A cooling air bleed air system and bleed air method

Abstract

The application discloses a cooling air bleeding system and a bleeding method. The cooling air bleeding system comprises a cooling passage, which is a closed passage and can transmit cooling air generated from a cooling air source generating device to a hot end component needing cooling in sequence; the cooling passage is located in a central closed area of a large system; the cooling air source generating device is a compressor; and the hot end component is a turbine wheel disc. The cooling passage comprises an intermediate shaft sleeve pipe. Cooling air is guided from the compressor to the center hole of the wheel disc in a radial direction, is guided along the center holes of the following stages after bleeding, is guided to the center hole of the turbine wheel disc through the intermediate shaft sleeve pipe, and is guided to the hot end component or a chamber needing cooling, so that the temperature of the turbine hot end component is reduced. The cooling air bleeding system and the bleeding method can be widely used in gas turbines and can effectively solve the technical problems of the existing bleeding difficulty and low efficiency of the bleeding from the compressor to the turbine.

Description

Technical Field

[0001] This invention pertains to air intake systems and methods, specifically a cooling gas intake system and method, and more particularly to a cooling gas intake system and method for a gas turbine tie rod combined rotor. This air intake system and method can be widely used in gas turbines. When a gas turbine uses this cooling gas intake system and method, it effectively solves the existing technical problems of difficulty and low efficiency in bleeding air from the compressor into the turbine. Background Technology

[0002] For gas turbine cycles, increasing the turbine's inlet temperature can improve operating efficiency and power, but it also causes the temperature of the turbine's hot-end components (mainly turbine blades) to exceed the material's tolerance range, resulting in higher thermal stress. This places higher demands on the cooling of the turbine's hot-end components.

[0003] During operation, the temperature of a gas turbine gradually increases, especially in the turbine itself. Taking an "F"-class gas turbine as an example, the turbine inlet temperature can reach approximately 1400°C. Therefore, in addition to using high-quality materials, the turbine components require complex cooling designs to ensure the turbine's safety and lifespan. Currently, gas turbines typically cool the turbine by drawing gas from the compressor; therefore, the rationality of the cooling design significantly impacts the gas turbine's efficiency.

[0004] In the prior art, CN104675440A discloses "a cooling gas induced draft structure for a gas turbine rotor". This cooling gas induced draft structure for a gas turbine rotor includes sealed chambers respectively disposed on the outer sides of the first-stage rotor and the last-stage rotor. No rotor cooling gas channels are provided on the rotors between the Nth and N+1th stage inter-stage chambers. Thus, the sealed chambers on the outer sides of the first-stage rotor sequentially introduce cooling gas in a forward direction from the first-stage inter-stage chamber to the Nth-stage inter-stage chamber; the sealed chambers on the outer sides of the last-stage rotor sequentially introduce cooling gas in a reverse direction from the last inter-stage chamber to the N+1th stage inter-stage chamber. While this prior art achieves the function of staged cooling gas induced draft for the turbine rotor, reducing the amount of high-pressure cooling gas used and improving the operating efficiency of the gas turbine to some extent, its structure is complex, and staged cooling gas induced draft is difficult due to the inherent complexity of the turbine rotor's structure. Summary of the Invention

[0005] Currently, there are technical problems such as the complexity of turbine rotor structure making staged air intake difficult, and the low efficiency of existing methods for absorbing air from the compressor into the turbine. The purpose of this invention is to overcome these shortcomings by providing a cooling gas evacuation system and method, thereby effectively solving the aforementioned technical problems.

[0006] The present invention provides a cooling gas induced draft system, characterized in that the system includes a cooling passage, which is a closed passage capable of sequentially transmitting cooling gas generated from a cold gas source generating device to the hot end component requiring cooling.

[0007] Furthermore, to further enhance the convenience of cooling gas transmission and control, the cooling gas induced draft system is characterized in that: the cooling passage included in the induced draft system is located in the central closed area of ​​the large system; the cooling gas source generating device is a compressor; the hot end component is a turbine wheel; wherein, the cooling passage includes an intermediate shaft sleeve; cooling gas is drawn from the compressor radially to the center hole of the wheel, along the center holes of each stage after induced draft, through the intermediate shaft sleeve, and from the center hole of the turbine wheel to the hot end component or the chamber to be cooled, thereby reducing the temperature of the turbine hot end component.

[0008] Furthermore, the cooling air induced draft system is characterized in that: the induced draft system includes a gas turbine, which has a rod-coupled rotor cooling air induced draft structure, which includes a first-stage compressor final stage impeller, a turbine first-stage impeller, an intermediate shaft sleeve and its inlet and outlet sealing assemblies.

[0009] Furthermore, the cooling air induced draft system is characterized in that: the induced draft system includes a gas turbine, which has a rod-coupled rotor cooling air induced draft structure, which includes a first-stage compressor final stage impeller and intermediate shaft, a turbine first-stage impeller and intermediate shaft, an intermediate shaft sleeve and its inlet and outlet sealing assemblies.

[0010] Furthermore, to ensure a good sealing effect, the cooling gas induced draft system is characterized in that: the induced draft system includes sealing components on both sides, and the impeller and the intermediate bushing of the induced draft system have the following connection and assembly relationship (one side sealing component, the other side sealing component), wherein, according to the assembly relationship between the impeller and the intermediate bushing, the sealing component on one side (the intermediate bushing is connected to the impeller), that is, the sealing component on the side where the intermediate bushing is connected to the impeller, is assembled with the impeller by a set of screws, and a locking washer is used to prevent the screws from loosening, or it is directly press-fitted onto the impeller using a heat fitting process; moreover, in terms of structural design, the positioning stop with a flange or the heat fitting curved surface can play a good sealing role; The other sealing assembly (the sealing assembly on the other side of the intermediate shaft sleeve) engages with the impeller via a C-shaped sealing ring, and is secured with a C-shaped sealing ring cover plate and cover plate screws. Furthermore, the structural design ensures that the C-shaped sealing ring engages with the corresponding impeller throughout the entire operation of the gas turbine, thus providing a seal. Simultaneously, the axial clearance between the C-shaped sealing ring cover plate screws and the impeller is less than the length of the screws (a loosening device or measure is provided), preventing the screws from falling out during operation and thus preventing loosening. This clearance also releases the axial expansion of the intermediate shaft sleeve after heating, ensuring that the stress of the intermediate shaft sleeve meets requirements.

[0011] Furthermore, in the gas turbine of the present invention, the inner dimensions of the intermediate shaft sleeve are highly flexible. From the perspective of improving efficiency, the cooling gas intake channel can be smoothly transitioned from the central hole of the compressor disk to the central hole of the turbine disk (such as gradual transition in each stage, or equal arithmetic or equal proportional setting in each stage), thereby reducing pressure loss.

[0012] Furthermore, the cooling gas ducting system is characterized in that: the cooling gas ducting channel flows from the central hole of the compressor disc to the central hole of the turbine disc in a transitional manner with each stage gradually changing, or with each stage being arithmetic or proportional, and finally flows to the hot end component.

[0013] Furthermore, the cooling gas evacuation system is characterized in that: the cooling gas extraction position of the compressor in the evacuation system can be freely set or matched automatically according to the needs of the hot end; due to the good sealing performance of the sealing components on both sides, and the fact that the entire cooling gas flow channel is located inside the rotor, the aerodynamic characteristics of the entire flow channel are easier to control, and the compressor supply pressure can be accurately matched with the cooling gas supply demand of the turbine hot end components, thereby locking the compressor extraction position, that is, determining the stage of the compressor radial extraction wheel.

[0014] Furthermore, the cooling gas induced draft system is characterized in that: the cooling gas extraction position can better avoid high-stress areas; the structures that mate with the intermediate shaft sleeve on the compressor disk exhaust side and turbine disk in the gas turbine need to be close to the exhaust side and intake side, respectively; due to the presence of blades and other structures, the compressor disk and turbine disk have high tangential stress in their central holes, so the parts of their structures that mate with the intermediate shaft sleeve must be far away from this high-stress area to reduce the stress level of the intermediate shaft sleeve.

[0015] Furthermore, the pressure and temperature of the cooling gas can also be adjusted. For example, the cooling gas is generally introduced from the compressor into the hot end components of the turbine. The induced gas locations in the compressor include after a specific stage of the compressor (divided into low, medium, and high pressure extraction according to the specific situation of the compressor) and induced gas from the compressor outlet diffuser. The hot end components of the turbine are generally supplied with air radially from the center hole of the turbine disc through the induced gas hole to the turbine groove and in front of the first stage turbine disc to the turbine groove. Cooling gas is supplied to the blades through the cooling holes at the bottom of the blade roots. In all the air supply channels, the air is radially drawn from a certain stage of the compressor to the center hole of the compressor disc, along the center hole of the compressor disc, through the intermediate shaft sleeve, and from the center hole of the turbine disc to the chamber that needs cooling gas. The air supply line is always inside the rotor. Compared with the flow path area, it is simpler, easier to solve the sealing problem, and can better control the supply pressure.

[0016] Furthermore, the cooling gas induced draft system is characterized in that: the induced draft system can be applied to a power generation system consisting of a gas turbine combination.

[0017] Another aspect of the present invention is: a method for bleed air into the cooling bleed air system, which improves the safety and efficiency of bleed air, characterized in that the method includes the following steps: Step A: Set up a cooling passage, which is a closed passage located in the central closed area of ​​the large system; Step B: The cold air source generating device is a compressor, and the hot end component is a turbine wheel. The cooling passage includes an intermediate shaft sleeve. Cooling air is drawn from the radial air intake wheel of the compressor to the center hole of the wheel, and then passes through the intermediate shaft sleeve along the center holes of each stage after air intake. Step C: Based on the actual cooling gas requirements of the hot end, the gas is delivered from the central hole of the turbine disk to the hot end component or the chamber requiring cooling, thereby reducing the temperature of the turbine hot end component.

[0018] Compared with the prior art, the present invention has the following features and beneficial effects: (1) The cooling gas induced draft system and induced draft method of the present invention can be widely used in gas turbines. When the gas turbine uses the cooling gas induced draft system and method, it can effectively solve the existing technical problems of difficulty and low efficiency in induced draft of the compressor into the turbine. (2) The cooling gas induced air system and induced air method of the present invention have a simple structure, are easy to operate, and are highly practical. Attached Figure Description

[0019] Figure 1 This is a schematic diagram of the cooling path of the gas turbine according to an embodiment of the present invention; Figure 2 This is a schematic diagram of the air intake side structure (interference fit) of the inter-shaft sleeve in an embodiment of the present invention. Figure 3 This is a schematic diagram of the air inlet side structure (threaded fastening) of the inter-shaft sleeve in an embodiment of the present invention; Figure 4 This is a schematic diagram of the exhaust side structure (sealing ring) of the inter-shaft sleeve in an embodiment of the present invention; In the diagram: 1-Compressor final stage impeller or intermediate shaft, 2-Turbine first stage impeller or intermediate shaft, 3-Intermediate shaft sleeve, 4-Compressor radial bleed impeller, 5-Intermediate shaft, 6-Hot end component (turbine second stage impeller), 7-Compressor impeller cooling hole, 8-Turbine impeller cooling hole 1, 9-Turbine impeller cooling hole 2, 10-Compressor impeller and sleeve mating structure, 11-Positioning stop, 12-Piece stop washer, 13-Intermediate shaft sleeve screw, 14-Turbine impeller and sleeve mating structure, 15-C-type sealing ring, 16-C-type sealing ring cover plate, 17-C-type sealing ring cover plate screw, 18-C-type sealing ring cover plate axial clearance. Detailed Implementation

[0020] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. However, this should not be construed as limiting the scope of the present invention to the above description. Various substitutions or modifications based on ordinary technical knowledge and conventional methods in the art should be included within the scope of the present invention without departing from the above-described spirit.

[0021] Example: See attached diagram.

[0022] See Figure 1 The present invention provides a cooling air induced draft structure for a gas turbine tie rod combined rotor, including a compressor final stage disc or intermediate shaft, a turbine first stage disc or intermediate shaft, an intermediate shaft sleeve, and an intermediate shaft sleeve inlet and outlet sealing assembly.

[0023] See Figure 3 The compressor's final stage impeller or intermediate shaft needs to be connected to the intermediate shaft sleeve. Its exhaust side is designed with a structure that mates with the sleeve. This structure includes a radial stop and a ring of evenly distributed threaded holes circumferentially. The diameter, pitch circle diameter, and number of threaded holes are determined based on the diameter of the compressor impeller's center hole. The intermediate shaft sleeve's inlet side has a flange structure, which also features a corresponding stop and threaded holes. During assembly, the compressor's final stage impeller or intermediate shaft and the intermediate shaft sleeve are positioned by the stop, assembled using intermediate shaft sleeve screws, and secured with locking washers to prevent loosening.

[0024] See Figure 4The turbine's first-stage rotor or intermediate shaft needs to be connected to the intermediate shaft sleeve. Its inlet side features a rotor-sleeve mating structure, while the intermediate shaft sleeve's exhaust side has a threaded hole. A C-shaped sealing ring is fixed to the outside using a C-shaped sealing ring cover plate and C-shaped sealing ring cover plate screws, ensuring the outer diameter of the C-shaped sealing ring is higher than the outer diameter of the intermediate shaft sleeve. The inner and exhaust sides are in close contact with the intermediate shaft sleeve and C-shaped sealing ring cover plate, respectively. During assembly, the turbine's first-stage rotor or intermediate shaft is mated with the C-shaped sealing ring, causing deformation and pressing tightly against the inner side of the rotor-sleeve mating structure, ensuring constant contact during rotor operation and thus achieving a seal. Because the inlet side of the sleeve is threaded to the compressor rotor, its axial expansion after heating needs to be released through the axial gap of the C-shaped sealing ring cover plate on the turbine side. This gap also prevents the C-shaped sealing ring cover plate from falling off, indirectly preventing loosening.

[0025] Based on the cooling requirements (supply pressure, flow rate, etc.) of the turbine hot-end components (first-stage turbine blades) and the dimensions and number of cooling channels (including cooling holes and the central hole of the turbine disc) allowed by the disc structure, and the determined parameters such as pressure loss of the cooling channels, the area before the compressor radial bleed turbine stage is selected as the evacuation point to provide the cooling air supply. Since the radial dimensions of the compressor disc and the turbine disc's central hole are equal, the channels inside the intermediate shaft sleeve transition parallel to each other.

[0026] During installation, install the intermediate shaft sleeve onto the intermediate shaft and secure it with intermediate shaft sleeve screws and locking washers to prevent loosening. Then, install the C-type seal ring in place and secure it with the C-type seal ring cover plate and C-type seal ring cover plate screws. Finally, assemble the turbine's first-stage impeller, ensuring the C-type seal ring fits snugly into the turbine impeller.

[0027] The technical contents of this invention and those not specifically described in the above embodiments are the same as those in the prior art.

[0028] The present invention is not limited to the above embodiments; all embodiments described herein can be implemented and have the aforementioned good effects.

Claims

1. A cooling gas evacuation system, characterized in that: The air intake system includes a cooling passage, which is a closed passage that can sequentially transmit cooling air generated from the cold air source generating device to the hot end components that need to be cooled. The cooling passage included in the air intake system is located in the central closed area of ​​the large system. The cold air source generating device is a compressor, and the hot end component is a turbine wheel. The cooling passage includes an intermediate shaft sleeve. Cooling air is drawn from the compressor radially to the central hole of the wheel, along the central holes of each stage after air intake, through the intermediate shaft sleeve, and from the central hole of the turbine wheel to the hot end component or the chamber to be cooled, thereby reducing the temperature of the turbine hot end component. The bleed air system includes sealing assemblies on both sides. One sealing assembly is assembled to the impeller with a set of screws, using locking washers to prevent screw loosening, or directly using a heat-fitting process for interference fitting onto the impeller. Structurally, the flanged positioning stop or heat-fitted curved surface provides a good seal. The other sealing assembly engages with the impeller using a C-type sealing ring, which is secured with a C-type sealing ring cover plate and cover plate screws. Structurally, by selecting an appropriate sealing ring size, the C-type sealing ring maintains its engagement with the corresponding impeller throughout the entire turbine operation, thus achieving a sealing function. Simultaneously, the axial clearance between the C-type sealing ring cover plate screws and the impeller is less than the screw length, preventing the screws from falling out during operation and thus preventing loosening. This clearance also releases the axial expansion of the intermediate shaft sleeve after heating, ensuring that the stress of the intermediate shaft sleeve meets requirements.

2. The cooling gas evacuation system according to claim 1, characterized in that: The bleed air system includes a gas turbine with a rod-coupled rotor cooling air bleed air structure, which includes a first-stage compressor final stage disc, a turbine first-stage disc, an intermediate shaft sleeve, and its inlet and outlet side sealing assemblies.

3. The cooling gas evacuation system according to claim 1, characterized in that: The bleed air system includes a gas turbine, which has a rod-coupled rotor cooling air bleed air structure. The rod-coupled rotor cooling air bleed air structure includes a first-stage compressor final stage disc and intermediate shaft, a turbine first-stage disc and intermediate shaft, an intermediate shaft sleeve and its inlet and exhaust side sealing components.

4. The cooling gas evacuation system according to claim 1, characterized in that: The cooling gas intake channel is set from the central hole of the compressor disc to the central hole of the turbine disc through a transition method with gradual changes, arithmetic progression, or proportional progression, and finally flows to the hot end components.

5. The cooling gas evacuation system according to any one of claims 1 to 4, characterized in that: The cooling gas extraction position of the compressor in the bleed air system can be freely set or matched automatically according to the needs of the hot end. Due to the good sealing performance of the sealing components on both sides and the fact that the entire cooling air flow channel is inside the rotor, the aerodynamic characteristics of the entire flow channel are easier to control. It can accurately match the compressor supply pressure from the cooling gas supply demand of the turbine hot end components, thereby locking the compressor extraction position, that is, determining the stage of the compressor radial extraction wheel.

6. The cooling gas evacuation system according to claim 5, characterized in that: The location of the cooling gas extraction can better avoid high-stress areas; the structure of the compressor disk exhaust side and turbine disk intake side of the gas turbine, which are connected to the intermediate shaft sleeve, should be close to the exhaust side and intake side, respectively.

7. The cooling gas evacuation system according to any one of claims 1 to 4, characterized in that: This bleed air system can be applied to a power generation system consisting of a gas turbine.

8. A method for induced draft in any of the cooling gas induced draft systems described in 1 to 7, characterized in that: The air-induction method includes the following steps: Step A: Set up a cooling passage. This cooling passage is a closed passage located in the central closed area of ​​the large system. This closed passage can transmit the cooling air generated from the cold air source equipment to the hot end components that need to be cooled in sequence. Step B: The cold air source generating device is a compressor, and the hot end component is a turbine wheel. The cooling passage includes an intermediate shaft sleeve. Cooling air is drawn from the radial air intake wheel of the compressor to the center hole of the wheel, and then passes through the intermediate shaft sleeve along the center holes of each stage after air intake. Step C: Based on the actual cooling gas requirements of the hot end, the gas is delivered from the central hole of the turbine disk to the hot end component or the chamber requiring cooling, thereby reducing the temperature of the turbine hot end component.

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