Cooling structure of gas turbine combustor in which cooling medium flows through grooves in wall is improved so that adjustment of flow velocity, pressure loss and heat transfer rate of cooling medium flow in the wall becomes possible and cooling effect thereof is enhanced. Wall of combustor tail tube is made in double structure in which outer plate (1) and inner plate (4) are jointed together being lapped one on another. The outer plate (1) has air inlet hole (3) and groove (2) formed therein. The groove (2) is closed by jointing of the inner plate (4) to the outer plate (1). The inner plate (4) has air outlet hole (5) formed therein. The groove (2) communicates with the air inlet hole (3) and the air outlet hole (5). Cross sectional shape of the groove (2) is changed two-dimensionally or three-dimensionally such that width enlarges toward the hole (5) from the hole (3) or depth is constant or changed in tapered form. Cooling air flows into the groove (2) from the air inlet hole (3) of tail tube surface to flow toward both sides along the groove (2) for cooling of the wall. The air is thereby heated to expand to increase flow velocity and pressure loss, but flow passage enlarges toward the hole (5) and flow velocity is suppressed and pressure loss is reduced.

Gas turbine combustor cooling structure

A cooling arrangement for a gas turbine power plant is disclosed in which cycle air is used in combination with atmospheric air as the cooling fluid. A plurality of eductor tubes with nozzles are disposed about the turbine casing and serve to direct pressurized cycle air into associated cooling holes. The cycle air gives momentum to the secondary atmospheric air and the cooling medium passes through the holes, removing heat therefrom. The use of cycle air in combination with atmospheric air also provides for compartment ventilation in a gas turbine of the enclosed type.

Eductor cooled gas turbine casing

One embodiment of a transition-to-turbine seal ( 300 ) comprises a first, flattened section ( 302 ) adapted to be received in a peripheral axial slot ( 320 ) of a transition ( 325 ), and a second, generally C-shaped section ( 301 ). The generally C-shaped section ( 301 ) comprises a flattened portion ( 305 ) near the first, flattened section ( 302 ), and a curved portion ( 306 ) extending to a free edge ( 307 ). A fiber metal strip component ( 309 ) may be attached to the flattened portion ( 305 ) to define a first engagement surface adapted to engage an upstream side ( 336 ) of an outer vane seal rail ( 337 ), and a second engagement surface 308 , adjacent the free edge ( 307 ), provides an opposed wear surface adapted to engage a downstream side ( 338 ) of the outer vane seal rail ( 337 ). System embodiments also are described, in which such transition-to-turbine seal ( 300 ) is isolated from a hot gas path ( 350 ) by provision of a plurality of cooling apertures ( 327 ) in the transition ( 325 ).

Slidable spring-loaded transition-to-turbine seal apparatus and heat-shielding system, comprising the seal, at transition/turbine junction of a gas turbine engine

A method of fabricating a turbomachinery blade is provided whereby high stress levels associated with stripe mode resonance can be avoided. Once the nodal pattern for an objectionable stripe mode resonance condition of a metallic blade is determined by conventional means, a portion of the blade parent metal near the tip is replaced with a composite patch, the composite patch comprising elongated filaments embedded in a matrix material. The chordwise extent of the patch includes the nodal points at the objectionable stripe mode resonance condition and the filament orientation is chordwise to enhance chordwise stiffness.

Composite reinforcement of metallic airfoils

A device for controlling the clearance between the blade tips of a gas turbine engine turbine rotor and associated segmented shrouds comprises a control ring which is secured to the engine casing by a plurality of dowels. The control ring is supplied with cooling fluid to its hollow interior and is covered with a thermally insulating layer, whereby the rate of movement of the shroud segments can be maintained substantially equal to that of the adjacent turbine rotor.

Gas turbine engines

Gas turbine nozzle structure

Gas turbine power plant with axial flow compressor

Power plant frame structure having air-cooling means for turbine rotors and exhaust frame struts

The present invention relates to surface type heat exchangers including a plurality of tubes or channels connected between inlet and outlet headers and means for directing air or like medium across the tubes to effect heat exchange between the air and a fluid flowing through the tubes. The...

Howard Alan

Papers

Gas turbine nozzle structure

Gas turbine power plant with axial flow compressor

Power plant frame structure having air-cooling means for turbine rotors and exhaust frame struts

Surface type heat exchanger

Axial flow compressor