Methods and systems for operating an engine having at least first and second variable geometry mechanisms are provided. A spool-specific ratio for at least one spool of the engine is determined, wherein the spool-specific ratio relates to an aerodynamic parameter. The spool-specific ratio is compared to a reference to determine a ratio discrepancy for the at least one spool. An engine-specific ratio relating to the aerodynamic parameter is determined. At least one of a position of the first variable geometry mechanism and a position of the second variable geometry mechanism is adjusted based on the engine-specific ratio and the ratio discrepancy to reduce the ratio discrepancy.

The method can include, in sequence: varying the flow rate of fuel from the first flow rate to a second flow rate, thereby varying the rotor speed from a first speed to a second speed, varying the flow rate of fuel back to the first flow rate, and rotating the rotor at the first speed for a given period of time.

A bleed valve includes a housing with an inlet coupled to an outlet by a duct, a guide tube with an orifice fixed in the housing between the inlet and the outlet, a piston, and baffle. The piston is slideably supported on the guide tube and is movable between an open and a closed position, the duct fluidly coupling the inlet and outlet in the open position, the duct fluidly separating the inlet and outlet in the closed position. The orifice fluidly couples the inlet and outlet in the open and closed positions to move piston between the open and closed positions according to differential pressure between the bleed valve inlet and outlet. The baffle is slideably supported by the guide tube to set the differential pressure at which the piston moves between the open and closed positions. Gas turbines and differential pressure adjustment methods are also described.

A method of maintaining rotor tip clearance and compressor operational line during a transient operation of a gas turbine engine is disclosed. In various embodiments, the method includes applying high spool auxiliary power to a high speed spool for a first time period, applying low spool auxiliary power to a low speed spool for a second time period, sensing one or more operational parameters of the gas turbine engine during the transient operation, and ceasing application of power to the high speed spool, based on the one or more operational parameters.

A bleed valve includes a housing with an inlet coupled to an outlet by a duct. A guide tube is fixed within the housing between the inlet and the outlet. A piston with a piston orifice is slideably supported on the guide tube and movable between an open position and a closed position. The duct fluidly couples the inlet to the outlet in the open position, the duct fluidly separates the inlet from the outlet in the closed position, and the piston orifice fluidly couples the inlet with the outlet in the open position and the closed position to move piston between the open position and the closed position according to differential in pressure between the inlet and the outlet of the bleed valve. Compressors, gas turbine engines, and methods of controlling fluid flow are also described.

A method of operation for a gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, reducing a rotational speed of a fan relative to a shaft through a gear train, driving the shaft with a low pressure turbine, driving a high pressure compressor with a high pressure turbine, communicating airflow from the fan through a bypass passage defined by a nacelle, the nacelle extending along an engine axis and surrounding the fan, discharging the airflow through a variable area fan nozzle defining a discharge airflow area, detecting an airfoil flutter condition associated with adjacent airfoils of the fan, and moving the variable area fan nozzle to vary the discharge airflow area and mitigate the airfoil flutter condition.

A turbocharger for an internal combustion engine has a bearing housing, in which a rotor shaft is mounted in a rotatable manner. A compressor having a compressor wheel is arranged for conjoint rotation on the rotor shaft. A fresh air supply channel conducts a fresh air mass flow to the compressor wheel. The fresh air supply channel has a first flow cross section upstream of the compressor wheel. A flow control device is provided and is adjustable between an open position, in which the first flow cross section is opened up, and a closed position, in which the first flow cross section is reduced to a second flow cross section. The flow control device is fluidically coupled to a compressor channel of the compressor downstream of the compressor wheel, such that the flow control device is adjusted in a manner dependent on a pressure prevailing in the compressor channel.

A gas turbine engine with a compressor supplying compressed air. A combustor receives the compressed air and fuel and generates a flow of combusted gas. A turbine receives a core flow of the combusted gas to rotate a turbine rotor. A turbine inlet nozzle directs the combusted gas to the turbine rotor. Vanes are disposed in the turbine inlet nozzle and rotate to vary a flow area through which the core flow passes. The vanes adjust a pressure ratio of the gas turbine engine to compensate for changing operational requirements of the gas turbine engine by rotating to positions matching the changing operational requirements.

The method can include, in sequence: varying the flow rate of fuel from the first flow rate to a second flow rate, thereby varying the rotor speed from a first speed to a second speed, varying the flow rate of fuel back to the first flow rate, and rotating the rotor at the first speed for a given period of time.

Instrumented airflow path components configured to determine airflow path distortion in an airflow path of a gas turbine engine (e.g., using for propulsion of an aircraft) are provided. In one embodiment, a gas turbine engine for an aircraft can include a compressor section, a combustion section, and turbine section in series flow. The compressor section, combustion section, and turbine section define at a portion of an engine airflow path for the gas turbine engine. The gas turbine engine further includes one or more members extending at least partially into the engine airflow path of the gas turbine engine and one or more pressure sensor devices at least partially integrated into the one or more members extending at least partially into the engine airflow path. The one or more pressure sensor devices are configured to obtain measurements for determining a distortion condition for the gas turbine engine.