Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: A combustor of a gas turbine engine and a turbine shroud for a turbine of a gas turbine engine are disclosed. The combustor is configured to ignite a mixture of compressed air and fuel in a combustion chamber included therein. The turbine shroud is configured to direct products of the combustion reaction toward a plurality of rotatable turbine blades of the turbine to cause the plurality of turbine blades to rotate.

Abstract: A gas turbine engine includes a turbine having turbine blades and a shroud having a blade track for positioning radially outside of the blades. At least one of the blades and the blade track are secured in place using a dovetail connection between a dovetail receiver and a dovetail assembly formed to include a contoured root.

Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: A gas turbine engine includes a turbine having turbine blades and a shroud having a blade track for positioning radially outside of the blades. At least one of the blades and the blade track are secured in place using a dovetail connection between a dovetail receiver and a dovetail assembly formed to include a contoured root.

Abstract: A combustor of a gas turbine engine and a turbine shroud for a turbine of a gas turbine engine are disclosed. The combustor is configured to ignite a mixture of compressed air and fuel in a combustion chamber included therein. The turbine shroud is configured to direct products of the combustion reaction toward a plurality of rotatable turbine blades of the turbine to cause the plurality of turbine blades to rotate.

Abstract: A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

Abstract: A relatively low cost, lightweight and thermal stress-free mechanical fastener system having particular application for reliably joining together high temperature structural members (e.g., a pair of flat fiber-reinforced ceramic composite plates). The mechanical fastener system includes a ceramic composite fastener having a semi-circular head at one end to be countersunk in the structural members to be joined together and a dove tail retention feature formed in the shank or root at the opposite end. The composite fastener has a 2-dimensional (i.e., flat) profile that facilitates an economic manufacture thereof from densified ceramic composite material. A matched pair of thread forms having external threads and an internal dove tail relief to match the dove tail retention feature at the root of the composite fastener is held in face-to-face mating engagement with one another so as to establish a mechanical interlock around the root of the composite fastener.

Abstract: A relatively low cost, lightweight and thermal stress-free mechanical fastener system having particular application for reliably joining together high temperature structural members (e.g., a pair of flat fiber-reinforced ceramic composite plates). The mechanical fastener system includes a ceramic composite fastener having a semi-circular head at one end to be countersunk in the structural members to be joined together and a dove tail retention feature formed in the shank or root at the opposite end. The composite fastener has a 2-dimensional (i.e., flat) profile that facilitates an economic manufacture thereof from densified ceramic composite material. A matched pair of thread forms having external threads and an internal dove tail relief to match the dove tail retention feature at the root of the composite fastener is held in face-to-face mating engagement with one another so as to establish a mechanical interlock around the root of the composite fastener.

Abstract: An actively-cooled, fiber-reinforced ceramic matrix composite thrust chamber for liquid rocket propulsion systems is designed and produced with internal cooling channels. The monocoque tubular structure consists of an inner wall, which is fully integrated to an outer wall via radial coupling webs. Segmented annular void spaces between the inner wall, outer wall and adjoining radial webs form the internal trapezoidal-shaped cooling channel passages of the tubular heat exchanger. The manufacturing method enables producing any general tubular shell geometry ranging from simple cylindrical heat exchanger tubes to complex converging-diverging, Delaval-type nozzle structures with an annular array of internal cooling channels. The manufacturing method allows for transitioning the tubular shell structure from a two-dimensional circular geometry to a three-dimensional rectangular geometry.

Abstract: A fiber-reinforced ceramic matrix composite material exhibiting increased matrix cracking strength and fracture toughness is produced by sequentially depositing a plurality of 5-500 nanometer-thick layers of a primary ceramic matrix material phase periodically separated by 1-100 nanometer-thick intermediate layers of a secondary matrix material phase onto the reinforcing fibers upon their consolidation. The resultant nanolayered matrix enhances the resistance to the onset of matrix cracking, thus increasing the useful design strength of the ceramic matrix composite material. The nanolayered microstructure of the matrix constituent also provides a unique resistance to matrix crack propagation. Through extensive inter-layer matrix fracture, debonding and slip, internal matrix microcracks are effectively diverted and/or blunted prior to their approach towards the reinforcing fiber, thus increasing the apparent toughness of the matrix constituent.

Abstract: An actively-cooled, fiber-reinforced ceramic matrix composite thrust chamber for liquid rocket propulsion systems is designed and produced with internal cooling channels. The monocoque tubular structure consists of an inner wall, which is fully integrated to an outer wall via radial coupling webs. Segmented annular void spaces between the inner wall, outer wall and adjoining radial webs form the internal trapezoidal-shaped cooling channel passages of the tubular heat exchanger. The manufacturing method enables producing any general tubular shell geometry ranging from simple cylindrical heat exchanger tubes to complex converging-diverging, Delaval-type nozzle structures with an annular array of internal cooling channels. The manufacturing method allows for transitioning the tubular shell structure from a two-dimensional circular geometry to a three-dimensional rectangular geometry.

Abstract: An actively-cooled, fiber-reinforced ceramic matrix composite thrust chamber for liquid rocket propulsion systems is designed and produced with internal cooling channels. The monocoque tubular structure consists of an inner wall, which is fully integrated to an outer wall via radial coupling webs. Segmented annular void spaces between the inner wall, outer wall and adjoining radial webs form the internal trapezoidal-shaped cooling channel passages of the tubular heat exchanger. The manufacturing method enables producing any general tubular shell geometry ranging from simple cylindrical heat exchanger tubes to complex converging-diverging, Delaval-type nozzle structures with an annular array of internal cooling channels. The manufacturing method allows for transitioning the tubular shell structure from a two-dimensional circular geometry to a three-dimensional rectangular geometry.

Abstract: An actively-cooled, fiber-reinforced ceramic matrix composite thrust chamber for liquid rocket propulsion systems is designed and produced with internal cooling channels. The monocoque tubular structure consists of an inner wall, which is fully integrated to an outer wall via radial coupling webs. Segmented annular void spaces between the inner wall, outer wall and adjoining radial webs form the internal trapezoidal-shaped cooling channel passages of the tubular heat exchanger. The manufacturing method enables producing any general tubular shell geometry ranging from simple cylindrical heat exchanger tubes to complex converging-diverging, Delaval-type nozzle structures with an annular array of internal cooling channels. The manufacturing method allows for transitioning the tubular shell structure from a two-dimensional circular geometry to a three-dimensional rectangular geometry.

Abstract: A fiber-reinforced ceramic composite material exhibiting high tensile strength, high fracture toughness and high-temperature oxidation resistance is produced by simultaneously depositing a thin coating layer of refractory metal carbide with fugitive carbon onto the fiber reinforcement prior to the subsequent densification with the ceramic matrix. The energy behind propagating matrix cracks in the resulting composite material are effectively dissipated by crack deflection/branching, fiber debonding and frictional slip through the relatively weak and compliant fiber coating layer. These energy release and arrest mechanisms sufficiently impede the driving force behind unstable crack propagation and render the cracks non-critical, thereby serving to blunt and/or divert propagating matrix cracks at or around the reinforcing fiber.

Abstract: A fiber-reinforced ceramic-matrix composite material exhibiting high tensile strength, high fracture toughness and high-temperature oxidation resistance is produced by alternatively depositing multiple thin layers of ceramic material separated by very thin intermediate layers of fugitive carbon onto the fiber reinforcement prior to the subsequent densification with the ceramic matrix. The energy behind propagating matrix cracks in the resulting composite material are effectively dissipated by the progressive increase in crack deflection/branching and frictional slip through the successive ceramic layers of the multilayer fiber coating system. These energy release and arrest mechanisms sufficiently impede the driving force behind unstable crack propagation and render the cracks non-critical, thereby serving to blunt and/or divert propagating matrix cracks at or around the reinforcing fiber.

Abstract: A fiber reinforced ceramic composite material exhibiting high tensile strength, high fracture toughness and high-temperature oxidation resistance is produced by simultaneously depositing a thin coating layer of refractory metal carbide with fugitive carbon onto the fiber reinforcement prior to the subsequent densification with the ceramic matrix. The energy behind propagating matrix cracks in the resulting composite material are effectively dissipated by crack deflection/branching, fiber debonding and frictional slip through the relatively weak and compliant fiber coating layer. These energy release and arrest mechanisms sufficiently impede the driving force behind unstable crack propagation and render the cracks non-critical, thereby serving to blunt and/or divert propagating matrix cracks at or around the reinforcing fiber.

Abstract: A fiber-reinforced ceramic-matrix composite material exhibiting high tensile strength, high fracture toughness and high-temperature oxidation resistance is produced by alternatively depositing multiple thin layers of ceramic material separated by very thin intermediate layers of fugitive carbon onto the fiber reinforcement prior to the subsequent densification with the ceramic matrix. The energy behind propagating matrix cracks in the resulting composite material are effectively dissipated by the progressive increase in crack deflection/branching and frictional slip through the successive ceramic layers of the multilayer fiber coating system. These energy release and arrest mechanisms sufficiently impede the driving force behind unstable crack propagation and render the cracks non-critical, thereby serving to blunt and/or divert propagating matrix cracks at or around the reinforcing fiber.

Abstract: An aircraft wing pivot structure is described which provides a flexible ring attached to the aircraft wing at flexible wing attachment points by spherical bearings. The ring is attached to the fuselage of the aircraft by pins fixed to the fuselage and fittings partially enclosing sliding shoes formed at the bottom edge of the ring to allow rotation of the wing. The wing is locked in the flight position by retractable pins connected to the primary fuselage attachment fittings. The ring is a unitary stainless structure which accepts wing strains induced by flight loads by flexing at the wing attachment points. The use of a flexible ring eliminates truss-like members mounted on spherical bearings used in conventional designs. The elimination of a substantial number of components by the use of a flexible ring substantially reduces the weight of the aircraft and improves the reliability.