Abstract: A method of operation is provided during which buffer air flows into a buffer cavity in a turbine engine. The buffer cavity surrounds a bearing compartment in the turbine engine. Leakage air flows across a buffer seal device in the turbine engine. The buffer air and the leakage air are combined to provide combined air. The combined air is vented to a pressure sink such that: a pressure of the buffer air within the buffer cavity is maintained at or below a maximum buffer air threshold during operation of the turbine engine; and a flowrate of the leakage air across the buffer seal device is maintained at or above a minimum leakage air threshold during the operation of the turbine engine.

Abstract: According to an aspect, a method includes predicting, by a processor, a projected oil-wetted metal temperature in a lubrication system of a gas turbine engine at shutdown based on one or more thermal models prior to shutdown of the gas turbine engine. The processor determines a coking index based on the projected oil-wetted metal temperature and a coking limit threshold associated with one or more engine components. An oil coking mitigation action is triggered as a shutdown management event of the gas turbine engine based on the coking index.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a shaft and first and second bearing structures that support the shaft. Each of the first bearing structure and the second bearing structure includes a bearing compartment that contains a lubricant and a seal that contains the lubricant within the bearing compartments. A buffer system is configured to pressurize the seals to prevent the lubricant from escaping the bearing compartments. The buffer system includes a first circuit configured to supply a first buffer supply air to the first bearing structure, a second circuit configured to supply a second buffer supply air to the second bearing structure, and a controller configured to select between at least two bleed air supplies to communicate the first buffer supply air and the second buffer supply air.

Abstract: A system for starting a gas turbine engine of an aircraft is provided. The system includes a pneumatic starter motor, a discrete starter valve switchable between an on-state and an off-state, and a controller operable to perform a starting sequence for the gas turbine engine. The starting sequence includes alternating on and off commands to an electromechanical device coupled to the discrete starter valve to achieve a partially open position of the discrete starter valve to control a flow from a starter air supply to the pneumatic starter motor to drive rotation of a starting spool of the gas turbine engine below an engine idle speed, where the controller modulates a duty cycle of the discrete starter valve via pulse width modulation.

Abstract: A system for controlling a start sequence of a gas turbine engine includes an electronic engine control system, a thermal model, memory, a model for determining a time period (tmotoring), and a controller. The thermal model synthesizes a heat state of the gas turbine engine. The memory records the current heat state at shutdown and a shutdown time of the gas turbine engine. The model for determining the time period is for motoring the gas turbine engine at a predetermined speed Ntarget that is less than a speed to start the gas turbine engine, where tmotoring is a function of the heat state recorded at engine shutdown and an elapsed time of an engine start request relative to a previous shutdown time. The controller modulates a starter valve to maintain the gas turbine engine within a predetermined speed range of NtargetMin to NtargetMax for homogenizing engine temperatures.

Abstract: A gas turbine engine includes a bearing compartment that has at least one seal, a bearing disposed in the bearing compartment, a deoiler fluidly connected with the bearing compartment, and a vacuum generator fluidly connected with the deoiler.

Abstract: A disclosed lubrication system for a gas turbine engine includes a primary passage defining a flow path for lubricant to a gas turbine engine and a bypass passage defining a flow path for lubricant around the gas turbine engine. The lubrication system further includes a primary lubrication pump including a reduced total flow capacity of lubricant with a reduced bypass lubricant flow relative to the total overall flow capacity.

Abstract: An oil supply system for a gas turbine engine has a lubricant pump delivering lubricant to an outlet line. The outlet line is split into at least a hot line and into a cool line, with the hot line directed primarily to locations associated with an engine that are not intended to receive cooler lubricant, and the cool line directed through one or more heat exchangers at which lubricant is cooled. The cool line then is routed to a fan drive gear system of an associated gas turbine engine. A method and apparatus are disclosed. The heat exchangers include at least an air/oil cooler wherein air is pulled across the air/oil cooler to cool oil. The air/oil cooler is provided with an ejector tapping compressed air from a compressor section to increase airflow across the air/oil cooler.

Abstract: A system for starting a gas turbine engine of an aircraft is provided. The system includes a pneumatic starter motor, a discrete starter valve switchable between an on-state and an off-state, and a controller operable to perform a starting sequence for the gas turbine engine. The starting sequence includes alternating on and off commands to an electromechanical device coupled to the discrete starter valve to achieve a partially open position of the discrete starter valve to control a flow from a starter air supply to the pneumatic starter motor to drive rotation of a starting spool of the gas turbine engine below an engine idle speed, where the controller modulates a duty cycle of the discrete starter valve via pulse width modulation.

Abstract: An oil supply system for a gas turbine engine has a lubricant pump delivering lubricant to an outlet line. The outlet line is split into at least a hot line and into a cool line, with the hot line directed primarily to locations associated with an engine that are not intended to receive cooler lubricant, and the cool line directed through one or more heat exchangers at which lubricant is cooled. The cool line then is routed to a fan drive gear system of an associated gas turbine engine. A method and apparatus are disclosed. The heat exchangers include at least an air/oil cooler wherein air is pulled across the air/oil cooler to cool oil. The air/oil cooler is provided with an ejector tapping compressed air from a compressor section to increase airflow across the air/oil cooler.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section, a shaft including a bearing system, a turbine section in communication with the shaft, a speed change mechanism coupling the fan section to the turbine section and a biasing device configured to apply a biasing force against the shaft.

Abstract: A turbine engine includes a turbine section with a low pressure turbine and a turbine case disposed about an axis. A frame assembly defines an outer cavity and an inner cavity with the outer cavity including at least one opening configured and adapted to communicate cooling air to the turbine case. A transfer tube is disposed within the outer cavity and is configured and adapted to receive cooling air. The transfer tube includes a bend configured to impart circumferential velocity to the cooling air within the outer cavity.

Abstract: A system for controlling a start sequence of a gas turbine engine includes an electronic engine control system, a thermal model, memory, a model for determining a time period (fmotoring), and a controller. The thermal model synthesizes a heat state of the gas turbine engine. The memory records the current heat state at shutdown and a shutdown time of the gas turbine engine. The model for determining the time period is for motoring the gas turbine engine at a predetermined speed Ntarget that is less than a speed to start the gas turbine engine, where tmotoring is a function of the heat state recorded at engine shutdown and an elapsed time of an engine start request relative to a previous shutdown time. The controller modulates a starter valve to maintain the gas turbine engine within a predetermined speed range of NtargetMin to NtargetMax for homogenizing engine temperatures.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a compressor section, a combustor in fluid communication with the compressor section and a turbine section in fluid communication with the combustor. A buffer system includes a first circuit that supplies a first buffer supply air and a second circuit that supplies a second buffer supply air. The first circuit includes a first bleed air supply and a second bleed air supply. The second circuit includes a third bleed air supply and a fourth bleed air supply and at least one of the first circuit and the second circuit includes a variable area ejector.

Abstract: A system for starting a gas turbine engine of an aircraft is provided. The system includes a pneumatic starter motor, a discrete starter valve switchable between an on-state and an off-state, and a controller operable to perform a starting sequence for the gas turbine engine. The starting sequence includes alternating on and off commands to an electromechanical device coupled to the discrete starter valve to achieve a partially open position of the discrete starter valve to control a flow from a starter air supply to the pneumatic starter motor to drive rotation of a starting spool of the gas turbine engine below an engine idle speed.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a shaft and first and second bearing structures that support the shaft. Each of the first bearing structure and the second bearing structure includes a bearing compartment that contains a lubricant and a seal that contains the lubricant within the bearing compartments. A buffer system is configured to pressurize the seals to prevent the lubricant from escaping the bearing compartments. The buffer system includes a first circuit configured to supply a first buffer supply air to the first bearing structure, a second circuit configured to supply a second buffer supply air to the second bearing structure, and a controller configured to select between at least two bleed air supplies to communicate the first buffer supply air and the second buffer supply air.

Abstract: According to an aspect, a method includes predicting, by a processor, a projected oil-wetted metal temperature in a lubrication system of a gas turbine engine at shutdown based on one or more thermal models prior to shutdown of the gas turbine engine. The processor determines a coking index based on the projected oil-wetted metal temperature and a coking limit threshold associated with one or more engine components. An oil coking mitigation action is triggered as a shutdown management event of the gas turbine engine based on the coking index.

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a shaft, a first bearing structure and a second bearing structure that support the shaft. Each of the first bearing structure and the second bearing structure includes a bearing compartment that contains a lubricant and a seal that contains the lubricant within the bearing compartments. A buffer system is configured to pressurize the seals to prevent the lubricant from escaping the bearing compartments. The buffer system includes a first circuit configured to supply a first buffer supply air to the first bearing structure, a second circuit configured to supply a second buffer supply air to the second bearing structure, and a controller configured to select between at least two bleed air supplies to communicate the first buffer supply air and the second buffer supply air.

Abstract: A system for starting a gas turbine engine of an aircraft is provided. The system includes a pneumatic starter motor, a discrete starter valve switchable between an on-state and an off-state, and a controller operable to perform a starting sequence for the gas turbine engine. The starting sequence includes alternating on and off commands to an electromechanical device coupled to the discrete starter valve to achieve a partially open position of the discrete starter valve to control a flow from a starter air supply to the pneumatic starter motor to drive rotation of a starting spool of the gas turbine engine below an engine idle speed.

Abstract: A gas turbine engine compressor section has a hub carrying a last row of compressor blades. A compressor exit guide vane is downstream of the last row of compressor blades. A housing is radially inward of the compressor exit guide vane. A non-contact seal is positioned on one of the housing and the hub.