Turbine disk rim temperature control system and method

Abstract

A turbine rim temperature control system can regulate cooling flow during different conditions of engine operation. The turbine disc edge temperature control method can save the air entrainment amount. The turbine disc edge temperature control system comprises an air guiding device used for guiding in cold air. The air entraining pipeline is used for receiving cold air output by the air entraining device; the circumferential distribution header pipe is communicated with the tail end of the air entraining pipeline and surrounds the peripheral side of the turbine casing; the turbine disc edge temperature control system further comprises a bypass pipeline and a valve device, the bypass pipeline is led out of the air entraining pipeline and communicated with at least one cooling pipeline, and the valve device is connected with the bypass pipeline and communicated with the air entraining pipeline. The cooling pipeline is arranged around the turbine outer casing; and the valve device is used for controlling the on-off of the bypass pipeline.

Description

Technical Field

[0001] The invention relates to a turbine disk rim temperature control system and method for a turbine. Background Art

[0002] The aircraft engine turbine is located behind the combustion chamber and is a hot end component. The high-temperature combustion gas drives the turbine to work and drives the compressor. As the pressure ratio and thrust indicators of modern civilian high-bypass turbofan engines continue to increase, the temperature before the turbine increases, which poses challenges to the turbine structure, life, cooling, sealing and material processing. The edge of the turbine disk close to the mainstream is affected by the heat conduction of the turbine blades and the mainstream ambient temperature, and the temperature level is relatively high. As the temperature of the turbine disk increases, the mechanical properties of the material drop sharply, which has an adverse effect on the strength and life of the turbine disk. The disk edge needs to be cooled and sealed to reduce or prevent the intrusion of gas into the turbine disk.

[0003] A common method of cooling and sealing turbine discs is to cool the discs by drawing cooling air from the compressor. The amount of cooling air used and the sealing effect are related to the temperature drop requirement and the gap of the sealing structure. The temperature drop requirement is related to the material performance and the inlet temperature of the turbine. For mature high-temperature alloy materials that can be used for aircraft engine turbine discs, secondary flow needs to be drawn for cooling and sealing in order to withstand the higher inlet temperature of modern civilian high-bypass turbofan engines. If the gap of the sealing structure is too large, more cooling air needs to be drawn to prevent a large amount of fuel gas from invading. Excessive use of secondary flow will cause loss of engine thrust. If the gap is too small, it is easy to cause a risk of friction during engine operation due to the different deformation responses of the rotor and stator, resulting in safety accidents. Summary of the invention

[0004] The object of the present invention is to provide a turbine disk rim temperature control system which can adjust the cooling flow rate in different operating states of the engine.

[0005] Another object of the present invention is to provide a turbine disk rim temperature control method, which adjusts the cooling flow rate in different engine operation states to save the amount of bleed air.

[0006] According to one aspect of the present invention, a turbine rim temperature control system includes an air bleed device for introducing cold air; an air bleed pipeline for receiving cold air output by the air bleed device; a circumferential distribution manifold connected to the end of the air bleed pipeline and surrounding the outer peripheral side of the turbine casing; and a plurality of manifolds connected to the circumferential distribution manifold, axially distributed on the outer peripheral side of the turbine casing, and used for impact cooling the turbine casing. The turbine rim temperature control system also includes a bypass pipeline and a valve device, wherein the bypass pipeline is led out from the air bleed pipeline and connected to at least one cooling pipeline, and the cooling pipeline is arranged around the outer casing of the turbine; the valve device is used to control the on and off of the bypass pipeline.

[0007] In one embodiment, the bypass pipeline has a plurality of outlets, and the plurality of outlets are respectively connected to a plurality of cooling pipelines.

[0008] In one embodiment, the plurality of cooling lines connected to the plurality of outlets are located on an upstream side of the turbine casing.

[0009] In one embodiment, the cooling line is selected from a plurality of the manifolds or is a line independent of the manifold.

[0010] In one embodiment, the valve device includes a valve and an actuator for driving the valve, and the actuator controls the opening and closing of the valve according to the working state of the engine or the aircraft.

[0011] According to another aspect of the present invention, a method for controlling the temperature of a turbine disk rim is provided, in which cold air is introduced through an air bleed line to impact and cool the turbine casing. The method adopts any of the above-mentioned turbine disk rim temperature control systems, and opens the valve device at takeoff to connect the bypass line. By directly drawing air from the air bleed line, the pressure of a part of the manifold is increased, the manifold cooling air volume is supplemented, and the casing surface of a specific stage is impacted, thereby reducing the wheel rim sealing gap and the inter-stage grate gap, reducing the air consumption of the compressor bleed air to cool the sealed turbine disk rim and improving the turbine disk rim temperature. The valve device is closed during climbing, cruising or / and slow speed to disconnect the bypass line.

[0012] In one embodiment, the air duct draws air from the outer circumference of the fan.

[0013] According to the embodiment of the present invention, by modifying the clearance control system, the outer surface of the turbine casing of a specific number of stages is subjected to adjustable impact strengthening heat exchange cooling casing, and the contraction of the turbine casing drives the contraction of the guide vanes to adjust the interstage and wheel rim sealing clearance formed by the guide vanes and the turbine disk. The impact pipe bleeds air, and the impact cooling flow rate can be adjusted by the valve device in different states of engine operation, so that the casing and the rotor deformation are matched, and the interstage and wheel rim clearances are optimized in various operating states of the engine, reducing the air consumption of the compressor bleed air cooling and sealing the turbine disk rim while improving the turbine disk rim temperature. BRIEF DESCRIPTION OF THE DRAWINGS

[0014] The above and other features, properties and advantages of the present invention will become more apparent through the following description in conjunction with the accompanying drawings and embodiments, in which:

[0015] Figure 1 It is a schematic diagram of an aircraft engine;

[0016] Figure 2 It is a schematic diagram of the tip clearance control system impact casing;

[0017] Figure 3It is a schematic diagram of the clearance adjustment or rim sealing of the first-stage turbine guide vane;

[0018] Figure 4 is a schematic diagram of a turbine disk rim temperature control system according to an embodiment. DETAILED DESCRIPTION

[0019] Reference will now be made in detail to embodiments of the present invention, one or more examples of which are illustrated in the accompanying drawings. Each example is provided to explain the present invention, not to limit the present invention. In fact, it will be apparent to those skilled in the art that various modifications and variations may be made in the present invention without departing from the scope or spirit of the present invention. For example, a feature illustrated or described as part of one embodiment may be used together with another embodiment to produce yet another embodiment. Therefore, the present invention is intended to cover these modifications and variations within the scope of the appended claims and their equivalents.

[0020] A turbine is a rotary power machine that converts the energy of a flowing medium into mechanical work. It is one of the main components in an aircraft engine. Figure 1 It is a schematic diagram of the basic structure of an aviation turbofan engine. The engine has a low-pressure compressor (including a fan 30 and a supercharging stage 31), a high-pressure compressor 32, a combustion chamber 33, a high-pressure turbine 34 and a low-pressure turbine 35 in the airflow direction. Among them, the low-pressure compressor and the low-pressure turbine 35 are connected to the low-pressure shaft (not shown), and the high-pressure compressor 32 and the high-pressure turbine 34 are connected to the high-pressure shaft (not shown). In the circumferential and axial directions, the above-mentioned engine components are surrounded by a segmented casing, the casing is placed in the core cabin 24, and the entire engine is placed in the nacelle 21. When the engine is working, the airflow 50 passes through the engine fan 30 and the diverter ring 22 on the supercharging stage 31. A part of the air 51 enters the inner flow channel through the supercharging stage 31, and then the air compressed by the high-pressure compressor 32 flows into the combustion chamber 33 and mixes with the fuel to burn. The combustion gas drives the high-pressure turbine 34 and the low-pressure turbine 35, the high-pressure turbine 34 drives the high-pressure compressor 32 to rotate, and the low-pressure turbine 35 drives the fan 30 to rotate. A portion of the air 52 directly enters the outer duct through the support plate 23 behind the fan 30, and is ejected backward through the outer duct composed of the engine nacelle 21 and the core cabin 24 to generate thrust to drive the engine.

[0021] Figure 1 and Figure 2An implementation of the turbine rim temperature control system is also schematically shown, and the turbine rim temperature control system can be understood as a part of the clearance control system. The turbine rim temperature control system includes an air bleed device 1, an air bleed pipeline 2, a circumferential distribution main pipe 4, a gas collecting cavity 5 and a plurality of manifolds 6. An air bleed device 1 installed on the fan culvert is used to collect the airflow 52 of the fan culvert. Part of the collected airflow 53 passes through the air bleed device 1 and a section of the air bleed pipeline 2, enters the circumferential distribution main pipe 4 above the casing where clearance control is required, and then enters the manifold 6 through the distribution main pipe 4 and the gas collecting cavity 5. The plurality of manifolds 6 surround the outer circumference of the casing 41 and are arranged in the axial direction. The gas collecting cavity 5 can be optionally provided in a plurality of circumferential directions, and is respectively connected to the distribution main pipe 4, so that the airflow outputted from each impact hole 7 of the manifold 6 tends to be uniform. The airflow is finally discharged from the impact hole 7 of the manifold 6, impacts the casing 41, and adjusts the tip clearance 43 between the casing 41 and the rotor 42 by changing the deformation of the casing 41. Impingement heat transfer is a form of heat transfer that utilizes the impact of airflow to enhance convective heat transfer and has the advantage of high heat transfer efficiency.

[0022] Figure 3 The schematic diagram shows the principle of the gap control system to improve the interstage and rim seal gaps. The compressor bleed air 54 cools the turbine rotor disk 42 and flows out through the rim seal gap 47 before and after the guide vane and the interstage grate gap 46. Since the guide vane 44 is hung on the turbine casing 41, when the gap control system impacts the casing 41 through the impact hole 7, the casing 41 will simultaneously drive the guide vane 44 to contract inward, reducing the rim seal gap 47 and the interstage grate gap 46.

[0023] like Figure 4 As shown, the turbine rim temperature control system also includes a bypass line 8 led out from the bleed air line 2, and a valve device 9 is arranged on the bypass line 8. An example of the valve device 9 includes a valve and an actuator. The actuator is actuated according to the working condition of the aircraft or the working condition of the engine to drive the valve to operate, thereby realizing the opening and closing of the bypass line 8. The bypass line 8 is connected to part of the cooling manifold 6, and it can be connected to one cooling manifold 6 through one outlet, or it can be connected to multiple cooling manifolds 6 through multiple outlets. In another embodiment, the bypass line 8 is the same as the cooling line independent of the cooling manifold 6, and the cooling line surrounds the outer circumference of the casing 41 like the cooling manifold 6. During takeoff, the valve device 9 is opened, the bypass line 8 is connected, and the pressure of the manifold 6 or the cooling line is increased by directly drawing air from the fan duct, so as to supplement the cooling air volume of the casing corresponding to the manifold 6, impact the casing surface of a specific stage, reduce the wheel rim sealing gap 47 and the inter-stage grate gap 46, reduce the air consumption of the turbine disk rim through the compressor bleed air cooling and sealing, and improve the turbine disk rim temperature; during climbing, cruising and / or slow speed, the valve / actuator 9 is closed, the bypass is disconnected, and the clearance control system works normally.

[0024] The above-mentioned embodiment can actively adjust the amount of impingement cooling air, draw cooling air in when the temperature before the turbine is high, such as during takeoff, to reduce the gap between the stages and the wheel rim, and reduce or not draw cooling air in other low-temperature states. By actively adjusting the cooling bleed air, the amount of bleed air can be saved and the engine performance can be improved. In this way, by drawing a small amount of cooling air from the outer duct, the gap between the turbine disc stages and the wheel rim seal can be reduced, thereby reducing the amount of cooling bleed air passing through the compressor and improving the engine performance.

[0025] The terms "upstream" and "downstream" refer to the relative directions of the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing. In alternative embodiments, such as Figure 4 As shown, with reference to the main direction of airflow through the turbine, the cooling pipeline connected to the bypass pipeline 8 is located on the upstream side of the turbine casing 41. The cooling pipeline is the cooling pipeline of the manifold 6 or the aforementioned independent manifold 6. Generally, the more front turbine casing 41 requires more cooling air. The cooling pipeline located on the upstream side of the turbine casing 41 can distribute the cooling air directly introduced from the air duct 2 to the part that needs cooling most, thereby improving the utilization rate of cooling air. However, the arrangement of the cooling pipeline connected to the bypass pipeline 8 is not limited to a certain stage of low-pressure turbine.

[0026] Although the present invention is disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications without departing from the spirit and scope of the present invention. For example, the above embodiment can be structurally adjusted and refined according to the actual structure of the engine and the need for turbine disc rim cooling. The core content and idea is to achieve engine performance benefits and disc temperature resistance requirements at a relatively low cost by setting a bypass modification structure in the gap control system. Therefore, any modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope defined by the claims of the present invention.

Claims

1. Turbine disk rim temperature control system, including: An air bleed device for introducing cold air; An air bleed pipeline for receiving cold air output by the air bleed device; a circumferential distribution manifold connected to the end of the bleed air pipeline and surrounding the outer circumference of the turbine casing; A plurality of manifolds connected to the circumferential distribution main pipe, axially distributed on the outer peripheral side of the turbine casing, and used for impingement cooling of the turbine casing, characterized in that the turbine rim temperature control system further comprises: a bypass pipeline, which is led out from the bleed air pipeline and connected to at least one cooling pipeline, wherein the cooling pipeline is arranged around the turbine outer casing; and The valve device is used to control the on-off of the bypass pipeline.

2. The turbine disk rim temperature control system according to claim 1, characterized in that: The bypass pipeline has a plurality of outlets, and the plurality of outlets are respectively connected to the plurality of cooling pipelines.

3. The turbine disk rim temperature control system according to claim 2, characterized in that: The plurality of cooling lines connected to the plurality of outlets are located on the upstream side of the turbine casing.

4. The turbine disk rim temperature control system according to claim 1, 2 or 3, characterized in that: The cooling line is selected from a plurality of the manifolds or is a line independent of the manifold.

5. The turbine disk rim temperature control system according to claim 1, 2 or 3, characterized in that: The valve device includes a valve and an actuator for driving the valve to move, and the actuator controls the opening and closing of the valve according to the working state of the engine or the aircraft.

6. A method for controlling the temperature of a turbine disk rim, wherein cold air is introduced through an air bleed pipe to impact and cool the turbine casing, characterized in that: A turbine rim temperature control system as described in any one of claims 1 to 5 is adopted, and the valve device is opened at takeoff to connect the bypass line, and air is directly bleed from the bleed duct to increase the pressure of a part of the manifold, supplement the manifold cooling air volume, impact the casing surface of a specific stage, reduce the rim sealing gap and the inter-stage grate gap, reduce the air consumption of the turbine rim sealed by the compressor bleed air, and improve the turbine rim temperature; the valve device is closed during climbing, cruising or / and idling to disconnect the bypass line.

7. The turbine disk rim temperature control method according to claim 6, characterized in that: The air duct draws air from the outer circumference of the fan.

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