A system for lubricating components of a gas turbine engine includes a supply pump to pump lubricant from a reservoir through a supply conduit to lubricated component. An outer body deicing circuit is fluidly coupled in series with the supply conduit. Moreover, a scavenge conduit extends from the lubricated component to the reservoir and a scavenge pump pumps the lubricant from the lubricated component through the scavenge conduit. In addition, a valve is fluidly coupled in series with the supply conduit and the scavenge conduit and a bypass conduit extends from the valve to a location on the supply conduit downstream of the deicing circuit. When the lubricant is supplied to the valve, the valve directs the lubricant flowing through the scavenge conduit into the reservoir. When the supply of lubricant to the valve is halted, the valve directs the lubricant flowing through the scavenge conduit into the bypass conduit.

An airfoil for a fan section of a turbine engine may include a fan blade or an outlet guide vane formed of a first material, and an edge guard disposed about an edge of the fan blade. The edge guard may include a matrix composite that has a toughness that is greater than a toughness of the first material. The airfoil may include a fan blade or an outlet guide vane. The first material of the airfoil may include a metal alloy and/or a matrix composite. A method of manufacturing an airfoil for a fan section of a turbine engine may include manufacturing an edge guard, attaching the edge guard to the airfoil.

An airfoil for a fan section of a turbine engine may include a fan blade or an outlet guide vane, and an edge guard attached thereto. The edge guard may include a heating conduit disposed within at least a portion of the edge guard. An anti-icing system for a plurality of fan blades or outlet guide vanes may include a fluid supply pathway configured to supply heating fluid to respective ones of a plurality of heating conduits within the edge guards attached to respective ones of a plurality of fan blades and/or to a plurality of outlet guide vanes. The heating fluid may include bleed air from a core air flowpath. A method of inhibiting icing on an airfoil may include flowing a heating fluid into a heating conduit disposed within an edge guard attached to the airfoil and heating the edge guard with the heating fluid.

An airfoil for a fan section of a turbine engine may include a fan blade or an outlet guide vane, and an edge guard attached thereto. The edge guard may include a heating conduit disposed within at least a portion of the edge guard. An anti-icing system for a plurality of fan blades or outlet guide vanes may include a fluid supply pathway configured to supply heating fluid to respective ones of a plurality of heating conduits within the edge guards attached to respective ones of a plurality of fan blades and/or to a plurality of outlet guide vanes. The heating fluid may include bleed air from a core air flowpath. A method of inhibiting icing on an airfoil may include flowing a heating fluid into a heating conduit disposed within an edge guard attached to the airfoil and heating the edge guard with the heating fluid.

A gas turbine engine includes: a fan rotating about an engine main axis; a core duct; an engine core; an Engine Section Stator (ESS) including a plurality of ESS vanes and arranged in the core duct downstream of the fan; and a plurality of variable inlet guide vanes (VIGV) adapted to rotate about a pivot axis and arranged in the core duct downstream of the ESS. The VIGV vanes are arranged angularly rotated with respect to the ESS vanes such that the VIGVs are shielded by the ESS, thereby protecting the VIGVs from icing and from ice shedding from the ESS vanes.

A gas turbine engine includes: a fan rotating about an engine main axis; a core duct; an engine core; an Engine Section Stator (ESS) including a plurality of ESS vanes and arranged in the core duct downstream of the fan; and a plurality of variable inlet guide vanes (VIGV) adapted to rotate about a pivot axis and arranged in the core duct downstream of the ESS. The VIGV vanes are arranged angularly rotated with respect to the ESS vanes such that the VIGVs are shielded by the ESS, thereby protecting the VIGVs from icing and from ice shedding from the ESS vanes.

A measurement system for an aircraft gas turbine engine includes a probe and a heated-gas source in fluid communication with the pressure probe. The probe includes a probe body defining an internal cavity of the probe. The probe further includes a plurality of sensor inlet ports extending through the probe body and configured to receive a sensed fluid flow. The probe further includes a plurality of probe conduits. Each probe conduit of the plurality of probe conduits is coupled to a respective sensor inlet port of the plurality of sensor inlet ports and extending from the respective sensor inlet port to an exterior of the probe body. The heated-gas source is configured to supply a heated gas flow to one or both of: the plurality of sensor inlet ports via the plurality of probe conduits and an interior of the probe body outside of the plurality of probe conduits.

A booster splitter for a gas turbine engine and a method of additively manufacturing the booster splitter are provided. The booster splitter includes an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine, an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path, an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port, and a passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end.

A booster splitter for a gas turbine engine and a method of additively manufacturing the booster splitter are provided. The booster splitter includes an annular inner wall defining a radially outer boundary of a compressor flow path defined through a compressor section of the gas turbine engine, an annular outer wall spaced apart from the annular inner wall along the radial direction, the annular outer wall adjacent to the annular inner wall at a forward end, the forward end defining an inlet to the compressor flow path, an annular bulkhead spanning between the annular inner wall and the annular outer wall substantially along the radial direction, the bulkhead defining an inlet port, and a passageway defined within the annular outer wall, the passageway extending from the inlet port, into the bulkhead, radially outward to the outer wall, and through the annular outer wall towards the inlet defined by the forward end.

A compressor cleaning apparatus to clean a compressor of a gas turbine is provided. The compressor cleaning apparatus includes a nozzle configured to inject a cleaning fluid into an interior of a compressor, a fluid supply tube connected to the nozzle to supply the cleaning fluid to the nozzle, a first cleaning fluid supply connected to the fluid supply tube to supply a first cleaning fluid, and a second cleaning fluid supply connected to the fluid supply tube to supply a second cleaning fluid having a temperature higher than that of the first cleaning fluid.