General Electric Patente

A gas turbine engine (100) includes a turbomachine (109) having a compressor section (112), a combustion section, and a turbine section arranged in serial flow order along a centerline axis. The compressor section includes a plurality of airfoils (206) . Each of the plurality of airfoils includes a leading edge (209) a trailing edge (212) opposite the leading edge, a first sidewall (301) extending between the leading edge and the trailing edge, a second sidewall extending (302) between the leading edge and the trailing edge opposite the first sidewall, an airfoil root, and an airfoil tip opposite the airfoil root. Each of the plurality of airfoils defines a span height between the airfoil root and the airfoil tip, a leading edge thickness (306) between the first sidewall and the second sidewall, an airfoil length (303) extending between the leading edge and the trailing edge, a maximum thickness (309) between the first sidewall and the second sidewall, and an airfoil thickness parameter.

A combustor (26, 26a, 26b) for a gas turbine engine (10) includes a dome (56), a cowl (60) connected with the dome (56) via an outer connection (106) and via an inner connection (112), an outer liner (54) having an outer liner first end (102) connected with the outer connection (106) and an outer liner second end (104) arranged downstream of the outer liner first end (102), an inner liner (52) having an inner liner first end (108) connected with the inner connection (112) and an inner liner second end (110) arranged downstream of the inner liner first end (108), and a combustion chamber (62) defined between the outer liner (54) and the inner liner (52). An outer piston ring assembly (114) is connected at a first end (118) to the outer connection (106) and includes a piston ring (122) at a second end (120), and an inner piston ring assembly (124) is connected at a first end (128) to the inner connection (112) and includes a piston ring (132) at a second end (130).

A cavity consolidation apparatus for consolidating a preform (12) of a ceramic matrix composite (CMC) component (10) defining a cavity (14) includes a pressurization device (32) and an expandable band (34) wrapped around the pressurization device (32), wherein the expandable band (34) has a first band portion (40) with a first thickness in a first region (18) of the cavity (14) and a second band portion (42) with a second thickness in a second region (22) of the cavity (14), wherein the first thickness of the expandable band (34) is different from the second thickness of the expandable band (34).

A method of manufacturing a composite component (161') having an outer shell (165), an inner hub (167), and a plurality of struts (158) connecting said outer shell and inner hub. An outer shell preform (308) and an inner hub preform (312) are arranged adjacent to one another along a first assembly reference direction (428), and strut warp fiber tows (458) are woven with said outer shell preform and inner hub preform to form strut preforms (306) extending in between. The distance between said outer shell preform and inner hub preform is increased to shift a strut portion (464) of the plurality of strut warp fiber tows to extend in a second assembly direction (429). The strut preforms are formed by weaving strut weft fiber tows (496) and strut interlocking fiber tows (498) with the strut warp fiber tows (458) to form a preform assembly (300).

Provided herein are systems and methods for preforming tactile profilometry. In some aspects, a system (100) for performing tactile profilometry includes a probe (290) (102) with a profilometry sensor apparatus (108). The profilometry sensor apparatus (108) comprises a touch medium (120) having a first side (176A) adjacent to the surface (106) being inspected and a second side (176B) opposite the first side (176A), the second side (176B) having a reflective layer (136). The profilometry sensor apparatus (108) also includes a first light source (124) coupled to a first deformable waveguide (126) to illuminate the reflective layer (136) of the touch medium (120) with a first light beam at a first incidence angle and a light sensor (122) to detect reflected light. A first shaping actuator (128) is operatively coupled to the first deformable waveguide (126) to displace the first deformable waveguide (126). A controller (150) is configured to receive optical data (158) from the light sensor (122) and operate the first shaping actuator (128) to adjust the first incidence angle based on the optical data (158).

A sensor probe (110) may include at least one ultrasound transducer (320) and a controller (350). The controller (350) may be configured to: receive, via the at least one ultrasound transducer (320), ultrasound signals (112, 114) from a component surface (125) of a component (120); detect a pattern indicative of a communication from an other sensor probe (110), associated with the component (120), based on the received ultrasound signals (112, 114); decode the received ultrasound signals (112, 114) to retrieve a message from the other sensor probe (110), in response to detecting the pattern indicative of the communication from the other sensor probe (110); obtain, via the at least one ultrasound transducer (320), ultrasound data indicative of ultrasound waves (112, 114) in the component (120), in response to the retrieved message; and cause data analysis for detecting defects in the component (120) to be performed based on the obtained ultrasound data.

A coated component (10), along with methods of its formation, is provided that includes a silicon-containing substrate (12) having a surface (14); a bondcoat (16) on the surface (14) of the silicon-containing substrate (12), wherein the bondcoat (16) comprises silicon; a thermally grown oxide layer (20) on the bondcoat (16), wherein the thermally grown oxide layer (20) comprises a borosilicate glass (24) and a mullite-based material (28); and an environmental barrier coating (18) on the thermally grown oxide layer (20).

A sensor probe (110) may include a plurality of sensor elements (310), wherein a sensor element (310), of the plurality of sensor elements (310) includes an ultrasound transducer (320). The sensor probe (110) includes a body formed at least in part from a flexible material (370). The flexible material (370) provides acoustic coupling between the ultrasound transducer (320) and a component surface (125), and wherein the flexible material (370) establishes a conformal surface between the plurality of sensor elements (310) and the component surface (125). The sensor probe (110) may further include a locomotion actuator (420) operable to move the sensor probe (110) along the component surface (125); a controller (250) configured to control operation of the plurality of sensor elements (310) and the locomotion actuator (420).

A combustor (80, 300, 400) for a turbine engine (10) includes a dome (86, 306), an outer liner (82, 302) having a plurality of outer liner holes (302A) and one or more outer liner dilution holes (302B), an inner liner (84, 304) having a plurality of inner liner holes (304A) and one or more inner liner dilution holes (304B), the dome (86, 306), the outer liner (82, 302) and the inner liner (84, 304) defining a combustion chamber (88, 308), an outer baffle (312, 412) connected to the outer liner (82, 302) to define an outer cavity (316, 416) therebetween, the outer baffle (312, 412) including a plurality of outer baffle holes (312A, 412A), and an inner baffle (314, 414) connected to the inner liner (84, 304) to define an inner cavity (318, 418) therebetween, the inner baffle (314, 414) including a plurality of inner baffle holes (314A, 414A). The one or more outer liner dilution holes (302B) and the one or more inner liner dilution holes (304B) are configured to provide pressure drop across the outer liner (82, 302) and the inner liner (84, 304) to enhance penetration of a jet airflow (328, 332) into the combustion chamber (88, 308).

Blade assemblies (300) for gas turbine engines (100) are described herein. The blade assembly (300) includes a shank (302), a platform (304), and an airfoil (306). The airfoil (306) has a plurality of cooling holes (334). Cooling fluid flows through internal conduits (700a-700e) of the airfoil (306) and is vented by the cooling holes (334). The exterior surface (324) of the airfoil (306) is divided into zones with certain metering areas defined by the cooling holes (334).

Blade assemblies (300) for gas turbine engines (100) are described herein. The blade assembly (300) includes a shank (302), a platform (304), and an airfoil (306). The airfoil (306) has a plurality of cooling holes (334). Cooling fluid flows through internal conduits (700a-700e) of the airfoil (306) and is vented by the cooling holes (334). The exterior surface (324) of the airfoil (306) is divided into zones with certain metering areas defined by the cooling holes (334).

An airfoil (200, 300) including a core (200B, 300B), a shell (200A, 300A) covering the core (200B, 300B), and one or more sensors (202, 302) provided within the core (200B, 300B) or the shell (200A, 300A), or both. The one or more sensors (202, 302) are configured to generate an electromagnetic signal when the airfoil (200, 300) is subjected to an impact force, the electromagnetic signal being processed to detect an impact of an object on the airfoil (200, 300).

An airfoil (92) for a turbine engine (10) has an outer wall (106) bounding an interior (116) and defining an outer surface extending between a leading edge (112) and a trailing edge (114) in a chordwise direction, and between a root (98) and a tip (100) in a spanwise direction. A cooling circuit (128) within the interior (116) and a plurality of elongated outlets (130) fluidly coupled to the cooling circuit (128) and opening onto the outer wall (106), the plurality of elongated outlets (130) having a first elongated outlet (130a) and a second elongated outlet (130b).

A core structure (110, 130, 140, 150, 160, 170, 180, 200, 300, 320, 340) for a composite panel (100), the core structure (110, 130, 140, 150, 160, 170, 180, 200, 300, 320, 340) including a plurality of hollow ceramic cells (116, 322, 342, 342A), each of the plurality of hollow ceramic cells (116, 322, 342, 342A) defined by a plurality of ceramic walls (118, 144), each of the plurality of ceramic walls (118, 144) extending from a top (120, 120A, 120B, 124, 124A, 124B, 312A, 312B) to a bottom (122, 122A, 122B, 126, 126A, 126B, 314A, 314B) and at least one composite ply extending from the top (120, 120A, 120B, 124, 124A, 124B, 312A, 312B) of a first one of the plurality of walls to the bottom (122, 122A, 122B, 126, 126A, 126B, 314A, 314B) of the first one of the plurality of walls and extending from the first wall of the plurality of ceramic walls (118, 144) to a second wall of the plurality of ceramic walls (118, 144) and across a seam (312) defined by the first wall and the second wall.

A core structure (110, 130, 140, 150, 160, 170, 180, 200, 300, 320, 340) for a composite panel (100), the core structure (110, 130, 140, 150, 160, 170, 180, 200, 300, 320, 340) including a plurality of hollow ceramic cells (116), each of the plurality of hollow ceramic cells (116) comprising a plurality of ceramic walls (118, 144), each of the plurality of ceramic walls (118, 144) extending from a top (120) to a bottom (122) and a ceramic connector (152) connecting one of the plurality of ceramic walls (118, 144) of a first hollow ceramic cell (116A, 116C, 306A, 322A, 342A) of the plurality of hollow ceramic cells (116) to one of the plurality of ceramic walls (118, 144) of a second hollow ceramic cell (116B, 116D, 166B, 306B, 322B, 342B) of the plurality of hollow ceramic cells (116), the ceramic connector (152) being spaced from at least one of: the top (124) of the first hollow ceramic cell (116A, 116C, 306A, 322A, 342A), the top (124) of the second hollow ceramic cell (116B, 116D, 166B, 306B, 322B, 342B), the bottom (126) of the first hollow ceramic cell (116A, 116C, 306A, 322A, 342A), or the bottom (126) of the second hollow ceramic cell (116B, 116D, 166B, 306B, 322B, 342B).

A gas turbine engine includes a turbomachine having a compressor section, a combustion section, and a turbine section in serial flow order. A fan section defines an axial centerline and a rotor and includes a fan and a spinner assembly drivingly coupled to the turbomachine. The rotor supports a plurality of fan blades being rotatable about the axial centerline in a circumferential direction. A cowl assembly is positioned aft of at least a portion of the spinner assembly. The cowl assembly and the spinner assembly define at least part of an airflow surface rotatable with the fan and exposed to a fan airflow provided to and through the fan during operation of the gas turbine engine. The cowl assembly is at least partly cantileverly supported by the rotor.

A build tank (100) includes a tank wall (102) including an inner surface (106) and an outer surface (108) opposite the inner surface (106), a movable table (104) including an outer edge (120), and a seal assembly (128) provided between the inner surface (106) of the tank wall (102) and the outer edge (120) of the movable table (104), the seal assembly (128) including a spring plate (158) including a plurality of springs (174) contacting the outer edge (120) of the movable table (104) and a rope seal (162) provided between the spring plate (158) and the inner surface (106) of the tank wall (102) and contacting the inner surface (106) of the tank wall (102).

An apparatus may include a control circuit (102) to: access textual requirements (106) for aviation software (104), input the textual requirements for the aviation software to a first machine learning model (108), output test case input and execution conditions from the first machine learning model, input the textual requirements and the test case input and execution conditions to a second machine learning model (110) that is different from the first machine learning model, output expected test results from the second machine learning model, and form at least one test case (112) for the aviation software using the test case input, execution conditions, and expected test results.

An assembly (200) for a gas turbine engine (100), the gas turbine engine (100) defining an axial direction and a radial direction, the assembly (200) including a hub shaft (202) configured to support a fan blade (154) of the gas turbine, a propellor shaft (204) configured to connect to a gearbox (155), the propellor shaft (204) disposed aft of the hub shaft (202) in the axial direction, a connector (206) fixing the hub shaft (202) to the propellor shaft (204), and a bearing (208) disposed on the hub shaft (202) forward of the connector (206) in the axial direction.

A apparatus for monitoring fluid flow includes an upstream sensor assembly having a first transceiver-based sensor and a second transceiver-based sensor. Furthermore, the apparatus includes a downstream sensor assembly having a third transceiver-based sensor and a fourth transceiver-based sensor. In this respect, the first transceiver-based sensor and the second transceiver-based sensor are oriented relative to the third transceiver-based sensor and the fourth transceiver-based sensor such that the third transceiver-based sensor and the fourth transceiver-based sensor receive a first sensor signal emitted by the first transceiver-based sensor and a second sensor signal emitted by the second transceiver-based sensor.

A gas turbine engine (100, 200, 300, 400, 420, 440, 500) defining an axial direction and a radial direction includes a spinner (112, 404, 504) defining a spinner duct (410, 506) and a spinner inlet (408, 506) to the spinner duct (410, 506), a turbomachine (110, 402, 502) comprising a compressor section (114), a combustion section (116), and a turbine section (118) arranged in serial flow order, the turbomachine (110, 402, 502) defining a fan duct (172), a fan duct (172) inlet to the fan duct (172), a core duct (142), and a core inlet (124) to the core duct (142), and a primary fan (152) driven by the turbomachine (110, 402, 502), wherein the spinner inlet (408, 506) is upstream of the primary fan (152).

An airfoil assembly (130, 230) having a trunnion (134, 234) and a spar (136, 236). The trunnion (134, 234) having an inner surface (152, 252) defining a socket (153, 253) with a flared section with an open top. The spar extending from the socket (153, 253) and through the open top (154, 254). Further, the spar (136, 236) being formed by a plurality of composite plies (163, 263) including a radially innermost ply (160, 260) having a respective centerline (176, 276), and a radially outermost ply (162, 262) having a respective centerline (176, 276) that is parallel to the respective centerline (176, 276) of the radially innermost ply (160, 260).

A heat shield (210, 310, 410, 510) for a fuel nozzle includes an aft wall (220, 320, 420, 520), a forward wall (222, 322, 422, 522) disposed between the aft wall (220, 320, 420, 520) and a cooling flow passage (224) partially defined by the forward wall (222, 322, 422, 522), an outer wall (230, 330, 430, 530) extending between the aft wall (220, 320, 420, 520) and the forward wall (222, 322, 422, 522), an inner wall (226, 326, 426, 526) extending between the aft wall (220, 320, 420, 520) and the forward wall (222, 322, 422, 522), and an internal cooling cavity (234, 334, 434, 534) defined and enclosed by the aft wall (220, 320, 420, 520), the forward wall (222, 322, 422, 522), the outer wall (230, 330, 430, 530), and the inner wall (226, 326, 426, 526). A plurality of first exhaust apertures (236, 336, 436, 536) opens into the cooling flow passage (224) and extends through the forward wall (222, 322, 422, 522) into the internal cooling cavity (234, 334, 434, 534), a plurality of second exhaust apertures (238, 338, 438, 538) opens into the internal cooling cavity (234, 334, 434, 534) and extends through the aft wall (220, 320, 420, 520), and a plurality of third exhaust apertures (240, 340, 440, 540) opens into the cooling flow passage (224) and extends through the forward wall (222, 322, 422, 522) to an outer surface (223, 233, 323, 333, 423, 433, 523, 533) of the outer wall (230, 330, 430, 530).

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). The combustion section (14) includes a combustion chamber (50) and a fuel-air mixer (66) fluidly coupled with the combustion chamber (50). The fuel-air mixer (66) includes an outer wall (68), a center body (70) disposed radially inward of the outer wall (68), a first splitter (72) including a first lobed trailing edge (74), a second splitter (76) including a second lobed trailing edge (78) axially offset from the first lobed trailing edge (74), a first swirler (82) disposed at least partially between the first splitter (72) and the second splitter (76); and a set of fuel orifices (80) located at the first lobed trailing edge (74), the second lobed trailing edge (78), or the first lobed trailing edge (74) and the second lobed trailing edge (78).

In some embodiments, apparatuses and methods are provided herein useful to inspect aircraft engine. In some embodiments, a rigidizable insertion tool (100) includes a plurality of links (102) arranged in a sequence. The plurality of links (102) includes at least one deformable link (104) that is structurally deformable. The rigidizable insertion tool (100) may include a tension assembly (600) configured to apply a first tensioning force on the plurality of links (102) to actuate the plurality of links (102) from a relaxed state to a rigidized state having a first shape. The tension assembly (600) may apply a second tensioning force greater than the first tensioning force on the plurality of links (102) while in the rigidized state to cause structural deformation of the at least one deformable link (104) and change a shape of the plurality of links (102) from the first shape. The second tensioning force may change the shape from the first shape to a second shape.

An apparatus for monitoring fluid flow includes a computing system configured to the control the operation of a first transceiver-based sensor such that the first transceiver-based sensor emits a downstream sensor signal directed at a third transceiver-based sensor and a fourth transceiver-based sensor. Furthermore, the computing system is configured to control the operation of the third transceiver-based sensor such that the third transceiver-based sensor emits an upstream sensor signal directed at the first transceiver-based sensor and a second transceiver-based sensor. Additionally, the computing system is configured to determine redundant values of a flow parameter of the fluid based on the downstream sensor signal and the upstream sensor signal.

A heat shield (300) for a fuel nozzle (200) of a turbine engine (100) includes an shield flange (320) and a ceramic coating (400, 430, 440, 470, 480). The shield flange (320) has an axial direction, a radial direction, and a circumferential direction. The shield flange (320) includes a flange hot-side surface (322). The ceramic coating (400, 430, 440, 470, 480) is formed on the flange hot-side surface (322). The ceramic coating (400, 430, 440, 470, 480) includes a coating hot-side surface (402), a bulk region (412), and one or more surface regions (414). The one or more surface regions (414) are a portion of the coating hot-side surface (402) of the ceramic coating (400, 430, 440, 470, 480). Each surface region (414) of the one or more surface regions (414) has a surface roughness less than a surface roughness of each bulk region of the one or more bulk regions (412), a density greater than the density of each bulk region of the one or more bulk regions (412), or both.

An apparatus for monitoring fluid flow includes a computing system configured to initiate a sensor calibration operation, determine a first time period between when a first transceiver-based sensor emits a first sensor signal directed at a second transceiver-based sensor and when the first transceiver-based sensor receives a first reflection signal that is reflected off of the second transceiver-based sensor. The computing system is configured to determine a second time period between when the second transceiver-based sensor emits a second sensor signal directed at the first transceiver-based sensor and when the second transceiver-based sensor receives a second reflection signal reflected off of the first transceiver-based sensor. The computing system is configured to determine a difference between the first time period and the second time period and calibrate at least one of the first transceiver-based sensor or the second transceiver-based sensor when the determined difference exceeds a threshold value.

An aerospace lifecycle management service-oriented architecture (SOA) server (200) includes one or more processors (202); and one or more non-transitory computer-readable media (204) that collectively store: a process module (208) that, when executed, provides a service to a client device (102) upon receipt of a request from the client device (102); a translation module (210) that, when executed, allows the request to be executed by the process module (208); and instructions that, when executed by the one or more processors (202), configure the SOA server (200) to perform operations, the operations comprising: receiving the request from the client device (102), the request associated with the service to be provided to the client device (102); implementing the translation module (210) to modify the request such that the request can be processed by the process module (208); and after modifying the request, implementing the process module (208) to provide the service to the client device (102).

A method for generating a live engineering bill of materials (EBOM) (200) in a project lifecycle management (PLM) system (10) is provided. The method includes generating a lifecycle status (208) for each part of one or more parts of a product. The method further includes generating, with a bill of materials (BOM) generation module, the live EBOM (200) based on product data (116) and the lifecycle status (208). The live EBOM (200) includes a table (202) having at least one line (204) representing each part of the one or more parts in the product. Each line (204) of the at least one line (204) includes the lifecycle status (208) and at least one of the part name data (118), part ID data (120), or part quantity data (122).

An assembly (300) to control a gap between a rotor blade (192) and a casing in a turbine engine (100), the assembly (300) including a casing (301) having an annular casing body (303) and a liner (302) positioned adjacent an inner radial surface of the annular casing body (303), the casing (301) being separated from the rotor blade (192) by the gap in a radial direction. The annular casing body (303) has a lower coefficient of thermal expansion than that of the liner (302) and that of the rotor blade (192). The liner (302) is provided with spring extensions (304) having a specific design and coefficient of thermal expansion to control the gap between the rotor blade (192) and the casing (301).

A rotating machine torque measurement system (10) includes a shaft assembly (12) and spaced sensors (14, 16). The shaft assembly (12) includes a first shaft (18) that is adapted to connect to a torque generating device (24), a second shaft (20) that has a second end that is secured to a second end of the first shaft (18), and a third shaft (22) that has a first end that is secured to a first end of the second shaft (20) and is adapted to connect to a torque receiving device (26). The first, second, and third shafts (18, 20, 22) are coaxial and can twist or angularly deform in relation to each other in response to a load applied to the shaft assembly (12). The sensors (14, 16) are positioned at the first and second ends of the shaft assembly (12) to measure angular deformation of the shaft assembly (12).

A directed energy deposition (DED) additive manufacturing system (100) for manufacturing a component (102) from a material includes a deposition assembly (120) having a deposition head (122) through which material is deposited to form a top surface (102c) of a component (102). The top surface (102c) defines a width (w) measured in millimeters (mm). The deposition assembly (120) also includes compression rig (140) having a compression head (246). The compression head (246) is configured to apply a compressive force (F) measured in kilonewtons (kN) to the top surface (102c) of the component (102). The compression rig (140) also defines a target load (Y).

A combustor dome assembly (100, 150, 200) includes an inner liner (102, 206), an outer liner (104, 204), a dome (116, 152, 202) disposed on the inner liner (102, 206) and the outer liner (104, 204), and a cowl (112) connected to the inner liner (102, 206) and to the outer liner (104, 204), wherein the dome (116, 152, 202) is disconnected from the cowl (112) and defines an annular surface facing the cowl (112), the annular surface extending from an outer diameter (124) of the dome (116, 152, 202) to an inner diameter (126) of the dome (116, 152, 202).

A method (200) of imaging a turbine engine component (68) with a thermographic sensor, the turbine engine component (68) having a first surface (86) and a second surface (89) spaced from the first surface (86), and a plurality of holes with inlets (94, 94a, 94b, 94c, 94d) formed in the second surface (89) and outlets (95, 95a, 95b, 95c, 95d) formed in the first surface (86). The method (200) including flowing air (202) through the turbine engine component (68) and obtaining, during the flowing of air (202), thermographic data (204) when the turbine engine component (68) and a thermograph sensor (126) are relatively orientated such that a viewing plane (140) and a centerline (110a, 110b, 100c, 110d) of the outlet (95, 95a, 95b, 95c, 95d) or a centerline (108a, 108b, 108c, 108d) of a cooling hole (92, 100, 100a, 100b, 100c, 100d) form and angle (152) in a range from 60° to 120°. A subset of the thermographic data is determined (206) and a flow score for the cooling hole (92, 100, 100a, 100b, 100c, 100d) is calculated (208), wherein the thermographic data subset, the flow score, or the thermographic data subset and the flow score are visually displayed (210).

A build material escapement assembly (300) for an additive manufacturing system (100) includes a base (302) defining a cavity (304) and an aperture (306) opening into the cavity (304), a diaphragm (312) extending across the aperture (306) and coupled to the base (302), and a retractable plate (316) disposed within the cavity (304) and coupled to the diaphragm (312). The build material escapement assembly (300) further includes a post (320) disposed within the cavity (304) beneath the diaphragm (312) and coupled to the retractable plate (316), the post (320) supported by the base (302) and movable with the retractable plate (316) between a retracted position and an extended position. The build material escapement assembly (300) further includes a drive assembly (328) coupled to the post (320) to move the post (320) between the retracted position and the extended position.

Systems, apparatus, articles of manufacture, and methods are disclosed for reference markers for engine components. An example turbine engine (100) includes a substrate (118, 142) and a blade (164, 168, 172, 176, 706, 708) coupled to the substrate (118, 142), the blade (164, 168, 172, 176, 706, 708) including a first side having an external surface (214) and a reference marker (216, 218, 302, 402, 502, 602, 604, 606, 722, 724) provided on the external surface (214), the reference marker (216, 218, 302, 402, 502, 602, 604, 606, 722, 724) including spatial marking features having predetermined dimensions, the spatial marking features including (a) a first spatial marking feature at a first location on the external surface (214) and (b) a second spatial marking feature at a second location on the external surface (214) that is different than the first location, wherein a combination of the first spatial marking feature and the second spatial marking feature provide a first measure of the blade (164, 168, 172, 176, 706, 708) based on the predetermined dimensions.

Ceramic matrix composite (CMC) fasteners and fastener systems, as well as methods of forming CMC fasteners and fastener systems, are provided. For example, a CMC fastener includes a CMC material forming a body portion having a first plurality of continuous fibers disposed in a first matrix, and a thread portion having a second plurality of continuous fibers disposed in a second matrix. The body portion has a body length in an axial direction. The thread portion is wound in a helix about the body portion such that the second plurality of continuous fibers wrap about the first plurality of continuous fibers along the axial direction. A CMC fastener system can include the CMC fastener and a CMC threaded nut.

An auto-visual data processing and insight generation system (100) for an aircraft engine includes a processor (720) and a memory (730) including processor executable instructions that cause the system to: perform video processing on image frames (501, 502), wherein the video processing includes selecting frames of the image frames as key frames based upon feature criteria including a change in pixels, a trend change in graphs, or a specified filter in the frames; perform an extraction of a feature included in the key frames, wherein the feature is extracted from a two-dimensional image to generate a three-dimensional model; identify an insight regarding the feature based on the three-dimensional model to predict performance of a component of the aircraft engine; verify an analytical or physical growth model based on the identified insight; and tune a design of the aircraft engine based on a verification of the analytical or physical growth model.

A three-dimensional woven fabric (400, 500, 600, 700) for a composite airfoil (300) for a turbine engine (100) and methods of manufacturing such a fabric. The method includes forming, during weaving a plurality of reinforcing fiber tows (202), a spar section (410) of the woven fabric (400, 500, 600, 700). The method also includes forming, during weaving the plurality of reinforcing fiber tows (202), a first edge section (420, 440) integrally woven with the spar section (410) and extending in a second direction therefrom and forming, during weaving the plurality of reinforcing fiber tows (202), a second edge section (430, 450) integrally woven with the spar section (410) and extending in the second direction therefrom, the second edge section (430, 450) being positioned opposite the first edge section (420, 440) with a gap (462, 464) formed therebetween.