System for preventing coking and carbon deposition of fuel nozzle of aviation gas turbine engine

Abstract

The invention provides a system for preventing coking and carbon deposition of a fuel nozzle of an aviation gas turbine engine. The system comprises an air entraining system, an axial flow fan assembly, a first air entraining pipe and a cooling air pipe, the outlet end of the cooling air pipe is communicated with the air entraining system, and the inlet end of the cooling air pipe is connected with the axial flow fan assembly; the outlet end of the first air guide pipe is connected with the axial flow fan assembly, and an air vent is formed in the outlet end and communicated with the outside. According to the invention, coking and carbon deposition of the fuel nozzle and nozzle atomization capacity change caused by coking are eliminated. Nozzle atomization capacity changes influence non-uniformity of a turbine inlet temperature field and influence the working capacity of a turbine, so that the performance of an engine is reduced, the service life of turbine blades is shortened, and the on-wing service time of the engine is shortened.

Description

Technical Field

[0001] This invention relates to the field of cooling technology for the system structure of aviation gas turbine engines, and particularly to a system for preventing coking and carbon buildup on the fuel nozzles of aviation gas turbine engines. Background Technology

[0002] In the field of cooling technology for aero gas turbine engine system structures, market competition and the trend towards green aviation are driving continuous increases in the design cycle parameters of gas turbine engines, resulting in increasingly higher ambient and space temperatures for components during operation and shutdown. Previously, aero gas turbine engine designs placed great emphasis on component cooling technology during engine operation. Thermal management systems designed for engine operation cooled components and systems by introducing ambient air and bleed air from the high / low compressors, such as cooling turbine rotor blades and cooling after lubricating oil circulation. The cooling flow path ensured that the component and ambient temperatures remained within acceptable ranges.

[0003] Although the engine design fully considers the cooling of component systems under various operating conditions, most of the engine's cooling systems cease to function after the engine completes flight shutdown, and the engine relies mainly on natural cooling, which brings new problems.

[0004] Currently, after a period of use, inspections of some new aviation gas turbine engines have revealed coking and carbon buildup in the fuel injectors. Analysis has shown that this overlooks scenarios where some engine components remain in a prolonged high-temperature environment after shutdown. This is particularly problematic at airports with high temperatures, where the engine is enclosed within systems such as nacelles, where residual heat from some components can raise the temperature of other engine parts.

[0005] At the same time, in order to reduce air pollution from combustion emissions, such as Figure 1 As shown, the fuel nozzle A in the combustion chamber is designed with a multi-stage fuel supply structure (such as a main combustion stage and a secondary combustion stage). It has many small fuel channels internally and many small-orifice fuel injectors externally. The fuel nozzle is in a "hot-wet" space and environment without cooling measures. Therefore, the nozzle area experiences high temperatures, leading to the coking and accumulation of unburned fuel.

[0006] Coking and carbon buildup on fuel injectors can cause numerous problems, such as poor atomization performance, uneven turbine inlet temperature, resulting in insufficient turbine power, reduced thrust, or even in-flight engine failure. While cold-running the engine can reduce the surface temperature of the fuel injectors to some extent, this requires adequate external logistical support and equipment at the airport, such as ground-based gas cylinders, maintenance personnel, aircraft crew, and onboard gas supply systems.

[0007] However, not all airports and flight schedules can meet these requirements. Existing solutions involve passively controlling and shortening the engine's service life (e.g., setting a limit of 1000 cycles) during actual operation. Once the engine has reached the specified number of cycles and service time, it is removed from the aircraft and returned to an engine repair shop or quick-service center. Carbon-deposited fuel nozzles are removed, cleaned, reinstalled, or have their assemblies replaced. In some cases, the engine may need to be retested before being sent back to the aircraft.

[0008] Technical problems caused by coking and carbon buildup on fuel injectors shorten the expected number of times and lifespan of engines, impose additional maintenance work on airlines, cause numerous interruptions and disruptions to fleet operations, increase modification and service costs for engine manufacturers, and reduce customer loyalty to the brand.

[0009] Because the combustion chamber of an aviation gas turbine engine is equipped with multiple fuel nozzles, which are connected to the fuel control system, when the engine is working, fuel and high-pressure air flow through it, carrying away a large amount of heat. During operation, the temperature of the metal surface of the fuel nozzle is lower than the temperature at which fuel coking occurs, so fuel nozzle coking will not occur.

[0010] However, after an engine shutdown, the engine and nacelle form a relatively enclosed space (without forced convection cooling). The fuel nozzles themselves are very hot, and some unburned aviation kerosene contains heavy aromatics, gums, and asphalt, which undergo dehydrogenation condensation with oxygen in the air under high temperatures, forming large molecules that deposit on the metal surface of the fuel nozzles, creating coke nuclei. These initially formed coke nuclei continue to aggregate and dehydrogenate, forming particulate carbon deposits that clog the fuel injection holes and fuel passages. The only way to remove these carbon deposits is by disassembling the engine's fuel nozzles. This shortens the engine's on-wing service time and overhaul intervals, and in extreme cases, can lead to in-flight engine shutdowns.

[0011] In view of this, the inventors of this application have designed a system for preventing coking and carbon buildup on fuel nozzles of aviation gas turbine engines, in order to overcome the aforementioned technical problems. Summary of the Invention

[0012] The technical problem to be solved by the present invention is to overcome the defects in the prior art, such as the easy coking and carbon deposits on the surface of the fuel nozzle after the engine is stopped, and to provide a system for preventing coking and carbon deposits on the fuel nozzle of an aviation gas turbine engine.

[0013] The present invention solves the above-mentioned technical problems through the following technical solution:

[0014] A system for preventing coking and carbon buildup on fuel nozzles of an aviation gas turbine engine, characterized in that the system includes an air intake system, an axial flow fan assembly, a first air intake duct, and a cooling duct, wherein the outlet end of the cooling duct is connected to the air intake system, and the inlet end is connected to the axial flow fan assembly.

[0015] The outlet end of the first air intake pipe is connected to the axial flow fan assembly, and the outlet end is provided with an air vent that communicates with the outside.

[0016] According to one embodiment of the present invention, the axial flow fan assembly includes an axial flow fan with a motor, the axial flow fan being installed between the cooling air duct and the first air duct.

[0017] According to one embodiment of the present invention, the axial flow fan assembly further includes a second air intake pipe and a switching valve, one end of the second air intake pipe being connected to the switching valve and the other end being connected to the axial flow fan, and the switching valve being installed between the cooling air duct and the second air intake pipe.

[0018] According to one embodiment of the present invention, the axial flow fan assembly further includes a third air intake pipe and a one-way valve, one end of the third air intake pipe being connected to the switching valve and the other end being connected to the one-way valve, the one-way valve being installed between the cooling air duct and the third air intake pipe.

[0019] According to one embodiment of the present invention, the system further includes an engine health monitoring device, which is connected to the bleed air monitoring computer of the bleed air system.

[0020] According to one embodiment of the present invention, the axial flow fan is communicatively connected to the air intake monitoring computer of the air intake system.

[0021] According to one embodiment of the present invention, the switching valve is communicatively connected to the bleed air monitoring computer of the bleed air system.

[0022] According to one embodiment of the present invention, the one-way valve is communicatively connected to the bleed air monitoring computer of the bleed air system.

[0023] According to one embodiment of the present invention, the air intake system includes a high-pressure shut-off valve and a high-pressure air intake pipe. The outlet end of the cooling air duct forms two paths, one of which is connected to the high-pressure shut-off valve and the other of which is connected to the high-pressure air intake pipe.

[0024] The positive and progressive effects of this invention are as follows:

[0025] The system for preventing coking and carbon buildup on fuel nozzles of aviation gas turbine engines, as described in this invention, has the following advantages:

[0026] First, it eliminates the effects of fuel nozzle coking and carbon buildup, as well as the resulting changes in nozzle atomization capability. Changes in nozzle atomization capability affect the uniformity of the turbine inlet temperature field, impacting the turbine's power output, leading to decreased engine performance, reduced turbine blade life, and shortened engine on-wing service time.

[0027] Second, it eliminates the possibility that fuel injectors may fail to function due to blockage, thus increasing the likelihood that gas turbine engines will exceed industry standards and have an environmental impact.

[0028] Third, it eliminates unplanned engine replacements for customers due to fuel coking, reduces operational disruptions to the aircraft fleet, and reduces the cost of spare parts / consumables and labor for disassembling and assembling engines on aircraft, disassembling engines in engine repair shops, and replacing fuel nozzles. Attached Figure Description

[0029] The above and other features, properties and advantages of the present invention will become more apparent from the following description taken in conjunction with the accompanying drawings and embodiments, in which the same reference numerals always denote the same features, wherein:

[0030] Figure 1 This is a schematic diagram of the structure of a typical aviation gas turbine engine's fuel nozzle.

[0031] Figure 2 This is a schematic diagram illustrating the working principle of the system for preventing coking and carbon buildup on fuel nozzles of aviation gas turbine engines according to the present invention.

[0032] Figure 3 This invention relates to a system for preventing coking and carbon buildup on fuel nozzles of an aviation gas turbine engine, showing the working principle of the bleed air system when the engine is at idle.

[0033] Figure 4 This invention relates to a system for preventing coking and carbon buildup on fuel nozzles of an aviation gas turbine engine, showing the working principle of the bleed air system when the engine is not at idle. Detailed Implementation

[0034] To make the above-mentioned objects, features and advantages of the present invention more apparent and understandable, the specific embodiments of the present invention will be described in detail below with reference to the accompanying drawings.

[0035] Embodiments of the invention will now be described in detail with reference to the accompanying drawings. Preferred embodiments of the invention will now be described in detail, examples of which are shown in the drawings. Wherever possible, the same reference numerals will be used in all the drawings to denote the same or similar parts.

[0036] Furthermore, although the terminology used in this invention is selected from commonly known and used terms, some terms mentioned in this specification may have been selected by the applicant in his or her judgment, and their detailed meanings are explained in the relevant sections of the description herein.

[0037] Furthermore, the invention should be understood not only through the actual terminology used, but also through the meaning implied by each term.

[0038] Figure 2 This is a schematic diagram illustrating the working principle of the system for preventing coking and carbon buildup on fuel nozzles of aviation gas turbine engines according to the present invention.

[0039] like Figure 2 As shown, this invention discloses a system for preventing coking and carbon buildup on fuel nozzles of an aviation gas turbine engine, comprising an air bleed system 100, an axial flow fan assembly 200, a first air bleed pipe 10, and a cooling air duct 20. The outlet end of the cooling air duct 20 is connected to the air bleed system 100, and the inlet end is connected to the axial flow fan assembly 200. The outlet end of the first air bleed pipe 10 is connected to the axial flow fan assembly 200, and the outlet end is provided with a vent 11, communicating with the outside.

[0040] Preferably, the axial flow fan assembly 200 includes an axial flow fan 210 with a motor, which is installed between the cooling air duct 20 and the first air intake duct 10.

[0041] The axial flow fan assembly 200 also includes a second air intake pipe 220 and a switching valve 230. One end of the second air intake pipe 220 is connected to the switching valve 230, and the other end is connected to the axial flow fan 210. The switching valve 230 is installed between the cooling air duct 210 and the second air intake pipe 220.

[0042] Furthermore, the axial flow fan assembly 200 may also include a third air intake pipe 240 and a one-way valve 250. One end of the third air intake pipe 240 is connected to the on / off valve 230, and the other end is connected to the one-way valve 250. The one-way valve 250 is installed between the cooling air duct 20 and the third air intake pipe 240. Here, if the high-pressure and medium-pressure sealing is good, the one-way valve 250 can be omitted.

[0043] In particular, the aforementioned axial flow fan 210, second air intake pipe 220, switching valve 230, third air intake pipe 240 and one-way valve 250 can be combined into a single device.

[0044] Preferably, the air intake system 100 includes a high-pressure shut-off valve 110 and a high-pressure air intake pipe 120. The outlet end of the cooling air duct 20 forms two paths, one of which is connected to the high-pressure shut-off valve 110 and the other of which is connected to the high-pressure air intake pipe 120.

[0045] In addition, the system also includes an engine health monitoring device 30, which is connected to the bleed air monitoring computer 130 of the bleed air system 100.

[0046] The axial flow fan 210 is communicatively connected to the induced draft monitoring computer 130 of the induced draft system 100. The switching valve 230 is communicatively connected to the induced draft monitoring computer 130 of the induced draft system 100. The one-way valve 250 is communicatively connected to the induced draft monitoring computer 130 of the induced draft system 100.

[0047] like Figure 2 As shown, in this embodiment, the bleed air system 100 (medium-pressure / high-pressure bleed air system) includes a compressor medium-pressure bleed air port 300, a compressor high-pressure bleed air port 140, a medium-pressure check valve 150, a high-pressure shut-off valve 110, a pressure regulating valve 170, a pressure sensor (upstream) 180, a pressure sensor (downstream) 181, an aircraft cabin environmental control interface 190, a bleed air monitoring computer (BMC) 130, a medium-pressure bleed air pipe 310, a high-pressure bleed air pipe 120, a first cabin environmental control bleed air pipe 160, and a second cabin environmental control bleed air pipe 161.

[0048] The bleed air system 100 consists of two branches. The first branch leads from the engine's high-pressure bleed air port 140 through the high-pressure bleed air pipe 120, connects to the high-pressure shut-off valve 110, and then connects to the first cabin environmental control bleed air pipe 160. A pressure sensor (upstream) 180, a pressure regulating valve 170, and a pressure sensor (downstream) 181 are sequentially connected to the first cabin environmental control bleed air pipe 160, which is then connected to the aircraft cabin environmental control interface 190 via the second cabin environmental control bleed air pipe 161.

[0049] The second branch line extends from the engine's intermediate-pressure bleed air inlet 300 through the intermediate-pressure bleed air pipe 310, connects to the total pressure check valve 320, and then connects to the first cabin environmental control bleed air pipe 160. A pressure sensor (upstream) 180, a pressure regulating valve 170, and a pressure sensor (downstream) 181 are sequentially connected to the first cabin environmental control bleed air pipe 160, and then the second cabin environmental control bleed air pipe 161 connects to the aircraft cabin environmental control interface 190.

[0050] The working principle of the air intake system is as follows:

[0051] First, when the aircraft is on the ground (engines not yet started): the aircraft cabin environmental control system requires bleed air to operate, and auxiliary power is used to supply air to the environmental control system.

[0052] Next, as Figure 3As shown, when the aircraft is taxiing on the ground (at idle after engine start): the Bleed Air Monitoring Computer (BMC) 130 controls the opening of the high-pressure shut-off valve 110, at which time bleed air is drawn from the engine's high-pressure bleed port 140 through the high-pressure bleed air pipe 120. Then, after passing through the opened high-pressure shut-off valve 110 (at which time the intermediate-pressure check valve 150 is closed to isolate the bleed air from the compressor's high-pressure bleed port 140), it enters the pressure regulating valve 170 through the first cabin environmental control bleed air pipe 160. The Bleed Air Monitoring Computer (BMC) 130 adjusts the opening of the pressure regulating valve 170 based on the pressure difference between the upstream and downstream pressure sensors 180 (upstream) and 181 (downstream). Once the bleed air pressure meets the requirements, it enters the second cabin environmental control bleed air pipe 161 and supplies air to the aircraft environmental control system through the aircraft cabin environmental control interface 190.

[0053] Then, as Figure 4 As shown, during aircraft takeoff (other than idle): the Bleed Air Monitoring Computer (BMC) 130 controls the closure of the high-pressure shut-off valve 110 (isolating bleed air from the compressor high-pressure bleed air port 140) and bleeds air from the engine compressor intermediate-pressure bleed air port 300.

[0054] Medium-pressure bleed air passes through medium-pressure bleed air pipe 310, then through medium-pressure one-way valve 320, through cabin bleed air pipe 160, and finally into pressure regulating valve 170. From there, it supplies air to the downstream cabin environmental control system via cabin bleed air pipe 160. Similarly, the opening of the pressure regulating valve is adjusted based on the pressure difference between upstream and downstream pressure sensors 180 (upstream) and 181 (downstream) of pressure regulating valve 170, thereby regulating the bleed air pressure at the aircraft cabin environmental control interface 190.

[0055] Combined Figure 2 As shown, the working principle of the system for preventing coking and carbon deposits on fuel nozzles of aviation gas turbine engines according to the present invention is as follows:

[0056] First, after the aircraft stops, the engine stops after landing: the Engine Electronic Controller (EEC) or Engine Health Monitoring Unit (EMU) 30 gives one or more status signals indicating that the engine has stopped. The Bleed Air Monitoring Computer (BMC) 130 obtains the status signal indicating that the engine has stopped from the Engine Electronic Controller (EEC) or Engine Health Monitoring Unit (EMU) 30, and at the same time obtains the aircraft's status information on the ground from other systems of the aircraft (e.g., wheel load signals or wing spoiler opening signals) to confirm that the aircraft is on the ground and the engine has stopped.

[0057] The Bleed Air Monitoring Computer (BMC) 130 first closes and confirms that the high-pressure shut-off valve 110 is closed. It then opens the switch valve 230 and confirms that it is in the open position. The motorized axial flow fan 210 is started; power can be supplied from the aircraft system or the BMC 130. After the axial flow fan starts, air is drawn from the vent (ambient atmosphere) 11 and the first bleed air pipe 10, pressurized, and then sequentially passes through the second bleed air pipe 220, the switch valve 230, the third bleed air pipe 240, the check valve 250, the cooling air duct 20, and the high-pressure bleed air pipe 120. Finally, the cooling airflow reaches the fuel nozzles 400 through the compressor's high-pressure bleed air port 140, cooling all the fuel nozzles.

[0058] When the motor-driven axial flow fan 210 runs for a set time, the bleed air monitoring computer (BMC) 130 shuts off the power to the motor-driven axial flow fan 210 to complete the cooling of the fuel nozzle.

[0059] The operating time of the motor-driven axial flow fan 210 can be set according to different ambient temperatures (e.g., atmospheric temperature sensor from the aircraft or total temperature sensor T12 at the fan bypass inlet), or it can be determined from the engine cross section or exhaust temperature related to the fuel injector temperature obtained from the engine electronic controller (EEC) or engine health monitoring device (EMU) 30 (e.g., airborne sensor T3 at the high pressure compressor outlet total temperature).

[0060] Next, before the aircraft takes off, during ground start-up and taxiing, and during flight: the engine is running, the bleed air monitoring computer (BMC) 130 controls the switch valve 230 to close, and at the same time the one-way valve 250 isolates the connection between the third bleed air pipe 240 and the cooling air pipe 20, the high-pressure bleed air pipe 120 and the compressor high-pressure bleed air port 140.

[0061] In this embodiment, the actuation mechanism of the switching valve 230 can be of a suitable type, such as pneumatic, hydraulic, or electric. The motor of the axial flow fan 210 can be of a DC or AC type, or it can adopt a "blowing" positive pressure mode and a "vacuuming" negative pressure mode.

[0062] According to the above structural description, the system for preventing coking and carbon buildup on fuel nozzles of aviation gas turbine engines automatically cools the fuel nozzles after the aviation gas turbine engine stops, thereby lowering the metal surface temperature of the fuel nozzles below the fuel coking temperature. The system requires no external maintenance work or additional equipment, automatically cooling the fuel nozzles to reduce the temperature below the fuel coking temperature, thus preventing the carbonization of unburned fuel at the fuel nozzles and avoiding coking and carbon buildup.

[0063] The system for preventing coking and carbon buildup on fuel nozzles of aviation gas turbine engines has the following characteristics:

[0064] I. This invention includes an axial flow fan with a motor that provides pressurized air (or suction). The axial flow fan can be DC or AC. The motor can be inside or outside the duct. The centerline of the motor shaft can be perpendicular to the centerline of the duct or at a suitable angle to the centerline of the duct. The power supply can be from an aircraft or ground-based power supply equipment (if applicable). The axial flow fan with a motor can be installed in the duct using a quick-release ring or flange connection.

[0065] Second, the present invention has an upstream connection to the atmosphere and a downstream connection to the engine's original air intake pipe in the axial flow fan with motor, and can draw in external ambient air and deliver cooling air to the vicinity of each fuel nozzle.

[0066] Third, this invention utilizes an ambient temperature sensor subordinate to the aircraft or engine to automatically execute an appropriate cooling time based on the difference between ambient temperature (e.g., the air temperature in different seasons and different temperate zones) and the high-pressure compressor exhaust temperature. A large temperature difference results in a longer fan operating time, while a small temperature difference results in a shorter fan operating time.

[0067] Fourth, the present invention utilizes position sensors on the aircraft and engine, such as wheel-mounted signals, aircraft diffuser opening signals, and aircraft throttle lever angles, to automatically determine and confirm that the engine is in a stopped state. Only after confirming that the engine is in a stopped state can the system for preventing coking and carbon buildup on the fuel injectors be automatically started.

[0068] V. This invention includes a one-way valve and a switching control valve, which can be electrically, pneumatically, or hydraulically operated. This ensures the system for preventing coking and carbon buildup on the fuel nozzles of aviation gas turbine engines operates normally after the engine is stopped, providing cooling air to cool the fuel nozzles. During engine operation, it maintains isolation from other high and low pressure bleed air systems, without affecting the operation of other engine systems.

[0069] VI. The present invention includes a control device, which may be an environmental control bleed air control computer belonging to the aircraft system, or an electronic controller belonging to the engine, or an engine health management device, or a separate system controller.

[0070] 7. After the aviation gas turbine engine completes its flight mission, it does not require additional power, fuel, external equipment and / or logistical support after the engine stops. The system automatically activates the cooling system to reduce the temperature of the fuel nozzles.

[0071] In summary, the system for preventing coking and carbon buildup on fuel nozzles of aviation gas turbine engines, as described in this invention, can prevent coking and carbon buildup on fuel nozzles after the aviation gas turbine engine is shut down. The system automatically operates after aircraft landing and engine shutdown, utilizing the aircraft cabin bleed air and the engine's medium / high pressure bleed air lines. It incorporates additional hardware such as a few bleed air pipes, an axial flow fan with a motor, on / off valves, and check valves to provide cooling air to cool the fuel nozzles. This system solves the technical problem that the surface temperature of the fuel nozzles after shutdown exceeds the coking temperature due to the "thermal wetting" phenomenon, thereby reducing and eliminating fuel nozzle coking and preventing unplanned engine replacements and aviation turbine engine return-to-factory repairs caused by fuel nozzle coking.

[0072] For those skilled in the art, the above disclosure is merely illustrative and does not constitute a limitation of this application. Although not explicitly stated herein, those skilled in the art may make various modifications, improvements, and corrections to this application. Such modifications, improvements, and corrections are suggested in this application and therefore remain within the spirit and scope of the exemplary embodiments of this application.

[0073] Furthermore, this application uses specific terms to describe embodiments of the application. For example, "an embodiment," "one embodiment," and / or "some embodiments" refer to a particular feature, structure, or characteristic related to at least one embodiment of the application. Therefore, it should be emphasized and noted that "an embodiment," "one embodiment," or "an alternative embodiment" mentioned twice or more in different locations in this specification do not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of the application can be appropriately combined.

[0074] Similarly, it should be noted that, in order to simplify the description of the embodiments disclosed in this application and thus aid in the understanding of one or more embodiments of the invention, the foregoing description of the embodiments of this application sometimes combines multiple features into a single embodiment, drawing, or description thereof. However, this disclosure method does not imply that the subject matter of this application requires more features than those mentioned in the claims. In fact, the embodiments have fewer features than all the features of the single embodiments disclosed above.

[0075] While specific embodiments of the present invention have been described above, those skilled in the art should understand that these are merely illustrative examples, and the scope of protection of the present invention is defined by the appended claims. Those skilled in the art can make various changes or modifications to these embodiments without departing from the principles and essence of the present invention, but all such changes and modifications fall within the scope of protection of the present invention.

Claims

1. A system for preventing coking and carbon buildup on fuel nozzles of an aviation gas turbine engine, characterized in that, The system includes an air intake system, an axial flow fan assembly, a first air intake duct, and a cooling duct. The outlet end of the cooling duct is connected to the air intake system, and the inlet end is connected to the axial flow fan assembly. The outlet end of the first air intake pipe is connected to the axial flow fan assembly, and the outlet end is provided with an air vent that communicates with the outside.

2. The system for preventing coking and carbon deposits on fuel nozzles of aviation gas turbine engines as described in claim 1, characterized in that, The axial flow fan assembly includes an axial flow fan with a motor, which is installed between the cooling air duct and the first air intake duct.

3. The system for preventing coking and carbon deposits on fuel nozzles of aviation gas turbine engines as described in claim 2, characterized in that, The axial flow fan assembly also includes a second air intake pipe and a switching valve. One end of the second air intake pipe is connected to the switching valve, and the other end is connected to the axial flow fan. The switching valve is installed between the cooling air duct and the second air intake pipe.

4. The system for preventing coking and carbon deposits on fuel nozzles of aviation gas turbine engines as described in claim 3, characterized in that, The axial flow fan assembly also includes a third air intake pipe and a one-way valve. One end of the third air intake pipe is connected to the switch valve, and the other end is connected to the one-way valve. The one-way valve is installed between the cooling air duct and the third air intake pipe.

5. The system for preventing coking and carbon deposits on fuel nozzles of aviation gas turbine engines as described in claim 1, characterized in that, The system also includes an engine health monitoring device, which is connected to the bleed air monitoring computer of the bleed air system.

6. The system for preventing coking and carbon deposits on fuel nozzles of an aviation gas turbine engine as described in claim 2, characterized in that, The axial flow fan is communicatively connected to the air intake monitoring computer of the air intake system.

7. The system for preventing coking and carbon deposits on fuel nozzles of an aviation gas turbine engine as described in claim 3, characterized in that, The switching valve is communicatively connected to the bleed air monitoring computer of the bleed air system.

8. The system for preventing coking and carbon deposits on fuel nozzles of an aviation gas turbine engine as described in claim 4, characterized in that, The one-way valve is communicatively connected to the bleed air monitoring computer of the bleed air system.

9. The system for preventing coking and carbon deposits on fuel nozzles of an aviation gas turbine engine as described in claim 1, characterized in that, The air intake system includes a high-pressure shut-off valve and a high-pressure air intake pipe. The outlet end of the cooling air duct forms two paths, one of which is connected to the high-pressure shut-off valve and the other of which is connected to the high-pressure air intake pipe.

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