Abstract: A gas turbine capable of minimizing the clearance between each of the rotor blades and partition ring, uses a structure in which the upstream thermally insulating ring is installed on the downstream blade ring.

Abstract: In order to provide an axial compressor, and a gas bleeding method to the thrust balance disks thereof, which are capable of preventing reduction of the mechanical strength of the thrust balance disks due to excessive heating up by bleed gas, a portion of the fuel gas F, after having passed through the diffuser, is conducted to the thrust balance disks as the bleed gas flow c.

Abstract: In order to provide a gas turbine and a gas bleeding method which can prevent the loss of drive power due to gas bleeding to the rotor disk, bleed gas is imparted with swirling flow in the same rotational direction as that of a first stage rotor disk by being passed through a set of TOBI nozzles which constitute a flow conduit therefor, and is supplied to this first stage rotor disk, with a portion of this bleed gas flow being bypassed and being supplied between first stage stationary blades and first stage moving blades.

Abstract: A plurality of moving blades are provided around rotor disks and rotate together with said rotor disks. Compressor rear case rings surround the periphery of these moving blades and form a compression flow path therein. A bleeding chamber is provided around the compressor rear case rings and introduces a portion of a main flow moving through the compression flow path as bleed air. Cooling flow path is formed between the compressor rear case rings and the bleeding chamber in which bleed air cf on its way to the bleeding chamber flows along the outer surface of the compressor rear case rings.

Abstract: In order to provide an axial compressor, and a gas bleeding method to the thrust balance disks thereof, which are capable of preventing reduction of the mechanical strength of the thrust balance disks due to excessive heating up by bleed gas, a portion of the fuel gas F, after having passed through the diffuser, is conducted to the thrust balance disks as the bleed gas flow c.

Abstract: The present invention reduces the amount of gas which leaks and flows from a high pressure side to a low pressure side of the axis and maintains good sealing performance even at high differential pressures. The invention discloses an axis sealing mechanism comprising casings which are supported inside of the stable portion, a plurality of planar plates, which are mutually spaced around the axis, each of which is fixed to the casings, positioned so that an inner end of each planar plate makes an acute angle with a peripheral surface of the axis and contacts the peripheral surface of the axis over a predetermined length in an axial direction, a high pressure side plate which is arranged at a high pressure side and a low pressure side plate which is arranged at a low pressure side so as to hold the planar plates therebetween and a flexible plate which is arranged between the planar plates and the high pressure side plate, and is flexible in an axial direction.

Abstract: The shaft seal structure is based on a leaf seal 25 that includes springs, disposed between the stator blades 24a and the leaf seal ring 26, to force the leaf seal ring 26 away from the rotation shaft 23, and pressure guiding grooves 34 for guiding the fluid pressure, through a boundary of planar plates 28, to a space between the outer peripheral surface of the leaf seal ring 26 and the inner peripheral surface of the stator blade 24a. Also, a turbine may be provided with the leaf seal so as to reduce gas leakage from the high-pressure-region to the low-pressure-region as well as to control the frictional wear between the planar plates and the rotation shaft.

Abstract: The interior of a gas turbine moving blade is sectioned by a rib into cooling passage portions. Cooling air enters one of the cooling passage portions and turns to flow into the other cooling passage portion. A stagnation area occurs in an end corner portion of the one cooling passage portion, but a cooling hole is provided so that air flow comes outside of the blade through the cooling hole, and thus cooling air flow occurs thereat. Also, a separation area occurs in a tip end portion of the rib due to separation of air flow, but another cooling hole is provided so that air flow comes outside of the blade and cooling air flow occurs thereat. Further, in a gas turbine moving blade having turbulators provided in multi-stages on a cooling passage inner wall, a film cooling hole structure for eliminating separation of cooling air flow between the turbulators is also provided.

Abstract: A seal ring separating surface for a gas turbine, which sets a shape of an end portion in the separation surface to a cutting surface and reduces the leaking amount of sealing air. A seal ring holding ring (1) fixes brush seals (3, 4) on an upstream side of a stationary blade by bolts (5, 6) respectively, so as to form a seal with respect to a rotor disc (69). Further, the seal ring holding ring (1) fixes and supports a seal ring (2) on a downstream side by a bolt (7) so as to form a seal with respect to a seal portion (8) on the disc (69) side. The brush seals (3, 4) and the seal ring (2) are structured so as to have a circular ring shape and a separation construction, and each of the separation pieces has a gap, so that sealing air can leak to a downstream side from an upstream side through the gap.

Abstract: An object of the present invention is to provide a shaft seal which reduces the leakage of gas from the high pressure side to the low pressure side, and has excellent corrosion-resistance, and a gas turbine which is equipped with the shaft seal; in order to achieve the object, the present invention provide a shaft seal comprising a plurality of flexible thin plates having width in the axial direction of a rotary shaft, tips thereof sliding against a peripheral face of said rotary shaft, and outer peripheral bases thereof being secured to a casing side maintaining a gap therebetween, said thin plates being laminated in the peripheral direction of said rotary shaft, enabling the outer periphery of said rotary shaft to be sealed, an acute angle being formed between the thin plates and the peripheral face of said rotary shaft; and a low pressure side side-plate and a high pressure side side-plate which are provided on both sides of said thin plates in the axial direction of said rotary shaft, wherein it further c

Abstract: A shaft seal having a high abrasion resistance is disclosed, by which the leakage of the gas from the high-pressure side to the low-pressure side can be reduced. In the shaft seal, flexible leaves are multi-layered to form a ring shape. The shaft seal is mainly arranged around the rotation shaft of a gas turbine or the like. The relevant turbine comprises a casing, a compressor, a rotation shaft, moving blades attached to the rotation shaft, and stationary blades attached to the casing in a manner such that the stationary blades face the moving blades, wherein the shaft seals are provided between a plurality of stationary blades and the rotation shaft wherein the leaves of each shaft seal contact the rotation shaft. Under the rated operating conditions, the top ends of the leaves slightly separate from the surface of the rotation shaft due to the dynamic pressure generated by the rotation of the rotation shaft.

Abstract: A sealing device for a gas turbine stator blade, in which an outer shroud (32) is mounted by heat insulating rings (32a,32b) on a blade ring (50). The blade ring (50) has a first air hole (1), which communicates with a space (53), and a second air hole (51), which communicates with a seal tube (2). The seal tube (2) is inserted into the second air hole (51), and a spring (6) is arranged between a projection (4) of the tube (2) and a retaining portion (5) of the air hole (51) to removably secure the seal tube (2). Cooling air (54) flows through the first air hole (1) to cool the shrouds and the inside of a stator blade (31) until it is released from the trailing edge of the blade. The cooling air also flows into a cavity (36) so that a high pressure can be maintained without a pressure loss because the tube (2) is independent of the space (53) in the blade ring.

Abstract: A gas turbine moving blade assembly in which a cooling effect at a leading edge of the blade, which is exposed to a high temperature combustion gas, is enhanced by an arrangement of turbulators in the leading edge. In particular, the turbulators are arranged on a slant and are arranged locally only on portions in which the cooling effect is to be reinforced, so that pressure loss of the cooling air is suppressed to a minimum level. Turbulators (8) are provided obliquely from the leading edge in a direction facing into a flow of the cooling air toward a downstream side of a combustion gas flow on both inner walls within a cooling passage (12). Also, to reduce pressure loss, short turbulators (18) may be similarly arranged between the adjacent turbulators (8). In this case, the ratio of the length (W) of the turbulators to that (Wr) of the short turbulators is Wr/W<0.5.