Abstract: A system for providing auxiliary power in an aircraft. A propulsion core comprises a compressor, a combustor, a turbine, and a shaft. An accessory unit comprises an accessory combustor, an accessory turbine, and an accessory shaft. A tank is configured to hold high pressure air and operably connected to the accessory unit. An electric generator comprises an electrical output and a mechanical input, with the mechanical input operably connected to the accessory shaft and the electrical output operably connected to an electric motor operably connected to the shaft. The electrical output is operably connected to an auxiliary power consuming device in the aircraft.

Abstract: A system for providing auxiliary power in an aircraft. A propulsion core comprises a compressor, a combustor, a turbine, and a shaft. An accessory unit comprises an accessory combustor, an accessory turbine, and an accessory shaft. A tank is configured to hold high pressure air and operably connected to the accessory unit. An electric generator comprises an electrical output and a mechanical input, with the mechanical input operably connected to the accessory shaft and the electrical output operably connected to an electric motor operably connected to the shaft. The electrical output is operably connected to an auxiliary power consuming device in the aircraft.

Abstract: A system for providing auxiliary power in an aircraft. A propulsion core comprises a compressor, a combustor, a turbine, and a shaft. An accessory unit comprises an accessory combustor, an accessory turbine, and an accessory shaft. A tank is configured to hold high pressure air and operably connected to the accessory unit. An electric generator comprises an electrical output and a mechanical input, with the mechanical input operably connected to the accessory shaft and the electrical output operably connected to an electric motor operably connected to the shaft. The electrical output is operably connected to an auxiliary power consuming device in the aircraft.

Abstract: A system for providing auxiliary power in an aircraft. A propulsion core comprises a compressor, a combustor, a turbine, and a shaft. An accessory unit comprises an accessory combustor, an accessory turbine, and an accessory shaft. A tank is configured to hold high pressure air and operably connected to the accessory unit. An electric generator comprises an electrical output and a mechanical input, with the mechanical input operably connected to the accessory shaft and the electrical output operably connected to an electric motor operably connected to the shaft. The electrical output is operably connected to an auxiliary power consuming device in the aircraft.

Abstract: A gas turbine engine system is disclosed which includes a core passage and a bypass passage which can be configured as a fan bypass duct or a third stream bypass duct. The core passage and bypass passage are routed to flow through a nozzle before exiting overboard an aircraft. The nozzle includes moveable members capable of changing a configuration of the nozzle. In one form the moveable members are capable of changing throat area for portions of the nozzle that receive working fluid from the core passage and the bypass passage. The bypass passage can include a branch. In one form the branch can include a heat exchanger. The bypass passage can also provide cooling to one or more portions of the nozzle, such as cooling to a deck of the nozzle.

Abstract: A gas turbine engine is disclosed which includes a bypass passage that in some embodiments are capable of being configured to act as a resonance space. The resonance space can be used to attenuate/accentuate/etc a noise produced elsewhere. The bypass passage can be configured in a number of ways to form the resonance space. For example, the space can have any variety of geometries, configurations, etc. In one non-limiting form the resonance space can attenuate a noise forward of the bypass duct. In another non-limiting form the resonance space can attenuate a noise aft of the bypass duct. Any number of variations is possible.

Abstract: A gas turbine engine is disclosed having a thermoelectric device capable of changing a tip clearance in a turbomachinery component. In one non-limiting form the turbomachinery component is a compressor. The thermoelectric device can be used in some forms to harvest power derived from a waste heat. The tip clearance control system can include a sensor used to determine a clearance between a tip and a wall of the turbomachinery component.

Abstract: One embodiment of the present invention is a unique aircraft propulsion gas turbine engine. Another embodiment is a unique gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines with heat exchange systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: An apparatus is disclosed that includes a gas turbine engine including a first rotor blade axially adjacent a second rotor blade and an aperture formed in one of the first rotor blade and the second rotor blade and structured to emit a fluid therefrom. A fluid source is in flow communication with the aperture and configured to flow the fluid through the aperture.

Abstract: A gas turbine engine system is disclosed which includes a core passage and a bypass passage which can be configured as a fan bypass duct or a third stream bypass duct. The core passage and bypass passage are routed to flow through a nozzle before exiting overboard an aircraft. The nozzle includes moveable members capable of changing a configuration of the nozzle. In one form the moveable members are capable of changing throat areas for portions of the nozzle that receive working fluid from the core passage and the bypass passage. The bypass passage can include a branch. In one form the branch can include a heat exchanger. The bypass passage can also provide cooling to one or more portions of the nozzle, such as cooling to a deck of the nozzle.

Abstract: An apparatus is disclosed that includes a gas turbine engine including a first rotor blade axially adjacent a second rotor blade and an aperture formed in one of the first rotor blade and the second rotor blade and structured to emit a fluid therefrom. A fluid source is in flow communication with the aperture and configured to flow the fluid through the aperture.

Abstract: One embodiment of the present invention is a unique aircraft propulsion gas turbine engine. Another embodiment is a unique gas turbine engine. Another embodiment is a unique gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines with heat exchange systems. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: A gas turbine engine is disclosed having a thermoelectric device capable of changing a tip clearance in a turbomachinery component. In one non-limiting form the turbomachinery component is a compressor. The thermoelectric device can be used in some forms to harvest power derived from a waste heat. The tip clearance control system can include a sensor used to determine a clearance between a tip and a wall of the turbomachinery component.

Abstract: A gas turbine engine is disclosed which includes a bypass passage that in some embodiments are capable of being configured to act as a resonance space. The resonance space can be used to attenuate/accentuate/etc a noise produced elsewhere. The bypass passage can be configured in a number of ways to form the resonance space. For example, the space can have any variety of geometries, configurations, etc. In one non-limiting form the resonance space can attenuate a noise forward of the bypass duct. In another non-limiting form the resonance space can attenuate a noise aft of the bypass duct. Any number of variations is possible.

Abstract: A gas turbine engine system is disclosed which includes a core passage and a bypass passage which can be configured as a fan bypass duct or a third stream bypass duct. The core passage and bypass passage are routed to flow through a nozzle before exiting overboard an aircraft. The nozzle includes moveable members capable of changing a configuration of the nozzle. In one form the moveable members are capable of changing throat areas for portions of the nozzle that receive working fluid from the core passage and the bypass passage. The bypass passage can include a branch. In one form the branch can include a heat exchanger. The bypass passage can also provide cooling to one or more portions of the nozzle, such as cooling to a deck of the nozzle.

Abstract: A gas turbine engine system is disclosed which includes a core passage and a bypass passage which can be configured as a fan bypass duct or a third stream bypass duct. The core passage and bypass passage are routed to flow through a nozzle before exiting overboard an aircraft. The nozzle includes moveable members capable of changing a configuration of the nozzle. In one form the moveable members are capable of changing throat area for portions of the nozzle that receive working fluid from the core passage and the bypass passage. The bypass passage can include a branch. In one form the branch can include a heat exchanger. The bypass passage can also provide cooling to one or more portions of the nozzle, such as cooling to a deck of the nozzle.