An integrated ignition system for aircraft engines

Abstract

The present application provides an integrated ignition system for an aircraft engine, comprising: an aircraft throttle lever; an aircraft numerical control system for obtaining changes in the aircraft throttle lever angle; an engine controller for controlling engine ignition, the engine controller generating a main combustion chamber ignition instruction and / or an afterburner combustion chamber ignition instruction for main combustion chamber ignition and / or afterburner combustion chamber ignition according to aircraft throttle lever angle information processed by the aircraft numerical control system; a main combustion chamber ignition nozzle arranged in the main combustion chamber and an afterburner combustion chamber ignition nozzle arranged in the afterburner combustion chamber; an ignition device connecting the engine controller with the main combustion chamber ignition nozzle and the afterburner combustion chamber ignition nozzle, the ignition device controlling the main combustion chamber ignition nozzle and / or the afterburner combustion chamber ignition nozzle to ignite the corresponding combustion chamber according to the main combustion chamber ignition instruction and / or afterburner combustion chamber ignition instruction generated by the engine controller.

Description

Technical Field

[0001] The present application belongs to the technical field of aircraft engine ignition control, and in particular relates to an integrated ignition system for an aircraft engine. Background Art

[0002] The aircraft engine is the heart of an aircraft, providing power for takeoff, climb, cruise, maneuvering, and landing, making it an indispensable component. Aircraft engines generate thrust through intake, compression, combustion, power generation, and exhaust. Combustion is a crucial step, and the reliability and stability of the ignition system, as well as its maintainability once installed, significantly impact the performance of the aircraft engine.

[0003] With the increase of aircraft engine thrust, aircraft engines are divided into main combustion chamber and afterburner. Therefore, the current ignition system mainly includes two ignition systems: main combustion chamber ignition system and afterburner ignition system. The two ignition systems work and are designed independently. Figure 1 As shown, the main combustion chamber ignition system realizes the ignition of the main combustion chamber through the main combustion chamber ignition device 1, such as electric ignition; the afterburner combustion chamber ignition system realizes the ignition of the afterburner combustion chamber through the afterburner combustion chamber ignition device 2, such as jet ignition, electric ignition, etc.

[0004] However, the ignition scheme of the prior art has the following problems:

[0005] 1) The existing main combustion chamber ignition system and afterburner ignition system are designed independently, resulting in poor system integration, increased complexity of the ignition system, and increased failure rate;

[0006] 2) The main combustion chamber ignition system and the afterburner ignition system are designed independently, and the related accessories, cables, and external pipes occupy a large external space, which is not conducive to the external design of the engine;

[0007] 3) The main combustion chamber ignition system and the afterburner ignition system are designed independently, which increases the total weight of the engine and reduces the engine's thrust-to-weight ratio;

[0008] 4) The main combustion chamber ignition device and the afterburner combustion chamber ignition device are located at different axial and circumferential positions of the engine, which is not conducive to field maintenance of the engine in the installed state. At the same time, the aircraft needs to add maintenance covers at the corresponding positions, which to a certain extent reduces the stealth performance of the aircraft. Summary of the Invention

[0009] The purpose of the present application is to provide an integrated ignition system for an aircraft engine to solve or alleviate at least one problem in the background technology.

[0010] The technical solution of this application is: an integrated ignition system for an aircraft engine, comprising:

[0011] Aircraft throttle lever;

[0012] Aircraft numerical control system used to obtain aircraft throttle angle changes;

[0013] an engine controller for controlling engine ignition, the engine controller generating a main combustion chamber ignition instruction and / or an afterburner ignition instruction for igniting the main combustion chamber and / or the afterburner according to the aircraft throttle lever angle information processed by the aircraft numerical control system;

[0014] A main combustion chamber ignition nozzle disposed in the main combustion chamber and an afterburner combustion chamber ignition nozzle disposed in the afterburner combustion chamber;

[0015] An ignition device is connected to the engine controller and the main combustion chamber ignition nozzle and the afterburner combustion chamber ignition nozzle. The ignition device controls the main combustion chamber ignition nozzle and / or the afterburner combustion chamber ignition nozzle to ignite the corresponding combustion chamber according to the main combustion chamber ignition instruction and / or the afterburner combustion chamber ignition instruction generated by the engine controller.

[0016] In an optional embodiment of the present application, the integrated ignition system also includes an aircraft electrical system, which is connected to the aircraft digital control system, the engine controller and the ignition device to provide power to the aircraft digital control system, the engine controller and the ignition device.

[0017] In an optional embodiment of the present application, the ignition device is connected to the engine controller, the main combustion chamber ignition nozzle, the afterburner combustion chamber ignition nozzle and the aircraft electrical system through a connector.

[0018] In an optional embodiment of the present application, a maintenance cover is provided on the aircraft skin adapted to the ignition device.

[0019] In an optional embodiment of the present application, there are one or more main combustion chamber ignition nozzles and afterburner combustion chamber ignition nozzles, which are evenly distributed or distributed as needed in the main combustion chamber and afterburner combustion chamber.

[0020] In an optional embodiment of the present application, the main combustion chamber ignition nozzle and the afterburner chamber ignition nozzle are electric ignition nozzles.

[0021] The engine integrated ignition system of the present application improves the integration of the ignition system, simplifies the design of the ignition system, reduces the number of related accessories, cables and external pipelines, increases the available space outside the engine, makes the external design of the engine more flexible, and reduces the total weight of the engine, thereby improving the thrust-to-weight ratio of the engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0022] In order to more clearly illustrate the technical solutions provided by this application, the following is a brief introduction to the accompanying drawings. Obviously, the accompanying drawings described below are only some embodiments of this application.

[0023] Figure 1 It is a schematic diagram of an engine ignition system in the prior art.

[0024] Figure 2 This is a schematic diagram of the integrated ignition system for an aircraft engine of this application.

[0025] Figure 3 This is a schematic diagram of the working process of the aircraft engine integrated ignition system of this application.

[0026] Figure 4 This is a maintenance diagram of the integrated ignition system of this application.

[0027] Figure 5 This is a schematic diagram of the connection between the ignition device of this application and other devices or equipment. DETAILED DESCRIPTION

[0028] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below in conjunction with the drawings in the embodiments of this application.

[0029] In order to overcome the problems caused by the independent design of the main combustion chamber ignition system and the afterburner combustion chamber ignition system in the prior art, the present application proposes an integrated ignition system for an aircraft engine, which integrates the main combustion chamber ignition device and the afterburner combustion chamber ignition device into one ignition device. When the engine is started, the ignition device plays the role of the main combustion chamber ignition device and ignites the main combustion chamber; it determines whether the engine needs to be connected to the afterburner according to the throttle lever angle. When the engine needs to be connected to the afterburner, the main combustion chamber is in a working state, and the ignition device is used as the afterburner combustion chamber ignition device to ignite the afterburner.

[0030] like Figure 2 As shown, the aircraft engine integrated ignition system provided by the present application includes an aircraft throttle lever 101 located on the aircraft side 100, an aircraft numerical control system 102, an aircraft electrical system 103 and an engine controller 201, an ignition device 202, a main combustion chamber ignition nozzle 203, an afterburner combustion chamber ignition nozzle 204 and a cable 205 located on the engine side 200.

[0031] On the aircraft side 100, the aircraft digital control system 102 and the aircraft electrical system 103 form the interactive system between the aircraft side 100 and the engine side 200. The aircraft digital control system 102 connects to the aircraft throttle lever 101, the aircraft electrical system 103, and the engine controller 201. It receives angle change signals generated by the pilot actively adjusting the angle of the aircraft throttle lever 101 based on the aircraft's flight status, processes these angle change signals, and transmits the information to the engine controller 201. Simultaneously, after the ignition device 102 on the engine side 200 executes the ignition command, it receives engine parameter change signals from the engine controller 201, processes these signals, and transmits them to the pilot.

[0032] The aircraft electrical system 102 is connected to the aircraft digital control system 102 , the engine controller 201 and the ignition device 202 , and is used to provide power to the aircraft digital control system 102 , the engine controller 201 and the ignition device 202 .

[0033] On the engine side 200, the engine controller 201 is connected to the aircraft digital control system 102, the aircraft electrical system 103 and the ignition device 202. It obtains the information sent by the aircraft digital control system 102 and determines whether the engine needs to be ignited under the current flight state based on the information. If main combustion chamber ignition is required, a main combustion chamber ignition instruction is generated and sent to the ignition device 202. If afterburner ignition is required, an afterburner ignition instruction is generated and also sent to the ignition device 202.

[0034] Ignition device 202 is connected to a main combustion chamber ignition nozzle 203 in the main combustion chamber and an afterburner ignition nozzle 204 in the afterburner via a cable 205. There may be one or more main combustion chamber ignition nozzles 203 and afterburner ignition nozzles 204. When there are multiple main combustion chamber ignition nozzles 203 and afterburner ignition nozzles 204, they may be distributed uniformly around the circumference or arranged as needed. In some embodiments of the present application, main combustion chamber ignition nozzle 203 is typically an electric ignition nozzle, and afterburner ignition nozzle 204 may be the same type of electric ignition nozzle as main combustion chamber ignition nozzle 203, or may be another type of nozzle. After receiving the ignition command of the main combustion chamber or afterburner from the engine controller 201, the ignition device 202 controls the main combustion chamber ignition nozzle 203 to ignite the main combustion chamber through the cable 205, or controls the afterburner ignition nozzle 204 to ignite the afterburner through the cable 205. The pilot can judge whether the ignition of the main combustion chamber or afterburner is successful based on the changes in the engine parameters.

[0035] like Figure 3As shown in the working principle diagram of the integrated ignition system, when the pilot controls the aircraft throttle lever 101 and causes the angle of the aircraft throttle lever 101 to change, the aircraft digital control system 102 receives the throttle lever angle information and transmits it to the engine controller 201. The engine controller 201 determines whether the engine needs to ignite the main combustion chamber or the afterburner based on the throttle lever angle and issues an ignition command to the ignition device 202. After completing the ignition command operation, the ignition device 202 feeds back to the engine controller 101. The pilot can determine whether the main combustion chamber ignition or the afterburner ignition is successful based on the changes in engine parameters.

[0036] like Figure 4 As shown, in the installed state, the engine is located in the engine compartment and is shielded by the aircraft skin 104. To improve the field maintainability of the aircraft engine integrated ignition system of this application, a maintenance access cover 105 is added to the aircraft skin 104. The maintenance access cover 105 facilitates maintenance of the ignition device 202. It should be noted that the maintenance access cover 105 can generally be in a circular, rectangular, hexagonal, or other shape.

[0037] In addition, Figure 5 As shown, ignition device 202 is connected to engine controller 201, aircraft electrical system 203, and main combustion chamber or afterburner ignition nozzles via connector 206. If ignition device 202 or another device or equipment fails, maintenance or replacement of ignition device 202 or the corresponding device or equipment can be achieved by simply unplugging connector 206.

[0038] The aircraft engine integrated ignition system of the present application has the following advantages:

[0039] 1) This application improves the integration of the ignition system by integrating the main combustion chamber ignition system with the afterburner ignition system, simplifies the design of the ignition system, reduces the number of related accessories, cables, and external piping, increases the available space outside the engine, makes the engine's external design more flexible, and reduces the total weight of the engine, thereby improving the engine's thrust-to-weight ratio;

[0040] 2) The present invention improves the engine's field maintainability by integrating the main combustion chamber ignition system with the afterburner ignition system. At the same time, it reduces the number and area of ​​aircraft maintenance covers, thereby improving the aircraft's stealth performance to a certain extent.

[0041] The above description is merely a specific embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in this application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. An integrated ignition system for an aircraft engine, characterized in that: include: Aircraft throttle lever (101); An aircraft numerical control system (102) for obtaining changes in the angle of the aircraft throttle lever; An engine controller (201) for controlling engine ignition, the engine controller (201) generating a main combustion chamber ignition instruction and / or an afterburner ignition instruction for main combustion chamber ignition and / or afterburner ignition according to aircraft throttle lever angle information processed by an aircraft numerical control system (102); A main combustion chamber ignition nozzle (203) disposed in the main combustion chamber and an afterburner combustion chamber ignition nozzle (204) disposed in the afterburner combustion chamber; An ignition device (202) is connected to an engine controller (201) and a main combustion chamber ignition nozzle (203) and an afterburner combustion chamber ignition nozzle (204). The ignition device (202) is connected to the main combustion chamber ignition nozzle (203) in the main combustion chamber and the afterburner combustion chamber ignition nozzle (204) in the afterburner combustion chamber via a cable (205). The main combustion chamber ignition nozzle (203) and the afterburner combustion chamber ignition nozzle (204) are electric ignition nozzles. The ignition device (202) controls the main combustion chamber ignition nozzle (203) and / or the afterburner combustion chamber ignition nozzle (204) to ignite the corresponding combustion chamber according to the main combustion chamber ignition instruction and / or the afterburner combustion chamber ignition instruction generated by the engine controller (201).

2. The aircraft engine integrated ignition system according to claim 1, characterized in that: The integrated ignition system further comprises an aircraft electrical system (103), wherein the aircraft electrical system (103) is connected to the aircraft digital control system (102), the engine controller (201) and the ignition device (202), and is used to provide power to the aircraft digital control system (102), the engine controller (201) and the ignition device (202).

3. The aircraft engine integrated ignition system according to claim 2, characterized in that: The ignition device (202) is connected to the engine controller (201), the main combustion chamber ignition nozzle (203), the afterburner chamber ignition nozzle (204) and the aircraft electrical system (103) through a connector (206).

4. The aircraft engine integrated ignition system according to claim 3, characterized in that: A maintenance cover (105) is provided on the aircraft skin (104) adapted to the ignition device (202).

5. The aircraft engine integrated ignition system according to claim 1, wherein: The main combustion chamber ignition nozzle (203) and the afterburner combustion chamber ignition nozzle (204) are one or more and are evenly distributed or distributed as needed in the main combustion chamber and the afterburner combustion chamber.

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