General Electric

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.

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

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).

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.

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.