A method of extracting work from a convertible gas turbine engine having a core flowpath and a bypass flowpath. The method comprises operating the convertible gas turbine engine at a first volumetric flow rate through the core flowpath and a second volumetric flow rate through the bypass flowpath to produce a first work output of the convertible gas turbine engine; extracting the first work output via an unshrouded fan and a shaft at a first fan to shaft extraction ratio; altering the second volumetric flowrate through the bypass flowpath while maintaining the first work output; and extracting the first work output via an unshrouded fan and a shaft at a second fan to shaft extraction ratio.

A turbomachinery engine includes a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine includes 3-5 rotating stages. The low-pressure turbine includes an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. In some instances, the area ratio is within a range of 3.1-5.1.

A hydraulic system for controlling orientation of fan blades in a turbine engine including at least one set of adjustable pitch fan blades actuated by actuators that are connected to a hydraulic pump, and a small pitch protection valve controlled by an associated solenoid valve to make it possible on the ground, when the orientation of the blades is substantially perpendicular to the axis of the turbine engine, to pass from a first operating zone in which a speed of pitch variation of the blades is low to a second operating zone in which a speed of pitch variation is considerably higher, wherein operating pressure of the hydraulic pump is controlled by the solenoid valve controlling the small pitch protection valve to define two outlet levels corresponding to the two operating zones.

An apparatus is disclosed that includes a gas turbine engine including a first rotor blade axially adjacent a second rotor blade and an aperture formed in one of the first rotor blade and the second rotor blade and structured to emit a fluid therefrom. A fluid source is in flow communication with the aperture and configured to flow the fluid through the aperture.

A turbine blade for a gas turbine engine comprises an airfoil having a pressure side, a suction side, a span direction and a chord-wise direction. The airfoil has an airfoil span on a pressure line being a projection of the stacking line onto the pressure side. The airfoil has a plurality of chords extending between a leading edge and a trailing edge of the airfoil. A generally round dimple is disposed on the pressure side. The dimple is contained in an area extending on the stacking line between 0% and 23% of the airfoil span from the inner end, and in the chord-wise direction between 0% of a first chord and 82% of a second chord from the leading edge. The dimple is configured to initiate fracture of the blade at a predetermined speed of rotation. A method of preventing rupture of a disk of a turbine rotor is also presented.

Systems and methods are provided for an air-driven augmentor fan equipped aircraft propulsor. The augmentor fan may increase the effective bypass ratio of the aircraft propulsor and reduce fuel consumption and carbon emissions of the aircraft. The augmentor fan may be driven by air energized by a ducted fan powered by the core engine of the aircraft propulsor. The energized air may be received by an inlet, flowed through a flow path, and exhausted out the outlet to drive the augmentor fan. The exhausted energized air may impart a torque on the augmentor fan or blades of the augmentor fan. One or more of the inlet, flow path, or outlet may be variable in size to control the volume of air flowed through the flow path.

A blade for a fan of a turbomachine, for example of unducted fan type, a corresponding fan, and a corresponding turbomachine. The blade includes a mechanism arranged, at a single location, to locally, as the fan rotates, disturb a distribution of flow around the blade so as to form two independent main vortices downstream.

A thrust scheduling method for a gas turbine engine that includes a plurality of blades having a variable pitch beta angle is provided. The method can include receiving into a control system at least one condition input from a respective sensor; receiving into a control system a low pressure shaft speed from a low pressure shaft speed sensor; receiving a control command from a full authority digital engine control (FADEC) in the control system; generating a low pressure shaft speed base reference from a first schedule logic in the control system based upon the at least one condition input received and the control command received; generating a beta angle base reference from a second schedule logic from the at least one condition input received, the low pressure shaft speed, and the control command received; and supplying the low pressure shaft speed base reference and the beta angle base reference to an engine control system, wherein the engine control system adjusts at least the pitch angle of the plurality of fan blades or a fuel flow to the engine.

The described reduction gearbox of a gas turbine engine includes a first gear reduction stage having an input gear adapted to be driven by a turbine output shaft. The input gear transfers power received from the turbine output shaft laterally away from the input gear to an input speed gear. Each input speed gear engages an output speed gear to define a main speed reduction gear set, and the main speed reduction gear sets are laterally spaced apart from one another to define a gap. The gearbox has a second gear reduction stage driven by the output speed gears, the second stage adapted to drive an engine output shaft.

A propulsion device including an unducted upstream propeller and a counter-rotating and coaxial downstream propeller. At least one blade of the upstream propeller is configured to generate a corotating secondary vortex for acting upstream of the downstream propeller to destabilize a blade tip vortex of the same at least one blade.