Abstract: The present invention relates to an impingement cooling mechanism that ejects a cooling gas toward a cooling target (2) from a plurality of impingement holes (3b) formed in a facing member (3) that is disposed facing the cooling target (2). Blocking members (5) that block a crossflow (CF), which is a flow formed by the cooling gas after being ejected from the impingement holes (3b), are installed on at least the upstream side of the crossflow (CF) with respect to at least a portion of the impingement holes (3b). Turbulent flow promoting portions (6) are provided in the flow path (R) of the crossflow (CF) regulated by the blocking members (5).

Abstract: The gas turbine engine (S1) includes: turbine blades (7b); and a cooling air supply unit (11) to supply cooling air to the turbine blades (7b). A flow path surface (31) is formed so as to be positioned in an upstream side of the turbine blades (7b) and so as to be connected to a base surface (32) in which the turbine blades (7b) are provided. The flow path surface (31) includes: depression portions (31a) depressed relative to the base surface (32), each depression portion (31a) including at least an area overlapping with a front end (7b1) of each turbine blade (7b), when viewed from a direction of the turbine axis (L); and protrusion portions (31b) protruding relative to the base surface (32), each protrusion portion (31b) being at least part of each area positioned between front ends (7b1) of the turbine blades (7b), when viewed from the above direction.

Abstract: The present invention is a turbine blade (1) having a hollow blade body (2). This turbine blade (1) is provided with: cooling air holes (5) that penetrate the blade body (2) from an internal wall surface (2e) to an external wall surface (2f) thereof, and are provided with a straight tube portion (5a) that is located on the internal wall surface (2e) side of the blade body (2), and an expanded diameter portion (5b) that is located on the external wall surface (2f) side of the blade body (2); and with a guide groove (6) that is located on an internal wall of the expanded diameter portion (5b) and that guides cooling air (Y) in the expanded diameter portion (5).

Abstract: A turbine blade has hollowness, and is provided with a back-side wall of which a portion of the inner wall surface is exposed at the rear edge portion, with cooling air flown along the inner wall surface at the exposed region; and a recess provided in the inner wall surface at the exposed region. The contour of the recess (5) viewed from the normal direction of the inner wall surface of the back-side wall is set to a shape that is symmetrical centered on a reference axis (L) that intersects the flow direction of cooling air, and that broadens along the reference axis (L).

Abstract: The invention is a combustor liner (12) of a dual wall cooling structure including an inner wall section (30) configured to surround a combustion region (13) and in which a plurality of effusion cooling holes (31) are formed, and an outer wall section (20) formed to be spaced apart from the inner wall section (30) and in which a plurality of impingement cooling holes (21) are formed, wherein the inner wall section (30) is constituted by a plurality of plate-shaped members (40), and a support guide member (50) is provided which is configured to guide the plurality of plate-shaped members (40) to enable free insertion and extraction and support the plurality of plate-shaped members (40) at intervals such that deformation by thermal expansion is able to be absorbed.

Abstract: It has a flat impingement hole (2) of which the opening width (D1) in the flow direction of a crossflow (F) in the gap between a cooling target (10) and an opposing member (20) is set greater than the opening width (D2) in a direction orthogonal with the flow direction of the crossflow F. Accordingly, the cooling efficiency is further improved by the impingement cooling mechanism.

Abstract: The film cooling structure includes: a wall surface along which a heating medium flows; and at least one pair of film cooling holes that open at the wall surface and that blow a cooling medium. The pair of film cooling holes are arranged to be adjacent to each other in a main flow direction of the heating medium. In addition, perforation directions of the pair of film cooling holes are set such that directions of swirls of the cooling medium formed by blowing are opposite to each other, a swirl of the cooling medium on a downstream side in the main flow direction is mixed and merged with another swirl of the cooling medium on an upstream side in the main flow direction, and the merged cooling medium flows along the wall surface in a direction intersecting with the main flow direction.

Abstract: A plurality of film cooling holes are formed, so as to communicate with a front cooling passage, in a vane surface on the front-edge side of a stator vane body of a turbine stator vane. The hole cross-section of each of the film cooling holes has a rectangular long-hole shape extending in a direction parallel to the cross-section along the span direction and having a rounded corner. The hole-center line of each of the film cooling holes is inclined with respect to the thickness direction in the cross-section along the span direction. The exit-side portion of the hole wall surface of each of the film cooling holes is inclined with respect to the thickness direction by a greater degree than that of the hole-center line.

Abstract: A combustor for gas turbine engine, wherein: a plurality of impingement-cooling holes are opened passing through an external wall of a liner for cooling air to blow out towards the outer surface of an internal wall of the liner; a plurality of pin-fins are formed on the outer surface of the internal wall of the liner; a plurality of effusion cooling holes are opened passing through the internal wall of the liner for cooling air to blow out along the inner surface of the internal wall of the liner. The top surface of each pin-fin is not in contact with the inner surface of the external wall of the liner and the ratio of the height of the pin-fin to the equivalent diameter of the impingement-cooling hole is set to be 1.0-3.0.

Abstract: A turbine blade (1) includes a plurality of rows of cooling portions (4, 4A, 4B) that have notch portions (41) that discharge cooling gas that has been introduced into an internal portion (25) of a blade portion (2) onto a ventral side blade surface (23), and that are formed in rows that are stacked in a direction between the blade front edge and the blade rear edge. This turbine blade (1) is provided with: turbulence promoting cooling portions (4B) that are provided in the row located furthest to the downstream side from among the plurality of rows, and that have turbulence promoting devices (44, 45, 46) in areas exposed by the notch portions (41); and film cooling portions (4A) that are provided in at least one of other rows from among the plurality of rows, and that form a film cooling layer in the areas exposed by the notch portions (41).

Abstract: A plurality of film cooling holes 13 is formed at regions exposed to a hot gas in a turbine component main body 3, each of the film cooling holes 13 is comprised of a straight hole portion 15 formed at a portion at a side of an inner wall surface 5a of the turbine component main body 3 and a diverging hole portion 17 formed at a portion at a side of an outer wall surface 5b of the turbine component main body 3, and further each of the film cooling holes 13 is so constituted that an outlet surface 17p of the diverging hole portion 17 from a center portion to both end sides in a lateral direction L3 gradually to a great extent leans toward a downstream side of the hot gas on the basis of a hypothetical plane VP parallel to two directions of an axial center direction L2 and a lateral direction L3 of the straight hole portion 15.

Abstract: A plurality of film cooling holes 13 is formed at regions exposed to a hot gas in a turbine component main body 3, each of the film cooling holes 13 is comprised of a straight hole portion 15 formed at a portion at a side of an inner wall surface 5a of the turbine component main body 3 and a diverging hole portion 17 formed at a portion at a side of an outer wall surface 5b of the turbine component main body 3, and further each of the film cooling holes 13 is so constituted that an outlet surface 17p of the diverging hole portion 17 from a center portion to both end sides in a lateral direction L3 gradually to a great extent leans toward a downstream side of the hot gas on the basis of a hypothetical plane VP parallel to two directions of an axial center direction L2 and a lateral direction L3 of the straight hole portion 15.

Abstract: A gas turbine engine includes a front cooling plate provided on a front stationary section in the vicinity of a front side of a turbine disc, a front cooling passage formed between an opposing surface of the front cooling plate and a front surface of a rim of the turbine disc, a rear cooling plate provided on a rear stationary section in the vicinity of a rear side of the turbine disc, and a rear cooling passage formed between an opposing surface of the rear cooling plate and a rear surface of the rim of the turbine disc.

Abstract: A cooled turbine rotor blade comprises a blade body having a cooling passage so that cooling air is introduced, and a film cooling hole which is formed on the surface of the blade body and includes an inlet opening and a wedge shaped spurt out opening. The film cooling hole is divided into at least two openings by a partition wall on the outlet side of the film cooling hole.

Abstract: A gas turbine engine includes a front cooling plate (85) provided on a front stationary section (69) in the vicinity of a front side of a turbine disc (55), a front cooling passage (87) formed between an opposing surface of the front cooling plate (85) and a front surface of a rim of the turbine disc (55), a rear cooling plate (93) provided on a rear stationary section (81) in the vicinity of a rear side of the turbine disc (55), and a rear cooling passage (95) formed between an opposing surface of the rear cooling plate (93) and a rear surface of the rim of the turbine disc (55).

Abstract: A cooled turbine rotor blade comprises a blade body having a cooling passage so that cooling air is introduced, and a film cooling hole which is formed on the surface of the blade body and includes an inlet opening and a wedge shaped spurt out opening. The film cooling hole is divided into at least two openings by a partition wall on the outlet side of the film cooling hole.