A gas turbine engine section includes a plurality of spaced rotor stages, with a static guide vane intermediate the spaced rotor stages. The static guide vane provides swirl into air passing toward a downstream one of the spaced rotor stages, and an outer housing surrounding the spaced rotor stages. A diverter diverts a portion of air radially outwardly through the outer housing, and across at least one heat exchanger. The diverted air passes back into a duct radially inwardly through the outer housing, and is exhausted toward the downstream one of the spaced rotor stages.

A gas turbine engine includes a first spool, a second spool, a primary combustor, and a diffuser. The first spool includes a first compressor rotationally driven by a first turbine via a first shaft. The second spool includes a second compressor driven by a second turbine via a second shaft. The first compressor, the diffuser, and the primary combustor are arranged in series to provide a compressed airflow discharged from the first compressor to the primary combustor via the diffuser, which includes walls that diverge towards the primary combustor. The second compressor is fluidly coupled to the diffuser to receive at least a portion of the compressed airflow from the diffuser. The second turbine is fluidly coupled to the diffuser to discharge an expanded airflow to the diffuser.

Systems and methods for adjusting airflow distortion in a gas turbine engine using a valved airflow passage assembly are provided. A gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. The compressor section can include a compressor. The gas turbine engine can further include a valved airflow passage assembly comprising a valve and a duct, the duct defining an inlet in airflow communication with the engine airflow path at a location downstream of the compressor and an outlet in airflow communication with the engine airflow path at a location upstream of the compressor, the duct comprising an airflow passage extending between the inlet and outlet. The valve can be operable with the airflow passage for controlling an airflow through the airflow passage to adjust airflow distortion.

Systems and methods for adjusting airflow distortion in a gas turbine engine using a valved airflow passage assembly are provided. A gas turbine engine can include a compressor section, a combustion section, and a turbine section in series flow and defining at least in part an engine airflow path. The compressor section can include a compressor. The gas turbine engine can further include a valved airflow passage assembly comprising a valve and a duct, the duct defining an inlet in airflow communication with the engine airflow path at a location downstream of the compressor and an outlet in airflow communication with the engine airflow path at a location upstream of the compressor, the duct comprising an airflow passage extending between the inlet and outlet. The valve can be operable with the airflow passage for controlling an airflow through the airflow passage to adjust airflow distortion.

Bleed air systems for use with aircraft and related methods are disclosed. An example apparatus includes a compressor having a compressor inlet and a compressor outlet. The compressor inlet to receive airflow from a first air supply source. An air mixing device having a first mixer inlet to receive compressed air from the compressor outlet and a second mixer inlet to receive bleed air from a bleed air system. The bleed air to provide a motive fluid to enable the air mixing device to mix the bleed air and the compressed air to produce mixed air for the anti-icing system.

Bleed air systems for use with aircraft and related methods are disclosed. An example apparatus includes a compressor having a compressor inlet and a compressor outlet. The compressor inlet to receive airflow from a first air supply source. An air mixing device having a first mixer inlet to receive compressed air from the compressor outlet and a second mixer inlet to receive bleed air from a bleed air system. The bleed air to provide a motive fluid to enable the air mixing device to mix the bleed air and the compressed air to produce mixed air for the anti-icing system.

A hybrid electric gas turbine engine includes a fan section having a fan, a turbine section having a turbine drivably connected to the fan through a main shaft that extends along a central longitudinal axis, a gas generating core extending along a first axis that is radially offset from the central longitudinal axis, a first electric motor drivably connected to the main shaft, wherein the electric motor is colinear with the main shaft, and an electric compressor extending along a second axis that is radially offset from the central longitudinal axis, the electric compressor in fluid communication with the second turbine section.

A hybrid electric gas turbine engine includes a fan section having a fan, a turbine section having a turbine drivably connected to the fan through a main shaft that extends along a central longitudinal axis, a gas generating core extending along a first axis that is radially offset from the central longitudinal axis, a first electric motor drivably connected to the main shaft, wherein the electric motor is colinear with the main shaft, and an electric compressor extending along a second axis that is radially offset from the central longitudinal axis, the electric compressor in fluid communication with the second turbine section.

A two-shaft gas turbine includes a compressor, a high-pressure turbine including a first shaft connected to a rotation shaft of the compressor, a low-pressure turbine including a second shaft separate from the first shaft, and is provided coaxially to the high-pressure turbine with an interval in an axis direction between the low-pressure turbine and the high-pressure turbine, an intermediate flow path provided between a final stage rotor blade of the high-pressure turbine and a first stage rotor blade of the low-pressure turbine in the axis direction, the intermediate flow path being configured to supply a combustion gas from the high-pressure turbine to the low-pressure turbine, and a strut disposed inside the intermediate flow path. The strut concurrently functions as a first stage stator blade of the low-pressure turbine.

A process for retrofitting an industrial gas turbine engine of a power plant where an old industrial engine with a high spool has a new low spool with a low pressure turbine that drives a low pressure compressor using exhaust gas from the high pressure turbine, and where the new low pressure compressor delivers compressed air through a new compressed air line to the high pressure compressor through a new inlet added to the high pressure compressor. The old electric generator is replaced with a new generator having around twice the electrical power production. One or more stages of vanes and blades are removed from the high pressure compressor to optimally match a pressure ratio split. Closed loop cooling of one or more new stages of vanes and blades in the high pressure turbine is added and the spent cooling air is discharged into the combustor.