An aero-engine core engine turbine working blade cooling air bleeding structure

Abstract

The application belongs to the technical field of aero-engine core turbine working blade cooling bleed air design, and particularly relates to a kind of aero-engine core turbine working blade cooling bleed air structure, when aero-engine works, current can be passed to electromagnetic converter through each electromagnetic control lead, electromagnetic converter generates magnetic force, acts with each magnetic seat, drives each valve stem to overcome the elastic force of spring to move, each valve piece removes the plugging of each bleed air groove, part of airflow in annular flow passage between inner wall 6 of combustion chamber and flame tube 8 is taken out as cooling air from each bleed air groove to cool high vortex working blade 10, and the size of current can be adjusted according to different working conditions, the plugging degree of each valve piece to each bleed air groove is adjusted, the flow of cooling air is adjusted, which is beneficial to ensure the performance optimization of aero-engine in the whole working condition range.

Description

Technical Field

[0001] This application belongs to the field of cooling bleed air design technology for turbine blades of aero-engine core engines, and specifically relates to a cooling bleed air structure for turbine blades of aero-engine core engines. Background Technology

[0002] The core of an aero-engine includes a compressor 1, a combustion chamber 2, and a turbine 3 connected in sequence. The compressor 1 and turbine 3 are connected by a high-pressure shaft 4. The compressor 1 pressurizes the airflow and supplies it to the combustion chamber 2 for combustion, producing high-temperature gas. This gas then drives the turbine 3 to rotate, performing work on the compressor 1 via the high-pressure shaft 4. Figure 1 As shown.

[0003] The combustion chamber 2 in the core of the aero-engine includes an outer combustion chamber wall 5, an inner combustion chamber wall 6, a diffuser 7, and a flame tube 8. The inner combustion chamber wall 6 is located inside the outer combustion chamber wall 5, and the flame tube 8 is located between the inner combustion chamber wall 6 and the outer combustion chamber wall 5. The diffuser 7 is located in front of the flame tube 8. After the high-pressure gas from the compressor 1 is diffused by the diffuser 7, part of it flows to the flame tube 8 for combustion to produce high-temperature gas, part of it flows into the annular flow channel between the outer combustion chamber wall 5 and the flame tube 8 to cool the outer wall of the flame tube 8, and part of it flows into the annular flow channel between the inner combustion chamber wall 6 and the flame tube 8 to cool the outer wall of the flame tube 8.

[0004] In the core of the aero-engine, a high-guide blade 9 is installed at the inlet of the turbine 3, and a high-vortex working blade 10 is installed after the high-guide blade 9. The high-temperature gas generated by the flame tube 8 enters the turbine 3, and after being rectified by the high-guide blade 9, it drives the high-vortex working blade 10 to rotate, and then performs work on the compressor 1 through the high-pressure shaft 4, thereby driving the compressor 1 to perform work.

[0005] The high-vortex working blade 10 of turbine 3 in the core engine of aero-engine operates in a harsh environment, subjected to the impact of high-temperature gas and subjected to temperatures of thousands of degrees. In order to ensure that the high-vortex working blade 10 is not damaged by high temperature, the following two methods are mainly used for cooling.

[0006] Method 1: A radial airflow channel is opened inside the high-guide-rate blade 9. The radial airflow channel connects to the outlet of the annular flow channel between the outer wall of the combustion chamber 5 and the flame tube 8. A pre-swirl nozzle connected to the radial airflow channel is installed at the root of the high-guide-rate blade 9. When the aero-engine is working, part of the airflow in the annular flow channel between the outer wall of the combustion chamber 5 and the flame tube 8 can pass through the radial airflow channel and be ejected by the pre-swirl nozzle, spraying towards the high-vortex working blade 10 to cool the high-vortex working blade 10. This technical solution has the following drawbacks:

[0007] 1) A radial airflow channel needs to be opened inside the high-guide-rate blade 9. The blade profile of the high-guide-rate blade 9 is usually narrow, making it difficult to set up a radial airflow channel, which is a great design challenge.

[0008] 2) The cooling airflow channel structure is fixed and the amount of cooling air cannot be adjusted. Under different operating conditions, the cooling air requirement of the high-vortex working blade 10 varies, and a large amount of cooling air will affect the working performance of the aero-engine, which is not conducive to the performance optimization of the aero-engine in the entire operating range.

[0009] 3) After the cooling airflow passes through the radial airflow channel opened inside the high-guide blade 9, it cools the high-vortex working blade 10. However, it is heated by the high-temperature gas inside the high-guide blade 9, resulting in poor cooling effect on the high-vortex working blade 10. To ensure the cooling effect on the high-vortex working blade 10, a cooler is usually required to cool the cooling air. In order to ensure the cooling effect of the cooling air in a short distance, the cooler usually needs to be designed with a large heat exchange area and volume, resulting in a large weight. This will increase the overall size and mass of the aero-engine and increase the airflow resistance, affecting the overall performance of the engine.

[0010] Method 2 involves inserting a pipe between the diffuser 7 and the flame tube 8 to guide part of the airflow from the inside of the annular channel between the outer wall of the combustion chamber 5 and the flame tube 8 to the high-vortex working blade 10 for cooling. A valve regulating mechanism is installed to adjust the cooling air flow rate according to the operating conditions. This technical solution has the following drawbacks:

[0011] 1) Inserting a pipe between the diffuser 7 and the flame tube 8 will occupy the length space between the diffuser 7 and the flame tube 8, which will increase the axial length of the combustion chamber 2 and affect the performance of the combustion chamber 2.

[0012] 2) Cooling air flow is regulated by valve adjustment mechanisms, which are mostly mechanical. After long-term use, wear and deformation will occur, resulting in reduced component fit accuracy and inability to guarantee adjustment accuracy. Moreover, the power for movement is mostly provided by booster pumps, which increases the complexity of the overall fuel system. To meet the fuel supply, the design of fuel pipelines is required, which increases the difficulty of laying external pipelines. It is also necessary to design multiple functional interfaces such as fuel interface, air interface, and circuit interface, which is highly complex. In addition, there are many components required, which are difficult to arrange within the limited and compact structure of the aero engine, and at the same time, it greatly increases the weight of the whole machine.

[0013] This application is made in view of the aforementioned technical deficiencies. Summary of the Invention

[0014] The purpose of this application is to provide a cooling bleed air structure for the turbine working blades of an aero-engine core, in order to overcome or mitigate at least one of the known technical defects.

[0015] The technical solution of this application is:

[0016] A cooling bleed air structure for turbine blades of an aero-engine core includes:

[0017] A support ring is installed between the inner wall of the combustion chamber and the high-pressure shaft. It has multiple support channels distributed circumferentially inside, and the rear end face has support holes that connect the various support channels. The rear end of the inner wall of the combustion chamber has multiple air venting grooves distributed circumferentially.

[0018] Multiple magnetic bases are installed within each support channel;

[0019] Multiple valve stems, with their front ends inserted into the support channels through the support holes, connect to the magnetic seats;

[0020] Multiple valve plates are installed between the combustion chamber wall and the flame tube, and are attached to the combustion chamber wall, with bent parts; each bent part passes through each air intake groove and connects to the rear end of each valve stem.

[0021] Multiple springs are installed in each support channel, with both ends abutting against the front end of each support channel and between the magnetic base. Relying on elastic force, each bent part is kept against the back of each air intake groove, and each valve plate blocks each air intake groove.

[0022] Multiple lead tubes, one end of which is a plugged end and the other end is an open end. The plugged end of each lead tube passes through the outer wall of the combustion chamber, the diffuser, and the outer wall of the combustion chamber from the front end to the front end of the support ring, and corresponds to the position of each valve stem in the circumferential direction.

[0023] Multiple electromagnetic transducers are installed inside the sealing end of each lead tube;

[0024] Multiple electromagnetic control leads are connected at one end to each electromagnetic converter, and the other end is led out from the opening end of each lead tube. When current is passed through, the electromagnetic converter can generate magnetic force, which interacts with each magnetic base, causing each valve stem to move against the elastic force of the spring, thereby releasing each valve plate from blocking each air vent.

[0025] According to at least one embodiment of this application, in the above-described cooling air bleed structure for the turbine working blades of the aero-engine core, the rear end of the combustion chamber wall has a plurality of circumferentially distributed air bleed holes.

[0026] The cooling bleed air structure for the turbine blades of the aero-engine core has the following features:

[0027] Sonic cruise bleed air status: All electromagnetic control leads are not energized, all bends are against the back of each bleed air slot, and all valve plates completely block each bleed air slot.

[0028] Subsonic and supersonic transition cruise bleed air state: A small current is passed through each electromagnetic control lead, each bend reaches the middle part of each bleed air slot, and each valve plate partially releases the blockage of each bleed air slot.

[0029] Supersonic cruise bleed air state: A large current is passed through each electromagnetic control lead, each bend is against the front of each bleed air slot, and each valve plate completely releases the blockage of each bleed air slot.

[0030] According to at least one embodiment of this application, in the above-described cooling bleed structure for the turbine working blades of the aero-engine core, each support ring is connected to the inner wall of the combustion chamber through multiple support plates.

[0031] According to at least one embodiment of this application, the above-described cooling bleed air structure for the turbine working blades of the aero-engine core engine further includes:

[0032] Multiple support plates are connected to the rear end of the support ring, extend from each air intake groove, abut against the front of each air intake groove, and abut against the inner side of each valve plate.

[0033] According to at least one embodiment of this application, the above-described cooling bleed air structure for the turbine working blades of the aero-engine core engine further includes:

[0034] The guide tube is installed between the inner wall of the combustion chamber and the high-pressure shaft, and its front end is connected to the rear end of the support ring through an outward folded edge.

[0035] According to at least one embodiment of this application, the above-described cooling bleed air structure for the turbine working blades of the aero-engine core engine further includes:

[0036] Multiple sets of metal sealing wires are wound around each valve stem to form a seal between the valve stem and each support channel.

[0037] According to at least one embodiment of this application, in the above-described cooling bleed structure for the turbine working blades of the aero-engine core, each lead tube has a sandwich layer inside its sidewall, and the sandwich layer is filled with heat insulation cotton.

[0038] This application has at least the following beneficial technical effects:

[0039] A cooling bleed air structure for the turbine blades of an aero-engine core is provided. During aero-engine operation, current can be supplied to the electromagnetic converter through various electromagnetic control leads, causing the electromagnetic converter to generate magnetic force. This force interacts with the magnetic bases, driving the valve stems to move against the elastic force of the springs. This causes the valve plates to release the blockage of the bleed air slots, allowing part of the airflow within the annular flow channel between the combustion chamber wall 6 and the flame tube 8 to be used as cooling air. This cooling air is then drawn out from the bleed air slots to cool the high-vortex turbine blades 10. Furthermore, the magnitude of the supplied current can be adjusted according to different operating conditions to regulate the degree of blockage of the bleed air slots by the valve plates and adjust the flow rate of the cooling air, thus facilitating performance optimization of the aero-engine across the entire operating range. Attached Figure Description

[0040] Figure 1 This is a structural schematic diagram of an existing aero-engine core.

[0041] Figure 2 This is a schematic diagram of the cooling bleed air structure for the turbine working blades of an aero-engine core provided in an embodiment of this application;

[0042] Figure 3 yes Figure 2 Sectional view along axis AA;

[0043] Figure 4 This is a schematic diagram of the cooling bleed air structure of the turbine working blades of the aero-engine core provided in this application embodiment in the pressure-sonic cruise bleed air state;

[0044] Figure 5 This is a schematic diagram of the cooling bleed air structure of the turbine working blades of the aero-engine core provided in this application embodiment in the subsonic and supersonic transition cruise bleed air state;

[0045] Figure 6 This is a schematic diagram of the cooling bleed air structure of the turbine working blades of the aero-engine core provided in this application embodiment in the supersonic cruise bleed air state;

[0046] in:

[0047] 1-Compressor; 2-Combustion chamber; 3-Turbine; 4-High-pressure shaft; 5-Outer wall of combustion chamber; 6-Inner wall of combustion chamber; 7-Diffuser; 8-Flame tube; 9-High-pressure guide vane; 10-High-pressure vortex working vane; 11-Support ring; 12-Valve stem; 13-Valve plate; 14-Spring; 15-Lead tube; 16-Electromagnetic converter; 17-Electromagnetic control lead; 18-Magnetic base; 19-Support plate; 20-Guide tube; 21-Metal sealing wire.

[0048] To better illustrate this embodiment, some parts in the accompanying drawings may be omitted, enlarged, or reduced, and do not represent the actual size of the product. Furthermore, the accompanying drawings are for illustrative purposes only and should not be construed as limiting this patent. Detailed Implementation

[0049] To make the technical solution and advantages of this application clearer, the technical solution of this application will be described in a clearer and more complete manner below with reference to the accompanying drawings. It should be understood that the specific embodiments described herein are only some embodiments of this application, and are only used to explain this application, not to limit this application. It should be noted that, for ease of description, only the parts related to this application are shown in the accompanying drawings, and other related parts can be referred to the general design.

[0050] Furthermore, unless otherwise defined, the technical or scientific terms used in this application description shall have the ordinary meaning understood by one of ordinary skill in the art to which this application pertains. The terms indicating direction used in this application description are used only to indicate relative direction or positional relationship; when the absolute position of the described object changes, its relative positional relationship may also change accordingly. The word "comprising" as used in this application description indicates that the element or object preceding the word encompasses the elements or objects listed following the word and their equivalents, but does not exclude other elements or objects.

[0051] Furthermore, it should be noted that, unless otherwise explicitly specified and limited, terms such as "installation" and "connection" used in the description of this application should be interpreted broadly. For example, a connection can be a fixed connection or a detachable connection; it can be a mechanical connection or an electrical connection; it can be a direct connection or an indirect connection through an intermediate medium. Those skilled in the art can understand its specific meaning in this application according to the specific circumstances.

[0052] A cooling bleed air structure for turbine blades in an aero-engine core, such as Figures 2-3 As shown, it includes:

[0053] The support ring 11 is disposed between the inner wall 6 of the combustion chamber and the high-pressure shaft 4. It has multiple support channels distributed circumferentially inside, and the rear end face has support holes connecting the various support channels; the rear end of the inner wall 6 of the combustion chamber has multiple air venting grooves distributed circumferentially.

[0054] Multiple magnetic seats 18 are provided in each support channel;

[0055] Multiple valve stems 12, with their front ends inserted into the support channels through the support holes, are connected to the magnetic seats 18.

[0056] Multiple valve plates 13 are disposed between the inner wall 6 of the combustion chamber and the flame tube 8, abutting against the inner wall 6 of the combustion chamber, and have bent portions; each bent portion passes through each air intake groove and connects to the rear end of each valve stem 12.

[0057] Multiple springs 14 are installed in each support channel, with both ends abutting against the front end of each support channel and between the magnetic base 18. Relying on elastic force, each bent part is kept against the back of each air intake groove, and each valve plate 13 blocks each air intake groove.

[0058] Multiple lead tubes 15, one end of which is a plugged end and the other end is an open end. The plugged end of each lead tube 15 passes through the outer wall 5 of the combustion chamber, the diffuser 7, and the outer wall 6 of the combustion chamber from the front end and reaches the front end of the support ring 11, corresponding to the position of each valve stem 12 in the circumferential direction.

[0059] Multiple electromagnetic converters 16 are disposed within the sealing end of each lead tube 15, and may be designed to include coils and iron cores disposed within the coils.

[0060] Multiple electromagnetic control leads 17 are connected at one end to each electromagnetic converter 16, and the other end is led out from the opening end of each lead tube 15. When current is passed through, the electromagnetic converter 16 can generate magnetic force, which interacts with each magnetic base 18, causing each valve stem 12 to move against the elastic force of the spring 14, so that each valve plate 13 can release the blockage of each air vent.

[0061] The above embodiment discloses a cooling bleed air structure for the turbine blades of an aero-engine core. When the aero-engine is working, current can be supplied to the electromagnetic converter 16 through each electromagnetic control lead 17, causing the electromagnetic converter 16 to generate magnetic force. This magnetic force interacts with each magnetic base 18, driving each valve stem 12 to move against the elastic force of the spring 14. This causes each valve plate 13 to release the blockage of each bleed air slot, allowing part of the airflow in the annular flow channel between the combustion chamber wall 6 and the flame tube 8 to be used as cooling air. This cooling air is drawn out from each bleed air slot to cool the high-vortex working blades 10. Furthermore, the magnitude of the supplied current can be adjusted according to different operating conditions to regulate the degree of blockage of each valve plate 13 on each bleed air slot and adjust the flow rate of the cooling air, which is beneficial to ensuring the optimal performance of the aero-engine across the entire operating range.

[0062] The cooling bleed air structure for the turbine working blades of the aero-engine core engine disclosed in the above embodiment does not require the opening of a radial airflow channel in the high-guide blade 9, thus avoiding the increase in design difficulty of the high-guide blade 9. It also eliminates the need for a heat exchanger to cool the cooling air, avoiding a series of problems caused by the installation of a heat exchanger. Furthermore, the lead tube 15 is designed to pass through the diffuser 7, thus avoiding occupying the length space between the diffuser 7 and the flame tube 8, increasing the axial length of the combustion chamber 2, and affecting the performance of the combustion chamber 2. In addition, the design utilizes the lead tube 15 to pass current to the electromagnetic converter 16 to generate magnetism, which interacts with the magnetic seat 18 to drive each valve stem 12 to move against the elastic force of the spring 14, thereby realizing the regulation of the cooling airflow. It is not prone to wear and deformation, and can maintain high regulation accuracy for a long time. The overall structure is simple and compact, making it easy to arrange on an aero-engine.

[0063] The electromagnetic control lead 17 can be connected to the aircraft engine control system, so that the aircraft engine control system can control the magnitude of the input current as needed.

[0064] Considering that the typical operating modes of an aero-engine in practice are composonic cruise, supersonic cruise, and a transitional state between composonic and supersonic cruise, the high-vortex turbine blade 10 has a lower demand for cooling air during composonic cruise, a higher demand during supersonic cruise, and an intermediate demand during the transitional state. To reduce the structural stress, the cooling bleed structure for the turbine blades of the aero-engine core engine is designed with multiple circumferentially distributed bleed holes at the rear end of the combustion chamber wall 6. The cooling bleed structure for the turbine blades of the aero-engine core engine is designed with the following features:

[0065] Sonic cruise bleed air status: All electromagnetic control leads 17 are de-energized, all bends are abutting the rear of the respective bleed air channels, and all valve plates 13 completely block the respective bleed air channels. Figure 4 As shown, at this time, the cooling air can only be drawn out from each air intake hole to cool the high-vortex working blade 10. The flow rate of the cooling air that can be drawn out is relatively small, which can be adapted to the supersonic cruise working mode of the aero engine.

[0066] Subsonic / supersonic transition cruise bleed air state: A small current is applied to each electromagnetic control lead 17, and each bend reaches the middle of each bleed air slot. Each valve plate 13 partially releases its blockage of each bleed air slot, such as... Figure 5 As shown, at this time, cooling air can be drawn out from the parts where the air vents and air vent slots are unblocked to cool the high-vortex working blades 10. The flow rate of cooling air that can be drawn out is relatively large and can be adapted to the transitional working mode of the aero-engine.

[0067] Supersonic cruise bleed air state: A large current is applied to each electromagnetic control lead 17, each bend is pressed against the front of each bleed air slot, and each valve plate 13 completely releases its blockage of each bleed air slot. Figure 6 As shown, at this time, cooling air can be drawn out from each air intake hole and each air intake slot to cool the high-speed vortex working blade 10. The flow rate of the cooling air that can be drawn out reaches the maximum, which can be adapted to the supersonic cruise working mode of the aero engine.

[0068] In some optional embodiments, in the above-described cooling bleed structure for the turbine working blades of the aero-engine core, each support ring 11 is connected to the inner wall 6 of the combustion chamber through multiple support plates to form a stable support structure.

[0069] In some optional embodiments, the above-described cooling bleed air structure for the turbine blades of the aero-engine core engine further includes:

[0070] Multiple support plates 19 are connected to the rear end of the support ring 11, extend from each air intake groove, abut against the front of each air intake groove, and abut against the inner side of each valve plate 13. This can lift the support plates 19 away from contact with the inner wall 6 of the combustion chamber, thereby reducing the friction between the valve plate 13 and the inner wall 6 of the combustion chamber.

[0071] In some optional embodiments, the above-described cooling bleed air structure for the turbine blades of the aero-engine core engine further includes:

[0072] The guide tube 20 is installed between the combustion chamber wall 6 and the high-pressure shaft 4. Its front end is connected to the rear end of the support ring 11 through an outward folded edge. It can guide the cooling air out and efficiently cool the high-pressure working blades 10.

[0073] In some optional embodiments, the above-described cooling bleed air structure for the turbine blades of the aero-engine core engine further includes:

[0074] Multiple sets of metal sealing wires 21 are wound around each valve stem 12 to form a seal with each support channel, preventing airflow from passing through.

[0075] In some optional embodiments, in the above-described cooling bleed structure for the turbine working blades of the aero-engine core, each lead tube 15 has a sandwich layer in its sidewall, and the sandwich layer is filled with heat insulation cotton to protect the electromagnetic control lead 17 from high temperature damage.

[0076] The various embodiments in the specification are described in a progressive manner, with each embodiment focusing on the differences from other embodiments. Similar or identical parts between embodiments can be referred to mutually. In the absence of conflict, the embodiments and technical features in the embodiments of this application can be combined to obtain new embodiments.

[0077] The technical solution of this application has been described in conjunction with the preferred embodiments shown in the accompanying drawings. Those skilled in the art should understand that the protection scope of this application is obviously not limited to these specific embodiments. Without departing from the principles of this application, those skilled in the art can make equivalent changes or substitutions to the relevant technical features, and the technical solutions after these changes or substitutions will all fall within the protection scope of this application.

Claims

1. An aircraft engine core turbine workpiece cooling bleed air structure, characterized by, Comprise: Support ring (11), set between combustion chamber inner wall (6), high pressure shaft (4), with multiple circumferential support channels in it, the rear end face has support holes connected with each support channel; the rear end of the combustion chamber inner wall (6) has multiple circumferential air guide grooves; Multiple magnetic seats (18) are arranged in each support channel; Multiple valve stems (12) are inserted into the support channel from each support hole at the front end, and are connected with each magnetic seat (18); Multiple valve plates (13) are arranged between the combustion chamber inner wall (6) and the flame tube (8), abutting against the combustion chamber inner wall (6), having bending parts; each bending part is connected to the rear end of each valve stem (12) through each air guide groove; Multiple springs (14) are arranged in each support channel, abutting against each support channel front end and magnetic seat (18) at both ends, relying on elastic force to keep each bending part abutting against the rear of each air guide groove, and each valve plate (13) blocks each air guide groove; Multiple lead pipes (15) have one end as a blocking end and the other end as an opening end, wherein the blocking end of each lead pipe (15) is arranged from the front end through the combustion chamber outer wall (5), the diffuser (7) and the combustion chamber inner wall (6) to the front end of the support ring (11), corresponding to the position of each valve stem (12) in the circumferential direction; Multiple electromagnetic converters (16) are arranged in the blocking end of each lead pipe (15); Multiple electromagnetic control leads (17) are connected with each electromagnetic converter (16) at one end and led out from the opening end of each lead pipe (15) at the other end, and can generate a magnetic force when current is passed through, so that the electromagnetic converter (16) can generate a magnetic force to act on each magnetic seat (18) to drive each valve stem (12) to move against the elastic force of the spring (14), so that each valve plate (13) unblocks each air guide groove; Part of the airflow in the annular flow passage between the combustion chamber inner wall (6) and the flame tube (8) is led out from each air guide groove as cooling air to cool the high vortex working blade (10).

2. The turbine working blade cooling air guide structure of the aero-engine core engine according to claim 1, wherein the rear end of the combustion chamber inner wall (6) has multiple circumferential air guide holes; The turbine working blade cooling air guide structure has: A subsonic cruise air guide state: each electromagnetic control lead (17) is not electrified, each bending part abuts against the rear of each air guide groove, and each valve plate (13) completely blocks each air guide groove; A subsonic to supersonic transition cruise air guide state: each electromagnetic control lead (17) is electrified with a small current, each bending part reaches the middle part of each air guide groove, and each valve plate (13) partially unblocks each air guide groove; A supersonic cruise air guide state: each electromagnetic control lead (17) is electrified with a large current, each bending part abuts against the front of each air guide groove, and each valve plate (13) completely unblocks each air guide groove.

3. The turbine working blade cooling air guide structure of the aero-engine core engine according to claim 1, wherein each support ring (11) is connected with the combustion chamber inner wall (6) through multiple support plates. ​ ​ 4. The aeroengine core engine turbine working blade cooling bleed air structure according to claim 1, characterized in that, Also includes: Multiple support plates (19) are connected to the rear end of the support ring (11), extend from each air intake groove, abut against the front of each air intake groove, and abut against the inner side of each valve plate (13).

5. The gas turbine engine core turbine workpiece cooling bleed air structure of claim 1, wherein, Also includes: The guide tube (20) is set between the inner wall (6) of the combustion chamber and the high-pressure shaft (4), and its front end is connected to the rear end of the support ring (11) through an outward folded edge.

6. The gas turbine engine core turbine workpiece cooling bleed air structure of claim 1, wherein, Also includes: Multiple sets of metal sealing wires (21) are wrapped around each valve stem (12) to form a seal with each support channel.

7. The cooling bleed air structure for the turbine blades of an aero-engine core engine according to claim 1, characterized in that, Each lead tube (15) has a sandwich layer inside its side wall, and the sandwich layer is filled with heat insulation cotton.

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