Abstract: The invention relates to a rotor for thermal turbomachines in a steam power process with a centrally arranged hollow space, several axis-symmetrical hollow spaces, at least two pipes with different diameters and lengths, and at least two through-openings in the mantle, whereby at least one opening is arranged in the turbine part and at least one opening in the compressor part or the middle part, and the openings of the various pipes overlap in the warm operating state in the turbine part, while they overlap in the compressor part and middle part in the cold operating state. At the waste gas side end of the rotor, the hollow space is in the cylindrical part limited by the inside wall of the through-pipe and is provided for supplying steam. The pipe is surrounded in the cylindrical part by another hollow space that is arranged concentrically around the pipe and is provided for removing steam.

Abstract: An air separator for a gas turbine, in which cracks are prevented from occurring at a flange portion of the air separator due to fretting fatigue. The air separator (20) has a cylindrical split structure formed by two separator members (20-1 and 20-2), of which one of the separators (20-1) is mounted on a rotor (1) whereas the other separator member (20-2) is mounted on a disc portion (7) on a side of a moving blade (2) by bolts (28) extending through bolt holes (23) formed in a flange portion (22). Cooling air from a compressor enters a space (6) from a duct (5) and passes to a passage (32) from a clearance (33) so that it is fed to air feed holes (43) and radial holes (44) of the disc portion (7). Alternatively, air holes (50) are provided in the flange portion (22) in the form of circumferential slots for covering a plurality of radial holes to feed the cooling air homogeneously so that a prior art air separator of an existing plant can be replaced.

Abstract: An apparatus (42) for inspecting an as cast turbine blade (40), with impingement cooling passages (36) (see FIG. 2 ), comprises a vacuum pump (52) arranged to evacuate the interior of the turbine blade (40) through pipe (48) and valve (54) and a supply of steam (46) arranged to supply steam to the interior of the turbine blade (40) through the pipe (48) and a valve (50). An infrared camera (56) is arranged to view the outer surface (32) of the turbine blade (40) to detect hot spots produced on the outer surface (32) of the turbine blade (40) by jets of steam impinging on the inner surface (35) of the turbine blade (40) from the cooling passages (36). A processor (60) records the images produced by the camera (56) and analyses the images to determine if the cooling passages (36) have been formed and if they have been formed in the correct position.

Abstract: In a turbine rotor, a thermal mismatch between various component parts of the rotor occurs particularly during transient operations such as shutdown and startup. A thermal medium flows past and heats or cools one part of the turbine which may have a deleterious thermal mismatch with another part. By passively controlling the flow of cooling medium past the one part in response to relative movement of thermally responsive parts of the turbine, the flow of thermal medium along the flow path can be regulated to increase or reduce the flow, thereby to regulate the temperature of the one part to maintain the thermal mismatch within predetermined limits.

Abstract: A turbine wall includes an outer surface for facing combustion gases, and an opposite inner surface for being impingement air cooled. A plurality of adjoining ridges and grooves are disposed in the inner surface for enhancing heat transfer by the impingement cooling air.

Abstract: A cooled platform for a gas turbine moving blade which enables the platform to be cooled when the moving blade is cooled by steam. The moving blade is provided with a plurality of steam passages. Steam is introduced from the steam passage at the trailing edge portion, flows in a serpentine passage composed of other steam passages to cool the blade, and flows out to a blade root portion from a base portion of the steam passage at the leading edge portion, being recovered. Part of steam flowing into the platform from the base portion of the steam passage at the trailing edge portion enters first and second steam passages in the platform. On one side, the steam passes through first, third, and fourth steam passages, and on the other hand, the steam passes through second, fifth, and sixth steam passages. The steam is recovered together with the steam having cooled the blade at the base portion of steam passage at the leading edge portion.

Abstract: A turbine engine airfoil includes a three-pass serpentine shaped cooling cavity including a leading edge chamber, an oversized intermediate chamber, and a trailing edge chamber in flow communication. The cavity further includes a combination of a plurality of ribs and a plurality of pins, a purge air swirler, a turning vane, and a metering partition which mixes the high pressure air and provides increased air cooling and uniform flow control in the airfoil. The metering partition provides a decrease of the air pressure in the trailing edge chamber allowing for an increase in the number of trailing edge slots.

Abstract: A turbine rotor blade comprises a shank portion, a tip portion and an airfoil. The airfoil has a pressure side wall and a suction side wall that are interconnected by a plurality of partition sidewalls, defining an internal cooling passageway within the airfoil. The internal cooling passageway includes at least one radial outflow passageway to direct a cooling medium flow from the shank portion towards the tip portion and at least one radial inflow passageway to direct a cooling medium flow from the tip portion towards the shank portion. A number of mixing ribs are disposed on the partition sidewalls within the radial outflow passageways so as to enhance the thermal mixing of the cooling medium flow, thereby producing improved heat transfer over a broad range of the Buoyancy number.

Abstract: In an air circulating fan assembly, the combination comprises a fan motor and fan blades rotated by the motor; a fluid reservoir and a thermo-electric cooler for cooling the water in the reservoir; and ducting associated with the blades to conduct fluid from the reservoir to heat exchange structure rotated to effect heat transfer between air relatively passing the blades and fluid being returned to the reservoir.

Abstract: A rotary assembly of a turbine stage of a gas turbine engine comprising a disc carrying internally air cooled blades around its periphery has a cover plate of one face of the disc to create a plenum for a cooling air supply to the blades. The cover plate is fitted to the disc by means of an annulus of bayonet connections at a first radius and is retained at a second radius, less than the first radius, by a split ring arrangement. The split ring is fitted into a groove in the disc and engages an inner rim on the cover plate to restrain axial movement. An anti-rotation key is engaged in a slot in the cover plate and in the disc and is also retained by the split ring.

Abstract: A gas turbine engine airfoil includes an axial serpentine cooling circuit therein. A plurality of the serpentine circuits are preferably stacked in a radial row along the airfoil trailing edge for cooling thereof.

Abstract: In the present invention, seal air passes through a tube extending in a stationary blade from an outside shroud to an inside shroud, flows into a cavity to keep the pressure in the cavity high so as to seal the high-temperature combustion gas, and is discharged to a passage. Part of cooling air flows into an air passage to cool the leading edge portion, passes through the peripheral portion of the inside shroud to cool the same, and is discharged to a passage. The remaining cooling air flows into a passage, passes through a serpentine cooling flow path consisting of air passages having turbulators, and is discharged through air cooling holes formed at the trailing edge. The cooling air passing through the passage at the leading edge portion cools the peripheral portion of the inside shroud as well as the leading edge portion of blade, by which the cooling efficiency is increased.

Abstract: A seal device between a fastening bolt and a bolthole in a gas turbine disc reduces cooling steam leaking from a gap between the fastening bolt and the bolthole. A gas turbine disc 31 carrying moving blades, a fastening bolt 33 for fixing the disc 31 passes through a bolthole 32. On a low pressure side of the disc 31, a bolthole diameter enlarged portion 39 having a larger diameter than that of the bolthole 32 is provided. Fastening bolt 33 is provided with a shaft diameter enlarged portion 33a. Seal piece 1 is fitted at one end around the shaft diameter enlarged portion 33a and is engaged at the other end with a stepped portion of the bolthole diameter enlarged portion 39. In operation of the gas turbine, the bolt 33 is biased outwardly be centrifugal force, and the gap between the bolt 33 and the bolthole 32 deforms. Nevertheless, the seal piece 1, having flexibility, deforms at bent portions R.sub.1 and R.sub.

Abstract: A gas turbine is equipped with a moving blade cooling apparatus arranged in association with a turbine rotor in which a plurality of discs are mounted, a plurality of moving blades are mounted each to an outer peripheral portion of each of the discs and a plurality of spacers are also disposed in spaces between the respective discs at portions corresponding to a location of stationary blades and in which the moving blades are formed each with a cooling medium flowing passage and the discs and the spacers are arranged with spaces therebetween for passing a cooling medium in a radial direction of the rotor. A closed-loop cooling unit is thus formed for supplying the cooling medium to the cooling medium flowing passage formed in the moving blade and recovering the same therefrom.

Abstract: In a cooled stationary blade assembly for a gas turbine, an interior of a blade and an inner shroud are cooled by steam to eliminate the use of air cooling. Steam passages 33A, 33B, 33C, 33D, 33E and 33F are provided in the stationary blade 30. The cooling steam 39 is introduced from the steam passage 33A on the front edge side through an outer shroud and passes, in order, through the steam passages 33B, 33C, 33D, and 33E to flow into the steam passage 33F at the rear edge side to cool the interior of the blade, and is recovered through the outer shroud from the upper portion of the steam passage 33F. A portion of the steam from the steam passage 33A is introduced into the inner shroud 31, enters to steam passages 20 from a steam introduction passage 22, branches to the right and left through both end portions, and flows out into the steam passage 33F at the rear edge from a steam discharge passage 21.

Abstract: A gas turbine blade having a platform and a turbine wheel plate, in which cooling passages are arranged in a plurality of rows and connected to one another in a blade trunk section of a moving blade and a supply-side passage and a discharge-side passage are formed in a blade root section. The gas turbine blade comprises a pocket having an inlet and an outlet formed in the blade root section. The supply-side passage is connected to one of the cooling passages and the inlet of the pocket, the outlet of the pocket is connected to another of the cooling passages, and the discharge-side passage is connected to one of the cooling passages other than the ones connected individually to the supply-side passage and the pocket.

Abstract: A gas turbine cooled blade has leading edge turbulators. A rounded tip portion of the leading edge portion cooling passage (3) is approximated by a triangular cooling passage (1) having orthogonal turbulators (11, 12) and a smoothly curved rear portion thereof of the leading edge portion cooling passage (3) is approximated by a square cooling passage 2 having oblique turbulators (13, 14).

Abstract: In the cooling system for the gas turbine blade, the apparatus of the present disclosure ensures cooling of the trailing edge part of the blade for which machining is difficult, while aiming at improvement of the thermal efficiency. The steam cooling structure for carrying out heat recovery-type steam cooling is employed to the leading edge part and the central part of the blade where machining is easy because of its large thickness, and the convection cooling end film cooling are employed for the trailing edge part of the blade where the thickness is small, using steam cooling and air cooling at the same time.

Abstract: In a cooling system for the leading-edge region of a hollow gas-turbine blade, a duct (3), through which flow occurs longitudinally, extends from the blade root up to the blade tip and is defined in the region of the blade body (4) by the inner walls of the leading edge (5), the suction side (6) and the pressure side (7) and by a web (8). The inner walls of the suction side and the pressure side are provided with a plurality of ribs (9), which run slantwise and at least approximately in parallel. The suction-side ribs and the pressure-side ribs are offset from one another over the blade height. The ribs (9) run radially inward from the web (8) in the direction of the leading edge (5), merge into the radial in the region of the leading edge and are led around the leading edge. The deviation of the ribs (9) from the slant into the radial is effected with the smallest possible radius.

Abstract: A gas turbine moving blade, which has a blade root section, a wing section, and a platform section, comprises a refrigerant supply port provided in the blade root section, a refrigerant recovery port provided in the blade root section, a serpentine passage provided in the wing section and communicating with the refrigerant supply port and the refrigerant recovery port, and a convection-cooling passage provided in the platform section and allowing sealing air and a refrigerant for cooling the platform section by convection to pass therein.

Abstract: A turbine nozzle includes radially outer and inner bands between which extend a plurality of vanes. The vanes include impingement baffles having impingement holes for impinging cooling air inside the vane for cooling thereof. An air circuit is disposed in the inner band and includes a plurality of outlet holes. A transfer tube is disposed in flow communication between respective ones of the baffles and the air circuit for channeling thereto a portion of the cooling air from inside the baffles as pre-impingement air for discharge through the outlet holes.

Abstract: A turbine (10) for a gas turbine engine includes an annular array of turbine aerofoil blades (12) which are mounted on a disc (19). Each of the aerofoil blades (12) is provided with a radially inner platform (21). Each platform (21) includes a passage (37) into which leaked cooling air flows. The passages (37) are disposed in a direction having a circumferential component so that cooling air is exhausted from them in a direction that is generally opposite to that in which the disc (19) operationally rotates.

Abstract: A monitor for detecting overheating of a critical component in a combustion turbine is provided. The monitor, when used in conjunction with a closed-loop cooling system, comprises a coating comprising an indicator material having an activation temperature. The coating is situated on the internal cooling passages of the critical component. The monitor further comprises a sensor connected to an outlet conduit of the cooling system for determining the amount of degradation of indicator material by monitoring the cooling fluid flowing through the outlet conduit. Embodiments further comprising "sniffer" tubes for use with open-loop air cooling systems also are provided.

Abstract: A turbine blade assembly including an airfoil portion, a root portion and a cooling air plenum tube. A cooling air flow path is formed in the root and airfoil portions of the blade and has first and second inlets and an outlet formed in the bottom of the root. The plenum tube has an open front end that forms a first supply port for receiving a flow of cooling air. Openings in the tube upper portion form first and second discharge ports and a second supply port. An open rear end of the tube forms a third discharge port. A baffle assembly within the plenum tube forms first, second, and third chambers and first and second passages. The first chamber receives cooling air from the first supply port and directs a first portion to the first discharge port, which then directs it to the first inlet of the cooling air flow path. The first chamber directs a second portion of the cooling air to the first passage which, in turn, directs it to the third chamber.

Abstract: In a cooling system for the trailing edge region of a hollow gas turbine blade, there extends from the blade root (1) to the blade tip (2) a duct (3) through which the flow passes longitudinally and which, in the region of the blade body (4), is delimited by the inner walls of the trailing edge (5), the suction side (6) and the pressure side (7) and by a web (9), the inner walls of the suction side and of the pressure side being provided with a plurality of ribs (8) running at least approximately parallel. The ribs (8) run obliquely from the web (9) in the direction of the trailing edge (5) and are directed radially outward on at least one of the two inner walls. The suction-side ribs and the pressure-side ribs are offset relative to one another over the blade height. The ratio of the height (h) of the ribs (8) to the local height (H) of the duct (4) is constant over the longitudinal extent of the ribs.

Abstract: A cooling steam circulation passage for a gas turbine rotor (30) having turbine discs (41.about.47) are composed of center line bores (73.about.77) open at an axial end of the rotor and extending through a central portion of the rotor; a steam inlet-outlet pipe (79) coaxially disposed therein so as to define an annular passage (81) for cooling steam at an outer side; steam cavities (89a, 89b) defined between and by facing side surfaces of said turbine discs; steam cavities (91a, 91b) each defined at non-facing side surface portions of said turbine discs (41, 43); axial steam holes (61, 63) formed to extend through the turbine discs and including a partition tube (99); and radial steam holes (97, 103a, 103b, 105, 107) extending from each of the steam cavities (91a, 101, 89a) to mounting portions for the rotor blades.

Abstract: A turbine shaft extends along a principal axis and has an outer surface. The turbine shaft is formed by a plurality of cylindrical shaft segments which are disposed axially one behind the other and are braced together by a bracing element. An axial gap which is formed between the bracing element and at least one shaft segment is connected in terms of flow to two axially spaced radial passages. The radial passages each open at the outer surface of the turbine shaft. A method for cooling a turbine shaft is also provided.

Abstract: A turbine blade which can be subjected to a hot gas flow includes a substrate, at least one interior space and a plurality of bores leading from the interior space out of the substrate. The substrate is at least partly covered by a heat-insulating-layer system at a suction side and/or a pressure side. At least one of the bores is closed by the heat-insulating-layer system and at least one further bore is open for developing film cooling.

Abstract: The present disclosure provides a gas turbine blade in which the tip end thereof is cooled effectively to decrease the metal temperature for the prevention of burning and the temperature gradient of blade metal is decreased to prevent the occurrence of cracking. In the gas turbine blade provided with a cooling passage therein, a protrusion is provided inwardly of the peripheral walls of the blade, which define the blade profile and the protrusion is positioned on the outer surface of blade tip end wall directly above a cooling passage within the blade.

Abstract: The present invention relates to a gas turbine cooled moving blade in order to cool the gas turbine using steam alone. In the moving blade, two cavities are installed in the lower part of the moving blade root section of the moving blade, a steam inlet is located in one cavity, and a steam outlet is installed in the other cavity. Steam is charged from the steam inlet and passes through a serpentine cooling passage consisting of a plurality of steam passages, and discharged to the steam outlet and collected. Bypasses are located near the base of the moving blade in the middle of each steam passage, so that cold steam on the leading edge side can be led into the bypass, mixed with the hot steam which comes from the tip section side and goes into the next downstream section of the steam passage, and flows to an upstream portion of the steam passage, thereby equalizing the steam temperature and cooling the steam over the entire range from the upstream to downstream of the steam flow.

Abstract: There is provided a cooled stationary blade of a gas turbine in which the portions which can be cooled sufficiently by air are air-cooled, and the portions which are difficult to cool by air are steam-cooled, by which high temperatures can be overcome. In a stationary blade 1, there is formed a serpentine passage 3 in which cooling steam flows and an air passage 10 adjacent to the trailing edge portion and separated from the serpentine passage 3. Also, an outside shroud 4 is formed with an air cooling passage 16 at the outer edge portion and a steam impingement cooling portion 17 and an air impingement cooling portion 18 on the inside of the air cooling passage 16. An inside shroud 11 is provided with an air cooling passage 19 at the outer edge portion and shaped holes 20 formed on the inside of the air cooling passage 19. The air flowing out through the shaped holes 20 performs film cooling.

Abstract: A gas turbine engine airfoil includes first and second sidewalls joined together at opposite leading and trailing edges, and extending longitudinally from a root to a tip. The sidewalls are spaced apart from each other to define in part first and second adjoining flow chambers extending longitudinally therein, and defined in additional part by corresponding first and second partitions disposed between the sidewalls. The second partition is common to both chambers, and both partitions include respective pluralities of first and second inlet holes sized to meter cooling air therethrough in series between the chambers.

Abstract: According to the invention, turbine blades are provided that have increased functional reliability, which is achieved in that the interior space (8) of the blade body (3) in the region of the suction-side wall (4), the pressure-side wall (5) and the trailing blade edge (7) has a closed steam-cooling system (9) having at least one cooling passage (10, 27, 28). An open cooling system (11) having at least one cooling passage (14, 15) and a plurality of film-cooling holes (22) which pass through the blade body (3) is formed in the region of the leading blade edge (6).

Abstract: In connection with a turbine rotor disk in whose disk grooves air-cooled turbine blades have been inserted, at least two cooling air channels, respectively extending from the same disk front face, terminate in each disk groove. The outlet openings of two cooling air channels in each disk groove preferably lie essentially next to each other in a common sectional plane, which is normal to the disk axis. Because of this it is possible to supply a larger cooling air flow without drastically increasing the weakening, or respectively the stress of the disk in the groove bottom.

Abstract: A gas turbine rotating blade comprises a serpentine passage provided in the blade width direction in plural rows in a blade profile portion except the portion near the trailing edge having a small blade thickness, a steam cooling passage provided around the outer periphery of a platform, an air passage provided in the blade width direction in the vicinity of the trailing edge of blade profile portion, an impingement plate provided under the platform on the inside of the steam cooling passage, slot holes formed so as to branch off from the air passage and be directed to the trailing edge, and slits formed in the platform above the impingement plate. Therefore, the blade profile portion is cooled by the steam passing through the serpentine passage and the air passing through the air passage, so that the cooling construction is not complicated, and cooling is performed effectively. For the platform, the outer periphery thereof is cooled by steam and the inside thereof by air effectively.

Abstract: In a steam cooled type gas turbine, taking the effect of efficiency and power output into consideration, a rear stage is sufficiently cooled by air in place of cooling all of the stages by steam. In a rotor of the gas turbine having a steam cooled moving blade, a moving blade disposed in a front stage is cooled by steam, and a moving blade disposed at the rear stage is cooled by cooling air through a stationary blade disposed in front of the moving blade at the rear stage.

Abstract: In a recovery type steam-cooled gas turbine, a steam passage leading to a moving blade is provided on a central side of a rotor for reducing leakage of the steam. Cooling steam 70 of a low temperature and high pressure is led into moving blades 1, 2 through supply side steam passages 12, 13 and steam passages 15, 16. After being used for cooling, it flows into a cavity 22 through steam passages 16, 17 to be recovered as recovery steam 71 of a high temperature and low pressure at a rotor end 20 through recovery side steam passages 19, 21. The cooling steam 70 of low temperature and high pressure passes on an inner side of the recovery steam 71 in the rotor, hence there are fewer places from where the high pressure steam leaks outside as compared with the prior art, in which the high pressure steam is supplied from the outer side in the rotor, and the leakage amount of steam is reduced.

Abstract: A gas turbine moving blade having a cooling medium flow path constructed so that a concave spherical surface (C) is formed on a inside surface of an end portion of the cooling medium path (B) formed in a blade root portion. A hollow pipe (6) is provided in the cooling medium flow path between the blade root portion and a disc. The hollow pipe has a convex spherical surface (D) at one end thereof which engages the concave spherical surface (C) and the other end of the pipe has a convex spherical surface (F) which engages an inside surface of a cooling medium path (E) formed in the disc. The hollow pipe (6) provides communication between the cooling medium path (B) on the blade side and the cooling medium path (E) on the disc side. A supporting structure (7, 8, 9) is provided for holding the hollow pipe in the communicating position.

Abstract: A coolable double walled structure, for use and exposure in a hot gas flow environment, includes a jet issuing wall, a target wall spaced from the jet issuing wall defining at least one cavity therebetween. The jet issuing wall includes at least one impingement rail having a greater thickness than the jet issuing wall such that a bottom portion of the impingement rail extends within the cavity formed between the jet issuing wall and the target wall. A number of impingement ports are disposed within each impingement rail to provide flow communication through the jet issuing wall to the target wall. Due to the relatively thick dimensions of the impingement rail, the impingement ports disposed within respective impingement rails can be directionalized so as to provide for local cooling needs. In one embodiment, the entry areas of the impingement ports are bell shaped so as to lower inlet pressure loss.

Abstract: The invention relates to a cooled shroud in a gas turbine stationary blade which is able to flow a cooling air in the entire area of an inner shroud for cooling thereof. Three stationary blades are fixed to the inner shroud 2, a cover 13, 14 is provided to form a space 21 and space 22a, 22b and 22c, respectively. The cooling air is introduced through an independent air passage 3A of a leading edge of each stationary blade into the spaces 22a, 22b and 22c and is flown therefrom through a tunnel 18 and air reservoirs 19-2, 19-3 and 19-4 to be blown out of a trailing edge while cooling surfaces of the shrouds and the trailing edges. Also, a portion of the cooling air from the space 22b is flown into the space 21 through a tunnel 11, a leading edge side passage 12 and an endmost tunnel 11 and is then blown out of the trailing edge through a tunnel 18 and an air reservoir 19-1, so that the leading edge portion, the endmost portion and the endmost trailing edge portion are cooled.

Abstract: In a turbo engine's welded rotor includes a number of disks, the radially or quasi-radially extending welding seams are intermediately interrupted by annular cavities through which a cooling medium flows, and these cavities are surrounded by a circumferentially extending insert ring. This insures the flow of the cooling medium through the entire rotor without reducing the strength of the welding bond.

Abstract: A gas-turbine blade of solid construction and having cooling passages is specified, which cooling passages run in its interior and through which the cooling medium flows essentially from the blade root through the gas-turbine blade (100), these cooling passages (20) having a relatively small cross section in relation to the blade thickness and being arranged so as to run close to the lateral surface (100.1) of the gas-turbine blade (100), and the cooling-medium feed and discharge being effected in the base (60) of the gas-turbine blade (100).

Abstract: A cooling arrangement for a bladed rotor in a gas turbine engine, wherein each of the blades includes cooling air passages and a cover with curved fins is mounted adjacent to but connected to the rotor and spaced apart slightly from the rotor disc to form a passageway for the cooling fluid. The cooling arrangement includes a tapered, conically shaped inlet formed in the cooling passage which then diverges to form a diffuser near the outer end of the passageway. The cover includes an enlarged inner portion and a thin outer wall portion beyond the free ring diameter. A hammerhead is formed at the outer periphery of the cover whereby the hammerhead will move closer to the disc in response to centrifugal forces, thus sealing the passage.

Abstract: A cooling medium path structure for cooling a gas turbine blade. The structure includes a disk-side cooling medium path, a blade-side cooling medium path formed at the root portion of the blade and a delivery block disposed between the two cooling medium paths so as to establish communication therebetween. The delivery block is provided with an elastic engaging section which comes into elastic and line-contact with the disk-side cooling medium path and the blade-side cooling medium path. Thus, the sealing property of the contact portions of the structure is secured so as to allow a cooling medium to be supplied without leaking from the cooling medium paths. The heat of the cooling medium, generated as a result of cooling the high-temperature portion of the gas turbine, can be recovered so as to make the best use of the heated medium.

Abstract: The apparatus is a heat pipe with an internal, multiple chamber evaporator for cooling a turbine engine stator vane. The evaporator comprises leading edge, middle, and trailing edge chambers within the stator vane, with the chambers defined by structural support ribs. Each chamber is constructed with a continuous fine pore metal powder wick coating the internal surfaces of the chamber and enclosing the chamber's central vapor space, except the wick at the very trailing edge of the vane is formed by screen embedded in the adjacent powder wick. The evaporator chambers have capillary arteries which extend through the adiabatic section of the heat pipe and terminate in a condenser wick within a heat sink structure exposed to cooler air. A capillary artery also interconnects the wick of the trailing edge chamber to the wick of the middle chamber.

Abstract: A gas turbine moving blade steam cooling system in which leakage of steam for cooling a moving blade is prevented and thermal stress at a blade root end portion is mitigated. Each end portion of a blade root portion (3) of the moving blade (1) is projected so as to form a projection portions (4a, 4b). A steam passage 5 is provided between the projection portions (4a, 4b) and a steam supply port 5a and a steam recovery port (5b) are provided downwardly with the respect to the steam passage 5. The steam supply port (5a) connects to a steam supply passage (20) and the steam recovery port (5b) connects to a steam recovery passage (21). Steam is supplied from the steam supply port (5a) into a blade interior and is recovered through the steam recovery port (5b). Side surface seal plates (6, 7 and 8) are provided for preventing steam leakage.

Abstract: In connection with a turbine rotating disk with disk grooves, constituted by disk fingers, for receiving turbine blades, as well as with measures for guiding a cooling air flow from a chamber located in front of the disk to one behind the disk, a cooling air transfer channel is provided in at least some of the disk fingers, which respectively starts at the front disk face, and makes a transition into a cooling air exhaust channel, also essentially extending in the radial direction in the disk finger, whose outlet opening on the rear disk face side lies closer to the disk axis than the disk ring section, which has the disk grooves and is widened in the disk axial directions. By means of this it is possible to convey a larger cooling air flow which, in an advantageous manner, indirectly aids the cooling of the disk in the area of the disk grooves.

Abstract: A gas turbine stationary blade, having a simple structure in which sufficient cooling is achieved and the drop in pressure of cooling vapor is decreased so that the turbine efficiency is prevented from lowering. The shape of a vapor passage is simplified to prevent the drop in pressure because an outer shroud (3) of the stationary blade and a blade unit (2) are cooled with vapor, while an inner shroud (4) is cooled with the air supplied from another system.

Abstract: An evaporatively cooled rotor for a gas turbine engine. Each rotor defines an internal cavity which includes a vaporization section that corresponds generally to the blade section of the rotor and a condensing section that corresponds generally to the hub section of the rotor. A radial array of circumferentially disposed capture shelves is provided in the vaporization section for capturing cooling fluid contained within the internal cavity and flowing radially outward under the centrifugal field generated during rotation of the rotor. A barrier disposed along the inner surface of the rotor wall in the condensing section slows or temporarily stops the flow of cooling fluid prior to reaching the vaporization section and a perforated baffle attached to the capture shelves prevents cooling fluid from splashing out of the shelves.

Abstract: A turbine cooling apparatus comprising a turbine disk having a plurality of moving blades, a torque tube coupled coaxially to one surface side of the turbine disk and having a thick stepped inner wall portion in the central portion thereof, an air separator fitted on the torque tube with the inner surface thereof in contact with the outer surface of the torque tube so that a passage through which cooling air is supplied to the moving blades via the turbine disk is defined between the air separator and the torque tube, and a torque tube cooling hollow portion provided along and in the vicinity of the outer surface of the thick stepped wall portion of the torque tube.