Pratt-whitney Canada Corp

Martin Drolet

Philippe Beauchesne-Martel

Yves Cloutier

Frederic Fortin

Axel Robert

Jason Nichols

Patrick Manoukian

Kevin Ng

Zachary Mounir Faty

Melanie Brillant

Francis Demers

Sylvain Desnoyers

Sebastien Lortie

A system and a method for configuring at least one variable geometry mechanism (VGM) of an aircraft engine are provided. Pass-off testing data for the aircraft engine is obtained, the pass-off testing data indicative of an actual value of at least one operating parameter of the aircraft engine. Based on the pass-off testing data, at least one trim value to be used to adjust a setting of the at least one VGM to bring the actual value of the at least one operating parameter towards a target value is determined, a running line of the aircraft engine being substantially constant when the actual value of the at least one operating parameter is at the target value. The setting of the at least one VGM is adjusted, during pass-off testing of the aircraft engine, using the at least one trim value.

System and method for variable geometry mechanism configuration

A system for a gas turbine engine includes an engine control system. The engine control system includes a processor and a non-transitory memory storing instructions, which instructions when executed by the processor, cause the processor to: obtain a current engine installation configuration for the gas turbine engine and the aircraft, determine an expected normalized value of the engine parameter based on the current engine installation configuration and one or more of a normalized engine power (SHPN) of the gas turbine engine, an engine inlet temperature, an airspeed, or an altitude, determine a fully deteriorated engine (FDE) value of the engine parameter using the expected normalized value of the engine parameter, determine a current value of the engine parameter for the gas turbine engine, and determine the engine operating margin for the engine parameter based on the FDE value of the engine parameter and the current value of the engine parameter.

Systems and methods for determining gas turbine engine operating margins

A gas turbine engine includes a compressor, a combustor and a turbine. A compressor section has an exit upstream of the combustor. A temperature probe extends through a compressor exit housing surrounding the compressor section exit. The temperature probe taps air downstream of the exit into a probe body having a passage leading to a temperature sensor. There is a controlled leakage also allowing air from within the compressor exit housing to leak. A method and a test probe are also disclosed.

Method and apparatus to measure gas turbine compressor exit temperature

Methods and systems for operating an aircraft engine having a compressor are described. The method comprises determining, based on actual operating parameters of the aircraft engine, a compressor mass flow limit for an aerodynamic stability of the aircraft engine; determining an actual compressor mass flow of the compressor of the aircraft engine, wherein the actual compressor mass flow is based on measured values of the aircraft engine; comparing the actual compressor mass flow to the compressor mass flow limit; and governing operation of the aircraft engine to cause an alternative compressor mass flow when the actual compressor mass flow reaches or is anticipated to reach the compressor mass flow limit.

Method and system for operating an engine to prevent high power engine surges

There are described methods and systems for operating an aircraft having two or more engines. One method comprises operating the two or more engines of the aircraft in an asymmetric operating regime, wherein a first of the engines is in an active mode to provide motive power to the aircraft and a second of the engines is in a standby mode to provide substantially no motive power to the aircraft; governing the first engine in the active mode using a first governing logic; and governing the second engine in the standby mode using a second governing logic, the second governing logic based on a target compressor speed and variable geometry mechanism (VGM) settings that are adjusted using trim values dependent on at least one parameter of the second engine in the standby mode.

Method and system for governing an engine at low power

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.

Method of controlling the geometrical configuration of a variable geometry element in a gas turbine engine compressor stage

In a method for operating an engine, a request for an increase in thrust generated by the engine is received. In response to receipt of the request, a determination is made as to whether at least one operating condition for heat application-based thrust is met. If so, a heat source is applied to heat bypass air flowing through the bypass duct towards the exhaust nozzle and the increase in thrust is generated from an increased temperature of mixed bypass air and core air at the exhaust nozzle.

Methods and systems for operating an engine to generate additional thrust

Systems and methods for operating a rotorcraft engine are described herein. Measurements indicative of at least one of current temperature and current pressure at an inlet of the engine are obtained from at least one sensor while the rotorcraft is in flight. At least one current inlet loss is determined from the measurements. Current available engine power of the rotorcraft engine is determined based on the at least one current inlet losses. A visual indication of the current available engine power is produced via a flight display.

Method and system for operating a rotorcraft engine

There is disclosed a centrifugal compressor including an impeller rotatable about an axis and a diffuser downstream of the impeller. The diffuser has walls delimiting flow passages. Plasma actuators are positioned adjacent the walls and are operatively connectable to a source of electricity. The plasma actuators have a first electrode, a second electrode, and a dielectric layer therebetween. The first electrode is upstream of the second electrode. The first electrode is exposed to the flow passage. The second electrode is shielded from the flow passage by the dielectric layer. The plasma actuators are operable to generate an electric field through the dielectric layer. The plasma actuators are located closer to inlets of the flow passage than to outlets of the flow passages. A method of operating the compressor is disclosed.

Compressor diffuser with plasma actuators

A gas turbine engine includes a compressor, an annular casing surrounding the compressor, and a bleed flow adapter mounted to an exterior side of the annular casing. The annular casing includes the exterior side and an interior side. The interior side surrounds a compressor bleed cavity located downstream of at least a portion of the compressor. The bleed flow adapter is in fluid communication with the bleed cavity. The bleed flow adapter includes an inlet end, an outlet end, and an inner diameter surface extending between the inlet end and the outlet end. The inner diameter surface defines a bleed passage. The bleed flow adapter further includes a fluid port formed through the inner diameter surface. The gas turbine engine further includes a bleed flow measurement system including a first pressure sensor in fluid communication with the bleed passage of the bleed flow adapter via the fluid port.

Method and apparatus for measuring compressor bleed flow

Systems and Methods are described for testing engine performance in-flight in an aircraft having a first engine and a second engine. The method comprises operating the first engine at a first power level in an output speed governing mode, operating the second engine at a second power level greater than the first power level in a core speed governing mode concurrently with the first engine operating at the first power level in the output speed governing mode, and performing an engine performance test on the second engine while the second engine is at the second power level in the core speed governing mode.

System and method for testing engine performance in-flight

A method and system for monitoring a temperature of a gas turbine engine are described. The method comprises obtaining individual sensor readings from functioning temperature sensors of a sensor array, where a number of the functioning temperature sensors is less than a total number of temperature sensors in the sensor array; applying correction factors to the individual sensor readings of the functioning temperature sensors based on deviations of the individual sensor readings from a total average temperature of all the temperature sensors in the array, to obtain corrected individual sensor readings; and determining a corrected total average temperature as a sum of the corrected individual sensor readings divided by the number of functioning temperature sensors.

System and method for monitoring temperature of a gas turbine engine

A system and method for detecting inlet temperature distortion of an engine are described. The method comprises obtaining an outside air temperature from at least one first sensor, obtaining an inlet temperature of the engine from at least one second sensor, determining an inlet temperature distortion based on a difference between the outside air temperature and the inlet temperature, comparing the inlet temperature distortion to a threshold, and issuing an alert when the inlet temperature distortion exceeds the threshold.

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Saint-Amable, Canada

Martin Drolet