Mitsubishi Heavy Industries Limited

Mitsubishi Jukogyo Kabushiki Kaisha

Kiyoshi Suenaga

Yasuoki Tomita

Hiroki Fukuno

Sunao Aoki

Kazuo Uematsu

Yukihiro Hashimoto

Koji Watanabe

Katsunori Tanaka

Shigemi Mandai

Masaaki Matsuura

Eiji Akita

Ichiro Fukue

Tatsuo Ishiguro

Taku Ichiryu

Eisaku Ito

Masaki Ono

Kuniaki Aoyama

Masamitsu Kuwabara

Kazuharu Hirokawa

Shunsuke Torii

Kazufumi Ikeda

Rintaro Chikami

Yoshichika Sato

Masahito Kataoka

Toshishige Ai

Kenichiro Takeishi

Yoshikuni Kasai

Hideki Murata

Shigeyuki Maeda

Hiroyuki Yamamoto

Satoshi Hada

Hideaki Sugishita

Yoichi Iwasaki

Kenichi Arase

Masanori Yuri

Tsuyoshi Kitamura

Junichiro Masada

Daigo Fujimura

Jun Kubota

Yoshiaki Tsukuda

Hisato Arimura

Toshiaki Sano

Akihiko Shirota

Yuko Suenaga

Akiko Aizawa

Youichi Iwasaki

A blade cooling structure of a gas turbine, which can reduce the pressure loss of a cooling medium without decreasing the heat transfer coefficient, is provided. The blade cooling structure comprises a cooling passage () for flowing cooling air (A) from a proximal end portion () toward a blade portion () of a moving blade (), a plurality of turbulators () arranged, on both wall surfaces of the cooling passage () opposing each other, in such a manner as to be inclined with respect to the flowing direction of the cooling air (A), and a plurality of dimples () formed in a downstream region (N) downstream of a center position (O) in the flowing direction of the cooling air (A) on the wall surface of the cooling passage () between the adjacent turbulators ().

Blade cooling structure of gas turbine

For the purpose of reduced NOx gas emission, a gas turbine engine comprises a cylinder having a combustion region inside of the cylinder; a resonator having a cavity and provided around the surface of the cylinder and sound absorption holes formed on the cylinder and having opening ends on the cylinder.

Gas turbine combustor, gas turbine, and jet engine

There is provided a combustor to burn fuel, comprising a bypass passage connected to one side of the combustor to supply air into the combustor; and an annular passage provided around the combustor and connected to the bypass passage, wherein air supplied through the bypass passage passes in the annular passage in the circumferential direction, and is uniformly supplied into the combustor in the circumferential direction thereof through an opening which connects the combustor and the annular passage. Accordingly, compressed air passing through the bypass passage can be supplied uniformly into a tail portion of the combustor, and unevenness of temperature distribution in a cross section of the combustor tail portion can be reduced.

Gas turbine combustor having bypass and annular gas passage for reducing uneven temperature distribution in combustor tail cross section

A gas turbine combustor includes a side wall, for defining a combustion volume, having upstream and downstream ends, a pilot nozzle, disposed adjacent the upstream end of the side wall, for discharging a pilot fuel to form a diffusion flame in the combustion volume, and a plurality of main nozzles, provided around the pilot nozzles, for discharging a fuel-air mixture to form premixed flames in the combustion volume. Film air is supplied into the combustion volume downstream of the main nozzles along the inner surface of the side wall to reduce the fuel-air ratio in a region adjacent the inner surface of the side wall and to restrain a combustion-driven oscillation in the combustion volume.

Gas turbine combustor

In a tail tube seal structure of gas turbine, a U-shaped groove is provided at one side of a tail tube seal where a flange of a tail tube outlet is fitted, and a pi-shaped groove is provided at other side of the tail tube seal where a gas pass side flange end is fitted, thereby composing the seal of the connection area. Inclined cooling holes are drilled in the tail tube seal in addition to the cooling holes existing conventionally. The cooling air flows in from the inclined holes and cools the gas pass side of the groove due to the film effect. Therefore, the difference in thermal expansion between the groove and flange end is decreased, the wear of this area is decreased, and the reliability of the seal is enhanced.

Tail tube seal structure of combustor and a gas turbine using the same structure

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 moving blade

A steam-cooled combustor for a gas turbine has cooling steam passages, the steam flowing in opposite axial directions in adjacent channels.

Gas turbine steam-cooled combustor with alternately counter-flowing steam passages

A combined cycle power plant has a gas turbine and an exhaust heat recovery boiler to which exhaust gas from the gas turbine is supplied. The exhaust heat recovery boiler is configured to generate steam for cooling. The combined cycle power plant has several elements: a cooling steam supply pipe; a cooling steam recovery pipe; and a cooling steam connecting pipe. The cooling steam connecting pipe has a cross-sectional area larger than cross-sectional areas of the cooling steam supply pipe and the cooling steam recovery pipe so that a pressure drop thereacross, as a whole, is small.

Gas turbine for combined cycle power plant

Turbine blade of gas turbine etc. and gas turbine equipment using the turbine blade suppress occurrence of thermal stress caused by temperature difference and provide high reliability. In gas turbine equipment comprising rotational portion of rotor ( ) and moving blade ( ), stationary portion of casing ( ), stationary blade ( ), various supporting members, etc. and combustor, thermal stress reducing portion is provided in any one or both of moving blade joint adjacent portion ( ) between moving blade trailing edge portion ( ) and platform ( ) and stationary blade joint adjacent portion ( ) between stationary blade trailing edge portion ( ) and shroud ( ). By the thermal stress reducing portion, undesirable thermal stress occurring in the blade joint adjacent portions ( ) is reduced and reliability of the turbine blade and the gas turbine equipment is enhanced.

Gas turbine equipment and turbine blade

In a stationary blade of a gas turbine, the pressure resisting strength can be decreased by using low-pressure cooling air, and effective cooling can be performed by means of a simple construction. An inside shroud,and an outside shroud,are cooled by cooling air passing through an impingement plate. A trailing edge portion of a blade portion,which is thin in shape, is cooled by cooling air flowing in an air passage,Part of this air is discharged through a hole,on the side of the inside shroud,as inside seal air for a combustion gas passage.

Stationary blade of gas turbine

A gas turbine cooled blade is constructed without an increase in the number of parts or time requirements, in which seal air is maintained at a lower temperature with the heat exchange rate being suppressed, and the heat transfer rate of the 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 the blade inner and outer peripheries and communicates with the second row cooling passage (B) through communication holes (,) and with the main flow gas path through film cooling holes (,). The second row cooling passage (B) communicates with the blade inner peripheral cavity (,) to form a seal air supply passage. A plurality of ribs (,) are disposed on the inner wall of cooling passage (,) with a predetermined pitch (P). The ribs are arranged alternately and are inclined against cooling medium flow with respective higher first end contacting lower side faces of an immediately upstream rib at a position on both side portions of cooling passage (,). High heat transfer rate areas are formed on both side portions of cooling passage (,), and the average heat transfer rate in cooling passage is enhanced.

Gas turbine cooled blade

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 split ring

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.

Gas turbine stationary blade

Growing community of inventors

Takasago, Japan

Kiyoshi Suenaga