A hydraulic position control arrangement with a base body which is rotatably supported on a shaft that includes a hydraulic fluid passage for the admission of hydraulic fluid to an operating chamber which is delimited by the base body and an annular piston which is axially movably supported by the base body, includes an annular sealing element disposed in the base body and seated on the shaft and a leakage return path which extends from the area outside the shaft to the operating chamber through the sealing element. The shaft has at least two axial sections with different outside diameters with the sealing element being seated on the axial shaft section of smaller diameter. The sealing element has a seal lip which abuts the outside of the shaft and is axially supported by a support ring which is provided in the base body.

An impeller (20, 120) for a ducted fan arrangement (10, 110), the impeller (20, 120) including a hub (24, 124) and a plurality of blades (26, 126) extending radially from the hub (24, 124), each of the plurality of blades (26, 126) including a root (28, 128) proximate the hub (24, 124) and a tip (30, 130). A camber of each of the plurality of blades (26, 126) is arranged to flatten or reduce between the root (28, 128) and the tip (30, 130). A fan arrangement (10, 110) including such an impeller (20, 120) is also disclosed.

A gas turbine engine defining an axial direction and a radial direction includes a spinner defining a spinner duct and a spinner inlet to the spinner duct, a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a fan duct, a fan duct inlet to the fan duct, a core duct, and a core inlet to the core duct, and a primary fan driven by the turbomachine, wherein the spinner inlet is upstream of the primary fan.

A turbofan engine for an aircraft includes a fan and a fan actuation system. The fan has a plurality of fan blades coupled to a fan shaft having one or more fan bearings. The fan blades are rotatable about a pitch axis. The fan actuation system is disposed within a fan hub and includes one or more actuators for rotating the fan blades about the pitch axis and one or more radial thrust bearings. The fan actuation system is characterized by a fan actuation system length envelope in a range from 8.5 to 24 and given

[00001]$\frac{N_{\mathrm{FB}}{D}_{\mathrm{FT}}}{L_{\mathrm{AXIAL}}\left(\frac{R_{\mathrm{TB}}}{N_{\mathrm{FB}}}\right)}.$

N.sub.FB is a number of the fan blades, D.sub.FT is a fan tip diameter of the fan blades, R.sub.TB is a thrust bearing radius of the radial thrust bearings, and L.sub.AXIAL is an axial length from a fan hub tip to the fan bearings.

A propulsor assembly for a turbine engine includes a propulsor and a propulsor actuation system. The propulsor has a plurality of propulsor blades. Each of the plurality of propulsor blades is rotatable about a blade pitch axis. The propulsor actuation system includes an actuator for rotating the plurality of propulsor blades and a trunnion mechanism that includes a plurality of trunnion assemblies. Each of the plurality of trunnion assemblies is coupled to a respective one of the plurality of propulsor blades and includes an outer sleeve coupled to a blade spar of the respective propulsor blade, an inner sleeve coupled to the outer sleeve, and an actuation member that engages the inner sleeve. The actuation member is rotatably engageable by the actuator to rotate the respective propulsor blade about the blade pitch axis.

A gas turbine engine includes a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order. A fan defining a fan axis and comprising a plurality of fan blades is rotatable about the fan axis. A pitch change mechanism is operable with the plurality of fan blades to control a pitch of the plurality of fan blades. A counterweight is connected to each respective fan blade. A spring mechanism is operably associated with each fan blade of the plurality of fan blades to create a spring twisting moment on the respective fan blade that is counter to a centrifugal twisting moment of the respective fan blade.

A low pitch protection control system and method that includes sensing hardware, an engine controller, and a pitch actuation system for variable pitch fan blades of a turbine engine is provided. The engine controller receives sensed data from the sensing hardware related to current flight conditions, determines, based on at least some of the sensed data, a first pitch angle threshold value for initiating a first mitigating action to prevent onset of excessive drag at the current conditions, monitors a current pitch angle of one or more variable pitch fan blades of the turbine engine relative to the first pitch angle threshold value, and commands the pitch actuation system to perform the first mitigating action in response to the current pitch angle being less than or equal to the first pitch angle value.

Trunnions having multiple fan blades for use with an aircraft engine are disclosed herein. An example comprises a trunnion rotatable within a hub of an aircraft engine, a first fan blade coupled to the trunnion, the first fan blade having a pitch axis, and a second fan blade coupled to the trunnion, the second fan blade having a longitudinal axis, the longitudinal axis laterally offset from the pitch axis.

An impeller (20, 120) for a ducted fan arrangement (10, 110), the impeller (20, 120) including a hub (24, 124) and a plurality of blades (26, 126) extending radially from the hub (24, 124), each of the plurality of blades (26, 126) including a root (28, 128) proximate the hub (24, 124) and a tip (30, 130). A camber of each of the plurality of blades (26, 126) is arranged to flatten or reduce between the root (28, 128) and the tip (30, 130). A fan arrangement (10, 110) including such an impeller (20, 120) is also disclosed.