Gas-film-free gas turbine blade adopting near-wall vertical channel for cooling

Abstract

The invention relates to the technical field of gas turbine blades, and discloses a gas-film-free gas turbine blade adopting a near-wall surface vertical channel for cooling, the gas turbine blade comprises a blade body, the blade body comprises a pressure side and a suction side, and a first cooling cavity and a second cooling cavity which vertically penetrate through the blade body are respectively formed in the blade height direction; the blade body is further provided with a plurality of first cooling passages formed in the pressure side, a plurality of second cooling passages formed in the suction side and a third cooling passage formed in the front edge, and the first cooling passages, the second cooling passages and the third cooling passage are all arranged in the blade height direction and vertically penetrate through the blade body. And the third cooling passage is spiral as a whole. The uniformity of surface temperature distribution of the gas turbine blade is improved, and the cooling effect of the front edge area of the blade is enhanced; and the endurance capacity of the gas turbine blade in a high-temperature environment is enhanced, so that the overall performance of a gas turbine is improved, and safe and reliable operation of the gas turbine is guaranteed.

Description

Technical Field

[0001] The present invention relates to the technical field of gas turbine blades, and particularly to a gas turbine blade without a gas film and cooled by a near-wall vertical channel. Background Art

[0002] During the operation of a gas turbine, it is affected by harsh conditions such as high temperature, high pressure, and corrosion. Therefore, effective cooling technologies need to be adopted to ensure its normal operation. Currently, the turbine inlet temperature is relatively high, which has exceeded the heat resistance limit of the blade material. High temperature will seriously affect the service life and safety performance of gas turbines and aero engines. Therefore, the development of efficient gas turbine blade cooling technologies is of great significance for improving the performance of gas turbines.

[0003] Common gas turbine blade cooling methods mainly include film cooling, impingement cooling, convective cooling, etc. The gas turbine blade structure can be divided into three parts according to position: the leading edge, the mid-chord, and the trailing edge. The leading edge region usually adopts a combination of impingement cooling and film cooling to cope with the direct impact of high-temperature gas; the mid-chord region strengthens heat transfer by setting structures such as turbulator ribs and depressions in the internal cooling channels; the trailing edge region mostly adopts the method of column rib turbulator cooling to enhance the heat transfer effect in the channels.

[0004] In the design of traditional gas turbine blades, film holes are usually arranged in a dense pattern on the pressure side and the suction side. The cold air from the internal cooling channels of the blade is discharged through these discretely distributed holes, and under the action of the high-temperature mainstream gas, a cooling gas film close to the wall surface is formed, thereby achieving thermal protection of the blade surface. However, there are still several technical limitations in this traditional film cooling method in practical applications: 1. There are significant problems in terms of cooling uniformity. Due to the discrete distribution characteristics of the film holes, the coverage of the cooling gas on the blade surface is uneven, and it is easy to generate local under-cooled areas, thereby forming hot spots with relatively large temperature gradients, significantly increasing the risk of thermal stress concentration on the blade. 2. The cooling efficiency is severely restricted by environmental factors. Although film cooling can theoretically form a protective layer, under actual high-temperature and high-pressure working conditions, the high-speed mainstream gas is extremely likely to cause a scouring effect on the cooling gas film, resulting in a decrease in the stability of the gas film and a significant reduction in the cooling effect. 3. From the perspective of system performance, this cooling method will have a negative impact. The continuously discharged cooling gas will inevitably reduce the temperature of the mainstream gas, resulting in insufficient gas temperature for the subsequent stages of blades, thus affecting the thermal cycle efficiency of the entire gas turbine system. 4. There are also challenges in manufacturing and maintenance. The processing of film holes requires extremely high precision requirements, significantly increasing the manufacturing cost; at the same time, during long-term operation, the tiny film holes are prone to blockage due to the accumulation of pollutants, not only reducing the cooling reliability but also increasing the maintenance difficulty. To improve the cooling efficiency of gas turbines, common cooling methods include film cooling, internal impingement cooling, ribbed channel cooling, and pin-fin cooling, etc. Multiple rows of film holes are arranged on the suction surface and pressure surface of the gas turbine blade, and the internal cooling air flows out through the film holes. Under the action of the mainstream gas, a layer of cold gas film is formed on the blade surface, which can effectively isolate the direct contact between the high-temperature gas and the wall surface. However, the mixing of the cooling air flow and the mainstream will cause a decrease in the temperature of the mainstream gas, resulting in insufficient mainstream temperature for the subsequent stages of blades and affecting the overall efficiency of the gas turbine. Therefore, the present invention pays more attention to the internal cooling of the channel and proposes a new blade structure. Compared with film cooling, this internal cooling channel arrangement will obtain a more uniform cooling effect. Summary of the Invention

[0005] In order to solve the problems in the above-mentioned prior art that the traditional film cooling method has uneven cooling, low cooling efficiency, high manufacturing cost, and will cause insufficient gas temperature, thus affecting the thermal cycle efficiency of the entire gas turbine system, the present invention proposes a gas turbine blade that uses near-wall vertical channels for cooling without film holes. This blade deletes the film holes on the pressure surface and suction surface of the blade body on the basis of the traditional blade mainly using film cooling, and arranges vertical pipes along the chord length inside the blade body wall surface, increasing the heat transfer area and improving the heat transfer efficiency, reducing the blade weight and also reducing the manufacturing cost.

[0006] The present invention is achieved through the following technical solutions: It includes a blade body, the blade body includes a pressure side and a suction side, the blade body is respectively provided with a first cooling chamber and a second cooling chamber that vertically penetrate the blade body along the blade height direction, the first cooling chamber is arranged near the leading edge of the blade body, and the second cooling chamber is arranged near the trailing edge of the blade body; The blade body is also provided with a plurality of first cooling passages opened on the pressure side, a plurality of second cooling passages opened on the suction side, and a third cooling passage opened on the leading edge. The first cooling passage, the second cooling passage, and the third cooling passage are all arranged along the blade height direction and vertically penetrate the blade body; The overall shape of the third cooling passage is spiral.

[0007] As a further preference, the leading edge portion of the blade body is also provided with a plurality of fourth cooling passages. The fourth cooling passages are equidistantly distributed circumferentially along the leading edge portion of the blade body, and a plurality of the fourth cooling passages are all arranged along the blade height direction and penetrate the blade body.

[0008] As a further preference, the trailing edge portion of the blade body is provided with a plurality of fifth cooling passages. The cross-sectional areas of the plurality of fifth cooling passages gradually decrease from the portion close to the leading edge of the blade body to the portion far from the leading edge of the blade body; A plurality of the fifth cooling passages are all arranged along the blade height direction of the blade body and penetrate the blade body.

[0009] As a further preference, the number of the fourth cooling passages is three, and the diameter is 1.8 mm.

[0010] As a further preference, the number of the fifth cooling passages is six, and the diameters of the six fifth cooling passages are 3.5 mm, 2.4 mm, 2 mm, 1.6 mm, 1.4 mm, and 1 mm respectively.

[0011] As a further preference, the first cooling passage and the second cooling passage are S-shaped.

[0012] As a further preference, a transverse channel transition is arranged between the arc segments of the S-shaped of the first cooling passage and the second cooling passage. The length of the arc segment of the S-shaped of the first cooling passage and the second cooling passage is 15 mm, and the length of the transverse channel is 4 mm.

[0013] As a further preference, an air inlet chamber is arranged on the same side of the first cooling passage and the second cooling passage.

[0014] As a further preference, air inlet chambers are arranged alternately on both sides of the first cooling passage and the second cooling passage.

[0015] As a further preference, there are twelve first cooling passages equally spaced along the profile direction of the pressure side of the blade body, and fourteen second cooling passages equally spaced along the profile direction of the suction side of the blade body. The diameters of both the first cooling passages and the second cooling passages are 2 mm.

[0016] Compared with the prior art, the beneficial effects of the present invention are as follows:

[0017] 1. The present invention uses cooling gas to cool the surface of the gas turbine blade through vertical cooling pipelines. Compared with traditional gas turbine blades, when deleting the surface film holes, vertical cooling pipelines close to the wall are added, which increases the heat transfer area inside the blade. Moreover, a spiral cooling channel is designed at the leading edge of the blade to enhance the cooling characteristics of the leading edge of the blade. At the trailing edge of the blade, a gradually shrinking vertical pipeline is designed. Such an arrangement will strengthen the cooling effect at the leading edge of the blade, and a relatively uniform temperature distribution can be observed on the pressure surface and the suction surface. Through the vertical simulation method, the parts with weak heat transfer can be observed in advance, and the heat transfer in this area can be enhanced, which can but is not limited to increasing the number of pipelines or the diameter of pipelines in this area to enhance the heat transfer capacity.

[0018] 2. The present invention deletes the film holes on the suction surface and the pressure surface of the gas turbine blade. Because the air film holes in the middle chord are replaced by longitudinal pipelines, and the air outlet is located on the upper wall of the blade without mixing with the mainstream, the amount of cold gas mixed with the hot gas is reduced. As Figure 4 , we also monitored the outlet temperature of the hot gas through numerical simulation and found that the outlet temperature of the new structure is higher than that of the other two structures, and the heat loss is smaller at different Reynolds numbers, which improves the overall efficiency of the gas turbine to a certain extent.

[0019] 3. Aiming at the problems of the traditional gas turbine blade with a relatively high surface temperature and uneven temperature distribution concentrated in the downstream area of the film holes, the present invention proposes a gas turbine blade with no film and cooled by near-wall vertical channels. This innovative design improves the uniformity of the surface temperature distribution of the gas turbine blade, and focuses on enhancing the cooling of the leading edge area of the blade. It enhances the tolerance of the gas turbine blade in a high-temperature environment, thereby improving the overall performance of the gas turbine and ensuring the safe and reliable operation of the gas turbine. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 It is a top view of the gas turbine blade with no film and cooled by near-wall vertical channels according to the present invention.

[0021] Figure 2 It is an overall structural schematic diagram of the gas turbine blade with no film and cooled by near-wall vertical channels according to the present invention.

[0022] Figure 3Schematic diagram of the flow directions of the cooling gas in the first cooling passage and the second cooling passage in the embodiments of the present invention.

[0023] Figure 4 Schematic diagram of the flow directions of the cooling gas and the mainstream gas during the operation of a gas turbine blade with near-wall vertical channels for cooling without an air film according to the present invention.

[0024] Figure 5 Schematic diagram of a preferred shape of the first cooling passage and the second cooling passage in the embodiments of the present invention.

[0025] Figure 6 Schematic diagram of a double-wall cooling blade with air film holes in a prior art in the comparative experiment of the present invention.

[0026] Figure 7 Schematic diagram of a single-wall cooling blade with air film holes in a prior art in the comparative experiment of the present invention.

[0027] Figure 8 Outlet temperatures of the mainstream gas of three types of structural blades at variable Reynolds numbers in the comparative experiment.

[0028] Figure 9 Internal streamline diagram of numerical simulation of a gas turbine blade with near-wall vertical channels for cooling without an air film according to the present invention.

[0029] Figure 10 Temperature contour maps of the pressure surface and the suction surface of numerical simulation of a gas turbine blade with near-wall vertical channels for cooling without an air film according to the present invention.

[0030] Labels in the figure:

[0031] 1. First cooling chamber; 2. Second cooling chamber; 3a. First cooling passage; 3b. Second cooling passage; 4. Third cooling passage; 5. Fourth cooling passage; 6. Fifth cooling passage; 100. Blade body. Detailed implementation manners

[0032] The advantages and features of the present invention will be illustrated and explained through non-limiting descriptions of the following preferred embodiments, which are given only as examples with reference to the accompanying drawings.

[0033] Such as Figure 1 And Figure 2As shown, the present invention provides a gas turbine blade with a gasless film and cooled by a near-wall vertical channel, including a blade body 100. The blade body 100 includes a pressure side and a suction side. The pressure side of the blade body 100 is the side with a higher static pressure on the blade, usually located on the curved surface of the blade facing the fluid flow direction. Its shape is generally relatively flat or convex because, according to Bernoulli's principle, a lower flow velocity results in a higher static pressure, so the fluid flow velocity is slower on the pressure side of the gas turbine blade. Opposite to the pressure side of the blade body 100, the suction side of the blade body 100 is the side with a lower static pressure on the blade, usually located on the curved surface of the blade facing away from the flow direction. Its shape is generally more curved or concave, and the fluid flow velocity is faster on this side. One end of the blade body 100 where the pressure side and the suction side transition smoothly is the leading edge, and the end far from the leading edge is the trailing edge. The included angle between the pressure side and the suction side of the blade body 100 at the trailing edge is smaller. Along the height direction of the blade body 100, a first cooling chamber 1 and a second cooling chamber 2 are respectively provided. The first cooling chamber 1 is arranged near the leading edge of the blade body 100, and the second cooling chamber 2 is arranged near the trailing edge of the blade body 100. Both the first cooling chamber 1 and the second cooling chamber 2 vertically penetrate the blade body 100.

[0034] The blade body 100 is also provided with a first cooling passage 3a, a second cooling passage 3b, and a third cooling passage 4. The first cooling passage 3a is opened on the pressure side of the blade body 100, the second cooling passage 3b is opened on the suction side of the blade body 100, and the third cooling passage 4 is opened on the leading edge of the blade body 100. The first cooling passage 3a, the second cooling passage 3b, and the third cooling passage 4 are all arranged along the height direction and penetrate the blade body 100. A plurality of the first cooling passages 3a are arranged along the profile line direction of the pressure side of the blade body 100, and a plurality of the second cooling passages 3b are arranged along the profile line direction of the suction side of the blade body 100. The third cooling passage 4 is integrally spiral, and such a setting can enhance the cooling characteristics of the leading edge of the blade body 100. The first cooling passage 3a is preferably equidistantly distributed along the profile line direction of the pressure side of the blade body 100, and the second cooling passage 3b is preferably equidistantly distributed along the profile line direction of the suction side of the blade body 100. Such an arrangement will strengthen the cooling effect at the leading edge of the blade, and a relatively uniform temperature distribution can be observed on the pressure surface and the suction surface. By means of vertical simulation, the part with weaker heat transfer can be observed in advance, and the heat transfer in this area can be enhanced. This can be achieved but is not limited to by increasing the number of pipes or the pipe diameter in this area to enhance the heat transfer capacity.

[0035] A plurality of fourth cooling passages 5 are also provided at the leading edge portion of the blade body 100. The fourth cooling passages 5 are equidistantly distributed circumferentially along the leading edge portion of the blade body 100, and the plurality of fourth cooling passages 5 are all arranged in the blade height direction and penetrate through the blade body 100. The number of the fourth cooling passages 5 is preferably three. A plurality of fifth cooling passages 6 are provided at the trailing edge portion of the blade body 100. The cross-sectional areas of the fifth cooling passages 6 gradually decrease from the portion close to the leading edge portion of the blade body 100 to the portion far from the leading edge portion of the blade body 100. The plurality of fifth cooling passages 6 are all arranged in the blade height direction of the blade body 100 and penetrate through the blade body 100.

[0036] As Figure 2 shown, it is a flow mode of the cooling gas of the gas turbine blade with near-wall vertical channels cooling without air film of the present invention. Among them, the cooling gas enters the blade body 100 from the same side of the first cooling passage 3a and the second cooling passage 3b, that is, Figure 2 the upper side in. After sufficient heat exchange in the first cooling passage 3a and the second cooling passage 3b inside the blade body 100, the cooling gas flows out of the blade body 100 from the other side of the first cooling passage 3a and the second cooling passage 3b, that is, Figure 2 the lower side in flows out of the blade body 100. In this structure, the intake chambers of the cooling gas are all on the same side in the blade height direction of the blade body 100. The connection relationship between the intake chamber of the cooling gas and the blade body 100 is the prior art and will not be elaborated in the present invention. Only the setting position of the intake chamber is changed.

[0037] As Figure 3 shown, it is another flow mode of the cooling gas of the gas turbine blade with near-wall vertical channels cooling without air film of the present invention. In some embodiments, the intake chambers of the cooling gas are staggered on both sides in the blade height direction of the blade body 100, that is, the cooling gas flows into the blade body 100 in a staggered manner from both ends of the first cooling passage 3a and the second cooling passage 3b for heat exchange, and then flows out of the blade body 100 from the other ends of the first cooling passage 3a and the second cooling passage 3b respectively. This staggered intake method can make the temperature distribution on the surface more uniform.

[0038] As Figure 4As shown, it is a schematic diagram of the flow directions of the cooling gas and the mainstream gas of a gas turbine blade with a near-wall vertical channel cooling for the airless film of the present invention in the working state. The blade body 100 of the present invention has no surface air film holes, and the impact chamber at the leading edge of the blade body 100 is replaced by the third cooling passage 4, and the impact air film holes at the leading edge of the blade body 100 are replaced by three fourth cooling passages 5. Such a setting enables the cooling gas to enter the blade body 100 from one side in the blade height direction of the blade body 100, and after sufficient heat exchange inside the blade body 100, then flow out of the blade body 100 from the other side in the blade height direction of the blade body 100, that is, enter the blade body 100 from the upper end face of the blade body 100 in Figure 4 and flow out of the blade body 100 from the lower end face of the blade body 100. The flow direction of the mainstream gas of the gas turbine is Figure 4 the horizontal direction in. Therefore, several cooling passages designed in the present invention will not mix with the mainstream gas to reduce the temperature at the mainstream outlet, and can significantly reduce the influence on the subsequent-stage blades caused by the reduction of the mainstream temperature due to the mixing of the cooling gas and the mainstream.

[0039] As Figure 5 shown, in some embodiments of the present invention, the first cooling passage 3a and the second cooling passage 3b are arranged in an S shape, that is, a serpentine shape. Such a setting can increase the effective length of the first cooling passage 3a and the second cooling passage 3b, thereby improving the heat exchange efficiency of the cooling gas. Preferably, a transverse channel transition is provided between the arc segments of the S shape of the first cooling passage 3a and the second cooling passage 3b. The arc segment length of the S shape of the first cooling passage 3a and the second cooling passage 3b is 15 mm, and the length of the transverse channel is 4 mm.

[0040] The diameter of the fourth cooling passage 5 is 1.8 mm. The number of the fifth cooling passages 6 is preferably six, and the diameters of the six fifth cooling passages 6 are 3.5 mm, 2.4 mm, 2 mm, 1.6 mm, 1.4 mm, and 1 mm respectively. The number of the first cooling passages 3a arranged at equal intervals on the pressure side of the blade body 100 is 12, and the number of the second cooling passages 3b arranged at equal intervals on the suction side of the blade body 100 is 14. The diameters of the first cooling passage 3a and the second cooling passage 3b are both 2 mm.

[0041] The gas turbine blades of the present invention that adopt airless film cooling by vertical channels near the wall have deleted the air film holes on the pressure side and suction side of the blade body on the basis of the traditional blades that mainly adopt air film cooling, and a number of vertical cooling channels are arranged inside the wall of the blade body along the blade height direction. Equidistant vertical cooling channels are arranged inside the blade body along the direction of the profile line to increase the heat exchange area between the cooling gas and the blade surface. The cooling gas takes away the wall heat through these pipes. The leading edge has deleted the traditional impingement cooling method and adopts internal spiral channel cooling, which can further increase the speed of the cooling fluid and increase the contact area with the blade, thereby improving the heat exchange efficiency. Three vertical pipes are arranged at the leading edge to enhance the heat exchange of the leading edge, the split seam is deleted at the trailing edge, and is replaced by a vertical channel. The middle of the blade is hollowed out to form two large cavities, which reduces the weight of the blade while also reducing the manufacturing cost.

[0042] like Figure 6 and Figure 7 The figures are schematic diagrams of double-wall cooling blades with air film holes and single-wall cooling blades with air film holes in the prior art. These two traditional air film cooling methods have problems such as uneven cooling, low cooling efficiency, high manufacturing cost, and insufficient gas temperature, thus affecting the thermal cycle efficiency of the entire gas turbine system. This is mainly reflected in the mixing of the cooling gas with the mainstream gas, which causes the temperature of the mainstream gas outlet to be too low, affecting the low thermal cycle efficiency of the entire gas turbine system.

[0043] In order to further illustrate the superiority of the structure of the gas turbine blades cooled by the airless membrane using the vertical channels near the wall of the present invention, the gas turbine blades cooled by the airless membrane using the vertical channels near the wall of the present invention and Figure 6 , Figure 7 The cooling blades in the two prior arts shown are modeled and numerically simulated respectively, and the boundary conditions of the numerical simulation are shown in Table 1. g is the inlet temperature of the mainstream gas, which is 423K; Q g is the flow rate of the mainstream gas, which is 0.231 kg·s -1 ;Re in is the Reynolds number. The present invention compares three types of Reynolds numbers, which are 5×10 4 ,7×10 4 and 1×10 5 ; T c is the temperature of the cooling gas, which is 298.5K; TR is the temperature ratio, which is 1.42; BR is the flow ratio, which is 4.8%.

[0044] Table 1

[0045]

[0046] Figure 8In the comparative experiment, for the mainstream gas outlet temperature of the three types of vane structures at variable Reynolds numbers, it can be seen from the figure that the mainstream gas outlet temperature of the gas-turbine vane structure (pipe network cooling) designed in the present invention without a gas film and using a near-wall vertical channel cooling is higher than that of the other two structures (single-wall cooling with gas film holes and double-wall cooling with gas film holes) under the three Reynolds number conditions. At the same time, the mainstream gas outlet temperature of the gas-turbine vane structure (pipe network cooling) designed in the present invention without a gas film and using a near-wall vertical channel cooling has less heat loss at different Reynolds numbers, improving the overall efficiency of the gas turbine.

[0047] Figure 9 And Figure 10 are respectively the numerical simulation internal streamline diagram of the gas-turbine vane without a gas film and using a near-wall vertical channel cooling and the numerical simulation temperature cloud diagram of the pressure side and the suction side of the present invention. Figure 10 In [figure], the temperature cloud diagram on the left side is the suction side, and the temperature cloud diagram on the right side is the pressure side. It can be found from the figure that the cooling effect at the leading edge is very significant. This is because of the spiral pipe design at the leading edge. After the cooling fluid enters the spiral pipe, its velocity gradually increases ( Figure 9 as can be seen in [figure]), making the turbulence degree of this part high, the heat transfer boundary layer thinner, and enhancing the heat transfer efficiency. It can be seen that the cooling effect is very good at the pressure surface and the suction surface of the leading edge.

[0048] Regarding traditional gas-turbine vanes, the cold air is ejected through discrete gas film holes on both sides of the vane, forming a cooling gas film on the vane surface to isolate the mainstream high-temperature gas, thereby reducing the temperature of the mainstream gas flowing to the vanes of the subsequent stages, resulting in an impact on the overall performance of the gas turbine. In the present invention, the gas film holes on the suction side and the pressure side of the vane are deleted because the air film holes in the middle chord are replaced by vertical pipes, and the air outlet is located on the upper wall of the vane without mixing with the mainstream, thus reducing the amount of cold gas mixed with the hot gas.

[0049] Figure 9 is the streamline diagram of the internal cooling gas of the numerical simulation result. It can be found from the figure that the coolant flows at a high speed in the pipe, and a higher flow phenomenon is observed in the spiral cooling channel at the leading edge. The ultra-high flow velocity, accompanied by intense turbulent motion, significantly enhances the heat transfer capacity of this part. As Figure 10 shown, a low-temperature region of the vane can also be observed at the spiral pipe on the pressure surface of the leading edge of the vane.

[0050] In summary, the gas turbine blade with a near-wall vertical channel cooling for the airless film of the present invention has superior performance, lower cost, and higher thermal conversion efficiency compared to the cooling blades with traditional structures in the prior art.

[0051] In addition to the above embodiments, the present invention may have other implementation manners. Any technical solutions formed by equivalent replacement or equivalent transformation fall within the protection scope required by the present invention.

Claims

1. A gas turbine blade with a gasless film cooled by a near-wall vertical channel, comprising a blade body (100), the blade body (100) including a pressure side and a suction side, characterized in that: The blade body (100) is respectively provided with a first cooling chamber (1) and a second cooling chamber (2) vertically penetrating the blade body (100) along the blade height direction. The first cooling chamber (1) is arranged near the leading edge of the blade body (100), and the second cooling chamber (2) is arranged near the trailing edge of the blade body (100). The blade body (100) is further provided with a plurality of first cooling passages (3a) opened on the pressure side, a plurality of second cooling passages (3b) opened on the suction side, and a third cooling passage (4) opened on the leading edge. The first cooling passages (3a), the second cooling passages (3b), and the third cooling passage (4) are all arranged along the blade height direction and vertically penetrate the blade body (100). The third cooling passage (4) is integrally spiral.

2. The gas turbine blade with an airless film and cooled by a near-wall vertical channel according to claim 1, wherein: The leading edge portion of the blade body (100) is further provided with a plurality of fourth cooling passages (5). The fourth cooling passages (5) are equidistantly distributed circumferentially along the leading edge portion of the blade body (100), and the plurality of fourth cooling passages (5) are all arranged along the blade height direction and penetrate the blade body (100).

3. The gas turbine blade with a gasless film and cooled by a near-wall vertical channel according to claim 2, characterized in that: The trailing edge portion of the blade body (100) is provided with a plurality of fifth cooling passages (6). The cross-sectional areas of the fifth cooling passages (6) gradually decrease from the portion near the leading edge of the blade body (100) to the portion far from the leading edge of the blade body (100). The plurality of fifth cooling passages (6) are all arranged along the blade height direction of the blade body (100) and penetrate the blade body (100).

4. The gas turbine blade with near-wall vertical channel cooling for airless film according to claim 2, characterized in that: The number of the fourth cooling passages (5) is three, and the diameter is 1.8 mm.

5. The gas turbine blade with a gasless film cooled by a near-wall vertical channel according to claim 3, wherein: The number of the fifth cooling passages (6) is six. The diameters of the six fifth cooling passages (6) are 3.5 mm, 2.4 mm, 2 mm, 1.6 mm, 1.4 mm, and 1 mm respectively.

6. The gas turbine blade with a gasless film and cooled by a near-wall vertical channel according to claim 1, characterized in that: The first cooling passages (3a) and the second cooling passages (3b) are S-shaped.

7. The gas turbine blade with a gasless film and cooled by a near-wall vertical channel according to claim 6, characterized in that: A transverse channel transition is provided between the arc segments of the S-shape of the first cooling passages (3a) and the second cooling passages (3b). The length of the arc segment of the S-shape of the first cooling passages (3a) and the second cooling passages (3b) is 15 mm, and the length of the transverse channel is 4 mm.

8. The gas turbine blade with a gasless film and cooled by a near-wall vertical channel according to claim 1, characterized in that: An air inlet chamber is arranged on the same side of the first cooling passages (3a) and the second cooling passages (3b).

9. The gas turbine blade with a non-air film and cooled by a near-wall vertical channel according to claim 1, characterized in that: Air inlet chambers are arranged alternately on both sides of the first cooling passages (3a) and the second cooling passages (3b).

10. The gas turbine blade with a wall-less film cooled by a near-wall vertical channel according to any one of claims 1 to 9, characterized in that: Twelve first cooling passages (3a) are equidistantly distributed along the profile line direction of the pressure side of the blade body (100), and fourteen second cooling passages (3b) are equidistantly distributed along the profile line direction of the suction side of the blade body (100). The diameters of the first cooling passages (3a) and the second cooling passages (3b) are both 2 mm.

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