Abstract: A method for operating a gas turbine engine having a starter-electric generator driven by one of a plurality of shafts of the gas turbine engine is provided. The method includes determining a desired amount of thrust to be produced by the gas turbine engine, as well as a desired amount of electrical power to be generated by the starter-electric generator of the gas turbine engine. The method operates the gas turbine engine to produce the desired amount of thrust, while producing less than the desired amount of electrical power using the starter-electric generator. Producing less than the desired amount of electrical power using the starter-electric generator allows for the desired amount of thrust production, or allows for the desired amount of thrust production more quickly.

Abstract: Methods and devices for cooling systems (700) are provided that are in fluid communication with bleed air from a jet engine compressor. The cooling system can include: a first precooler (210) receiving bleed air from the jet engine compressor; a heat exchanger (730) downstream from the first precooler (210); a cooling system compressor (220) downstream from the first precooler (210), wherein the heat exchanger (730) and the cooling system compressor (220) are in separate flow paths from the first precooler (210); a cooling system precooler (230) downstream from the cooling system compressor (220); a VGT cooling system turbine (240) downstream from the cooling system precooler (230); and a discharge conduit (245) downstream from the cooling system turbine (240) and the heat exchanger (730). A bypass line (290) for bypassing the turbine can also be included.

Abstract: Methods and devices for cooling systems (100, 700) are provided that are in fluid communication with bleed air from a jet engine compressor. The cooling systems include: a first precooler (210) receiving bleed air from the jet engine compressor; a heat exchanger (730) downstream from the first precooler (210); a cooling system compressor (220) downstream from the first precooler (210), wherein the heat exchanger (730) and the cooling system compressor (220) are in separate flow paths from the first precooler (210); a cooling system precooler (230) downstream from the cooling system compressor (220); a cooling system turbine (240) with variable guide vanes—VGT—and downstream from the cooling system precooler (230); and a discharge conduit (245) downstream from the cooling system turbine (240) and the heat exchanger (730). A bypass line (290) can also be included that bypasses the cooling system turbine (240).

Abstract: Airplanes and jet engines are provided that includes an engine compressor; a combustor in flow communication with the engine compressor; an engine turbine in flow communication with the combustor to receive combustion products from the combustor; and a bleed air cooling system in fluid communication with bleed air from the engine compressor. The bleed air cooling system can include a first precooler in fluid communication with the bleed air from the engine compressor; a cooling system turbine in fluid communication with and downstream from the first precooler; and a discharge conduit from the cooling system turbine that is configured to be in fluid communication with at least one of an aircraft thermal management system and an aircraft environmental control system. Methods are also described for providing cooling fluid from a jet engine.

Abstract: Methods and devices for cooling systems (700) are provided that are in fluid communication with bleed air from a jet engine compressor. The cooling system can include: a first precooler (210) receiving bleed air from the jet engine compressor; a heat exchanger (730) downstream from the first precooler (210); a cooling system compressor (220) downstream from the first precooler (210), wherein the heat exchanger (730) and the cooling system compressor (220) are in separate flow paths from the first precooler (210); a cooling system precooler (230) downstream from the cooling system compressor (220); a VGT cooling system turbine (240) downstream from the cooling system precooler (230); and a discharge conduit (245) downstream from the cooling system turbine (240) and the heat exchanger (730). A bypass line (290) for bypassing the turbine can also be included.

Abstract: Methods and devices for cooling systems (100, 700) are provided that are in fluid communication with bleed air from a jet engine compressor. The cooling systems include: a first precooler (210) receiving bleed air from the jet engine compressor; a heat exchanger (730) downstream from the first precooler (210); a cooling system compressor (220) downstream from the first precooler (210), wherein the heat exchanger (730) and the cooling system compressor (220) are in separate flow paths from the first precooler (210); a cooling system precooler (230) downstream from the cooling system compressor (220); a cooling system turbine (240) with variable guide vanes—VGT—and downstream from the cooling system precooler (230); and a discharge conduit (245) downstream from the cooling system turbine (240) and the heat exchanger (730). A bypass line (290) can also be included that bypasses the cooling system turbine (240).

Abstract: A method for operating a gas turbine engine having a starter-electric generator driven by one of a plurality of shafts of the gas turbine engine is provided. The method includes determining a desired amount of thrust to be produced by the gas turbine engine, as well as a desired amount of electrical power to be drawn from the starter-electric generator of the gas turbine engine. The method operates the gas turbine engine to produce the desired amount of thrust, while drawing less than the desired amount of electrical power from the starter-electric generator. Drawing less than the desired amount of electrical power from the starter-electric generator allows for the desired amount of thrust production, allows for the desired amount of thrust production more quickly, or allows for maintenance of a stall margin for any purpose (such as to increase an efficiency of the engine or to allow for certain engine designs).

Abstract: Airplanes and jet engines are provided that includes an engine compressor; a combustor in flow communication with the engine compressor; an engine turbine in flow communication with the combustor to receive combustion products from the combustor; and a bleed air cooling system in fluid communication with bleed air from the engine compressor. The bleed air cooling system can include a first precooler in fluid communication with the bleed air from the engine compressor; a cooling system turbine in fluid communication with and downstream from the first precooler; and a discharge conduit from the cooling system turbine that is configured to be in fluid communication with at least one of an aircraft thermal management system and an aircraft environmental control system. Methods are also described for providing cooling fluid from a jet engine.

Abstract: A method for assembling a turbine engine includes assembling a heat exchanger assembly that includes at least a radially inner plate, a radially outer plate, and a heat exchanger coupled between the radially inner and outer plates, forming the heat exchanger assembly such that the heat exchanger assembly has a substantially arcuate shape, and coupling the heat exchanger to a fan casing such that the heat exchanger is positioned upstream or downstream from the fan assembly.

Abstract: A bleed system for a gas turbine engine includes: (a) a bleed air turbine having a turbine inlet adapted to be coupled to a source of compressor bleed air at a first pressure; (b) a bleed air compressor mechanically coupled to the bleed air turbine, and having a compressor inlet adapted to be coupled to a source of fan discharge air at a second pressure substantially lower than the first pressure; and (c) a mixing duct coupled to a turbine exit of the bleed air turbine and to a compressor exit of the bleed air compressor.

Abstract: A bleed system for a gas turbine engine includes: (a) a bleed air turbine having a turbine inlet adapted to be coupled to a source of compressor bleed air at a first pressure; (b) a bleed air compressor mechanically coupled to the bleed air turbine, and having a compressor inlet adapted to be coupled to a source of fan discharge air at a second pressure substantially lower than the first pressure; and (c) a mixing duct coupled to a turbine exit of the bleed air turbine and to a compressor exit of the bleed air compressor.

Abstract: A method for assembling a turbine engine includes assembling a heat exchanger assembly that includes at least a radially inner plate, a radially outer plate, and a heat exchanger coupled between the radially inner and outer plates, forming the heat exchanger assembly such that the heat exchanger assembly has a substantially arcuate shape, and coupling the heat exchanger to a fan casing such that the heat exchanger is positioned upstream or downstream from the fan assembly.

Abstract: A coating system and method for reducing the tendency for hydrocarbon fluids, such as fuels and oils, to form carbonaceous deposits that adhere to a wall of a containment article. Of particular concern are carbonaceous deposits that form at temperatures below about 650° F. (about 345° C.). The coating system combines an outermost layer of platinum with a ceramic barrier layer. The coating system significantly reduces the formation of carbonaceous deposits and the adhesion of such deposits. To further reduce wall and hydrocarbon fluid temperatures and formation of carbonaceous deposits, the coating system is preferably applied to the surface of the wall wetted by the fluid, as well as the opposite surface of the wall exposed to a surrounding environment. The outermost layers serve as radiation shields to reduce heat transfer from the surrounding environment to the wall, and from the wall to the hydrocarbon fluid.

Abstract: A coating system and method for reducing the tendency for hydrocarbon fluids, such as fuels and oils, to form carbonaceous deposits that adhere to the walls of a containment article. Of particular concern are carbonaceous deposits that form at temperatures below about 650° F. (about 345° C.). The coating system combines an outermost layer of platinum with a ceramic barrier layer. The coating system has been shown to significantly reduce the formation of carbonaceous deposits at temperatures between about 220° F. and 650° F. (about 105° C. to about 345° C.), as well as reduce the adhesion of such deposits. The platinum outermost layer also serves as a radiation shield to reduce heat transfer from the containment article to the hydrocarbon fluid. The outermost layer is preferably deposited as an extremely thin film by chemical vapor deposition. The barrier layer is deposited to a thickness sufficient to prevent interdiffusion of the platinum outermost layer with the containment wall.

Abstract: A method and system for monitoring the condition of a bearing mounted on a rotating shaft, particularly one present in an aircraft gas turbine engine, where the vibration sensor that monitors the bearing is remote therefrom but proximate to the rotating shaft. The vibration sensor obtains a broadband signal having frequencies that include the bearing defect peak of the monitored bearing. The broadband signal is analyzed to identify the presence of the bearing defect peak. If the bearing defect peak is present, the amplitude of this peak is quantified to determine whether degradation of the monitored bearing has at least reached a threshold criteria previously established.

Abstract: A coating system and method for reducing the tendency for hydrocarbon fluids, such as fuels and oils, to form carbonaceous deposits that adhere to the walls of a containment article. Of particular concern are carbonaceous deposits that form at temperatures below about 650° F. (about 345° C.). The coating system combines an outermost layer of platinum with a ceramic barrier layer. The coating system has been shown to significantly reduce the formation of carbonaceous deposits at temperatures between about 220° F. and 650° F. (about 105° C. to about 345° C.), as well as reduce the adhesion of such deposits. The platinum outermost layer also serves as a radiation shield to reduce heat transfer from the containment article to the hydrocarbon fluid. The outermost layer is preferably deposited as an extremely thin film by chemical vapor deposition. The barrier layer is deposited to a thickness sufficient to prevent interdiffusion of the platinum outermost layer with the containment wall.

Abstract: A method and system for monitoring the condition of a bearing mounted on a rotating shaft, particularly one present in an aircraft gas turbine engine, where the vibration sensor that monitors the bearing is remote therefrom but proximate to the rotating shaft. The vibration sensor obtains a broadband signal having frequencies that include the bearing defect peak of the monitored bearing. The broadband signal is analyzed to identify the presence of the bearing defect peak. If the bearing defect peak is present, the amplitude of this peak is quantified to determine whether degradation of the monitored bearing has at least reached a threshold criteria previously established.

Abstract: A method and system for monitoring engine performance in a gas turbine engine uses a plurality of sensors to sense data related to the operation and performance of the gas turbine engine. Selected data parameters from the sensed data are continuously sampled prior to completion of an engine start sequence. The selected data parameters are then evaluated to determine whether specific criteria have been met, and evaluation data are captured whenever the criteria are met. The evaluation data are used to normalize engine performance data to a particular thermal state of the engine. The normalized engine performance data are then trended by comparing it to a parametric baseline for the gas turbine engine.

Abstract: A method and system for monitoring engine performance in a gas turbine engine uses a plurality of sensors to sense data related to the operation and performance of the gas turbine engine. Selected data parameters from the sensed data are continuously sampled prior to completion of an engine start sequence. The selected data parameters are then evaluated to determine whether specific criteria have been met, and evaluation data are captured whenever the criteria are met. The evaluation data are used to normalize engine performance data to a particular thermal state of the engine. The normalized engine performance data are then trended by comparing it to a parametric baseline for the gas turbine engine.

Abstract: A method and system for monitoring engine starting system performance in a gas turbine engine using a plurality of sensors to sense data related to the operation and performance of the gas turbine engine. Selected data parameters from the sensed data are continuously sampled during an engine start sequence. The data parameters are then evaluated to determine whether predetermined sets of criteria have been met. Selected engine starting system performance data and data sampling times are captured when each of the sets of criteria are met. The captured engine starting system performance data are normalized and then compared to a parametric baseline for the gas turbine engine.