Cavity baffle; 7-Flow guide baffle; 8-Diffuser; 9-Compressor; 10-Turbine;

[0029] A-Air intake port;

[0030] B - Fuel injection port.

[0031] To better illustrate this embodiment, some parts in the accompanying drawings may be omitted, enlarged, or reduced, and do not represent the actual size of the product. Furthermore, the accompanying drawings are for illustrative purposes only and should not be construed as limiting this patent. Detailed Implementation

[0032] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings. Other related parts can be referred to the general design. In the absence of conflict, the embodiments and technical features in the embodiments of this application can be combined with each other to obtain new embodiments.

[0033] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The terms "upper," "lower," "left," "right," "center," "vertical," "horizontal," "inner," and "outer," etc., used in this application description to indicate relative direction or positional relationship are used only to indicate relative orientation or positional relationship, and do not imply that the device or component must have a specific orientation, or be constructed and operated in a specific orientation. When the absolute position of the described object changes, its relative positional relationship may also change accordingly, and therefore should not be construed as a limitation on this application. The terms "first," "second," "third," and similar terms used in this application description are used only for descriptive purposes to distinguish different components, and should not be construed as indicating or implying relative importance. The terms "a," "one," or "the," etc., used in this application description should not be construed as an absolute limitation on quantity, but should be construed as indicating the existence of at least one. The terms "including," "comprising," etc., used in this application description mean that the element or object preceding the word covers the element or object listed after the word and its equivalents, without excluding other elements or objects.

[0034] Furthermore, it should be noted that, unless otherwise explicitly specified and limited, terms such as “installation,” “connection,” and “linkage” used in the description of this application should be interpreted broadly. For example, a 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 or an indirect connection through an intermediate medium; or it can be a connection within two components. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0035] The following is in conjunction with the appendix Figures 1 to 3 This application will be described in further detail.

[0036] A fuel cooling structure for an aircraft engine combustion chamber includes:

[0037] The combustion chamber outer wall 1 has an outer wall fuel inlet;

[0038] The inner wall 2 of the combustion chamber is installed inside the outer wall 1 of the combustion chamber;

[0039] The flame tube 3 is located between the outer wall 1 and the inner wall 2 of the combustion chamber, at the front end of the turbine 10. It has multiple air inlet holes A formed at its head and fuel flow holes formed at the upper end of its outer wall. Its side wall is hollow inside. Each air inlet hole A has multiple fuel injection holes B distributed along its circumference on its side wall.

[0040] Fuel passage 4 is connected between the outer wall of the combustion chamber 1 and the outer wall of the flame tube 3, and connects to the fuel inlet and fuel flow hole on the outer wall.

[0041] An internal partition 5 is provided inside the side wall of the flame tube 3, dividing the interior of the flame tube 3 into an outer cavity and an inner cavity; the inner cavity and the outer cavity are connected at the ends of the outer and inner walls of the flame tube 3; the inner cavity is connected to each fuel injection hole B;

[0042] The partition plate 6 is installed in the inner cavity and located at the head of the flame tube 3, dividing the inner cavity into two parts circumferentially.

[0043] The diffuser 8 is connected at the inlet between the outer wall 1 and the inner wall 2 of the combustion chamber, and is located at the rear end of the compressor 9.

[0044] In the above-described embodiment of the fuel cooling structure for the combustion chamber of an aero-engine, during operation, high-pressure air from the compressor 9 flows through the diffuser 8 into the inlet of the outer wall 1 and the inner wall 2 of the combustion chamber. Part of the airflow flows between the outer wall 1 and the outer wall of the flame tube 3, and part of the airflow flows between the inner wall 2 and the inner wall of the flame tube 3. These two parts of the airflow flow backward in the circumferential direction, cooling the inner and outer walls of the flame tube 3 along the way, and then are discharged into the turbine 10. The remaining high-pressure air flows into the interior of the flame tube 3 through the air inlet A at the head of the flame tube 3. This part of the airflow is relatively large.

[0045] The fuel cooling structure for the combustion chamber of the aero-engine disclosed in the above embodiments allows fuel to enter the interior of the side wall of the flame tube 3 via the fuel channel 4 during operation. At the end of the outer wall of the flame tube 3, the fuel splits into two paths. One path flows axially forward in the inner cavity, cooling the outer wall of the flame tube 3 along the way. It then turns and enters the exterior of the cavity partition 6 at the head of the flame tube 3, entering the corresponding air inlet A through the fuel injection port B. There, it mixes with the compressed air in the air inlet A and is injected into the flame tube 3 for combustion. The other path flows axially forward in the outer cavity. The fuel flows forward, then turns and flows past the head of the flame tube 3, then turns again and flows axially backward. At the end of the inner wall of the flame tube 3, it turns and flows into the inner cavity, where it flows axially forward. Along the way, it cools the inner wall of the flame tube 3. Afterward, it turns and enters the cavity partition 6 of the head section of the flame tube 3, and enters the corresponding air inlet A through the fuel injection port B. It mixes with the compressed air in the air inlet A and is injected into the flame tube 3 for combustion. The combustion produces high-temperature gas, which is discharged from the tail of the flame tube 3 and drives the turbine 10 to do work. The fuel mentioned above can be hydrogen.

[0046] Regarding the fuel cooling structure of the aero-engine combustion chamber disclosed in the above embodiments, those skilled in the art will understand that its combustion chamber design adopts a multi-point injection combustion chamber, which can have a more uniform combustion temperature field and reduce pollutant emissions. The air inlet A and fuel injection port B are integrated in the head of the flame tube 3, which has a simple structure and is not prone to leakage. Furthermore, the heat sink of the fuel is used to cool the inner and outer walls of the flame tube 3, which can prevent the sidewalls of the flame tube 3 from being ablated and can fully preheat the fuel, making the combustion more uniform and complete.

[0047] In some alternative embodiments, in the above-described fuel cooling structure for the aero-engine combustion chamber,

[0048] Each air intake A is divided into multiple rows radially and multiple columns circumferentially;

[0049] Multiple crisscrossing circumferential and radial flow channels are formed between the internal partition 5 and the inner sidewall of the flame tube 3 head.

[0050] Each circumferential airflow channel and each row of air inlet holes A are arranged alternately in the radial direction;

[0051] Each radial guide channel and each row of air inlet holes A are distributed alternately in the circumferential direction and are connected to the fuel injection holes B on the side wall of the air inlet holes A.

[0052] The fuel cooling structure of the aero-engine combustion chamber disclosed in the above embodiments can utilize multiple crisscrossing circumferential and radial flow channels to divert fuel, ensuring uniform fuel flow throughout and thus guaranteeing the uniformity of the combustion temperature field.

[0053] In some optional embodiments, the above-described fuel cooling structure for the aero-engine combustion chamber further includes:

[0054] Multiple flow guide baffles 7 are arranged circumferentially in the inner cavity and intersect with the cavity partition baffles 6 at the head of the flame tube 3, which cut off each circumferential flow guide channel radially so as to guide the fuel flow and reduce loss.

[0055] The various embodiments in the specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. The same or similar parts between the various embodiments can be referred to each other.

[0056] The technical solution of this 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 this application is obviously not limited to these specific embodiments. Without departing from the principles of this 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 all fall within the scope of protection of this application.

Claims

1. A fuel cooling structure for an aircraft engine combustion chamber, characterized in that, include: The combustion chamber exterior wall (1) has an outer wall fuel inlet; The combustion chamber wall (2) is installed inside the combustion chamber wall (1); The flame tube (3) is set between the outer wall (1) of the combustion chamber and the inner wall (2) of the combustion chamber. It has multiple air inlets (A) formed at its head and fuel flow holes formed at the upper end of its outer wall. Its side wall is hollow. Each air inlet (A) has multiple fuel injection holes (B) distributed along its circumference on its side wall. Fuel passage (4) is connected between the outer wall of the combustion chamber (1) and the outer wall of the flame tube (3), and connects the fuel inlet and fuel flow hole of the outer wall; An internal partition (5) is provided inside the side wall of the flame tube (3) to divide the inside of the flame tube (3) into an outer cavity and an inner cavity; the inner cavity and the outer cavity are connected at the end of the outer wall and the end of the inner wall of the flame tube (3); the inner cavity is connected to each fuel injection hole (B); A partition plate (6) is installed in the inner cavity and located at the head of the flame tube (3), dividing the inner cavity into two parts along the circumference; Each air intake (A) is divided into multiple rows radially and multiple columns circumferentially; Multiple crisscrossing circumferential and radial flow channels are formed between the internal partition (5) and the inner wall of the flame tube (3) head. Each circumferential airflow channel and each row of air inlets (A) are arranged alternately in the radial direction; Each radial guide channel and each row of air inlet holes (A) are distributed alternately in the circumferential direction and are connected to the fuel injection holes (B) on the side wall of the air inlet holes (A).

2. The fuel cooling structure for the combustion chamber of an aero-engine according to claim 1, characterized in that, Also includes: Multiple flow guide baffles (7) are arranged circumferentially in the inner cavity and intersect with the cavity partition baffle (6) at the head of the flame tube (3), cutting off each circumferential flow guide channel radially.

3. The fuel cooling structure for the combustion chamber of an aero-engine according to claim 1, characterized in that, Also includes: The diffuser (8) is connected at the inlet between the outer wall (1) of the combustion chamber and the inner wall (2) of the combustion chamber.

Images