APPARATUS STATE ESTIMATION DEVICE, APPARATUS STATE ESTIMATION METHOD AND PROGRAM

A state quantity acquisition unit acquires the state quantity of an apparatus to be estimated that includes the temperature of the apparatus to be estimated. A load specification unit specifies the load history of the apparatus to be estimated on the basis of the state quantity. A remaining life calculation unit calculates a parameter that pertains to the remaining life of the apparatus to be estimated for each of multiple types of degradation on the basis of the load history specified by the load specification unit.

PLANT ANALYSIS DEVICE, PLANT ANALYSIS METHOD, AND PROGRAM

A function-of-state acquisition unit acquires the function of state of a turbine, including turbine temperature. A variable calculation unit calculates past variables pertaining to the history of the function of state. A time calculation unit calculates the time period of possible operation of a turbine in an over-firing operation on the basis of the calculated past variables, and a past variable that corresponds to the design life of the turbine.

PLANT ANALYSIS DEVICE, PLANT ANALYSIS METHOD, AND PROGRAM

A function-of-state acquisition unit acquires the function of state of a turbine. A time calculation unit calculates the time period of possible operation of a turbine in an over-firing operation, on the basis of the function of state and the design life of the turbine. A distance calculation unit calculates the Mahalanobis distance on the basis of the function of state. A determination unit determines whether a turbine over-firing operation is possible by using the Mahalanobis distance and the time period of possible operation.

GAS TURBINE COOLING SYSTEM, GAS TURBINE PLANT EQUIPPED WITH THE SAME AND METHOD OF COOLING HIGH-TEMPERATURE SECTION OF GAS TURBINE

A gas turbine cooling system includes an organic Rankine cycle in which a cycle medium repeatedly circulates through condensation and evaporation, an air-extraction line configured to extract compressed air from a compressor of a gas turbine, a cooling apparatus for cooling the compressed air while heating and evaporating the cycle medium condensed in the organic Rankine cycle using heat of the compressed air passing through the air-extraction line, and a cooling air line configured to guide the compressed air cooled by the cooling apparatus to a high temperature section of the gas turbine. 1. A gas turbine cooling system configured to cool a high temperature section in contact with a combusted gas in a gas turbine using compressed air extracted from a compressor of the gas turbine , the gas turbine cooling system comprising:an organic Rankine cycle configured to repeatedly circulate a cycle medium through condensation and evaporation in a cycle including the turbine and drive the turbine using the evaporated cycle medium;an air-extraction line configured to extract the compressed air from the compressor;a cooling apparatus configured to cool the compressed air while heating and evaporating the cycle medium condensed in the organic Rankine cycle using heat of the compressed air passing through the air-extraction line; anda cooling air line configured to guide the compressed air cooled by the cooling apparatus to the high temperature section.2. The gas turbine cooling system according to claim 1 , wherein the cooling apparatus has an indirect type heat exchange apparatus configured to exchange heat between the compressed air passing through the air-extraction line and the cycle medium condensed in the organic Rankine cycle via an intermediate medium.3. The gas turbine cooling system according to claim 2 , wherein the indirect type heat exchange apparatus has:a first heat exchanger configured to exchange heat between the compressed air passing through the air- ...

GAS TURBINE COOLING SYSTEM, GAS TURBINE PLANT EQUIPPED WITH SAME, AND METHOD FOR COOLING HIGH-TEMPERATURE SECTION OF GAS TURBINE

This gas turbine cooling system (40) is equipped with: an organic Rankine cycle (50) in which a cycle medium (CM) is circulated while being condensed and evaporated repeatedly; an extraction line (41) for extracting compressed air (CA) from a compressor (11) of a gas turbine (10); a cooling device (60) that, by utilizing the heat of the compressed air (CA) that came in through the extraction line (41), heats and evaporates the cycle medium (CM) that has been condensed in the organic Rankine cycle (50) and cools the compressed air (CA) at the same time; and a cooled air line (42) that introduces the compressed air (CA) that has been cooled by the cooling device (60) to a high-temperature section of the gas turbine (10).

PLANT ANALYZER, PLANT ANALYSIS METHOD, AND PROGRAM THEREOF

A state quantity acquiring unit is configured to acquire a state quantity of a turbine. A time calculation unit is configured to calculate an operable time at an over firing operation of the turbine based on a design life of the turbine and a state quantity. A distance calculation unit configured to calculate a Mahalanobis distance based on the state quantity. A determination unit configured to determine whether or not the over firing operation of the turbine is possible by the Mahalanobis distance and the operable time. 1. A plant analyzer comprising:a state quantity acquiring unit configured to acquire a state quantity of a turbine;a time calculation unit configured to calculate an operable time at an over firing operation of the turbine based on a design life of the turbine and the state quantity;a distance calculation unit configured to calculate a Mahalanobis distance based on the state quantity; anda determination unit configured to determine whether or not the over firing operation of the turbine is possible by the Mahalanobis distance and the operable time.2. The plant analyzer according to claim 1 ,wherein the determination unit configured to determine whether or not a power selling price is less than a predetermined threshold value and whether or not the over firing operation of the turbine is possible by the Mahalanobis distance and the operable time.3. The plant analyzer according to claim 1 , further comprising:a variable calculation unit configured to calculate a history variable with respect to a history of the state quantity; andwherein the time calculation unit is configured to calculate the operable time at the over firing operation of the turbine based on a history variable corresponding to the design life of the turbine and the calculated history variable.4. The plant analyzer according to claim 1 ,wherein the time calculation unit is configured to calculate the operable time to prevent the turbine from reaching the product life until the turbine ...

PLANT ANALYZER, PLANT ANALYSIS METHOD, AND PROGRAM THEREOF

A state quantity acquiring unit is configured to acquire a state quantity of a turbine including a temperature of the turbine. A variable calculation unit is configured to calculate a history variable with respect to a history of the state quantity. A time calculation unit is configured to calculate the operable time at an over firing operation of the turbine based on a history variable corresponding to a design life of the turbine and a calculated history variable. 1. A plant analyzer comprising:a state quantity acquiring unit configured to acquire a state quantity of a turbine including a temperature of the turbine;a variable calculation unit configured to calculate a history variable with respect to a history of the state quantity; anda time calculation unit configured to calculate an operable time at an over firing operation of the turbine based on a history variable corresponding to a design life of the turbine and a calculated history variable.2. The plant analyzer according to claim 1 , further comprising:a first determination unit configured to determine whether or not the over firing operation of the turbine is possible based on the operable time calculated by the time calculation unit.3. The plant analyzer according to claim 1 ,wherein the time calculation unit is configured to calculate the operable time to prevent the turbine from reaching a product life until an inspection timing of the turbine.4. The plant analyzer according to claim 1 , further comprising:an inspection timing determination unit configured to determine the inspection timing of the turbine based on the operable time calculated by the time calculation unit.5. The plant analyzer according to claim 1 , further comprising:a distance calculation unit configured to calculate a Mahalanobis distance based on the state quantity; anda second determination unit configured to determine whether or not the over firing operation of the turbine is possible by the Mahalanobis distance.6. The plant ...

INTAKE AIR COOLING SYSTEM

An intake air cooling system for a gas turbine is provided with: an intake duct for leading intake air from an intake-air inlet to a compressor of the gas turbine, the intake duct having a vertical duct and a manifold part disposed on a downstream side of the vertical duct; a cooling part provided in the intake duct to cool the intake air by heat exchange with a cooling medium which is introduced from an outside; a filter part provided on an inlet side of the manifold part to remove impurities contained in the intake air introduced through the vertical duct; and a drain catcher constituted by a gutter member provided immediately above the filter part along inner wall surfaces of the vertical duct, the drain catcher being configured to collect drain water flowing along the inner wall surfaces of the vertical duct. 1. An intake air cooling system for a gas turbine , the system comprising:an intake duct configured to lead intake air taken in from an intake-air inlet to a compressor of the gas turbine, the intake duct having a vertical duct and a manifold part disposed on a downstream side of the vertical duct;a cooling part provided in the intake duct and configured to cool the intake air by heat exchange with a cooling medium which is introduced from an outside;a filter part provided on an inlet side of the manifold part and configured to remove impurities contained in the intake air introduced through the vertical duct; anda drain catcher constituted by a gutter member provided immediately above the filter part along an inner wall surface of the vertical duct, the drain catcher being configured to collect drain water flowing along the inner wall surface of the vertical duct.2. The intake air cooling system according to claim 1 ,wherein the vertical duct has a rectangular cross sectional shape having a long side and a short side,wherein the inner wall surface of the vertical duct comprises a first inner wall surface along the long side and a second inner wall surface ...

INTAKE AIR COOLING SYSTEM

An intake air cooling system for a gas turbine is provided with: an intake duct for leading intake air taken in from an intake-air inlet to a compressor of the gas turbine; a cooling part provided in the intake duct and configured to cool the intake air by heat exchange with a cooling medium which is introduced from an outside; a protruding step part formed in a convex shape protruding from bottom surfaces of the intake duct disposed on a downstream side of the cooling part; and at least one drain hole formed in the bottom surfaces of the intake duct disposed on the downstream side of the cooling part so as to discharge drain water, generated on a surface of the cooling part and dropping from the surface, to an outside of the intake duct. 1. An intake air cooling system for a gas turbine , the system comprising:an intake duct configured to lead intake air taken in from an intake-air inlet to a compressor of the gas turbine;a cooling part provided in the intake duct and configured to cool the intake air by heat exchange with a cooling medium which is introduced from an outside;a protruding step part formed in a convex shape protruding from a bottom surface of the intake duct disposed on a downstream side of the cooling part; andat least one drain hole formed in the bottom surface of the intake duct disposed on the downstream side of the cooling part, the at least one drain hole being configured to discharge drain water, generated on a surface of the cooling part and dropping from the surface, to an outside of the intake duct.2. The intake air cooling system according to claim 1 ,wherein the at least one drain hole is formed in the bottom surface of the intake duct which is disposed between the cooling part and the protruding step part.3. The intake air cooling system according to claim 2 ,wherein a silencer is provided in the protruding step part to muffle noise generated when taking in the air, andwherein the at least one drain hole is further formed in a bottom ...

TURBINE BLADE, GAS TURBINE, INTERMEDIATE PRODUCT OF TURBINE BLADE, AND METHOD OF MANUFACTURING TURBINE BLADE

Provided are a turbine blade, a gas turbine, an intermediate product of the turbine blade, and a method of manufacturing the turbine blade. This turbine blade has a blade body having hollow shape, cavities provided inside the blade body, and a cooling passage that opens from the cavities to the rear end portion of the blade body. The cooling passage includes: a first passage provided on the third cavity side and having a width that becomes narrower from the third cavity side toward the rear end portion of the blade body; and a second passage provided on the rear end portion side of the blade body and having a width that is constant from the third cavity side toward the rear end portion of the blade body. 1. A turbine blade , comprising:a blade body having a hollow shape;a cavity provided in the interior of the blade body; anda cooling passage that is open from the cavity to a rear end portion of the blade body, whereinthe cooling passage includesa first passage provided on the cavity side and having a width that becomes narrower from the cavity side toward the rear end portion of the blade body, anda second passage provided on the rear end portion side of the blade body and having a width that is constant from the cavity side toward the rear end portion of the blade body.2. The turbine blade according to claim 1 , wherein a flow rate adjustment mechanism is provided in the second passage.3. The turbine blade according to claim 2 , wherein the flow rate adjustment mechanism has a plurality of columns provided at predetermined intervals in the second passage of the blade body along a longitudinal direction.4. The turbine blade according to claim 1 , wherein the cooling passage includes a third passage having a constant width from the cavity side toward the rear end portion of the blade body claim 1 , a first end portion of the third passage communicating with the second passage and a second end portion of the third passage opening at the rear end portion of the blade ...

DIAGNOSTIC DEVICE, DIAGNOSTIC METHOD, AND PROGRAM

A diagnostic device includes a storage unit that stores first information including a first abnormal event which occurred in the past in a plant, a first attribution event that is a cause of the first abnormal event, and a first occurrence probability of the first attribution event, in which a causal relationship between the first abnormal and attribution events is indicated by a tree structure, and second information including a second abnormal event which is supposed to occur in the plant but does not occur yet, a second attribution event that is a cause of the second abnormal event, and a second occurrence probability of the second attribution event, in which a causal relationship between the second abnormal and attribution events is indicated by a tree structure; and an estimation unit that estimates the cause of the sign of the abnormality, based on the first and second information. 1. A diagnostic device which diagnoses a sign and a cause of abnormality of a plant , comprising:a reception unit that receives from the plant, operation data indicating an operation state of the plant;a storage unit that stores first information including a first abnormal event which occurred in the past in the plant, at least one or more first attribution events that are causes of the first abnormal event, and a first occurrence probability that is an occurrence probability of the first attribution event, in which a causal relationship between the first abnormal event and the first attribution event is indicated by a tree structure, and second information including a second abnormal event which is supposed to occur in the plant but does not occur yet, at least one or more second attribution events that are causes of the second abnormal event, and a second occurrence probability that is an occurrence probability of the second attribution event, in which a causal relationship between the second abnormal event and the second attribution event is indicated by a tree structure;a ...

Gas turbine rotor blade

Gas turbine rotor blade

Gas turbine blade

(57)【要約】 【課題】 本発明は、ガスタービン翼列の後段側に設置 され、内部から冷却を行うようにした、薄肉長大された ガスタービン翼に関する。翼プロフィル部の翼付根部に 隣接する部分に空洞を設けて、この空洞を流れる冷却空 気により内部から冷却するようにしたガスタービン動翼 を先に提案したが、冷却効果が不十分であった。本発明 は、このような不具合を解消できるガスタービン翼を提 供することを課題とする。 【解決手段】 本発明のガスタービン翼は、ハブ部ばか りでなく、ハブ部の外周側に設けられている冷却通路の 少くとも一部を、ハブ部に形成されている空洞と略相似 の横断面形状にされ、内壁から内部を流れる冷却媒体中 に突出するピンフィンを設けた空洞部にした。これによ り、冷却効率がさらに向上して、クリープ寿命がさらに 延びるとともに、ねじり剛性が増したガスタービン翼と することができる。また、加工の難しかったマルチホー ル等の冷却通路の形成が容易になる。

Seal device of gas turbine stationary blade

(57)【要約】 【課題】 ガスタービン静翼のシール装置に関し、内側 シュラウドのシール圧力を高め、シール効果を増す。 【解決手段】 翼環５０には遮熱環３２ａ，３２ｂを介 して外側シュラウド３２が取付けられ、翼環５０には空 気穴１，５１を設け、１は空間５３へ、５１はシールチ ューブ２へ連通する。シールチューブ２は先端部３が空 気穴５１内へ挿入され、バネ６がチューブ２の突起部４ と空気穴５１の係止部５との間に配設されてシールチュ ーブ２を着脱可能に固定する。冷却空気５４は一方では 空気穴１より空間５３へ入り、シュラウドと静翼３１内 を冷却し、後縁より放出し、他方ではキャビティ３６に 入り、キャビティ３６内はチューブ２が空間５３とは独 立しているので圧損を受けず高圧を維持する。キャビテ ィ３６からの高圧空気はＳ１，Ｓ２のようにシール部４ ０ａ，４０ｂから流出し、燃焼ガス通路からの高温ガス 侵入を防止する。

Moving blade and gas turbine using the same

In a gas turbine having a plurality of moving blades provided on a rotary shaft in a circumferentially adjoining condition, a seal pin is provided in a spacing between the shanks of the adjacent moving blades for preventing leakage of cooling air from a blade root portion side to an airfoil side; an arcuately depressed portion is formed on the shank of each of the moving blades; and vibration of each of the moving blades is suppressed in such a manner that the seal pin serves as a spring system while the airfoil portion, the platform, the shank, and the blade root portion serve as a mass system.

Division wall and shroud of gas turbine

The division wall is made up of a plurality of division wall sections forming a passage wall of high temperature gas which are connected in the direction of arrangement of blades to form a wall surface having a roughly circular cross section as a whole, a gas flow restricting structure for preventing high temperature gas from passing through a gap formed at a connecting portion between the division wall sections in the flow direction of the high temperature gas from the opening on the upstream side of the high temperature gas in the gap, or a gas flow restricting structure for preventing the high temperature gas from being embraced in the gap, for example, a sealing member formed into a prism having a T-shape cross section as a whole composed of a plane portion as a sealing portion and a projected portion for filling the gap is provided.

Sealing device for gas turbine stator blades

A sealing device for a gas turbine stator blade, in which a sealing pressure of an inner shroud is raised to enhance a sealing effect. An outer shroud 32 is mounted through heat insulating rings 32a and 32b on a blade ring 50, and this blade ring 50 has air holes 1 and 51, of which the air hole 1 communicates with a space 53 whereas the air hole 51 communicates with a seal tube 2. This seal tube 2 is inserted at its leading end portion 3 into the 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 fix the seal tube 2 removably. Cooling air 54 flows on one side from the air hole 1 into the space 53 to cool the shrouds and the inside of a stator blade 31 until it is released from the trailing edge, and on the other hand into a cavity 36 so that the inside of the cavity 36 is kept at a high pressure without receiving a pressure loss because the tube 2 is independent of the space 53. The high pressure air from the cavity 36 flows out from seal portions 40a and 40b, as indicated by arrows S1 and S2, to prevent the inflow of the hot gas from a combustion gas passage.

Turbuletor for gaz turbine cooling blades

Leading edge portion turbulators in gas turbine cooled blade are improved so as to enhance cooling performance. Rounded tip portion of leading edge portion cooling passage 3 is approximated by triangular cooling passage 1 having orthogonal turbulators 11, 12 and smoothly curved rear portion thereof of the leading edge portion cooling passage 3 is approximated by square cooling passage 2 having oblique turbulators 13, 14, thus the leading edge portion cooling passage 3 is formed so as to have combination of orthogonal turbulators 21 of the rounded tip portion and oblique turbulators 22, 23 of the smoothly curved rear portion thereof. These rounded portion and rear portion thereof are provided with the respective turbulators having excellent heat transfer characteristics and cooling performance of the leading edge portion is enhanced.

Gas turbine moving blade

A gas turbine moving blade in which the moving blade operating in a high temperature operating gas is cooled from its interior by steam and the steam after being used for the cooling is recovered, by use of a simple structure. A supply side passage (10), through which steam (6) is supplied directly into a cooling passage (51) of a blade leading edge side (11), is provided within a blade root portion (2) and a pocket (19). One end of the pocket is connected to the supply side passage and the other end of which is connected to a cooling passage (52) of a blade trailing edge side (12). The pocket is covered with a plate, which extends in a blade chord direction on an outer peripheral side face of the blade root portion between a blade platform and a rotor plate. Thus, efficiency of the gas turbine can be effectively enhanced.

Cooling structure for a gas turbine

Gas turbine stationary blade

Gas turbine stationary blade is improved in shapes of blade leading edge and fillets, in supporting of inserts and in blowing of cooling air, so that blade cooling efficiency is enhanced, insert supporting structure is simplified and clogging of cooling holes is prevented, thus reliability of the stationary blade is enhanced. Passages (23, 24) are provided in stationary blade (10). Front insert (2) is provided in the passage (23) and rear insert (5) in the passage (5) to be supported at two points of insert supporting portions (3a, 3b), (6a, 6b), respectively. Projection (1) is provided at blade leading edge so that portion where thermal load is high is made smaller in size and number of rows of cooling holes (11a) in this portion is lessened. Air blowing holes (4b) on dorsal side of the front insert (2) and film cooling holes (12) of the blade have diameters larger than those of other holes, so that dusts in cooling air are caused to flow out to prevent clogging of the holes. Curved surface of the blade leading edge is formed to elliptical curve, so that flow of the cooling air is made smooth. Curved surfaces of fillets are also formed to elliptical curve and thermal stress concentration is avoided.

Gas turbine moving blade

In a gas turbine which is constructed such that there is provided a shroud at a terminal end of a blade and cooling air is led into the blade to flow through a multiplicity of cooling holes provided in the blade to be then led into the shroud and is flown out of a multiplicity of cooling passages provided in the shroud, a uniform flow of cooling air in both side portions of the shroud and thus a more uniform cooling of the shroud with a facilitated flow control of the cooling air is achieved by the following measures: The multiplicity of cooling holes of the blade (221) and the multiplicity of cooling passages (70-73,74-77) of the shroud (92) are sectioned into two groups, respectively. There are formed in the shroud (92) two cavities (80,81), each connecting to each one of the groups of cooling holes of the blade as well as connecting to each one of the groups of cooling passages of the shroud (92). The groups of the cooling passages of the shroud (92) are arranged so that the cooling air flowing therethrough is flown out of mutually opposing side portions of the shroud (92).

Cooling structure for a gas turbine

In a cooling structure for a gas turbine, cooling air diffusion holes are formed from inner surface to outer surface of a platform so as to open from high pressure side blade surface of a moving blade offset in a direction toward low pressure side blade surface of adjacent moving blade confronting the high pressure side blade surface, in a direction of primary flow.

Jacket ring for gas turbines

Gas turbine moving blade

A gas turbine moving blade is provided which prevents occurrence of cracks caused by thermal stresses due to temperature differences between a blade and a platform while the gas turbine is being stopped. During steady operation time of the moving blade, cooling air enters cooling passages to flow through other cooling passages for cooling the blade, and then to flow out of the blade. A recessed portion having a smooth curved surface is provided in the platform near a blade fitting portion on the blade trailing edge side. A fillet of the blade fitting portion on the blade trailing edge side has a curved surface with curvature larger than that of a conventional blade fitting portion. A hub slot below the fillet, for blowing air, has a cross sectional area larger than other slots of the blade trailing edge. A thermal barrier coating is applied to the blade surface. By the above construction, thermal stresses due to temperature differences between the blade and the platform during gas turbine stoppage are made smaller and occurrence of cracks is prevented.

Gas turbine split ring

In the gas turbine split ring, on an outer peripheral surface 1b between two cabin attachment flanges, a circumferential rib which extends in the circumferential direction and an axial rib which extends in the axial direction and has a height taller than that of the circumferential rib are, respectively, formed in plural lines, so that it is possible to suppress heat deformation in the axial direction which largely contributes to reduction of the tip clearance compared to head deformation in the circumferential direction more efficiently.

Cooling steam supply/discharge structure for gas turbine

(57)【要約】 【課題】 ガスタービンにおいて、動翼冷却用の蒸気の 給排部の構造をコンパクトにする。 【解決手段】 ガスタービンの冷却用蒸気給排部構造 は、ロータ１を囲んで設けられた軸受箱３の一端に固定 された本体ケース５と、ロータ１内を延びる蒸気導入外 管１１の端部１１ａ外面のラビリンスシール用凹凸面１ ３と、蒸気導入外管１の内部を延びる導入内管１５の端 部１５ａ外面のラビリンスシール用凹凸面１７と、凹凸 面１３，１７をそれぞれ取り囲むラビリンスリング１９ と、ラビリンスリング１９を保持し本体ケース５に摺動 自在に設けられたシールリテーナ２１と、本体ケース５ の端部にロータ１の端面に臨んで設けられた伸差センサ ２５と、伸差センサ２５とシールリテーナ２１との間に 介装されたサーボ機構３０とを有する。

Cooled gas turbine blade with turbulators

Cooled gas turbine blade

The present invention relates to a cooled blade of a gas turbine, and in particular an object of the present invention is to reduce thermal stress around air-transpiration holes (4) provided for shower-head cooling of a leading edge portion of a blade. A number of air-transpiration holes (4) are provided at a leading edge portion (2) of a cooled blade (1). Cooling air flowing through a cooling air passage (15) formed inside of the blade blowout to the blade surface of the leading edge portion (2) of the cooled blade (1) by way of the air-transpiration holes (4), to thereby shower-head cool the surface of the leading edge portion (2). In the conventional blade, the air-transpiration holes (14) are formed obliquely to the leading edge portion (2), whereby acute-angled portions (30) are formed at inlet/outlet ports of the air-transpiration holes (14) in the leading edge portion (2), and thus cracks due to thermal stress develop around the air-transpiration holes (14). In contrast, according to the present invention, the air-transpiration holes (4) are formed so as to extend substantially orthogonal to the leading edge surface of the cooled blade (1) such that the acute-angled portions are eliminated, whereby thermal stress can be reduced and thus the cracks can be prevented.

Cooling for the blade tips of a turbine

Gas turbine cooled blade turbulators

Leading edge portion turbulators in gas turbine cooled blade are improved so as to enhance cooling performance. Rounded tip portion of leading edge portion cooling passage 3 is approximated by triangular cooling passage 1 having orthogonal turbulators 11, 12 and smoothly curved rear portion thereof of the leading edge portion cooling passage 3 is approximated by square cooling passage 2 having oblique turbulators 13, 14, thus the leading edge portion cooling passage 3 is formed so as to have combination of orthogonal turbulators 21 of the rounded tip portion and oblique turbulators 22, 23 of the smoothly curved rear portion thereof. These rounded portion and rear portion thereof are provided with the respective turbulators having excellent heat transfer characteristics and cooling performance of the leading edge portion is enhanced.

Turbine blade and gas turbine

A turbine blade applicable to a gas turbine has a turbine blade body having film cooling holes, the interior space of which is partitioned into two cavities by a rib. Hollow inserts each having impingement holes are respectively arranged in the cavities to form cooling spaces therebetween. Communication is ensured between the cavities by a communication means, so that the impingement cooling is interrupted with respect to the prescribed side having a good heat transmission in the turbine blade body. A partition wall is further arranged between the rib and the insert arranged in the trailing-edge side, thus providing a separation between the cooling spaces respectively arranged in the rear side and front side. Thus, it is possible to noticeably reduce the amount of cooling air in the turbine blade body; and it is possible to reduce temperature differences entirely over the turbine blade body as small as possible.

Blade cooling air supplying system of gas turbine

In the present invention, an air pipe extends through a stationary blade between outer and inner shrouds. Further, an air passage is directed to a lower portion of the stationary blade and is communicated with the air pipe so that a serpentine cooling passage is formed. The air enters a cavity from the air passage and is discharged to a gas passage through an air hole, a passage and a seal. Thus, the cavity is sealed at a high pressure. Cooling air is supplied from the air passage to a rotating blade through a cooling air hole, a cooling air chamber, a radial hole and a lower portion of a platform. The stationary blade is cooled by the air through the air passage. The cooling air can be supplied to the rotating blade at a low temperature and a high pressure as they are. Accordingly, the air can be also supplied to the rotating blade when a rotor is cooled by vapor.

Gas turbine

12 One object of the present invention is to provide a gas turbine where cooli ng failure attributable to the occurrence of a horseshoe vortex produced in the vicinity of the stationary blades of the turbine, can be prevented. In order to achieve the object, the present invention provides a gas turbine comprising moving blades provided o n a rotor side which rotate together with the rotor, and stationary blades provided on a stationary side which cover the periphery of the moving blades and form a combustion ga s how path in the interior, and which are arranged alternately with the moving blades i n the rotation axis direction of the rotor, and where the stationary blades have a blade portion arranged inside the combustion gas flow path, an outside shroud provided on an outer peripheral end side of the blade portion, and an inside shroud provided on an inner peripheral end side of the blade portion, in one or both of the outside shroud and the insi de shroud, corresponding to a leading edge of the blade portion, there is provided a first cooling air flow path which blows out cooling air into the combustion gas flow path, fro m downstream to upstream in the flow direction of the combustion gas.

Gas turbine blade

Ring segment of gas turbine

An object of the present invention is to provide a ring segment of a gas turbine in which the temperature is maintained low, damage due to high temperature oxidation is prevented, and high temperature deformation is prevented. In order to achieve the object, the present invention provides a ring segment of a gas turbine which comprises a blade ring, a main shaft and moving blades comprising a plurality of individual units which define an annular form by being arranged around the peripheral direction of the main shaft, and disposed so that its inner peripheral surface is maintained at a constant distance from the tips of the moving blades, wherein grooves which extend along the axial direction of the main shaft of the turbine are formed upon of the individual units so as mutually to confront one another; a seal plate which is inserted into each mutually confronting pair of the grooves so as to connect together the adjacent pair of individual units; and contact surfaces which are formed at positions more radially inward than the seal plates, which extend in the axial direction and the peripheral direction and which mutually contact one another.

Rotary blade of gas turbine

A wound through passage is provided for cooling a blade profile area (2), with steam (S) being allowed to flow into it. The through passage is connected to a steam feed aperture and a steam outlet aperture which are formed at a blade foot area. Several flow paths (91-96) are arranged in the direction of the blade (1) width. A cooling passage for steam for cooling a platform (3) is provided and a flow path is formed in connection with the wound through passage around the outer edge of the platform of the rotary blade. Slots are so formed and are so inclined that air from the upper surface of the platform is emitted into the main gas flow. The slots are directed from the widened side of the blade profile area in the rotary direction of the rotating blade. Also air is emitted from the rear side of the blade profile area in the direction of the flow of the main gas current in the central area of the blade profile area.

Gas turbine cooling blade

A number of air-transpiration holes ( 4 ) are provided at a leading edge portion ( 2 ) of a cooled blade ( 1 ). Cooling air flowing through a cooling air passage ( 15 ) formed inside of the blade blows out to the blade surface of the leading edge portion ( 2 ) of the cooled blade ( 1 ) by way of the air-transpiration holes ( 4 ), to thereby shower-head cool the surface of the leading edge portion ( 2 ). The air-transpiration holes ( 4 ) are formed so as to extend substantially orthogonal to the leading edge surface of the cooled blade ( 1 ) such that the acute-angled portions are eliminated, whereby thermal stress is reduced and cracking is prevented.

Gas turbine shroud

In gas turbine split rings (1a, 1b), end faces (3a-1, 3b-1) having bent surfaces are formed in the flanges (4a, 4b). Adjoining split rings (1a, 1b) are coupled together with a groove (5-1) therebetween to form a cylindrical split ring. Notches (26a, 26b) are formed in the flanges (4a, 4b). These notches (26a, 26b) are sealed by inserting a seal plate (25) into the notches (26a, 26b) of adjoining split rings. A hole (2) for passing cooling air is drilled obliquely in the flange (4a). Cooling air (100) is allowed to flow out along the direction of rotation (of the turbine) (R). This cooling air (100) cools the outlet of the groove (5-1) due to the effect of film cooling. Because of such cooling, high temperature gas is prevented from staying in this area, cooling effect is enhanced, and hence burning of the end portions can be prevented.

Gas turbine stationary blade

In cooling a gas turbine stationary blade, steam and air are used as cooling media, the steam is recovered surely without leakage and used effectively, and the amount of air required for cooling is decreased to provide a margin for combustion air, by which the gas turbine efficiency is improved. A steam cooling section is provided at the rear from the blade leading edge, and an air cooling section is provided at the blade trailing edge. The steam cooling is effected by cooling the blade by the cooling steam flowing in the serpentine flow path having turbulators after impingement cooling of an outside shroud and by impingement-cooling an inside shroud during the cooling process, the cooling steam being led to a recovery section from the outside shroud. On the other hand, the air cooling section consists of an air flow path extending from the outside shroud to the inside shroud and slot cooling at the blade trailing edge. Thus, the stationary blade is cooled by both of the steam cooling section and air cooling section.

Jacket ring for gas turbines

Cooled shroud of gas turbine stationary blade

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 endmosttunnel 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. By use of the independent air passage, the cooling air is introduced to cool the entire area of the shroud.

SEAL STRUCTURE FOR GAS TURBINES

Seal of a turbine shroud ring

Gas turbine stationary blade double cross type seal device

Seal plates for gas turbine stationary blade inner shrouds are made in double cross type seal structure with view to enhance sealing ability. Seal plates 1, 2 are mutually lapped and disposed in turbine axial direction between inner shrouds 12 of stationary blades 11. End portion seal plate 5 is lapped on end portion of the seal plate 2 and end portion seal plate 6 is lapped under end portion of the seal plate 1. All these seal plates are fitted with their side end portions being inserted into groove 9a provided in the inner shrouds 12. Seal plates 3, 4 and seal plates 7, 8 engaged with the seal plates 3, 4 are also fitted between flange portions of the inner shrouds 12 with their side end portions being inserted into grooves 10a, 10b and grooves 9b, 9c, respectively. All the seal plates 1 to 8 are fitted between mutually opposing inner shrouds in turbine circumferential direction so as to cover cavity 24 between mutually adjacent shrouds, so that gaps between engaged portions of each seal plates and between the seal plates and seal ring support ring 13 are eliminated, thereby seal air 20 is prevented from leaking from the cavity 24.

A method for managing lifespans of high temperature gas turbine components

The lifespan of the high temperature component being evaluated is set within a component life limit that is determined based on field data giving a correlation between operating time and operating cycles for the states of fatigue from past high temperature components. An independent claim is also included for stored software.

ROTOR FOR GAS TURBINES

Sealing apparatus for gas turbine

A seal ring (1) securing inner shroud members (32) of stationary blades (31) is provided with arm portions (2, 3) projecting along lower surfaces of end portions of the inner shroud members (32). Honeycomb seals (4a, 4b) are mounted on the arm portions (2, 3), respectively. The honeycomb seal (4a) is disposed opposite fins (11a) provided on a rotor arm portion (11) of a platform (22) of a moving blade (21) so that a predetermined clearance (t) can be maintained between the honeycomb seal and the fins. On the other hand, the honeycomb seal (4b) is disposed opposite fins (14b) provided on a seal portion (14a) of a sealing plate (14) of the moving blade (21) so that a predetermined clearance (t) can be maintained between the honeycomb seal and the fins. The inner shroud members (32) undergo deformation after operation of the gas turbine. However, because the honeycomb seals (4a, 4b) are mounted on the arm portions (3, 2), respectively, of the seal ring (1) disposed separately and independently from the inner shroud members (32), the honeycomb seals (4a, 4b) can remain unaffected by the deformation of the inner shroud members (32), whereby the predetermined clearances (t) can be consistently maintained.

Gas turbine blade

The present invention 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 a crack. In the gas turbine blade provided with a cooling passage therein, a protrusion is provided on the inside from the extension of blade profile on the outer surface of blade tip end.

Gas turbine segmental ring

Gas turbine segmental ring has an increased rigidity to suppress a thermal deformation and enables less cooling air leakage by less number of connecting portions of segment structures. Cooling air (70) from a compressor flows through cooling holes (61) of an impingement plate (60) to enter a cavity (62) and to impinge on a segmental ring (1) for cooling thereof. The cooling air (70) further flows into cooling passages (64) from openings (63) of the cavity (62) for cooling an interior of the segmental ring (1) and is discharged into a gas path from openings of a rear end of the segmental ring (1). Waffle pattern (10) of ribs arranged in a lattice shape is formed on an upper surface of the segmental ring (1) to thereby increase the rigidity. A plurality of slits (6) are formed in flanges (4, 5) extending in the turbine circumferential direction to thereby absorb the deformation and thermal deformation of the segmental ring (1) is suppressed.

Gas turbine cooled blade tip shroud

A tip shroud for a moving blade, which is made thin and light to be used at a gas turbine downstream stage which is improved in the flow of cooling air to enhance the cooling efficiency. Cooling air holes 13 to 16 of a slot shape are formed in a tip shroud 11 of a moving blade 10 and are opened in two side faces to release the cooling air from the inside of the moving blade 10. In the upper face of the tip shroud 11, there are formed cooling air holes 20 which communicate with the higher pressure side in a combustion gas flow direction R to release the cooling air so that the cooling air flows from the higher pressure side to the lower pressure side to cool a high stress portion Y. A high stress portion X is likewise cooled with the cooling air coming from the adjoining moving blade. The slope shapes of the cooling air holes 13 to 16 allow the cooling air to flow widely thereby to cool the face of the tip shroud 11, and the cooling air holes ?0 cools the high stress portions X and Y effectively.

Gas turbine moving blade platform

Gas turbine moving blade platform having simple cooling structure and effecting uniform cooling is provided. Cavities 2, 3, 4 are formed in platform 1 with impingement plate 11 being provided below the cavities 2, 3, 4. Cooling hole 5 communicates with cavity 2, cooling hole 6 with cavity 3 and cooling holes 7, 8 with cavity 4 and all these cooling holes pass through the platform 1 inclinedly upwardly. Cooling air 70 flows into the cavities 2, 3, 4 through holes 12 of the impingement plate 11 for effecting impingement cooling of platform 1 plane portion. The cooling air 70 further flows through the cooling holes 5, 6, 7 to blow outside inclinedly upwardly for cooling platform 1 peripheral portions. Thus, the platform 1 is cooled uniformly, no lengthy and complicated cooling passage is provided and workability is enhanced.

Gas turbine rotor

A gas turbine rotor, in which a plurality of discs having teeth of a bevel gear are juxtaposed to engage the teeth and are integrally fastened by a bolt extending through the discs, so that the discs may be cooled by feeding cooling air to the air passages of the individual discs sequentially at the running time, is provided whereby the cooling air is prevented from leaking by means having no wear. Radially outward of an air passage through hole of the faces of the adjoining discs, there are provided arms which are made lower than the dedendums of the teeth and protruded in an annular shape to confront each other; one of the arms has a tip made to have an elastically deformable thickness and a sectional shape bent inward or outward, whereas there is welded to the other arm an extension which has a tip made to have an elastically deformable thickness and a sectional shape bent inward or outward; and the end face of the tip of the one arm and the end face of the tip of the extension of the other arm are held in abutment against each other so that the two end faces may be forced, when the discs are integrated, into contact with each other to prevent leakage of cooling air.

Cooling structure for a gas turbine

In a cooling structure for a gas turbine, cooling air diffusion holes are formed from inner surface to outer surface of a platform so as to open from high pressure side blade surface of a moving blade offset in a direction toward low pressure side blade surface of adjacent moving blade confronting the high pressure side blade surface, in a direction of primary flow.

Cooled stationary blade for a gas turbine

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.

Ring segment of gas turbine

An object of the present invention is to provide a ring segment of a gas turbine in which the temperature is maintained low, damage due to high temperature oxidization is prevented, and high temperature deformation is prevented. In order to achieve the object, the present invention provides a ring segment of a gas turbine which comprises a blade ring, a main shaft and moving blades comprising a plurality of individual units which define an annular form by being arranged around the peripheral direction of the main shaft, and disposed so that its inner peripheral surface is maintained at a constant distance from the tips of the moving blades, wherein grooves which extends along the axial direction of the main shaft of the turbine are formed upon of the individual units so as mutually to confront one another; a seal plate which is inserted into each mutually confronting pair of the grooves so as to connect together the adjacent pair of individual units; and contact surfaces which are formed at positions more radially inward than the seal plates, which extend in the axial direction and the peripheral direction and which mutually contact one another.

Double cross type seal device for stationary gas turbine blades

Seal plate for a gas turbine moving blade

The present invention relates to a seal plate at the platform portion of a gas turbine moving blade. The seal plate is inserted in a gap between the adjacent platforms at each end portion of a seal pin to prevent air from leaking to the outside. A groove is provided at each of four corners of the end portion of platform of the moving blade, and the seal plate is inserted to cover the gap so as to extend the end portions of the adjacent platforms, by which the gap between the platforms is blocked. A seal pin and end seal pins are inserted between the platforms to seal this portion. However, there is a gap at the end portions, so that cooling air leaks from the lower part of platform. Since the seal plate is inserted in the groove to block the gap, the sealing property is increased.

Cooled moving blade for gas turbines

With a gas turbine systems, there is an increasing trend to use a high temperature combustion gas to enhance the operating efficiency of the gas turbine. This, however, is accompanied by the generation of cracks at an increased frequency in a base portion of the blade where thermal stress of large magnitude is likely to occur. An object of the invention is to provide a cooled moving blade for a gas turbine which has a blade profile capable of more effectively reducing thermal stress in a blade base portion, and thus, prevent cracks from occurring. A moving blade (1) is fixedly secured to a platform (2). On the other hand, a cooling air passage (3) is formed in a serpentine pattern inside of the blade for cooling with cooling air. The moving blade (1) has a base portion of a profile formed by an elliptically curved surface (11) and a rectilinear surface portion (12), wherein the rectilinear surface portion (12) is provided at a hub portion of the blade where thermal stress is large. In the conventional moving blade, the base portion is formed as an elliptic fillet R presenting a arcuate profile protruding convexly inward. The cross-sectional area of the blade is increased by providing the rectilinear surface portion (12). The heat capacity is increased compared with the conventional blade, due to the increased cross-sectional area of the blade. This in turn results in a decrease of the temperature difference due to the thermal stress. Thus, the thermal stress can be suppressed more effectively than with the conventional blade.

Gas turbine rotor blade

Tail tube seal structure for the combustor of a gas turbine

Gas turbine cooled blade

A gas turbine cooled blade is constructed without an increase in number of parts or time requirements, in which seal air is maintained at a lower temperature with heat exchange rate being suppressed and heat transfer rate of cooling medium in cooling passage is enhanced. A plurality of cooling passages (A, B, C, D, E) is provided in a blade, and the first row cooling passage (A) is covered at blade inner and outer peripheries and communicates with second row cooling passage (B) through communication holes (6) and with main flow gas path through film cooling holes (7). Second row cooling passage (B) is separated from the first row cooling passage (A) by a partition wall, and is closed at the inner peripheral side. Second row cooling passage (B) has at least one communication hole in the partition wall for allowing communication between an interior of the second row cooling passage and the interior of the first row cooling passage. At least one seal air passage (C, D or E) arranged downstream of the second row cooling passage has a seal air supply hole at the inner peripheral side for allowing communication between an interior of the seal air passage (C, D or E) and an inner cavity of the turbine adjacent to the inner peripheral side.

Moving blade of gas turbine

(57)【要約】 【課題】 翼部に長さ方向に穿設され冷却用気体を通す 複数の第一の冷却穴と、この第一の冷却穴に連通してシ ュラウドに面方向に穿設され冷却用気体を通す複数の第 二の冷却穴とを備えたガスタービンの動翼において、冷 却用気体を均等に分配して流せるように構成したガスタ ービン動翼を提供する。 【解決手段】 翼部２には冷却用気体を流す複数の第一 の冷却穴３が穿設され、シュラウド１には冷却用気体を 流す複数の第二の冷却穴５がシュラウド１の面方向に穿 設されている。第二の冷却穴５は、二段穴４を介して第 一の冷却穴３と一対一に連通されており、かつ、相隣る ものが動翼の背側及び腹側に交互に穿設されている。

Cooling system for cooling platform of gas turbine moving blade

An effective cooling structure that is intended to cool a platform of a gas turbine moving blade. Cooling air passages (B1, C1, D1; B2, C2, D2; and B3, C3 and D3) are provided through the interior and along the peripheral edge of the platform (2) to thereby cause a cooling air to pass therethrough.

Gas turbine moving blade steam cooling system

A gas turbine moving blade steam cooling system is disclosed wherein 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 a moving blade 1 is projected to form a projection portion 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 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 a secure prevention of steam leakage. The steam cools the moving blade 1 and can be recovered without leakage and stress concentration due to heat at the projection portions 4a, 4b of the blade root end portion is mitigated.

Gas turbine stationary blade

Gas turbine cooling static blade

(57)【要約】 【課題】 ガスタービン冷却静翼に関し、内部を空気冷 却すると共に内側シュラウドを冷却し、静翼全体の冷却 効果を高める。 【解決手段】 静翼１０内部には外側シュラウド１１か ら内側シュラウド１２へチューブ１３が貫通し、シール 空気２００がキャビティ１４へ流入し、キャビティ１４ 内部を高圧に保持して高温燃焼ガスをシールし、通路１ ６へ放出する。空気通路１９Ａには冷却空気１００の１ 部が流入し、前縁部を冷却すると共に内側シュラウド１ ２の翼部外周囲を通り、冷却して通路１８に放出され る。冷却空気１００の残りは通路１９Ｂより流入し、タ ービュレータ５５付きの空気通路１９Ｃ〜１９Ｆのサー ペンタイン冷却系路を通り、後縁の空気冷却穴２０より 放出される。前縁部の通路１９Ａの冷却空気により前縁 部と共に内側シュラウド１２外周部も冷却し、冷却効率 を高める。

Gas turbine rotor blade

Provided is a gas turbine rotor blade comprising a plurality of first cooling holes for flow of a cooling gas bored in a blade portion along its lengthwise direction and a plurality of second cooling holes for flow of the cooling gas bored in a shroud along its plane direction so as to communicate with the first cooling holes, and being constructed such that the cooling gas can be flown in a uniform distribution. The plurality of the first cooling holes 3 for flow of the cooling gas are bored in the blade portion 2 and the plurality of the second cooling holes 5 for flow of the cooling gas are bored in the shroud 1 along its plane direction. The second cooling holes 5 communicate with the first cooling holes 3, hole to hole, via two-step holes 4, and the second cooling holes 5 are bored alternately on the dorsal side and the ventral side of the rotor blade.

Manufacture of oxide superconductive bulk material

(57)【要約】 【課題】 ガスタービン車室から静翼を経て動翼に冷却 空気を供給するに際し、動翼前後の圧力差を確保して確 実に供給を行い得るようにしたものを提供することを課 題とする。 【解決手段】 相隣接するディスクの翼根位置に設けら れ互いに軸方向に対向して張り出して当接する一対のデ ィスク腕と、同腕の外側で動翼の上流側端部底位置に形 成した動翼翼溝キャビティと、静翼の内周端上流側に対 向して形成した静翼上流側キャビティを有し、前記一対 のディスク腕にはこれを軸方向に貫通して前記静翼上流 側キャビティと動翼翼溝キャビティとを連通する連通孔 を設け、静翼上流側キャビティの圧力に相当する静翼の 圧力差を動翼翼溝キャビティに確保し、これを動翼上流 側の圧力として動翼翼溝キャビティに続く翼根部への冷 却空気を流す力として用いるようにした。

Sealing device for gas turbine

(57)【要約】 【課題】 ガスタービンのシール装置に関し、シール部 のクリアランスコントロールを可能とし、もれ空気量を 減少する。 【解決手段】 静翼３１の内側シュラウド３２にはシー ルリング１が設けられ、シールリング１にはアーム部 ２，３がシュラウド端の下面に沿って突出して設けら れ、それぞれハニカムシール４ａ，４ｂが取付けられ る。ハニカムシール４ａは、動翼２１のプラットフォー ム２２のロータアーム部１１のフィン１１ａと対向し、 ４ｂは動翼２１のシール板１４のシール部１４ａに設け られたフィン１４ｂと対向してそれぞれ所定のクリアラ ンスを保って配置される。内側シュラウド３２は運転後 に変形するが、ハニカムシール４ａ，４ｂは円形状のシ ールリング１のアーム部３，２にそれぞれ取付けられて おり、内側シュラウド３２の変形には影響されずに所定 のクリアランスを保つことができる。

Turbine blade, gas turbine, and method of manufacturing turbine blade

Seal structure for gas turbines

Disclosed is a sealing apparatus for a gas turbine, comprising arm portions projecting from a seal ring, respectively, the seal ring fixedly securing inner shroud members of stationary blades, the arm portions extending along front end portions and rear end portions of the inner shroud members, respectively, as viewed in an axial direction thereof, wherein the seal ring, which is provided with the arm portions, is constructed separately and independently from the inner shroud members, and sealing members mounted on the arm portions, respectively, to constitute sealing mechanisms through cooperation with end portions of platforms of moving blades disposed adjacent to the front end portion and the rear end portion, respectively, of the inner shroud member, wherein the sealing mechanism seals off an interior of the inner shroud members from a combustion gas passage.

Cooling structure to cool platform for drive blades of gas turbine

A mechanism for cooling the platform for the drive blades of a gas turbine, which uses a simple configuration which will reliably cool the platform, the mechanism being constituted by cooling channels in the interior of the platform which open out from one of the cooling air channels for cooling the turbine blades and which exit the platform through the edge nearest the tail. In a preferred embodiment cooling channels 6a and 6b in platform 2 open out from the entrance to blade cooling channel 5a and travel from the head of the blade along the blade sides 3c and 3d and exit through the edge 3e near the tail of the blade. This structure diverts a portion of the cooling air 4a entering the blade 3 from the cooling channel in base 1 in order to cool platform 2. Alternatively or in addition, at least one of the following types of cooling passages may be provided: (a) cooling air channels 7, which extend from enclosed air space 11 below platform 2 to the upper surface of the platform; (b) convection cooling channels 8 which extend from the leading edge of the platform to the upper surface of the platform at the front or rear side of the blade, and/or (c) air channels 9 which extend on the rear of the turbine blade obliquely from the underside of the platform to the trailing edge of the platform. These various channels or combinations thereof constitute a cooling structure through which air can flow to cool a platform for the drive blades of a gas turbine in an efficient and effective manner.

Cooling structure to cool platform for drive blades of gas turbine

A mechanism for cooling the platform for the drive blades of a gas turbine uses a simple configuration which reliably cools the platform. The mechanism includes cooling channels in the interior of the platform which open out from one of the cooling air channels for cooling the turbine blades and which exit the platform through the edge nearest the tail. Cooling channels in the platform open out from the entrance to blade cooling channels, travel from the head of the blade along the blade sides, and exit through the edge near the tail of the blade. This structure diverts a portion of the cooling air entering the blade from the cooling channel in the base in order to cool the platform. Cooling air channels may extend from an enclosed air space below the platform to the upper surface of the platform at the front or rear side of the blade. Air channels may also extend on the rear of the turbine blade obliquely from the underside of the platform to the trailing edge of the platform. These channels or combinations thereof constitute a cooling structure through which air can flow to cool a platform for the drive blades of a gas turbine in an efficient and effective manner.

Turbine moving blade, turbine stationary blade, turbine split ring, and gas turbine

The present invention provides a turbine moving blade, a turbine stationary blade, and a turbine split ring which are capable of restraining the deterioration and peeling-off of a thermal barrier coating easily and surely, and a gas turbine capable of enhancing the energy efficiency by increasing the temperature of combustion gas. The turbine moving blade provided in a turbine constituting the gas turbine includes a platform having a gas path surface extending in the combustion gas flow direction, and a blade portion erecting on the platform. The thermal barrier coating covering the gas path surface is formed so as to go around from the gas path surface to an upstream-side end face and a downstream-side end face of the outer peripheral faces of the platform.

Gas turbine rotor

A gas turbine rotor, in which a plurality of discs having teeth of a bevel gear are juxtaposed to engage the teeth and are integrally fastened by a bolt extending through the discs, so that the discs may be cooled by feeding cooling air to the air passages of the individual discs sequentially at the running time, is provided whereby the cooling air is prevented from leaking by means having no wear. Radially outward of an air passage through hole of the faces of the adjoining discs, there are provided arms which are made lower than the dedendums of the teeth and protruded in an annular shape to confront each other; one of the arms has a tip made to have an elastically deformable thickness and a sectional shape bent inward or outward, whereas there is welded to the other arm an extension which has a tip made to have an elastically deformable thickness and a sectional shape bent inward or outward; and the end face of the tip of the one arm and the end face of the tip of the extension of the other arm are held in abutment against each other so that the two end faces may be forced, when the discs are integrated, into contact with each other to prevent leakage of cooling air. 16

Gas turbine stationary blade

The gas turbine stationary blade comprises a stationary blade section provided therein with a passage for cooling air, an inner shroud for supporting the stationary blade section on the side of a discharge port of the cooling air, and a plurality of segments each of which includes at least one stationary blade section and at least one inner shroud. A flow passage is pulled out from the discharge port of the cooling air, and the flow passage is introduced to a front edge corner section of the inner shroud and is extended rearward along a side edge of the inner shroud.

Cooled platform for a gas turbine rotor blade

The present invention relates to a cooled platform for a gas turbine moving blade, and enables a 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. Therefore, the peripheral portion of platform can be cooled by steam, air is not needed, and the platform can be cooled by steam when the steam cooling system is used to cool the blade.

Gas turbine stationary blade unit

Gas turbine stationary blade unit in which two stationary blades are built in a segment by shrouds is provided with object to lessen occurrence of cracks. Two stationary blades 1a, 1b are fixed respectively by outer shroud and inner shroud, each divided into two parts 2a, 2b and 3a, 3b. Flanges 4a, 4b are provided to divided end portions of the outer shrouds 2a, 2b to be jointed together by bolts via boltholes 7. Likewise, flanges 5a, 5b are provided to divided end portions of the inner shrouds 3a, 3b to be so jointed. If the two stationary blades 1a, 1b are fixed in a segment by the shrouds which are not divided, restraining force becomes larger, local stress occurs due to thermal stress and frequency of crack occurrence increases, but as the shrouds are divided respectively into two parts and jointed together by bolts, crack occurrence is lessened. Also, pinholes are provided in face of divided portion of the shrouds and pins are inserted thereinto for connection of the divided shrouds, thereby relative movement between the divided shrouds is prevented and a strong jointed blade unit is provided.

Stationary blade shroud of a gas turbine

A base plate and honeycomb member disposed at the inner circumference side of an inside shroud are fixed to the inside shroud, at a phase deviation in the peripheral direction, with respect to the inside shroud, so as to plug the missing range of seal member, out of gaps between adjacent inside shrouds, and therefore leakage of purge air from the missing range of seal member is prevented without adding new constituent members.

Gas turbine cooled moving blade

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. A turbulator is installed in the middle of steam passage in order to generate turbulence and enhance heat transfer. Since the moving blade is evenly cooled by steam alone without using air, and the steam is then collected, the performance of the gas turbine can be enhanced.

Gas turbine split ring

In the gas turbine split ring, on an outer peripheral surface 1b between two cabin attachment flanges, a circumferential rib which extends in the circumferential direction and an axial rib which extends in the axial direction and has a height taller than that of the circumferential rib are, respectively, formed in plural lines, so that it is possible to suppress heat deformation in the axial direction which largely contributes to reduction of the tip clearance compared to head deformation in the circumferential direction more efficiently.

Stationary blade shroud of a gas turbine

A base plate and honeycomb member disposed at the inner circumference side of an inside shroud are fixed to the inside shroud, at a phase deviation in the peripheral direction, with respect to the inside shroud, so as to plug the missing range of seal member, out of gaps between adjacent inside shrouds, and therefore leak of purge air from the missing range of seal member is prevented without adding new constituent members.

Double cross seal for gas turbine guide vanes

Gas turbine split ring

In gas turbine split rings, end faces having bent surfaces are formed in flanges. Adjoining split rings are coupled together with a groove therebetween to form a cylindrical split ring. Notches are formed in the flanges. These notches are sealed by inserting a seal plate into the notches of adjoining split rings. A hole for passing cooling air is drilled obliquely in the flange. Cooling air is allowed to flow out along the direction of rotation (of the turbine) . This cooling air cools the outlet of the groove due to the effect of film cooling. Because of such cooling, high temperature gas is prevented from staying in this area, cooling effect is enhanced, and hence burning of the end portions can be prevented.

Gas turbine rotor for steam cooling

Device for sealing gas turbine stator blades

A device for sealing gas turbine stator blades, wherein the sealing pressure of the inner shroud is enhanced to increase the sealing effect; an external shroud (32) is fitted to a blade ring (50) via heat insulating rings (32a, 32b), and the blade ring (50) is provided with air holes (1, 51), with one hole (1) communicating to a space (53) and the other (51), to a seal tube (2); a tip (3) of the seal tube (2) is inserted into the air hole (51), and a spring (6) is arranged between a protrusion (4) of the tube (2) and an engaging part (5) of the air hole (51) to fix the seal tube (2) detachably; cooling air (54), on one hand, enters the space (53) through the air hole (1), cools the inside of the shroud and a stator blade (31), and is discharged from the trailing edge, and, on the other hand, enters a cavity (36), the inside of which, as the tube (2) being independent of the space (53), is not subjected to pressure loss and kept at a high pressure; and high pressure air from the cavity (36) flows out of seal parts (40a, 40b), as indicated by S1 and S2, to prevent invasion of high temperature gas from the combustion gas passage.

Stationary blade structure of a gas turbine

Inspiratory cooling system

Seal of a turbine shroud ring

Gas turbine moving blade

In gas turbine moving blade, convection of cooling air is promoted to enhance heat transfer rate, cooling effect of shroud is enhanced and entire cooling effect of the blade is enhanced. Inner cavity 10 is formed in the blade in entire length thereof. Multiplicity of pin fins 5 are provided in the inner cavity being fixed to wall thereof. Enlarged cavity 6 is formed in shroud 2 of terminal end of the blade 1. Cooling air entering the inner cavity 10 of the blade 1 flows into the enlarged cavity 6 and is flown out of the shroud 2 downwardly through holes 7 of peripheral portion of the enlarged cavity 6. Entire portion of the shroud 2 is cooled uniformly and cooling effect of entire blade is enhanced by enhanced heat transfer rate in the blade and by uniform cooling of entire shroud.

Seal structure for gas turbines

Disclosed is a sealing apparatus for a gas turbine, comprising arm portions projecting from a seal ring, respectively, the seal ring fixedly securing inner shroud members of stationary blades, the arm portions extending along front end portions and rear end portions of the inner shroud members, respectively, as viewed in an axial direction thereof, wherein the seal ring, which is provided with the arm portions, is constructed separately and independently from the inner shroud members, and sealing members mounted on the arm portions, respectively, to constitute sealing mechanisms through cooperation with end portions of platforms of moving blades disposed adjacent to the front end portion and the rear end portion, respectively, of the inner shroud member, wherein the sealing mechanism seals off an interior of the inner shroud members from a combustion gas passage.

Gas turbine cooling system, gas turbine plant equipped with same, and method for cooling high-temperature section of gas turbine

吸気冷却システム

【課題】吸気ダクトの下流側に有する鉛直ダクトの壁面を伝って流れるドレンを捕集して、ドレンの圧縮機への侵入を防止する。 【解決手段】ガスタービン１８の吸気冷却システム１００であって、吸気入口２２から取り込まれた吸気をガスタービンの圧縮機１４に導く吸気ダクト１２と、吸気ダクト内に設けられ、外部から導入された冷却媒体との熱交換により吸気を冷却する冷却部２６と、吸気ダクトの下流側に有する鉛直ダクト１２ｃを介して設けられ、吸気を圧縮機に導くマニホールド部１２ｄと、マニホールド部の入口側に設けられ、鉛直ダクトを介して導入される吸気に含まれる不純物を取り除くフィルタ部４２と、フィルタ部の直上に鉛直ダクトの内壁面１２ｃ１、１２ｃ２に沿うように設けられる樋部材であって、内壁面を伝って流れるドレンを回収可能なドレンキャッチャ１１０と、を備える。 【選択図】図１

吸氣冷卻系統

以回收未被冷卻盤管(冷卻部)之排水皿接住之排水為目的，本發明之燃氣渦輪機28之吸氣冷卻系統100之特徵在於包括：吸氣管12，其將自吸氣入口22獲取之吸氣導入燃氣渦輪機之壓縮機14；冷卻部26，其設於吸氣管內，藉由與自外部導入之冷卻介質之熱交換而對吸氣進行冷卻；凸階差部13，其相對於冷卻部之下游側具有之吸氣管之底面12a1、12a2而形成為凸狀；及排水排出孔部110，其形成於冷卻部之下游側具有之吸氣管之底面12a1、12a2，可將冷卻部之表面產生並滴下之排水排出至吸氣管之外部。

構造部材の温度推定方法及び構造部材の保全方法

【課題】ガスタービンの動翼をはじめとするＮｉ基耐熱合金で構成された構造部材の温度を、広い温度範囲にわたって推定可能な構造部材の温度推定方法、及びこの温度推定方法を用いた構造部材の保全方法を提供する。 【解決手段】Ｎｉ基耐熱合金で構成された構造部材の温度を推定する温度推定方法であって、前記構造部材の表層に形成されたγ’相が消失した変質層の厚さと、温度及び時間の関係を示した指標を定め、前記構造部材における前記変質層の厚さを測定し、該測定値を前記指標と対応させて、前記構造部材の温度を算出する。 【選択図】図２

Stationary blade in gas turbine and gas turbine comprising the same

A stationary blade (133) of a gas turbine (10), which can reduce thermal stress produced at a portion in the vicinity of a rear edge of an inner shroud (133c) of the stationary blade. The stationary blade is positioned adjacent to at least one of moving-blade disks (32e, 34e, 36e, 38e) in an axial direction of the gas turbine. A concave portion (40) is provided in the inner shroud in a manner such that the concave portion is formed in the vicinity of a rear edge of the inner shroud and on an inner-peripheral face of the inner shroud, where cooling air passes along the inner-peripheral face which faces a rotation shaft of the moving-blade disks; and a protruding portion (133i) which protrudes towards the rotation shaft is formed at the rear edge of the inner shroud.

ガスタービン用吸気冷却装置、ガスタービンプラント、既設ガスタービンプラントの再構築方法、及び、ガスタービンの吸気冷却方法

【課題】吸気冷却を行うことで発電用ガスタービンによる発電の高出力化と経済性向上との両立を図る。 【解決手段】ガスタービン用吸気冷却装置１０は、冷凍機１１を駆動させて発電用ガスタービン２の吸気冷却を行う第一の冷却手段１２と、加湿することで発電用ガスタービン２の吸気冷却を行う第二の冷却手段１３と、第一の冷却手段１２及び第二の冷却手段１３のそれぞれを稼動または非稼動に切替可能な冷却制御部２０とを備え、冷却制御部２０は、前記第一の冷却手段で吸気冷却を行って発電することによる発電収支と、第二の冷却手段で吸気冷却を行って発電することによる発電収支との大小比較に基づいて、第一の冷却手段及び第二の冷却手段のそれぞれの稼動または非稼動を決定する。 【選択図】図１

吸氣冷卻系統

以捕獲經過吸氣管之下游側具有之鉛直管之壁面而流動之排水，防止排水向壓縮機之滲入為目的，本發明之燃氣渦輪機18之吸氣冷卻系統100包括：吸氣管12，其將自吸氣入口22獲取之吸氣導入燃氣渦輪機之壓縮機14，且包含鉛直管12c、及鉛直管之下游之歧管部12d；冷卻部26，其設於吸氣管內，藉由與自外部導入之冷卻介質之熱交換而對吸氣進行冷卻；過濾部42，其設於歧管部之入口側，去除經由鉛直管而導入之吸氣所含之雜質；及排水收集器110，其係於過濾部之正上方沿鉛直管之內壁面12c1、12c2而設之水收集引導構件，可回收經過內壁面而流動之排水。

機器状態推定装置、機器状態推定方法およびプログラム

【課題】対象機器に係る負荷による寿命を劣化要因ごとに適切に管理する。 【解決手段】状態量取得部は、対象機器の温度を含む前記対象機器の状態量を取得する。負荷特定部は、状態量に基づいて、対象機器の負荷の履歴を特定する。余寿命算出部は、負荷特定部が特定した負荷の履歴に基づいて、複数の劣化種別のそれぞれについて対象機器の余寿命に関するパラメータを算出する。 【選択図】図１

プラント分析装置、プラント分析方法、およびプログラム

【課題】オーバーファイアリング運転での運転の可否を正確に判定する。【解決手段】状態量取得部は、タービンの状態量を取得する。時間算出部は、タービンの設計寿命と状態量とに基づいて、タービンのオーバーファイアリング運転での運転可能時間を算出する。距離算出部は、状態量に基づいてマハラノビス距離を算出する。判定部は、マハラノビス距離および運転可能時間によりタービンのオーバーファイアリング運転の可否を判定する。【選択図】図２

タービン部材の温度推定方法

【課題】メタル温度が低い場合でも、その温度を精確に推定可能なタービン部材の温度推定方法を提供する。【解決手段】温度推定方法は、加熱前のタービン部材中に含まれるγ’相を検出することと、加熱前の検出の結果に基づいて、第１粒径においてその個数が最大となる第１分布データ、及び第１粒径よりも大きい第２粒径においてその個数が最大となる第２分布データを有し、加熱前のγ’相の粒径とその粒径のγ’相の単位面積当たりの個数との関係を示す加熱前粒径分布データを導出することと、加熱後のタービン部材中に含まれるγ’相を検出することと、加熱後の検出の結果に基づいて、加熱後のγ’相の粒径とその粒径のγ’相の単位面積当たりの個数との関係を示す加熱後粒径分布データを導出することと、加熱前粒径分布データと加熱後粒径分布データとを比較することと、比較の結果に基づいて、加熱後のタービン部材の温度を推定することと、を含む。【選択図】図３

Pale fixe de refroidissement pour turbine a gaz