Turbine working blade and turbofan device with same

Abstract

The invention relates to the technical field of turbine blades, and discloses a turbine working blade and a turbine fan device with the same. The turbine working blade comprises a blade body, a flow inlet channel is formed in the blade body, a first middle cavity and a second middle cavity are formed in the blade body, and the first middle cavity and the second middle cavity are both communicated with the flow inlet channel; the front cavity is formed in the front edge area of the blade body, and the front cavity communicates with the first middle cavity; and the trailing edge cavity is formed in the trailing edge area of the blade body, one end of the trailing edge cavity communicates with the second middle cavity, the other end of the trailing edge cavity communicates with a trailing edge air film hole, the trailing edge air film hole communicates with the external environment, and a honeycomb type airflow channel is formed in the trailing edge cavity. The interior of the trailing edge cavity is arranged to be the honeycomb type airflow channel, and the honeycomb bionic channel is adopted, so that the cooling effect is greatly improved, and the metal temperature of the surface of the blade body can be obviously reduced.

Description

Technical Field

[0001] The present invention relates to the technical field of turbine blades, and in particular to a turbine working blade and a turbofan device having the same. Background Art

[0002] The demand for high power, low fuel consumption and other indicators of aircraft engines is constantly increasing. The temperature before the engine turbine has reached above 2000K, which poses a severe test to the stable operation of turbine blades in high temperature and long-term environment. It is urgent to develop a new type of turbine blade cooling structure to meet the demand of advanced aircraft engines for high cooling effect of turbine blades.

[0003] The cooling of turbine blades mainly includes cooling forms such as impingement, serpentine channel, air film, spoiler column, etc., and a composite cooling design with multiple cooling structures is adopted according to the structure of turbine blades and environmental influences. For high cycle parameter turbine blades, a combined cooling form of "serpentine channel + impingement + air film + rib + slit" is usually adopted. At present, this technology is very mature and has been widely used in large, medium and small aircraft engines.

[0004] However, as the temperature before the turbine of an aircraft engine further increases, this cooling form has limited ability to reduce the metal temperature of the blades by optimizing the inner cavity cooling structure to enhance heat exchange. It can only passively use full air film cooling to reduce the blade temperature, while also increasing the cold air bleed ratio, which will lead to a decrease in engine performance. Specifically, while the cold air takes away the heat from the high-temperature wall of the blades during the flow of the serpentine channel, its own temperature continues to rise, resulting in a continuous weakening of the heat dissipation capacity along the flow direction, and ultimately causing uneven temperature distribution on the blade wall and concentrated thermal stress. Moreover, most of the serpentine channel structures in the prior art are not much different from the prototype after optimization, and the heat exchange capacity is limited. While the cooling effect is improved by adjusting the geometric parameters of the serpentine channel and the rib column structure, the channel flow resistance will also increase, making it difficult to achieve a balance. Summary of the invention

[0005] In view of this, the present invention provides a turbine working blade and a turbofan device having the same, so as to solve the problem that the turbine blades in the prior art cannot meet the cooling demand of the turbine working blades of a high cycle parameter engine.

[0006] In a first aspect, the present invention provides a turbine working blade, comprising:

[0007] A blade body, wherein the blade body is provided with an inlet channel, wherein a first intermediate chamber and a second intermediate chamber are provided in the blade body, and both the first intermediate chamber and the second intermediate chamber are communicated with the inlet channel;

[0008] A front chamber, disposed inside the front edge region of the blade body, the front chamber being communicated with the first intermediate chamber;

[0009] The trailing edge chamber is arranged inside the trailing edge area of ​​the blade body. One end of the trailing edge chamber is connected to the second intermediate chamber, and the other end is connected to a trailing edge air film hole. The trailing edge air film hole is connected to the external environment. A honeycomb airflow channel is arranged in the trailing edge chamber.

[0010] When the turbine blade is working, the cold air enters the inner cavity of the blade body from the inlet channel and is divided into two streams, entering the first intermediate chamber and the second intermediate chamber respectively, wherein the cold flow in the first intermediate chamber flows to the front chamber and then is discharged from the front chamber to cool the leading edge area of ​​the blade body and the middle part near the leading edge; the cold flow in the second intermediate chamber flows toward the trailing edge chamber, flows to the trailing edge air film hole in the trailing edge chamber and then is discharged to cool the trailing edge area of ​​the blade body and the middle part near the trailing edge area. By setting the interior of the trailing edge chamber as a honeycomb airflow channel, a honeycomb bionic channel is adopted, the cooling effect is greatly improved, the metal temperature on the blade surface can be significantly reduced, and the required amount of cold air is small. The honeycomb airflow channel can control the cooling effect in different areas by adjusting its structural parameters, changing the heat exchange area and heat exchange coefficient in different areas, and realizing the refined control of the cooling of the high temperature area and low temperature area of ​​the blade, thereby reducing the temperature gradient of the blade metal, improving the cooling effect of the vortex working blade, and enabling the vortex working blade to better meet the cooling requirements of the turbine working blade of the high cycle parameter engine. This provides a new improvement idea for improving the cooling effect of subsequent vortex working blades.

[0011] In an optional embodiment, a first basin-side chamber and a first back-side chamber are further provided in the blade body, the first basin-side chamber is provided on the side of the first intermediate chamber facing the blade basin of the blade body, the first back-side chamber is provided on the side of the first intermediate chamber facing the back of the blade body, and a honeycomb-type airflow channel is provided in the first basin-side chamber and / or the first back-side chamber. By respectively providing the first basin-side chamber and the first back-side chamber on both sides of the first intermediate chamber and providing the honeycomb-type airflow channel therein, the cooling effect of the cold flow on the blade body in the first intermediate chamber area can be enhanced.

[0012] In an optional embodiment, a second basin side chamber and a second back side chamber are further provided in the blade body, the second basin side chamber is provided on the side of the second intermediate chamber facing the blade basin of the blade body, the second back side chamber is provided on the side of the second intermediate chamber facing the back of the blade body, a honeycomb airflow channel is provided in the second basin side chamber and / or the second back side chamber, and the honeycomb airflow channel provided in the second basin side chamber and / or the second back side chamber is arranged perpendicular to the honeycomb airflow channel provided in the trailing edge chamber. By respectively providing the second basin side chamber and the second back side chamber on both sides of the second intermediate chamber and providing the honeycomb airflow channel therein, the cooling effect of the cold flow on the blade body in the second intermediate chamber area can be enhanced.

[0013] In an optional embodiment, the first intermediate chamber is connected to the front chamber via a front chamber impact hole, and a plurality of the front chamber impact holes are arranged at intervals.

[0014] In an optional embodiment, one end of the front cavity impact hole connected to the front chamber is arranged toward the inner wall of the back side of the blade of the front chamber, thereby constructing a leading edge swirl impact structure, which enables the airflow to form a vortex in the process of radially upward flow when entering the front chamber from the front cavity impact hole, thereby enhancing the heat exchange capacity of the cold air inside the front chamber.

[0015] In an optional embodiment, at least two layers of honeycomb-type airflow channels are provided on the inner wall of the front chamber to enhance the cooling effect of the cold flow on the blade body in the front chamber area.

[0016] In an optional embodiment, a plurality of trailing edge chambers are spaced apart in the radial direction of the blade body, and the plurality of trailing edge chambers are separated from each other to avoid an increase in metal temperature gradient caused by uneven flow distribution in the honeycomb channels in the trailing edge region in the radial direction. The trailing edge chambers in the trailing edge region can be divided into a plurality of sections in radial height, and the honeycomb channels in different trailing edge chambers are separated by partitions. The airflows in different regions in the radial direction will not cross-flow with each other, thereby avoiding the phenomenon of uneven flow distribution caused by uneven static pressure distribution in the radial direction at the trailing edge cold air outlet.

[0017] In an optional embodiment, a front chamber blade tip hole is provided at one end of the front chamber facing the blade tip, and the front chamber blade tip hole is connected to the external environment.

[0018] In an optional embodiment, a tenon is provided on the blade body, and the inlet channel is provided in an inner cavity of the tenon.

[0019] In a second aspect, the present invention further provides a turbofan device having the turbine blades of the present invention, wherein the turbofan device may be an aircraft engine or a gas turbine, because the turbofan device includes the turbine blades, which have the same effect as the turbine blades, and will not be described in detail here. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] In order to more clearly illustrate the specific implementation methods of the present invention or the technical solutions in the prior art, the drawings required for use in the specific implementation methods or the description of the prior art will be briefly introduced below. Obviously, the drawings described below are some implementation methods of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying creative work.

[0021] Figure 1 A cross-sectional view of a turbine blade provided for an embodiment of the present invention.

[0022] Figure 2 A top view of a turbine blade provided in accordance with an embodiment of the present invention.

[0023] Figure 3 A schematic diagram of the flow path of cold air in the inner cavity of a turbine working blade provided in an embodiment of the present invention, wherein the red arrow indicates the flow direction of the cold air.

[0024] Figure 4 A schematic structural diagram of a honeycomb-type airflow channel provided for an embodiment of the present invention, wherein the red arrow indicates the flow direction of cold air.

[0025] Explanation of the accompanying drawings: 1. front chamber; 2. first intermediate chamber; 3. second intermediate chamber; 4. front chamber impact hole; 5. first basin side chamber; 6. first dorsal chamber; 7. second basin side chamber; 8. second dorsal chamber; 9. trailing edge chamber; 10. trailing edge air film hole; 11. front chamber blade tip hole; 12. first intermediate blade basin tip outlet; 13. first intermediate blade back tip outlet; 14. second intermediate blade basin tip outlet; 15. second intermediate blade back tip outlet; 16. tenon inner cavity; 17. extended root inner cavity. DETAILED DESCRIPTION

[0026] In order to make the purpose, technical solution and advantages of the embodiments of the present invention clearer, the technical solution in the embodiments of the present invention will be clearly and completely described below in conjunction with the drawings in the embodiments of the present invention. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of the present invention.

[0027] Combine the following Figures 1 to 4, describing an embodiment of the present invention.

[0028] According to an embodiment of the present invention, on the one hand, a turbine working blade is provided, including a blade body and a first middle chamber 2, a second middle chamber 3, a front chamber 1 and a trailing edge chamber 9 arranged inside the blade body.

[0029] The blade body is provided with an inlet channel. In this embodiment, a tenon is provided on the blade body, and the inlet channel is provided in the inner cavity 16 of the tenon. The middle area of ​​the blade body is provided with a first intermediate chamber 2 and a second intermediate chamber 3, and both the first intermediate chamber 2 and the second intermediate chamber 3 are connected to the inlet channel. The front chamber 1 is provided inside the leading edge area of ​​the blade body, and the front chamber 1 is connected to the first intermediate chamber 2. The trailing edge chamber 9 is provided inside the trailing edge area of ​​the blade body, and one end of the trailing edge chamber 9 is connected to the second intermediate chamber 3, and the other end is connected to the trailing edge air film hole 10, and the trailing edge air film hole 10 is connected to the external environment. A honeycomb airflow channel is provided in the trailing edge chamber 9.

[0030] When the turbine blade is working, the cold air enters the inner cavity of the blade body from the inlet channel and is divided into two streams, entering the first intermediate chamber 2 and the second intermediate chamber 3 respectively. The cold flow in the first intermediate chamber 2 flows to the front chamber 1 and then is discharged from the front chamber 1 to cool the leading edge area of ​​the blade body and the middle part close to the leading edge; the cold flow in the second intermediate chamber 3 flows toward the trailing edge chamber 9, flows to the trailing edge air film hole 10 in the trailing edge chamber 9 and then is discharged to cool the trailing edge area of ​​the blade body and the middle part close to the trailing edge area. By setting the interior of the trailing edge chamber 9 as a honeycomb-type airflow channel and adopting a honeycomb bionic channel, the cooling effect is greatly improved, the metal temperature on the blade surface can be significantly reduced, and the required amount of cold air is small. The honeycomb airflow channel can control the cooling effect in different areas by adjusting its structural parameters and changing the heat exchange area and heat exchange coefficient in different areas, so as to achieve refined control of the cooling of the high-temperature and low-temperature areas of the blade, thereby reducing the temperature gradient of the blade metal and improving the cooling effect of the vortex working blade, so that the vortex working blade can better meet the cooling requirements of the turbine working blade of the high-cycle parameter engine. This provides a new improvement idea for the subsequent improvement of the cooling effect of the vortex working blade.

[0031] In one embodiment, a first basin-side chamber 5 and a first back-side chamber 6 are further provided in the blade body. The first basin-side chamber 5 is provided on the side of the first intermediate chamber 2 facing the blade basin of the blade body, and the first back-side chamber 6 is provided on the side of the first intermediate chamber 2 facing the back of the blade body. Honeycomb-type airflow channels are provided in the first basin-side chamber 5 and the first back-side chamber 6. By respectively providing the first basin-side chamber 5 and the first back-side chamber 6 on both sides of the first intermediate chamber 2 and providing the honeycomb-type airflow channels therein, the cooling effect of the cold flow on the blade body in the area of ​​the first intermediate chamber 2 can be enhanced. In some other embodiments, if the volume of the blade body is small, only one of the first basin-side chamber 5 or the first back-side chamber 6 may be provided.

[0032] Furthermore, a second basin side chamber 7 and a second back side chamber 8 are also provided in the blade body. The second basin side chamber 7 is provided on the side of the blade basin of the second intermediate chamber 3 facing the blade body, and the second back side chamber 8 is provided on the side of the blade back of the second intermediate chamber 3 facing the blade body. A honeycomb type airflow channel is provided in the second basin side chamber 7 and / or the second back side chamber 8. The honeycomb type airflow channel provided in the second basin side chamber 7 and / or the second back side chamber 8 is arranged vertically with the honeycomb type airflow channel provided in the trailing edge chamber 9. By respectively providing the second basin side chamber 7 and the second back side chamber 8 on both sides of the second intermediate chamber 3 and providing the honeycomb type airflow channel therein, the cooling effect of the cold flow on the blade body in the second intermediate chamber 3 region can be enhanced. In some other embodiments, if the volume of the blade body is small, only one of the second basin side chamber 7 or the second back side chamber 8 may be provided.

[0033] In one embodiment, in order to enhance the cooling effect of the cold flow on the front chamber 1 , the first intermediate chamber 2 is connected to the front chamber 1 via a front chamber impact hole 4 , and a plurality of front chamber impact holes 4 are arranged at intervals.

[0034] Specifically, the end of the front cavity impact hole 4 connected to the front chamber 1 is arranged toward the inner wall of the back side of the front chamber 1, so as to construct a leading edge swirl impact structure, which can enable the airflow to form a vortex in the process of radially upward flow when entering the front chamber 1 from the front cavity impact hole 4, thereby enhancing the heat exchange capacity of the cold air inside the front chamber 1. At least two layers of honeycomb airflow channels are arranged on the inner wall of the front chamber 1 to enhance the cooling effect of the cold flow on the blade body in the front chamber 1 area. A front cavity tip hole 11 is arranged at one end of the front chamber 1 facing the blade tip, and the front cavity tip hole 11 is connected to the external environment to discharge the cold flow in the front chamber 1 from the front cavity tip hole 11 to the outside of the blade body.

[0035] In one embodiment, a plurality of trailing edge chambers 9 are arranged at intervals in the radial direction of the blade body, and the plurality of trailing edge chambers 9 are separated from each other to avoid an increase in the metal temperature gradient caused by uneven flow distribution in the honeycomb channels in the trailing edge area in the radial direction. The trailing edge chambers 9 in the trailing edge area can be divided into a plurality of sections in radial height, and the honeycomb channels in different trailing edge chambers 9 are separated by partitions. The airflows in different areas in the radial direction will not cross-flow with each other, thereby avoiding the uneven flow distribution caused by uneven static pressure distribution in the radial direction at the trailing edge cold air outlet.

[0036] This embodiment proposes a novel design for cooling flow path in the inner cavity of a turbine working blade, which takes the honeycomb bionic channel in a confined space as the cooling structure unit and constructs a combined cooling form for the inner cavity of the turbine working blade in combination with the blade profile characteristics. Compared with the traditional cooling design of "winding channel + impact + air film + rib + slit", it can greatly improve the heat exchange capacity of the inner cavity of the blade under the condition of a small amount of cold air, improve the cooling effect of the blade, and meet the cooling needs of the turbine working blades of the new generation of high cycle parameter engines.

[0037] The turbine working blade provided in this embodiment is a combined cooling structure of leading edge swirl impact + middle honeycomb channel + trailing edge honeycomb channel + trailing edge air film hole 10. The front cavity impact hole 4 of the present invention is a swirl impact between the front chamber 1 of the blade and the first intermediate chamber 2, the first basin side chamber 5 is between the first intermediate chamber 2 and the blade basin wall of the blade body, the first back side chamber 6 is between the first intermediate chamber 2 and the blade back wall, the second basin side chamber 7 is between the second intermediate chamber 3 and the blade basin wall, the second back side chamber 8 is between the second intermediate chamber 3 and the blade back wall, the trailing edge chamber 9 is connected to the second intermediate chamber 3, and the trailing edge air film hole 10 is connected to the trailing edge chamber 9.

[0038] The secondary flow cold air is extracted from the two airflows in the combustion chamber. The cold air enters the tenon inner cavity 16 of the tenon through the cold air inlet on the blade tenon, and continues to flow upward through the root inner cavity 17. At this time, part of the cold air enters the first intermediate cavity 2, and the other part of the cold air enters the second intermediate cavity 3. The cold air entering the first intermediate cavity 2 is divided into three airflows, one of which directly enters the first intermediate cavity 2 radially, and then impacts the front cavity 1 through the front cavity impact hole 4. At the same time, the angle of the front cavity impact hole 4 is biased toward the back of the leading edge blade, constructing a leading edge swirl impact structure, which can form a vortex in the process of the airflow flowing radially upward in the front cavity, enhancing the heat exchange capacity of the cold air inside the front cavity, and the cold air entering the front cavity 1 will eventually be discharged from the front cavity blade tip hole 11. The other two streams of cold air entering the first intermediate chamber 2 directly enter the first basin side chamber 5 and the first back side chamber 6 respectively. The honeycomb channel network is arranged on the blade basin side and the back side of the blade close to the wall. The wall cooling plus honeycomb channel method can effectively reduce the metal wall temperature of the blade under the condition of small cold air volume. The cold air entering the honeycomb channel is discharged from the first intermediate blade basin tip outlet 12 and the first intermediate blade back tip outlet 13 respectively after flowing through multiple honeycomb channel units. Similarly, the cold air entering the second intermediate chamber 3 is divided into three airflows, one of which directly enters the second intermediate chamber 3 radially, then flows toward the trailing edge through the transverse honeycomb channel, and finally is discharged from the trailing edge air film hole 10. The other two airflows are distributed and directly enter the second basin side chamber 7 and the second back side chamber 8, and are discharged from the second intermediate blade basin tip outlet 14 and the second intermediate blade back tip outlet 15 respectively after passing through the multi-layer honeycomb channel units.

[0039] The shape of the honeycomb channel is adaptively improved based on the honeycomb structure in nature. The gaps between multiple hexagons form a stable honeycomb network channel to construct a honeycomb channel bionic cooling structure for turbine blade cooling design. On the turbine working blades, referring to the design concept of double-wall cooling of the blades, honeycomb channel units are arranged on the basin side and the back side of the turbine working blades by wall cooling, so as to efficiently cool the turbine working blade body with a small amount of cold air.

[0040] In the first middle chamber 2 of the blade body, a part of the cold air directly enters the first middle chamber 2, the first middle chamber 2 of the blade is hollowed out, and the first middle chamber 2 has no cold air outlet at the blade tip. The cold air in the first middle chamber 2 enters the front chamber 1 through the front chamber impact hole 4, generating a swirl impact effect, and the cold air spirals upward in the front chamber and is discharged from the front chamber tip hole 11. The outlet of the front chamber impact hole 4 provided in this embodiment is close to the back side of the leading edge blade, the aperture of the front chamber impact hole 4 is 0.3mm-1.0mm, and the number of the front chamber impact holes 4 is 8-20. The specific values ​​of the above parameters are determined according to one-dimensional flow calculation and three-dimensional temperature field calculation.

[0041] In the first middle chamber 2 of the blade, a part of the cold air directly enters the first basin side chamber 5 and the first back side chamber 6, and the cold air flows upward along the radial wall surface. At the corner of the honeycomb channel, the airflow will mix with each other, and the heat transfer coefficient is high. Multiple honeycomb channel units are interconnected, which increases the heat transfer area, thereby greatly improving its cooling effect, and finally the cold air will flow out from the tip outlet. The advantages of setting honeycomb channels on the blade basin side and the back side of the blade are: the wall cooling method is used to fit the blade shape, and honeycomb channels are arranged on the blade basin side and the back side of the blade, with a large heat transfer area and good cooling effect; by changing the structural parameters of the honeycomb channel units at different positions on the blade, the cooling effect at different positions can be adjusted, so that the temperature in the high temperature area is reduced, the temperature in the low temperature area is increased, and the temperature gradient of the blade is reduced; compared with the traditional winding channel, the structural parameters and channel seam width of the honeycomb channel unit are adjusted, and the heat transfer coefficient of the inner cavity is greatly improved, which can meet the cooling requirements of the new generation of wide-operating aircraft engines for blades; the structural parameters of the honeycomb channel unit can be adaptively changed in combination with the characteristics of the blade shape, and the structure is simple and changeable. Compared with the complex bionic cooling unit, it can be better applied to turbine cooling blades. In this embodiment, the width F of all honeycomb channel seams is 0.4mm-0.7mm, the angle γ is 60°-120°, D1 is 0.8mm-1.4mm, D2 is 0.4-0.7mm, and D3 is 0.8mm-2.5mm, wherein γ, D1, D2, and D3 are all structural parameters of the honeycomb channel unit. Figure 4 The honeycomb channel unit framed by the dotted line is a single layer of honeycomb channel unit. On the blade basin side of the front chamber 1, 5 to 12 layers of honeycomb channel units are arranged radially, and 2 to 4 layers of honeycomb channel units are arranged laterally. On the blade back side of the front chamber 1, 5 to 12 layers of honeycomb channel units are arranged radially, and 2 to 4 layers of honeycomb channel units are arranged laterally. The above parameters need to be determined by one-dimensional flow iterative calculation and three-dimensional temperature field iterative calculation under the condition of considering the rotation effect.

[0042] In the second intermediate chamber 3 of the blade, a part of the cold air directly enters the second intermediate chamber 3. The middle of the second intermediate chamber 3 of the blade is hollowed out. The second intermediate chamber 3 has no cold air outlet at the tip of the blade, but is directly connected to the trailing edge chamber 9. The layout of the trailing edge chamber 9 is rotated 90 degrees compared with the honeycomb channel of the first intermediate chamber 2. The cold air entering the honeycomb channel in the trailing edge area continues to flow along the trailing edge direction of the blade profile and flows out from the trailing edge air film hole 10. In the present invention, the parameter ranges of the honeycomb channel unit structure parameters γ, D1, D2, D3 and the channel seam width F of the second intermediate chamber 3 are consistent with those of the leading edge area. In the trailing edge area, the honeycomb channel is arranged with 6 to 18 layers of honeycomb channel units in the radial direction and 2 to 10 layers of honeycomb channel units in the transverse direction. The above parameters need to be determined by one-dimensional flow iterative calculation and three-dimensional temperature field iterative calculation under the condition of considering the rotation effect. Since the static pressure distribution along the radial height of the outlet area of ​​the trailing edge air film hole 10 is uneven, in order to avoid the increase of the metal temperature gradient caused by the uneven flow distribution along the radial direction of the honeycomb channel in the trailing edge area, the trailing edge area can be divided into 3 to 5 areas along the radial height, and the honeycomb channels in the area are separated by partitions, and the airflows in different areas along the radial direction will not cross each other, thereby avoiding the uneven flow distribution caused by the uneven static pressure distribution along the radial direction of the trailing edge cold air outlet. In the present invention, the diameter of the trailing edge air film hole 10 is 0.3mm to 0.5mm, and the number of air film holes is 6 to 20.

[0043] The high cycle parameter engine puts the turbine blades in a high temperature and harsh environment. The demand for high performance requires the blades to minimize the amount of air bleed. The turbine blades provided in this embodiment can take into account both small air bleed and high-efficiency cooling. The metal temperature in different areas of the blades can also be adjusted by adjusting the local cooling structure, which has a very high engineering application value. By adopting the honeycomb bionic channel, the cooling effect is greatly improved, the metal temperature on the blade surface can be significantly reduced, and the amount of cold air required is small. The honeycomb channel can control the cooling effect in different areas by adjusting its structural parameters and changing the heat exchange area and heat exchange coefficient in different areas, thereby realizing refined control of the cooling of the high-temperature and low-temperature areas of the blades, thereby reducing the temperature gradient of the blade metal.

[0044] According to an embodiment of the present invention, on the other hand, a turbofan device is also provided, which has the turbine working blades described in the present invention. The turbofan device can be an aircraft engine or a gas turbine, because the turbofan device includes turbine working blades, which have the same effect as the turbine working blades. The cooling effect of the turbine working blades is greatly improved by adopting honeycomb bionic channels, which can significantly reduce the metal temperature on the blade surface, and the required amount of cold air is small. The honeycomb channel can control the cooling effect in different areas by adjusting its structural parameters and changing the heat exchange area and heat exchange coefficient in different areas, so as to achieve refined control of the cooling of the high temperature zone and low temperature zone of the blade, thereby reducing the temperature gradient of the blade metal.

[0045] Although the embodiments of the present invention have been described in conjunction with the accompanying drawings, those skilled in the art may make various modifications and variations without departing from the spirit and scope of the present invention, and such modifications and variations are all within the scope defined by the appended claims.

Claims

1. A turbine blade, characterized in that: include: A blade body, wherein the blade body is provided with an inlet channel, wherein a first intermediate chamber (2) and a second intermediate chamber (3) are provided in the blade body, and both the first intermediate chamber (2) and the second intermediate chamber (3) are communicated with the inlet channel; A front chamber (1) is arranged inside the front edge region of the blade body, and the front chamber (1) is communicated with the first middle chamber (2); The trailing edge chamber (9) is arranged inside the trailing edge area of ​​the blade body, one end of the trailing edge chamber (9) is connected to the second intermediate chamber (3), and the other end is connected to the trailing edge air film hole (10), and the trailing edge air film hole (10) is connected to the external environment. A honeycomb airflow channel is arranged in the trailing edge chamber (9).

2. The turbine blade according to claim 1, characterized in that: A first basin-side chamber (5) and a first back-side chamber (6) are also provided in the blade body. The first basin-side chamber (5) is provided on the side of the first intermediate chamber (2) facing the basin of the blade body, and the first back-side chamber (6) is provided on the side of the first intermediate chamber (2) facing the back of the blade body. A honeycomb-type airflow channel is provided in the first basin-side chamber (5) and / or the first back-side chamber (6).

3. The turbine blade according to claim 1 or 2, characterized in that: A second basin-side chamber (7) and a second back-side chamber (8) are also provided in the blade body. The second basin-side chamber (7) is provided on the side of the second intermediate chamber (3) facing the basin of the blade body, and the second back-side chamber (8) is provided on the side of the second intermediate chamber (3) facing the back of the blade body. A honeycomb-type airflow channel is provided in the second basin-side chamber (7) and / or the second back-side chamber (8). The honeycomb-type airflow channel provided in the second basin-side chamber (7) and / or the second back-side chamber (8) is arranged perpendicular to the honeycomb-type airflow channel provided in the trailing edge chamber (9).

4. The turbine blade according to claim 1 or 2, characterized in that: The first intermediate chamber (2) is connected to the front chamber (1) via a front chamber impact hole (4), and a plurality of the front chamber impact holes (4) are arranged at intervals.

5. The turbine blade according to claim 4, characterized in that: One end of the front cavity impact hole (4) communicating with the front cavity (1) is arranged toward the inner wall of the front cavity (1) on the blade back side.

6. The turbine rotor blade according to claim 1 or 2, characterized in that: At least two layers of honeycomb-shaped airflow channels are arranged on the inner wall of the front chamber (1).

7. The turbine rotor blade according to claim 1 or 2, characterized in that: A plurality of the trailing edge chambers (9) are arranged at intervals in the radial direction of the blade body, and the plurality of trailing edge chambers (9) are separated from each other.

8. The turbine rotor blade according to claim 1 or 2, characterized in that: A front chamber blade tip hole (11) is provided at one end of the front chamber (1) facing the blade tip, and the front chamber blade tip hole (11) is in communication with the external environment.

9. The turbine rotor blade according to claim 1 or 2, characterized in that: The blade body is provided with a tenon, and the inlet channel is arranged in the inner cavity of the tenon.

10. A turbofan device, characterized in that: A turbine rotor blade according to any one of claims 1 to 9.

Images