An auxiliary power unit (APU) control system for an aircraft is disclosed and includes an APU, an air inlet having an effective area, an air inlet door moveable to vary the effective area of the air inlet, an actuator configured to move the air inlet door into a set position, one or more processors, and a memory coupled to the one or more processors. The memory stores data comprising a database and program code that, when executed by the one or more processors, causes the APU control system to receive one or more ambient signals indicative of an air density value. The system also determines the effective area of the air inlet based on the air density value. The system is further caused to instruct the actuator to move the air inlet door into the set position.

An auxiliary power unit (APU) control system for an aircraft is disclosed and includes an APU, an air inlet having an effective area, an air inlet door moveable to vary the effective area of the air inlet, an actuator configured to move the air inlet door into a set position, one or more processors, and a memory coupled to the one or more processors. The memory stores data comprising a database and program code that, when executed by the one or more processors, causes the APU control system to receive one or more ambient signals indicative of an air density value. The system also determines the effective area of the air inlet based on the air density value. The system is further caused to instruct the actuator to move the air inlet door into the set position.

A bushing includes cylindrical body depending from preferably annular polygonal flange, bore extending through flange and body, and at least one overload indicator. Overload indicator may be frangible such as frangible tab at distal end of indicator tab depending from body or frangible coupons extending between forward and aft sections of polygonal flange. Frangible tab and coupons may include distinctive mark discernible or visible under visible or ultraviolet light. Bushing may be used in unison ring of high pressure compressor assembly to connect lever arms, coupled to variable vanes, to unison ring. Pins disposed in bore of bushing couple arm to unison ring. Forward and aft overhangs may depend from forward and aft sections of polygonal flange and overhang forward and aft annular sides of unison ring.

A bushing includes cylindrical body depending from preferably annular polygonal flange, bore extending through flange and body, and at least one overload indicator. Overload indicator may be frangible such as frangible tab at distal end of indicator tab depending from body or frangible coupons extending between forward and aft sections of polygonal flange. Frangible tab and coupons may include distinctive mark discernible or visible under visible or ultraviolet light. Bushing may be used in unison ring of high pressure compressor assembly to connect lever arms, coupled to variable vanes, to unison ring. Pins disposed in bore of bushing couple arm to unison ring. Forward and aft overhangs may depend from forward and aft sections of polygonal flange and overhang forward and aft annular sides of unison ring.

A dual volute turbocharger for use with an internal combustion engine includes a valve for controlling exhaust gas flow to a turbine housing interior of the dual volute turbocharger. The dual volute turbocharger also includes a first volute and a second volute each adapted for fluid communication with the internal combustion engine. The dual volute turbocharger further includes a wall separating the first and second volutes and a valve seat. The valve seat and the wall collectively define a valve cavity. The valve is movable between a closed position and an open position. The valve and the wall of the turbine housing collectively define a first cross-sectional flow area. The valve and the valve seat collectively define a second cross-sectional flow area. A method of controlling the valve of the dual volute turbocharger is also disclosed.

Disclosed is a probe including a housing; a plurality of axially stacked caps disposed within the housing, the plurality of stacked caps including an inner cap and an outer cap, a probe sensor fixedly disposed between the inner cap and an axial bottom of the housing, and an electronic lead connected to the probe sensor and axially extending through the plurality of caps to an exterior of the probe.

Disclosed is a probe including a housing; a plurality of axially stacked caps disposed within the housing, the plurality of stacked caps including an inner cap and an outer cap, a probe sensor fixedly disposed between the inner cap and an axial bottom of the housing, and an electronic lead connected to the probe sensor and axially extending through the plurality of caps to an exterior of the probe.

A system and a method for detecting a shaft shear event in a gas turbine engine are disclosed. The system comprises a sensor configured to detect a shaft shear force exerted by a shaft on a support structure supporting the shaft where the shaft shear force is indicative of shearing of the shaft. The system also comprises a controller operatively coupled to the sensor and configured to initiate a shutdown of the gas turbine engine in response to the detection of the shaft shear force by the sensor.

A method for selectively modulating bleed air used for cooling a downstream turbine section in a gas turbine engine. The method including: measuring an engine and/or aircraft performance parameter by an engine sensor device; comparing the engine and/or aircraft performance parameter to a performance threshold; determining a bleed trigger condition, if the engine and/or aircraft performance parameter crosses the performance threshold; determining a non-cooling condition, if the engine and/or aircraft performance parameter is below the performance threshold; actuating a flow control valve to an open position, in response to the bleed trigger condition, so that bleed air is extracted from the compressor section and flowed to the downstream turbine section; and terminating, in response to the non-cooling condition, the flow of the bleed air to the downstream turbine section of the engine by actuating the flow control valve to a closed position.

A method for selectively modulating bleed air used for cooling a downstream turbine section in a gas turbine engine. The method including: measuring an engine and/or aircraft performance parameter by an engine sensor device; comparing the engine and/or aircraft performance parameter to a performance threshold; determining a bleed trigger condition, if the engine and/or aircraft performance parameter crosses the performance threshold; determining a non-cooling condition, if the engine and/or aircraft performance parameter is below the performance threshold; actuating a flow control valve to an open position, in response to the bleed trigger condition, so that bleed air is extracted from the compressor section and flowed to the downstream turbine section; and terminating, in response to the non-cooling condition, the flow of the bleed air to the downstream turbine section of the engine by actuating the flow control valve to a closed position.