Abstract: A gas turbine engine may include a core. A gas turbine engine may include a nacelle at least partially surrounding the core. A gas turbine engine may include a flow bypass channel defined between an inner surface of the nacelle and an outer surface of the core. A gas turbine engine may include an axial shaft extending along a major axis of the core. A gas turbine engine may include a fan positioned at a leading end of the axial shaft. A gas turbine engine may include a reduction gear assembly coupling the fan to the axial shaft. A gas turbine engine may include a turbine directly coupled to the axial shaft and positioned downstream relative to the combustion chamber, wherein the reduction gear assembly functions to reduce rotation of the fan relative to the axial shaft.

Abstract: A pump assembly may include a drive mechanism and a rotatable shaft coupled to the drive mechanism. A pump may be carried by the rotatable shaft. The pump may include at least one pump stage for communication of a lubricant. A separator may be driven by the rotatable shaft. The separator may be operable to separate oil from a mixture of air and oil. A method of assembly is also disclosed.

Abstract: A method of operation is provided during which a heating fluid is directed through a first passage into a combustor section to heat a component within the combustor section when an aircraft engine is in a first standby mode. The aircraft engine includes a flowpath, a compressor section, the combustor section, a turbine section and an exhaust section. The flowpath extends through the compressor section, the combustor section, the turbine section and the exhaust section from an inlet into the flowpath to an exhaust from the flowpath. The heating fluid is directed through a second passage into the exhaust section when the aircraft engine is in a second mode. The heating fluid is directed through the second passage bypasses the turbine section.