Turbine case and gas turbine

Abstract

To provide a turbine case that does not cause thermal deformation due to heat and pressure.SOLUTION: A turbine case 2a has a first flange F1 to which a shroud case 9 and an exhaust diffuser 10 are connected, and a second flange F2 that is provided on an outer wall side surface 17 and in which a combustor 14 is provided, where the turbine case has an asymmetrical shape with respect to a central axial line Z. Around the first flange, a plurality of radial ribs 20 and circular ribs 30 connecting the radial ribs are provided. Since the circular ribs are provided around the first flange that is significantly affected by heat and pressure, it is possible to suppress thermal deformation of the turbine case at a low cost compared to measures such as changing to an expensive material, cooling the shroud case or the like.SELECTED DRAWING: Figure 4

Description

Technical Field

[0001] The present invention relates to a turbine case for assembling various components including a turbine and housing them inside, in a gas turbine that rotates a turbine by combustion gas generated by a combustor. In particular, the present invention relates to a turbine case that suppresses thermal deformation due to heat and pressure during driving, and a gas turbine using the same.

Background Art

[0002] Patent Document 1 discloses an invention of a joined structure in which a tubular structural member and a substrate such as a base plate or a joining flange are welded and reinforced using reinforcing ribs. The reinforcing rib in this invention has a U-shaped structure that wraps around the outer periphery of the tubular structural member. With this rib structure, both the fatigue strength of the welded end portion on the outer side of the rib bending side and the fatigue strength of the rib open end close to the base plate or joining flange side can be simultaneously increased, thereby further enhancing the strength and fatigue performance.

[0003] Patent Document 2 discloses an invention related to the cooling of a low-pressure turbine casing in a gas turbine engine. This invention is not aimed at suppressing the deformation of the casing, but rather aims to cool the long and slender casing peculiar to a jet engine by flowing cooling air between the shroud and the casing.

[0004] Patent Document 3 discloses an invention related to the rib structure of an exhaust manifold of an internal combustion engine. The reinforcing rib of this invention is provided by connecting between two flanges at both ends of the manifold, and it is said that the thermal deformation of the manifold can be suppressed thereby.

Prior Art Documents

Patent Documents

[0005]

Patent Document 1

Patent Document 2

[0006] A gas turbine is an internal combustion engine that supplies high-temperature gas generated by a combustor to a turbine provided in a turbine case, and rotates the turbine and the rotor attached thereto to obtain rotational kinetic energy. The inventors of the present application are engaged in the research and development of gas turbines. In the gas turbines known to the inventors of the present application before the present invention, the turbine case may be deformed by the heat and pressure of the high-temperature combustion gas as the power source. When the turbine case is deformed, the clearance (tip clearance) between the turbine rotating at high speed and other components such as the shroud case changes, and the components may be damaged or the engine performance may deteriorate. For this reason, the inventors of the present application have made it an issue of recent research and development to suppress the deformation of the turbine case due to heat and pressure. The present invention has been made to solve such problems, and an object thereof is to provide a turbine case that does not cause thermal deformation due to heat and pressure when driving a gas turbine, and a gas turbine using the same. [Means for Solving the Problems]

[0007] The turbine case according to claim 1 is a turbine case having a flange to which other parts are connected, When viewed from the axial direction of the central axis of the flange surrounding the flange and provided in a pattern connecting the apexes of the circumferential outer wall curved surfaces that bulge in the direction in which combustion gas is discharged outward along the direction of the central axis from the outer periphery of the seating surface of the flange characterized by comprising continuous ribs.

[0008] The turbine case according to claim 2 is the turbine case according to claim 1, wherein it comprises a plurality of radial ribs surrounding the flange at equal intervals in the circumferential direction, the continuous rib and the radial rib are connected to intersect at a right angle at the connection part, In the connection part, the space between the continuous rib and the radial rib is formed by a curved surface.

[0009] The turbine case according to claim 3 is the turbine case according to claim 2, wherein the thickness of the continuous rib is 25 mm or more or is greater than the thickness of the radial rib.

[0010] The turbine case according to claim 4 is the turbine case according to any one of claims 1 to 3, wherein a first flange to which a shroud case is connected opens on one end face, and a second flange to which a combustor is connected opens on the peripheral wall, the first flange of central axis a cylindrical shape having an asymmetric structure, and the continuous rib at least When viewed from the axial direction of the central axis surrounds the first flange and provided in a pattern connecting the apexes of the circumferential outer wall curved surfaces that bulge in the direction in which combustion gas is discharged outward along the direction of the central axis from the outer periphery of the seating surface of the first flange and is characterized in that.

[0011] The gas turbine according to claim 5 is a turbine case in which a first flange opens on one end face and a second flange opens on the peripheral wall, the first flange of central axis a cylindrical shape having an asymmetric structure, 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 by the combustor and supplied through the shroud case, and at least surrounds the first flange When viewed from the axial direction of the central axis and is provided with a continuous rib provided on the turbine case and is characterized in that. and provided in a pattern connecting the apexes of the circumferential outer wall curved surfaces that bulge in the direction in which combustion gas is discharged outward along the direction of the central axis from the outer periphery of the seating surface of the first flange

Advantages of the Invention

[0012] ​According to the turbine case described in claim 1, since other components are connected, continuous ribs are provided around the flange that is greatly affected by heat and pressure. Therefore, compared with the measures of changing to a more expensive material with higher durability than before, or the conventional measures such as cooling the components connected to the turbine case, for example, the shroud case and the turbine scroll, the thermal deformation of the turbine case can be suppressed at a low cost.

[0013] Also, according to the turbine case described in claim 1, by providing continuous ribs, the surface area of the turbine case cooled by normal-temperature air increases, and since the continuous ribs function as cooling fins, the cooling performance of the turbine case by outside air can be improved.

[0014] According to the turbine case described in claim 2, at the connection part where the continuous ribs and the radial ribs intersect at a right angle, since the space between the continuous ribs and the radial ribs is configured by a curved surface, while suppressing the stress generated in the turbine case, the rigidity can be increased, and thereby the thermal expansion concentrated on the surface of the turbine case can be suppressed.

[0015] According to the turbine case described in claim 3, by optimizing the thickness of the continuous ribs, permanent deformation can be prevented from occurring in the turbine case composed of an inexpensive material that has been conventionally used.

[0016] The turbine case according to claim 4 and the turbine case used in the turbine of claim 5 are cylindrical with a first flange to which the shroud case is connected opening at one end face. However, since the second flange to which the combustor is connected opens in the peripheral wall, it has an asymmetric structure with respect to the center line of the cylinder. Therefore, if no measures are taken, due to the high heat and pressure generated in the turbine case when the gas turbine is driven, stress concentration will occur at specific locations in the asymmetric turbine case, and the turbine case may be deformed. However, since the turbine case is provided with continuous ribs surrounding at least the first flange, the thermal deformation of the turbine case can be suppressed at low cost, and the cooling performance by the outside air of the turbine case is also improved. Therefore, a gas turbine with high durability of the turbine case can be realized.

Brief Description of the Drawings

[0017]

Figure 1

Figure 2

Figure 3

Figure 4

Figure 5

Figure 6

Mode for Carrying Out the Invention

[0018] To explain the characteristic configuration of the gas turbine of the embodiment and the effects that cannot be obtained with the gas turbine in the current shape due to this, first, the structure common to the gas turbine in the current shape and the gas turbine of the embodiment will be described with reference to FIGS. 1 to 4. Note that the gas turbine in the current shape does not mean what was known before the present invention, but means the prior art known to the inventors of the present invention before the present invention.

[0019] As shown in FIG. 1, the gas turbine 1 includes an outer casing composed of a substantially cylindrical turbine case 2 and a substantially cylindrical compressor case 3 coaxially connected to the turbine case 2. At the center inside this outer casing, a common rotor 4 is rotatably supported via a plurality 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 thereto. Further, a duct 8 for taking in intake air is attached to the outside of the compressor case 3.

[0020] As shown in FIGS. 1, 3(a) and 4(a), the turbine case 2 is a substantially cylindrical member as described above. A first flange F1 for attaching various members to be described later opens at one end surface, and a second flange F2 opens on the peripheral wall. For this reason, although the turbine case 2 is substantially cylindrical, it has an asymmetric structure with respect to its center line (central axis Z).

[0021] As shown in FIGS. 1 and 2, a shroud case 9 and an exhaust diffuser 10 are attached to the first flange F1 of the turbine case 2. In particular, as shown enlarged in FIG. 2, the first flange F1 of the turbine case 2, the attachment portion of the shroud case 9, and the attachment portion of the exhaust diffuser 10 are overlapped and integrally fastened by a common bolt 11. The shroud case 9 and the exhaust diffuser 10 constitute an exhaust passage 12 for discharging the combustion gas after driving the turbine 7.

[0022] As shown in FIG. 1, an outer cylinder 13 is attached to the second flange F2 of the turbine case 2. Inside the outer cylinder 13, a can-type combustor 14 for supplying combustion gas for driving the turbine 7 is provided at a predetermined interval from the inner wall of the outer cylinder 13. The can-type combustor 14 is formed with a plurality of through holes for taking in air. Although not shown in detail, a fuel pipe for supplying fuel is connected to the can-type combustor 14, and a spark plug is provided.

[0023] As shown in FIG. 1, inside the turbine case 2, a turbine scroll 15 that constitutes a combustion gas flow path for guiding the combustion gas sent from the can-type combustor 14 to the turbine 7 is provided. 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. Also, an air flow path 16 for supplying 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 communicates the compressor case 3 provided with the compressor 6 and the space between the outer cylinder 13 and the can-type combustor 14.

[0024] According to the above configuration, when the compressed air compressed by the compressor 6 is supplied from the air flow path 16 to the can-type combustor 14 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, and an ignition plug ignites and burns this 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. After driving the turbine 7 to rotate the rotor 4, it is discharged to the outside from the exhaust flow path 12 formed by the shroud case 9 and the exhaust diffuser 10.

[0025] Next, problems caused by the structure of the current-shaped gas turbine 1 will be described. For the turbine case of the embodiment, the reference numeral 2a is used, for the current-shaped gas turbine, the reference numeral 2b is used, and when there is no particular distinction, the reference numeral 2 is used for the description. As described above, the turbine case 2 is a complex structure that serves multiple roles, such as forming the air flow path 16 between the shroud case 9, shielding heat from the turbine scroll 15 where combustion gas at 1000 °C flows through the internal combustion gas flow path, housing the rotor 4 and the turbine 7, and holding the can-type combustor 14 and the shroud case 9 in predetermined positions.

[0026] In such a complex structure, when the turbine case 2 is deformed by heat or internal pressure, the clearance (tip clearance) between the outer wall of the shroud case 9 and the tip of the turbine 7 changes. If this clearance becomes narrow, the outer wall of the shroud case 9 may come into contact with the turbine 7 and be damaged. If it becomes wide, the aerodynamic performance of the turbine 7 may be significantly reduced, which may deteriorate the engine performance. In addition, when the turbine case 2 is deformed, the holding position of the bearing 5 of the rotor 4 changes, resulting in misalignment of the rotor 4 and an increase in vibration. Furthermore, due to the deformation of the turbine case 2, the sealing performance of the seating surfaces of the first flange F1 and the second flange F2, which are the connection surfaces with the shroud case 9 and the can-type combustor 14 shown in FIGS. 3(a) and 4(a), is significantly reduced, and compressed air may leak out of the turbine case 2, making engine operation difficult. Moreover, if the turbine case 2 is deformed, the compressor case 3 connected to it may also be deformed.

[0027] The main factor causing the deformation of the shape of the turbine case 2 is heat transfer from the high-temperature fluid. The outer wall of the turbine case 2 is constantly cooled by ambient air, while compressed air, which is a high-temperature fluid close to the discharge air temperature of the compressor 6 of 300°C or higher, flows along its inner wall. In addition, combustion gas at 1000°C flows through the turbine scroll 15 provided inside the turbine case 2, and the turbine case 2 receives strong radiation. Furthermore, the first flange F1 and the second flange F2 are respectively connected to the shroud case 9 that partitions the exhaust passage 12 and the can-type combustor 14 that becomes hot due to combustion, and a temperature close to 700°C is directly transmitted to the turbine case 2 by heat conduction.

[0028] Thus, the turbine case 2 is exposed to high temperatures in each part and is in a harsh environment where there is a large temperature difference between the parts. However, since the turbine case 2 has a structure in which a large number of components are connected and assembled integrally as described above, it is difficult to thermally expand, and stress is generated in a non-uniform manner in various places. In particular, as shown in FIGS. 3(a) and 4(a), the portion between the seating 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 portion that bulges and curves semi-cylindrically in the central axis Z direction of the cylindrical turbine case 2b from the outer periphery of the seating surface of the first flange F1, although the outside thereof is cooled by ambient temperature air, it is prone to temperature differences because it directly receives heat conduction from the high-temperature components inside, and is prone to deformation due to heat. The vicinity of the seating surface of the second flange F2 to which the can-type combustor 14 is connected is the same.

[0029] As shown in FIG. 3(a), in order to suppress this thermal deformation, a plurality of radial ribs 20 are continuously provided on the current-shaped turbine case 2b from the outer peripheral portion of the first flange F1 to the outer wall side surface 17 via the outer wall curved surface 18. 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. The radial ribs 20 are parallel to the central axis Z of the circular first flange F1 on the outer wall side surface 17. Further, the end portions of the radial ribs 20 on the outer wall side surface 17 are continuous with a slope having a certain angle with respect to the outer wall side surface 17 of the turbine case 2b.

[0030] However, in the current-shaped turbine case 2b, heat distortion cannot be suppressed by this radial rib 20, and large stresses are locally generated. Fig. 3(b) is a diagram showing the stress applied to the current-shaped turbine case 2b when the gas turbine 1 having the current-shaped turbine case 2b shown in Fig. 3(a) is driven under predetermined conditions, calculated by simulation using the finite element method, and the result is shown as a shaded graphic on a 3D image. As shown in Fig. 3(b), a part of the outer wall curved surface 18 near 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 at a high temperature and has a significant temperature difference, so high stress is generated as represented by a dark gray in the figure. Stress is also concentrated at a position on the opposite side (referred to as the second concentration point P2.) that is rotated 180° in the circumferential direction of the circular first flange F1 with respect to the first concentration point P1 on the outer wall curved surface 18 where this high stress is generated. At these locations, due to the concentration of stress, the surface of the turbine case 2b is thermally deformed.

[0031] When the gas turbine 1 repeatedly starts and stops, the turbine case 2 repeatedly heats and cools, so fatigue accumulates and there is concern about a reduction in fatigue life. Although no cracks or the like have occurred in the current-shaped turbine case 2b, since an increase in the pressure ratio is assumed in the future development of gas turbines, the inventors of the present application considered that some countermeasures against thermal deformation are necessary for the turbine case 2. Furthermore, the above-described current-shaped turbine case 2b is of a type having a can-type combustor 14, and the turbine case 2a of the present invention for improving this is also considered to be suitably applied to a type having a can-type combustor 14. However, the turbine case 2 of this type is different from a cylindrical turbine case that is axisymmetric like an annular combustor typified by a jet engine. That is, the turbine case 2 used in the gas turbine 1 of the type having a can-type combustor 14 has an asymmetric structure in which a cylinder having a smaller radius than the main body is embedded in the side peripheral surface of the cylindrical main body in order to connect the can-type combustor 14. Therefore, the difficulty of arranging and designing the above-described radial ribs 20 is high, and as described with reference to FIG. 3(b), it is hard to say that the thermal deformation of the turbine case 2b is effectively suppressed.

[0032] As described above, the structure common to the current-shaped gas turbine 2b and the gas turbine 2a of the embodiment, and the problems of the radial ribs 20 in the current-shaped gas turbine 2b have been described. Based on these, the structure of the continuous ribs, which is a feature of the present invention, and the effects of solving the problems of the current shape will be described below with reference to FIGS. 4 to 6.

[0033] As shown in Fig. 4(a), in the turbine case 2a of the embodiment, in addition to the radially rib 20 in the current shape, a circular rib 30 (hereinafter also referred to as a ring rib 30) as a continuous rib intersecting therewith is provided. The circular rib 30 is provided in a closed circular pattern connecting the vertices of the outer wall curved surface 18 surrounding the first flange F1, intersects all of the plurality of radially ribs 20, and is connected to each of them. That is, the "continuity" of the continuous rib indicates a state in which one line is closed and forms a loop regardless of its shape. For example, it may be circular like the circular rib 30, or may be elliptical or rectangular depending on the outer shape of the flange.

[0034] As shown in Fig. 4(a), the height from the surface of the turbine case 2a (that is, the surface of the outer wall curved surface 18) to the surface of the ring rib 30 is the same as that of the radially rib 20, and the ring rib 30 has a plane parallel to the surface of the turbine case 2a and having a constant thickness T on the surface. The surface of the ring rib 30 has a ring shape with an inner circumference of a constant radius and an outer circumference of a radius obtained by adding the thickness T to the constant radius. The radially rib 20 and the ring rib 30 intersect and are connected at right angles on the outer wall curved surface 18, and the connection portion is formed of a curved surface having a constant radius R. At the connection portion, since the circular rib 30 and the radially rib 20 are connected by a curved surface, when viewed from a line of sight parallel to the axial direction Z of the first flange F1, the connection portion exhibits a shape similar to a rhombus or a star shape.

[0035] Figure 4(b) is a diagram in which the stress applied to the turbine case 2a of the embodiment when the gas turbine 1 is driven under predetermined conditions is calculated by simulation using the finite element method (FEM), and the result is shown as a shaded graphic on a 3D image. As described above with reference to Figure 3(b), in the turbine case 2b of the current shape, among the outer wall curved surfaces 18 of the turbine case 2, stress was particularly concentrated at the first concentration point P1 where the outer wall curved surface 18 was the shortest due to the presence of the second flange F2, and at the second concentration point P2 on the opposite side by 180° from this. 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 rib 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 additionally providing the ring rib 30, it was possible to suppress the stress generated by heat and internal pressure, increase the rigidity, and thereby suppress the thermal expansion that had concentrated at specific locations on the surface of the turbine case 2b.

[0036] In the turbine case 2a of the embodiment, since the ring rib 30 is additionally provided in addition to the radial rib 20, 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, increases compared to the turbine case 2b of the current shape. Therefore, heat transfer from the surface of the turbine case 2a to the air is promoted, and the effect of suppressing the thermal expansion of the turbine case 2a becomes higher.

[0037] Referring to Figure 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 are compared during the operation of the gas turbine 1. Fig. 5 shows the stresses generated in the turbine cases 2a and 2b under predetermined operating conditions for the turbine case 2b in its current shape and two types of turbine cases 2a with different thicknesses T of the ring rib 30, which are calculated by simulation using the FEM method, and the results are shown as the relationship between the internal pressure ratio and the stress ratio. Sub-figure (a) is a table showing the numerical values of the calculation results, and sub-figure (b) is a graph of the calculation results. Here, the stress ratio is the ratio of the compressive stress to the 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 the compressive stress is generated to the pressure applied inside the position where the tensile / compressive stress is generated.

[0038] Here, the thickness of the radial rib 20 of the turbine case 2b in its current shape is 10 mm. Also, the two types of turbine cases 2a of the embodiment are a turbine case 2a (ring rib 30R20T20) provided with a ring rib 30 having a radius R of the curved surface of the connection part of 20 mm and a thickness T of 20 mm in addition to the radial rib 20 of the turbine case 2b in its current shape, and a turbine case 2a (ring rib 30R20T25) provided with a ring rib 30 having a radius R of the curved surface of the connection part of 20 mm and a thickness T of 25 mm in addition to the radial rib 20 of the turbine case 2b in its current shape.

[0039] As the material of the turbine case 2, FCD (spheroidal graphite cast iron) has been conventionally known as a common and inexpensive material. The limit line at which the stress ratio becomes 1 regardless of the internal pressure ratio even at a predetermined high temperature for this FCD is shown in Fig. 5 together with the calculation results. If the stress ratio of the generated stress is lower than this limit line, it is possible to avoid thermal deformation in which stress exceeding the 0.2% proof stress is generated and permanent strain remains even if the turbine case 2 is made of FCD.

[0040] According to FIGS. 5(a) and 5(b), for example, when looking at the stress ratio at an internal pressure ratio of 1.0, in the current shape (shown by the solid line in sub - figure (b)), it becomes 1.126 which exceeds the limit line of the stress ratio of 1.0, and thermal deformation occurs where permanent strain remains due to the stress ratio exceeding the 0.2% proof stress. In contrast, for the turbine case 2a of the embodiment (ring rib 30R20T20, shown by the dashed line in sub - figure (b)), the stress ratio is 1.046. Although the stress ratio slightly exceeds the limit line of the 0.2% proof stress, it is recognized that the possibility of avoiding thermal deformation where permanent strain remains is there, and even if permanent strain remains, it is very small. Further, for the turbine case 2a of the embodiment (ring rib 30R20T25, shown by the chain - dotted line in sub - figure (b)), the stress ratio is 0.950, and since the stress ratio is below the limit line of the 0.2% proof stress, thermal deformation where permanent strain remains does not occur.

[0041] Thus, in the turbine case 2a of the embodiment, by setting the thickness T of the continuous rib to an optimal value that is at least greater than the value (T20) of the radial rib 20 in the current shape or 25 mm or more, even if the turbine case 2a is made of an inexpensive material such as FCD that has been conventionally used, permanent deformation due to heat and pressure can be substantially prevented from occurring.

[0042] Referring to FIG. 6, during the operation of the gas turbine 1, the displacement occurring at the first flange F1 of the current - shape turbine case 2b is compared with the displacement occurring at the first flange F1 of the turbine case 2a of the embodiment. FIG. 6 shows, for the current - shape turbine case 2b and the turbine case 2a of the embodiment, the axial displacement ratio at the peripheral edge of the first flange F1 when the gas turbine 1 is driven under predetermined conditions, calculated by simulation using the finite - element method, and shown for each phase angle (°) of the peripheral edge of the first flange F1. In FIGS. 3(a) and 4(a), the axial direction serving as the reference for representing the displacement ratio is the central axis Z of the circular first flange F1, and the phase angle (°) of the peripheral edge of the first flange F1 means the circumferential angle with TOP being 0°, RIGHT being 45°, BOTTOM being 180°, and LEFT being 270°.

[0043] As shown in FIG. 6, during the operation of the gas turbine 1, the axial displacement ratio (indicated by the broken line) occurring in the first flange F1 of the turbine case 2b in the current shape exceeds 0.9 for all phase angles. In particular, at the position between LEFT (270°) and TOP (0°), that is, at the first concentrated point P1 shown in FIG. 3(a), a value close to 1.1 is shown. This indicates that, as described above, at the first concentrated point P1, large local deformation occurs due to high heat.

[0044]

[0043] In contrast, as shown in FIG. 6, during the operation of the gas turbine 1, the axial displacement ratio (indicated by the solid line) occurring in the first flange F1 of the turbine case 2a of the embodiment is less than 0.7 for all phase angles and shows an almost constant value, and no situation where the axial displacement ratio is particularly large and prominent at a specific phase angle is observed. This indicates that, according to the embodiment, the deformation in the central axis Z direction occurring around the first flange F1 due to heat and internal pressure is smaller than that of the turbine case 2b in the current shape and is averaged in the circumferential direction of the first flange F1.

[0045] In the embodiment described above, in addition to the radial rib 20, the ring rib 30 is provided. However, even when only the ring rib 30 is provided without providing the radial rib 20, the effect of suppressing thermal deformation can be obtained. Therefore, in the embodiment, the ring rib 30 is provided on the first flange F1, but the ring rib 30 may be provided on the circumferential surface around the second flange F2.

[0046] According to the turbine case 2a of the present embodiment, since the ring rib 30 is added as a reinforcing rib as described above, the thermal deformation of the turbine case 2a caused by the stress generated by heat and pressure can be suppressed, and thus the following effects can be obtained. (1) The engine performance of the gas turbine 1 can be stably maintained (2) Since the thermal deformation of the turbine case 2a can be suppressed even with the inexpensive materials currently used, there is no particular need to use high-grade heat-resistant materials. (3) Since the thermal deformation of the seating surfaces of the flanges F1 and F2 that hold the shroud case 9 can be suppressed, the holding force of the shroud case 9 is improved. (4) Since the change in the position of the bearing 5 that holds the rotor 4 in the turbine case 2a and the compressor case 3 can be suppressed, the 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 heat insulation of the heat received from the turbine 7. (6) The ring rib 30 has a relatively simple structure, and it is possible to take measures against thermal deformation at low cost. (7) Since the pressure resistance performance of the turbine case 2a is improved, the life of the turbine case 2a is extended. (8) By keeping the tip clearance of the turbine 7 constant, the performance of the turbine 7 can be maintained over a long period. (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 becomes small. For this reason, performance management during driving and assembly after disassembly becomes easy. (11) Since the ring rib 30 serves as a cooling fin, the cooling performance by the outside air is improved.

[0047] Further, the turbine case 2a of the present embodiment can be adopted not only for the turbine cases of gas turbine engines including marine, land, and aviation use and industrial turbines, but also for the turbine cases of superchargers and turbochargers, and the housings of pressure vessels that also serve to support pipes such as boilers. In these various applications, the thermal deformation of the turbine case or the housing can be suppressed at low cost.

Description of Reference Numerals

[0048] 1…Gas turbine 2…Turbine case 2a…Turbine case of the embodiment 2b…Current turbine case 7…Turbine 9… Shroud case 14… Cylindrical combustor 20… Radial rib 30… Circular rib (ring rib) as continuous rib F1… First flange F2… Second flange T… Thickness of circular rib

Claims

1. A turbine case having a flange to which other components are connected, comprising continuous ribs provided in a pattern connecting the vertices of a circumferential outer wall curved surface that surrounds the flange when viewed from the axial direction of the central axis of the flange and bulges in a direction in which combustion gas is discharged outward along the direction of the central axis from the outer periphery of the seating surface of the flange. The turbine case is characterized by this.

2. Comprising a plurality of radial ribs surrounding the flange at equal intervals in the circumferential direction, the continuous rib and the radial rib are connected at a right angle at the connection part, The turbine case according to claim 1, wherein between the continuous rib and the radial rib at the connection part is formed of a curved surface.

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

4. One end face of the turbine case has an opening for a first flange to which a shroud case is connected, and the circumferential wall has an opening for a second flange to which a combustor is connected. The turbine case is cylindrical with a structure asymmetric with respect to the central axis of the first flange, The continuous rib at least surrounds the first flange when viewed from the axial direction of the central axis, and is provided in a pattern connecting the vertices of a circumferential outer wall curved surface that bulges in a direction in which combustion gas is discharged outward along the direction of the central axis from the outer periphery of the seating surface of the first flange. The turbine case according to any one of claims 1 to 3 is characterized by this.

5. A turbine case in which a first flange has an opening at one end face and a second flange has an opening in the circumferential wall, and the turbine case is cylindrical with a structure asymmetric with respect to the 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 by the combustor and supplied through the shroud case, A gas turbine characterized by comprising continuous ribs provided on the turbine case in a pattern connecting the vertices of a circumferential outer wall curved surface that at least surrounds the first flange when viewed from the axial direction of the central axis and bulges in a direction in which combustion gas is discharged outward along the direction of the central axis from the outer periphery of the seating surface of the first flange.

Images