Turbine blade
By designing a turbine blade structure with a spiral rising blade tail and a cooling gas channel, combined with a thermal barrier coating, the problems of airflow loss and mechanical strength of turbine blades in high-temperature environments were solved, achieving efficient cooling and strength improvement.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Utility models(China)
- Current Assignee / Owner
- Filing Date
- 2025-04-22
- Publication Date
- 2026-03-06
AI Technical Summary
Existing turbine blades are difficult to simultaneously reduce airflow loss, improve efficiency, provide mechanical strength, and withstand centrifugal force and thermal stress in high-temperature environments.
The turbine blade structure is designed with a spirally rising blade tail, cooling gas passages, and a thermal barrier coating. A cooling gas film is formed through the air cooler and gas passages to reduce temperature and enhance mechanical strength.
It effectively reduces airflow loss, improves turbine efficiency, enhances the mechanical strength of the blades, and can withstand the centrifugal force and thermal stress generated by high-speed rotation.
Smart Images
Figure CN223975162U_ABST
Abstract
Description
Technical Field
[0001] This utility model relates to the field of gas turbine technology, specifically to a turbine blade. Background Technology
[0002] Because the combustion chamber outlet temperature of a gas turbine exceeds 1300℃, the turbine blades are constantly subjected to damage from the high-temperature airflow. To ensure the safe and reliable operation of the turbine, the turbine blades need to be cooled to maintain a reasonable temperature and stress distribution. However, reducing airflow losses and improving turbine efficiency while providing sufficient mechanical strength to withstand the centrifugal force and thermal stress generated by high-speed rotation makes the proper cooling of turbine blades a challenging problem that needs to be overcome in current technology. Utility Model Content
[0003] The purpose of this invention is to address the aforementioned shortcomings by providing a turbine blade that can reduce airflow loss, improve turbine efficiency, and provide sufficient mechanical strength to withstand the centrifugal force and thermal stress generated by high-speed rotation.
[0004] To achieve the above objectives, the present invention adopts the following technical solution:
[0005] A turbine blade includes a frame, a turbine shaft rotatably connected to the frame, multiple blade bodies mounted on the turbine shaft, a gas passage provided inside the turbine shaft, a rotating seat rotatably sleeved on the turbine shaft, the rotating seat being fixed to the frame, an annular air chamber provided inside the rotating seat, an air inlet pipe connected to the air chamber on one side of the rotating seat, multiple air inlets radially opened on the turbine shaft connecting the gas passage and the air chamber, a heat dissipation cavity provided inside the blade body, an air inlet passage connecting the heat dissipation cavity and the gas passage, multiple through holes connected to the heat dissipation cavity on both sides of the blade body, and a cooler connected to the air inlet pipe.
[0006] The blade body includes a blade root, one end of the blade root is a fixed part, and the other end of the blade root is a blade tail. The fixed part is connected to the turbine shaft. The blade tail is spirally rising along the rotation direction of the turbine blade. The front end of the blade tail is provided with a circular curved surface one, and the rear end of the blade tail is provided with a circular curved surface two. The radius of the circular curved surface one is larger than the radius of the circular curved surface two.
[0007] Furthermore, along the rotation direction of the blade body, the thickness of the front end of the blade tail is greater than the thickness of the rear end of the blade tail.
[0008] Furthermore, the helix angle of the blade body ranges from 30 to 40°.
[0009] Furthermore, the heat dissipation cavity is spirally ascending in shape, and the cross-sectional area of the heat dissipation cavity at the front end of the blade tail is larger than the cross-sectional area of the heat dissipation cavity at the rear end of the blade tail.
[0010] Furthermore, a thermal barrier coating is sprayed onto both the root and the tail of the blade.
[0011] The beneficial effects of this utility model are:
[0012] In practical applications, the turbine shaft is connected by a fixed part, and the spirally ascending blade tail adapts to changes in airflow at different radii, improving aerodynamic efficiency. The radius of the first circular curved surface at the front end of the blade tail is larger than the radius of the second circular curved surface at the rear end, reducing airflow resistance and losses. Cool air flows out from the inside through a cooler, intake pipe, air chamber, intake hole, gas channel, intake passage, heat dissipation cavity, and through-hole, forming a cooling gas film on the blade surface, effectively reducing the blade's temperature and enabling it to withstand the centrifugal force and thermal stress generated by high-speed rotation. This invention not only reduces airflow losses and improves turbine efficiency but also provides sufficient mechanical strength to withstand the centrifugal force and thermal stress generated by high-speed rotation. Attached Figure Description
[0013] Figure 1 This is a schematic diagram of the overall structure of this utility model;
[0014] Figure 2 This is a top view of the present invention;
[0015] Figure 3 This is a schematic diagram of the internal structure of the blade body in this utility model;
[0016] Reference numerals: 1. Turbine shaft; 2. Blade body; 201. Heat dissipation cavity; 202. Through hole; 2031. Circular curved surface one; 2032. Circular curved surface two; 204. Intake channel; 3. Rotating seat; 4. Intake pipe. Detailed Implementation
[0017] like Figure 1 , Figure 2 and Figure 3As shown, a turbine blade includes a frame, a turbine shaft 1 rotatably connected to the frame, multiple blade bodies 2 mounted on the turbine shaft 1, a gas passage provided inside the turbine shaft 1, a rotating seat 3 rotatably sleeved on the turbine shaft 1, the rotating seat 3 being fixed to the frame, an annular air chamber provided inside the rotating seat 3, an air inlet pipe 4 communicating with the air chamber on one side of the rotating seat 3, multiple air inlet holes radially opened on the turbine shaft 1 communicating with the gas passage and the air chamber, a heat dissipation cavity 201 provided inside the blade body 2, and an air inlet channel communicating with the heat dissipation cavity 201 and the gas passage. 204. The blade body 2 has multiple through holes 202 on both sides that connect to the heat dissipation cavity 201. A cooler is connected to the air intake pipe 4. The blade body 2 includes a blade root, one end of which is a fixed part, and the other end of which is a blade tail. The fixed part is connected to the turbine shaft 1. The blade tail is spirally ascending along the rotation direction of the turbine blade. The front end of the blade tail is provided with a circular curved surface 2031, and the rear end of the blade tail is provided with a circular curved surface 2032. The radius of the circular curved surface 2031 is larger than the radius of the circular curved surface 2032.
[0018] The turbine shaft 1 is connected by a fixed part, and the spirally rising blade tail adapts to changes in airflow at different radii, improving aerodynamic efficiency. The radius of the circular curved surface 2031 at the front end of the blade tail is larger than the radius of the circular curved surface 2032 at the rear end of the blade tail, reducing airflow resistance and reducing airflow loss. Cool air is drawn out from the inside by the cooler, intake pipe 4, air chamber, intake hole, gas channel, intake channel 204, heat dissipation cavity 201 and through hole 202, forming a cooling gas film on the surface of the blade body 2, effectively reducing the temperature of the blade body 2, and enabling the blade body 2 to withstand the centrifugal force and thermal stress generated by high-speed rotation. This utility model can not only reduce airflow loss and improve turbine efficiency, but also provide sufficient mechanical strength to withstand the centrifugal force and thermal stress generated by high-speed rotation.
[0019] like Figure 1 , Figure 2 and Figure 3 As shown, along the rotation direction of the blade body 2, the thickness of the front end of the blade tail is greater than the thickness of the rear end of the blade tail. In this embodiment, when the thickness of the front end of the blade tail is greater than the thickness of the rear end of the blade tail, airflow resistance can be reduced and airflow loss can be reduced.
[0020] like Figure 1 , Figure 2 and Figure 3 As shown, the helix angle of the blade body 2 is in the range of 30 to 40°; in this embodiment, when the helix angle of the blade body 2 is in the range of 30 to 40°, the overall airflow loss of the blade body 2 is minimized.
[0021] like Figure 1 , Figure 2 and Figure 3 As shown, the heat dissipation cavity 201 is spirally ascending in shape, and the cross-sectional area of the heat dissipation cavity 201 at the front end of the blade tail is larger than the cross-sectional area of the heat dissipation cavity 201 at the rear end of the blade tail. In this embodiment, when the heat dissipation cavity 201 is spirally ascending in shape, the shape of the heat dissipation cavity 201 is adapted to the blade body 2, and the wider the coverage of the cooling gas film, the better the cooling effect.
[0022] like Figure 1 , Figure 2 and Figure 3 As shown, a thermal barrier coating is sprayed on both the root and the tail of the blade. In this embodiment, the thermal barrier coating made by spraying yttrium-stabilized zirconium oxide (YSZ) can effectively reduce the surface temperature of the blade body 2 and reflect heat.
[0023] The specific embodiments described herein are merely illustrative examples of this utility model. Those skilled in the art to which this utility model pertains may make various modifications or additions to the described specific embodiments or use similar methods to replace them, without departing from the scope defined by this utility model.
Claims
1. A turbine blade, characterized by: The turbine comprises a frame, a turbine shaft rotatably connected to the frame, a plurality of blade bodies mounted on the turbine shaft, a gas passage arranged in the turbine shaft, a rotating seat rotatably sleeved on the turbine shaft, the rotating seat being fixed on the frame, an annular air cavity arranged in the rotating seat, an air inlet pipe arranged on one side of the rotating seat and communicating with the air cavity, a plurality of air inlet holes radially arranged on the turbine shaft and communicating with the gas passage and the air cavity, a heat dissipation cavity arranged in the blade body, an air inlet passage communicating with the heat dissipation cavity and the gas passage, a plurality of through holes arranged on both sides of the blade body and communicating with the heat dissipation cavity, and a cold air blower connected to the air inlet pipe. The blade body comprises a blade root, one end of the blade root being a fixed portion, the other end of the blade root being a blade tail portion, the fixed portion being connected to the turbine shaft, the blade tail portion being helically upward, the blade tail portion being in the rotating direction of the turbine blade, a circular curved surface one being arranged at the front end of the blade tail portion, a circular curved surface two being arranged at the rear end of the blade tail portion, the radius of the circular curved surface one being greater than the radius of the circular curved surface two.
2. A turbine blade according to claim 1, wherein In the rotating direction of the blade body, the thickness of the front end of the blade tail portion is greater than the thickness of the rear end of the blade tail portion.
3. A turbine blade according to claim 1, wherein The helical angle of the blade body ranges from 30 to 40 degrees.
4. A turbine blade according to claim 1, wherein The heat dissipation cavity is helically upward, the cross-sectional area of the heat dissipation cavity at the front end of the blade tail portion is greater than the cross-sectional area of the heat dissipation cavity at the rear end of the blade tail portion.
5. A turbine blade according to claim 1, wherein The blade root and the blade tail portion are both sprayed with a thermal barrier coating.