Abstract: A gas turbine engine includes a plurality of rotating components housed within a main compressor section and a turbine section. A first tap is connected to the main compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A cooling air valve is configured to selectively block flow of cooling air across the heat exchanger. A cooling compressor is connected downstream of the heat exchanger. A shut off valve stops flow between the heat exchanger and the cooling compressor. A second tap is configured to deliver air at a second pressure which is higher than the first pressure. A mixing chamber is connected downstream of the cooling compressor and the second tap. The mixing chamber is configured to deliver air to at least one of the plurality of rotating components. A system stops flow between the cooling compressor and the plurality of rotating components.

Abstract: A heat exchanger array includes a first row of heat exchangers, a second row of heat exchangers, and side curtains. The first row heat exchangers are spaced apart to define first gaps. The second row heat exchangers are spaced apart to define second gaps and are positioned downstream of and staggered from the first row heat exchangers such that the second row heat exchangers are aligned with the first gaps and the first row heat exchangers are aligned with the second gaps. Each side curtain is in close proximity to a first row heat exchanger and a second row heat exchanger. The side curtains define a neck region upstream of and aligned with each first row heat exchanger and each second row heat exchanger. Each neck region has a neck area that is less than a frontal area of the heat exchanger with which it is aligned.

Abstract: A gas turbine engine comprises a fan rotor having fan blades received within an outer nacelle, and the outer nacelle having an inner surface. At least a portion of the nacelle inner surface extends radially inwardly to be radially inward of an outer diameter of the fan blades. The inner surface of the nacelle is formed with a trench, which extends into the inner surface to a radially outer extent that is spaced radially outward of the outer diameter of the fan blades.

Abstract: A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap and configured to deliver air to an aircraft fuselage. A cooling compressor is connected downstream of the heat exchanger. A high pressure feed is configured to deliver air at a second pressure which is higher than the first pressure. The cooling compressor is configured to deliver air to at least one of the plurality of rotating components. A valve assembly that can select whether air from the first tap or air from the high pressure feed is delivered to the aircraft pneumatic system.

Abstract: A system is provided for alternating starter use during multi-engine motoring in an aircraft. The system includes a first engine starting system of a first engine. A first controller is in communication with a second controller that controls a second engine starting system of a second engine, the first controller being configured to intermittently direct power to the first engine starting system to alternately accelerate and decelerate the first engine during motoring with respect to the second engine.

Abstract: A gas turbine engine includes a plurality of rotatable components housed within a main compressor section and a turbine section. A cooling system is connected to tap air from said main compressor section. A first tap is connected to a first heat exchanger. The first heat exchanger is connected to a cooling compressor for raising a pressure of the tapped air downstream of the first heat exchanger. A second heat exchanger is downstream of the cooling compressor, and a connection is downstream of the second heat exchanger for delivering air to a bearing compartment. A connection intermediate the cooling compressor and the second heat exchanger delivers cooling air to at least one of the rotatable components.

Abstract: A method of monitoring a gas turbine engine includes the steps of: (a) receiving information from actual flights of an aircraft including an engine to be monitored, and including at least one of the ambient temperature at takeoff, and internal engine pressures, temperatures and speeds; (b) evaluating the damage accumulated on an engine component given the data received in step (a); (c) storing the determined damage from step (b); (d) repeating steps (a)-(c); (e) recommending a suggested future use for the component based upon steps (a)-(d). A system is also disclosed.

Abstract: A recuperated gas turbine engine includes an engine core that has a compressor section, a combustor section, and a turbine section. An exhaust duct is located downstream of the turbine section for receiving a hot turbine exhaust stream from the turbine section. The exhaust duct includes a heat exchanger and a temperature-control module upstream of the heat exchanger. A first compressor bleed line portion leads into the heat exchanger, and a second compressor bleed lie portion leads into the exhaust duct upstream of the heat exchanger. A compressor return line leads from the heat exchanger into the engine core upstream of the combustor section. The compressor bleed line is operable to selectively feed compressed air to the heat exchanger, and the temperature-control module is operable to selectively modulate at least one of temperature and flow of the hot turbine exhaust stream with respect to the heat exchanger.

Abstract: A gas turbine engine having an engine axis and method of manufacturing the same is disclosed. The gas turbine engine may comprise a fan configured to drive air, a low pressure compressor section having a core flow path and configured to draw in and compress air flowing along the core flow path, a spool configured to drive the fan, and geared architecture configured to adjust the fan speed. The gas turbine engine may also include a housing defining a compartment that encloses the geared architecture. The housing is disposed between the core flow path and the axis, and includes a shielded mid-section that is in thermal communication with the core flow path of the low pressure compressor section. The shielded mid-section includes an outer layer and an insulator adjacent to the outer layer.

Abstract: Assemblies are provided for rotational equipment. One of these assemblies includes a bladed rotor assembly, a stator vane assembly, a fixed stator structure and a seal assembly. The bladed rotor assembly includes a rotor disk structure. The stator vane assembly is disposed adjacent the bladed rotor assembly. The fixed stator structure is connected to and radially within the stator vane assembly. The seal assembly is configured for sealing a gap between the stator structure and the rotor disk structure, wherein the seal assembly includes a non-contact seal.

Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

Abstract: A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A flowpath is defined between a rotating surface and a non-rotating surface. The flowpath is connected downstream of the heat exchanger and is configured to deliver air to at least one of the plurality of rotating components. At least a portion of the non-rotating surface and the rotating surface includes a base metal. An insulation material is disposed on a surface along the flowpath.

Abstract: A system is provided for alternating starter use during multi-engine motoring in an aircraft. The system includes a first engine starting system of a first engine and a controller. The controller is operable to coordinate control of the first engine starting system to alternate use of power from a power source relative to use of the power by one or more other engine starting systems of the aircraft while maintaining a starting spool speed of the first engine below a resonance speed of the starting spool during motoring of the first engine.

Abstract: A gas turbine engine assembly includes a fan section delivering air into a main compressor section. The main compressor section compresses air and delivers air into a combustion section. Products of combustion pass from the combustion section over a turbine section to drive the fan section and main compressor sections. A gearbox is driven by the turbine section to drive the fan section. A pylon supports the gas turbine engine. An environmental control system includes a higher pressure tap at a higher pressure location in the main compressor section, and a lower pressure tap at a lower pressure location. The lower pressure location being at a lower pressure than the higher pressure location. The lower pressure tap communicates to a first passage leading to a downstream outlet and a compressor section of a turbocompressor.

Abstract: A gas turbine engine installed on an aircraft includes a fan rotor, a turbine rotor, a gearbox, an auxiliary pump, and an electric motor. The gearbox couples the fan rotor to the turbine rotor, the turbine rotor being adapted to drive the fan rotor via the gearbox. The auxiliary pump is configured to circulate lubricating fluid in an auxiliary lubrication system and supplies the gearbox. The electric motor is configured to receive electricity when the aircraft is parked an adapted to drive the auxiliary pump such that the auxiliary pump circulates lubricating fluid while the aircraft is parked.

Abstract: A gas turbine engine for an aircraft includes a fan section, a turbine section, a compressor section, and an engine bleed system. The compressor section includes a low compressor stage proximate to the fan section, a high compressor stage axially downstream from the low compressor stage and proximate to the turbine section, and a mid-compressor stage including variable vane assemblies distributed axially between the low and high compressor stage. The engine bleed system includes engine bleed taps with a mid-compressor bleed tap axially between two of the variable vane assemblies, at least one low stage bleed tap axially upstream from the mid-compressor bleed tap, and at least one high stage bleed tap axially downstream from the mid-compressor bleed tap. An external manifold is in pneumatic communication with the mid-compressor bleed tap. A valve system can select one engine bleed tap as a bleed air source for an aircraft use.

Abstract: A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including a low pressure compressor and a second compressor section, and a turbine section including a fan drive turbine and a high pressure turbine. The fan drive turbine drives the low pressure compressor and a gear arrangement to drive the fan section. A core split power ratio is provided by power input to the high pressure compressor divided by a power input to the low pressure compressor measured in horsepower.

Abstract: A turbine engine system includes a first lubricant circuit, a second lubricant circuit, a plurality of engine stages and a shaft. The first lubricant circuit includes a first turbine engine component that is fluidly coupled with a first lubricant heat exchanger. The first turbine engine component includes a gear train, which connects a first of the engine stages to a second of the engine stages. The second lubricant circuit includes a second turbine engine component that is fluidly coupled with a second lubricant heat exchanger. The second lubricant circuit is fluidly coupled with the first lubricant circuit, and the second turbine engine component includes a bearing. The shaft is supported by the bearing, and connected to one of the engine stages.

Abstract: A ratio of an outer diameter of a fan hub at a leading edge of the blades to an outer tip diameter of the blades at the leading edge is greater than or equal to about 0.24 and less than or equal to about 0.38. The fan tip diameter is greater than or equal to about 84 inches (213.36 centimeters) and a fan tip speed is less than or equal to about 1050 ft/second (320.04 meters/second). The fan drive turbine has between three and six stages.

Abstract: A bowed rotor prevention system for a gas turbine engine includes a core turning motor operable to drive rotation of an engine core of the gas turbine engine. The bowed rotor prevention system also includes an auxiliary electric motor control operable to control the core turning motor to drive rotation of the engine core using electric power while a full authority digital engine control that controls operation of the gas turbine engine is either depowered or in a power state that is less than a power level used by the full authority digital engine control in flight operation.