Turbine stator blade of gas turbine and gas turbine using the same

Abstract

A turbine stator blade of a gas turbine and a gas turbine using the same, the turbine stator blade comprising: a blade body, the blade body comprising a pressure surface and a suction surface; wherein an internal cavity of the blade is formed by the pressure surface and the suction surface; an impact guide bushing, arranged in the internal cavity of the blade; the area between the internal cavity of the blade and the impact guide bushing is defined as an annular cavity; at least one flow limiting structure, arranged in the annular cavity; wherein the impact guide bushing is provided with impact cooling jet holes, and the blade body is provided with cooling air holes; the cooling air inside the impact guide bushing impacts the cooling jet through the impact cooling jet holes to the annular cavity, forming an impact cooling exhaust gas which is discharged outside the blade body through the cooling air holes; wherein the flow limiting structure is used to limit the free flow of the impact cooling exhaust gas in the annular cavity, thereby significantly improving the impact cooling effect.

Description

Technical Field

[0001] The present invention relates to the technical field of gas turbine design, and in particular to a turbine stator blade of a gas turbine and a gas turbine using the same. Background Art

[0002] With the improvement of gas turbine design technology, the temperature of gas at the inlet of gas turbine turbine is increasing, and the thermal load faced by turbine components is extremely high, which has long exceeded the limit that high-temperature materials can withstand. In order to ensure the safe and reliable operation of turbine blades, complex cooling design is required to keep the temperature and stress distribution of the blade body at a reasonable level.

[0003] When designing the cooling of the stator blades of a gas turbine with a temperature level of 1000-1200 degrees Celsius (corresponding to the first stage of an E-class gas turbine or the second stage of an F-class gas turbine), the most common cooling design scheme is to use impingement cooling inside the blade, and a cooling scheme that uses trailing edge injection (or trailing edge injection on the pressure surface) and film cooling (or leading edge spray film cooling) near the leading edge of the suction surface. The internal impingement cooling medium needs to be discharged into the mainstream combustion gas from the film holes or injection holes on the blade surface. When the number of film hole rows on the blade surface is small, the characteristic parameters and position of the film holes will have a significant impact on the distribution of the internal cooling medium between the film hole rows, thereby affecting the lateral flow of the fluid on the inner surface of the blade. This lateral flow will have a significant impact on the heat exchange on the inner surface of the blade. How to reasonably control the lateral flow of the cold air inside the blade according to the design intent so that it has a positive impact on cooling is a problem that gas turbine cooling designers are concerned about. Summary of the invention

[0004] In view of this, the main object of the present invention is to provide a turbine stator blade of a gas turbine and a gas turbine using the same, in order to at least partially solve at least one of the above-mentioned technical problems.

[0005] In order to achieve the above object, the technical solution of the present invention includes:

[0006] As one aspect of the present invention, a turbine stator blade of a gas turbine is provided, comprising:

[0007] A blade body, wherein the blade body comprises a pressure surface and a suction surface; wherein the pressure surface and the suction surface surround and form an internal cavity of the blade;

[0008] An impact guide bushing is arranged in the internal cavity of the blade; the area between the internal cavity of the blade and the impact guide bushing is defined as an annular cavity;

[0009] at least one flow limiting structure disposed in the annular cavity;

[0010] Wherein, the impact guide bushing is provided with impact cooling jet holes, and the blade body is provided with cold air holes; the cooling air inside the impact guide bushing is impact cooled by jets through the impact cooling jet holes to the annular cavity, forming impact cooling exhaust gas which is discharged out of the blade body through the cold air holes; wherein the flow limiting structure is used to limit the free flow of the impact cooling exhaust gas in the annular cavity.

[0011] As another aspect of the present invention, a gas turbine is provided, comprising the turbine stator blade as described above.

[0012] Based on the above technical solution, it can be known that the present invention has at least one or part of the following beneficial effects compared with the prior art:

[0013] (1) At least one flow limiting structure is arranged in the annular cavity between the inner surface of the blade body and the impingement guide bushing. The flow limiting structure is closely matched with the inner surface of the blade body and the impingement guide bushing to achieve the purpose of limiting the free flow of the impingement cooling exhaust gas;

[0014] (2) Impingement cooling of the exhaust gas can significantly improve the impingement cooling effect by purposefully changing or restricting the flow according to the design objectives, while also improving the convective cooling level, optimizing the flow parameters in the cooling air holes on the blade surface, and optimizing the air film coverage effect on the blade surface, thereby achieving the best cooling effect and engine performance without increasing the amount of cooling air;

[0015] (3) Simplify the turbine blade structure and reduce costs. BRIEF DESCRIPTION OF THE DRAWINGS

[0016] Figure 1 is a schematic diagram of the design of the turbine stator blades of the gas turbine of Comparative Example 1;

[0017] Figure 2 yes Figure 1 AA expanded view;

[0018] Figure 3 is a schematic diagram of the design of a turbine stator blade of a gas turbine according to Embodiment 1 of the present invention;

[0019] Figure 4 yes Figure 3 BB expanded view;

[0020] Figure 5 is a schematic diagram of the design of the turbine stator blades of the gas turbine of Comparative Example 2;

[0021] Figure 6 It is a schematic diagram of the design of the turbine stator blades of the gas turbine according to Example 2 of the present invention.

[0022] In the above drawings, the meanings of the reference numerals are as follows:

[0023] 1-suction surface; 2-pressure surface; 3-leading edge; 4-trailing edge; 5-leading edge stagnation line; 6-impact guide bushing; 7-air film hole at the leading edge of the suction surface; 8-pressure surface trailing edge injection hole; 9-annular cavity; 10-impact cooling jet; 11-impact cooling exhaust gas stagnation line; 12, 13-impact cooling exhaust gas; 14, 17-high-pressure cavity; 15-flow limiting structure; 16-cooling cavity partition plate; 18-through hole. DETAILED DESCRIPTION

[0024] The turbine stator blade provided by the present invention reasonably designs the internal cooling channel of the turbine stator blade without increasing the total cooling air volume, controls the lateral flow to develop in a direction that is beneficial to the cooling design, and thus maximizes the blade cooling efficiency.

[0025] In order to make the objectives, technical solutions and advantages of the present invention more clearly understood, the present invention is further described in detail below in conjunction with specific embodiments and with reference to the accompanying drawings.

[0026] As one aspect of the present invention, a turbine stator blade of a gas turbine is provided, comprising:

[0027] The blade body includes a pressure surface and a suction surface; wherein the pressure surface and the suction surface surround and form an internal cavity of the blade;

[0028] An impact guide bushing is arranged in the internal cavity of the blade; the area between the internal cavity of the blade and the impact guide bushing is defined as an annular cavity;

[0029] at least one flow limiting structure disposed in the annular cavity;

[0030] Among them, the impact guide bushing is provided with impact cooling jet holes, and the blade body is provided with cold air holes; the cooling air inside the impact guide bushing is impact cooled by the impact cooling jet through the impact cooling jet holes to the annular cavity, forming impact cooling exhaust gas which is discharged out of the blade body through the cold air holes; wherein the flow limiting structure is used to limit the free flow of the impact cooling exhaust gas in the annular cavity.

[0031] In an embodiment of the present invention, a length of the flow limiting structure in a blade height direction is less than or equal to a blade height of a blade body.

[0032] More specifically, in the annular cavity between the inner surface of the blade body and the impact guide bushing, the flow limiting structure can occupy the entire blade height direction or a portion of the blade height direction, and the length of the flow limiting structure in the blade height direction is adjusted according to the arrangement of the cooling air holes.

[0033] In an embodiment of the present invention, the cross-sectional shape of the current limiting structure includes a rectangle, a trapezoid or an arc.

[0034] More specifically, in the annular cavity between the inner surface of the blade body and the impact guide bushing, the cross-sectional shape of the flow limiting structure can be any polygon, including but not limited to a rectangle, a trapezoid, an arc segment, etc., while ensuring good sealing.

[0035] In an embodiment of the present invention, two ends of the flow limiting structure are in sealing contact with the inner surface of the blade body and the outer surface of the impact guide bushing respectively.

[0036] It is worth mentioning that the sealing contact specifically refers to a close fit to achieve a sealing effect.

[0037] In an embodiment of the present invention, the flow limiting structure and the blade body are integrally cast.

[0038] In other embodiments of the present invention, the flow limiting structure and the impact guide bushing are integrally formed.

[0039] In other embodiments of the present invention, the flow limiting structure is an independent structure formed by welding with the blade body or the impact guide bushing.

[0040] In the embodiments of the present invention, the position, quantity, shape, size and arrangement of the current limiting structure can be determined according to the design requirements of the cooling design profession.

[0041] In an embodiment of the present invention, the stator blades are processed by precision casting, and the blade body can be an independent casting structure of a single blade, or an integral casting structure of multiple blades.

[0042] In an embodiment of the present invention, the cooling air holes on the blade body include a plurality of rows of injection holes and / or air film holes arranged in an array.

[0043] In an embodiment of the present invention, the injection holes include trailing edge injection holes and / or pressure surface trailing edge injection holes; the film holes include suction surface film holes and / or suction surface front edge film holes.

[0044] In an embodiment of the present invention, the number of the impact guide bushing is one or more;

[0045] When a plurality of impingement guide bushings are arranged in the internal cavity of the blade body, a cooling cavity partition plate with through holes is arranged between two adjacent impingement guide bushings.

[0046] In an embodiment of the present invention, when the turbine stator blade includes a single blade internal cavity, an impact guide bushing is arranged in the blade internal cavity, and the cooling air hole includes a trailing edge injection hole and / or a pressure surface trailing edge injection hole; after the cooling air impacts and cools the inner surface of the blade body through the impact guide bushing, the impact cooling exhaust gas is discharged into the mainstream combustion gas through the trailing edge injection hole and / or the pressure surface trailing edge injection hole. The impact cooling exhaust gas in the annular cavity between the inner surface of the blade body and the impact guide bushing is separated by a flow limiting structure arranged in the annular cavity, so that the impact cooling exhaust gas in the annular cavity area cannot pass through the flow limiting structure, and can only be discharged into the mainstream combustion gas from the trailing edge injection hole and / or the pressure surface trailing edge injection hole according to the flow path formed by the flow limiting structure.

[0047] In a preferred embodiment of the present invention, when the turbine stator blade includes a single blade internal cavity, an impact guide bushing is arranged in the blade internal cavity, and the cold air hole includes a row of air film holes on the front edge of the suction surface and a row of injection holes on the trailing edge of the pressure surface, the flow limiting structure includes two, one of which is arranged between the suction surface near the trailing edge of the blade and the impact guide bushing, and the other flow limiting structure is arranged between the pressure surface near the leading edge of the blade and the impact guide bushing.

[0048] More specifically, after the cooling air impinges and cools the inner surface of the blade body through the impingement guide bushing, the impingement cooling exhaust gas is discharged into the mainstream combustion gas through at least one exhaust film hole on the blade body surface and the injection hole at the trailing edge of the pressure surface. The impingement cooling exhaust gas in the annular cavity between the inner surface of the blade body and the impingement guide bushing is separated by at least two flow limiting structures arranged in the annular cavity, so that the impingement cooling exhaust gas in a part of the annular cavity area can only be discharged from the cooling air hole near the blade body, and the impingement cooling exhaust gas in another part of the annular cavity area can only be discharged into the mainstream combustion gas from the injection hole at the trailing edge of the pressure surface.

[0049] In a preferred embodiment of the present invention, when there are multiple impact guide bushings and the cold air holes include a row of air film holes on the leading edge of the suction surface and a row of injection holes on the trailing edge of the pressure surface, the flow limiting structure includes two flow limiting structures, which are respectively arranged between the suction surface near the leading edge of the blade and the impact guide bushing and between the pressure surface near the leading edge of the blade and the impact guide bushing.

[0050] More specifically, the blade has a plurality of internal cavities, and at least one internal cavity of the blade is provided with an impact guide bushing. After the cooling air impacts and cools the inner surface of the blade body corresponding to the internal cavity of the blade through the impact guide bushing, the impact cooling exhaust gas is discharged into the mainstream combustion gas through the cooling air holes arranged on the surface of the blade body, and at least one flow limiting structure is arranged in the annular cavity between the blade body and the impact guide bushing, so that the flow direction of the impact cooling exhaust gas in the annular cavity area is changed, and is discharged into the mainstream combustion gas from the cooling air holes arranged on the blade body.

[0051] As another aspect of the present invention, a gas turbine is provided, comprising the turbine stator blade as described above.

[0052] The technical solution of the present invention is further described below in conjunction with comparative examples and specific embodiments. However, it should be noted that the following embodiments are only used to illustrate the technical solution of the present invention, but the present invention is not limited thereto.

[0053] Comparative Example 1

[0054] Figure 1 This is a schematic diagram of the design of the turbine stator blades of the gas turbine of Example 1. Figure 2 yes Figure 1 The expanded view along the AA dividing line. The gas turbine stator blade blade includes a suction surface 1, a pressure surface 2, a leading edge 3 and a trailing edge 4; the high-temperature combustion gas forms a leading edge stagnation line 5 in the leading edge 3 area on the outer surface of the blade blade; an impact guide bushing 6 is arranged inside the blade blade; at the same time, a row of suction leading edge air film holes 7 are arranged on the suction surface 1; a row of pressure surface trailing edge injection holes 8 are arranged on the pressure surface 2 of the blade blade; a high-pressure cavity 14 is formed inside the impact guide bushing 6; and an annular cavity 9 is formed between the impact guide bushing 6 and the blade blade.

[0055] When the high-pressure cooling air in the high-pressure cavity 14 impinges on the inner surfaces of the suction surface 1, the pressure surface 2, the leading edge 3, and the trailing edge 4 of the blade body through the impingement cooling jet 10 formed by the impingement cooling jet holes on the impingement guide bushing 6, the impingement cooling exhaust gas fills the annular cavity 9 and then is discharged into the mainstream combustion gas through the air film holes 7 at the front edge of the suction surface and the injection holes 8 at the trailing edge of the pressure surface. According to the principle of mass conservation, the cooling air mass flow rates of the impingement cooling jet 10, the annular cavity 9, the air film holes 7 at the front edge of the suction surface, and the injection holes 8 at the trailing edge of the pressure surface should be equal. The impingement cooling exhaust gas in the annular cavity 9 will automatically be divided into two parts, one part is the impingement cooling exhaust gas 12 flowing to the air film holes 7 at the front edge of the suction surface, and the other part is the impingement cooling exhaust gas 13 flowing to the injection holes 8 at the trailing edge of the pressure surface, and the two are separated by the impingement cooling exhaust gas stagnation line 11. The position of the impingement cooling exhaust gas stagnation line 11 can be obtained based on flow analysis, and the position is affected by the distribution of the impingement cooling jet 10, the geometric characteristics of the annular cavity 9, and the size and position of the air film holes 7 at the front edge of the suction surface and the injection holes 8 at the trailing edge of the pressure surface. When the conditions of the high-temperature combustion gas change, the impingement cooling exhaust gas stagnation line 11 will also fluctuate in time and space. This brings great uncertainty to the impingement cooling and convection cooling design in the annular cavity 9, which may weaken or enhance the internal cooling, thereby changing the original intention of the designer; it will also cause the cooling air injection flow rate of the air film holes 7 at the front edge of the suction surface and the injection holes 8 at the trailing edge of the pressure surface to change, thereby changing the air film coverage effect on the blade body surface.

[0056] The above problem is an inherent problem in the design of turbine stator blades of this type of traditional gas turbine, which may lead to substandard blade cooling design and threaten the safe and stable operation of the gas turbine equipped with such turbine stator blades.

[0057] Example 1

[0058] Figure 3 The invention is an improved design of a turbine stator blade of a gas turbine according to the first embodiment of the invention. Figure 4 yes Figure 3 The expanded view along the BB dividing line. The blade body of the turbine stator blade of the gas turbine includes a suction surface 1, a pressure surface 2, a leading edge 3 and a trailing edge 4; the high-temperature combustion gas forms a leading edge stagnation line 5 in the leading edge 3 area on the outer surface of the blade body; an impact guide bushing 6 is arranged inside the blade body; at the same time, a row of suction front edge air film holes 7 are arranged on the suction surface 1; a row of pressure surface trailing edge injection holes 8 are arranged on the pressure surface 2 of the blade body; a high-pressure cavity 14 is formed inside the impact guide bushing 6; an annular cavity 9 is formed between the impact guide bushing 6 and the blade body; the pressure surface near the leading edge and the suction surface near the trailing edge are arranged as follows Figure 3 , Figure 4 The current limiting structure 15 is shown.

[0059] When the high-pressure cooling air in the high-pressure cavity 14 impinges on the inner surfaces of the suction surface 1, the pressure surface 2, the leading edge 3 and the trailing edge 4 of the blade body through the impingement cooling jet 10 formed by the impingement cooling jet holes on the impingement guide bushing 6, the impingement cooling exhaust gas fills the annular cavity 9 and then is discharged into the mainstream combustion gas through the air film holes 7 at the front edge of the suction surface and the injection holes 8 at the trailing edge of the pressure surface. According to the principle of mass conservation, the cooling air mass flow rates of the impingement cooling jet 10, the annular cavity 9, the air film holes 7 at the front edge of the suction surface and the injection holes 8 at the trailing edge of the pressure surface should be equal. Due to the existence of the flow limiting structure 15, the impingement cooling exhaust gas in the annular cavity 9 is divided into two parts at the location of the flow limiting structure 15, one part is the impingement cooling exhaust gas 12 flowing to the air film holes 7 at the front edge of the suction surface, and the other part is the impingement cooling exhaust gas 13 flowing to the injection holes 8 at the trailing edge of the pressure surface, and the two are separated by the flow limiting structure 15. The impingement cooling exhaust gas stagnation line 11 automatically determined according to the mass conservation principle is replaced by the manually set flow limiting structure 15, so that the position is no longer affected by the distribution of the impingement cooling jet 10, the geometric characteristics of the annular cavity 9, and the size and position of the air film hole 7 at the front edge of the suction surface and the injection hole 8 at the rear edge of the pressure surface. When the conditions of the high-temperature combustion gas change, the impingement cooling exhaust gas stagnation line 11 (coinciding with the flow limiting structure 15) no longer fluctuates in time and space. By setting the flow limiting structure 15, the distribution ratio of the impingement cooling exhaust gas 12 flowing to the air film hole 7 at the front edge of the suction surface and the impingement cooling exhaust gas 13 flowing to the injection hole 8 at the rear edge of the pressure surface can be limited, thereby increasing the flexibility of the cooling design.

[0060] Comparative Example 2

[0061] Figure 5 The comparative example shows another conventional method for solving the problem of lateral flow of exhaust gas in single-cavity impingement cooling, that is, solving this problem by setting multiple chambers. In this case, the high-pressure cavity 14 is divided into two independent parts, namely, the high-pressure cavity 14 and the high-pressure cavity 17, by the cooling cavity partition plate 16. However, the problem of lateral flow of exhaust gas in single-cavity impingement cooling still exists in the single independent high-pressure cavity 14 and the high-pressure cavity 17.

[0062] Example 2

[0063] Figure 6 Shown for Figure 5 A specific embodiment of the improvement. By setting a flow limiting structure 15 in the high-pressure cavity 14 corresponding to the leading edge 3 area, the impingement cooling exhaust gas is divided into the impingement cooling exhaust gas 12 flowing to the air film hole 7 at the front edge of the suction surface and the impingement cooling exhaust gas 13 flowing to the injection hole 8 at the trailing edge of the pressure surface, wherein the impingement cooling exhaust gas 13 flowing to the injection hole 8 at the trailing edge of the pressure surface flows into the annular cavity 9 corresponding to the trailing edge 4 area through the through hole 18 set on the cooling cavity partition plate 16.

[0064] Through the above cooling design, the temperature level in the end wall area can be effectively reduced, the temperature gradient level can be reduced, and problems such as ablation, oxidation, and cracks caused by high temperature can be significantly alleviated.

[0065] The specific embodiments described above further illustrate the objectives, technical solutions and beneficial effects of the present invention in detail. It should be understood that the above description is only a specific embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principles of the present invention should be included in the scope of protection of the present invention.

Claims

1. A turbine stator blade of a gas turbine, characterized in that: include: A blade body, wherein the blade body comprises a pressure surface and a suction surface; wherein the pressure surface and the suction surface surround and form an internal cavity of the blade; An impact guide bushing is arranged in the internal cavity of the blade; the area between the internal cavity of the blade and the impact guide bushing is defined as an annular cavity; at least one flow limiting structure disposed in the annular cavity; The impingement guide bushing is provided with an impingement cooling jet hole, and the blade body is provided with a cooling air hole; the cooling air inside the impingement guide bushing is impinged by the impingement cooling jet hole to the annular cavity, forming an impingement cooling exhaust gas which is discharged out of the blade body through the cooling air hole; wherein the flow limiting structure is used to limit the free flow of the impingement cooling exhaust gas in the annular cavity; When the number of the impact guide bushing is one, and the cold air holes include a row of air film holes at the front edge of the suction surface and a row of injection holes at the trailing edge of the pressure surface, the flow limiting structures include two, one of which is arranged between the suction surface near the trailing edge of the blade and the impact guide bushing, and the other is arranged between the pressure surface near the leading edge of the blade and the impact guide bushing; The flow limiting structure divides the impact cooling exhaust gas in the annular cavity into impact cooling exhaust gas flowing to the air film holes at the front edge of the suction surface and impact cooling exhaust gas flowing to the injection holes at the trailing edge of the pressure surface at the location of the flow limiting structure, so that the impact cooling exhaust gas in the annular cavity area can only be discharged from the air film holes at the front edge of the suction surface and the injection holes at the trailing edge of the pressure surface according to the flow path formed by the flow limiting structure.

2. The turbine stator blade according to claim 1, characterized in that: The length of the flow-limiting structure in the blade height direction is less than or equal to the blade height of the blade body.

3. The turbine stator blade according to claim 1, characterized in that: The cross-sectional shape of the current limiting structure includes a rectangle, a trapezoid or an arc.

4. The turbine stator blade according to claim 1, characterized in that: The two ends of the flow limiting structure are in sealing contact with the inner surface of the blade body and the outer surface of the impact guide bushing respectively.

5. The turbine stator blade according to claim 4, characterized in that: The flow limiting structure is integrally cast with the blade body; The flow limiting structure and the impact guide bushing are integrally formed; or The flow limiting structure is formed by welding with the blade body or the impact guide bushing.

6. The turbine stator blade according to claim 1, characterized in that: The cooling air holes on the blade body include a plurality of rows of injection holes and / or air film holes arranged in an array.

7. The turbine stator blade according to claim 6, characterized in that: The injection holes include trailing edge injection holes and / or pressure surface trailing edge injection holes; The air film holes include suction surface air film holes and / or suction surface front edge air film holes.

8. The turbine stator blade according to claim 7, characterized in that: The number of the impact guide bushing is one or more; When a plurality of impingement guide bushings are arranged in the internal cavity of the blade body, a cooling cavity partition plate with through holes is arranged between two adjacent impingement guide bushings.

9. The turbine stator blade according to claim 8, characterized in that: When there are multiple impact guide bushings, and the cold air holes include a row of air film holes on the leading edge of the suction surface and a row of injection holes on the trailing edge of the pressure surface, the flow limiting structures include two, and the two flow limiting structures are respectively arranged between the suction surface close to the leading edge of the blade and the impact guide bushing and between the pressure surface close to the leading edge of the blade and the impact guide bushing.

10. A gas turbine, characterized in that: The invention comprises a turbine stator blade according to any one of claims 1 to 9.

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