General Electric Patente

A method including applying a mask to at least a portion of a surface of a component less a defect surface, removing at least one metal oxide from the defect surface, electrodepositing metallic powder particles including a superalloy material on the defect surface by immersing the defect surface in an electrolyte bath having metal ions and the metallic powder particles including the superalloy material, agitating the electrolyte bath, and applying a current, removing the mask, and heat treating at least a portion of the component including the defect surface.

A variable stator vane (200) includes a shaft (202) defining a longitudinal axis (204) of the variable stator vane (200), the shaft (202) having a first metallic end (202A) and a second metallic end (202B) coupled to a shaft portion (202C), the shaft portion (202C) being made essentially of a first composite material to reduce weight of the variable stator vane (200). The variable stator vane (200) further includes an airfoil portion (206) coupled to the shaft portion (202C), the airfoil portion (206) being made essentially of a second composite material. The shaft (202) and the airfoil portion (206) are rotatable around the longitudinal axis (204) of the variable stator vane (200).

Systems (200) and methods (10, 20) for forecasting damage in engine components using advanced machine learning techniques, including hybrid frameworks and spatiotemporal modeling integrate annotated inspection images (13a) and operational engine data to generate forecasted damage states and operational parameters for future engine cycles. The machine learning model (17) includes convolutional layers to extract spatial features, recurrent layers to analyze temporal patterns, and an output layer that predicts future damage states, operational conditions, and distress ranks of engine components. Physics-informed constraints are incorporated to ensure predictions adhere to physical laws, such as irreversibility of damage progression. By leveraging these forecasts, the system optimizes maintenance schedules, reduces unnecessary inspections, and extends time-on-wing of engine components. This novel approach provides a significant improvement over traditional methods by automating damage analysis, capturing multi-failure mechanisms, and enhancing predictive maintenance strategies.

Aspects of the present disclosure provide techniques for additive manufacturing anomaly detection. An example method includes obtaining, from an electromagnetic energy sensor (202, 204), a time-series signal (410) comprising a plurality of samples, wherein the samples correspond to electromagnetic emissions from an interaction between a first energy source (120) and powder (142) at a first focal point on a layer (172) of additive material (142) and the samples are generated at a sampling rate greater than or equal to a first sampling rate; decomposing the time-series signal into at least a high-frequency signal variation component (430); computing one or more spatiotemporal statistical values from the high-frequency signal variation component of the time-series signal; identifying, with a machine learning model (526), one or more perturbations, from the one or more spatiotemporal statistical values, that exceeds a predetermined threshold corresponding to the one or more spatiotemporal statistical values; and generating an alert in response to identifying the one or more perturbations.

A swirler (92, 92a) for a fuel nozzle assembly (58) of a gas turbine (10) includes a first wall (114), a second wall (118) extending circumferentially about a swirler centerline axis (94), a splitter wall (122), a first plurality of radial swirl vanes (124) extending between the first wall (114) and the splitter wall (122), a second plurality of radial swirl vanes (126) extending between the splitter wall (122) and the second wall (118), and a plurality of vortex generating elements (132, 152, 176, 198) including at least one of (i) a first plurality (134) of vortex generating elements (132, 152) arranged on the first wall (114), (ii) a second plurality (136) of vortex generating elements (132, 176) arranged on a radial wall portion (168) of the splitter wall (122), or (iii) a third plurality (138) of vortex generating elements (132, 198) arranged on an axial wall portion (170) of the splitter wall (122).

A fuel deoxygenator system includes a stripping gas line, a contactor including a fuel inlet fluidly connected to a fuel tank, a gas inlet, and a mixture outlet, a separator including a mixture inlet fluidly connected to the mixture outlet, a fuel outlet, and a gas outlet fluidly connected to the stripping gas line, a recuperator including a first fluid inlet, a first fluid outlet, a second fluid inlet, and a second fluid outlet, and a catalyst including an inlet and an outlet. The first fluid inlet of the recuperator is fluidly connected to the gas outlet of the separator, the first fluid outlet of the recuperator is fluidly connected to the inlet of the catalyst, the second fluid inlet is fluidly connected to the outlet of the catalyst, and the second fluid outlet is fluidly connected to the gas inlet of the contactor.

An annular vortex combustor (200, 300, 400, 500, 600, 700, 800, 900, 1000, 1100, 1200) for a turbine engine (10). The annular vortex combustor (200, 300, 400, 500, 600, 700, 800, 900, 1000, 1100, 1200) defines a forward end (202, 212, 220, 320, 420, 502, 512, 620, 720, 920, 1020) and an aft end (204, 222, 322, 422, 504, 622, 722, 922, 1022). The annular vortex combustor (200, 300, 400, 500, 600, 700, 800, 900, 1000, 1100, 1200) includes a combustor liner (214, 314, 414, 614, 714, 814, 914, 1014, 1214) and a combustion chamber (224, 324, 424, 624, 724, 824, 924, 1024). The combustor liner (214, 314, 414, 614, 714, 814, 914, 1014, 1214) includes an outer liner (216, 316, 416, 616, 716, 816, 916, 1016, 1216), an inner liner (218, 318, 418, 618, 718, 918, 1018, 1218), and a dome (221, 321, 421, 621, 721, 921, 1021) at the forward end (202, 212, 220, 320, 420, 502, 512, 620, 720, 920, 1020). The dome (221, 321, 421, 621, 721, 921, 1021) has a generally arc shape and extends from the outer liner (216, 316, 416, 616, 716, 816, 916, 1016, 1216) to the inner liner (218, 318, 418, 618, 718, 918, 1018, 1218). The combustion chamber (224, 324, 424, 624, 724, 824, 924, 1024) is defined by the outer liner (216, 316, 416, 616, 716, 816, 916, 1016, 1216), the inner liner (218, 318, 418, 618, 718, 918, 1018, 1218), and the dome (221, 321, 421, 621, 721, 921, 1021), and extends from the forward end (202, 212, 220, 320, 420, 502, 512, 620, 720, 920, 1020) to the aft end (204, 222, 322, 422, 504, 622, 722, 922, 1022). The combustion chamber (224, 324, 424, 624, 724, 824, 924, 1024) has a primary combustion zone (252, 352, 452, 652, 752, 952, 1052) defined by the dome (221, 321, 421, 621, 721, 921, 1021). The combustion chamber (224, 324, 424, 624, 724, 824, 924, 1024) is configured to combust a mixture of a fuel flow (250, 350, 450, 650, 750) and a compressed air flow (240) in the primary combustion zone (252, 352, 452, 652, 752, 952, 1052) to generate a primary zone vortex (254, 354, 454, 654, 754, 954, 1054), the generally arc shape of the dome (221, 321, 421, 621, 721, 921, 1021) shaping the primary zone vortex (254, 354, 454, 654, 754, 954, 1054).

A turbine engine having an engine core extending along an engine centerline. An engine component includes an internal cooling passage defined by an inner surface and a turbulator located within the internal cooling passage. The turbulator includes a front face, an upstream face, a downstream face, a first top face, and a second top face. The turbulator is curved to define a turbulator nose having a radius of curvature greater than zero. The engine component also includes a cavity width, a turbulator length, and a cavity aspect ratio

An airfoil for a gas turbine engine can include a cooling aperture (150) for local thermal heat management and to provide a cooling film along an exterior of the airfoil. The cooling aperture is positioned within an outer wall defining an interior for the airfoil and a cooling circuit provides a cooling fluid flow to the cooling aperture. The cooling aperture (150) includes a cavity (152) at least partially defined by an impingement portion (154) , a bend (146), and a diffusion portion (156) hat is arranged at an angle relative to the impingement portion. A pin (170) is provided within the cooling aperture.

An airfoil (100) for a turbine engine (10) has an exterior wall (102), a set of cooling conduits (120), a tip rail (122) and an outlet. The exterior wall (102, 202, 302) bounds an interior (103). The exterior wall (102) defines extends between a leading edge (104) and a trailing edge (106) in a chordwise direction (Cd). The exterior wall (102) extends between a root (108) and a tip (110) in a spanwise direction (Sd). The set of cooling conduits (120) are located within the interior (103). The tip rail (122) extends in the spanwise direction (Sd). The outlet is located along the tip rail (122).

Systems and methods for processing effluent (105) from a cleaning operation performed on a gas turbine engine (102) are provided herein. In some embodiments, such a system (100) comprises a foam collapsing unit (201A) (110), a particulate removal unit (201B), and/or a microbe treatment unit (201D). The system (100) may also include a detergent digestion unit (201C). The foam collapsing unit (201A) (110) is in fluid communication with an effluent (105) reservoir (104) and is operative to reduce aeration of the effluent (105) and generate a defoamed effluent (107A, 107B, 107C). The particulate removal unit (201B) is operative to separate particulates (239) from the defoamed effluent (107A, 107B, 107C). The microbe treatment unit (201D) is operative to reduce microbial load of the defoamed effluent (107A, 107B, 107C). The particulate removal unit (201B), and the microbe treatment unit (201D) are disposed in serial flow order to treat the defoamed effluent (107A, 107B, 107C) and output a recovered fluid (109) stream.

A turbine vane frame assembly (200) defining an axial direction, a radial direction, and a circumferential direction includes an outer frame (202) extending from a forward end to an aft end (106, 212, 216, 220, 224), the outer frame (202) including a first connector (234) at the forward end and a second connector (236, 280, 290, 300) at the aft end (106, 212, 216, 220, 224), an outer fairing (204) extending from a forward end to an aft end (106, 212, 216, 220, 224), the outer fairing (204) connected to the outer frame (202) by the first connector (234) and the second connector (236, 280, 290, 300), an inner frame (204, 206) extending from a forward end to an aft end (106, 212, 216, 220, 224), the inner frame (204, 206) including a third connector (238) at the forward end and a fourth connector (240) at the aft end (106, 212, 216, 220, 224), and an inner fairing (208) extending from a forward end to an aft end (106, 212, 216, 220, 224), the inner fairing (208) connected to the inner frame (204, 206) by the third connector (238) and the fourth connector (240), wherein the first connector (234) provides support in the circumferential direction to the outer fairing (204), the second connector (236, 280, 290, 300) provides support in the axial direction to the outer fairing (204), and the fourth connector (240) provides support in both the circumferential direction and axial direction to the inner fairing (208).

A turbofan engine defining a circumferential direction is provided. The turbofan engine includes a turbomachine; an unducted rotor assembly drivingly coupled to the turbomachine, the unducted rotor assembly including a plurality of unducted rotor blades; and a plurality of outlet guide vanes positioned downstream of the plurality of unducted rotor blades, the plurality of outlet guide vanes each defining a span, the spans of the plurality of outlet guide vanes being nonuniform along the circumferential direction.

A turbine blade assembly (100) for a gas turbine engine (10) can include a cooling aperture (150) for local thermal heat management and to provide a cooling film along an exterior surface (134) of an airfoil (104) of the turbine blade assembly (100). The cooling aperture (150) is positioned within an outer wall (118) defining an interior (130) for the airfoil (104) and a cooling circuit (132) provides a cooling fluid flow to the cooling aperture (150). The cooling aperture (150) is at least partially defined by an outer surface (160), an inner surface (162), a first side (164), a second side (166), and a flared portion (170). The cooling aperture (150) exhausts at a diffuser outlet (156) extending between the first side (164) and the second side (166).

A system (1), method (400), and program for generating self-consistent assurance cases (60) using one or more computer models (20) configured to receive a query (24) from a user, to process the query (24) and determine the nature of the assurance case to be generated, to determine and retrieve from a normalized, curated database (10) the evidence to support the assurance case claims, and to populate an assurance case candidate (62) with the same. The assurance case candidate (62) may then be evaluated by a critic (40) against one or more constraints (32) and then either returned to the computer model (20) for re-synthesis or forwarded to a translator (50) to be converted into a human-readable format.

A system (1), method (400), and program for generating self-consistent assurance cases (60) using one or more computer models (20) configured to receive a query (24) from a user, to process the query (24) and determine the nature of the assurance case to be generated, to determine and retrieve from a normalized, curated database (10) the evidence to support the assurance case claims, and to populate an assurance case candidate (62) with the same. The assurance case candidate (62) may then be evaluated by a critic (40) against one or more constraints (32) and then either returned to the computer model (20) for re-synthesis or forwarded to a translator (50) to be converted into a human-readable format.

Turbine engine with a nozzle having cooling holes is disclosed herein. An example nozzle vane (184) includes a platform (214) and an airfoil (186) extending from the platform (214), the airfoil (186) including a leading edge (222); a trailing edge (224); a rib (306); an outer wall (218) defining an interior (302) adjacent to the trailing edge (224); and an interior wall (310) extending through the interior (302), the interior wall (302) and the outer wall (218) defining a near-wall passage (318) extending in a chordwise direction of the airfoil (186), the interior wall (318) including a first end (324A) proximate to the leading edge (222), the near-wall passage (318) including an inlet (322) between the first end (324A) and the rib (306); and a second end (324B) coupled to the outer wall (218), the second end (324B) proximate to the trailing edge (224).

A casing (200) for a turbine engine (100) including a composite section (230). The composite section (230) has (i) an arcuate shape or an annular shape and (ii) a circumferential direction. The composite section (230) includes a matrix (236) and a plurality of circumferential reinforcing fiber tows (250) embedded in the matrix (236). Each circumferential reinforcing fiber tow (250) of the plurality of circumferential reinforcing fiber tows (250) extends in the circumferential direction and has a plurality of undulations (252) in the circumferential direction to allow the matrix material of the composite section (230) to expand circumferentially.

A turbine blade assembly (100) for a gas turbine engine (10) can include a cooling aperture (140) for local thermal heat management and to provide a cooling film along an exterior (134) of an airfoil (104). The cooling aperture (140) is positioned within an outer wall (118) of the airfoil (104) defining an interior (130) for the airfoil (104) and a cooling circuit (132) provides a cooling fluid flow to the cooling aperture (140). The cooling aperture (140) includes a cavity (142) at least partially defined by an outer surface (148), an inner surface (156), a first side (150), and a second side (152). The cavity (142) exhausts at a diffuser outlet (144) extending between the first side (150) and the second side (152).

Turbine engine with a nozzle having cooling holes is disclosed herein. An example nozzle vane (184) includes an airfoil (186) including a trailing edge (224), a leading edge (222), an outer wall (218) defining a cavity (312) adjacent to the trailing edge (224), an interior wall (310) in the cavity (312), the interior wall (310) and the outer wall (218) defining a near-wall passage (318) extending in a chordwise direction of the airfoil (186); a first cooling passage array (1204) extending from the near-wall passage (318) to an exterior of the airfoil (186), the first cooling passage array (1204) including a first cooling passage (814, 1110) having a first inlet (1140) and a first outlet (1142); and a second cooling passage array (1202) extending from the near-wall passage (318) to the exterior of the airfoil (186), the second cooling passage array (1202) including a second cooling passage having a second inlet (1134) and a second outlet (1136), the second inlet (1134) downstream of the first inlet (1140), the second outlet (1136) upstream of the first outlet (1142).

An apparatus for detecting a ground fault of an electric power system (12) of an aircraft (10), the apparatus including a first variable resistor (102) configured to be connected to the electric power system (12), a second variable resistor (104) configured to be connected to the electric power system (12), a switch (110) movable from an open position to a closed position, the switch (110) configured to connect an electrical load (20) to the first variable resistor (102) and the second variable resistor (104) in the closed position, and a controller (112) configured to detect the ground fault based on a first voltage across the first variable resistor (102) when the switch (110) is in the opened position, a second voltage across the second variable resistor (104) when the switch (110) is in the opened position, a third voltage across the first variable resistor (102) when the switch (110) is in the closed position, and a fourth voltage across the second variable resistor (104) when the switch (110) is the closed position.

Turbine engine with a nozzle having cooling holes is disclosed herein. An example nozzle vane (184) includes an airfoil including an outer wall (216) defining a cavity (312); a first extension (816) fluidly coupled to the cavity (312); an interior wall (310) extending through the cavity (312), the interior wall (310) and the outer wall (216) defining a near-wall passage (318) extending in a chordwise direction of the airfoil (186); a second extension (802) fluidly coupled to the near-wall passage (318), the second extension (802) extending chordwise from the near-wall passage (318) toward the trailing edge (224); a first cooling passage (818, 1104) fluidly coupled to the first extension (816); and a second cooling passage (812, 1108) fluidly coupled to the second extension (802), the first cooling passage (818, 1104) and the second cooling passage (812, 1108) having respective outlets (1122, 1136) in the outer wall (216).

Engines of an aircraft (120) are controlled during a mission including a takeoff phase, a climb phase, and a cruise phase, the climb phase including a top of climb (ToC) period prior to transitioning to the cruise phase. A controller may control the one or more engines (600) to output a takeoff thrust during the takeoff phase, the takeoff thrust corresponding to the maximum thrust during the takeoff phase. The controller may further control the one or more engines (600) to output a ToC thrust during the ToC period, the ToC thrust corresponding to the average thrust during the ToC period, wherein the one or more engines (600) is controlled such that a thrust ratio is between 1.1 to 1.6, the thrust ratio being defined as a ratio of the ToC thrust as normalized by a ToC ambient static pressure to the takeoff thrust as normalized by a takeoff ambient static pressure.

An unducted turbine engine system (200) may include an emitter (212) positioned to direct an energy beam (214) towards a surface (205) of an airfoil (204) of an engine (100) to produce a reflected beam (216). The unducted turbine engine system (200) may include a detector (218) positioned to capture data from the reflected beam (216). The unducted turbine engine system (200) may include a processor (222) configured to determine a condition of the engine (100) based on data captured by the detector (218).

A gas turbine engine (10) including a compressor section (12), a combustion section (14), and a turbine section (16) in serial flow arrangement, with the combustion section (14) having a fuel nozzle (101, 201, 301, 401, 501, 601, 701). The fuel nozzle (101, 201, 301, 401, 501, 601, 701) includes a fuel nozzle body (102, 302, 402, 502, 602, 702) defining a central channel (112, 312, 412, 512, 612, 712) comprising a compressed air passage (114, 314, 514, 614) and a mixing tube (116, 316, 516, 616), the fuel nozzle body (102, 302, 402, 502, 602, 702) defining a fuel nozzle centerline (104, 304, 404, 504, 604, 704). The fuel nozzle (101, 201, 301, 401, 501, 601, 701) includes an inner air flow path (134, 334, 335, 434, 534) supplying a first air flow (Ai) to the mixing tube (116, 316, 516, 616), an outer air flow path (138, 238, 338, 339, 438, 538, 638), circumscribing the inner air flow path (134, 334, 335, 434, 534), and supplying a second air flow (A₂) to the mixing tube (116, 316, 516, 616), a set of fluid diodes (140, 240, 340, 440, 540, 640, 740) including a subset (141, 241, 342, 343, 541, 647) of the set of fluid diodes (140, 240, 340, 440, 540, 640, 740) located within the outer air flow path (138, 238, 338, 339, 438, 538, 638) and generating a set of vortices within the outer air flow path (138, 238, 338, 339, 438, 538, 638), and a fuel supply (150, 350, 550, 650) including a set of fuel orifices (156, 556, 658) through which a fuel (F) flow is supplied to the mixing tube (116, 316, 516, 616).

An airfoil for a gas turbine engine (10) includes a leading edge (220), a trailing edge (225), a pressure sidewall (210) extending between the leading edge (220) and the trailing edge (225), a suction sidewall (215) opposite the pressure sidewall (210) and extending between the leading edge (220) and the trailing edge (225), and a flared portion (303, 403, 503, 603, 703, 803, 903) extending from the suction sidewall (215). The flared portion (303, 403, 503, 603, 703, 803, 903) extends from a flare lead point (310) towards the trailing edge (225). The flare lead point (310) is downstream of a throat point (313) on the suction sidewall (215). The throat point (313) is at a minimum distance between the suction sidewall (215) of the airfoil and another pressure sidewall (210) of an adjacent airfoil.

A method (500) for cleaning one or more deposited layers (116) of an additively manufactured part is disclosed. The method (500) includes curing a first layer (116) of a first material on a build plate (110). The first material includes a first component, a second component, and a third component. The method (500) further includes placing the first layer (116) in a bath (200) of a cleaning solution (204) that includes the third component, agitating the bath (200) such that an excess of the first material is removed, and curing a second layer (116) of a second material. The second material includes the third component.

A gas turbine engine (10) includes a turbomachine (16) having a compressor section (21), a combustion section (26), and a turbine section (27). The turbine section (27) includes a band (270, 275, 280, 285) having an upstream end (320) and a downstream end (328). The band (270, 275, 280, 285) extends between the upstream end (320) and the downstream end (328), and the band (270, 275, 280, 285) at least partially defines the working gas flow path (33). A plurality of airfoils (325) extend into the working gas flow path (33) from the band (270, 275, 280, 285). Each airfoil of the plurality of airfoils (325) includes a leading edge (330), a trailing edge (335), a first side (340), and a second side (345) opposite the first side (340). Each of the plurality of airfoils (325) is substantially symmetric across an airfoil centerline (310) extending through a center of each of the plurality of airfoils (325). The band (270, 275, 280, 285) defines a valley portion (355) adjacent the leading edge (330) of each of the plurality of airfoils (325) and a pair of hill portions (350) on opposing sides of the valley portion (355).

A method of forming a ceramic matrix composite (CMC) component (230) includes providing a ceramic structure (242) defining at least one wall (246) where the at least one wall (246) includes one or more bores (262) extending through the wall (246). At least one bore (262) of the one or more bores (262) has an inlet end (264) proximate a first side (248) of the wall (246) and an outlet end (266) proximate a second side (250) of the wall (246). One or more CMC plies (272) are disposed on at least a portion of the second side (250) to cover the outlet end (266). The one or more CMC plies (272) are densified and bonded to the ceramic structure (242). An environmental barrier coating (282) is applied on least a portion of the one or more CMC plies (272) covering the outlet end (266) of the at least one bore (262).

An electronic component (100) includes a first subassembly (102) and a second subassembly (130). The first subassembly (102) includes a first power overlay (POL) (104) that provides electrical connections to at least one semiconductor device (112) included in the first subassembly (102). The second subassembly (130) includes a second POL (132) that provides electrical connections to at least one semiconductor device (140) included in the second subassembly (130). The second subassembly (130) is disposed over the first subassembly (102) with the second POL (132) oriented toward the first POL (104) and with electrical connections between the first and second subassemblies (102)(130) that form a power module.

Coated components (10) and their methods of formation are provided. The coated component (10) may include: a substrate (12) having a surface (14); a bondcoat (16) on the surface (14) of the substrate (12); and an environmental barrier coating (18) on the bondcoat (16). At least one of the substrate (12), the bondcoat (16) or an environmental barrier coating (18) includes at least one dopant source. The at least one dopant source comprises at least one of germanium, phosphorus, vanadium, arsenic, antimony, or a mixture thereof.

Methods are described herein for assisting with coating (604A) removal. In some approaches, a method (200) for removing a coating (604A) from a coated component (100, 700, 800, 900) is provided. The coated component (100, 700, 800, 900) comprises a substrate (102, 302, 602, 702, 802, 902) with a coating (604A) disposed on the substrate (102, 302, 602, 702, 802, 902). The coating (604A) may have one or more defects therein. The method includes forming a plurality of recesses (314, 414, 614, 714, 814, 914) in the coating (104) that extend from a surface (322, 422) of the coating (104) through one or more layers of the coating (104) to a target layer (320, 420, 720, 820, 920) to produce an etched coating (304, 404, 604B, 704, 804, 904, 604B), and removing the one or more layers of the coating (104) by subjecting the etched coating (304, 404, 604B, 704, 804, 904, 604B) to at least one of a mechanical removal process or a chemical removal process to expose the target layer (320, 420, 720, 820, 920).

Methods are provided herein for coating removal. In some embodiments, the methods can be used to remove an environmental barrier coating (EBC) (104). The methods of coating removal described herein can be used to remove a coating from a component using chemical digestion of the coating by a liquid reagent (122) in a closed vessel (124, 150). The chemical digestion is performed in the closed vessel (124, 150) at elevated temperatures and/or elevated pressures. Heat is applied to the closed vessel (124, 150) to achieve elevated temperatures to accelerate the digestion reaction. In some embodiments, heating is accomplished by applying microwave energy to the closed vessel (124, 150).

An airfoil (154) for an aircraft engine (100) includes a blade body (194) and a leading edge protector (225) operably coupled to the blade body (194). The leading edge protector (225) includes a preformed sheet (220) defining a nose portion (222) and two flange portions (224). The nose portion (222) and the flange portions (224) are arranged in the chordwise direction. The preformed sheet (220) further defines a defined point (230) located on the nose portion (222). An X-axis and a Y-axis are defined relative to the defined point (230). A deposit (232) is operably coupled with the nose portion (222) of the preformed sheet (220) and extends from the nose portion (222) along the flange portions (224) on opposing sides of the defined point (230) in the X-axis direction.

A hybrid propulsion system includes an engine which includes a metal-air battery pack, a combustion chamber, and a propulsion power bus. The propulsion power bus is configured to receive energy from the metal-air battery pack and the combustion chamber, regulate voltage of the engine, manage an electric load of the engine, monitor electrical system faults or abnormalities, manage the heat generated within the engine, and distribute a stable and reliable power supply to propulsion-related components, such as the propeller or turbine.

An apparatus (300) may include processors (305), memory modules (340), and machine-readable instructions stored in the memory modules (340). When executed by the processors (305), the instructions may cause the apparatus (300) to receive build data associated with a part (122) to be built by an additive manufacturing machine (100) using two or more lasers, wherein a first portion (402) of the part (122) is to be built by a first laser (120) and a second portion (404) of the part (122) is to be built by a second laser (121), identify a contour of the part (122) based on the build data, receive image data of a layer of the part (122) while the part is being built by the additive manufacturing machine (100), identify pixel locations of an edge of the part (122) based on the image data, determine distances between the pixel locations of the edge and the contour, and determine misalignment between the two or more lasers based on the determined distances.

Provided herein are conformable sensor assemblies and methods for manufacturing conformable sensor assemblies. In some embodiments, a method of manufacturing a conformable sensor assembly (100A, 100B, 100C) includes printing a high-resolution sensor element (106, 204, 604) on a first layer (102, 102B, 206, 306) using at least one of an electrohydrodynamic inkjet or an ultra-high resolution micro-dispensing process. The first layer (102, 102B, 206, 306) comprises a flexible or a stretchable substrate. The method also includes printing a mid-resolution sensor element (116, 304) on a second sensor layer (206, 306) using at least one of an aerosol, inkjet jet, screen-printing, or copper etching process. The second sensor layer (206, 306) is made from a flexible or a stretchable substrate. The mid-resolution sensor element (116, 304) having a trace spacing (606) in the range of about 20 microns to about 100 microns. The high-resolution sensor element (106, 204, 604) has a trace spacing (606) of less than about 20 microns. The first layer (102, 102B, 206, 306) is coupled with the second sensor layer (206, 306) to form a sensor assembly (100A, 100B, 100C).

A sensor assembly (100) may include at least one stretchable layer (102) including a stretchable substrate (103) having a modulus of less than about 5 megapascals (MPa) and an elongation of break of greater than about 25 percent about at least one axis, and at least one flexible layer (104) including a flexible substrate (105) having an elongation of break of greater than about 10 percent about at least one axis, the at least one flexible layer (104) having a plurality of cuts (106, 206) formed therein. The sensor assembly (100) may include an actuation layer (116) including at least one contact pressure actuator (118) for adjusting contact pressure of a target surface (154) on the sensor assembly (100).

Eddy current sensors are provided herein that are stretchable. In some aspects, the stretchable eddy current sensors have a sensor (108) element array that is designed to act as an eddy current inspection sensor (108) and multiaxial strain sensor (108) to enable distortion compensation for measurement signals. Using strain detected by the eddy current sensors and, in some aspects, combining information about the geometry of a component (104) and position information of the sensor (108) on the component (104), corrections can be determined to make the eddy current sensor (108) readings invariant to deformations of sensor (108) elements (111) of the eddy current sensors. In other aspects, the stretchable eddy current sensors can be installed on a component (104) to be inspected for health monitoring. A baseline measurement can be acquired after installation, for example, after the stretchable eddy current sensor (108) is initially stretched, and defects detected based on changes from the baseline measurement.

An airfoil assembly (100) for a turbine engine (10) having a composite spar (120), a sleeve (130), and a set of lobes (150). The composite spar (120) includes an exterior spar portion (128), where at least part of the exterior spar portion (128) of the composite spar (120) is received at a sleeve inner surface (134) of the sleeve (130). The set of lobes (150) extends from the composite spar (120) and is received by a set of recesses (198) at the sleeve inner surface (134).