Blade rapid cooling air entraining control method and air entraining system

Abstract

The rapid cooling air entraining control method comprises the following steps that S1, a full-authority digital engine controller is used for judging the state and the flight height of an aircraft engine in real time, and whether the risk of rotor locking occurs or not is judged; s2, when the risk of rotor locking exists, an air entraining system is started, and an engine hot end component is cooled; s3, after the full-authority digital engine controller recognizes that the airplane enters the air starting envelope, the full-authority digital engine controller starts an air starting program and recognizes whether the engine is in a slow state or not; and S4, when the air start is successful and the engine is in a slow state, closing the air entraining system, and when the air start of the engine is not successful, repeating the steps S2-S3. The invention further provides a rapid cooling air entraining system. The hot end component is effectively cooled under the condition of high-altitude flameout of the engine, and the engine is prevented from entering a rotor locking state.

Description

Technical Field

[0001] The present invention relates to the technical field of aero-engine cooling, and particularly to the field of bleed air control systems. Background Art

[0002] During the high-altitude flight of an aircraft, if the engine accidentally stalls or is forced to stop for other reasons, in order to ensure the flight power and safe operation of the aircraft, an in-flight engine restart is required. However, since the aircraft is above the in-flight start envelope at this time, a direct restart is generally not possible. To be able to restart, the aircraft needs to glide down to the start envelope while keeping the flight speed within the start range.

[0003] During the glide of the aircraft, the engine is in the windmilling state, the core engine speed drops rapidly, and there is still a large amount of residual heat inside. Due to the different materials and heat transfer conditions of the rotating and stationary components, the contraction rates of the rotor and stator components are also different. Since the internal flow rate is relatively small after the engine stalls, the heat dissipation of the rotating components is slow, and the thermal expansion deformation persists, while the stationary components, such as the casing and stator, will cool and contract relatively quickly, reducing the clearance between the rotating and stationary components, and even causing interference such as frictional contact and jamming, which generally occurs at the tip of the high-pressure turbine blade and the seal. Once rotor locking occurs, the engine cannot be started when the aircraft glides down to the start envelope, which will endanger the flight safety of the aircraft.

[0004] Since the engine is in the stopped state at this time, the air system cannot bleed air from the high-pressure compressor to cool the rotating components such as turbine blades and seals. At the same time, since the contraction rate of the turbine casing is greater than that of the blades at this time, resulting in a reduced clearance, the active clearance control system cannot function either. To reduce the harm caused by rotor locking in the engine and ensure the normal start of the engine, a cooling method capable of cooling the blades is required to accelerate the contraction rate of the hot-end components such as the rotor blades and maintain a reasonable tip clearance size. Summary of the Invention

[0005] An object of the present invention is to provide a rapid cooling bleed air control method that can effectively cool the hot-end components in the event of an engine stall at high altitude.

[0006] The quick-cooling bleed air control method for achieving the above object includes the following steps: S1. Use the full-authority digital engine controller to continuously judge the state of the aircraft engine and the flight altitude, and judge whether there is a risk of rotor lock; S2. When there is a risk of rotor lock, turn on the bleed air system to cool the hot-end components of the engine; S3. When the full-authority digital engine controller recognizes that the aircraft enters the in-flight start envelope, the full-authority digital engine controller starts the in-flight start program and recognizes whether the engine is in the idle state; S4. When the in-flight start is successful and the engine is in the idle state, turn off the bleed air system. When the in-flight start of the engine is unsuccessful, repeat steps S2-S3.

[0007] In one or more embodiments, in step S1, when it is judged that the aircraft engine is in the flameout state and the aircraft flight altitude is at least 10,000 ft higher than the upper line of the in-flight start envelope, it is determined that there is a risk of rotor lock.

[0008] In one or more embodiments, part of the air flow that enters the main flow path from the last-stage collector chamber of the high-pressure compressor is made to flow towards the combustion chamber, and the other part of the air flow is made to flow towards the rotor blades.

[0009] In one or more embodiments, the gas in the other part of the air flow path is pressurized and then impinged on the rotor blades.

[0010] Another object of the present invention is to provide a quick-cooling bleed air system for performing the above method, including a bleed air pipeline. The starting end of the bleed air pipeline is connected to the fan bypass, and the tail end includes multiple branches. Part of the branches are used to cool the hot-end components in the radially inward direction, and the other part of the branches are used to cool the hot-end components in the radially outward direction.

[0011] In one or more embodiments, the system further includes a diffuser interface provided at the tail end of the branch for forming a jet air flow.

[0012] In one or more embodiments, the system further includes a main pipeline valve and branch pipeline valves provided on each branch.

[0013] In one or more embodiments, the system further includes a full-authority digital engine controller, and the main pipeline valve and the branch pipeline valves are bidirectionally signal-connected to the full-authority digital engine controller.

[0014] In one or more embodiments, the system further includes a check valve provided on the branch.

[0015] In one or more embodiments, the hot-end components further include a seal and a casing.

[0016] The above-mentioned bleed air system and control method for rapid cooling of engine blades can cool the hot-end components, including turbine blades and seals, in the event of an engine flameout at high altitude, avoid rotor locking during a descent, which may affect the engine restart, reduce or even avoid the harm of rotor locking, and ensure flight safety. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] The above and other features, properties, and advantages of the present invention will become more apparent from the following description in conjunction with the drawings and embodiments, where:

[0018] Figure 1 is a schematic diagram of the gas flow path of the bleed air control system;

[0019] Figure 2 is a schematic diagram of the bleed air system pipeline;

[0020] Figure 3 is the control logic diagram of the bleed air system;

[0021] Figure 4 is the flowchart of the rapid cooling bleed air control method.

[0022] DESCRIPTION OF SYMBOLIC MARKINGS

[0023] 1 Bleed air device

[0024] 2 Main pipe valve

[0025] 3 Transmission pipeline

[0026] 4 Second branch pipe valve

[0027] 5 Branch pipe

[0028] 6 Check valve

[0029] 7 Diffuser interface

[0030] 8 Third branch pipe valve

[0031] 9 ACC component

[0032] 10 First branch pipe valve

[0033] 11 Interface component

[0034] 100 Turbine engine

[0035] 110 Fan group

[0036] 120 Booster stage component

[0037] 130 Turbine component

[0038] 140 Inlet duct component

[0039] 150 High-pressure compressor assembly

[0040] 160 Combustion chamber assembly Detailed implementation manners

[0041] The present invention will be further described below in conjunction with specific embodiments and the accompanying drawings. More details are set forth in the following description to facilitate a full understanding of the present invention. However, the present invention is clearly capable of being implemented in many other different ways than described herein. Those skilled in the art can make similar generalizations and deductions according to the actual application situation without departing from the connotation of the present invention. Therefore, the protection scope of the present invention should not be limited by the content of this specific embodiment.

[0042] It should be noted that these and subsequent other drawings are only examples, and they are not drawn under the condition of equal proportion, and should not be used to limit the actual required protection scope of the present invention.

[0043] During the high-altitude flight of an aircraft, after the engine stalls or shuts down, in order to be able to restart, the aircraft needs to glide to the start envelope. Once rotor lock occurs, the engine cannot start when the aircraft glides to the start envelope, endangering the flight safety of the aircraft.

[0044] Rotor lock refers to the situation where the core rotor speed drops to zero after the engine shuts down or stalls in the air, and it still cannot rotate under windmilling or assisted starting. When studying rotor lock, rotor drag is usually considered together. Rotor drag is the phenomenon that the rotor speed decreases under windmilling conditions due to the connection load of engine accessories and the high friction between the rotating and stationary components.

[0045] Therefore, in the case where there is a risk of rotor lock when the engine stalls at high altitude, by accelerating the heat dissipation of hot-end rotating components such as turbine blades and seals, and maintaining the clearance between the rotating and stationary components to avoid jamming, the normal in-air start of the engine can be ensured.

[0046] The present invention provides an air extraction control system for rapid cooling of engine blades, which can promote the cooling of hot-end components in the case of high-altitude engine stall and avoid the occurrence of rotor lock during the gliding process.

[0047] Figure 1 and Figure 2 FIG. shows a schematic diagram of the gas flow path of the air extraction control system for rapid cooling of engine blades. The turbofan engine 100 includes an inlet duct assembly 140, a fan assembly 110, a booster stage assembly 120, a high-pressure compressor assembly 150, a combustion chamber assembly 160, a turbine assembly 130 including a high-pressure turbine assembly and a low-pressure turbine assembly, and an exhaust nozzle assembly 170.

[0048] When the engine is running, air G enters the engine through the intake duct and flows through the fan assembly 110 towards the booster stage assembly 120. The compressed air enters the high-pressure compressor assembly from the booster assembly, is further compressed and then flows towards the combustion chamber assembly, where it mixes with fuel. The mixture burns in the combustion chamber assembly to form high-temperature and high-pressure gases. The gases pass through the high-pressure turbine assembly and the low-pressure turbine assembly in sequence, and finally are discharged through the tail nozzle assembly 170.

[0049] The engine needs to cool the turbine blades. Generally, the cooling of the blades is supplied by the air system and the cooling air of the turbine cooling blades is controlled to ensure that the blade cold air inlet has the required pressure and temperature. The solid line A represents the flow path of the active clearance control system (ACC). The active clearance control system is a system that actively adjusts the clearance between the turbine blades and the casing. By means of the cooling pipelines arranged outside the high-pressure turbine and low-pressure turbine casings, bleed air from the bypass duct is used to cool the high-pressure / low-pressure turbine casings, thereby achieving the purpose of adjusting the tip clearance of the moving blades and improving the engine efficiency and performance.

[0050] However, during the aircraft's descent glide, the engine is in the windmilling state, and there is a problem that the heat dissipation speeds of the rotor and stator components are different, and it is easy to cause the situation of rotor locking, which affects the restart of the engine. Therefore, the rapid cooling bleed air control system and method described in the present invention can increase the heat dissipation effect on the hot-end components of the engine, promote the rapid heat dissipation of the hot-end components, and achieve the purpose of accelerating cooling and reducing the expansion deformation of the hot-end rotating components.

[0051] The schematic flow path diagram of the bleed air pipeline described in the present invention is shown by the dashed line B, and it includes a bleed air pipeline. The starting end of the bleed air pipeline is connected to the bypass duct of the fan, and the tail end includes multiple branches. One part of the branches is used to cool the hot-end components in the radially inward direction, and the other part of the branches is used to cool the hot-end components in the radially outward direction.

[0052] Preferably, the bleed air pipeline, as an air extraction and transmission system, shares the air extraction device and part of the pipeline with the active clearance control system (ACC), as Figure 1 shown. The gas in line A is used for the active clearance control system, and line B is used as the flow path of the bleed air control system of the present invention. The high-pressure turbine cooling air system shares the diversion interface and pipeline, introduces cooling air from the bypass duct of the fan, controls the bleed air flow through a valve, and shunts the cooling gas through the pipeline.

[0053] Preferably, part of the air path that enters the main flow path from the last-stage plenum chamber of the high-pressure compressor is made to flow towards the combustion chamber, and the other part of the air path is made to flow towards the rotor blades.

[0054] For example, a part of the gas flows from the diffuser to the combustion chamber from the gas collection chamber at the end stage of the high-pressure compressor, taking away part of the heat. After a part of the gas is pressurized by the diffuser, it enters the main flow path to impact and strengthen the heat transfer and cooling of the rotor blades such as turbine blades, so as to accelerate the cooling and reduce the expansion deformation of the hot-end rotating components, thereby reducing the risk of jamming caused by clearance loss.

[0055] By sharing the air extraction device and part of the pipeline with the active clearance control system, and sharing the diversion interface and pipeline with the high-pressure turbine cooling air system, the rotor blades such as turbine blades are cooled by extracting air from the fan outer duct, avoiding excessive design changes that increase the engine weight.

[0056] Specifically, as Figure 2 shown, it includes an air extraction device 1, a main pipe valve 2, a transmission pipeline 3, a second branch pipe valve 4, a branch pipe 5, a check valve 6, a diffuser interface 7, a third branch pipe valve 8, an ACC assembly 9, a first branch pipe valve 10, and an interface assembly 11. Among them, the air extraction device 1 and the transmission pipeline 3 form an air extraction pipeline, and the three flow paths where the first branch pipe valve 10, the branch pipe 5, and the third branch pipe valve 8 are located form three branch paths.

[0057] Cooling air is introduced from the fan outer duct through the air extraction device 1, the main pipe valve 2 controls the air extraction flow rate, and the cooling gas is shunted through the transmission pipeline 3.

[0058] A part of the gas enters the ACC assembly 9 through the third branch pipe valve 8 to cool the turbine casing; a part of the gas enters the diffuser interface 6 through the branch pipe 5 via the second branch pipe valve 4, and the diffuser interface 6 pressurizes the cooling gas to perform impact strengthening heat transfer and cooling on the turbine blades. The check valve 6 is used to prevent backflow caused by excessive pressure at the turbine and other places during normal engine operation; a part of the gas enters the main flow path from the gas collection chamber at the end stage of the high-pressure compressor through the interface assembly 11 via the first branch pipe valve 10, flows from the diffuser to the combustion chamber, taking away part of the heat. The interface assembly 11 can prevent backflow caused by excessive internal pressure of the engine during normal engine operation.

[0059] Each valve is bidirectionally signal-connected to the Full Authority Digital Engine Control (FADEC), and its opening and closing are controlled by the FADEC system.

[0060] Figure 2 Only three branch paths are shown. Those skilled in the art can understand that the number of branch paths can be determined according to the specific parts to be cooled.

[0061] Combined with the introduction of the above rapid cooling air extraction system, the present invention also provides a rapid cooling air extraction control method, referring to Figure 3 andFigure 4 and the following steps for understanding.

[0062] S1. Use the Full Authority Digital Engine Controller (FADEC) to continuously monitor the aircraft engine status and flight altitude in real time, and determine whether there is a risk of rotor lock. When the aircraft engine is in the flameout state and the flight altitude of the aircraft is far above the upper limit of the in-flight starting envelope (in-flight start boundary), it is determined that there is a risk of rotor lock before the engine starts in the air. For example, in some embodiments, the flight altitude of the aircraft is at least 10,000 ft higher than the in-flight starting envelope to determine the risk of rotor lock.

[0063] S2. When there is a risk of rotor lock, activate the above-mentioned bleed air system to cool the hot section components of the engine. Optionally, open the branch line to cool the hot section components that need to be cooled. The FADEC system activates the bleed air system, controls the bleed air pipeline valves, opens the main pipeline valve 2, the second branch pipeline valve 4, and the first branch pipeline valve 10, closes the third branch pipeline valve 8, and introduces cooling air to cool the hot section rotating components.

[0064] S3. When the Full Authority Digital Engine Controller (FADEC) identifies that the aircraft has entered the in-flight starting envelope, the FADEC starts the in-flight starting procedure and identifies whether the engine is in the idle state.

[0065] S4. When the in-flight start is successful and the engine is in the idle state, turn off the bleed air system. When the in-flight start of the engine is unsuccessful, repeat steps S2 - S3, that is, use the FADEC system to control the above-mentioned rapid cooling bleed air system to cool the hot section components multiple times until the engine can start in the air to the idle state, and then close all the bleed air system valves.

[0066] Thus, based on the high-altitude engine shutdown condition, the above method controls the bleed air system through the FADEC system to cool the hot section rotating components, accelerate the heat dissipation of the hot section components, maintain the clearance between the rotating and stationary components, reduce or even avoid the harm of rotor lock, and ensure flight safety.

[0067] This application uses specific terms to describe the embodiments of this application. For example, "one embodiment", "an embodiment", and / or "some embodiments" mean a certain feature, structure, or characteristic related to at least one embodiment of this application. Therefore, it should be emphasized and noted that the "one embodiment" or "an embodiment" or "an alternative embodiment" mentioned twice or more at different positions in this specification does not necessarily refer to the same embodiment. In addition, certain features, structures, or characteristics in one or more embodiments of this application can be appropriately combined.

[0068] Although the present invention is disclosed above in preferred embodiments, it is not intended to limit the present invention. Any person skilled in the art can make possible changes and modifications without departing from the spirit and scope of the present invention. Therefore, all modifications, equivalent changes and decorations made to the above embodiments based on the technical essence of the present invention without departing from the technical solution of the present invention fall within the protection scope defined by the claims of the present invention.

Claims

1. Quick cooling bleed air control method, Characterized in that, It includes the following steps: S1. Use the full-authority digital engine controller to continuously judge the state of the aircraft engine and the flight altitude, and judge whether there is a risk of rotor lock; S2. When there is a risk of rotor lock, turn on the bleed air system to cool the hot section components of the engine; S3. When the full-authority digital engine controller recognizes that the aircraft enters the in-air start envelope, the full-authority digital engine controller starts the in-air start procedure and recognizes whether the engine is in the idle state; S4. When the in-air start is successful and the engine is in the idle state, turn off the bleed air system. When the in-air start of the engine is unsuccessful, repeat steps S2 - S3.

2. The quick cooling bleed air control method according to claim 1, Characterized in that, In step S1, when it is judged that the aircraft engine is in the flameout state and the aircraft flight altitude is at least 10,000 ft higher than the upper line of the in-air start envelope, it is determined that there is a risk of rotor lock.

3. The quick cooling bleed air control method according to claim 1, Characterized in that, Make a part of the air flow that enters the main flow path from the last-stage collector cavity of the high-pressure compressor flow to the combustion chamber, and make another part of the air flow flow to the rotor blades.

4. The quick cooling bleed air control method according to claim 3, Characterized in that, Make the gas in the other part of the air flow be pressurized and then impinge on the rotor blades.

5. Quick cooling bleed air system, Characterized in that, Used to perform the method according to any one of claims 1 - 4, including: The bleed air pipeline, the starting end is connected to the fan outer duct, and the tail end includes multiple branches, where a part of the branches is used to cool the hot section components in the radially inward direction, and another part of the branches is used to cool the hot section components in the radially outward direction.

6. The quick cooling bleed air system according to claim 5, Characterized in that, The system further includes a diffuser interface arranged at the tail end of the branch for forming a jet air flow.

7. The quick cooling bleed air system according to claim 6, Characterized in that, The system further includes a main pipe valve and branch pipe valves arranged on each branch.

8. The quick cooling bleed air system according to claim 7, Characterized in that, The system further includes a full-authority digital engine controller, and the main pipe valve and the branch pipe valves are bidirectionally signal-connected to the full-authority digital engine controller.

9. The quick cooling bleed air system according to claim 5, Characterized in that, The system further includes a check valve arranged on the branch.

10. The quick cooling bleed air system according to claim 5, Characterized in that, The hot section components further include a seal and a casing.

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