A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

A power transmission system includes a shaft, a stator disposed within the shaft and substantially concentric with the shaft; and at least one supporting element positioned between the stator and the shaft and configured to support the shaft on the stator to reduce a vibration of the shaft and allow the shaft to rotate relative to the stator. A gas turbine engine including the power transmission system is also described.

A power transmission system includes a shaft, a stator disposed within the shaft and substantially concentric with the shaft; and at least one supporting element positioned between the stator and the shaft and configured to support the shaft on the stator to reduce a vibration of the shaft and allow the shaft to rotate relative to the stator. A gas turbine engine including the power transmission system is also described.

A pressure relief latch includes a housing configured to be mounted to a pressure relief door. A bolt is coupled to the housing about a first axis and is rotatable about the first axis between a first bolt position and a second bolt position. The bolt includes a bearing centered about a second axis. The bearing is configured to rotate about the second axis. A spring assembly includes a first assembly end mounted to the bearing and a second assembly end in rotational communication with the housing about a third axis. The spring assembly is rotatable about the second and third axes and further includes a spring extending between the first and second assembly ends. The bolt is configured to permit movement of the pressure relief door from the closed position to the open position as the bolt rotates from the first bolt position to the second bolt position.

A pressure relief latch includes a housing configured to be mounted to a pressure relief door. A bolt is coupled to the housing about a first axis and is rotatable about the first axis between a first bolt position and a second bolt position. The bolt includes a bearing centered about a second axis. The bearing is configured to rotate about the second axis. A spring assembly includes a first assembly end mounted to the bearing and a second assembly end in rotational communication with the housing about a third axis. The spring assembly is rotatable about the second and third axes and further includes a spring extending between the first and second assembly ends. The bolt is configured to permit movement of the pressure relief door from the closed position to the open position as the bolt rotates from the first bolt position to the second bolt position.

A variable guide vane assembly is provided for a turbine defining a core air flowpath. The variable guide vane assembly includes an airfoil band defining a flowpath surface and a cavity. The variable guide vane assembly further includes an airfoil including a first end extending at least partially into the cavity of the airfoil band and an opposite second end, the airfoil extending generally along an axis between the first end and the second end and being moveable generally about the axis relative to the airfoil band. The variable guide vane assembly further includes a sealing element operable to form a seal between the first end of the airfoil and the airfoil band.

The present invention relates to a systems and methods for improved removal of one or more species in a process stream, such as combustion product stream formed in a power production process. The systems and methods particularly can include contacting the process stream with an advanced oxidant and with water.

The present invention relates to a systems and methods for improved removal of one or more species in a process stream, such as combustion product stream formed in a power production process. The systems and methods particularly can include contacting the process stream with an advanced oxidant and with water.

A burner assembly includes a plurality of burners for mixing fuel and air. Each of the plurality of burners includes: a fuel nozzle for injecting the fuel; a mixing passage supplied with the fuel and the air; and a support portion connecting a passage wall of the mixing passage and the fuel nozzle to support the fuel nozzle.

A seal system includes a ceramic component, a metallic component, a silicon-containing layer, and a barrier layer. The ceramic component has a first surface region that defines a first surface roughness. The metallic component is situated adjacent to the first surface region and has a second surface region facing the first surface region. The silicon-containing layer is on the first surface region of the ceramic component and has a contact surface that defines a second surface roughness which is less than the first surface roughness. The barrier layer is on the metallic component and in contact with the silicon-containing layer and serves to limit interaction between silicon of the silicon-containing layer and the metallic component. The barrier layer includes at least one of alumina or MCrAlY.