A turbofan engine is disclosed that includes a fan rotatable about an axis, a compressor section including a high pressure compressor, a medium pressure compressor and a low pressure compressor and a turbine section including a high pressure turbine, an intermediate turbine and a fan drive turbine. A fan drive gear system is driven by the fan drive turbine for driving the fan. A compressor drive gear system is driven by the intermediate turbine for driving the low pressure compressor. A gear controller controls rotation of at least one of the fan drive gear system and the compressor drive gear system relative to a static structure to vary an effective speed reduction ratio of one of the fan drive gear system and the compressor drive gear system.

A system for discharging compressed air from a compressor includes a air distribution manifold that is in fluid communication with the compressor via a conduit and at least one discharge line that is in fluid communication with the air distribution manifold. The discharge line defines a flow path from the air distribution manifold to atmosphere. The discharge line comprises a coupling pipe that is coupled to the air distribution manifold, a sparger section that is disposed downstream from the coupling pipe and at least one restrictor plate that is disposed between the coupling pipe and the sparger section within the flow path. The restrictor plate comprises at least one aperture that provides a pressure drop of the compressed air between the air distribution manifold and the sparger section.

The present disclosure is directed to a rotor thrust balanced turbine engine that may increase engine performance and efficiency while managing thrust mismatch or imbalance in a low pressure (LP) spool between a fan assembly and a turbine rotor assembly. The gas turbine engine defines a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a turbine rotor assembly and a turbine frame. The turbine rotor assembly defines a first flowpath radius and a second flowpath radius each extended from the axial centerline. The first flowpath radius is disposed at the upstream end of the turbine rotor assembly, and wherein the second flowpath radius is disposed at the downstream end of the turbine rotor assembly. The turbine frame and the turbine rotor assembly together define a seal interface radius inward of the turbine rotor assembly along the radial direction and concentric to the axial centerline, and wherein the turbine rotor assembly defines a ratio of the first flowpath radius to the seal interface radius less than or equal to approximately 1.79.

A unique low cost and efficient submersible, hermetically sealed, variable speed system intended to drive submersible boosting units. The system includes a unique combination of a liquid filled electrical motor connected to a hydraulic coupling and a magnetic coupling driver section, in a hermetically sealed container, with a magnetic coupling follower driving a booster unit. The system further includes integrated cooling, lubrication and control functionality. The drive unit has an actuating system connected to internal guide vanes which controls the liquid flow between the pump impeller and turbine wheel of the hydrodynamic coupling and hence the torque and speed. The combined system is a sealed seal-less and topside-less submersible drive unit that can operate in harsh subsea environments. The drive unit opens up for use of thin walled pressure casings and low pressure electrical penetrators.

An aircraft system includes, among other things, an aircraft and a gas turbine engine coupled to the aircraft. The gas turbine engine includes a propulsor section including a propulsor, a compressor section, a turbine section including a first turbine and a second turbine, and a gear reduction between the propulsor and the second turbine. The second turbine includes a number of turbine blades in each of a plurality of rows of the second turbine. The second turbine blades operating at least some of the time at a rotational speed. The number of blades and the rotational speed being such that the following formula holds true for a majority of the blade rows of the second turbine: 5500 Hz(number of bladesspeed)/60 sec10000 Hz. The gas turbine engine is rated to produce 15,000 pounds of thrust or more.

A compressor recirculation valve having a noise-suppressing muffler. The muffler has a double-walled construction, including a perforated outer wall and a non-perforated inner wall. The muffler is arranged surrounding the RCV such that when the RCV valve member is in an open position, compressed recirculation air is constrained to flow along the non-perforated inner wall and past an upper end thereof before reaching and flowing radially outwardly through the perforated outer wall to a transverse flow passage for delivery back into an inlet of the compressor.

A control system for a gas turbine engine includes a processing system operable to control a speed of the gas turbine engine and a memory system. The memory system is operable to store instructions executable by the processing system to determine at least one performance parameter associated with a stall condition of the gas turbine engine and to incrementally adjust an acceleration rate of the gas turbine engine based on detecting a degraded stall line limit according to the at least one performance parameter.

A turbo-machine, which can be operated in an optimized driving range is provided. To this end, a method for operating a turbo-machine having at least one turbo-machine stage, which has at least one rotary shaft is disclosed. According to the method, the following method steps are carried out: a) determining a desired efficiency characteristic value .sub.soll of the turbomachine stage; b) determining an actual efficiency characteristic value .sub.ist of the turbo-machine-stage; c) determining a comparison efficiency characteristic value of the turbo-machine stage by comparing the actual efficiency characteristic value .sub.ist and the desired efficiency characteristic value .sub.soll to one another; and d) changing at least one operating parameter of the turbo-machine stage subject to the comparison efficiency characteristic value .sub.vgl, wherein in order to determine the actual efficiency characteristic value .sub.ist, a measuring of a torque of the rotary shaft of the turbo-machine-stage is carried out.

A reverse flow gas turbine engine has a low pressure (LP) spool and a high pressure (HP) spool arranged sequentially in an axial direction. The LP spool comprises an LP compressor disposed forward of an LP turbine and drivingly connected thereto via an LP compressor gear train. The HP spool comprises an HP compressor in flow communication with the LP compressor, and an HP turbine disposed forward of the HP compressor and drivingly connected thereto via an HP shaft.

A method for controlling an inlet-adjustment mechanism in an air inlet for a compressor so as to switch the mechanism in a binary fashion between two positions P1 and P2 for adjusting a flow area of the inlet. The method includes identifying a threshold line on a compressor map of pressure ratio versus corrected flow rate for the compressor. The threshold line is a line on which the pressure ratio and flow rate of the compressor are the same for the P1 and P2 positions of the inlet-adjustment mechanism at equal speeds. When the operating point of the compressor on the compressor map crosses the threshold line, the inlet-adjustment mechanism is switched from one of its binary positions to the other.