A cooling bleed air structure for turbine blades of a gas turbine engine core.

Abstract

This application specifically relates to a cooling bleed air structure for the turbine blades of a gas turbine engine core, comprising: a gas collecting shroud, fitted onto the outer side of the combustion chamber outer wall and communicating with the interior of the combustion chamber outer wall; a bleed air ring, fitted onto the outer circumference of the high-pressure shaft, with an annular bleed air passage formed therein; the annular bleed air passage forming an annular bleed air cavity at the front end of the bleed air ring and an annular conical opening at the rear end of the bleed air ring; an annular adjusting plate, the front end of which is inserted into the annular bleed air passage through the annular conical opening and extends into the annular bleed air cavity, and the rear end of which is formed with an annular conical block; a bleed air pipe, one end of which communicates with the interior of the gas collecting shroud, and the other end of which passes through the outer wall of the combustion chamber, the diffuser, and the outer wall of the combustion chamber from the front end and communicates with the annular bleed air cavity; a transmission rod, disposed within the bleed air pipe, one end of which extends into the annular bleed air cavity and is toothedly engaged with the front end of the annular adjusting plate; and an actuator cylinder, disposed outside the combustion chamber outer wall, with its piston rod extending into the bleed air pipe and toothedly engaged with the other end of the transmission rod.

Description

Technical Field

[0001] This application belongs to the field of cooling bleed air design technology for turbine blades of gas turbine engine core, specifically relating to a cooling bleed air structure for turbine blades of gas turbine engine core. Background Technology

[0002] The core of a gas turbine 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 gas turbine 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 gas turbine 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 the gas turbine 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 11, which connects to the radial airflow channel, is installed at the root of the high-guide-rate blade 9. When the aero-engine is operating, 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 11, spraying it towards the high-vortex working blade 10 to cool it. Figure 2 As shown, 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 demand of the high-vortex working blade 10 varies, and a large amount of cooling air will affect the working performance of the gas turbine engine, which is not conducive to the performance optimization of the gas turbine 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 volume and mass of the gas turbine 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) The cooling air flow is regulated by a valve regulating mechanism, which has a complex structure and low regulation accuracy.

[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 a gas turbine 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 the turbine blades of a gas turbine engine core includes:

[0017] The gas collection hood is fitted onto the outer side of the outer wall of the combustion chamber, located at the rear end of the outer wall of the combustion chamber, and connects to the interior of the outer wall of the combustion chamber.

[0018] The bleed air ring is fitted around the outer periphery of the high-pressure shaft and is located on the inner side of the combustion chamber wall. An annular bleed air passage is formed inside it. The annular bleed air passage forms an annular bleed air chamber at the front end of the bleed air ring and an annular conical opening at the rear end of the bleed air ring.

[0019] The annular regulating plate has its front end inserted into the annular air intake channel through an annular conical opening and extends into the annular air intake chamber, while its rear end is formed with an annular conical block.

[0020] The air intake pipe has one end connected to the inside of the air collection hood, and the other end is set from the front end through the outer wall of the combustion chamber, the diffuser, and the outer wall of the combustion chamber, and is connected to the annular air intake chamber.

[0021] The transmission rod is installed inside the air duct, with one end extending into the annular air duct chamber and connecting with the front toothed part of the annular adjusting plate.

[0022] The actuator is installed outside the outer wall of the combustion chamber. Its piston rod extends into the air intake pipe and is connected to the other end of the transmission rod by a toothed engagement. This allows the transmission rod to rotate through the extension and retraction of the piston rod, thereby driving the annular adjusting plate to move axially and adjusting the effective flow area between the annular cone orifice and the annular cone block.

[0023] According to at least one embodiment of this application, in the above-described gas turbine engine core turbine working blade cooling bleed air structure, the annular conical opening at the rear end of the bleed air ring and the annular conical block at the rear end of the annular adjusting plate have a straight guide section.

[0024] According to at least one embodiment of this application, the above-described gas turbine engine core turbine blade cooling bleed air structure further includes:

[0025] Multiple pairs of support blocks are installed inside the annular air intake channel, supporting the annular adjusting plate and the two side walls of the annular air intake channel.

[0026] According to at least one embodiment of this application, in the above-described gas turbine engine core turbine working blade cooling bleed air structure, the part where the piston rod of the actuator cylinder extends into the bleed air pipe is in the shape of a boss.

[0027] According to at least one embodiment of this application, in the above-described gas turbine engine core turbine working blade cooling bleed air structure, the bleed air ring and the inner wall of the combustion chamber are connected by fins and bolt fasteners.

[0028] According to at least one embodiment of this application, in the above-described gas turbine engine core turbine working blade cooling bleed air structure, the actuator is a hydraulic actuator.

[0029] The cooling bleed air structure for the turbine blades of the gas turbine engine core also includes:

[0030] A hydraulic control device is connected to the actuator cylinder and integrated into the gas turbine engine control system.

[0031] According to at least one embodiment of this application, in the above-described gas turbine engine core turbine working blade cooling bleed air structure, the annular adjusting plate is divided into multiple arc-shaped segments, and there are multiple bleed air pipes, transmission rods, and actuators corresponding to each arc-shaped segment.

[0032] The gas turbine engine core turbine blade cooling bleed air structure also includes:

[0033] Multiple partitions are installed within the annular air intake channel and are distributed circumferentially at intervals with each arc segment. Attached Figure Description

[0034] Figure 1 This is a structural schematic diagram of an existing gas turbine engine core.

[0035] Figure 2 This is a schematic diagram of an existing method for cooling the turbine blades of a gas turbine engine core.

[0036] Figure 3 This is a schematic diagram of the cooling bleed air structure for the turbine working blades of the gas turbine engine core provided in the embodiments of this application;

[0037] Figure 4 yes Figure 3 A partial schematic diagram of part A;

[0038] Figure 5 yes Figure 3 A partial schematic diagram of part B;

[0039] Figure 6 This is a schematic diagram illustrating the adjustment of the effective flow area between the annular cone opening and the annular cone blocks provided in an embodiment of this application;

[0040] Figure 7 This is a partial cross-sectional view of the cooling bleed air structure of the turbine working blades of the gas turbine engine core provided in the embodiments of this application;

[0041] in:

[0042] 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-Pre-swirl nozzle; 12-Gas collection hood; 13-Inlet ring; 14-Annular regulating plate; 15-Inlet pipe; 16-Drive rod; 17-Actuator cylinder; 18-Support block; 19-Hydraulic control device; 20-Divider plate.

[0043] 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

[0044] 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.

[0045] 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.

[0046] 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.

[0047] A cooling bleed air structure for the turbine blades of a gas turbine engine core, such as... Figure 3 Shown, including:

[0048] The gas collection hood 12 is fitted onto the outer side of the outer wall of the combustion chamber 5, located at the rear end of the outer wall of the combustion chamber 5, and connects to the interior of the outer wall of the combustion chamber 5.

[0049] The air intake ring 13 is sleeved on the outer periphery of the high-pressure shaft 4 and located inside the inner wall 6 of the combustion chamber, forming an annular air intake channel; the annular air intake channel forms an annular air intake cavity in the front end of the air intake ring 13 and an annular conical opening in the rear end of the air intake ring 13.

[0050] The annular adjusting plate 14 has its front end inserted into the annular air intake channel through the annular conical opening and extends into the annular air intake chamber, and its rear end is formed with an annular conical block.

[0051] The air intake pipe 15 has one end connected to the inside of the air collection hood 12, and the other end is set from the front end through the outer wall 5 of the combustion chamber, the diffuser 7, and the inner wall 6 of the combustion chamber, and is connected to the annular air intake chamber.

[0052] The transmission rod 16 is installed inside the air duct 15, with one end extending into the annular air duct chamber and connecting with the front toothed part of the annular adjusting plate 14, such as... Figure 4 As shown;

[0053] The actuator 17 is installed outside the outer wall 5 of the combustion chamber. Its piston rod extends into the air intake pipe 15 and is connected to the other end of the transmission rod 16 by a toothed engagement. This allows the piston rod to extend and retract, driving the transmission rod 16 to rotate, thereby causing the annular adjusting plate 14 to move axially and adjust the effective flow area between the annular cone opening and the annular cone block. Figure 5 As shown.

[0054] The above embodiment discloses a cooling bleed air structure for the turbine blades of a gas turbine engine core. During operation, a portion of the airflow from the annular channel between the combustion chamber outer wall 5 and the flame tube 8 is introduced into the gas collection shroud 12. This airflow is then sprayed onto the high-pressure turbine blades 10 through the bleed air pipe 15, annular bleed air chamber, annular bleed air passage, and annular cone, cooling the blades. Furthermore, the effective flow area between the annular cone and the annular cone block can be adjusted by extending or retracting the piston rod of the actuator cylinder 17, thus regulating the cooling air flow rate, depending on the operating conditions. For example, during initial startup of the gas turbine engine, the annular cone block can be retracted into the annular cone opening, sealing it and reducing the cooling air flow rate to zero. As the operating conditions increase, the annular cone block is gradually pushed out of the annular cone opening, increasing the effective flow area between the annular cone opening and the annular cone block, thereby increasing the cooling air flow rate. Figure 6 As shown, this helps to ensure the optimal performance of the gas turbine engine across the entire operating range.

[0055] The above-described embodiment discloses a cooling bleed air structure for the turbine working blades of a gas turbine engine core. This structure eliminates the need for a radial airflow channel within the high-guide blade 9, avoiding increased design complexity for the high-guide blade 9. It also eliminates the need for a heat exchanger to cool the air, preventing a series of problems associated with heat exchanger installation. Furthermore, the bleed air pipe 15, which passes through the diffuser 7, avoids occupying the length space between the diffuser 7 and the flame tube 8, thus preventing an increase in the axial length of the combustion chamber 2 and its performance. In addition, the design utilizes a toothed transmission to adjust the effective flow area between the annular cone opening and the annular cone block, achieving high adjustment accuracy. The entire adjustment mechanism is located inside the cooling flow path, without occupying additional space. The structure is simple, compact, and easy to arrange on a gas turbine engine.

[0056] In some optional embodiments, the above-described cooling bleed structure for the turbine blades of the gas turbine engine core has a straight guide section behind the annular conical opening at the rear end of the bleed ring 13 and the annular conical block at the rear end of the annular adjusting plate 14, so as to guide the outflowing cooling gas and enhance the reliability of the fit between the structures.

[0057] In some optional embodiments, the above-described gas turbine engine core turbine blade cooling bleed air structure further includes:

[0058] Multiple pairs of support blocks 18 are installed inside the annular air intake channel, supporting the annular adjusting plate 14 between the two side walls of the annular air intake channel.

[0059] In some optional embodiments, in the above-described gas turbine engine core turbine working blade cooling bleed air structure, the part where the piston rod of the actuator cylinder 17 on the bleed air pipe 15 extends is in the shape of a boss.

[0060] In some alternative embodiments, in the above-described gas turbine engine core turbine working blade cooling bleed air structure, the bleed air ring 13 is connected to the combustion chamber wall 6 by fins and bolt fasteners.

[0061] In some optional embodiments, in the above-described gas turbine engine core turbine working blade cooling bleed air structure, the actuator 17 is a hydraulic actuator.

[0062] The cooling bleed air structure for the turbine blades of the gas turbine engine core also includes:

[0063] The hydraulic control device 19 is connected to the actuator 17 and connected to the gas turbine engine control system. Thus, the gas turbine engine control system can send control commands to the hydraulic control device 19 to control the piston rod of the actuator 17 to extend and retract, thereby adjusting the cooling air flow.

[0064] In the above embodiment, the cooling bleed structure for the turbine blades of a gas turbine engine core machine discloses a correspondence between the extension and retraction length of the piston rod of the actuator cylinder 17 and the effective flow area between the annular cone orifice and the annular cone block. In specific applications, the correspondence between the required cooling air flow rate and the effective flow area between the annular cone orifice and the annular cone block under different operating conditions can be determined. This leads to the correspondence between the required cooling air flow rate and the extension and retraction length of the piston rod of the actuator cylinder 17 under different operating conditions. This correspondence is then built into the gas turbine engine control system. When the gas turbine engine is working, by identifying the operating conditions, corresponding control parameters are sent to the actuator cylinder 17 to adjust the effective flow area between the annular cone orifice and the annular cone block, thereby controlling the cooling air flow rate and ensuring the performance of the gas turbine engine across the entire operating range.

[0065] To ensure the high reliability of the cooling bleed air structure for the turbine blades of the gas turbine engine core and the precise adjustment of the cooling air volume, avoid redundant cooling air volume, and reduce structural complexity, the design further divides the annular adjusting plate 14 into multiple arc-shaped segments. Corresponding to each arc-shaped segment, there are multiple bleed air pipes 15, transmission rods 16, and actuator cylinders 17. Multiple partition plates 20 are installed within the annular bleed air passage, with each partition plate 20 and each arc-shaped segment spaced circumferentially to partially block the annular bleed air passage. Figure 7 As shown.

[0066] 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.

[0067] 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 scope of protection 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 scope of protection of this application.

Claims

1. A cooling bleed air structure for the turbine blades of a gas turbine engine core, characterized in that, include: The gas collection hood (12) is fitted onto the outside of the outer wall of the combustion chamber (5), located at the rear end of the outer wall of the combustion chamber (5), and connected to the inside of the outer wall of the combustion chamber (5); The air intake ring (13) is sleeved on the outer periphery of the high-pressure shaft (4) and located inside the inner wall (6) of the combustion chamber. An annular air intake channel is formed inside it. The annular air intake channel forms an annular air intake cavity in the front end of the air intake ring (13) and an annular cone opening in the rear end of the air intake ring (13). The annular regulating plate (14) is inserted into the annular air intake channel through the annular cone opening at its front end and extends into the annular air intake chamber. The rear end is formed with an annular cone block. The air intake pipe (15) is connected to the inside of the air collection hood (12) at one end, and the other end is set through the outer wall (5) of the combustion chamber, the diffuser (7) and the inner wall (6) of the combustion chamber from the front end, and is connected to the annular air intake chamber. The transmission rod (16) is installed inside the air duct (15), with one end extending into the annular air duct cavity and connected to the front toothed part of the annular adjusting plate (14); The actuator (17) is installed outside the outer wall (5) of the combustion chamber. Its piston rod extends into the air intake pipe (15) and is connected to the other end of the transmission rod (16) by toothed engagement. The piston rod can drive the transmission rod (16) to rotate by extension and retraction, thereby driving the annular adjustment plate (14) to move axially and adjust the effective flow area between the annular cone and the annular cone block.

2. The cooling bleed air structure for the turbine blades of the gas turbine engine core as described in claim 1, characterized in that, The annular conical opening at the rear end of the air intake ring (13) and the annular conical block at the rear end of its annular adjusting plate (14) have a straight guide section.

3. The cooling bleed air structure for the turbine blades of the gas turbine engine core as described in claim 1, characterized in that, Also includes: Multiple pairs of support blocks (18) are installed inside the annular air intake channel, supporting the annular adjustment plate (14) and the two side walls of the annular air intake channel.

4. The cooling bleed air structure for the turbine blades of the gas turbine engine core as described in claim 1, characterized in that, The part where the piston rod of the actuator cylinder (17) on the air intake tube (15) extends is in the shape of a boss.

5. The cooling bleed air structure for the turbine blades of the gas turbine engine core as described in claim 1, characterized in that, The bleed air ring (13) is connected to the combustion chamber wall (6) by fins and bolt fasteners.

6. The cooling bleed air structure for the turbine blades of the gas turbine engine core as described in claim 1, characterized in that, The actuator (17) is a hydraulic actuator; The cooling bleed air structure for the turbine blades of the gas turbine engine core also includes: A hydraulic control device (19) is connected to an actuator (17) and connected to a gas turbine engine control system.

7. The cooling bleed air structure for the turbine blades of the gas turbine engine core as described in claim 1, characterized in that, The annular regulating plate (14) is divided into multiple arc segments. Corresponding to each arc segment, there are multiple air intake pipes (15), transmission rods (16), and actuators (17). The gas turbine engine core turbine blade cooling bleed air structure also includes: Multiple partition plates (20) are set in the annular air intake channel and are distributed circumferentially at intervals with each arc segment.

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