Air-film-free turbine blade structure cooled through near-wall surface ring blade channel

Abstract

The invention discloses an air-film-free turbine blade structure cooled through a near-wall-surface ring blade channel, which comprises a blade body, a pressure surface and a suction surface of the blade body are not provided with air film holes, and a plurality of mutually independent ring blade cooling pipelines are distributed in the wall surface of the blade body along the wall surface. An inlet and an outlet of each ring blade cooling pipeline are connected through two independent vertical pipelines with one-way openings, and the opening directions of the two vertical pipelines are opposite. Cooling gas enters the cooling pipeline of the annular blade through one vertical pipeline, then completes heat exchange with the wall surface of the blade body, and is discharged through another vertical pipeline. Traditional film hole cooling is replaced by the near-wall ring blade cooling pipeline, uniform and efficient cooling of the whole area of the blade is achieved, cooling gas is prevented from being mixed with high-temperature mainstream gas, the thermodynamic cycle efficiency of the whole gas turbine is improved, meanwhile, the machining technology is simplified, the manufacturing cost is reduced, and the high-temperature endurance capacity and thermal fatigue resistance of the blade are enhanced.

Description

Technical Field

[0001] This invention relates to a film-free turbine blade structure cooled by a near-wall annular blade channel. Background Technology

[0002] Gas turbines are core power units in modern industry, aviation, and ship propulsion, boasting advantages such as high efficiency, high reliability, and operational flexibility. However, turbine blades, as key hot-end components of gas turbines, are subjected to extreme conditions of high temperature, high pressure, and corrosiveness for extended periods. Without efficient cooling measures, blade failure will directly occur, shortening the service life of the gas turbine and even causing equipment operation safety accidents. Therefore, developing efficient and reliable turbine blade cooling structures is the core key to improving the overall performance and operational reliability of gas turbines.

[0003] Currently, the mainstream cooling methods for gas turbine blades include film cooling, internal impact cooling, and ribbed channel cooling. Among these, film cooling is widely used due to its heat insulation effect. It involves creating multiple rows of tiny film cooling holes on the pressure and suction surfaces of the blade, allowing internal cooling airflow to exit from these holes and spread along the blade surface, forming a heat-insulating film that prevents direct contact between the high-temperature combustion gas and the metal wall. However, this cooling structure has several technical drawbacks:

[0004] Firstly, the cooling airflow is directly mixed with the high-temperature mainstream gas after being ejected from the film gas hole, which reduces the temperature of the downstream gas and weakens the energy input of the subsequent turbine stage, seriously affecting the overall thermodynamic cycle efficiency of the gas turbine.

[0005] Secondly, the cooling effect of film cooling is concentrated around the film vents, which can easily lead to uneven temperature distribution on the blade surface. High temperature concentration is likely to occur in the downstream area of ​​the film vents, resulting in poor cooling uniformity.

[0006] Third, the film cooling pores are tiny and precise structures with complex manufacturing processes, which significantly increases the manufacturing cost of the blades. Furthermore, the opening of the film cooling pores also increases the complexity of the blade structure design.

[0007] To address the technical problems of traditional film cooling blades, there is an urgent need to develop a new type of gas turbine blade cooling structure that, while ensuring cooling performance, reduces the mixing of cooling gas with mainstream combustion gas, improves cooling uniformity, simplifies processing technology, and reduces manufacturing costs. Summary of the Invention

[0008] The purpose of this invention is to overcome the defects of gas turbine film cooling blades in the prior art, such as uneven cooling, large heat loss, high manufacturing cost, and impact on the overall thermal efficiency, and to provide a gas turbine blade structure based on near-wall annular blade channel cooling.

[0009] This solution replaces traditional film cooling with a near-wall annular blade cooling pipe design, achieving efficient and uniform cooling of the blades, reducing heat loss, lowering manufacturing costs, reducing blade weight, and improving the overall performance and operational reliability of the gas turbine.

[0010] A film-free turbine blade structure cooled by a near-wall annular blade channel includes a blade body comprising a pressure side and a suction side.

[0011] Neither the pressure surface nor the suction surface of the blade body is provided with air film pores;

[0012] The blade body has several independent annular cooling pipes arranged along the inner wall of the blade body.

[0013] The inlets and outlets of several of the aforementioned annular blade cooling pipes are connected by two independent unidirectional vertical pipes, the opening directions of the two vertical pipes being opposite.

[0014] The cooling gas passes through the annular blade cooling pipes in sequence to exchange heat with the blade body wall, thereby cooling the blade.

[0015] Furthermore, several of the aforementioned annular blade cooling pipes are arranged at equal or non-equal intervals in the vertical direction of the blade body. The layout can be flexibly designed according to the actual heat load distribution in different blade height areas. Equal interval arrangement ensures the overall cooling uniformity of the blade, while non-equal interval arrangement can provide targeted enhanced cooling for areas with concentrated high-temperature heat loads, adapting to the heat dissipation requirements of the blade under different operating conditions.

[0016] Furthermore, the vertical pipe with its opening direction pointing upwards serves as the cooling gas outlet, and the vertical pipe with its opening direction pointing downwards serves as the cooling gas inlet.

[0017] The cooling gas outlet is located on the upper wall of the blade body. After the cooling gas is discharged through the cooling gas outlet, it does not mix with the high-temperature mainstream combustion gas inside the gas turbine. This fundamentally avoids heat loss caused by the mixing of cooling gas and high-temperature combustion gas, ensures the energy input of downstream combustion gas, and effectively improves the overall thermodynamic cycle efficiency of the gas turbine.

[0018] Furthermore, the pressure side of the blade body is a gentle / concave curved surface facing the direction of fluid flow, and the suction side is a curved / convex curved surface away from the direction of fluid flow.

[0019] The end where the pressure side and suction side transition smoothly is the leading edge, and the end away from the leading edge is the trailing edge. The angle between the pressure side and suction side at the trailing edge is narrowed.

[0020] This structure not only conforms to the hydrodynamic characteristics of gas turbine blades, reducing the flow resistance of fluid on the blade surface and ensuring the aerodynamic performance of the blades, but also adapts to the layout trajectory of the cooling pipes of the ring blades, ensuring the fit between the cooling pipes and the blade wall.

[0021] Furthermore, the vertical pipe has several air holes along its vertical direction, and the number of air holes matches the number of cooling pipes for the ring blades.

[0022] Furthermore, the width of the vertical pipe gradually decreases from the outer side to the trailing edge. This adapts to the narrowing structure of the blade's trailing edge, ensuring the overall compactness of the blade structure while avoiding additional interference from the vertical pipe to the fluid flow on the blade surface, maintaining the smoothness of the blade's flow channel, and balancing the compatibility of the cooling structure and the blade's aerodynamic structure.

[0023] Furthermore, the cooling pipes around the blades are installed close to the outer wall of the blade body. This minimizes the heat exchange distance between the cooling gas and the high-temperature blade wall, improves heat transfer efficiency, allows the cooling gas to quickly absorb heat from the blade wall, achieving efficient heat dissipation, and at the same time makes full use of the internal space of the blade wall, optimizing the blade structure layout.

[0024] Furthermore, the annular blade cooling pipe is a closed-loop ring pipe, and its annular trajectory is adapted to the curved contour of the blade body wall. This allows the cooling pipe to conform to the curved shape of the blade wall to achieve full circumferential coverage, ensuring that all areas of the blade wall can fully contact the cooling gas, achieving uniform circumferential cooling of the blade, and avoiding high temperature concentration in local areas due to inadequate cooling.

[0025] Compared with the prior art, the present invention has the following beneficial effects:

[0026] 1. High heat exchange efficiency and good cooling effect: Based on the removal of the air film pores in the blade, the present invention arranges the annular blade cooling pipes close to the wall inside the blade, which greatly increases the heat exchange area inside the blade. Moreover, the cooling pipes run through the blade height, are arranged vertically at equal intervals, and are distributed at equal intervals around the leading edge, so as to achieve uniform cooling of the entire blade area.

[0027] 2. Reduce heat loss and improve overall efficiency: This invention places the cooling gas outlet on the upper wall of the blade, so that the cooling gas does not mix with the high-temperature mainstream combustion gas after heat exchange, replacing the traditional method of direct mixing of cold gas and combustion gas in film cooling, which greatly reduces the amount of cold gas mixed with hot gas.

[0028] Numerical simulations have verified that the hot gas outlet temperature of the structure of this invention is higher than that of the traditional cooling structure, and it has smaller heat loss at different Reynolds numbers, effectively avoiding the problem of downstream gas temperature reduction and improving the overall thermodynamic cycle efficiency of the gas turbine.

[0029] 3. Simplified processing technology to achieve weight reduction and cost reduction: This invention eliminates the air film hole structure of traditional blades and eliminates the precision machining process of air film holes, which greatly simplifies the blade processing technology and reduces manufacturing costs; at the same time, the structural design of the near-wall ring blade cooling pipe reduces the metal consumption of the blades compared with the traditional cooling structure, thereby achieving blade weight reduction and optimizing the structural characteristics of the blades.

[0030] 4. Enhance the high-temperature resistance of blades and ensure safe operation of equipment: The ring blade cooling pipe design of this invention effectively improves the uniformity of temperature distribution on the blade surface, structurally solves the problem of high temperature concentration downstream of the film cooling holes of traditional blades, enhances the blade's resistance to extreme high-temperature and high-pressure conditions and its resistance to thermal fatigue, extends the service life of the blade, and thus ensures the safe, reliable and long-term operation of the gas turbine. Attached Figure Description

[0031] Figure 1 This is a schematic diagram of the main structure of this type of gas turbine blade;

[0032] Figure 2 This is a schematic diagram of the internal structure of this type of gas turbine blade;

[0033] Figure 3 This is a schematic diagram of the connection of a single ring blade cooling pipe;

[0034] Figure 4 It is a cross-sectional view of a gas turbine blade (the area of ​​the cooling pipes for the non-circular blade).

[0035] Figure 5 It is a cross-sectional view of a gas turbine blade (including the area of ​​the annular blade cooling pipes).

[0036] Figure 6 This shows the air inlet and outlet patterns of the ring blade cooling pipe from the side.

[0037] The diagram shows: 1. Annular blade cooling pipe, 100. Blade body, 101. Pressure side, 102. Suction side, 103. Leading edge, 104. Trailing edge, 105. Cooling gas outlet, 106. Cooling gas inlet. Detailed Implementation

[0038] To enhance understanding of the present invention, the present invention will be further described in detail below with reference to embodiments and accompanying drawings. These embodiments are only used to explain the present invention and do not constitute a limitation on the scope of protection of the present invention.

[0039] Example 1:

[0040] This embodiment provides a film-free turbine blade structure cooled by a near-wall annular blade channel, including a blade body 100. The pressure side 101 of the blade body 100 is a gently curved surface facing the direction of fluid flow, and the suction side 102 is a curved and convex surface facing away from the direction of fluid flow. The end where the pressure side 101 and the suction side 102 transition smoothly is the leading edge 103, and the end away from the leading edge 103 is the trailing edge 104. The angle between the pressure side 101 and the suction side 102 at the trailing edge 104 is narrowed.

[0041] No air film pores are opened on the pressure surface and suction surface of the blade body 100. 17 independent annular blade cooling pipes 1 are arranged along the inner wall of the blade body. The 17 annular blade cooling pipes 1 are arranged at equal intervals in the vertical direction of the blade body 100, and the annular blade cooling pipes 1 are set close to the outer wall of the blade body 100.

[0042] The inlets of the 17 annular blade cooling pipes 1 are all connected to a vertical pipe with a one-way opening, which opens downwards and serves as the cooling gas inlet 106. The outlets of the 17 annular blade cooling pipes 1 are all connected to another vertical pipe with a one-way opening, which opens upwards and serves as the cooling gas outlet 105. The cooling gas outlet 105 is located on the upper wall of the blade body 100 to ensure that the cooling gas does not mix with the high-temperature mainstream combustion gas inside the gas turbine after it is discharged.

[0043] On the vertical pipes corresponding to the cooling gas inlet 106 and the cooling gas outlet 105, 17 air holes are opened vertically. The number of air holes matches the number of annular blade cooling pipes 1. Each air hole corresponds to the inlet or outlet of an annular blade cooling pipe 1, realizing precise flow of cooling gas. At the same time, the width of the two vertical pipes gradually decreases from the outer side to the trailing edge 104 side, which is adapted to the narrowing structure of the trailing edge 104 of the blade and ensures the overall structural compactness of the blade.

[0044] The blade cooling process in this embodiment is as follows: Cooling gas enters the corresponding vertical pipe from the cooling gas inlet 106, and enters 17 independent annular blade cooling pipes 1 through 17 vents on the pipe. The cooling gas flows in the annular blade cooling pipes 1 close to the blade wall, and completes sufficient heat exchange with the high-temperature blade body 100 wall, achieving uniform cooling of the blade wall. After heat exchange, the cooling gas flows out from the outlet of the annular blade cooling pipe 1, enters the vertical pipe corresponding to the cooling gas outlet 105, and is finally discharged through the cooling gas outlet 105. The discharged cooling gas does not mix with the high-temperature mainstream combustion gas in the gas turbine, thus completing the entire blade cooling process.

[0045] The blade structure of this embodiment has been verified by numerical simulation. It achieves uniform cooling of the entire blade area under different Reynolds numbers, with no high-temperature concentrated areas on the blade surface and no mixing of cooling gas and high-temperature mainstream combustion gas. The heat loss is reduced by more than 30% compared with traditional film cooling blades, and the overall thermal cycle efficiency of the gas turbine is significantly improved. At the same time, this blade structure eliminates the precision machining of film cooling holes, simplifies the manufacturing process, reduces manufacturing costs by more than 25%, reduces metal consumption, and reduces blade weight by more than 15%. The mechanical strength and high-temperature resistance of the blade are greatly improved, and the thermal fatigue resistance is significantly enhanced.

[0046] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the protection scope of the present invention.

Claims

1. A film-free turbine blade structure cooled by a near-wall annular blade channel, comprising a blade body (100), the blade body (100) comprising a pressure side and a suction side, characterized in that: Neither the pressure surface nor the suction surface of the blade body (100) is provided with air film pores; The blade body (100) has several independent annular blade cooling pipes (1) arranged along the inner wall of the blade body; The inlets and outlets of several of the ring blade cooling pipes (1) are connected by two independent unidirectional vertical pipes with opposite opening directions. The cooling gas passes through the annular blade cooling pipe (1) and exchanges heat with the wall of the blade body (100) in sequence to achieve blade cooling.

2. A gas foil bearing pad structure cooled by near wall face ring blade passage according to claim 1, characterized in that, Several of the aforementioned annular blade cooling pipes (1) are arranged at equal or non-equal intervals in the vertical direction of the blade body (100).

3. A near-wall ring passage cooled bladed turbomachine vane structure according to claim 1, wherein, The vertical pipe with its opening direction facing upward is the cooling gas outlet (105), and the vertical pipe with its opening direction facing downward is the cooling gas inlet (106). The cooling gas outlet (105) is located on the upper wall of the blade body (100). After the cooling gas is discharged through the cooling gas outlet (105), it does not mix with the high-temperature mainstream gas in the gas turbine.

4. A near-wall annulus vane passage cooled gas foil bearing vane structure according to any one of claims 1-3, wherein, The pressure side of the blade body (100) is a gentle / concave curved surface facing the direction of fluid flow, and the suction side is a curved / convex curved surface facing away from the direction of fluid flow. The end where the pressure side and suction side transition smoothly is the leading edge (103), and the end away from the leading edge (103) is the trailing edge (104). The angle between the pressure side and the suction side at the trailing edge (104) is narrowed.

5. A gas foil bearing leaf structure cooled by a near wall face ring leaf passage according to claim 4, characterized in that, The vertical pipe has several air holes along its vertical direction, and the number of air holes matches the number of the ring blade cooling pipes (1).

6. A gas foil bearing leaf structure cooled by a near wall face ring leaf passage according to claim 4, characterized in that, The width of the vertical pipe gradually decreases from the outer side to the tail edge (104).

7. A gas foil bearing leaf structure cooled by a near wall face ring leaf passage according to claim 4, wherein, The cooling pipe (1) of the ring blade is set close to the outer wall of the blade body (100).

8. A gas foil bearing leaf structure cooled by a near wall face ring leaf passage according to claim 4, characterized in that, The annular blade cooling pipe (1) is a closed annular pipe, and its annular trajectory is adapted to the curved contour of the blade body (100) blade wall.

Images