A turboshaft engine for a rotorcraft includes a first spool and an independently rotatable second spool. The first spool includes a low pressure compressor, a first set of variable guide vanes disposed at an entry of the low pressure compressor, and a low pressure turbine drivingly engaged to the low pressure compressor. The second spool includes a high pressure compressor, a second set of variable guide vanes disposed at an entry of the high pressure compressor, the second set of variable guide vanes independently operable relative to the first set of variable guide vanes, and a high pressure turbine drivingly engaged to the high pressure compressor. One or both of the low pressure compressor and the high pressure compressor includes a mixed flow rotor.

The gas turbine engine for an aircraft includes at least a low pressure spool with a low pressure turbine shaft operatively connected to at least one turbine, the low pressure turbine shaft rotatable about an engine axis, and a low pressure compressor operatively connected to a low pressure compressor shaft that is independently rotatable relative to the low pressure turbine shaft. A differential gearbox has an input operatively connected to the low pressure turbine shaft, a first output and a second output, the first output of the differential gearbox operatively connected to the low pressure compressor shaft and the second output of the differential gearbox operatively connected to an output shaft of the gas turbine engine. The differential gearbox permits the output shaft, the low pressure compressor shaft and the low pressure turbine shaft to rotate at different speeds.

Turnbuckle linkages are disclosed herein. A turnbuckle linkage includes an inside rod including a first opening on a first longitudinal end and a second opening on a second longitudinal end of the inside rod, the first opening and the second opening extending at least partially through the inside rod, the inside rod including tabs extending from the first longitudinal end and the second longitudinal end of the inside rod around the first opening and the second opening, an exterior surface of the tabs including conical threads, outside rods to be positioned within the openings of the inside rod, the tabs of the inside rod at least partially surrounding the outside rods, the outside rods including end bearings to couple to a first structure and a second structure, and conical nuts to constrict the tabs around the outside rods and couple the inside rod and the outside rods.

Methods and related apparatus for improving synchronization of two or more engines on an aircraft are disclosed. Such method may be used where each engine comprises a first spool and a second spool, and, where a rotational speed of a first spool of a first engine has been substantially synchronized with a rotational speed of a first spool of a second engine. An exemplary method comprises receiving a value of a sensed parameter useful in controlling the first engine; adding a bias to the value; and using the biased value for controlling the first engine to cause a change in rotational speed of the second spool of the first engine in relation to the rotational speed of the first spool of the first engine.

Systems and methods for controlling an unducted turbofan engine to limit noise. The unducted turbofan engine may include an unducted fan drivingly coupled with a low-pressure turbine and a plurality of unducted outlet guide vanes. The unducted fan may include a plurality of fan blades and a pitch angle of the fan blades may be variable. A pitch angle of the unducted outlet guide vanes may be variable. A controller is configured to control the engine to limit noise based on a noise sensitive condition.

The invention relates to a guide vane assembly as well as to a method for mounting a guide vane assembly of a turbomachine, including a number of adjustable guide vanes, the guide vane platform of which has a guide vane journal that is mounted in a receiving opening of a housing of the turbomachine, wherein a guide vane head of the guide vane has a bearing pin, which is mounted on an inner ring arranged on a rotor of the turbomachine. In the method, a guide vane journal, which is arranged on a guide vane platform of a guide vane, is brought into a receiving opening of a housing, which is arranged radially with respect to a rotor axis of the turbomachine, and an inner ring is provided on a rotor of the turbomachine.

A unison ring of a gas turbine engine includes: an annular body including fiber-reinforced resin or circular-arc bodies including the fiber-reinforced resin, the fiber-reinforced resin including resin and reinforced fibers; and pin holes in which a pin is in a radial direction orthogonal to an axial direction of the unison ring. A main orientation of the reinforced fibers of the fiber-reinforced resin is directed in a circumferential direction of the unison ring.

The method can include determining a mass flow rate W of working fluid circulating through the compressor stage, determining a control parameter value associated to the geometrical configuration of the variable geometry element based on the determined value of mass flow rate W; and changing the geometrical configuration of the variable geometry element in accordance with the determined control parameter value.

The method can include determining a mass flow rate W of working fluid circulating through the compressor stage, determining a control parameter value associated to the geometrical configuration of the variable geometry element based on the determined value of mass flow rate W; and changing the geometrical configuration of the variable geometry element in accordance with the determined control parameter value.

A jet engine which includes: a fan provided with a plurality of stages of rotor blades; a compressor which compresses air which is sent from the fan; a combustor which generates combustion gas by using compressed air generated by the compressor; a turbine which generates a driving force from the combustion gas; and a nozzle which discharges the combustion gas, the jet engine further includes: a variable guide vane which is disposed upstream of the rotor blades of a second and later stage of the rotor blades of the fan and which adjusts an inlet angle of air flow against the second and later stage of the rotor blades; a fluid resistance adjusting device which adjusts a fluid resistance at the nozzle; and a controller which controls the variable guide vane such that the inlet angle at the time of cruise flight is smaller than the inlet angle at the time of acceleration and controls the fluid resistance adjusting device such that increase in the fluid resistance, due to an increase in a volume flow at an outlet of the fan corresponding to a reduction in the inlet angle, is suppressed.