A helicoidal structure promotes cooling of a liner applied in an annular space of a double-shell structure formed of the liner and a flow sleeve to cool a duct assembly, by increasing the residence time and cooling area of cooling compressed air on the surface of the liner. The helicoidal structure includes a helicoidal rib protruding from a surface of the liner to guide the cooling compressed air at a predetermined angle with respect to an axial direction of the liner, the helicoidal rib having cooling holes and forming a first cooling passage along which main-cooling compressed air flows and a second cooling passage along which auxiliary-cooling compressed air flows. The second cooling passage induces a flow of the auxiliary-cooling compressed air through the cooling holes, and the flow of the auxiliary-cooling compressed air is derived from a flow of the main-cooling compressed air in the first cooling passage.

A helicoidal structure promotes cooling of a liner applied in an annular space of a double-shell structure formed of the liner and a flow sleeve to cool a duct assembly, by increasing the residence time and cooling area of cooling compressed air on the surface of the liner. The helicoidal structure includes a helicoidal rib protruding from a surface of the liner to guide the cooling compressed air at a predetermined angle with respect to an axial direction of the liner, the helicoidal rib having cooling holes and forming a first cooling passage along which main-cooling compressed air flows and a second cooling passage along which auxiliary-cooling compressed air flows. The second cooling passage induces a flow of the auxiliary-cooling compressed air through the cooling holes, and the flow of the auxiliary-cooling compressed air is derived from a flow of the main-cooling compressed air in the first cooling passage.

Systems and methods for dual-fuel operation of a gas turbine combustor are provided. An exemplary gas turbine combustor may comprise one or more components, such as a cylindrical combustion liner, a flow sleeve, a main mixer, a radial inflow swirler, a combustor dome, and a fuel cartridge assembly, one or more of which may be configured to supply either a gaseous or a liquid fuel to the combustion liner, depending on whether gaseous fuel operation or liquid fuel operation of the combustor is desired.

Systems and methods for dual-fuel operation of a gas turbine combustor are provided. An exemplary gas turbine combustor may comprise one or more components, such as a cylindrical combustion liner, a flow sleeve, a main mixer, a radial inflow swirler, a combustor dome, and a fuel cartridge assembly, one or more of which may be configured to supply either a gaseous or a liquid fuel to the combustion liner, depending on whether gaseous fuel operation or liquid fuel operation of the combustor is desired.

This disclosure provides tools for installing and removing a combustion liner in a combustion casing. A mounting frame has fasteners positioned circumferentially around an end casing surface of the combustion casing and removably attaches the mounting frame to the combustion casing. A positioning member has pushing surfaces and pulling surfaces distributed circumferentially around the combustion liner. An axial positioning mechanism engages the mounting frame and the positioning member along a common axis. The axial positioning mechanism incrementally positions the positioning member relative to the mounting frame by adjusting a positioning distance between the mounting frame and the positioning member.

This disclosure provides tools for installing and removing a combustion liner in a combustion casing. A mounting frame has fasteners positioned circumferentially around an end casing surface of the combustion casing and removably attaches the mounting frame to the combustion casing. A positioning member has pushing surfaces and pulling surfaces distributed circumferentially around the combustion liner. An axial positioning mechanism engages the mounting frame and the positioning member along a common axis. The axial positioning mechanism incrementally positions the positioning member relative to the mounting frame by adjusting a positioning distance between the mounting frame and the positioning member.

A turbine engine including a combustion chamber having an inner annular shroud and an outer annular shroud that are coaxial with each other and that are connected at their downstream ends respectively to an inner annular link wall and to an outer annular link wall, for linking respectively to an inner casing and to an outer casing. At least a first one of the inner and outer annular link walls includes at least one coolant fluid circuit extending between the radially inner and outer ends of said first annular link wall.

A turbine engine including a combustion chamber having an inner annular shroud and an outer annular shroud that are coaxial with each other and that are connected at their downstream ends respectively to an inner annular link wall and to an outer annular link wall, for linking respectively to an inner casing and to an outer casing. At least a first one of the inner and outer annular link walls includes at least one coolant fluid circuit extending between the radially inner and outer ends of said first annular link wall.

A gas turbine engine combustion chamber includes upstream and downstream ring structures and a plurality of circumferentially arranged combustion chamber segments. Each segment extends the full length of the combustion chamber and each segment is secured to the upstream ring structure and is mounted on the downstream ring structure. The upstream end of each combustion chamber segment includes a surface having a plurality of circumferentially spaced radially extending holes and the upstream ring structure having a plurality of circumferentially spaced holes extending radially through a portion abutting the surface of the upstream end of each combustion chamber segment. Each combustion chamber segment being removably secured to the upstream ring structure by a plurality of fasteners locatable in the holes in the combustion chamber segment and corresponding holes in the upstream ring structure.

A gas turbine engine combustion chamber includes upstream and downstream ring structures and a plurality of circumferentially arranged combustion chamber segments. Each segment extends the full length of the combustion chamber and each segment is secured to the upstream ring structure and is mounted on the downstream ring structure. The upstream end of each combustion chamber segment includes a surface having a plurality of circumferentially spaced radially extending holes and the upstream ring structure having a plurality of circumferentially spaced holes extending radially through a portion abutting the surface of the upstream end of each combustion chamber segment. Each combustion chamber segment being removably secured to the upstream ring structure by a plurality of fasteners locatable in the holes in the combustion chamber segment and corresponding holes in the upstream ring structure.