Abstract: Systems and methods for providing localized heat treatment of gas turbine components are provided. In this regard, an exemplary method includes: identifying an area of a gas turbine component to which localized heat treatment is to be performed; positioning a resistance heating element, sized and shaped to transfer heat to the area, adjacent the area; electrically coupling a current-limiting power controller to the resistance heating element; and locally heating the area using the resistance heating element, with power being provided to the resistance heating element via the current-limiting power controller.

Abstract: Gas turbine engine systems and related methods involving multiple gas turbine cores are provided. In this regard, a representative gas turbine engine includes: an inlet; a blade assembly mounted to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assembly, each of the multiple gas turbine cores being independently operative in a first state, in which rotational energy is provided to rotate the blade assembly, and a second state, in which rotational energy is not provided to rotate the blade assembly.

Abstract: Gas turbine engine systems involving hydrostatic face seals with integrated back-up seals are provided. In this regard, a representative seal assembly for a gas turbine engine includes: a stator assembly and a rotor assembly configured to operatively engage each other to form a first seal and a second seal; the first seal being provided by a hydrostatic seal having a seal face and a seal runner; and the second seal being provided by a back-up seal such that responsive to a failure of the first seal, the back-up seal maintains at least a portion of a pressure differential established by the first seal prior to the failure.

Abstract: Gas turbine engine systems involving hydrostatic face seals are provided. In this regard, representative turbine assembly for a gas turbine engine includes a turbine having a hydrostatic seal formed by a seal face and a seal runner.

Abstract: Systems and methods involving engine models are provided. A representative method for updating a gas turbine engine reference model includes: determining whether an absolute difference between a dominant time constant of a sensor and an assumed sensor model time constant is outside a tolerance value; and responsive to determining that the absolute difference is outside the tolerance value, updating the assumed sensor model time constant with the dominant time constant of the sensor.

Abstract: Gas turbine engine systems involving hydrostatic face seals with anti-fouling provisioning are provided. In this regard, a representative turbine assembly for a gas turbine engine comprises: a turbine having a hydrostatic seal; the hydrostatic seal having a seal face, a seal runner, a carrier, and a biasing member; the seal face and the seal runner defining a high-pressure side and a lower-pressure side of the seal; the carrier being operative to position the seal face relative to the seal runner; and the biasing member being located on the lower-pressure side of the seal and being operative to bias the carrier such that interaction of the biasing member and gas pressure across the seal causes the carrier to position the seal face relative to the seal runner.

Abstract: Gas turbine engine systems involving hydrostatic face seals with back-up seals are provided. In this regard, a representative seal assembly for a gas turbine engine includes: a hydrostatic seal having a seal face and a seal runner; and a back-up seal; wherein, in a normal mode of operation of the hydrostatic seal, interaction of the seal face and the seal runner maintains a pressure differential within the gas turbine engine and, in a failure mode of operation of the hydrostatic seal, the back-up seal maintains a pressure differential within the gas turbine engine.

Abstract: Masks and related methods for repairing a gas turbine engine components are provided. A representative method includes using a self-coiling strip of material as a mask during application of a coating to a surface of the component.

Abstract: Methods for repairing a gas turbine engine knife edge seals are provided. A representative method includes: providing a knife edge of a gas turbine component, the component comprising IN-100, the knife edge being degraded; directing a laser beam toward the knife edge; and dispensing IN-100 powder such that the IN-100 powder is melted by the laser beam and is deposited on the knife edge.

Abstract: Systems and methods for defining information provided by aircraft are provided. An exemplary method, involving an aircraft that is operative to provide information in a first report format, includes: communicating information corresponding to a second report format definition to the aircraft, wherein the information corresponding to the second report format definition is received at the aircraft via wireless communication; and receiving, from the aircraft, information complying with the second report format definition.

Abstract: Systems and methods for cooling vane platforms are provided. In this regard, a representative method for cooling a vane platform includes: providing a cooling channel on a platform from which a vane airfoil extends, the cooling channel being defined by a cooling surface and a channel cover, the channel wall being spaced from the cooling surface and located such that the cooling surface is positioned between a gas flow path of the vane and the channel cover; and directing a flow of cooling air through the cooling channel such that heat is extracted from the cooling surface of the platform by the flow of cooling air.

Abstract: Systems and methods involving struts are provided. In this regard, a representative system includes a gas turbine engine defining a gas flow path. The gas turbine engine comprises a strut extending into the gas flow path. The strut has an interior operative to receive pressurized air, an outer surface, and ports communicating between the outer surface and the interior of the strut, the ports are operative to receive the pressurized air from the interior and emit the pressurized air into the gas flow path such that an effective aerodynamic length of the strut along the gas flow path is increased.

Abstract: Methods for performing manual laser deposition are provided. In this regard, a representative method includes: directing a laser beam at a first deposit location of a substrate to re-melt: at least a portion of a first deposit of weld filler, at least a portion of a second deposit of weld filler located at a second deposit location that partially overlaps the first deposit, and at least a portion of the substrate.

Abstract: Systems and methods for providing localized heat treatment of metal components are provided. In this regard, a representative method includes: identifying a portion of a metal component to which localized heat treatment is to be performed; shielding an area in a vicinity of the portion of the metal component; and directing electromagnetic energy in the infrared (IR) spectrum toward the portion of the metal component such that the portion is heated to a desired temperature and such that the area in the vicinity of the portion that is subjected to shielding does not heat to the temperature desired for the heat treatment.

Abstract: Insulated supports are provided. In this regard, representative insulated support includes: a framing member having opposing faces and opposing sidewalls extending between the faces, the framing member further having a cutout positioned along a length thereof; and an insulated plug sized and shaped to be received within the cutout such that the insulated plug, when inserted into the cutout, increases an insulation (R) value of the support.

Abstract: Systems involving feather seals are provided. A representative vane assembly for a gas turbine engine includes: a first mounting platform having a first slot; a first airfoil extending from the first mounting platform; and a feather seal having opposing faces, a first side extending between the faces, and a first tab, the first tab extending outwardly beyond the first side; the first slot being sized and shaped to receive the feather seal including the first tab.

Abstract: Floatwall panel assemblies and related systems are provided. A floatwall panel assembly includes a panel formed of porous ceramic material, the porous ceramic material exhibiting a porosity gradient along at least one of a length, a width and a depth of the panel, the panel lacking a substrate, formed of a material other than porous ceramic material, for supporting the porous ceramic material.

Abstract: Systems and methods involving vanes are provided. In this regard, a representative method for modifying the throat area between vanes of a gas turbine engine includes: directing a gas flow path of the gas turbine engine between a first vane and a second vane, wherein each of the first vane and the second vane has an outer surface and an interior; and emitting pressurized air from outlet ports communicating between the outer surface and the interior of the first vane, wherein the emitted pressurized air from the first vane modifies a throat area between the first vane and the second vane.

Abstract: Systems and methods for monitoring gas turbine engines are provided. In this regard, a representative method includes: monitoring lubrication oil at multiple locations of the gas turbine engine to detect a presence of debris in the oil; determining a characteristic of the debris in the oil; and correlating the characteristic of the debris with the location of detection to determine, while the gas turbine engine is operating, whether the engine is operating within predetermined limits.

Abstract: An representative gas turbine includes: a first annular gas flow path; a second annular gas flow path located radially outwardly from the first gas flow path; in series order, a multi-stage fan, radially inboard portions of a first set of high pressure compressor blades, a second set of high pressure compressor blades, a combustion section and a high pressure turbine, positioned along the first gas flow path; and in series order, the multi-stage fan and radially outboard portions of the first set of high pressure compressor blades, positioned along the second gas flow path such that the inboard portions of the first set of high pressure compressor blades, the second set of high pressure compressor blades, the combustion section and the high pressure turbine are not positioned along the second gas flow path.