Cooling structure for high-temperature area of turbine first-stage stationary blade of heavy gas turbine

Abstract

The invention relates to a cooling structure for a high-temperature area of a turbine first-stage stationary blade of a heavy-duty gas turbine, and aims to solve the problem that accurate cooling coverage can be realized according to a structural form for cooling the high-temperature area of a blade, a plurality of combustion chamber outlets are connected end to end in the circumferential direction, and the left upper corner and the right upper corner of each combustion chamber outlet are both unfilled; a cold air flow channel is formed by the adjacent unfilled corner positions of the outlets of the two adjacent combustion chambers and an area defined by the inner wall of the heavy-duty gas turbine, the flow channel comprises a main path and a plurality of branch paths communicated with the main path, the main path of the flow channel is an inlet end, the branch paths of the flow channel are outlet ends, the inlet end of the flow channel is connected with the cold air flow channel, and the outlet end of the flow channel is connected with the cold air flow channel. According to the scheme, the relative high temperature of the front edges of the first-stage stationary blades can be reduced in a targeted mode, local overheating is avoided, and the running safety of the gas turbine is guaranteed.

Description

Technical Field

[0001] This invention relates to the field of heavy-duty gas turbine technology, specifically to a cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine. Background Technology

[0002] The leading edge of the first-stage stator blades in heavy-duty gas turbines typically withstands the impact of high-temperature combustion gases ranging from 1100 to 1400°C. This is especially true in the area directly in front of the stator blades, near the upper endwall where the airflow is directed, leading to thermal stress deformation and ablation. This makes this area a critical zone for ablation damage, necessitating a structural design capable of precisely cooling this high-temperature region while maintaining structural strength and manufacturing feasibility to address the shortcomings of existing technologies. Summary of the Invention

[0003] The purpose of this invention is to address the urgent need for a structural design that can achieve precise cooling coverage for the high-temperature region of the blades. This invention proposes a cooling structure for the high-temperature region of the first stage stator blades in a heavy-duty gas turbine.

[0004] The objective of this invention is achieved as follows: a cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine, comprising multiple combustion chamber outlets, multiple flow channels, and multiple first-stage stator blades;

[0005] Multiple combustion chamber outlets are connected end to end along the circumference. Each combustion chamber outlet has a notch at the upper left and upper right corners. The adjacent notch positions of two adjacent combustion chamber outlets and the area enclosed by the inner wall of the heavy-duty gas turbine form a cold air flow channel. The flow channel includes a main channel and multiple branch channels connected to the main channel. The main channel is the inlet end, and the branch channels are the outlet ends. The inlet end of the flow channel is connected to the cold air flow channel. Each outlet end of the flow channel faces the upper part of a first-stage stator vane. The combustion chamber outlets, flow channels, and first-stage stator vanes are arranged sequentially from front to back in the heavy-duty gas turbine.

[0006] Furthermore, a cold air passage is provided inside the first stage stationary blade, and the branch of the flow passage is connected to the first stage stationary blade, and the branch of the flow passage is connected to the cold air passage of the first stage stationary blade.

[0007] Furthermore, a gap is left between the branch channels of the flow channel and the primary stationary blade.

[0008] Furthermore, the ratio of the number of combustion chamber outlets to the number of first-stage stator vanes is 1:2.

[0009] Furthermore, the two first-stage stator vanes are located near the 25% and 75% circumferential percentages of the combustion chamber outlet, respectively.

[0010] Furthermore, the flow channel is a Y-shaped flow channel.

[0011] Furthermore, the profiles at the upper left and upper right corners of the combustion chamber outlet are both sloped surfaces, and the area enclosed by the adjacent corners of the adjacent combustion chamber outlets and the inner wall of the heavy-duty gas turbine is triangular.

[0012] Furthermore, the cross-sections of both the main and branch channels are triangular.

[0013] Furthermore, the upper left and upper right corners of the combustion chamber outlet are concave arc surfaces, and the area enclosed by the adjacent corners of the adjacent combustion chamber outlets and the inner wall of the heavy-duty gas turbine is semi-circular.

[0014] Furthermore, the main channel cross-section is semi-circular, while the branch channel cross-sections are approximately fan-shaped.

[0015] Beneficial effects:

[0016] This invention precisely introduces cooling gas to directly cool a specific area of ​​the first-stage stator blade. Unlike previous methods that dispersed cooling gas through a ring of cooling holes at the combustion chamber outlet, where the gas reaches the high-temperature zone of the first-stage stator blade due to its already high temperature, this invention specifically reduces the relatively high temperature generated at the leading edge of the first-stage stator blade, avoiding localized overheating and ensuring the safety of the gas turbine operation. By adjusting the ratio of the combustion chamber outlet and the number of first-stage stator blades to one-to-many, the hot spots at the combustion chamber outlet are directed to enter the flow channel between the two first-stage stator blades, preventing them from blowing directly onto the blade body and causing overheating. The overall design is compact and reasonable, exhibiting high cooling efficiency and adaptability. Attached Figure Description

[0017] Figure 1 This is a schematic diagram of the cold air flow direction in the flow channel of the present invention;

[0018] Figure 2 This is a schematic diagram showing the combustion chamber outlet notch profile with a straight edge;

[0019] Figure 3 This is a schematic diagram showing the arc-shaped notch at the combustion chamber outlet.

[0020] Figure 4 This is a schematic diagram showing that the flow channel has a triangular cross-section;

[0021] Figure 5 This is a schematic diagram showing the connection between the flow channel of the present invention and the internal cooling air passage of the first-stage blade;

[0022] Figure 6 This is a perspective view showing the flow channel of the present invention communicating with the internal cooling air passage of the first-stage blade;

[0023] Figure 7 This is a schematic diagram showing that the flow channel and the first-stage blade of the present invention have a gap;

[0024] Figure 8This is a three-dimensional view of the flow channel and the first-stage blade of the present invention with a gap. Detailed Implementation

[0025] Specific implementation method 1: A cooling structure for the high-temperature zone of the first stage stator blade of a heavy-duty gas turbine, which includes multiple combustion chamber outlets 1, multiple flow channels 2 and multiple first stage stator blades 3;

[0026] Multiple combustion chamber outlets 1 are connected end to end along the circumference. Each combustion chamber outlet 1 has a notch at the upper left and upper right corners. The adjacent notch positions of two adjacent combustion chamber outlets 1 and the area enclosed by the inner wall of the heavy-duty gas turbine form a cold air flow channel. The flow channel 2 includes a main channel and multiple branch channels connected to the main channel. The main channel of the flow channel 2 is the inlet end, and the branch channels of the flow channel 2 are the outlet ends. The inlet end of the flow channel 2 is connected to the cold air flow channel. Each outlet end of the flow channel 2 faces the upper part of a first-stage stator vane 3. The combustion chamber outlets 1, the flow channel 2, and the first-stage stator vane 3 are arranged sequentially from front to back in the heavy-duty gas turbine.

[0027] In this embodiment, the upper left and upper right corners of the combustion chamber outlet are changed from conventional right-angle structures to corner-cut structures with a gap. The cold air enters the cold air through the cold air flow channel. After entering the main flow channel, the cold air enters the branch flow channel and finally blows towards the upper part of the first stage stator vane. The hot air flowing out of the combustion chamber outlet blows towards the middle and lower part of the first stage stator vane.

[0028] Other implementation methods are the same as those in Specific Implementation Method 1.

[0029] Specific implementation method 2: A cooling structure for the high temperature zone of the first stage stationary blade of a heavy-duty gas turbine, wherein a cold air channel is provided inside the first stage stationary blade 3, and a branch of the flow channel 2 is connected to the first stage stationary blade 3, and the branch of the flow channel 2 is connected to the cold air channel of the first stage stationary blade 3.

[0030] In this embodiment: Optionally, the cold air passage is set inside the first stage stationary blade along the length direction of the first stage stationary blade. The cold air flowing through the cold air passage is blown into the upper part of the first stage stationary blade and enters the cold air passage inside the first stage stationary blade. The hot air flowing out of the combustion chamber outlet is blown towards the middle and lower part of the first stage stationary blade.

[0031] Other implementation methods are the same as those in Specific Implementation Method 1.

[0032] Specific implementation method three: A cooling structure for the high-temperature zone of the first stage stator blade of a heavy-duty gas turbine, wherein a gap is left between the branch of the flow channel 2 and the first stage stator blade 3.

[0033] In this embodiment: the cold air flowing through the cold air passage blows towards the upper part of the first stage stator blade to cool its outer surface, and the hot air flowing out of the combustion chamber outlet blows towards the middle and lower part of the first stage stator blade.

[0034] Other implementation methods are the same as those in Specific Implementation Method 1.

[0035] Specific implementation method four: A cooling structure for the high-temperature zone of the first stage stationary blade of a heavy-duty gas turbine, wherein the ratio of the number of combustion chamber outlet 1 to the number of first stage stationary blades 3 is 1:2.

[0036] In this embodiment, the number of combustion chamber outlets and first stage stationary blades is arranged in a 1:2 ratio to avoid hot spots.

[0037] Other implementation methods are the same as those in Specific Implementation Method 1.

[0038] Specific implementation method five: A cooling structure for the high-temperature zone of the first stage stator blades of a heavy-duty gas turbine, wherein the two first stage stator blades 3 are located near the 25% and 75% circumferential percentages of the combustion chamber outlet 1, respectively.

[0039] In this embodiment, the two first-stage stator blades are located near the 25% and 75% circumferential percentages of the combustion chamber outlet, respectively. This allows the hot spot at the combustion chamber outlet to enter the flow channel primarily between the two first-stage stator blades, preventing it from being blown onto the first-stage stator blade body and causing overheating.

[0040] Other implementation methods are the same as those in Specific Implementation Method Four.

[0041] Specific implementation method six: A cooling structure for the high-temperature zone of the first stage stator blade of a heavy-duty gas turbine, wherein the flow channel 2 is a Y-shaped flow channel.

[0042] In this embodiment: the flow channel has a Y-shaped structure, the main flow channel is welded to the combustion chamber outlet, and the branch flow channel faces the first stage stator vane.

[0043] Other implementation methods are the same as those in Specific Implementation Method Four.

[0044] Specific Implementation Method Seven: A cooling structure for the high-temperature zone of the first stage stator blade of a heavy-duty gas turbine, wherein the profiles at the upper left and upper right corners of the combustion chamber outlet 1 are both inclined planes, and the area enclosed by the adjacent corner positions of the adjacent combustion chamber outlet 1 and the inner wall of the heavy-duty gas turbine is triangular.

[0045] In this embodiment, the upper left and upper right corners of the combustion chamber outlet are changed from the conventional right-angle structure to a notched corner structure with a gap. There will be an additional near-triangular gap at the top of the outlets of two adjacent combustion chambers, through which some cooling gas flows out and enters the flow channel.

[0046] Other implementation methods are the same as those in Specific Implementation Method 1.

[0047] Specific Implementation Method 8: A cooling structure for the high-temperature zone of the first stage stator blade of a heavy-duty gas turbine, wherein the cross-section of the main path and the branch path of the flow channel 2 are both triangular.

[0048] In this embodiment, the flow channel is designed to fit the additional triangular gaps at the top of the outlets of two adjacent combustion chambers.

[0049] Other implementation methods are the same as those in Specific Implementation Method Seven.

[0050] Specific Implementation Method Nine: A cooling structure for the high-temperature zone of the first stage stator blade of a heavy-duty gas turbine, wherein the notch at the upper left and upper right corners of the combustion chamber outlet 1 is a concave arc surface, and the area enclosed by the adjacent notch positions of the adjacent combustion chamber outlet 1 and the inner wall of the heavy-duty gas turbine is a semi-circle.

[0051] In this embodiment, the upper left and upper right corners of the combustion chamber outlet are changed from the conventional right-angle structure to a notched corner structure with a gap. There will be an extra semi-circular gap at the top of the outlets of two adjacent combustion chambers, through which some cooling gas flows out and enters the flow channel.

[0052] Other implementation methods are the same as those in Specific Implementation Method 1.

[0053] Specific Implementation Method 10: A cooling structure for the high-temperature zone of the first stage stator blade of a heavy-duty gas turbine, wherein the main cross-section of the flow channel 2 is semi-circular and the branch cross-section of the flow channel 2 is approximately fan-shaped.

[0054] In this embodiment, the main inlet of the flow channel is designed to accommodate the additional semi-circular gaps above the outlets of the two adjacent combustion chambers.

[0055] Other implementation methods are the same as those in specific implementation method nine.

[0056] Working principle:

[0057] In turbine design, as can be seen from the zweifel number, too few blades will increase the load on a single blade, easily leading to fatigue damage; too many blades will result in narrow flow channels, increasing manufacturing difficulty and cost. Compared with a 1:1 design of the number of combustion chamber flame tubes and the number of first-stage stator blades, this 1:2 design is closer to the common blade number design and can bear sufficient blade load.

[0058] A stream of cold air is introduced from near the combustion chamber outlet transition section and directly connected to the area in front of the first-stage stationary blades of the gas turbine, near the upper endwall, to prevent localized overheating in this area. Alternatively, the air stream can be directed away from the first-stage blades and only cool the outer surface of the area near the upper endwall.

[0059] The two primary stator blades are located near the 25% and 75% circumferential percentages of the combustion chamber outlet, respectively. This allows the hot spot at the combustion chamber outlet to enter the flow channel primarily between the two primary stator blades, preventing it from being blown onto the primary stator blade body and causing overheating.

Claims

1. A cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine, characterized in that: It includes multiple combustion chamber outlets (1), multiple flow channels (2), and multiple first stage stator vanes (3); Multiple combustion chamber outlets (1) are connected end to end along the circumference. Each combustion chamber outlet (1) has a missing corner at the upper left and upper right corners. The adjacent missing corner positions of two adjacent combustion chamber outlets (1) and the area enclosed by the inner wall of the heavy-duty gas turbine form a cold air flow channel. The flow channel (2) includes a main road and multiple branch roads connected to the main road. The main road of the flow channel (2) is the inlet end, and the branch roads of the flow channel (2) are the outlet ends. The inlet end of the flow channel (2) is connected to the cold air flow channel. Each outlet end of the flow channel (2) faces the upper part of a first stage stator (3). The combustion chamber outlets (1), flow channels (2) and first stage stator (3) are arranged sequentially from front to back in the heavy-duty gas turbine.

2. The cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine according to claim 1, characterized in that: The first stage stationary blade (3) is equipped with a cold air passage, and the branch of the flow channel (2) is connected to the first stage stationary blade (3), and the branch of the flow channel (2) is connected to the cold air passage of the first stage stationary blade (3).

3. The cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine according to claim 1, characterized in that: There is a gap between the branch of the flow channel (2) and the first stage stationary blade (3).

4. The cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine according to claim 1, characterized in that: The ratio of the number of combustion chamber outlets (1) to the number of first stage stator vanes (3) is 1:

2.

5. The cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine according to claim 4, characterized in that: The two first stage stator vanes (3) are located near the circumferential percentages of the combustion chamber outlet (1) at 25% and 75%, respectively.

6. The cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine according to claim 4, characterized in that: The flow channel (2) is a Y-shaped flow channel.

7. The cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine according to claim 1, characterized in that: The upper left and upper right corners of the combustion chamber outlet (1) are both sloped, and the adjacent corners of the combustion chamber outlet (1) and the area enclosed by the inner wall of the heavy-duty gas turbine form a triangle.

8. The cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine according to claim 7, characterized in that: The cross-sections of the main path and branch path of the flow channel (2) are both triangular.

9. The cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine according to claim 1, characterized in that: The upper left and upper right corners of the combustion chamber outlet (1) are concave arc surfaces, and the adjacent corners of the combustion chamber outlet (1) and the area enclosed by the inner wall of the heavy-duty gas turbine are semi-circular.

10. The cooling structure for the high-temperature zone of the first-stage stator blades of a heavy-duty gas turbine according to claim 9, characterized in that: The main cross section of the flow channel (2) is semi-circular, and the branch cross section of the flow channel (2) is approximately fan-shaped.

Images