A hydrogen-fueled interstage combustor for an aeroengine

Abstract

The application belongs to the technical field of inter-stage combustion chamber design of an aero-engine, and particularly relates to an aero-engine hydrogen-fueled inter-stage combustion chamber, which comprises a combustion chamber outer wall in the shape of a ring, a combustion chamber inner wall in the shape of a ring arranged in the combustion chamber outer wall, a plurality of guide vanes supported between the combustion chamber outer wall and the combustion chamber inner wall in a circumferential direction, each guide vane being hollow, the leeward side wall of each guide vane being provided with a row of leeward side injection holes distributed along the height of the guide vane, the windward side wall of each guide vane being provided with a row of windward side injection holes distributed along the height of the guide vane, and each guide vane being provided with a row of pilot injection holes distributed along the height of the guide vane and a pilot slot extending along the height of the guide vane, the pilot slot being located behind the windward side injection holes and being connected to the pilot injection holes, a plurality of hydrogen gas injection rods connected to a hydrogen gas source and penetrating through the combustion chamber outer wall and being inserted into each guide vane, and each hydrogen gas injection rod being provided with three rows of hydrogen gas nozzles on the side wall of the hydrogen gas injection rod and being inserted into the leeward side injection holes, the windward side injection holes and the pilot injection holes of the corresponding guide vane.

Description

Technical Field

[0001] This application belongs to the field of interstage combustion chamber design technology for aero-engines, specifically relating to an interstage combustion chamber for hydrogen combustion in aero-engines. Background Technology

[0002] Setting up an interstage combustion chamber between the high-pressure turbine and the low-pressure turbine of an aero-engine can maintain a low fuel consumption rate and increase the thrust of the aero-engine.

[0003] Currently, most interstage combustors in aero engines use fuel-fired interstage combustors with kerosene as fuel and employ a concave cavity vortex combustion scheme. However, due to the low flame propagation speed of kerosene, it is difficult to efficiently organize combustion in the high-speed airflow within the limited length of the interstage combustor, resulting in low combustion efficiency. To address this, current methods often involve increasing the length of the interstage combustor to improve combustion efficiency. However, this increases the overall length and weight of the aero engine, thereby reducing the thrust-enhancing effect of the interstage combustor.

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

[0005] The purpose of this application is to provide an interstage combustion chamber for hydrogen combustion in an aircraft engine to overcome or mitigate at least one of the known technical defects.

[0006] The technical solution of this application is:

[0007] An interstage combustion chamber for hydrogen combustion in an aircraft engine, comprising:

[0008] The outer wall of the combustion chamber is ring-shaped;

[0009] The inner wall of the combustion chamber is annular and is installed inside the outer wall of the combustion chamber;

[0010] Multiple guide vanes are circumferentially supported between the outer wall of the combustion chamber and the inner wall of the combustion chamber. Each guide vane is hollow inside. The leeward side wall has a row of leeward side injection holes distributed along the blade height, the windward side wall has a row of windward side injection holes distributed along the blade height, and an ignition injection hole distributed along the blade height, as well as an ignition groove extending along the blade height. The ignition groove is located behind the windward side injection holes and connects to the ignition injection holes.

[0011] Multiple hydrogen injectors, connected to a hydrogen source, are installed through the outer wall of the combustion chamber and inserted into each guide vane. Each hydrogen injector has three rows of hydrogen nozzles on its side wall. One row of hydrogen nozzles is inserted into the leeward side injection hole on the corresponding guide vane, one row of hydrogen nozzles is inserted into the windward side injection hole on the corresponding guide vane, and one row of hydrogen nozzles is inserted into the ignition injection hole on the corresponding guide vane.

[0012] According to at least one embodiment of this application, in the above-mentioned hydrogen-fueling interstage combustion chamber of the aero-engine, the leeward side injection holes and the windward side injection holes on each guide vane are located near the leading edge, and the ignition groove is adjacent to the windward side injection hole.

[0013] According to at least one embodiment of this application, in the above-mentioned interstage combustion chamber of the aero-engine hydrogen combustion, the cross-section of the ignition groove on each guide vane is V-shaped, and the included angle of the V-shape is 100° to 150°.

[0014] The ignition injection holes on each guide vane are located on the front side wall of the ignition groove.

[0015] According to at least one embodiment of this application, in the above-mentioned hydrogen-fueling interstage combustion chamber of the aero-engine, L>V*(D / v);

[0016] in,

[0017] L is the distance from the ignition slot on each guide vane to the working blade of the low-pressure turbine front stage;

[0018] V is the airflow velocity in the interstage combustion chamber;

[0019] D is the spacing between each guide vane;

[0020] v represents the propagation speed of the hydrogen combustion flame within the interstage combustion chamber.

[0021] According to at least one embodiment of this application, in the above-mentioned hydrogen-fueling interstage combustion chamber of the aero-engine, the outer wall of the combustion chamber and the inner wall of the combustion chamber are provided with multiple bleed air cooling holes, which are connected to the compressor through pipelines to bleed air for cooling each guide vane. Attached Figure Description

[0022] Figure 1 This is a schematic diagram of the hydrogen combustion chamber of an aero-engine provided in an embodiment of this application;

[0023] Figure 2 This is a schematic diagram of the guide vanes and the working blades of the low-pressure turbine front stage in the interstage combustion chamber of an aero-engine provided in an embodiment of this application;

[0024] Figure 3 This is a schematic diagram of a hydrogen injector in the hydrogen combustion chamber of an aero-engine provided in an embodiment of this application;

[0025] in:

[0026] 1-Outer wall of the combustion chamber; 2-Inner wall of the combustion chamber; 3-Guide vane; 4-Hydrogen injector; 5-Working blade of the low-pressure turbine inlet stage;

[0027] L is the distance from the ignition slot on each guide vane to the working blade of the low-pressure turbine front stage;

[0028] D represents the spacing between the guide vanes.

[0029] 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

[0030] 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, and other related parts can be referred to the general design.

[0031] 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 indicating direction used in this application description are used only to indicate relative direction or positional relationship; when the absolute position of the described object changes, its relative positional relationship may also change accordingly. The word "comprising" as used in this application description indicates that the element or object preceding the word encompasses the elements or objects listed following the word and their equivalents, but does not exclude other elements or objects.

[0032] Furthermore, it should be noted that, unless otherwise explicitly specified and limited, terms such as "installation" and "connection" used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection or a detachable connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0033] The following is in conjunction with the appendix Figures 1 to 3 This application will be described in further detail.

[0034] An interstage combustion chamber for hydrogen combustion in an aircraft engine, comprising:

[0035] The outer wall of the combustion chamber is annular;

[0036] The inner wall 2 of the combustion chamber is annular and is installed inside the outer wall 1 of the combustion chamber;

[0037] Multiple guide vanes 3 are circumferentially supported between the outer wall 1 and the inner wall 2 of the combustion chamber; each guide vane 3 is hollow, with a row of leeward side injection holes distributed along the blade height on the leeward side wall, a row of windward side injection holes distributed along the blade height on the windward side wall, and a row of ignition injection holes distributed along the blade height, as well as an ignition groove extending along the blade height, the ignition groove being located behind the windward side injection holes and connected to the ignition injection holes;

[0038] Multiple hydrogen injectors 4 are connected to a hydrogen source, pass through the outer wall 1 of the combustion chamber, and are inserted into each guide vane 3. Each hydrogen injector 4 has three rows of hydrogen nozzles on its side wall. One row of hydrogen nozzles is inserted into the leeward side injection hole on the corresponding guide vane 3, one row of hydrogen nozzles is inserted into the windward side injection hole on the corresponding guide vane 3, and one row of hydrogen nozzles is inserted into the ignition injection hole on the corresponding guide vane 3.

[0039] For the hydrogen-fueling interstage combustion chamber of the aero-engine disclosed in the above embodiments, those skilled in the art will understand that hydrogen can be injected into the flow channel between the outer wall 1 and the inner wall 2 of the combustion chamber through various hydrogen injectors 4 via hydrogen nozzles, leeward side injection holes, windward side injection holes, and ignition injection holes. The airflow in the flow channel between the outer wall 1 and the inner wall 2 of the combustion chamber comes from the high-pressure turbine exhaust and forms a local low-speed backflow zone at the ignition slots of multiple guide vanes 3. Due to the high temperature of the airflow, the hydrogen ejected from the ignition injection holes will spontaneously combust. After the hydrogen spontaneously combusts, the flame will spread in the circumferential direction, realizing the combustion of the entire interstage combustion chamber. It does not require the use of an outer wall cavity stabilizer or an ignition electrode for ignition, which greatly reduces the complexity of the interstage combustion chamber.

[0040] The hydrogen-fueled interstage combustor of the aero-engine disclosed in the above embodiments is designed to use hydrogen as fuel. Hydrogen combustion has a high propagation speed, which is more than 6 times that of kerosene combustion. It does not require atomization and evaporation and can efficiently organize combustion in the high-speed airflow within a limited-length interstage combustor. The combustion efficiency is high, and there is no need to significantly increase the length of the interstage combustor to improve combustion efficiency. This can reduce the overall length and weight of the aero-engine while ensuring the thrust-enhancing effect of the interstage combustor.

[0041] Although hydrogen combustion has a high propagation speed, reaching tens of meters per second, the airflow velocity in the interstage combustion chamber is greater than hundreds of meters per second, which is much greater than the propagation speed of hydrogen combustion. If hydrogen is directly injected into the flow channel between the outer wall 1 and the inner wall 2 of the combustion chamber, the spontaneously combusted hydrogen will be extinguished immediately. In the hydrogen-fueling interstage combustion chamber of the aero-engine disclosed in the above embodiment, each guide vane 3 is designed with an ignition groove extending along the blade height on the windward side wall, and a row of corresponding ignition injection holes is designed. The low-speed recirculation zone formed by the ignition groove can be used to enable the hydrogen ejected from the ignition injection holes to stably spontaneously combust, thereby carrying out flame propagation and flame connection, and enabling efficient organized combustion in the interstage combustion chamber.

[0042] Considering the high propagation speed of hydrogen combustion, in the hydrogen-fueling interstage combustion chamber of the aero-engine disclosed in the above embodiments, ignition grooves are designed only on the windward sidewall of each guide vane 3 to ensure the ignition, flame propagation, flame connection and combustion organization of the interstage combustion chamber. Ignition grooves are not designed on the leeward sidewall, thereby reducing the number of ignition grooves on the guide vane 3, avoiding large aerodynamic losses, and preventing flame from hitting the wall and causing the guide vane 3 to be ablated.

[0043] In some optional embodiments, in the above-mentioned hydrogen-fueling interstage combustion chamber of the aero-engine, the leeward side injection holes and the windward side injection holes on each guide vane 3 are located close to the leading edge, and the ignition groove is adjacent to the windward side injection hole, so as to extend the flame propagation and flame connection length of hydrogen in the flow channel between the outer wall 1 and the inner wall 2 of the combustion chamber as much as possible, thereby reducing the overall length of the interstage combustion chamber and its aero-engine, avoiding a significant increase in the weight of the aero-engine and reducing the thrust enhancement efficiency of the interstage combustion chamber.

[0044] In some optional embodiments, in the above-mentioned interstage combustion chamber of the aero-engine hydrogen combustion, the cross-section of the ignition groove on each guide vane 3 is V-shaped, and the included angle of the V-shape is 100° to 150°.

[0045] The ignition injection holes on each guide vane 3 are located on the front side wall of the ignition slot, which ensures stable combustion of hydrogen injected into the ignition slot and reduces aerodynamic losses in the flow channel between the outer wall 1 and the inner wall 2 of the combustion chamber.

[0046] In some alternative embodiments, in the aforementioned interstage combustion chamber of the aero-engine hydrogen combustion system,

[0047] L>V*(D / v);

[0048] in,

[0049] L is the distance from the ignition groove on each guide vane 3 to the working blade 5 of the low-pressure turbine front stage;

[0050] V is the airflow velocity in the interstage combustion chamber;

[0051] D represents the spacing between the three guide vanes;

[0052] v represents the propagation speed of the hydrogen combustion flame within the interstage combustion chamber.

[0053] Regarding the hydrogen-fueling interstage combustion chamber of the aero-engine disclosed in the above embodiments, those skilled in the art will understand that the distance L from the ignition groove on each guide vane 3 to the working blade 5 of the low-pressure turbine inlet stage is designed to be greater than the length required for the hydrogen injected into the ignition injection hole on each guide vane 3 to propagate between the guide vanes 3 after self-ignition at the ignition groove. This ensures the combustion efficiency in the interstage combustion chamber, allowing for complete combustion when it reaches the working blade 5 of the low-pressure turbine inlet stage. Each guide vane 3 can be integrated with the guide vane located before the working blade 5 of the low-pressure turbine inlet stage, and the entire interstage combustion chamber is located between the working blade 5 of the low-pressure turbine inlet stage and the guide vane.

[0054] In some optional embodiments, the above-mentioned hydrogen-fueling interstage combustion chamber of the aero-engine has multiple bleed air cooling holes on the outer wall 1 and the inner wall 2 of the combustion chamber. Bleed air is drawn through pipelines to the compressor to cool each guide vane 3, and at the same time, it can cool the inner walls of the outer wall 1 and the inner wall 2 of the combustion chamber to supply hydrogen for combustion.

[0055] The various embodiments in the specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. Similar or identical parts between embodiments can be referred to mutually. In the absence of conflict, the embodiments and technical features in the embodiments of this application can be combined to obtain new embodiments.

[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 hydrogen-fueling interstage combustion chamber for an aero-engine, characterized in that, include: The outer wall of the combustion chamber (1) is annular; The inner wall of the combustion chamber (2) is annular and is installed inside the outer wall of the combustion chamber (1); Multiple guide vanes (3) are circumferentially supported between the outer wall (1) and the inner wall (2) of the combustion chamber; each guide vane (3) is hollow, the leeward side wall has a row of leeward side injection holes distributed along the blade height, the windward side wall has a row of windward side injection holes distributed along the blade height, and has a row of ignition injection holes distributed along the blade height, and an ignition groove extending along the blade height, the ignition groove is located behind the windward side injection holes and connects to the ignition injection holes; Multiple hydrogen injectors (4) are connected to a hydrogen source, pass through the outer wall (1) of the combustion chamber, and are inserted into each guide vane (3). Each hydrogen injector (4) has three rows of hydrogen nozzles on its side wall. One row of hydrogen nozzles is inserted into the leeward side injection hole on the corresponding guide vane (3), one row of hydrogen nozzles is inserted into the windward side injection hole on the corresponding guide vane (3), and one row of hydrogen nozzles is inserted into the ignition injection hole on the corresponding guide vane (3).

2. The interstage combustion chamber for hydrogen combustion in an aero-engine according to claim 1, characterized in that, Each guide vane (3) has a leeward side injection hole and a windward side injection hole near the leading edge, and an ignition groove is adjacent to the windward side injection hole.

3. The interstage combustion chamber for hydrogen combustion in an aero-engine according to claim 1, characterized in that, The cross-section of the ignition groove on each guide vane (3) is V-shaped, and the included angle of the V-shape is 100° to 150°. The ignition injection holes on each guide vane (3) are located on the front side wall of the ignition groove.

4. The interstage combustion chamber for hydrogen combustion in an aero-engine according to claim 1, characterized in that, L>V*(D / v); in, L is the distance from the ignition groove on each guide vane (3) to the working blade (5) of the low-pressure turbine front stage; V is the airflow velocity in the interstage combustion chamber; D is the distance between each guide vane (3); v represents the propagation speed of the hydrogen combustion flame within the interstage combustion chamber.

5. The interstage combustion chamber for hydrogen combustion in an aero-engine according to claim 1, characterized in that, The outer wall (1) and inner wall (2) of the combustion chamber have multiple bleed air cooling holes, which are connected to the compressor through pipelines to bleed air to cool each guide vane (3).

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