Conventional gas turbine engines are generally optimized to operate at nearly a fixed speed with fixed blade geometries for the design operating condition. When the operating condition of the engine changes, the flow incidence angles may not be optimum with the blade geometries resulting in reduced off-design performance. By contrast, according to embodiments of the present invention, articulating the pitch angle of turbine blades in coordination with adjustable nozzle vanes improves performance by maintaining flow incidence angles within the optimum range at all operating conditions of a gas turbine engine. Maintaining flow incidence angles within the optimum range can prevent the likelihood of flow separation in the blade passage and also reduce the thermal stresses developed due to aerothermal loads for variable speed gas turbine applications.

A propeller oil control system for a turboprop engine of an aircraft includes an engine control unit and a propeller oil controller. The engine control unit is operable to determine a flight phase of the aircraft and is configured to supply control commands. The propeller oil controller is coupled to receive a supply of oil and to discharge the oil at a discharge oil pressure. The propeller oil controller includes an electrohydraulic servo valve that receives the control commands moves to a plurality of positions between a first position and a mid-position, and a plurality of positions between the mid-position and a second position. The engine control unit only commands the electrohydraulic servo valve to move out of the first position when the engine control unit determines the aircraft is conducting a take-off roll or the aircraft is in flight.

Hydraulic pitch actuation mechanism includes vaned rotor within timing chamber surrounded by annular timing chamber wall within fan hub and variable area and volume retarding and advancing chambers within timing chamber. Timing pocket walls extend inwardly from timing chamber wall and interdigitated with timing vanes extending outwardly from vane shaft of vaned rotor. Hydraulic retarding and advancing fluid passages extend through fan drive shaft and through the fan hub to the retarding and advancing chambers respectively. Fluid passages include annular axially spaced apart retarding and advancing passage discharge sections, annular and axially spaced apart retarding and advancing passage entry sections spaced apart from and aft of retarding and advancing passage discharge sections respectively. Retarding and advancing connecting passage sections extend through fan drive shaft and fluidly connect retarding and advancing passage entry sections to retarding and advancing passage discharge sections. Hub passages through fan hub connect passage discharge sections to chambers.

A propeller assembly of an aircraft includes a hub, a plurality of propeller blades extending from the hub and secured thereto and a propeller blade pitch change system located at at least one propeller blade of the plurality of propeller blades. The propeller blade pitch change system includes a pitch change actuator located in the propeller blade, and a drive mechanism operably connected to the pitch change actuator and to the propeller blade to urge rotation of the propeller blade about a propeller blade axis.

Rotor assembly apparatus are disclosed. An example rotor assembly includes a twist actuator to drive a first rotation of a first shaft about a first axis, the twist actuator positioned at a center of rotation of the rotor assembly. A first gear assembly to convert the first rotation into a plurality of second rotations of a plurality of second shafts. Each of the second shafts to provide torque to a respective blade coupled to the rotor assembly.

A turbine blade is provided with a rotatable leading edge to reduce pressure loss as a flow direction of combustion gas varies. The turbine blade includes an airfoil including as separate bodies a trailing-edge portion and a leading-edge portion being linked to the trailing-edge portion and disposed upstream of the trailing-edge portion, the leading-edge portion including a front surface arranged on an upstream side of the leading-edge portion; and a rotary unit connected to the leading-edge portion and configured to rotate the leading-edge portion according to an inflow angle of the combustion gas such that the front surface faces a flow of the combustion gas. A barrier wall extends from a side surface of the leading-edge portion toward an open end of the trailing edge portion, and when the leading-edge portion is rotated, the barrier wall portion prevents formation of a gap between the side surface and the open end.

A system and method for determining a position of a feedback ring of a propeller of an aircraft engine are provided. The feedback ring is coupled to the propeller to rotate with the propeller and to be displaced along a longitudinal axis with adjustment of a blade angle. An engagement member is configured to engage the feedback ring and to be displaced along a longitudinal direction substantially parallel to the longitudinal axis with displacement of the feedback ring. A sensor comprises a first member coupled to the engine and a second member coupled to the engagement member. The second member is moveable relative to the first member along the longitudinal direction as the engagement member is displaced. The sensor generates a signal indicative of a longitudinal position of the second member relative to the first member. A controller determines an axial position of the feedback ring from the sensor signal.

A device for controlling variable-pitch turbomachine members such as stator vanes, wherein the movement of a control mechanism is imparted to a ring that rotates the levers for pivoting the vanes about a stator, and takes place with a possible change in the length of the bottle screw connecting the ring to the control mechanism, thereby allowing the blade pitch angle law to be changed according to the extension of the control device. A greater freedom of adjustment is thus obtained. The device can be useful when a single control mechanism is used for multiple rings arranged side by side, making it possible to have different control laws for the rings.

A variable pitch fan assembly includes variable pitch fan blades circumscribed about engine centerline axis coupled to a drive shaft centered about the engine centerline axis. Each blade pivotable about pitch axis perpendicular to centerline axis and having blade turning lever connected thereto. One or more linear actuators non-rotatably mounted parallel to engine centerline axis and operably linked to fan blades for pivoting fan blades and connected to spider ring through thrust bearings for transmission of axial displacement of non-rotatable actuator rods of actuators while the fan blades are rotating. Spider arms extending away from spider ring towards blade roots and each spider arm connected to one of the turning levers. Turning levers may be connected and caromed to spider arms by pin and slot joint. Each spider arm may include joint pin disposed through joint slot of turning lever. Joint slot may be angled or curved.

Systems and methods are disclosed for controlling the pitch angle of a propeller and rotor assembly that minimizes circumferential loads and stresses to a pitch angle control system. The system may generally include an annular actuator, load transfer bearing (LTB), and a guide shaft is pivotally attached to the LTB to direct the LTB along an arcuate path relative to a rotor frame.