A gas turbine with adjustable high-pressure air cooling

Abstract

The present invention discloses an adjustable high-pressure air-cooled gas turbine, comprising a booster pump, a radiator, an external pipeline, a control valve, a flow meter, a turbine casing volute, and a first-stage turbine guide device. The first-stage turbine guide device comprises an outer ring connected to the turbine casing volute, an inner ring, and an annular turbine guide blade installed between the inner ring and the outer ring. The annular turbine guide blade is formed by multiple groups of triple guide blades and a group of double guide blades. The triple guide blades and the double guide blades are both cast from the first-stage turbine guide blades. The first-stage turbine guide blade comprises a blade body, a blade-shaped impact tube 1, a blade-shaped impact tube 2, an upper edge plate, and a lower edge plate. The new cold air inlet and cold air outlet added by the present invention both enter and exit from the turbine casing volute, and this space is utilized as a cold air circulation air path to ensure the air tightness of the cold air. The flow ratio can be adjusted from 1% to 6% according to different working conditions to avoid cold air waste at low power and cold air shortage at high power.

Description

Technical Field

[0001] The present invention relates to the technical field of gas turbines, in particular to an adjustable high-pressure air-cooled gas turbine. Background Art

[0002] A certain type of gas turbine, with a power of 1-2 MW, is used as a fuel for combined cooling, heating, and power (CCHP) systems. Currently, an air-cooled first-stage guide vane (DIV) is used to cool the turbine blades. This air-cooled first-stage guide vane is standard equipment for long-life gas turbines. It draws high-pressure air from the compressor outlet into the guide vane's outer ring, where it enters the inner cavities of each guide vane and is ejected through inclined holes at the rear of the blade basin, where it joins the main flow. However, the diameter and number of holes in the cooling air flow path of the current air-cooled first-stage guide vane are fixed, making the flow rate unadjustable. Regardless of power level, it only provides a fixed cooling air flow rate of approximately 4% of the total airflow, which inevitably results in cooling air waste at low power levels and insufficient cooling air at high power levels. Furthermore, the cooling air holes are located on both the inner and outer rings, creating a lack of airtightness between the inner ring, the first-stage turbine guide vanes, and the outer ring. This results in the original cooling air being dispersed, reducing cooling efficiency.

[0003] Patent application publication number CN116104587A discloses a gas turbine guide vane for an aircraft engine. The guide vane comprises an annular inner ring, an annular outer ring, and guide vanes connecting the two rings. The inner ring is provided with a first air film hole, the outer ring is provided with a second air film hole, and the guide vane has a hollow cavity connecting the inner surface of the inner ring with the outer surface of the outer ring. An impingement tube is snap-fitted into the hollow cavity to provide vibration damping and airflow cooling for the guide vane. The front end of the guide vane is sleeved with the flame tube, and the rear end of the guide vane is connected to the inner casing of the turbine. The inner ring, outer ring, and guide vanes are integrally formed. The aperture and number of holes in the gas turbine guide vane are fixed, making it impossible to adjust the flow rate. Regardless of the power state, it can only provide a fixed cooling air flow rate of approximately 4% of the total airflow. This cannot avoid cooling air waste at low power and cooling air shortage at high power. Moreover, the cooling air holes are provided on both the inner and outer rings, resulting in insufficient airtightness between the inner ring, the first-stage turbine guide vanes, and the outer ring. As a result, the original cooling air is dispersed, reducing cooling efficiency. Summary of the Invention

[0004] In order to solve the problems of the prior art, the present invention provides a gas turbine with adjustable high-pressure air cooling.

[0005] The technical solution adopted in the present invention is:

[0006] A gas turbine with adjustable high-pressure air cooling, comprising a boost air pump, a radiator connected to the boost air pump, an external pipeline having one end connected to the radiator, a control valve and a flow meter provided on the external pipeline, a turbine casing volute connected to the other end of the external pipeline, and a first-stage turbine guide connected to the turbine casing volute;

[0007] The first-stage turbine guide vane includes an outer ring connected to the turbine casing volute, an inner ring, and an annular turbine guide vane installed between the inner ring and the outer ring. The annular turbine guide vane is formed by multiple sets of triple guide vanes and one set of double guide vanes. The triple guide vanes are integrally cast from three first-stage turbine guide vanes, and the double guide vanes are integrally cast from two first-stage turbine guide vanes. The first-stage turbine guide vane includes a blade body containing an inner cavity, two blade-shaped impact tubes 1 and 2 of different sizes disposed in the inner cavity of the blade body, an upper edge plate disposed above the blade body, and a lower edge plate disposed below the blade body. The blade-shaped impact tube 1 and the blade-shaped impact tube 2 both directly connect to the upper edge plate and the lower edge plate.

[0008] The first blade-shaped impact tube is larger than the second blade-shaped impact tube. The first blade-shaped impact tube is arranged at the front edge of the inner cavity, and the second blade-shaped impact tube is arranged at the rear edge of the inner cavity. The front edge and wall of the first blade-shaped impact tube are provided with impact cooling holes. The blade body is provided with double rows of cooling air discharge holes, and the impact cooling holes and the cooling air discharge holes are staggered in position.

[0009] A front upper support and a rear upper support connected to the inner side of the outer ring are provided above the upper edge plate, and a cold air inlet hole is provided on the side of the rear upper support. An annular cavity is formed between the outer ring and all the upper edge plates, the front upper support and the rear upper support. The cavity is communicated with all the leaf-shaped impact tubes 1, 2 and the cold air inlet hole.

[0010] Furthermore, a front lower support and a rear lower support connected to the outer side of the inner ring are provided at the front and rear below the lower edge plate.

[0011] Furthermore, the blade body is formed by stretching a smooth blade profile, and the stretching direction is perpendicular to the turbine shaft, which deflects the axial airflow at the combustion chamber outlet by 60 degrees, reduces the pressure and accelerates the ejection.

[0012] Furthermore, the leaf-shaped impact tube 1 and the leaf-shaped impact tube 2 have the same wall thickness of 0.3 mm, and the thickness between the two leaf-shaped impact tubes 1 is about 1±0.5 mm, which improves the rigidity and natural frequency. This design can improve the qualification rate and reduce the manufacturing cost.

[0013] Furthermore, the first-stage turbine guide blades are cast from a nickel-based high-temperature alloy and coated with a cobalt-chromium-aluminum-yttrium alloy to improve high-temperature oxidation resistance.

[0014] Furthermore, the lower edge plate is provided with an angular positioning groove, and the inner ring is connected to the lower edge plate by inserting the square tenon into the angular positioning groove.

[0015] Furthermore, the upper edge plate is provided with an outer ring positioning pin hole, and the outer ring is connected to the upper edge plate by inserting a positioning pin into the outer ring positioning pin hole.

[0016] Furthermore, the interface between the outer ring and the turbine casing volute is sealed and positioned with an expansion ring, and iron-based high-temperature alloy strips are inserted between the first-stage turbine guide blades for sealing, which also has a vibration damping function.

[0017] Furthermore, a gap is provided between the triple guide vane and the double guide vane, and between the triple guide vane and the triple guide vane, and the gap is 0.2±0.05 mm, and the gap is a compensation gap for thermal expansion and contraction.

[0018] Furthermore, an optical fiber is provided on the outer front wall of the blade-shaped impact tube 1 or the blade-shaped impact tube 2, and the light emitted by the red-hot guide vane is transmitted to the outside of the engine through the optical fiber, and the wall temperature is measured by a radiation temperature measuring instrument at the output end of the optical fiber.

[0019] Compared with the prior art, the present invention has the following beneficial effects:

[0020] (1) The present invention removes the original cold air inlet of the inner ring and the outer ring, and adds new cold air inlet and cold air outlet, both of which enter and exit from the turbine casing volute. The space between the turbine casing volute and the first-stage turbine guide is sealed and idle, and this space is used as a cold air circulation air path to ensure the air tightness of the cold air. In addition, the pressurized cold air is supplied by an external pipeline, and the flow ratio can be adjusted from 1% to 6% according to different working conditions, avoiding waste of cold air at low power and insufficient cold air at high power.

[0021] (2) The present invention uses triple guide vanes and double guide vanes for combined whole-ring casting and whole-ring turning and milling, which improves the qualification rate and productivity. It can be applied to higher gas temperatures (up to 1K℃) and thermal deformation stress, can reduce the number of dozens of parts, reduce production costs, reduce cold air leakage, improve efficiency, and extend the service life of the first-stage turbine guide vanes.

[0022] (2) The present invention coats the surface of the high-temperature alloy guide vane with a cobalt-chromium-aluminum-yttrium alloy to improve high-temperature oxidation resistance, increase the gas temperature, reduce fuel consumption, and improve fuel economy.

[0023] (4) The four arc legs (front upper support, rear upper support, front lower support, and rear lower support) of the present invention are stably supported in the inner and outer rings. Small gaps are provided between the triple guide vanes and the double guide vanes, and between the triple guide vanes and the triple guide vanes to allow for thermal expansion and contraction. Multiple damping locations reduce vibration stress. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Figure 1 The figure is a schematic diagram of the structure of a gas turbine with adjustable high-pressure air cooling;

[0025] Figure 2 It is a structural diagram of the first-stage turbine guide vane;

[0026] Figure 3 This is a schematic diagram of the double guide vane structure when viewed from above;

[0027] Figure 4 Schematic diagram of the top view of the double guide vane;

[0028] Figure 5 Schematic diagram of the structure of triple guide vanes;

[0029] Figure 6 It is a structural schematic diagram of annular turbine guide blades;

[0030] Figure 7 Schematic diagram of the structure of the impact tube.

[0031] Among them: 1-turbine casing volute; 2-first-stage turbine guide vane; 3-outer ring; 4-inner ring; 5-annular turbine guide blade; 6-triple guide vane; 7-double guide vane; 8-optical fiber; 9-blade body; 10-outer ring positioning pin hole; 11-leaf-shaped impact tube one; 12-leaf-shaped impact tube two; 13-upper edge plate; 14-lower edge plate; 15-impact cooling air hole; 16-cooling air exhaust hole; 17-front upper support; 18-rear upper support; 19-cooling air inlet hole; 20-cavity; 21-front lower support; 22-rear lower support; 23-angular positioning groove. DETAILED DESCRIPTION

[0032] In order to clearly illustrate the technical features of this solution, the present invention will be described in detail below by way of specific implementation methods and in conjunction with the accompanying drawings. In the following description, many specific details are described to facilitate a full understanding of the present application. However, the present application can also be implemented in other ways different from those described herein. Therefore, the scope of protection of the present application is not limited by the specific embodiments disclosed below. Unless otherwise defined, all technical terms used hereinafter have the same meaning as those generally understood by those skilled in the art. The technical terms used herein are only for the purpose of describing specific embodiments and are not intended to limit the scope of protection of the present invention. Unless otherwise specified, the various raw materials, reagents, instruments and equipment used in the present invention, etc. can be purchased from the market or can be prepared by existing methods.

[0033] Example 1

[0034] See also Figures 2 to 7 The present invention provides an embodiment of a gas turbine with adjustable high-pressure air cooling, comprising a boost air pump, a radiator connected to the boost air pump, an external pipeline having one end connected to the radiator, a control valve and a flow meter provided on the external pipeline, a turbine casing volute 1 connected to the other end of the external pipeline, and a first-stage turbine guide vane 2 connected to the turbine casing volute 1;

[0035] The first-stage turbine guide vane 2 includes an outer ring 3 connected to the turbine casing volute 1, an inner ring 4, and an annular turbine guide blade 5 installed between the inner ring 4 and the outer ring 3. The annular turbine guide blade 5 is surrounded by multiple groups of triple guide vanes 6 and a group of double guide vanes 7. The triple guide vanes 6 are integrally cast by three first-stage turbine guide blades 8, and the double guide vanes 7 are integrally cast by two first-stage turbine guide blades 8.

[0036] The first-stage turbine guide blade 8 comprises a blade body 9 containing an inner cavity 10, two blade-shaped impingement tubes 11 and 12 of different sizes disposed within the inner cavity 10 of the blade body 9, an upper edge plate 13 disposed above the blade body 9, and a lower edge plate 14 disposed below the blade body 9. The first-stage turbine guide blade 8 is cast from a nickel-based high-temperature alloy and coated with a cobalt-chromium-aluminum-yttrium alloy to enhance its high-temperature oxidation resistance.

[0037] The blade body 9 is formed by stretching a smooth blade profile, and the stretching direction is perpendicular to the turbine shaft, which deflects the axial airflow at the combustion chamber outlet by 60 degrees, reduces the pressure and accelerates the ejection.

[0038] The blade-shaped impact tube 11 and the blade-shaped impact tube 2 12 are both directly connected to the upper edge plate 13 and the lower edge plate 14. The blade-shaped impact tube 11 is larger than the blade-shaped impact tube 2 12. The blade-shaped impact tube 11 is arranged at the front edge of the inner cavity 10, and the blade-shaped impact tube 2 12 is arranged at the rear edge of the inner cavity 10. The front edge and wall of the blade-shaped impact tube 11 are provided with impact cooling holes 15. The impact cooling holes 15 at the front edge of the blade-shaped impact tube 11 blow towards the inner back side of the blade body 9. Before cooling, the temperature of this part is the highest. The blade body 9 is provided with double rows of cooling air exhaust holes 16. The impact cooling holes 15 and the cooling air exhaust holes 16 are staggered, and the cooling air moves forward in a zigzag manner.

[0039] A front upper support 17 and a rear upper support 18 are provided above the upper edge plate 13, connected to the inner side of the outer ring 3. A cold air inlet hole 19 is provided on the side of the rear upper support 18. An annular cavity 20 is formed between the outer ring 3 and the entire upper edge plate 13, the front upper support 17, and the rear upper support 18. The cavity 20 communicates with all of the leaf-shaped impingement tubes 11 and 12, and the cold air inlet hole 19. The upper edge plate 13 is provided with an outer ring locating pin hole 10, and the outer ring 3 is connected to the upper edge plate 13 by inserting a locating pin into the outer ring locating pin hole 10.

[0040] The lower edge plate 14 is provided with a front lower support 21 and a rear lower support 22 connected to the outer side of the inner ring 4. The lower edge plate 14 is provided with an angular positioning groove 23, and the inner ring 4 is connected to the lower edge plate 14 by inserting a square tenon into the angular positioning groove 23.

[0041] The interface between the outer ring 3 and the turbine casing volute 1 is sealed and positioned by an expansion ring, and the first-stage turbine guide blades 8 are sealed by inserting iron-based high-temperature alloy strips, which also have a vibration damping function.

[0042] There is a gap between the triple guide vane 6 and the double guide vane 7, and between the triple guide vane 6 and the triple guide vane 6, the gap is 0.2mm, and multiple gaps leave space for thermal expansion and contraction, reducing thermal stress and increasing cycle life. Since the adjustable cold air pressure is significantly higher than the gas pressure, higher requirements are placed on the sealing performance, and the number of leakage points must be reduced.

[0043] During operation, the booster pump and radiator generate high-pressure cold air. The flow of high-pressure cold air can be adjusted through the control valve. The flow meter observes the flow rate and adjusts the amount of cold air according to the power of the gas turbine at that time. After that, the high-pressure cold air enters the turbine casing volute 1 from the external pipeline, pours into the cavity 20 from the cold air inlet hole 19, enters each blade-shaped impact tube 11, enters the inner cavity 10 of each first-stage turbine guide blade 8 blade body 9, and then is ejected from the two rows of cold air discharge holes 16 on the trailing edge of each blade body 9 and merges into the mainstream turbine casing volute 1.

[0044] Example 2

[0045] See also Figures 2 to 7 The present invention provides an embodiment of a gas turbine with adjustable high-pressure air cooling, comprising a boost air pump, a radiator connected to the boost air pump, an external pipeline having one end connected to the radiator, a control valve and a flow meter provided on the external pipeline, a turbine casing volute 1 connected to the other end of the external pipeline, and a first-stage turbine guide vane 2 connected to the turbine casing volute 1;

[0046] The first-stage turbine guide vane 2 includes an outer ring 3 connected to the turbine casing volute 1, an inner ring 4, and an annular turbine guide blade 5 installed between the inner ring 4 and the outer ring 3. The annular turbine guide blade 5 is surrounded by multiple groups of triple guide vanes 6 and a group of double guide vanes 7. The triple guide vanes 6 are integrally cast by three first-stage turbine guide blades 8, and the double guide vanes 7 are integrally cast by two first-stage turbine guide blades 8.

[0047] The first-stage turbine guide blade 8 comprises a blade body 9 containing an inner cavity 10, two blade-shaped impingement tubes 11 and 12 of different sizes disposed within the inner cavity 10 of the blade body 9, an upper edge plate 13 disposed above the blade body 9, and a lower edge plate 14 disposed below the blade body 9. The first-stage turbine guide blade 8 is cast from a nickel-based high-temperature alloy and coated with a cobalt-chromium-aluminum-yttrium alloy to enhance its high-temperature oxidation resistance.

[0048] The blade body 9 is formed by stretching a smooth blade profile, and the stretching direction is perpendicular to the turbine shaft, which deflects the axial airflow at the combustion chamber outlet by 60 degrees, reduces the pressure and accelerates the ejection.

[0049] The blade-shaped impact tube 11 and the blade-shaped impact tube 2 12 are both directly connected to the upper edge plate 13 and the lower edge plate 14. The blade-shaped impact tube 11 is larger than the blade-shaped impact tube 2 12. The blade-shaped impact tube 11 is arranged at the front edge of the inner cavity 10, and the blade-shaped impact tube 2 12 is arranged at the rear edge of the inner cavity 10. The front edge and wall of the blade-shaped impact tube 11 are provided with impact cooling holes 15. The impact cooling holes 15 at the front edge of the blade-shaped impact tube 11 blow towards the inner back side of the blade body 9. Before cooling, the temperature of this part is the highest. The blade body 9 is provided with double rows of cooling air exhaust holes 16. The impact cooling holes 15 and the cooling air exhaust holes 16 are staggered, and the cooling air moves forward in a zigzag manner.

[0050] Furthermore, the leaf-shaped impact tube 1 1 and the leaf-shaped impact tube 2 12 have the same wall thickness of 0.3 mm. The thickness between the leaf-shaped impact tube 1 1 and the leaf-shaped impact tube 2 12 is about 1 mm, which improves the rigidity and natural frequency. This design can improve the qualification rate and reduce the manufacturing cost.

[0051] A front upper support 17 and a rear upper support 18 are provided above the upper edge plate 13, connected to the inner side of the outer ring 3. A cold air inlet hole 19 is provided on the side of the rear upper support 18. An annular cavity 20 is formed between the outer ring 3 and the entire upper edge plate 13, the front upper support 17, and the rear upper support 18. The cavity 20 communicates with all of the leaf-shaped impingement tubes 11 and 12, and the cold air inlet hole 19. The upper edge plate 13 is provided with an outer ring locating pin hole 10, and the outer ring 3 is connected to the upper edge plate 13 by inserting a locating pin into the outer ring locating pin hole 10.

[0052] The lower edge plate 14 is provided with a front lower support 21 and a rear lower support 22 connected to the outer side of the inner ring 4. The lower edge plate 14 is provided with an angular positioning groove 23, and the inner ring 4 is connected to the lower edge plate 14 by inserting a square tenon into the angular positioning groove 23.

[0053] The interface between the outer ring 3 and the turbine casing volute 1 is sealed and positioned by an expansion ring, and the first-stage turbine guide blades 8 are sealed by inserting iron-based high-temperature alloy strips, which also have a vibration damping function.

[0054] There is a gap between the triple guide vane 6 and the double guide vane 7, and between the triple guide vane 6 and the triple guide vane 6, the gap is 0.2mm, and multiple gaps leave space for thermal expansion and contraction, reducing thermal stress and increasing cycle life. Since the adjustable cold air pressure is significantly higher than the gas pressure, higher requirements are placed on the sealing performance, and the number of leakage points must be reduced.

[0055] During operation, the booster pump and radiator generate high-pressure cold air. The flow of high-pressure cold air can be adjusted through the control valve. The flow meter observes the flow rate and adjusts the amount of cold air according to the power of the gas turbine at that time. After that, the high-pressure cold air enters the turbine casing volute 1 from the external pipeline, pours into the cavity 20 from the cold air inlet hole 19, enters each blade-shaped impact tube 11, enters the inner cavity 10 of each first-stage turbine guide blade 8 blade body 9, and then is ejected from the two rows of cold air discharge holes 16 on the trailing edge of each blade body 9 and merges into the mainstream turbine casing volute 1.

[0056] Example 3

[0057] See also Figures 2 to 7 The present invention provides an embodiment of a gas turbine with adjustable high-pressure air cooling, comprising a boost air pump, a radiator connected to the boost air pump, an external pipeline having one end connected to the radiator, a control valve and a flow meter provided on the external pipeline, a turbine casing volute 1 connected to the other end of the external pipeline, and a first-stage turbine guide vane 2 connected to the turbine casing volute 1;

[0058] The first-stage turbine guide vane 2 includes an outer ring 3 connected to the turbine casing volute 1, an inner ring 4, and an annular turbine guide blade 5 installed between the inner ring 4 and the outer ring 3. The annular turbine guide blade 5 is surrounded by multiple groups of triple guide vanes 6 and a group of double guide vanes 7. The triple guide vanes 6 are integrally cast by three first-stage turbine guide blades 8, and the double guide vanes 7 are integrally cast by two first-stage turbine guide blades 8.

[0059] The first-stage turbine guide blade 8 comprises a blade body 9 containing an inner cavity 10, two blade-shaped impingement tubes 11 and 12 of different sizes disposed within the inner cavity 10 of the blade body 9, an upper edge plate 13 disposed above the blade body 9, and a lower edge plate 14 disposed below the blade body 9. The first-stage turbine guide blade 8 is cast from a nickel-based high-temperature alloy and coated with a cobalt-chromium-aluminum-yttrium alloy to enhance its high-temperature oxidation resistance.

[0060] The blade body 9 is formed by stretching a smooth blade profile, and the stretching direction is perpendicular to the turbine shaft, which deflects the axial airflow at the combustion chamber outlet by 60 degrees, reduces the pressure and accelerates the ejection.

[0061] The blade-shaped impact tube 11 and the blade-shaped impact tube 2 12 are both directly connected to the upper edge plate 13 and the lower edge plate 14. The blade-shaped impact tube 11 is larger than the blade-shaped impact tube 2 12. The blade-shaped impact tube 11 is arranged at the front edge of the inner cavity 10, and the blade-shaped impact tube 2 12 is arranged at the rear edge of the inner cavity 10. The front edge and wall of the blade-shaped impact tube 11 are provided with impact cooling holes 15. The impact cooling holes 15 at the front edge of the blade-shaped impact tube 11 blow towards the inner back side of the blade body 9. Before cooling, the temperature of this part is the highest. The blade body 9 is provided with double rows of cooling air exhaust holes 16. The impact cooling holes 15 and the cooling air exhaust holes 16 are staggered, and the cooling air moves forward in a zigzag manner.

[0062] A front upper support 17 and a rear upper support 18 are provided above the upper edge plate 13, connected to the inner side of the outer ring 3. A cold air inlet hole 19 is provided on the side of the rear upper support 18. An annular cavity 20 is formed between the outer ring 3 and the entire upper edge plate 13, the front upper support 17, and the rear upper support 18. The cavity 20 communicates with all of the leaf-shaped impingement tubes 11 and 12, and the cold air inlet hole 19. The upper edge plate 13 is provided with an outer ring locating pin hole 10, and the outer ring 3 is connected to the upper edge plate 13 by inserting a locating pin into the outer ring locating pin hole 10.

[0063] The lower edge plate 14 is provided with a front lower support 21 and a rear lower support 22 connected to the outer side of the inner ring 4. The lower edge plate 14 is provided with an angular positioning groove 23, and the inner ring 4 is connected to the lower edge plate 14 by inserting a square tenon into the angular positioning groove 23.

[0064] The interface between the outer ring 3 and the turbine casing volute 1 is sealed and positioned by an expansion ring, and the first-stage turbine guide blades 8 are sealed by inserting iron-based high-temperature alloy strips, which also have a vibration damping function.

[0065] There is a gap between the triple guide vane 6 and the double guide vane 7, and between the triple guide vane 6 and the triple guide vane 6, the gap is 0.2mm, and multiple gaps leave space for thermal expansion and contraction, reducing thermal stress and increasing cycle life. Since the adjustable cold air pressure is significantly higher than the gas pressure, higher requirements are placed on the sealing performance, and the number of leakage points must be reduced.

[0066] Furthermore, an optical fiber 8 is provided on the outer wall at the leading edge of the impact tube. The light emitted by the red-hot guide vane is transmitted to the outside of the engine through the optical fiber 8. The wall temperature is measured by a radiation temperature measuring instrument at the output end of the optical fiber 8. The low-temperature impact tube serves as a mounting bracket for the optical fiber 8 sensor to protect the optical fiber 8. This is a non-destructive temperature measurement method. Ordinary thermocouple sensors do not have a sufficient service life and hinder the airflow.

[0067] During operation, the booster pump and radiator generate high-pressure cold air. The flow of high-pressure cold air can be adjusted through the control valve. The flow meter observes the flow rate and adjusts the amount of cold air according to the power of the gas turbine at that time. After that, the high-pressure cold air enters the turbine casing volute 1 from the external pipeline, pours into the cavity 20 from the cold air inlet hole 19, enters each blade-shaped impact tube 11, enters the inner cavity 10 of each first-stage turbine guide blade 8 blade body 9, and then is ejected from the two rows of cold air discharge holes 16 on the trailing edge of each blade body 9 and merges into the mainstream turbine casing volute 1.

[0068] It will be apparent to those skilled in the art that the present invention is not limited to the details of the exemplary embodiments described above and that the invention can be embodied in other specific forms without departing from the spirit or essential characteristics of the invention. Therefore, the embodiments should be considered in all respects as illustrative and non-restrictive, and the scope of the invention is defined by the appended claims rather than the foregoing description. It is intended that all variations within the meaning and range of equivalents of the claims be embraced herein, and any reference signs in the claims should not be construed as limiting the claims to which they relate.

Claims

1. A gas turbine with adjustable high-pressure air cooling, characterized in that: It includes a booster air pump, a radiator connected to the booster air pump, an external pipeline connected to the radiator at one end, a control valve and a flow meter provided on the external pipeline, a turbine casing volute connected to the other end of the external pipeline, and a first-stage turbine guide connected to the turbine casing volute; The first-stage turbine guide vane includes an outer ring connected to the turbine casing volute, an inner ring, and an annular turbine guide blade installed between the inner ring and the outer ring. The annular turbine guide blade is formed by multiple sets of triple guide vanes and one set of double guide vanes. The triple guide vanes are integrally cast from three first-stage turbine guide blades. The first-stage turbine guide blades are cast from a nickel-based high-temperature alloy and coated with a cobalt-chromium-aluminum-yttrium alloy. The double guide vanes are integrally cast from two first-stage turbine guide blades. The first-stage turbine guide blade includes a blade body containing an inner cavity, two blade-shaped impact tubes of different sizes, a first blade-shaped impact tube and a second blade-shaped impact tube, an upper edge plate above the blade body, and a lower edge plate below the blade body. The blade-shaped impact tubes are directly connected to the upper edge plate and the lower edge plate. The lobe-shaped impact tube 1 is larger than the lobe-shaped impact tube 2. The lobe-shaped impact tube 1 and the lobe-shaped impact tube 2 have the same wall thickness of 0.3 mm. The thickness between the lobe-shaped impact tube 1 and the lobe-shaped impact tube 2 is 1±0.5 mm. The lobe-shaped impact tube 1 is arranged at the front edge of the inner cavity, and the lobe-shaped impact tube 2 is arranged at the rear edge of the inner cavity. The front edge and wall of the lobe-shaped impact tube 1 are provided with impact cooling holes, and the rear edges of the lobe-shaped impact tube 1 and the lobe-shaped impact tube 2 are provided with cooling air discharge holes. A front upper support and a rear upper support connected to the inner side of the outer ring are provided above the upper edge plate, and a cold air inlet hole is provided on the side of the rear upper support. An annular cavity is formed between the outer ring and all the upper edge plates, the front upper support and the rear upper support. The cavity is communicated with all the leaf-shaped impact tubes 1, 2 and the cold air inlet hole.

2. The adjustable high-pressure air-cooled gas turbine according to claim 1, characterized in that: A front lower support and a rear lower support connected to the outer side of the inner ring are provided at the front and rear sides below the lower edge plate.

3. The adjustable high-pressure air-cooled gas turbine according to claim 1, characterized in that: The blade body is formed by stretching a smooth blade profile.

4. The adjustable high-pressure air-cooled gas turbine according to claim 1, characterized in that: The lower edge plate is provided with an angular positioning groove, and the inner ring is connected to the lower edge plate by inserting the square tenon into the angular positioning groove.

5. The adjustable high-pressure air-cooled gas turbine according to claim 1, characterized in that: The upper edge plate is provided with an outer ring positioning pin hole, and the outer ring is connected to the upper edge plate by inserting a positioning pin into the outer ring positioning pin hole.

6. The adjustable high-pressure air-cooled gas turbine according to claim 1, characterized in that: The interface between the outer ring and the turbine casing volute is sealed and positioned by an expansion ring, and iron-based high-temperature alloy strips are inserted between the first-stage turbine guide blades for sealing.

7. The adjustable high-pressure air-cooled gas turbine according to claim 1, characterized in that: There is a gap between the triple guide vane and the double guide vane, and between the triple guide vane and the triple guide vane, and the gap is 0.2±0.05mm.

8. The adjustable high-pressure air-cooled gas turbine according to claim 1, characterized in that: An optical fiber is provided on the outer front wall of the first or second blade-shaped impact tube. The light emitted by the red-hot guide vane is transmitted to the outside of the engine through the optical fiber. The wall temperature is measured by a radiation temperature measuring instrument at the output end of the optical fiber.

Images