Aero-engine combustion chamber and applicable fuel spray rod thereof

Abstract

The invention belongs to the technical field of aero-engine combustion chamber design, and particularly relates to an aero-engine combustion chamber and an applicable fuel spray rod thereof, an integrated multi-point injection oil nozzle design is adopted, multi-degree-of-freedom control can be carried out on the flow of a plurality of fuel nozzles of a single oil nozzle, and a bluff body flame stabilizing combustion organization mode is combined, so that the combustion efficiency of the aero-engine combustion chamber is improved. The turbine given design requirements such as the combustion chamber outlet temperature profile shape and the hot spot temperature are met through the partition oil control function of the oil nozzle.

Description

Technical Field

[0001] This application belongs to the field of aero-engine combustor design technology, specifically relating to an aero-engine combustor and its applicable fuel injector. 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 fuel injector and enters the space between the outer casing of the combustor and the outer wall of the flame tube, forming an outer annular 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 annular airflow. The third part enters the flame tube through the head of the flame tube, forming a head airflow. The outer annular 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 annular 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 annular 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] The typical fuel injector in an aero-engine combustion chamber usually uses a centrifugal + air atomizing injector. Each injector has 2 to 3 fuel lines. Fuel is injected from the injector into the flame tube in a conical spray pattern. This makes it difficult to precisely control the fuel distribution in the combustion field within the flame tube. This can easily lead to the combustion chamber outlet temperature profile and hot spot temperature exceeding design limits, even when the fuel-air ratio at the combustion chamber design point is close to the stoichiometric ratio.

[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 a suitable fuel injector thereon to overcome or mitigate at least one of the known technical defects.

[0007] The technical solution of this application is:

[0008] A fuel injector for an aircraft engine combustion chamber, the fuel injector having five fuel nozzles inserted into mounting holes in the head of the flame tube, arranged radially from the outside to the inside as an outer main fuel nozzle, an outer auxiliary fuel nozzle, a main fuel nozzle, an inner auxiliary fuel nozzle, and an inner main fuel nozzle.

[0009] The main injector, outer main injector, and inner main injector include a main nozzle and a fuel distribution cone. The inlet of the main nozzle is connected to the fuel injection rod, and the fuel distribution cone is installed inside the outlet. A small vortex is installed between the fuel distribution cone and the outlet of the main nozzle, and the fuel distribution cone and the end of the main nozzle outlet are narrowed. A large vortex is installed outside the outlet of the main nozzle.

[0010] The main nozzle's main nozzle inlet is connected to the first main fuel passage opened inside the fuel injector rod, and the main nozzle inlets of the outer and inner main nozzles are connected to the second main fuel passage opened inside the fuel injector rod.

[0011] The external and internal auxiliary fuel injectors include an external nozzle, an internal nozzle, and an internal cone. The inlet of the external nozzle is connected to the fuel injector rod. The internal nozzle is set inside the external nozzle. The outlets of the external and internal nozzles are conical, and a small swirler is set between them. The internal cone is set inside the internal nozzle. It is hollow inside, with its cylindrical end connected to the fuel injector rod. The conical end is located inside the outlet of the internal nozzle, and a small swirler is set between it and the outlet of the internal nozzle. The side wall of the conical end has multiple circumferentially distributed fuel injection holes.

[0012] The outer nozzle and inner nozzle have their outer nozzle inlets connected to the first secondary fuel passage opened inside the fuel injector rod, and their inner cones connected to the second secondary fuel passage opened inside the fuel injector rod.

[0013] According to at least one embodiment of this application, in the fuel injector applicable to the combustion chamber of the above-mentioned aircraft engine, the first main fuel passage and the second main fuel passage are connected to the aircraft engine fuel source through the first oil passage and the second oil passage;

[0014] The first and second auxiliary fuel passages are connected to the fuel source of the aircraft engine via a third fuel line.

[0015] According to at least one embodiment of this application, in the fuel injector applicable to the combustion chamber of the above-mentioned aircraft engine, a fuel quantity adjustment mechanism is provided between the first auxiliary fuel passage, the second auxiliary fuel passage and the third fuel passage. The fuel quantity adjustment mechanism includes a housing, a piston seat, a piston sleeve, a piston, a clamp, and a spring.

[0016] The outer casing has an oil inlet, a first oil outlet, and a second oil outlet. The oil inlet is connected to a third oil passage, the first oil outlet is connected to a first auxiliary fuel passage, and the second oil outlet is connected to a second auxiliary fuel passage.

[0017] The piston seat is disposed inside the housing, forming a normally open flow channel between the piston seat and the housing. The normally open flow channel is connected to the second oil outlet. The piston seat has an annular cavity and a flow port connected to the annular cavity. The flow port is connected to the first oil outlet.

[0018] The piston sleeve is stepped and hollow, with an oil inlet on the stepped part. The large end of the piston sleeve is located inside the annular cavity, and the small end faces the oil inlet.

[0019] The piston head is located inside the large end of the piston sleeve, and the piston rod extends out from the small end of the piston sleeve. The piston head has an overflow chamber that communicates with the large end of the piston sleeve. The overflow chamber is annular and surrounds the piston rod. The side wall of the piston head has multiple overflow holes that are distributed circumferentially and communicate with the overflow chamber.

[0020] The clamp is connected to the end of the piston rod;

[0021] The spring is sleeved on the outer periphery of the small end of the piston sleeve, and its two ends abut against the stepped part of the clamp and the piston sleeve. Relying on its elastic force, the head of the piston is kept inside the large end of the piston sleeve, and each overflow hole is submerged in the large end of the piston sleeve and blocked by the large end of the piston sleeve.

[0022] 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 suitable for the aforementioned aircraft engine combustion chamber.

[0023] 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.

[0024] 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.

[0025] The head of the flame tube is located at the flame tube inlet and has an air inlet.

[0026] There are multiple fuel injector booms, which are circumferentially connected to the outer casing of the combustion chamber. Five fuel injectors are radially distributed on the boom and inserted into the mounting holes in the head of the flame tube.

[0027] 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.

[0028] According to at least one embodiment of this application, in the aforementioned aircraft engine combustion chamber, the five fuel injectors on the fuel injector boom have a recessed distance relative to the head of the flame tube.

[0029] According to at least one embodiment of this application, in the above-mentioned aero-engine combustion chamber, the retraction distance of the main fuel injector, outer main fuel injector, and inner main fuel injector on the fuel injector rod relative to the head of the flame tube is L1, and the retraction distance of the outer auxiliary fuel injector and inner auxiliary fuel injector relative to the head of the flame tube is L2, where L2 > L1.

[0030] 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.

[0031] The outer ring airflow flows into the flame tube through the mixing holes 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 holes 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 holes, 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;

[0032] The reflux zone is located behind the bluff body reflux zone and overlaps with it in space.

[0033] According to at least one embodiment of this application, in the aforementioned aircraft engine combustion chamber, the head airflow entering the interior of the flame tube through the air inlet on the flame tube head is divided into six paths that are radially alternately distributed with the five fuel injectors on the fuel injector rod.

[0034] 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.

[0035] When operating under low 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 operating conditions in the combustion chamber.

[0036] When operating under medium conditions, 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 medium conditions in the combustion chamber.

[0037] Under high operating conditions, the main fuel injector, external main fuel injector, internal main fuel injector, external auxiliary fuel injector, and internal auxiliary fuel injector are all in operation, which is suitable for high operating conditions in the combustion chamber. Attached Figure Description

[0038] Figure 1This is a schematic diagram of a typical combustion chamber design for current aero engines;

[0039] Figure 2 This is a schematic diagram of an aircraft engine combustion chamber provided in an embodiment of this application;

[0040] Figure 3 This is a partial schematic diagram of an aero-engine combustion chamber provided in an embodiment of this application;

[0041] Figure 4 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;

[0042] Figure 5 This is a schematic diagram of the operation of the main fuel injector, outer main fuel injector, and inner main fuel injector provided in the embodiments of this application;

[0043] Figure 6 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;

[0044] Figure 7 This is a schematic diagram of the operation of the external auxiliary fuel injector and the internal auxiliary fuel injector provided in the embodiments of this application;

[0045] Figure 8 This is a schematic diagram of the oil quantity regulating mechanism provided in the embodiments of this application;

[0046] Figure 9 This is a partial schematic diagram of the oil quantity regulating mechanism provided in the embodiments of this application;

[0047] Figure 10 This is a schematic diagram of the piston sleeve provided in an embodiment of this application;

[0048] Figure 11 This is a schematic diagram of the piston provided in an embodiment of this application;

[0049] in:

[0050] 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-Fuel injector; 7-Diffuser; 8-Compressor; 9-High-pressure turbine guide vane;

[0051] 61-Main fuel injector; 62-Outer main fuel injector; 63-Inner main fuel injector; 64-Outer auxiliary fuel injector; 65-Inner auxiliary fuel injector;

[0052] 10-Main nozzle; 11-Distribution cone; 12-Outer nozzle; 13-Inner nozzle; 14-Inner cone; 15-Outer casing; 16-Piston seat; 17-Piston sleeve; 18-Piston; 19-Clamp; 20-Spring;

[0053] A - First main fuel passage; B - Second main fuel passage;

[0054] C - First auxiliary fuel passage; D - Second auxiliary fuel passage;

[0055] E - Oil inlet; F - First oil outlet; G - Second oil outlet; H - Normally open flow channel.

[0056] 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

[0057] 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.

[0058] 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.

[0059] 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.

[0060] An aircraft engine combustion chamber, such as Figure 2 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 a fuel injection boom 6.

[0061] 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.

[0062] 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.

[0063] 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.

[0064] The head 5 of the flame tube is located at the inlet of the flame tube and has an air inlet.

[0065] There are multiple fuel injector booms 6, which are circumferentially connected to the outer casing 1 of the combustion chamber. Each boom has five radially distributed fuel injectors that are inserted into the mounting holes in the head of the flame tube 5.

[0066] Of the five fuel injectors on the fuel injector boom 6, the one in the middle is the main fuel injector 61, the one near the outer wall 3 of the flame tube is the outer main fuel injector 62, the one near the inner wall 4 of the flame tube is the inner main fuel injector 63, the one between the main fuel injector 61 and the outer main fuel injector 62 is the outer auxiliary fuel injector 64, and the one between the main fuel injector 61 and the inner main fuel injector 63 is the inner auxiliary fuel injector 65.

[0067] The five fuel injectors on the fuel injector boom 6 have a recessed distance relative to the flame tube head 5. Specifically, the recessed distance between the main fuel injector 61, the outer main fuel injector 62, and the inner main fuel injector 63 and the flame tube head 5 is L1; the recessed distance between the outer auxiliary fuel injector 64 and the inner auxiliary fuel injector 65 and the flame tube head 5 is L2, where L2 > L1. Figure 3 As shown.

[0068] 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.

[0069] 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 recirculation 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 recirculation 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.

[0070] 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 alternated with the five fuel injectors on the fuel injector rod 6, and are specifically divided by airflow baffles.

[0071] 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.

[0072] The aero-engine combustion chamber disclosed in the above embodiments can be designed to have low operating conditions, medium operating conditions, and high operating conditions.

[0073] 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.

[0074] 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.

[0075] 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.

[0076] The main fuel injector 61, outer main fuel injector 62, and inner main fuel injector 63 can be further designed, such as... Figure 4 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, 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.

[0077] The inlet of the main nozzle 10 of the main injector 61 is connected to the first main fuel passage A opened inside the fuel injector 6. The inlets of the main nozzles 10 of the outer main injector 62 and the inner main injector 63 are connected to the second main fuel passage B opened inside the fuel injector 6. The first main fuel passage A and the second main fuel passage B are connected to the fuel source of the aircraft engine through the first oil passage and the second oil passage.

[0078] When the aero-engine is operating, fuel enters the main nozzle 10 through the first main fuel passage A and the main nozzle 61 through the second main fuel passage B, passing through the outer main nozzle 62 and the inner main nozzle 63. The fuel distribution cone 11 transforms the fuel entering the main nozzle 10 into a ring-shaped flow, which is then swirled by a small vortex generator before being ejected from the outlet of the main nozzle 10. Part of the head airflow can be transformed into swirling air by a large vortex generator, further atomizing and breaking up the fuel ejected from the outlet of the main nozzle 10. This facilitates thorough mixing and efficient combustion of fuel and air under high operating conditions. Figure 5 As shown.

[0079] External auxiliary fuel injector 64, internal auxiliary fuel injector 65, etc. Figure 6 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. The inner nozzle 13 is disposed inside the outer nozzle 12. The outlets of the outer nozzle 12 and the inner nozzle 13 are conical, and a small swirler is disposed between them. The inner cone 14 is disposed 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 swirler is disposed between it and the outlet of the inner nozzle 13. The side wall of the conical end has multiple circumferentially distributed fuel injection holes.

[0080] The inlet of the outer nozzle 12 of the outer auxiliary nozzle 64 and the inner auxiliary nozzle 65 is connected to the first auxiliary fuel passage C opened inside the fuel injector 6, and the inner cone 14 is connected to the second auxiliary fuel passage D opened inside the fuel injector 6. The first auxiliary fuel passage C and the second auxiliary fuel passage D are connected to the fuel source of the aircraft engine through the third oil passage.

[0081] When the aero-engine is operating, fuel can enter the outer nozzle 12 of the outer auxiliary injector 64 and inner auxiliary injector 65 through the first auxiliary fuel passage C, and enter the inner cone 14 of the outer auxiliary injector 64 and inner auxiliary injector 65 through the second auxiliary fuel passage D. The fuel entering the outer auxiliary injector 64 is swirled by a small vortex generator before being ejected, while the fuel entering the inner cone 14 enters the inner auxiliary injector 65 through the injection hole, and is then swirled by the small vortex generator before being ejected. The fuel ejected from the outer auxiliary injector 64 and inner auxiliary injector 65 is centrifugally atomized, which facilitates stable combustion during ignition and under low operating conditions, thus playing a role in flame stabilization. Figure 7 As shown.

[0082] A fuel quantity adjustment mechanism can be installed between the first auxiliary fuel passage C, the second auxiliary fuel passage D, and the third fuel line. The fuel quantity adjustment mechanism includes a housing 15, a piston seat 16, a piston sleeve 17, a piston 18, a clamp 19, and a spring 20. Figures 8-11 As shown.

[0083] The outer casing 15 has an oil inlet E, a first oil outlet F, and a second oil outlet G. The oil inlet E is connected to the third oil passage, the first oil outlet F is connected to the first auxiliary fuel passage C, and the second oil outlet G is connected to the second auxiliary fuel passage D.

[0084] The piston seat 16 is disposed inside the housing 15, and forms a normally open flow channel H between itself and the housing 15. The normally open flow channel H is connected to the second oil outlet G. The piston seat 16 has an annular cavity and a flow port connected to the annular cavity. The flow port is connected to the first oil outlet F.

[0085] The piston sleeve 17 is stepped and hollow, with an oil passage on the stepped part. The large end of the piston sleeve 17 is located inside the annular cavity, and the small end faces the oil inlet E.

[0086] The head of piston 18 is located inside the large end of piston sleeve 17, and the rod of piston 18 extends out from the small end of piston sleeve 17. The head of piston 18 has an overflow cavity that communicates with the large end of piston sleeve 17. The overflow cavity is annular and surrounds the rod of piston 18. The side wall of the head of piston 18 has multiple overflow holes that are distributed circumferentially and communicate with the overflow cavity.

[0087] Clamp 19 is connected to the end of the piston rod.

[0088] Spring 20 is sleeved on the outer periphery of the small end of piston sleeve 17, and both ends abut against the stepped part of clamp 19 and piston sleeve 17. Relying on its elastic force, the head of piston 18 is kept inside the large end of piston sleeve 17, and each overflow hole is submerged in the large end of piston sleeve 17 and blocked by the large end of piston sleeve 17.

[0089] When the aero-engine is operating, when the flow rate of fuel entering the casing 15 through the third oil passage via the inlet E is small, the oil pressure is low. The pressure generated by the fuel on both sides of the piston 18 is insufficient to overcome the elastic force of the spring 20 and push the piston 18 to move within the piston sleeve 17. The fuel entering the casing 15 can only flow through the normally open flow channel H and the second outlet G to the second auxiliary fuel passage D. When the flow rate of fuel entering the casing 15 through the third oil passage via the inlet E is large, the oil pressure is large. The pressure generated by the fuel on both sides of the piston 18 can overcome the elastic force of the spring 20 and push the piston 18 to move within the piston sleeve 17. This causes the overflow holes to protrude from the large end of the piston sleeve 17 into the overflow chamber. The fuel entering the casing 15 can flow into the overflow chamber through the oil inlet, and then into the annular cavity through the overflow holes, and finally flow into the first auxiliary fuel passage C through the first outlet F.

[0090] 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.

[0091] The aero-engine combustor disclosed in the above embodiments adopts a combination of radially parallel two-way centrifugal and three-way air-atomizing fuel injectors, and supplies fuel to the fuel injectors through three fuel lines. With separate fuel supply and control, the fuel quantity can be controlled according to the test results of the combustor components for various typical operating conditions of different aero-engines, so as to meet the limiting requirements of combustor outlet temperature profile shape and hot spot temperature under various typical operating conditions, especially under large operating conditions.

[0092] 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 injector suitable 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), 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 fuel 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 main nozzle (61) has its main nozzle pipe (10) inlet connected to the first main fuel passage (A) opened inside the fuel injector (6), and the main nozzle pipe (10) inlets of the outer main nozzle (62) and inner main nozzle (63) have their main nozzle pipe (10) inlet connected to the second main fuel passage (B) opened inside the fuel injector (6). The external auxiliary fuel injector (64) and the internal auxiliary fuel injector (65) include an external injector pipe (12), an internal injector pipe (13), and an internal cone (14). The inlet of the external injector pipe (12) is connected to the fuel injector rod (6). The internal injector pipe (13) is set inside the external injector pipe (12). The outlets of the external injector pipe (12) and the internal injector pipe (13) are conical, and a small vortex is set between them. The internal cone (14) is set inside the internal injector pipe (13). It is hollow inside, with its cylindrical end connected to the fuel injector rod (6). The cone end is located inside the outlet of the internal injector pipe (13), and a small vortex is set between it and the outlet of the internal injector pipe (13). The cone end has multiple circumferentially distributed injection holes on its side wall. The outer nozzle (64) and inner nozzle (65) have their outer nozzle pipe (12) inlet connected to the first auxiliary fuel passage (C) opened inside the fuel injector rod (6), and their inner cone (14) is connected to the second auxiliary fuel passage (D) opened inside the fuel injector rod (6).

2. The fuel injector suitable for an aircraft engine combustion chamber according to claim 1, characterized in that, The first main fuel passage (A) and the second main fuel passage (B) are connected to the fuel source of the aircraft engine through the first fuel line and the second fuel line; The first auxiliary fuel passage (C) and the second auxiliary fuel passage (D) are connected to the fuel source of the aircraft engine through the third fuel line.

3. The fuel injector suitable for an aircraft engine combustion chamber according to claim 2, characterized in that, A fuel quantity adjustment mechanism is provided between the first auxiliary fuel passage (C), the second auxiliary fuel passage (D) and the third fuel line. The fuel quantity adjustment mechanism includes a housing (15), a piston seat (16), a piston sleeve (17), a piston (18), a clamp (19), and a spring (20). The outer casing (15) has an oil inlet (E), a first oil outlet (F), and a second oil outlet (G), wherein the oil inlet (E) is connected to the third oil passage, the first oil outlet (F) is connected to the first auxiliary fuel passage (C), and the second oil outlet (G) is connected to the second auxiliary fuel passage (D). The piston seat (16) is disposed inside the housing (15) and forms a normally open flow channel (H) with the housing (15). The normally open flow channel (H) is connected to the second oil outlet (G). The piston seat (16) has an annular cavity and a flow port connected to the annular cavity. The flow port is connected to the first oil outlet (F). The piston sleeve (17) is stepped and hollow. It has an oil passage on the stepped part. The large end of the piston sleeve (17) is located in the annular cavity, and the small end faces the oil inlet (E). The head of the piston (18) is located inside the large end of the piston sleeve (17), and the rod of the piston (18) passes through the small end of the piston sleeve (17). The head of the piston (18) has an overflow cavity that communicates with the large end of the sleeve (17). The overflow cavity is annular and surrounds the rod of the piston (18). The side wall of the head of the piston (18) has multiple overflow holes that are distributed circumferentially and communicate with the overflow cavity. Clamp (19) is connected to the end of the piston rod; The spring (20) is sleeved on the outer periphery of the small end of the piston sleeve (17), and both ends abut against the step of the clamp (19) and the piston sleeve (17). Relying on its elastic force, the head of the piston (18) is kept inside the large end of the piston sleeve (17), and each overflow hole is submerged inside the large end of the piston sleeve (17) and blocked by the large end of the piston sleeve (17).

4. An aircraft engine combustion chamber, characterized in that, 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 a fuel injector (6) suitable for the combustion chamber of an 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 bars (6), which are circumferentially connected to the outer casing (1) of the combustion chamber. Five fuel injectors are radially distributed on the injector bars and inserted into the mounting holes opened in the flame tube head (5). 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, The five fuel injectors on the fuel injector boom (6) have a recessed distance relative to the head of the flame tube (5).

6. The aero-engine combustion chamber according to claim 5, characterized in that, The indentation distance of the main fuel injector (61), the outer main fuel injector (62), and the inner main fuel injector (63) of the fuel injector rod (6) relative to the head of the flame tube (5) is L1, and the indentation distance of the outer auxiliary fuel injector (64) and the inner auxiliary fuel injector (65) relative to the head of the flame tube (5) is L2, where L2 > L1.

7. The aero-engine combustion chamber according to claim 6, 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 rod (6), burns to form gas, and flows into the high-pressure turbine guide vane (9). The reflux zone is located behind the bluff body reflux zone and overlaps with it in space.

8. The aircraft engine combustion chamber according to claim 7, 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).

9. The aircraft engine combustion chamber according to claim 8, characterized in that, It has low working condition, medium working condition and high working condition modes; When operating under low 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. When operating under medium conditions, 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 medium conditions in the combustion chamber. When operating under high conditions, the main fuel injector (61), the external main fuel injector (62), the internal main fuel injector (63), the external auxiliary fuel injector (64), and the internal auxiliary fuel injector (65) operate, which is suitable for high conditions in the combustion chamber.

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