An aircraft engine and a method of cooling an aircraft engine
By installing maintenance access panels and air supply pipes on aircraft engines, external cooling airflow is used to quickly cool the core components of the engine, solving the problems of fuel nozzle coking and low maintenance efficiency caused by engine overheating after shutdown, and achieving low-cost and high-efficiency cooling.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- COMMERCIAL AIRCRAFT CORP OF CHINA LTD
- Filing Date
- 2026-05-28
- Publication Date
- 2026-07-14
AI Technical Summary
In the existing technology, the fuel nozzles of aircraft engines coke due to the warming phenomenon after shutdown, which increases the replacement cost and the maintenance efficiency is low. Existing cooling equipment is expensive and occupies a lot of space.
An aircraft engine comprising a high-pressure stage compressor, a transient venting valve exhaust pipe, a rear core cowl, and a cooling system was designed. By installing a maintenance cover and an air supply pipe on the rear core cowl, the engine's core components are rapidly cooled using external cooling airflow, preventing high-pressure stage airflow leakage. The design is simple and safe.
It achieves rapid and low-cost engine cooling, reduces the risk of fuel injector coking, improves the efficiency of engine maintenance after shutdown, and reduces operating and maintenance costs.
Smart Images
Figure CN122383504A_ABST
Abstract
Description
Technical Field
[0001] This application relates to the field of aircraft structural design, and particularly to the design of aircraft engines, specifically to the cooling system of aircraft engines. Background Technology
[0002] In aircraft, such as civil aircraft, the engine is a crucial component. The core components of an engine include the compressor, combustion chamber, and turbine. When the engine is running, the high-energy airflow from the combustion chamber drives the turbine, converting the energy of the airflow into rotational kinetic energy, which in turn powers the compressor, thus maintaining engine operation. The operation of the engine combustion chamber is critical and vital to the overall operation of the engine. After the aircraft lands or after a day of flight operations, the engine must be shut down.
[0003] After the engine is shut down, there will be no significant inflow of outside air into the aforementioned core components. Therefore, these components will experience a noticeable warm-up period after shutdown. On the other hand, the engine's fuel injectors spray atomized aviation fuel for combustion during engine operation. Inevitably, some of this fuel will remain on the fuel injectors. The warm-up of the engine after shutdown causes the fuel injector temperature to rise above the fuel coking threshold again shortly after shutdown. This leads to accelerated coking of the fuel injectors, necessitating frequent fuel injector replacements and increasing aircraft operating costs.
[0004] On the other hand, during aircraft operation, if an engine malfunctions and urgent troubleshooting work such as borescope inspection is required, the work must be carried out after the engine has cooled down. However, as mentioned above, the relevant components in the engine will continue to heat up after the engine is shut down, requiring several hours for them to cool completely. Therefore, the warm-up process after the engine is shut down can also affect the efficiency of engine maintenance.
[0005] Currently, the common method for quickly cooling engines is to use a dedicated cooling vehicle to cool the engine when it is parked. However, these dedicated cooling vehicles are expensive and require a large amount of space.
[0006] Therefore, in the field of aircraft design, there is a need for improved engine cooling systems that are low-cost and easy to operate. Summary of the Invention
[0007] This application is made to address the problems existing in the prior art described above. The purpose of this application is to provide an aircraft engine with a cooling system and a related cooling method, which has low manufacturing and operating costs and can effectively and rapidly cool the engine after shutdown.
[0008] This application discloses an aircraft engine, comprising: a high-pressure stage compressor; a transient bleed valve exhaust pipe, on which a transient bleed valve is disposed; and a rear core cowling. The aircraft engine further includes a cooling system comprising: a maintenance access cover located on the rear core cowling; an air supply pipe, one end of which is located at the maintenance access cover, and the other end connected to the transient bleed valve exhaust pipe; and a cooling airflow source, which can be connected to one end of the air supply pipe for delivering cooling airflow to the transient bleed valve exhaust pipe.
[0009] In this aircraft engine structure, cooling airflow is input to the vicinity of core engine components, such as fuel nozzles, through a transient bleed valve exhaust pipe. This allows for efficient delivery of cooling airflow and is structurally simple. Furthermore, during operation, only the maintenance hatch on the rear core cowl needs to be opened, and the cooling airflow source needs to be connected to the air supply pipe. This avoids contact with the engine's heat-generating components, thus offering high safety and convenience.
[0010] In one example, the cooling airflow source includes a fan and a fan duct connected to the fan at one end, wherein the other end of the fan duct can be connected to one end of the air supply pipe at a maintenance cover.
[0011] Preferably, a clamping device is provided at one end of the gas supply pipe. This clamping device facilitates the connection of the gas supply pipe to a cooling airflow source, such as a fan duct for a blower.
[0012] Preferably, a one-way valve is provided on the gas supply pipe. The one-way valve is configured to allow cooling airflow from the maintenance cover to the transient vent valve exhaust pipe, but prevents the airflow from flowing back to the maintenance cover in the opposite direction. By providing this one-way valve, high-pressure stage airflow in the high-pressure stage compressor of the aircraft engine can be prevented from flowing out of the gas supply pipe, thereby improving the overall airtightness of the aircraft engine.
[0013] Preferably, a control valve is installed downstream of the connection point between the gas supply pipe and the transient vent valve exhaust pipe. This control valve can close before supplying cooling gas flow to the transient vent valve exhaust pipe, thereby preventing cooling gas flow from escaping from the transient vent valve outlet of the transient vent valve exhaust pipe. In other words, this structure helps to improve the utilization efficiency of the cooling gas flow.
[0014] Preferably, a pressure relief door is provided on the aft core cowling, wherein the maintenance access cover and the pressure relief door are integrally formed. This integrated design of the maintenance access cover and the pressure relief door helps to reduce the structural complexity of the aircraft engine.
[0015] A method for cooling an aircraft engine as described above is also provided. This method includes the following steps: After the aircraft engine stops, open the maintenance access cover located on the rear core cowling; The cooling airflow source is connected to the air supply pipe via the maintenance cover; and Turn on the cooling airflow source and supply the cooling airflow to the exhaust pipe of the transient vent valve via the air supply pipe.
[0016] Preferably, for a preferred configuration in which a control valve is provided downstream of the connection point between the gas supply pipe and the transient vent valve exhaust pipe, the method may further include the step of closing the control valve before turning on the cooling airflow source. Attached Figure Description
[0017] Preferred embodiments of the present invention are shown in the accompanying drawings, from which the specific implementation of the invention can be more clearly understood, wherein: Figure 1 A schematic structural diagram of the aircraft engine of this application is shown.
[0018] (Symbol Explanation)
[0019] 10 Aircraft engine; 11 High-pressure stage compressor; 12 Tail nozzle; 13 Transient bleed valve; 14 Transient bleed valve exhaust pipe; 15 Transient bleed valve outlet; 16 Control valve; 17 Aft core cowling; 20 Cooling system; 21 Maintenance port cover; 22 Air supply pipe; 23 Check valve; 24 Fan; 25 Fan piping. Detailed Implementation
[0020] The specific embodiments of this application will now be described in detail with reference to the accompanying drawings. It should be understood that the embodiments shown in the drawings are merely preferred embodiments of this application and do not constitute a limitation on the scope of this application. Those skilled in the art can make various obvious modifications, variations, and equivalent substitutions to this application based on the embodiments shown in the drawings. Furthermore, without contradiction, the technical features in the different embodiments described below can be arbitrarily combined with each other, all of which fall within the protection scope of this application.
[0021] Figure 1The diagram shows a schematic structural diagram of an aircraft engine 10 according to this application. The aircraft engine 10 includes a high-pressure stage compressor 11 and a tail nozzle 12. A transient bleed valve 13 is disposed between the high-pressure stage compressor 11 and the tail nozzle 12. The transient bleed valve 13 is mounted on a transient bleed valve exhaust pipe 14 extending from the high-pressure stage compressor 11. The transient bleed valve exhaust pipe 14 may extend from the high-pressure stage compressor 11 to the vicinity of the tail nozzle 12, for example, and form a transient bleed valve outlet 15 at its end adjacent to the tail nozzle 12.
[0022] When the aircraft engine 10 is started, the transient vent valve 13 can be opened to quickly release pressure, thereby achieving pressure balance of the aircraft engine 10. After the aircraft engine 10 is started and running stably, the transient vent valve 13 is closed, and the transient vent valve 13 remains closed throughout the normal operation of the aircraft engine 10.
[0023] When a low surge margin is detected in the aircraft engine 10, for example, when the detected engine exhaust temperature (EGT) is too high, the aircraft's control system will open the transient bleed valve 13 to perform anti-surge operation. In addition, the transient bleed valve 13 remains open after the aircraft lands and the aircraft engine 10 shuts down.
[0024] In the scheme of this application, in order to cool the core of the aircraft engine 10 after the aircraft engine 10 stops, a cooling system 20 is further provided on the aircraft engine 10. For example... Figure 1 As shown, the cooling system 20 includes a maintenance access cover 21 on the rear core cowling 17 of the aircraft engine 10, and an additional air supply pipe 22 connected to the transient vent valve exhaust pipe 14. One end of the air supply pipe 22 is connected to the transient vent valve exhaust pipe 14, and the other end is located at the maintenance access cover 21, where it can be connected to the fan duct 25 of the fan 24. Preferably, a locking device is provided at the end of the air supply pipe 22 located at the maintenance access cover 21 to reliably connect the fan duct 25 to the air supply pipe 22 and substantially fix it in the position of the maintenance access cover 21.
[0025] The cooling system 20 also includes a fan 24 and fan ducts 25 connected to the fan 24. After the aircraft lands and the aircraft engines 10 shut down, ground operators can open the maintenance access cover 21 on the aft core cowl 17 and connect the fan ducts 25 of the fan 24 to the maintenance access cover 21. Afterward, ground operators can turn on the fan 24 to blow cooling airflow from the external environment into the core area of the aircraft engine 10 through the transient vent valve exhaust pipe 14. Preferably, the maintenance access cover 21 is located on top of the aircraft engine 10, allowing cooling gas to be blown into the core area of the aircraft engine 10 from the top, thereby further improving the cooling efficiency of the core components of the aircraft engine 10.
[0026] Preferably, a one-way valve 23 is provided on the gas supply pipe 22. This one-way valve 23 is configured to allow cooling airflow to flow from the maintenance cover 21 into the transient vent valve exhaust pipe 14, but prevents airflow from flowing out in the opposite direction. This structure ensures that the high-pressure stage airflow from the high-pressure stage compressor 11 does not leak from the gas supply pipe 22 due to the addition of the gas supply pipe 22.
[0027] Preferably, a control valve 16 is provided downstream of the connection point between the gas supply pipe 22 and the transient vent valve exhaust pipe 14. When cooling components in the core area of the aircraft engine 10, the control valve 16 closes to prevent the cooling airflow entering the transient vent valve exhaust pipe 14 from the gas supply pipe 22 from flowing out of the transient vent valve outlet 15. This improves the utilization rate of the cooling airflow and ensures more thorough cooling of the core components of the aircraft engine 10.
[0028] Furthermore, in this application, the maintenance cover 21 is preferably integrally formed with the pressure relief door provided on the aft core cowl 17. In other words, the maintenance cover 21 can utilize the existing pressure relief door structure on the aft core cowl 17, thereby helping to reduce the structural complexity of the aircraft engine 10 system.
[0029] The following describes the cooling method for the aircraft engine 10 after it has been shut down, specifically the core components of the aircraft engine 10.
[0030] After the aircraft lands and the aircraft engine 10 shuts down, the operator can open the maintenance access cover 21 located on the rear core cowling 17 of the aircraft engine 10. Then, the fan duct 25 connected to the fan 24 is connected to the air supply pipe 22 via the maintenance access cover 21.
[0031] Next, the fan 24 is turned on, and the cooling airflow is blown into the transient vent valve exhaust pipe 14 through the fan pipe 25 and the air supply pipe 22. The cooling airflow flows through the transient vent valve exhaust pipe 14 to the core components in the core area of the aircraft engine 10 and the fuel nozzles, thereby effectively cooling the core components of the aircraft engine 10 and preventing coking on the fuel nozzles.
[0032] Preferably, when a control valve 16 is provided on the transient vent valve exhaust pipe 14, the control valve 16 is closed before the fan 24 is turned on to blow cooling airflow into the transient vent valve exhaust pipe 14. This prevents or at least significantly reduces the amount of cooling airflow flowing out through the transient vent valve outlet 15.
[0033] As can be seen, in this application, an air supply pipe 22 is added and connected to the transient bleed valve exhaust pipe 14. Cooling airflow is delivered to the core area of the aircraft engine 10 through the air supply pipe 22 via the transient bleed valve exhaust pipe 14. Thus, a simple structure can be used to continuously deliver cooling airflow to the core components of the aircraft engine 10 to reduce temperature, and ensure that the temperature of the fuel nozzle is lower than the fuel coking temperature, thereby slowing down or even avoiding coking on the fuel nozzle, which helps to extend the service life of the aircraft engine 10 and reduce the operating cost of the aircraft and the maintenance cost of the engine.
[0034] Moreover, as can be seen from the above disclosure, the cooling part of the aircraft engine with cooling system in this application has a simple structure, occupies little space, and has the characteristics of speed and mobility, thus enabling the cooling of the core area of the aircraft engine to be achieved quickly at a lower cost.
[0035] The preferred embodiments of the engine cooling system and related cooling methods of this application have been described in detail above. Those skilled in the art can make various obvious variations and modifications based on the above disclosure, and these are still within the scope of this application.
[0036] For example, as described above, valve structures such as one-way valve 23 and control valve 16 are preferred structures. Even without these structures, the cooling system 20 of the aircraft engine 10 of this application can still basically achieve its purpose of cooling the core components in the core area of the aircraft engine 10.
[0037] For example, the above disclosure describes a device for supplying cooling airflow to the aircraft engine 10 using fan 24 as an example. However, in addition to fan 24 and fan duct 25 connected thereto, other forms of cooling airflow sources can also be used to supply cooling airflow to the aircraft engine 10 through air duct 22, and these are also within the scope of this application.
Claims
1. An aircraft engine, the aircraft engine comprising: High-pressure stage compressor; Transient venting valve exhaust pipe, wherein a transient venting valve is provided on the transient venting valve exhaust pipe; And a rear core shield; characterized in that the aircraft engine further includes a cooling system, the cooling system comprising: A maintenance cover is provided on the rear core cover; An air supply pipe, one end of which is located at the maintenance port cover, and the other end of which is connected to the exhaust pipe of the transient vent valve; and A cooling airflow source is provided, which can be connected to one end of the air supply pipe to deliver cooling airflow to the exhaust pipe of the transient vent valve.
2. The aircraft engine as described in claim 1, characterized in that, The cooling airflow source includes a fan and a fan duct connected to the fan at one end, wherein the other end of the fan duct can be connected to the end of the air supply pipe at the maintenance cover.
3. The aircraft engine as described in claim 1, characterized in that, A clamping device is provided on one end of the gas pipeline.
4. The aircraft engine as described in claim 1, characterized in that, A one-way valve is provided on the gas supply pipe. The one-way valve is configured to allow cooling gas flow from the maintenance port cover to the transient vent valve exhaust pipe, but prevents the gas flow from flowing back to the maintenance port cover in the opposite direction.
5. The aircraft engine as described in claim 1, characterized in that, A control valve is provided downstream of the connection point between the gas supply pipe and the transient vent valve exhaust pipe.
6. The aircraft engine as claimed in claim 1, characterized in that, A pressure relief door is provided on the rear core cover, wherein the maintenance cover is integrally formed with the pressure relief door.
7. A method for cooling an aircraft engine as described in claim 1, characterized in that, The method includes the following steps: After the aircraft engine stops, open the maintenance access cover located on the aft core cowling; The cooling airflow source is connected to the air supply pipe via the maintenance cover; and Turn on the cooling airflow source and supply the cooling airflow to the exhaust pipe of the transient vent valve via the air supply pipe.
8. The method as described in claim 7, characterized in that, A control valve is provided downstream of the connection point between the gas supply pipe and the transient vent valve exhaust pipe, and the method further includes: Close the control valve before turning on the cooling airflow source.