RTX Corporation Patente

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An aircraft system (10) is provided that includes a first turbine engine (18), a second turbine engine (20) and an electrical system (14). The electrical system (14) includes a clutch (152), a first electric machine (118) and a second electric machine (120) electrically coupled to the first electric machine (118). The first electric machine (118) is mechanically coupled to a first rotating assembly (52) of the first turbine engine (18). The second electric machine (120) is mechanically coupled to a second rotating assembly (94) of the second turbine engine (18) through the clutch (152).

A gas turbine engine includes a fan received within a fan case (102), the fan delivers bypass air into a bypass duct between an inner radial portion of the fan case (102) and an outer radial surface of an inner core housing (104). The fan also delivers air into the inner core housing (104). A compressor section, a combustor and a turbine are within the inner core housing (104). A bifurcation (106) extends radially to connect the inner core housing (104) to the fan case (102). The bifurcation (106) has a forward opening defining an air scoop (110). The air scoop (110) separates air into a first airflow chamber connected downstream to a precooler (112). The precooler (112) also has a connection to receive compressed air (113) from the compressor section. The air in the bypass duct is delivered to the precooler (112) for cooling the compressed air (113) in the precooler (112). The air scoop (110) also defines a second airflow chamber for delivering cooling air to components within the inner core housing (104). An aircraft supply system is also disclosed.

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A core is provided for fabricating a blade to include an airfoil (220). The airfoil (220) includes pressure and suction surfaces (221, 222), leading and trailing edges (223, 224) extending along the pressure and suction surfaces (221, 222) and a tip shelf (225) with a first sweep configuration (701) and a wall (226, 310). The core includes channel sections configured to form internal channels (702) within the airfoil (220) by casting processes and tip rods extending from respective portions of the channel sections proximate to a tip shelf (225) location. The respective portions of the channel sections have a second sweep configuration (430) corresponding to the first sweep configuration (701). The tip rods are configured to extend through the wall (226, 310) at an angle of about 5-12 degrees inclusive relative to a normal angle of the wall (226, 310) during the casting processes to form through-holes angled at about 5-12 degrees inclusive in the wall (226, 310).

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A wear-reducing coating system (300) for a substrate (100) includes an innermost layer (304) disposed on a substrate (100) and an outermost layer (302) disposed on the innermost layer (304). The innermost layer (304) has higher load-carrying capacity than the outermost layer (302) and the outermost layer (302) has a lower coefficient of friction than the innermost layer (304). A method of applying a wear-reducing coating (300) to a substrate (100) and a method of applying a coating (300) to a split ring casing (100) for a high pressure compressor are also disclosed.

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A hybrid electric propulsion system (100) including: a gas turbine engine (20; 120) comprising a low speed spool (30) and a high speed spool (32), the low speed spool comprising a low pressure compressor (44) and a low pressure turbine (46), and the high speed spool comprising a high pressure compressor (52) and a high pressure turbine (54); an electric motor (212A, 212B; 282) configured to augment rotational power of the high speed spool (32) or the low speed spool (30); and a controller (256) to: detect an in-flight windmill re-start condition of the gas turbine engine (20; 120); and cause power to be supplied from a power source to the electric motor (212A, 212B; 282) in order to augment rotational power of the high speed or the low speed spool during the detected in-flight windmill re-start condition.

Embodiments of the present disclosure generally relate to aircraft engines (20) and, more particularly to data acquisition for aircraft engines (20) using neuromorphic sensors (206). In some embodiments, an event associated with the aircraft engine (20) may be identified based on a change in a data characteristic measured from a neuromorphic sensor (206) and, in response to identifying the event associated with the aircraft engine (20), at least one other sensor (206) coupled to the aircraft engine (20) may be activated. Other embodiments may be disclosed or claimed.

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Airfoils (60, 62; 100; 400; 500; 600; 700; 800; 900; 1000; 1100) for gas turbine engines have internal walls and internal cavities. A leading edge cavity (510) is defined within the airfoil body (500) along a leading edge (502) and a leading edge feed cavity (512) is arranged aft of the leading edge cavity (510). A bent leading edge rib (528) is arranged between the leading edge cavity (510) and the leading edge feed cavity (512). A main body cavity (522a) is arranged aft of the leading edge feed cavity (512) and defined at least in part by two interior ribs (534) that define a part of the leading edge feed cavity (512). The main body cavity (522a) is fluidly connected to the leading edge feed cavity (512) by an interior fluid connection (524) through the intersection of the two interior ribs (534). A shield cavity (530) is arranged to thermally shield the leading edge feed cavity (512) from heat pickup along the suction side (508) of the airfoil body (500).

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A method of joining polymer components includes additively manufacturing first and second mating features on first (12) and second (14) polymer components such that a mechanical lock is created through undercut geometric features of an adhesive material when the polymer components are joined. Adhesive (17) is added between the mating components to strengthen the joint.

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An airfoil (60) includes an airfoil wall (62) that defines a leading end (62a), a trailing end (62b), a first side (62c), and a second side (62d). Radially-extending ribs (64) partition the interior cavity of the airfoil into first and second cooling channels (72) and a radial cooling passage that is situated between the first and second cooling channels. The cooling channels extend to respective first and second channel ends. A turn channel connects the first and second channel ends. The turn channel splits at the first channel end into first and second channel legs (80, 82) such that there is a region between the first and second channel legs. The channels legs merge at the second channel end. The radial cooling passage extends through the region between the first and second channel legs.