JP7704655B2 - Turbine case and gas turbine - Google Patents

Description

The present invention relates to a turbine case that houses various parts, including the turbine, assembled in a gas turbine that rotates the turbine using combustion gas generated in a combustor, and in particular to a turbine case that suppresses thermal deformation caused by heat and pressure during operation, and a gas turbine that uses the turbine case.

Patent Document 1 discloses an invention for a joint structure in which a reinforcing rib is used to weld and reinforce the space between a tubular structural member and a base plate such as a base plate or a joining flange. The reinforcing rib in this invention has a U-shaped structure that wraps around the outer circumference of the tubular structural member, and this rib structure is said to be able to simultaneously increase the fatigue strength of both the outer weld toe on the bent side of the rib and the fatigue strength of the open end of the rib close to the base plate or joining flange, thereby further improving strength and fatigue performance.

Patent Document 2 discloses an invention related to cooling the low-pressure turbine casing in a gas turbine engine. The purpose of this invention is not to suppress deformation of the casing, but to cool the long and slender casing that is unique to jet engines by flowing cooling air between the shroud and the casing.

Patent document 3 discloses an invention related to the rib structure of an exhaust manifold for an internal combustion engine. The reinforcing rib of this invention is provided by connecting two flanges at both ends of the manifold, which is said to suppress thermal deformation of the manifold.

JP 2006-2464 A JP 2004-60656 A JP 2007-2726 A

A gas turbine is an internal combustion engine that supplies high-temperature gas generated in a combustor to a turbine installed in a turbine case, and rotates the turbine and the rotor to which it is attached to obtain rotational kinetic energy. The present inventors are engaged in research and development of gas turbines, but in gas turbines that the present inventors knew before the present invention, the turbine case could be deformed by the heat and pressure of the high-temperature combustion gas that is the power source. If the turbine case is deformed, the gap (tip clearance) between the turbine rotating at high speed and other components such as the shroud case changes, and there is a possibility that the components will be damaged or the engine performance will deteriorate. For this reason, the present inventors have made it a goal of research and development in recent years to suppress deformation of the turbine case due to heat and pressure. The present invention has been made to solve such problems, and aims to provide a turbine case that does not undergo thermal deformation due to the heat and pressure when the gas turbine is driven, and a gas turbine using the same.

The turbine case according to claim 1 comprises:
A turbine case having a flange to which other components are connected,
The flange is surrounded when viewed in the axial direction of the central axis of the flange , and is provided with continuous ribs arranged in a pattern connecting the vertices of a circumferential outer wall curved surface that bulges from the outer periphery of the flange seating surface along the direction of the central axis in a direction in which combustion gas is discharged to the outside .

The turbine case according to claim 2 is the turbine case according to claim 1,
A plurality of radial ribs are provided around the flange at equal intervals in the circumferential direction,
The continuous rib and the radial rib are connected to each other at right angles at a connection portion,
The connecting portion between the continuous rib and the radial ribs is characterized in that it is formed of a curved surface.

The turbine case according to claim 3 is the turbine case according to claim 2,
The thickness of the continuous rib is 25 mm or more or is greater than the thickness of the radial ribs.

A turbine case according to a fourth aspect of the present invention is a turbine case according to any one of the first to third aspects,
The turbine case has a first flange to which a shroud case is connected, the first flange having an opening at one end surface, and a second flange to which a combustor is connected, the second flange having an opening at a peripheral wall, the turbine case being cylindrical and asymmetric with respect to a central axis of the first flange ,
The continuous rib is characterized in that it surrounds the first flange when viewed at least in the axial direction of the central axis, and is arranged in a pattern connecting the vertices of a circumferential outer wall curved surface that bulges from the outer periphery of the seating surface of the first flange along the direction of the central axis in a direction in which combustion gas is discharged to the outside .

The gas turbine according to claim 5 further comprises:
a turbine case having a cylindrical shape with a first flange opening on one end surface and a second flange opening on a peripheral wall, the turbine case being asymmetric with respect to a central axis of the first flange ;
a shroud case housed in the turbine case and connected to the first flange;
a combustor connected to the second flange;
a turbine housed in the turbine case and driven by combustion gas generated in the combustor and supplied through the shroud case;
The turbine case is characterized in that it has a continuous rib that surrounds the first flange when viewed in the axial direction of the central axis and is provided on the turbine case in a pattern connecting vertices of a circumferential outer wall curved surface that bulges from the outer periphery of the seating surface of the first flange along the direction of the central axis in a direction in which combustion gas is discharged to the outside.

According to the turbine case described in claim 1, since other parts are connected to the flange, which is greatly affected by heat and pressure, continuous ribs are provided around the flange, so that thermal deformation of the turbine case can be suppressed at low cost compared to conventional measures such as changing to a more expensive material with higher durability or cooling the parts connected to the turbine case, such as the shroud case or turbine scroll.

In addition, according to the turbine case described in claim 1, the surface area of the turbine case that is cooled by air at room temperature is increased by providing continuous ribs, and the continuous ribs function as cooling fins, improving the cooling performance of the turbine case by outside air.

According to the turbine case described in claim 2, at the connection where the continuous rib and the radial rib intersect at right angles, the area between the continuous rib and the radial rib is formed with a curved surface, so that the stress generated in the turbine case can be suppressed while the rigidity can be increased, and the thermal expansion concentrated on the surface of the turbine case can be suppressed.

According to the turbine case described in claim 3, by optimizing the thickness of the continuous rib, it is possible to prevent permanent deformation from occurring in turbine cases made of inexpensive materials that have traditionally been used.

The turbine case described in claim 4 and the turbine case used in the turbine described in claim 5 have a cylindrical shape with a first flange to which the shroud case is connected that opens on one end face, but the second flange to which the combustor is connected opens on the peripheral wall, resulting in an asymmetric structure with respect to the center line of the cylinder. Therefore, if no measures are taken, the high heat and pressure generated in the turbine case when the gas turbine is running may cause stress concentration in specific locations in the asymmetric turbine case, resulting in deformation of the turbine case. However, since the turbine case is provided with a continuous rib surrounding at least the first flange, thermal deformation of the turbine case can be suppressed at low cost, and the cooling performance of the turbine case by the outside air is also improved, resulting in a gas turbine with a highly durable turbine case.

FIG. 1 is a schematic perspective view showing a gas turbine according to an embodiment, with a part of the gas turbine cut away to illustrate its structure. 2 is an enlarged cross-sectional view showing a connecting portion between a first flange of a turbine case, a shroud case, and an exhaust diffuser in a gas turbine of an embodiment. FIG. FIG. 1A is a perspective view of a turbine case in its current shape, and FIG. 1B is a diagram showing the results of a simulation using the finite element method of the stress applied to the turbine case in its current shape when the gas turbine is operated under specified conditions. FIG. 4A is a perspective view of a turbine case according to an embodiment of the present invention, and FIG. 4B is a diagram showing the results of calculations, by simulation using the finite element method, of stresses applied to the turbine case of the embodiment when a gas turbine is driven under specified conditions. FIG. 1A is a table showing the results of calculations performed by a simulation using the finite element method regarding the relationship between the stress ratio and the internal pressure ratio for a turbine case of an existing shape and a turbine case according to an embodiment when the gas turbine is operated under specified conditions, and FIG. 1B is a graph showing the calculation results shown in the table of FIG. 1A. 10 is a graph showing the results of a calculation performed by simulation using a finite element method of an axial displacement ratio at a peripheral portion of a first flange to which a shroud case and the like are attached when a gas turbine is operated under specified conditions for a turbine case of a current shape and a turbine case according to an embodiment.

In order to explain the characteristic configuration of the gas turbine of the embodiment and the resulting effects that cannot be obtained with gas turbines of the current configuration, the structure common to gas turbines of the current configuration and the gas turbine of the embodiment will first be explained with reference to Figures 1 to 4. Note that the gas turbine of the current configuration does not mean something that was publicly known before the present invention, but rather means prior art that was known to the inventors of the present application before the present invention.

As shown in FIG. 1, the gas turbine 1 has an outer casing consisting of a roughly cylindrical turbine case 2 and a roughly cylindrical compressor case 3 coaxially connected to the turbine case 2. A common rotor 4 is rotatably supported in the center inside this outer casing via a number of bearings 5. A compressor 6 housed in the compressor case 3 is attached to the rotor 4, and a turbine 7 housed in the turbine case 2 is also attached. In addition, a duct 8 for taking in intake air is attached to the outside of the compressor case 3.

As shown in Figures 1, 3(a) and 4(a), the turbine case 2 is a substantially cylindrical member as described above, with a first flange F1 for mounting various members described later opening at one end face and a second flange F2 opening at the peripheral wall. Therefore, although the turbine case 2 is substantially cylindrical, it has an asymmetric structure with respect to its center line (center axis Z).

As shown in Figures 1 and 2, a shroud case 9 and an exhaust diffuser 10 are attached to the first flange F1 of the turbine case 2. As shown in an enlarged view in Figure 2 in particular, the first flange F1 of the turbine case 2, the mounting portion of the shroud case 9, and the mounting portion of the exhaust diffuser 10 are overlapped and fastened together with a common bolt 11. The shroud case 9 and the exhaust diffuser 10 form an exhaust flow passage 12 that exhausts the combustion gas after driving the turbine 7.

As shown in FIG. 1, an external cylinder 13 is attached to the second flange F2 of the turbine case 2. Inside the external cylinder 13, a can-type combustor 14 is provided at a predetermined distance from the inner wall of the external cylinder 13, supplying combustion gas for driving the turbine 7. The can-type combustor 14 has multiple through holes for taking in air. In addition, although not shown in detail, the can-type combustor 14 is connected to a fuel pipe for supplying fuel, and is also provided with an ignition plug.

As shown in FIG. 1, inside the turbine case 2, a turbine scroll 15 is provided that constitutes a combustion gas flow path that guides the combustion gas sent from the can-type combustor 14 to the turbine 7. That is, one end of the turbine scroll 15 is connected to the open lower end of the can-type combustor 14, and the other end opens toward the turbine 7. In addition, an air flow path 16 that supplies compressed air to the can-type combustor 14 is formed between the inner wall of the turbine case 2 and the outer wall of the turbine scroll 15. That is, the air flow path 16 connects the compressor case 3, in which the compressor 6 is provided, to the space between the outer cylinder 13 and the can-type combustor 14.

According to the above configuration, when compressed air compressed by the compressor 6 is supplied to the can-type combustor 14 from the air flow path 16 and fuel is supplied to the can-type combustor 14 from the fuel pipe, an air-fuel mixture is generated inside the can-type combustor 14, which is ignited by the spark plug and combusted to generate combustion gas. The combustion gas generated in the can-type combustor 14 is guided to the turbine 7 through the combustion gas flow path, which is the internal space of the turbine scroll 15, and drives the turbine 7 to rotate the rotor 4, and then discharged to the outside from the exhaust flow path 12 formed by the shroud case 9 and the exhaust diffuser 10.

Next, a description will be given of problems caused by the structure of the current shape of the gas turbine 1. Note that the turbine case of the embodiment will be denoted by reference symbol 2a, the gas turbine of the current shape will be denoted by reference symbol 2b, and the reference symbol 2 will be used when no particular distinction is made.
As described above, the turbine case 2 is a complex structure that serves multiple roles, such as forming an air flow path 16 between itself and the shroud case 9, acting as a heat insulator to block heat from the turbine scroll 15 through which 1000° C. combustion gas flows in an internal combustion gas flow path, housing the rotor 4 and the turbine 7, and holding the can-type combustor 14 and shroud case 9 in predetermined positions.

In such a complex structure, if the turbine case 2 is deformed by heat or internal pressure, the gap (tip clearance) between the outer wall of the shroud case 9 and the tip of the turbine 7 changes. If this gap narrows, the outer wall of the shroud case 9 and the turbine 7 may come into contact and break, and if it widens, the aerodynamic performance of the turbine 7 may be significantly reduced and the engine performance may deteriorate. In addition, the deformation of the turbine case 2 may change the holding position of the bearing 5 of the rotor 4, causing misalignment of the rotor 4 and increasing vibration. Furthermore, the deformation of the turbine case 2 may significantly reduce the sealing performance of the connection surface with the shroud case 9 and the seat surface of the first flange F1 and the second flange F2, which are the connection surfaces of the can-type combustor 14, as shown in Figures 3(a) and 4(a), and may cause compressed air to leak outside the turbine case 2, making it difficult to operate the engine. Furthermore, if the turbine case 2 deforms, the compressor case 3 connected to it may also deform.

The main factor that deforms the shape of the turbine case 2 is heat transfer from high-temperature fluid. While the outer wall of the turbine case 2 is always cooled by room temperature air, compressed air, which is a high-temperature fluid close to the discharge air temperature of the compressor 6, which is over 300°C, flows along its inner wall. In addition, combustion gas at 1000°C flows through the turbine scroll 15 installed inside the turbine case 2, and the turbine case 2 is subjected to strong radiation. Furthermore, the first flange F1 and the second flange F2 are respectively connected to the shroud case 9 that defines the exhaust flow path 12 and the can-type combustor 14, which becomes hot due to combustion, and a temperature of nearly 700°C is transmitted to the turbine case 2 by direct thermal conduction.

As described above, the turbine case 2 is exposed to high temperatures at each part and is in a harsh environment with large temperature differences between each part. However, as described above, the turbine case 2 is a structure in which many parts are connected and assembled into one unit, so it is difficult to expand thermally, and stress occurs in a non-uniform manner at each part. In particular, as shown in Figures 3(a) and 4(a), the part between the seat surface of the first flange F1 to which the shroud case 9 is connected and the outer wall side surface 17 that contacts the circumferential surface of the cylindrical turbine case 2b, that is, the outer wall curved surface 18, which is a circumferential part that bulges and curves in a semi-cylindrical shape from the outer periphery of the seat surface of the first flange F1 in the direction of the central axis Z of the cylindrical turbine case 2b, is cooled on the outside by air at room temperature, but is easily subjected to temperature differences and deformation due to heat because it receives heat directly from high-temperature parts inside. The same is true for the vicinity of the seat surface of the second flange F2 to which the can-type combustor 14 is connected.

As shown in FIG. 3(a), in order to suppress this thermal deformation, the turbine case 2b in its current shape has multiple radial ribs 20 continuously provided from the outer periphery of the first flange F1 through the outer wall curved surface 18 to the outer wall side surface 17. The radial ribs 20 extend in the radial direction of the circular first flange F1 on the outer wall curved surface 18, and are provided so as to surround the first flange F1 at equal intervals in the circumferential direction of the circular first flange F1. On the outer wall side surface 17, the radial ribs 20 are parallel to the central axis Z of the circular first flange F1. In addition, the ends of the radial ribs 20 on the outer wall side surface 17 are continuous with an inclined surface having a certain angle with respect to the outer wall side surface 17 of the turbine case 2b.

However, in the turbine case 2b of the current shape, the radial ribs 20 cannot suppress the thermal deformation, and large stress occurs locally. Figure 3(b) is a diagram showing the stress applied to the turbine case 2b of the current shape when the gas turbine 1 having the turbine case 2b of the current shape shown in Figure 3(a) is driven under a specified condition, calculated by a simulation using the finite element method, and the result is shown in a 3D image with a shading graphic. As shown in Figure 3(b), a part of the outer wall curved surface 18 close to the second flange F2 to which the can-type combustor 14 is attached, that is, a part of the outer wall curved surface 18 sandwiched between the first flange F1 and the second flange F2 (referred to as the first concentration point P1), is high in temperature and has a significant temperature difference, so high stress is generated as shown in dark gray in the figure. Stress is also concentrated at a position (referred to as a second concentration point P2) on the opposite side of the first concentration point P1 on the outer wall curved surface 18, rotated 180 degrees in the circumferential direction of the circular first flange F1, where high stress occurs. At these locations, the surface of the turbine case 2b is thermally deformed due to the concentration of stress.

When the gas turbine 1 is repeatedly started and stopped, the turbine case 2 is repeatedly heated and cooled, which causes fatigue to accumulate and raises concerns about a shortened fatigue life. Although no cracks have occurred in the current shape of the turbine case 2b, the inventors of the present application have considered that some kind of thermal deformation countermeasure is necessary for the turbine case 2, since an increase in the pressure ratio is expected in future gas turbine developments. Furthermore, the current shape of the turbine case 2b described above is a type having a can-type combustor 14, and the turbine case 2a of the present invention, which improves on this, is also considered to be suitably applied to a type having a can-type combustor 14, but this type of turbine case 2 is different from the cylindrical turbine case that is axially symmetric, such as an annular combustor typified by a jet engine. That is, the turbine case 2 used in the gas turbine 1 with a can-type combustor 14 has an asymmetric structure in which a cylinder with a smaller radius than the cylindrical main body is embedded in the side surface of the cylindrical main body in order to connect the can-type combustor 14. Therefore, it is difficult to design the arrangement of the radial ribs 20 described above, and as explained with reference to Figure 3 (b), it cannot be said that the thermal deformation of the turbine case 2b is necessarily effectively suppressed.

The above describes the structure common to the current gas turbine 2b and the gas turbine 2a of the embodiment, and the problems with the radial ribs 20 in the current gas turbine 2b. Based on this, the structure of the continuous rib, which is a feature of the present invention, and its effect of solving the problems with the current design will be described below with reference to Figures 4 to 6.

As shown in FIG. 4(a), in the turbine case 2a of the embodiment, in addition to the radial ribs 20 of the current shape, a circular rib 30 (hereinafter also referred to as a ring rib 30) is provided as a continuous rib that intersects with the radial ribs 20. The circular rib 30 is provided in a closed circular pattern that connects the vertices of the outer wall curved surface 18 surrounding the first flange F1, and intersects with and is connected to all of the multiple radial ribs 20. In other words, the "continuous" of the continuous rib refers to a state in which a line is closed and forms a loop regardless of its shape. For example, it may be circular like the circular rib 30, or it may be elliptical or rectangular depending on the outer shape of the flange.

As shown in FIG. 4(a), the height from the surface of the turbine case 2a (i.e., the surface of the outer wall curved surface 18) to the surface of the ring rib 30 is the same as that of the radial rib 20, and the ring rib 30 has a flat surface on its surface that is parallel to the surface of the turbine case 2a and has a constant thickness T. The surface of the ring rib 30 is in a ring shape with an inner circumference of a constant radius and an outer circumference of a constant radius plus thickness T. The radial rib 20 and the ring rib 30 are connected at right angles on the outer wall curved surface 18, and the connection is composed of a curved surface with a constant radius R. At the connection, the circular rib 30 and the radial rib 20 are connected by a curved surface, so that when viewed from a line of sight parallel to the axial direction Z of the first flange F1, the connection has a shape similar to a diamond or star.

Figure 4(b) shows the stress applied to the turbine case 2a of the embodiment when the gas turbine 1 is driven under a predetermined condition, calculated by a simulation using the finite element method (FEM), and the result is shown in a 3D image with a shading graphic. As previously described with reference to Figure 3(b), in the turbine case 2b of the current shape, stress was particularly concentrated at the first concentration point P1, where the outer wall curved surface 18 of the turbine case 2 is the shortest due to the presence of the second flange F2, and at the second concentration point P2, which is 180° opposite to the first concentration point P1. On the other hand, as shown in Figure 4(b), in the turbine case 2a of the embodiment having the ring rib 30 in addition to the radial ribs 20, the stress at the position corresponding to the first concentration point P1, where the compressive stress was the highest, can be reduced by about 20% compared to the turbine case 2b of the current shape. By adding the ring rib 30, the stress caused by heat and internal pressure is suppressed while rigidity is increased, which makes it possible to suppress the thermal expansion that was concentrated in specific areas on the surface of the turbine case 2b.

In addition, in the turbine case 2a of the embodiment, ring ribs 30 are added in addition to the radial ribs 20, so the surface area of the outer wall curved surface 18 of the turbine case 2a, which is particularly exposed to high heat and has a complex heat distribution, is increased compared to the current shape of the turbine case 2b, and heat transfer from the surface of the turbine case 2a to the air is promoted, thereby improving the effect of suppressing thermal expansion of the turbine case 2a.

With reference to FIG. 5, the stress generated in the turbine case 2b of the current shape and the stress generated in the turbine case 2a of the embodiment during operation of the gas turbine 1 will be compared.
5 shows the results of a simulation using an FEM technique to calculate stresses generated in the turbine cases 2a and 2b under specified operating conditions for the turbine case 2b of the current shape and two types of turbine cases 2a having different thicknesses T of the ring rib 30, and shows the results as a relationship between the internal pressure ratio and the stress ratio. Diagram (a) is a table showing the numerical values of the calculation results, and diagram (b) is a graph showing the calculation results. Note that the stress ratio here is the ratio of compressive stress to tensile stress generated in the turbine cases 2a and 2b, and the internal pressure ratio is the ratio of the pressure applied inside the position where compressive stress occurs to the pressure applied inside the position where tensile-compressive stress occurs.

Here, the thickness of the radial ribs 20 of the turbine case 2b in its current shape is 10 mm. The two types of turbine case 2a in the embodiment are a turbine case 2a (ring rib 30R20T20) in which, in addition to the radial ribs 20 of the turbine case 2b in its current shape, a ring rib 30 with a curved surface radius R of 20 mm and a thickness T of 20 mm at the connection portion is provided, and a turbine case 2a (ring rib 30R20T25) in which, in addition to the radial ribs 20 of the turbine case 2b in its current shape, a ring rib 30 with a curved surface radius R of 20 mm and a thickness T of 25 mm at the connection portion is provided.

FCD (spheroidal graphite cast iron) has traditionally been known as a common and inexpensive material for the turbine case 2, and the limit line at which the stress ratio of this FCD becomes 1 regardless of the internal pressure ratio even at a specified high temperature is shown along with the calculation results in Figure 5. If the stress ratio of the generated stress is lower than this limit line, then even if the turbine case 2 is made of FCD, it will be possible to avoid thermal deformation that generates stress exceeding the 0.2% yield strength and leaves permanent strain.

5(a) and (b), for example, when looking at the stress ratio at an internal pressure ratio of 1.0, the current shape (shown by the solid line in sub-diagram (b)) has a stress ratio of 1.126, which exceeds the limit line of 1.0, and a stress ratio exceeding the 0.2% yield strength will cause thermal deformation that leaves permanent strain. In contrast, the turbine case 2a of the embodiment (ring rib 30R20T20, shown by the dashed line in sub-diagram (b)) has a stress ratio of 1.046, which is slightly above the limit line of 0.2% yield strength, but it is possible to avoid thermal deformation that leaves permanent strain, and even if permanent strain remains, it will be very small. Furthermore, in the embodiment of the turbine case 2a (ring rib 30R20T25, shown by the three-dot chain line in FIG. (b)), the stress ratio is 0.950, which is below the limit line of the 0.2% yield strength, so thermal deformation that leaves permanent strain does not occur.

In this way, in the turbine case 2a of the embodiment, by setting the thickness T of the continuous rib to an optimal value such as a value (T20) at least larger than the radial rib 20 of the current shape or 25 mm or more, it is possible to substantially prevent permanent deformation due to heat or pressure even if the turbine case 2a is constructed from an inexpensive material such as FCD, which has been used conventionally.

With reference to FIG. 6, the displacement occurring in the first flange F1 of the turbine case 2b of the current shape during operation of the gas turbine 1 will be compared with the displacement occurring in the first flange F1 of the turbine case 2a of the embodiment.
6 shows the axial displacement ratio of the peripheral portion of the first flange F1 calculated by a simulation using the finite element method for the turbine case 2b of the current shape and the turbine case 2a of the embodiment when the gas turbine 1 is driven under predetermined conditions, and is shown for each phase angle (°) of the peripheral portion of the first flange F1. In Fig. 3(a) and Fig. 4(a), the axial direction serving as the reference for expressing the displacement ratio is the central axis Z of the circular first flange F1, and the phase angle (°) of the peripheral portion of the first flange F1 means the circumferential angle with TOP at 0°, RIGHT at 45°, BOTTOM at 180°, and LEFT at 270°.

As shown in Figure 6, when the gas turbine 1 is in operation, the axial displacement ratio (shown by the dashed line) that occurs in the first flange F1 of the turbine case 2b in its current shape exceeds 0.9 for all phase angles, and in particular, at the position between LEFT (270°) and TOP (0°), i.e., the first concentration point P1 shown in Figure 3(a), it shows a value close to 1.1. This indicates that, as mentioned above, large local deformation occurs at the first concentration point P1 due to high heat.

In contrast, as shown in FIG. 6, the axial displacement ratio (shown by the solid line) generated in the first flange F1 of the turbine case 2a of the embodiment during operation of the gas turbine 1 is below 0.7 for all phase angles and is approximately constant, with no situation in which the axial displacement ratio is exceptionally large at any particular phase angle. This indicates that, according to the embodiment, the deformation in the direction of the central axis Z caused by heat and internal pressure around the first flange F1 is smaller than that of the turbine case 2b in its current shape, and is averaged in the circumferential direction of the first flange F1.

In the embodiment described above, the ring rib 30 is provided in addition to the radial ribs 20, but the effect of suppressing thermal deformation can also be obtained when only the ring rib 30 is provided without providing the radial ribs 20. Therefore, in the embodiment, the ring rib 30 is provided on the first flange F1, but the ring rib 30 may also be provided on the circumferential surface around the second flange F2.

According to the turbine case 2a of this embodiment, since the ring rib 30 is added as a reinforcing rib as described above, it is possible to suppress thermal deformation of the turbine case 2a caused by stress generated by heat and pressure, thereby obtaining the following effects.
(1) The engine performance of the gas turbine 1 can be stably maintained.
(2) Since thermal deformation of the turbine case 2a can be suppressed even with the inexpensive materials currently used, there is no need to use particularly high-grade heat-resistant materials.
(3) Since thermal deformation of the bearing surfaces of the flanges F1, F2 that hold the shroud case 9 can be suppressed, the holding force of the shroud case 9 is improved.
(4) Since the position of the bearings 5 that hold the rotor 4 inside the turbine case 2a and the compressor case 3 can be prevented from changing, vibration of the turbine case 2a and the compressor case 3 can be suppressed.
(5) Since the thermal deformation of the turbine case 2a can be prevented, there is no need to worry about blocking the heat from the turbine 7.
(6) The ring rib 30 has a relatively simple structure, and measures against thermal deformation can be taken at low cost.
(7) The pressure resistance of the turbine case 2a is improved, thereby extending the life of the turbine case 2a.
(8) By keeping the tip clearance of the turbine 7 constant, the performance of the turbine 7 can be maintained for a long period of time.
(9) Contact between the turbine 7 and the shroud case 9 can be avoided.
(10) The tip clearance during operation can be kept constant, and the difference in tip clearance before and after operation is reduced. This makes it easier to manage performance during operation and when reassembling after disassembly.
(11) The ring ribs 30 function as cooling fins, improving the cooling performance using outside air.

The turbine case 2a of this embodiment can be used not only for gas turbine engines and industrial turbines, including those for marine, land and aviation use, but also for turbine cases of superchargers and turbochargers, and the housings of pressure vessels that also serve to support piping, such as boilers, and can inexpensively suppress thermal deformation of the turbine case or housing in these various applications.

Reference Signs List 1 gas turbine 2 turbine case 2a turbine case according to an embodiment 2b current turbine case 7 turbine 9 shroud case 14 can-type combustor 20 radial rib 30 circular rib (ring rib) as continuous rib
F1: First flange F2: Second flange T: Thickness of circular rib

Claims (5)

A turbine case having a flange to which other components are connected,
A turbine case comprising: a continuous rib that surrounds the flange when viewed in the axial direction of the central axis of the flange , and is provided in a pattern connecting vertices of a circumferential outer wall curved surface that bulges from the outer periphery of the seating surface of the flange along the direction of the central axis in a direction in which combustion gas is discharged to the outside. A plurality of radial ribs are provided around the flange at equal intervals in the circumferential direction,
The continuous rib and the radial rib are connected to each other at right angles at a connection portion,
2. The turbine case according to claim 1, wherein a curved surface is formed between the continuous rib and the radial rib at the connection portion. The turbine case according to claim 2, characterized in that the thickness of the continuous rib is 25 mm or more or is greater than the thickness of the radial rib. The turbine case has a first flange to which a shroud case is connected, the first flange having an opening at one end surface, and a second flange to which a combustor is connected, the second flange having an opening at a peripheral wall, the turbine case being cylindrical and asymmetric with respect to a central axis of the first flange ,
4. The turbine case according to claim 1, wherein the continuous rib surrounds the first flange when viewed in the axial direction of the central axis, and is provided in a pattern connecting vertices of a circumferential outer wall curved surface that bulges from an outer periphery of the seating surface of the first flange along the direction of the central axis in a direction in which combustion gas is discharged to the outside . a turbine case having a cylindrical shape with a first flange opening on one end surface and a second flange opening on a peripheral wall, the turbine case being asymmetric with respect to a central axis of the first flange ;
a shroud case housed in the turbine case and connected to the first flange;
a combustor connected to the second flange;
a turbine housed in the turbine case and driven by combustion gas generated in the combustor and supplied through the shroud case;
a continuous rib provided on the turbine case in a pattern that surrounds at least the first flange when viewed in the axial direction of the central axis , and connects vertices of a circumferential outer wall curved surface that bulges from an outer periphery of the seating surface of the first flange along the direction of the central axis in a direction in which combustion gas is discharged to the outside.

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