A guide cooling blade and a turbine guide

Abstract

The present application discloses a guide cooling blade and a turbine guide. The guide cooling blade includes a blade base, a cavity is provided inside the blade base, an impact tube is provided inside the cavity, and an impact hole is provided on the impact tube toward the inner wall of the leading edge of the blade base and a cold air inlet for introducing cold air. The turbine guide includes the above-mentioned guide cooling blade, and also includes an inner ring and an outer ring coaxially arranged from the inside to the outside, an annular airflow channel is formed between the outer ring and the inner ring, an outer ring hole is penetrated on the outer ring, and an inner ring hole is penetrated on the inner ring to the annular airflow channel. The guide cooling blade is sequentially penetrated by the inner ring hole, the annular airflow channel, and the outer ring hole, and the blade base is arranged in the annular airflow channel. The present application is used to solve the technical problems of poor cooling stability and low cold air utilization rate in the existing cooling technology for guide blades.

Description

Technical Field

[0001] The present application relates to the technical field of gas turbine engines, and in particular to a guide cooling blade and a turbine guide vane. Background Art

[0002] The aviation auxiliary power unit (APU) is a branch of the gas turbine engine. Its main function is to start the main engine of the aircraft and provide electricity to the aircraft. In the early stage, the APU used a non-cooled turbine guide vane to meet the requirements of aircraft manufacturers. However, with the development of science and technology, aircraft manufacturers have higher and higher requirements not only for the performance of the main engine, but also for the APU. The previous non-cooled turbine guide vane can no longer meet the development requirements of advanced APU. In order to adapt to the higher turbine inlet temperature of advanced APU, it is necessary to take appropriate cooling measures for the turbine guide vane. A major development trend of gas turbine engines at present is to develop towards higher power-to-weight ratio and higher thermodynamic cycle parameters. In order to improve the power-to-weight ratio of the engine, the gas temperature inlet of the turbine increases year by year. The turbine guide vane is usually set behind the high-temperature gas. The working conditions of the turbine guide vane are very bad. It is the component with the largest heat load in the engine. The temperature distribution of the guide vane is uneven, and it has to bear large aerodynamic forces and unstable pulsating loads. Therefore, adopting effective cooling measures, reducing the working temperature of the blades, and reducing the thermal stress of the blades are a strong guarantee for the safe and reliable operation of the engine.

[0003] In order to cool the guide blades, the conventional method is to enhance convection cooling. Enhanced convection cooling is to arrange ribs or spoiler columns in the internal channel of the blade to increase the heat exchange area between the cold air and the inner wall of the guide blade, and increase the turbulence of the cold air flow. Turbulence refers to the degree of turbulence of the flow. The more chaotic the flow, the stronger the heat exchange, so that the guide blades can be quickly and evenly heat exchanged. However, it is difficult to quickly cool the leading edge of the blade by enhanced convection cooling, and it is difficult to quickly eliminate the thermal stress at the leading edge of the blade. After that, in order to effectively reduce the thermal stress at the leading edge of the blade, the blade can be cooled by impact cooling. Impact cooling uses one or more jets of cold air to impact the inner surface of the blade. Its characteristic is that it has a high heat transfer coefficient in the jet stagnation area, which can quickly reduce the temperature of the impact area, but the jet structure will weaken the structural strength of the inner surface of the blade. At present, the most effective cooling method is film cooling. Film cooling refers to opening one or more exhaust film holes on the blades, and cold air flows out from the nozzle to prevent direct contact between the gas and the blades. Although film cooling can quickly eliminate the thermal stress on the leading edge of the blade and is not easy to weaken the structural strength of the inner surface of the blade, the cooling effect of film cooling is affected by factors such as the flow rate and flow direction of the high-temperature gas airflow, and the working stability is poor. At the same time, more cold air is required, and the utilization rate of cold air is low. Summary of the invention

[0004] Therefore, the present application aims to solve the technical problems of poor cooling stability and low cooling air utilization rate in the existing cooling technology for guide blades, thereby providing a guide cooling blade and a turbine guide.

[0005] In order to solve the above technical problems, the technical solution of the present application is as follows: a guide cooling blade, comprising a blade base, a cavity is provided inside the blade base, an impact tube is provided in the cavity, and the impact tube is provided with an impact hole facing the inner wall of the leading edge of the blade base and a cold air inlet for introducing cold air.

[0006] Preferably, a cold air channel is reserved between the outer wall of the impact tube and the inner wall of the cavity to allow the cold air after impact to flow out.

[0007] Preferably, a first disturbance structure is provided on the inner wall of the cavity, and a gap less than or equal to 0.1 mm is reserved between the first disturbance structure and the outer wall of the impact tube.

[0008] Preferably, a second disturbance structure is provided on an inner wall of an end surface of the cavity, and an end surface of the second disturbance structure is arranged closely against an end surface of the impact tube.

[0009] Preferably, a process boss is provided on an outer wall of one end surface of the cavity, and a gap of less than or equal to 0.1 mm is reserved between an outer wall of one end surface of the impact tube and an inner wall of the process boss.

[0010] Preferably, the tail of the blade base is provided with an "L"-shaped notch and a trailing edge hole connected to the "L"-shaped notch, the inlet of the trailing edge hole is connected to the cavity, the outlet of the trailing edge hole is located on the short side of the "L"-shaped notch, and one long side of the trailing edge hole is coplanar with the long side of the "L"-shaped notch.

[0011] The present application also provides a turbine guide, comprising the guide cooling blades described in any one of the above items, and also comprising an inner ring and an outer ring coaxially arranged from the inside to the outside, an annular airflow channel is formed between the outer ring and the inner ring, the outer ring is penetrated by an outer ring hole, the inner ring is penetrated by an inner ring hole extending to the annular airflow channel, the guide cooling blades are penetrated by the inner ring hole, the annular airflow channel and the outer ring hole in sequence, and the blade base is arranged in the annular airflow channel.

[0012] Preferably, the outer ring is provided with air holes, the outer edges of the outer ring and the inner ring overlap with the gas outlet end of the flame tube, and a gap is reserved at the overlapping parts of the outer ring, the inner ring and the gas outlet end of the flame tube.

[0013] Preferably, it also includes an enclosing ring coaxially arranged on the inner side of the inner ring, the enclosing ring is connected to an outer edge of the inner ring, and the other outer edge of the inner ring overlaps the gas outlet end of the flame tube.

[0014] Preferably, the thickness of the portion where the inner ring contacts the containing ring is greater than the thickness of other portions of the inner ring, and the inner ring and the containing ring are connected via fasteners.

[0015] The technical solution of this application has the following advantages:

[0016] 1. The guide cooling blade provided by the present application has an impact hole facing the inner wall of the leading edge of the blade base after the impact tube is provided. When the cold air enters the impact tube and flows out from the impact hole, the cold air directly exchanges heat with the inner wall of the leading edge of the blade base, and accurately and quickly cools the leading edge of the blade base, thereby achieving the purpose of reducing thermal fatigue damage to the leading edge of the blade base. In addition, compared with the heat exchange method of directly impacting the inner wall of the leading edge of the blade base with cold air, in the heat exchange method provided by the present application, the cold air enters the impact tube and then flows out from the impact hole, and the impact strength of the cold air has been greatly reduced, which will not cause a strong impact on the inner wall of the leading edge of the blade base, and thus will not damage the structural strength of the inner surface of the leading edge of the blade base. Moreover, compared with the air film cooling method mentioned in the background technology, the cooling method provided by the present application is not affected by the external high-temperature combustion gas, and can provide a stable cooling effect for the leading edge of the blade base. At the same time, in the technical solution of the present application, after the cold air cools the inside of the leading edge of the blade base, it continues to cool the inside of the blade base cavity, thereby improving the utilization rate of the cold air and reducing waste.

[0017] 2. The guide cooling blade provided in the present application cools the inner wall of the leading edge of the blade base with cold air and then evenly cools the inner wall of the blade base cavity along the cold air channel, which can cool the entire blade base and help reduce the thermal stress of the blade base.

[0018] 3. The guide cooling blade provided by the present application has a first disturbance structure in the cavity that disturbs the cold air flow, which can increase the heat exchange area between the cold air flow and the inner wall of the cavity, increase the turbulence of the cold air flow, and thus improve the heat exchange efficiency. The gap between the first disturbance structure and the outer wall of the impact tube is set to be less than or equal to 0.1 mm to prevent the first disturbance structure from getting stuck in the impact tube and affecting the assembly between the impact tube and the guide cooling blade.

[0019] 4. The guide cooling blade provided in the present application and the second disturbance structure can limit the impact tube and also disturb the flow of cold air inside the impact tube, thereby accelerating the flow speed of the cold air and improving the heat exchange efficiency.

[0020] 5. The guide cooling blade provided in the present application has a process boss on the outer wall of the cavity end face, which is then connected to the impact tube via the process boss, thereby reducing the difficulty of connecting the impact tube to the process boss and facilitating assembly.

[0021] 6. In the guide cooling blade provided by the present application, all the cold air in the cavity is discharged from the trailing edge hole, and the discharged cold air flows closely to the long side of the L-shaped notch, and air film cooling is performed on the trailing edge of the blade base. The long side of the trailing edge hole is coplanar with the long side of the L-shaped notch, which can ensure the stability of the cold air flow field and make the air film cover uniformly.

[0022] 7. In the turbine guide vane provided in the present application, the high-temperature combustion gas flows along the annular airflow channel, and the cold air can quickly cool the guide blades, outer ring, and inner ring to prevent the high-temperature combustion gas from damaging the structural strength of the turbine guide vane.

[0023] 8. The turbine guide provided by the present application has two cold air sources to cool the front and back of the turbine guide. After the two cold air sources preliminarily cool the front and back of the turbine guide, they flow into the annular air flow channel from the gap between the outer ring, the inner ring and the gas outlet of the flame tube to perform secondary cooling on the inner ring and the outer ring. The inner ring and the outer ring are the main heat-receiving parts, and secondary cooling of the inner ring and the outer ring can prevent their structural strength from being damaged, which helps to extend the service life of the turbine guide.

[0024] 9. The turbine guide provided by the present application has a containment ring surrounded by cold air, which has a lower temperature. The containment ring can absorb the heat transferred by the inner ring, which is conducive to reducing the working temperature of the inner ring. In addition, the contact part between the inner ring and the containment ring is thickened, which helps to improve the heat transfer efficiency between the inner ring and the containment ring. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] In order to more clearly illustrate the specific implementation methods of the present application 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 application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0026] Figure 1 It is a schematic diagram of the cooperation between the turbine guide and the flame tube in the embodiment of the present application;

[0027] Figure 2 Schematic diagram of the overall structure of the turbine guide device in the embodiment of the present application;

[0028] Figure 3 A schematic diagram of the positional relationship between the guide cooling blades and the turbine guide vane in an embodiment of the present application;

[0029] Figure 4 is a schematic diagram of the overall structure of the guide cooling blade in the embodiment of the present application;

[0030] Figure 5 It is a schematic diagram of the overall structure of the impact tube in the embodiment of the present application;

[0031] Figure 6 is a longitudinal sectional view of a guide cooling blade in an embodiment of the present application;

[0032] Figure 7 In the embodiment of this application Figure 6 An enlarged schematic diagram of the C portion;

[0033] Figure 8 It is a cross-sectional view of the guide cooling blade along line B in an embodiment of the present application;

[0034] Fig. 9 It is a cross-sectional view of the guide cooling blade along line A in an embodiment of the present application.

[0035] Explanation of the reference numerals: 1. blade base; 2. impact tube; 3. outer ring; 4. inner ring; 5. air vent; 6. annular air flow channel; 7. impact hole; 8. first disturbance structure; 9. second disturbance structure; 10. spoiler column; 11. trailing edge hole; 12. middle arc surface; 13. process boss; 14. partition rib; 15. flame tube outer ring; 16. flame tube inner ring; 17. flame tube bird's beak; 18. containment ring; 19. inner ring hole; 20. outer ring hole. DETAILED DESCRIPTION

[0036] The technical solution of the present application will be described clearly and completely below in conjunction with the accompanying drawings. Obviously, the described embodiments are part of the embodiments of the present application, rather than all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present application.

[0037] In the description of this application, it should be noted that, unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be indirectly connected through an intermediate medium, or it can be the internal communication of two components. For ordinary technicians in this field, the specific meanings of the above terms in this application can be understood according to specific circumstances.

[0038] Example 1

[0039] The present embodiment provides a guide cooling blade, including a blade base 1, the shape of the blade base 1 is as follows: Figure 1 , Figure 4-Figure 9 As shown, Figure 6 Line A is the middle arc surface 12 of the cavity. Figure 1The arrow a in the figure indicates the flow direction of the high-temperature combustion gas, and the arrows b and c indicate the flow direction of the cold air source. The leading edge of the blade base 1 is thicker than the trailing edge of the blade base 1. A cavity is provided inside the blade base 1. The shape of the cavity is similar to that of the blade base 1. An impact tube 2 is provided in the cavity. The impact tube 2 is inserted into the blade base 1 from one end face of the blade base 1. The cross-sectional shape of the impact tube 2 is similar to that of the cavity. An impact hole 7 and a cold air inlet for introducing cold air are provided on the impact tube 2. The impact hole 7 is arranged toward the inner wall of the leading edge of the blade base 1, and the cold air inlet is arranged toward the end face where the impact tube 2 is inserted into the blade base 1. The impact holes 7 are arranged in three rows, and the number of impact holes 7 in each row varies from 4 to 8. The middle row of impact holes 7 in the three rows of impact holes 7 faces the middle of the leading edge of the blade base 1, the upper row of impact holes 7 faces the upper side wall of the leading edge of the blade base 1, and the lower row of impact holes 7 faces the lower side wall of the leading edge of the blade base 1. In addition, the diameter, number of holes, arrangement, etc. of the impact holes 7 may be changed according to different guides, and the impact holes 7 may also be provided in other parts of the impact tube 2 .

[0040] The working process of the guide cooling blade provided in this embodiment is as follows: the guide cooling blade is usually used in conjunction with a guide, and the leading edge of the guide cooling blade is set toward the high-temperature gas. When the high-temperature gas rushes toward the guide blade, the high-temperature gas thermally impacts the leading edge of the blade substrate 1. The temperature of the leading edge of the blade substrate 1 is higher than that of other parts of the blade substrate 1. If this state is maintained for a long time, it is easy to cause thermal fatigue at the leading edge of the blade substrate 1. At this time, cold air can be introduced into the impact tube 2 through the cold air inlet. When the cold air enters the impact tube 2 and flows out from the impact hole 7, the cold air directly exchanges heat with the inner wall of the leading edge of the blade substrate 1, and accurately and quickly cools the leading edge of the blade substrate 1, thereby reducing the thermal fatigue damage of the leading edge of the blade substrate 1. In addition, compared with the heat exchange method of directly impacting the inner wall of the leading edge of the blade substrate, in the heat exchange method provided in this application, the cold air enters the impact tube 2 and then flows out from the impact hole 7. The impact strength of the cold air has been greatly reduced, and will not cause a strong impact on the inner wall of the leading edge of the blade substrate 1, so it will not damage the structural strength of the inner surface of the leading edge of the blade substrate 1. Moreover, compared with the air film cooling method mentioned in the background technology, the cooling method provided by the present application is not affected by external high-temperature combustion gas, and can provide a stable cooling effect for the leading edge of the blade base 1. At the same time, in the technical solution of the present application, the cold air cools the inside of the leading edge of the blade base 1 and then continues to cool the inside of the cavity of the blade base 1, thereby improving the utilization rate of the cold air and reducing waste.

[0041] Example 2

[0042] As a further improvement of Example 1, a cold air channel is reserved between the outer wall of the impact tube 2 and the inner wall of the cavity to allow the cold air to flow out after impact. After the cold air cools the inner wall of the leading edge of the blade base 1, it evenly cools the inner wall of the cavity of the blade base 1 along the cold air channel, cooling the entire blade base 1, which helps to reduce the thermal stress of the blade base 1.

[0043] The inner wall of the cavity is provided with a first disturbance structure 8, and a gap of less than or equal to 0.1 mm is reserved between the first disturbance structure 8 and the outer wall of the impact tube 2. The first disturbance structure 8 is in the shape of a hemispherical boss, and the hemispherical boss in the cavity plays a role in disturbing the cold air flow, which can increase the heat exchange area between the cold air flow and the inner wall of the cavity, increase the turbulence of the cold air flow, and thus improve the heat exchange efficiency. The gap between the hemispherical boss and the outer wall of the impact tube 2 is set to be less than or equal to 0.1 mm to prevent the hemispherical boss from getting stuck in the impact tube 2 and thus affecting the assembly between the impact tube 2 and the guide cooling blade.

[0044] A second disturbance structure 9 is provided on the inner wall of the end surface of the cavity away from the cold air inlet, and one end surface of the second disturbance structure 9 is arranged close to the end surface of the impact tube 2. The structure of the second disturbance structure 9 is a semi-cylindrical boss, and the second disturbance structure 9 can limit the impact tube 2 and also disturb the cold air flow inside the impact tube 2, thereby accelerating the flow speed of the cold air and improving the heat exchange efficiency.

[0045] The outer wall of the end surface of the cavity corresponding to the cold air inlet is provided with a process boss 13, and one end surface of the impact tube 2 is connected to the process boss 13. In general processes, the impact tube 2 is connected to the blade base 1 by welding. The provision of the process boss 13 reduces the difficulty of welding the impact tube 2 to the blade base 1, and facilitates the assembly of the impact tube 2 and the blade base 1.

[0046] A spoiler column 10 is provided near the trailing edge of the cavity. The two trailing edge side walls of the blade base 1 are one long and one short, so the trailing edge of the blade base 1 forms an L-shaped notch. A trailing edge hole 11 connected to the "L"-shaped notch is provided at the tail of the blade base 1. The inlet of the trailing edge hole 11 is connected to the cavity, and the outlet of the trailing edge hole 11 is located on the short side of the "L"-shaped notch. One long side of the trailing edge hole 11 is coplanar with the long side of the "L"-shaped notch. The trailing edge hole 11 is a plurality of rectangular holes distributed radially along the blade base 1. The holes are separated by partition ribs 14. The trailing edge hole 11 is connected to the cavity. The cold air in the cavity passes through the trailing edge hole 11 and is discharged along the L-shaped notch. After being discharged from the L-shaped notch, the cold air maintains a uniform and stable flow along the long side of the trailing edge of the blade base 1, which helps to quickly discharge the heat in the cavity. The disturbance of the spoiler column 10 implements enhanced convection cooling inside the cavity, and part of the cold air passes through the cold air channel to implement convection cooling on the cavity, and finally all the cold air is discharged from the trailing edge hole 11, and the discharged cold air flows closely to the long side of the L-shaped notch, and implements air film cooling on the trailing edge of the blade base 1. The long side of the trailing edge hole 11 is coplanar with the long side of the L-shaped notch, which can ensure the stability of the cold air flow field and make the air film cover uniform.

[0047] Example 3

[0048] This embodiment provides a turbine guide vane, which includes the guide cooling vane described in Embodiment 1 or 2. Figure 1-Figure 3 As shown, the turbine guide vane further includes an inner ring 4 and an outer ring 3 coaxially arranged from the inside to the outside, and an annular airflow channel 6 is formed between the outer ring 3 and the inner ring 4. Figure 1 As shown, the annular airflow channel 6 is generally in an inverted L shape when viewed from the central cross-sectional view. After the annular airflow channel 6 extends a certain length along the axial direction at the turbine guide inlet, it turns 90 degrees to form an airflow channel distributed along the radial direction. Figure 3 As shown, the outer ring 3 is provided with an outer ring hole 20, the inner ring 4 is provided with an inner ring hole 19 which is provided to the annular air flow channel 6, the guide cooling blades are provided with the inner ring hole 19, the annular air flow channel 6, and the outer ring hole 20 in sequence, and the blade base 1 is placed in the annular air flow channel 6. The high-temperature combustion gas impacts the annular air flow channel 6, and the high-temperature combustion gas flows along the annular air flow channel 6. The cold air can quickly cool the guide blades, the outer ring 3, and the inner ring 4, so as to prevent the high-temperature combustion gas from damaging the structural strength of the turbine guide.

[0049] Example 4

[0050] This embodiment is a further supplement to the embodiment 3. The outer ring 3 is provided with a plurality of air vents 5 along the circumferential direction. The outer edges of the outer ring 3 and the inner ring 4 of the turbine guide overlap with the gas outlet end of the flame tube. A gap is reserved at the overlap between the outer ring 3, the inner ring 4 and the gas outlet end of the flame tube. The gas outlet end of the flame tube includes the outer ring 15 of the flame tube, the inner ring 16 of the flame tube, and the bird's beak 17 of the flame tube. The outer ring 3 overlaps with the outer ring 15 of the flame tube, and the overlap gap △1=0.2-0.4mm. The two sides of the inner ring 4 overlap with the inner ring 16 of the flame tube and the bird's beak 17 of the flame tube, respectively. The overlap gap △2 between the inner ring 4 and the inner ring 16 of the flame tube is 0.2-0.4mm, and the overlap gap between the inner ring 4 and the bird's beak 17 of the flame tube is not specifically limited. Figure 1 As shown, there are two cold air sources b and c to cool the front and back of the turbine guide. After the cold air source c performs preliminary cooling on the back of the outer ring 3, it flows into the annular air flow channel 6 from the overlap between the outer ring 3 and the outer ring 15 of the flame tube to perform secondary cooling on the outer ring 3. A part of the cold air source b flows into the impact tube 2 through the cold air inlet, and the other part flows into the annular air flow channel 6 from the overlap between the inner ring 4 and the inner ring 16 of the flame tube and the bird beak 17 of the flame tube to perform secondary cooling on the inner ring 4. As the main heat-receiving parts, the inner ring 4 and the outer ring 3 are subjected to secondary cooling to prevent their structural strength from being damaged, which helps to extend the service life of the turbine guide. A part of the cold air source c also enters the annular air flow channel 6 through the vent 5 to implement air film cooling on the front of the outer ring 3, which helps to reduce the damage to the turbine guide caused by the impact of high-temperature combustion gas.

[0051] The turbine guide also includes an enclosing ring 18 coaxially arranged on the inner side of the inner ring 4. The enclosing ring 18 is connected to an outer edge of the inner ring 4, and the other outer edge of the inner ring 4 overlaps the gas outlet end of the flame tube. The thickness of the portion where the inner ring 4 contacts the enclosing ring 18 is greater than the thickness of other portions of the inner ring 4. The inner ring 4 and the enclosing ring 18 are connected by fasteners, and the fasteners can be bolts. The enclosing ring 18 is surrounded by a cold air source b and has a lower temperature. The enclosing ring 18 can absorb the heat transferred by the inner ring 4, which is beneficial to reducing the working temperature of the inner ring 4. Thickening the portion where the inner ring 4 contacts the enclosing ring 18 helps to improve the heat transfer efficiency between the inner ring 4 and the enclosing ring 18.

[0052] Obviously, the above embodiments are merely examples for the purpose of clear explanation, and are not intended to limit the implementation methods. For those skilled in the art, other different forms of changes or modifications can be made based on the above description. It is not necessary and impossible to list all the implementation methods here. The obvious changes or modifications derived therefrom are still within the scope of protection created by this application.

Claims

1. A turbine guide vane, characterized in that: The guide cooling blade comprises a blade base (1), a cavity is provided inside the blade base (1), an impact tube (2) is provided in the cavity, an impact hole (7) facing the inner wall of the leading edge of the blade base (1) and a cold air inlet for introducing cold air are provided on the impact tube (2); a second disturbance structure (9) is provided on the inner wall of the end surface of the cavity away from the cold air inlet, and one end surface of the second disturbance structure (9) is arranged in close contact with the end surface of the impact tube (2); The turbine guide vane further comprises an inner ring (4) and an outer ring (3) coaxially arranged from the inside to the outside, an annular airflow channel (6) is formed between the outer ring (3) and the inner ring (4), an outer ring hole (20) is penetrated on the outer ring (3), an inner ring hole (19) penetrated to the annular airflow channel (6) is provided on the inner ring (4), the guide cooling blade is penetrated through the inner ring hole (19), the annular airflow channel (6), and the outer ring hole (20) in sequence, and the blade base (1) is arranged in the annular airflow channel (6), the outer ring (3) is provided with an air vent (5), the outer edges of the outer ring (3) and the inner ring (4) overlap with the air outlet end of the flame tube, a gap is reserved at the overlapped portion between the outer ring (3), the inner ring (4) and the air outlet end of the flame tube, and two cold air sources cool the front and back sides of the turbine guide vane.

2. The turbine guide vane according to claim 1, characterized in that: A cold air channel is reserved between the outer wall of the impact tube (2) and the inner wall of the cavity to allow the cold air after impact to flow out.

3. The turbine guide vane according to claim 1, characterized in that: The inner wall of the cavity is provided with a first disturbance structure (8), and a gap less than or equal to 0.1 mm is reserved between the first disturbance structure (8) and the outer wall of the impact tube (2).

4. The turbine guide vane according to claim 1, characterized in that: An outer wall of one end surface of the cavity is provided with a process boss, and a gap of less than or equal to 0.1 mm is reserved between an outer wall of one end surface of the impact tube (2) and an inner wall of the process boss.

5. The turbine guide vane according to any one of claims 1 to 4, characterized in that: The blade base (1) is provided at the rear portion with an "L"-shaped notch and a trailing edge hole (11) connected to the "L"-shaped notch, the inlet of the trailing edge hole (11) being in communication with the cavity, the outlet of the trailing edge hole (11) being located on the short side of the "L"-shaped notch, and one long side of the trailing edge hole (11) being coplanar with the long side of the "L"-shaped notch.

6. The turbine guide vane according to claim 1, characterized in that: It also comprises an enclosing ring (18) coaxially arranged on the inner side of the inner ring (4), the enclosing ring (18) being connected to an outer edge of the inner ring (4), and the other outer edge of the inner ring (4) being overlapped with the gas outlet end of the flame tube.

7. The turbine guide vane according to claim 6, characterized in that: The thickness of a portion of the inner ring (4) in contact with the containing ring (18) is greater than the thickness of other portions of the inner ring (4), and the inner ring (4) and the containing ring (18) are connected via a fastener.

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