Aero-engine combustion chamber and oil injection rod thereof
By employing a five-injector design and a complex cooling structure, the problems of insufficient fuel injector cooling and unstable fuel-air ratio in the combustion chamber have been solved, achieving efficient and stable combustion and cooling under ultra-high temperature rise conditions, thereby improving the reliability of the aero-engine and the adaptability of the combustion chamber.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- AECC SHENYANG ENGINE RES INST
- Filing Date
- 2026-03-16
- Publication Date
- 2026-05-26
Smart Images
Figure CN122083373A_ABST
Abstract
Description
Technical Field
[0001] This application belongs to the field of aero-engine combustion chamber design technology, specifically relating to an aero-engine combustion chamber and its fuel injector rod. Background Technology
[0002] Future aircraft will demand higher thrust-to-weight ratios, wider operating envelopes, and longer ranges, requiring aero-engine combustors to possess ultra-high temperature rise, ultra-wide operating range, and high-efficiency, stable combustion capabilities. The ultra-high temperature rise combustor design point requires an air-fuel ratio close to the stoichiometric ratio, resulting in a significantly higher turbine inlet temperature compared to existing models. This necessitates more stringent design requirements for the combustor outlet temperature profile and hot spot temperature limits to prevent turbine erosion. Furthermore, the near-stoichiometric air-fuel ratio at the combustor design point further expands the required stable air-fuel ratio range. While meeting the overall combustion performance at the design point, ensuring proper flameout performance becomes challenging, ultimately impacting the reliable operation of the aero-engine.
[0003] Currently, the typical combustion chamber form of aero engines is as follows: Figure 1 As shown, the airflow from the compressor is divided into three parts by the diffuser. The first part bypasses the injector and enters the space between the outer casing of the combustor and the outer wall of the flame tube, forming an outer ring airflow. The second part enters the space between the inner casing of the combustor and the inner wall of the flame tube, forming an inner ring airflow. The third part enters the flame tube through the flame tube head, forming a head airflow. The outer ring airflow flows into the flame tube through mixing holes on the outer wall of the flame tube, while another part is used to cool the outer wall of the flame tube. The inner ring airflow flows into the flame tube through mixing holes on the inner wall of the flame tube, while another part is used to cool the inner wall of the flame tube. The head airflow, together with the outer and inner ring airflows that enter the flame tube through the mixing holes, forms a circumferentially continuous main combustion zone. The combustion gas in the main combustion zone flows into the high-pressure turbine guide vane.
[0004] Currently, in typical aero-engine combustion chamber fuel injector rods, the fuel injector cooling structure is as follows: Figure 2 As shown, the main cooling methods employed are impact and film cooling. The design point of the ultra-high temperature combustion chamber has an air-fuel ratio close to the stoichiometric ratio, resulting in a significant increase in the thermal load on the fuel injector. Simultaneously, more cooling air needs to be used for flame tube cooling, reducing the amount of cooling air available for the fuel injector. The current simple impact and film cooling methods can no longer meet the long-term reliable operation requirements of the fuel injector.
[0005] In view of the aforementioned technical deficiencies, this application is hereby filed. Summary of the Invention
[0006] The purpose of this application is to provide an aircraft engine combustion chamber and its fuel injector rod to overcome or mitigate at least one of the known technical defects.
[0007] The technical solution of this application is:
[0008] A fuel injector rod for an aircraft engine combustion chamber has five fuel injectors that are inserted into mounting holes in the head of the flame tube. The injectors are arranged radially from the outside to the inside as an outer main fuel injector, an outer auxiliary fuel injector, a main fuel injector, an inner auxiliary fuel injector, and an inner main fuel injector.
[0009] The main fuel injector, outer main fuel injector, and inner main fuel injector on the fuel injector rod protrude from the head of the flame tube, while the outer auxiliary fuel injector and inner auxiliary fuel injector are retracted into the head of the flame tube.
[0010] The main injector, outer main injector, and inner main injector include a main nozzle and a distributor cone. The inlet of the main nozzle is connected to the injector rod, and the distributor cone is installed inside the outlet. A small vortex is installed between the distributor cone and the outlet of the main nozzle, and the distributor cone narrows at the end of the outlet of the main nozzle. A large vortex is installed outside the outlet of the main nozzle.
[0011] The rear end of the outer wall of the large cyclone protrudes backward from the main nozzle and contracts inward. The outer wall of the large cyclone is divided into an outer wall and an inner wall. The front end of the outer wall and the inner wall form an annular air inlet, and the rear end has multiple outer exhaust holes and inner exhaust holes.
[0012] The large hydrocyclone has a sandwich wall inside the outer and inner walls. The sandwich wall has multiple air holes. The rear end is connected to the rear end of the outer wall of the large hydrocyclone, and the front end is connected to the front side of the inward bending part of the outer wall.
[0013] The outer and inner vents connect the space between the outer wall, the inner wall, and the interlayer wall, and the outer wall has multiple diverging holes on the inwardly bent part; the diverging holes have a non-axisymmetric gradually expanding structure, with a backward-bending guide wall on the front side and a backward-expanding expansion wall on the rear side.
[0014] According to at least one embodiment of this application, in the fuel injector rod of the above-mentioned aircraft engine combustion chamber, the outer auxiliary fuel injector and the inner auxiliary fuel injector include an outer nozzle, an inner nozzle, and an inner cone;
[0015] The external nozzle inlet is connected to the fuel injector rod;
[0016] The inner nozzle is installed inside the outer nozzle. The outlets of the outer and inner nozzles are conical, and a small vortex generator is installed between them.
[0017] The inner cone is installed inside the inner nozzle. It is hollow inside, with its cylindrical end connected to the fuel injection rod. The cone end is located inside the outlet of the inner nozzle, and a small vortex is installed between it and the outlet of the inner nozzle. The side wall of the cone end has multiple fuel injection holes distributed circumferentially.
[0018] According to at least one embodiment of this application, in the fuel injector rod of the above-mentioned aircraft engine combustion chamber, the distance by which the main fuel injector, the outer main fuel injector, and the inner main fuel injector protrude from the head of the flame tube is L1, and the distance by which the outer auxiliary fuel injector and the inner auxiliary fuel injector are recessed into the head of the flame tube is L2, where L2 > L1.
[0019] An aircraft engine combustion chamber includes an outer combustion chamber casing, an inner combustion chamber casing, an outer wall of a flame tube, an inner wall of a flame tube, a flame tube head, and a fuel injector rod for the aforementioned aircraft engine combustion chamber.
[0020] The combustion chamber casing is installed inside the combustion chamber casing, forming the combustion chamber wall structure. The inlet of the combustion chamber wall structure is connected to the diffuser outlet, and the diffuser inlet is connected to the compressor outlet.
[0021] The outer and inner walls of the flame tube have mixing holes, which are set within the combustion chamber wall structure. The inner wall of the flame tube is located inside the outer wall of the flame tube, forming the flame tube. The flame tube outlet is directly opposite the high-pressure turbine guide.
[0022] The head of the flame tube is located at the flame tube inlet and has an air inlet.
[0023] There are multiple fuel injection rods, which are circumferentially connected to the outer casing of the combustion chamber.
[0024] The head of the flame tube has a flared mouth at the nozzles of the outer and inner auxiliary fuel injectors, forming a triangular blunt body.
[0025] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber, during operation, the airflow from the compressor flows through the diffuser and is divided into three parts. The first part bypasses the fuel injector and enters the space between the outer casing of the combustion chamber and the outer wall of the flame tube, forming an outer ring airflow. The second part enters the space between the inner casing of the combustion chamber and the inner wall of the flame tube, forming an inner ring airflow. The third part enters the interior of the flame tube through the air inlet on the head of the flame tube, forming a head airflow. The head airflow forms a blunt body return zone after the triangular blunt body.
[0026] The outer ring airflow flows into the flame tube through the mixing hole on the outer wall of the flame tube, and the other part is used to cool the outer wall of the flame tube; the inner ring airflow flows into the flame tube through the mixing hole on the inner wall of the flame tube, and the other part is used to cool the inner wall of the flame tube; the head airflow, together with the outer ring airflow and the inner ring airflow that enter the flame tube through the mixing hole, forms the main return flow zone, and mixes with the fuel injected by the fuel injector on the fuel injector rod, burns to form gas, and flows into the high-pressure turbine guide vane;
[0027] The main reflux region is located after the bluff body reflux region and overlaps with it in space.
[0028] According to at least one embodiment of this application, the head airflow entering the interior of the combustion chamber of the aforementioned aircraft engine through the air inlet on the head of the flame tube is divided into six paths that are radially alternately distributed with the five fuel injectors on the fuel injector rod.
[0029] According to at least one embodiment of this application, the above-mentioned aero-engine combustion chamber has a low operating condition, a medium operating condition, and a high operating condition.
[0030] In low-operation conditions, the external auxiliary fuel injector and the internal auxiliary fuel injector are working, while the main fuel injector, the external main fuel injector, and the internal main fuel injector are not working. This is suitable for low-operation conditions in the combustion chamber. At this time, the fuel injected into the flame tube by the external auxiliary fuel injector and the internal auxiliary fuel injector will enter the bluff body recirculation zone for stable combustion.
[0031] In the medium-condition operation, the main fuel injector, external auxiliary fuel injector, and internal auxiliary fuel injector are working, while the external main fuel injector and internal main fuel injector are not working. This is suitable for the medium-condition operation in the combustion chamber. At this time, the fuel injected into the flame tube by the external auxiliary fuel injector and internal auxiliary fuel injector will enter the bluff body recirculation zone for stable combustion, and the fuel injected into the flame tube by the main fuel injector will enter the main recirculation zone. The flame in the bluff body recirculation zone can ignite the fuel in the main recirculation zone to form the main combustion zone for combustion.
[0032] In high-operation conditions, the main injector, external main injector, internal main injector, external auxiliary injector, and internal auxiliary injector operate, suitable for high combustion conditions in the combustion chamber. The fuel injected into the flame tube by the external auxiliary injector and internal auxiliary injector will enter the bluff body recirculation zone for stable combustion, while the fuel injected into the flame tube by the main injector, external main injector, and internal main injector will enter the main recirculation zone. The flame in the bluff body recirculation zone can ignite the fuel in the main recirculation zone, thus forming three main combustion zones for combustion. Attached Figure Description
[0033] Figure 1 This is a schematic diagram of a typical combustion chamber design for current aero engines;
[0034] Figure 2 This is a schematic diagram of the fuel injector cooling structure in a typical combustion chamber fuel injector rod of an aero-engine.
[0035] Figure 3 This is a schematic diagram of an aircraft engine combustion chamber provided in an embodiment of this application;
[0036] Figure 4 This is a partial schematic diagram of an aero-engine combustion chamber provided in an embodiment of this application;
[0037] Figure 5 This is a schematic diagram of the structure of the main fuel injector, outer main fuel injector, and inner main fuel injector provided in the embodiments of this application;
[0038] Figure 6 yes Figure 5 A partial schematic diagram;
[0039] Figure 7 yes Figure 6 Sectional view along axis AA;
[0040] Figure 8 yes Figure 6 BB-direction sectional view;
[0041] Figure 9 This is a schematic diagram of the structure of the external auxiliary fuel injector and the internal auxiliary fuel injector provided in the embodiments of this application;
[0042] in:
[0043] 1-Outer combustion chamber casing; 2-Inner combustion chamber casing; 3-Outer wall of the flame tube; 4-Inner wall of the flame tube; 5-Flame tube head; 6-Injection rod; 7-Diffuser; 8-Compressor; 9-High-pressure turbine guide vane;
[0044] 61-Main fuel injector; 62-Outer main fuel injector; 63-Inner main fuel injector; 64-Outer auxiliary fuel injector; 65-Inner auxiliary fuel injector;
[0045] 10-Main nozzle; 11-Oil distribution cone; 12-Outer nozzle; 13-Inner nozzle; 14-Inner cone; 15-Outer wall; 16-Inner wall; 17-Interlayer wall.
[0046] To better illustrate this embodiment, some content in the accompanying drawings may be omitted, enlarged, or reduced. They are for illustrative purposes only and should not be construed as limiting the scope of this application. Detailed Implementation
[0047] 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.
[0048] 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 word "comprising" as used in this application description indicates that the concept preceding the word encompasses the concepts listed following the word and their equivalents, without excluding other related concepts.
[0049] Furthermore, the terms indicating location used in the description of this application are only used to indicate relative directions or positional relationships. When the absolute position of the described object changes, its relative positional relationship may also change accordingly. It should also 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.
[0050] An aircraft engine combustion chamber, such as Figure 3 As shown, it includes an outer combustion chamber casing 1, an inner combustion chamber casing 2, an outer wall of the flame tube 3, an inner wall of the flame tube 4, a flame tube head 5, and an injection rod 6.
[0051] The outer combustion chamber 1, the inner combustion chamber 2, the outer wall of the flame tube 3, the inner wall of the flame tube 4, and the head of the flame tube 5 are all rotating bodies, and the axis of rotation is the axis of the aero-engine.
[0052] The combustion chamber casing 2 is installed inside the combustion chamber casing 1, forming the combustion chamber wall structure. The inlet of the combustion chamber wall structure is connected to the outlet of the diffuser 7, and the inlet of the diffuser 7 is connected to the outlet of the compressor 8.
[0053] The outer wall 3 and the inner wall 4 of the flame tube have mixing holes, which are set in the combustion chamber wall structure. The inner wall 4 of the flame tube is located inside the outer wall 3 of the flame tube, forming the flame tube. The outlet of the flame tube is directly opposite the high-pressure turbine guide 9.
[0054] The head 5 of the flame tube is located at the inlet of the flame tube and has an air inlet.
[0055] There are multiple fuel injector rods 6, which are circumferentially connected to the outer casing 1 of the combustion chamber. Each rod has five radially distributed fuel injectors that are inserted into the mounting holes in the head of the flame tube 5.
[0056] Of the five nozzles on the fuel injector 6, the one in the middle is the main nozzle 61, the one near the outer wall 3 of the flame tube is the outer main nozzle 62, the one near the inner wall 4 of the flame tube is the inner main nozzle 63, the nozzle between the main nozzle 61 and the outer main nozzle 62 is the outer auxiliary nozzle 64, and the nozzle between the main nozzle 61 and the inner main nozzle 63 is the inner auxiliary nozzle 65.
[0057] The main injector 61, outer main injector 62, and inner main injector 63 on the injector rod 6 are air-assisted atomizing nozzles, protruding from the head of the flame tube 5 by a distance L1. The outer auxiliary injector 64 and inner auxiliary injector 65 are centrifugal atomizing nozzles, retracted into the head of the flame tube 5 by a distance L2, where L2 > L1. Figure 4 As shown.
[0058] The flame tube head 5 has a flared mouth formed at the nozzles of the outer auxiliary fuel injector 64 and the inner auxiliary fuel injector 65, forming a triangular blunt body.
[0059] When the combustion chamber of the aero-engine disclosed in the above embodiment is working, the airflow from the compressor 8 is divided into three parts by the diffuser 7. The first part bypasses the fuel injector 6 and enters the space between the outer casing 1 of the combustion chamber and the outer wall 3 of the flame tube, forming an outer ring airflow. The second part enters the space between the inner casing 2 of the combustion chamber and the inner wall 4 of the flame tube, forming an inner ring airflow. The third part enters the interior of the flame tube through the air inlet on the flame tube head 5, forming a head airflow. The head airflow forms a blunt body return zone after passing through a triangular blunt body. The outer ring airflow flows into the interior of the flame tube through the mixing hole on the outer wall 3 of the flame tube, and another part is used to cool the outer wall 3 of the flame tube. The inner ring airflow flows into the interior of the flame tube through the mixing hole on the inner wall 4 of the flame tube, and another part is used to cool the inner wall 4 of the flame tube. The head airflow, the outer ring airflow that enters the interior of the flame tube through the mixing hole, and the inner ring airflow form the main return zone, and mix with the fuel injected by the fuel injector on the fuel injector 6, which is then burned to form combustion gas, which flows into the high-pressure turbine guide vane 9.
[0060] The head airflow that enters the interior of the flame tube through the air inlet on the flame tube head 5 can be designed to be divided into six paths that are radially alternately distributed with the five fuel injectors on the fuel injector rod 6, and are specifically divided by airflow baffles.
[0061] The shape of the combustion chamber outlet temperature profile can be adjusted by regulating the fuel injection ratio of the main injector 61, the outer main injector 62, and the inner main injector 63, so as to approximate the design profile and thus achieve active adjustment of the shape of the combustion chamber outlet temperature profile.
[0062] In the aero-engine combustion chamber disclosed in the above embodiments, the fuel injected into the flame tube by the external auxiliary fuel injector 64 and the internal auxiliary fuel injector 65 burns in the bluff body recirculation zone, playing the role of ignition and flame stabilization. The fuel injected into the flame tube by the main fuel injector 61, the external main fuel injector 62, and the internal main fuel injector 63 can form a radially parallel main combustion zone, which is used to achieve efficient and uniform combustion under the condition that the fuel-air ratio at the combustion chamber design point is close to the chemical appropriate ratio. In this way, the aero-engine combustion chamber can be designed to have low operating conditions, medium operating conditions, and high operating conditions.
[0063] In low-operation conditions, the external auxiliary fuel injector 64 and the internal auxiliary fuel injector 65 are working, while the main fuel injector 61, the external main fuel injector 62, and the internal main fuel injector 63 are not working. This is suitable for low-operation conditions in the combustion chamber. At this time, the fuel injected into the flame tube by the external auxiliary fuel injector 64 and the internal auxiliary fuel injector 65 will enter the bluff body recirculation zone for stable combustion.
[0064] In the medium-condition operation, the main injector 61, the external auxiliary injector 64, and the internal auxiliary injector 65 are working, while the external main injector 62 and the internal main injector 63 are not working. This is suitable for the medium-condition operation in the combustion chamber. At this time, the fuel injected into the flame tube by the external auxiliary injector 64 and the internal auxiliary injector 65 will enter the bluff body recirculation zone for stable combustion, which has a flame stabilizing effect on the combustion of the entire combustion chamber. The fuel injected into the flame tube by the main injector 61 will enter the main recirculation zone. The main recirculation zone and the bluff body recirculation zone overlap in space. The flame in the bluff body recirculation zone can ignite the fuel in the main recirculation zone to form the main combustion zone for combustion.
[0065] In high-operation conditions, the main injector 61, external main injector 62, internal main injector 63, external auxiliary injector 64, and internal auxiliary injector 65 operate, suitable for high combustion conditions in the combustion chamber. The fuel injected into the flame tube by the external auxiliary injector 64 and internal auxiliary injector 65 will enter the bluff body recirculation zone for stable combustion, which has a flame stabilizing effect on the combustion of the entire combustion chamber. The fuel injected into the flame tube by the main injector 61, external main injector 62, and internal main injector 63 will enter the main recirculation zone. The main recirculation zone and the bluff body recirculation zone overlap in space. The flame in the bluff body recirculation zone can ignite the fuel in the main recirculation zone, thus forming three main combustion zones for combustion.
[0066] The main fuel injector 61, outer main fuel injector 62, and inner main fuel injector 63 can be further designed, such as... Figure 5 As shown, it includes a main nozzle 10 and a fuel distribution cone 11. The inlet of the main nozzle 10 is connected to the fuel injection rod 6, which is connected to the fuel passage opened in the fuel injection rod 6 and connected to the fuel source of the aircraft engine through the fuel line. The fuel distribution cone 11 is set in the outlet. A small vortex is set between the fuel distribution cone 11 and the outlet of the main nozzle 10, and the fuel distribution cone 11 and the outlet end of the main nozzle 10 are narrowed. A large vortex is set outside the outlet of the main nozzle 10.
[0067] When the aircraft engine is working, fuel can be introduced into the main nozzle 10 through the fuel passage, including the main injector 61, the outer main injector 62, and the inner main injector 63. The fuel distribution cone 11 turns the fuel entering the main nozzle 10 into a ring-shaped flow, which is then swirled by a small vortex generator and ejected through the outlet of the main nozzle 10. Part of the head airflow can be turned into swirling air by a large vortex generator, which breaks up and atomizes the fuel ejected from the outlet of the main nozzle 10, so as to achieve full mixing and efficient combustion of fuel and air under high operating conditions.
[0068] To prevent the main injector 61, outer main injector 62, and inner main injector 63 from high-temperature erosion, the rear end of the outer wall of the large vortex is designed to protrude backward from the main nozzle 10 and contract inward. The outer wall of the large vortex is divided into an outer wall 15 and an inner wall 16. The front ends of the outer wall 15 and the inner wall 16 form an annular air inlet, and the rear ends have multiple outer exhaust holes and inner exhaust holes.
[0069] A sandwich wall 17 is provided within the outer wall 15 and inner wall 16 of the large hydrocyclone. The sandwich wall 17 has multiple vent holes, its rear end connected to the rear end of the outer wall of the large hydrocyclone, and its front end connected to the front side of the inwardly bent portion of the outer wall 15. Figure 6 As shown.
[0070] The outer vent connects to the space between the outer wall 15 and the interlayer wall 17, the inner vent connects to the space between the inner wall 16 and the interlayer wall 17, and the outer wall 15 has multiple vent holes on the inwardly bent part.
[0071] When the aero-engine is working, part of the head airflow can enter between the outer wall 15 and the inner wall 16 through the annular air inlet on the outer wall of the large vortex generator. Part of the airflow is discharged directly through the inner exhaust port and enters the interior of the flame tube. The other part of the airflow enters the space between the outer wall 15 and the interlayer wall 17 through the vent, and performs efficient impact cooling on the outer wall 15. Most of this part of the airflow is discharged through the divergence hole, and a small part is discharged through the outer exhaust port and enters the interior of the flame tube.
[0072] The diverging aperture has a non-axisymmetric, gradually expanding structure, such as... Figure 7 As shown, the front side is a backward-curving guide wall, and the rear side is a backward-expanding expansion wall. While gradually expanding, the direction of the exhaust airflow is changed to be roughly consistent with the direction of the local airflow field inside the flame tube. At the same time, it can reduce the flow velocity of the exhaust airflow, improve the adhesion effect between the exhaust airflow and the outer wall 15, and thus improve the cooling efficiency.
[0073] The outer and inner vent holes are axisymmetric circular holes with uniform cross-sections, such as... Figure 8 As shown, by maintaining a certain angle with the rear end of the outer wall of the large cyclone separator, the direction of the exhaust airflow is changed to be roughly consistent with the direction of the local airflow field inside the flame tube, thereby improving the cooling efficiency.
[0074] External auxiliary fuel injector 64, internal auxiliary fuel injector 65, etc. Figure 9 As shown, it includes an outer nozzle 12, an inner nozzle 13, and an inner cone 14. The inlet of the outer nozzle 12 is connected to the fuel injector 6, and it connects to the fuel passage opened inside the fuel injector 6, and is connected to the fuel source of the aircraft engine through the fuel line. The inner nozzle 13 is set inside the outer nozzle 12. The outlets of the outer nozzle 12 and the inner nozzle 13 are conical, and a small vortex is set between them. The inner cone 14 is set inside the inner nozzle 13. It is hollow inside, with its cylindrical end connected to the fuel injector 6. The conical end is located inside the outlet of the inner nozzle 13, and a small vortex is set between it and the outlet of the inner nozzle 13. The side wall of the conical end has multiple fuel injection holes distributed circumferentially.
[0075] When the aero-engine is working, fuel can be introduced into the outer nozzle 12 and inner cone 14 of the outer auxiliary fuel injector 64 and the inner auxiliary fuel injector 65 through the fuel passage. The fuel introduced into the outer auxiliary fuel injector 64 is sprayed out after being swirled by a small vortex generator. The fuel introduced into the inner cone 14 enters the inner auxiliary fuel injector 65 through the fuel injection hole, and is then sprayed out after being swirled by a small vortex generator. The fuel sprayed from the outer auxiliary fuel injector 64 and the inner auxiliary fuel injector 65 is centrifugally atomized, which facilitates stable combustion during ignition and low operating conditions, and plays a role in flame stabilization.
[0076] The aero-engine combustor disclosed in the above embodiments adopts an integrated multi-point injection nozzle design. It can control the flow rate of multiple fuel nozzles of a single nozzle with multiple degrees of freedom. Combined with the bluff body flame stabilization combustion organization mode, the combustion chamber outlet temperature profile shape, hot spot temperature and other turbine-given design requirements can be achieved through the zoned fuel control function of the nozzle.
[0077] The aero-engine combustion chamber disclosed in the above embodiments uses plate cooling in the fuel injector cooling structure. It utilizes a combination of impact and divergent cooling to improve the overall cooling efficiency. The divergent holes are designed with a non-axisymmetric gradually expanding structure to enhance the wall adhesion effect of the cooling airflow, thereby reducing the amount of cooling air used in the fuel injector and meeting the requirements for long-term reliable operation of the fuel injector structure.
[0078] 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 protection scope 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 protection scope of this application.
Claims
1. A fuel injector rod for an aircraft engine combustion chamber, characterized in that, The fuel injector (6) has five fuel injectors, which are inserted into the mounting holes opened in the flame tube head (5). The injectors are arranged radially from the outside to the inside as follows: outer main fuel injector (62), outer auxiliary fuel injector (64), main fuel injector (61), inner auxiliary fuel injector (65), and inner main fuel injector (63). The main injector (61), outer main injector (62), and inner main injector (63) on the injector rod (6) protrude from the head of the flame tube (5), while the outer auxiliary injector (64) and inner auxiliary injector (65) are recessed inside the head of the flame tube (5). The main injector (61), the outer main injector (62), and the inner main injector (63) include a main nozzle (10) and a fuel distribution cone (11). The inlet of the main nozzle (10) is connected to the injector rod (6), and the fuel distribution cone (11) is installed inside the outlet. A small vortex is installed between the fuel distribution cone (11) and the outlet of the main nozzle (10), and the fuel distribution cone (11) and the outlet end of the main nozzle (10) are narrowed. A large vortex is installed outside the outlet of the main nozzle (10). The rear end of the outer wall of the large cyclone protrudes backward from the main nozzle (10) and contracts inward. The outer wall of the large cyclone is divided into an outer wall (15) and an inner wall (16). The front ends of the outer wall (15) and the inner wall (16) form an annular air inlet, and the rear ends have multiple outer exhaust holes and inner exhaust holes. A sandwich wall (17) is provided inside the outer wall (15) and inner wall (16) of the large hydrocyclone. The sandwich wall (17) has multiple ventilation holes. The rear end is connected to the rear end of the outer wall of the large hydrocyclone, and the front end is connected to the front side of the inward bending part of the outer wall (15). The outer vent and inner vent connect the space between the outer wall (15), the inner wall (16) and the interlayer wall (17), and the outer wall (15) has multiple diverging holes on the inward bending part; the diverging holes are non-axisymmetric gradually expanding structures, with a backward bending guide wall on the front side and a backward expanding expansion wall on the rear side.
2. The fuel injector rod for the combustion chamber of an aero-engine according to claim 1, characterized in that, The external auxiliary fuel injector (64) and the internal auxiliary fuel injector (65) include an external nozzle (12), an internal nozzle (13), and an internal cone (14); The inlet of the external nozzle (12) is connected to the fuel injector (6); The inner nozzle (13) is installed inside the outer nozzle (12). The outlets of the outer nozzle (12) and the inner nozzle (13) are conical, and a small vortex is installed between them. The inner cone (14) is set inside the inner nozzle (13), and is hollow inside. Its cylindrical end is connected to the fuel injection rod (6), and the cone end is located inside the outlet of the inner nozzle (13). A small vortex is set between the cone end and the outlet of the inner nozzle (13), and the cone end has multiple fuel injection holes distributed circumferentially on its side wall.
3. The fuel injector rod for the combustion chamber of an aero-engine according to claim 2, characterized in that, The distance between the main injector (61), the outer main injector (62), and the inner main injector (63) protruding from the head of the flame tube (5) on the injector rod (6) is L1, and the distance between the outer auxiliary injector (64) and the inner auxiliary injector (65) retracted into the head of the flame tube (5) is L2, where L2 > L1.
4. An aircraft engine combustion chamber, characterized in that, It includes an outer combustion chamber casing (1), an inner combustion chamber casing (2), an outer wall of the flame tube (3), an inner wall of the flame tube (4), a flame tube head (5), and an injection rod (6) for the combustion chamber of the aero-engine as described in claim 3. The combustion chamber casing (2) is installed inside the combustion chamber casing (1) to form the combustion chamber wall structure. The inlet of the combustion chamber wall structure is connected to the outlet of the diffuser (7), and the inlet of the diffuser (7) is connected to the outlet of the compressor (8). The outer wall (3) and inner wall (4) of the flame tube have mixing holes, which are set in the combustion chamber wall structure. The inner wall (4) of the flame tube is located inside the outer wall (3) of the flame tube, forming the flame tube. The outlet of the flame tube is directly opposite the high-pressure turbine guide (9). The head of the flame tube (5) is located at the inlet of the flame tube and has an air inlet. There are multiple fuel injector rods (6), which are circumferentially connected to the outer casing (1) of the combustion chamber; The flame tube head (5) has a flared mouth formed at the nozzles of the outer auxiliary oil nozzle (64) and the inner auxiliary oil nozzle (65), forming a triangular blunt body.
5. The aero-engine combustion chamber according to claim 4, characterized in that, During operation, the airflow from the compressor (8) is divided into three parts by the diffuser (7). The first part goes around the fuel injector (6) and enters the space between the outer casing (1) of the combustion chamber and the outer wall (3) of the flame tube, forming an outer ring airflow. The second part enters the space between the inner casing (2) of the combustion chamber and the inner wall (4) of the flame tube, forming an inner ring airflow. The third part enters the interior of the flame tube through the air inlet on the head (5) of the flame tube, forming a head airflow. The head airflow forms a blunt body return zone after the triangular blunt body. The outer ring airflow flows into the interior of the flame tube through the mixing hole on the outer wall (3) of the flame tube, and the other part is used to cool the outer wall (3) of the flame tube; the inner ring airflow flows into the interior of the flame tube through the mixing hole on the inner wall (4) of the flame tube, and the other part is used to cool the inner wall (4) of the flame tube; the head airflow, together with the outer ring airflow and the inner ring airflow that enter the interior of the flame tube through the mixing hole, forms the main return flow zone, and mixes with the fuel injected by the fuel injector on the fuel injector (6), and burns to form gas, which flows into the high-pressure turbine guide vane (9). The main reflux region is located after the bluff body reflux region and overlaps with it in space.
6. The aero-engine combustion chamber according to claim 5, characterized in that, The head airflow entering the interior of the flame tube through the air inlet on the flame tube head (5) is divided into six paths that are radially interspersed with the five fuel injectors on the fuel injector rod (6).
7. The aero-engine combustion chamber according to claim 6, characterized in that, It has low working condition, medium working condition and high working condition modes; In low operating conditions, the external auxiliary fuel injector (64) and the internal auxiliary fuel injector (65) are working, while the main fuel injector (61), the external main fuel injector (62), and the internal main fuel injector (63) are not working. This is suitable for low operating conditions in the combustion chamber. At this time, the fuel injected into the flame tube by the external auxiliary fuel injector (64) and the internal auxiliary fuel injector (65) will enter the bluff body recirculation zone for stable combustion. In the medium working condition, the main fuel injector (61), the external auxiliary fuel injector (64), and the internal auxiliary fuel injector (65) are working, while the external main fuel injector (62) and the internal main fuel injector (63) are not working. This is suitable for the medium working condition in the combustion chamber. At this time, the fuel injected into the flame tube by the external auxiliary fuel injector (64) and the internal auxiliary fuel injector (65) will enter the bluff body recirculation zone for stable combustion, and the fuel injected into the flame tube by the main fuel injector (61) will enter the main recirculation zone. The flame in the bluff body recirculation zone can ignite the fuel in the main recirculation zone to form the main combustion zone for combustion. In high-operational-condition operation, the main fuel injector (61), the outer main fuel injector (62), the inner main fuel injector (63), the outer auxiliary fuel injector (64), and the inner auxiliary fuel injector (65) are in operation. This is suitable for high-operational-condition combustion chambers. The fuel injected into the flame tube by the outer auxiliary fuel injector (64) and the inner auxiliary fuel injector (65) will enter the bluff body recirculation zone for stable combustion. The fuel injected into the flame tube by the main fuel injector (61), the outer main fuel injector (62), and the inner main fuel injector (63) will enter the main recirculation zone. The flame in the bluff body recirculation zone can ignite the fuel in the main recirculation zone, thus forming three main combustion zones for combustion.