摘要:
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.
摘要:
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.
摘要:
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 (6) and with the main flow gas path through film cooling holes (7). The second row cooling passage (B) communicates with the blade inner peripheral cavity (10) to form a seal air supply passage. A plurality of ribs (31) are disposed on the inner wall of cooling passage (22) 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 (22). High heat transfer rate areas are formed on both side portions of cooling passage (22), and the average heat transfer rate in cooling passage is enhanced.
摘要:
A gas turbine moving blade platform having a simplified cooling structure for effecting uniform cooling of the platform. The platform (1) includes cavities (2, 3, 4) and an impingement plate (11) provided below the cavities (2, 3, 4). A cooling hole (5) communicates with cavity (2), cooling hole (6) communicated with cavity (3) and cooling holes (7, 8) communicate with cavity (4) and all of the cooling holes pass through the platform (1) at an inclined angle. Cooling air (70) flows into the cavities (2, 3, 4) through holes (12) in 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 angularly upward for cooling peripheral portions of the platform. Thus, the platform is cooled uniformly, no lengthy and complicated cooling passage is provided, and workability is enhanced.
摘要:
A collision plate having plural small holes is provided at an interval from a bottom surface of an inner shroud to form a chamber and guides cooling air from the small holes into the chamber. A leading edge flow path is provided at a leading edge side along a width direction and introduces the cooling air. A side flow path is provided along both sides of the inner shroud and guides the cooling air to a trailing edge side. A header is formed along the width direction near the trailing edge and guides the cooling air from the side flow path. Plural trailing edge flow paths are formed at the trailing edge side at intervals along a width direction, in which one end of each flow path is connected to the header and the other end is open at the trailing edge, and the cooling air in the header is ejected from the trailing edge.
摘要:
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.
摘要:
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.
摘要:
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.
摘要:
A gas turbine stationary blade having passages (23, 24) that are provided in the stationary blade (10). A front cylindrical insert (2) is provided in the passage (23) and a rear cylindrical insert (5) is provided in the passage (24), and the inserts are supported at two supporting portions (3a, 3b), (6a, 6b), respectively. A projection (1) is provided at a leading edge portion of the blade so that the leading edge, where the thermal loads are high, is made smaller in size and the number of rows of cooling holes (11a) in the leading edge portion is reduced. Air blowing holes (4b) are provided on the dorsal side rear portion of the front insert (2) and film cooling holes (12) are provided on the dorsal side of the blade, both have diameters that are larger than other air blowing and cooling holes provided in the insert (2) and the blade (10), so that dust in the cooling air is caused to flow out, thereby preventing clogging of the holes. The curved surface of the blade leading edge portion is formed on an elliptical curve, so that the cooling air is caused to flow smoothly. Also, curved surfaces of fillets are formed on an elliptical curve so that thermal stresses concentrated near the fillets are avoided.
摘要:
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 bypass hole and slits, 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.