Abstract: An assembly for a gas turbine engine includes a combustor and a vane assembly disposed downstream thereof. A portion of an outer platform of the vane assembly defines an axial sliding joint connection with the combustor, and includes a plurality of depressions located in an outer circumferential surface thereof opposite the combustor. The depressions are disposed in regions of expected higher thermal growth about the circumference of the outer platform such that thermal growth of the entire outer platform is substantially uniform circumferentially therearound.

Abstract: A method for operating a pulse detonation engine, wherein the method includes channeling air flow from a pulse detonation combustor into a flow mixer having an inlet portion, an outlet portion, and a body portion extending therebetween. The method also includes channeling ambient air past the flow mixer and mixing the air flow discharged from the pulse detonation combustor with the ambient air flow such that a combined flow is generated from the flow mixer that has less flow variations than the air flow discharged from the pulse detonation combustor.

Abstract: An injection assembly for use with a combustor is provided. The injection assembly includes an effusion plate that has a plurality of plate openings and a plate sleeve having a sidewall portion that includes a forward edge. The forward edge is coupled to the effusion plate such that the effusion plate is oriented obliquely with respect to a centerline extending through the combustor. The injection assembly also includes a plurality of ring extensions where each of the ring extensions is coupled to one of the plurality of plate openings. Each ring extension extends rearwardly into the plate sleeve.

Abstract: A combustor heat shield comprises a heat shield member defining at least one opening for receiving a fuel nozzle. A louver is received in the opening. The louver has a flow diverting portion extending radially outwardly from the opening for directing air along the hot side of the heat shield member.

Abstract: A screw shaft turbine compressor comprising (i) a compressor section, (ii) a turbine section, (iii) a combustion section coupling to the compressor section and the turbine section, and (iv) a grooved shaft. The grooved shaft in one embodiment extends from a portion of the compressor section, through the combustion section, and to a portion of the turbine section.

Abstract: One or more Helmholtz-type resonators (270) is/are provided at the junction (260) of a combustor (220) and a combustion chamber (240) of a gas turbine engine (100). In one embodiment, adjacent Helmholtz-type resonators (290, 291, 292), which may be separated by respective baffles (285), have different volumes that help provide for damping different undesired combustion-generated acoustic pressure waves. In some embodiments, a structural member (435) may be provided between adjacent Helmholtz-type resonators (425, 426, 427, 428) at the junction. At least one of the plurality of Helmholtz-type resonators comprises one or more inlet openings (480), and one or more exit openings (482). Embodiments (370, 425-429) are described in which Helmholtz-type resonators provided at the junction are enlarged in size using various approaches.

Abstract: A method of operating an exhaust nozzle assembly includes translating a translatable structure between a plurality of operational positions to open and close a flow diverting port extending in an exhaust duct. The longitudinal position of a throat constriction varies with translation of the translatable structure. Additionally, the cross-sectional area of the exhaust duct at the throat constriction may vary with translation of the translatable structure. An exemplary exhaust nozzle is operable in-flight for providing at least partial reverse thrust to at least partly control the velocity of an aircraft. An exemplary exhaust nozzle is operable on the ground to spoil ground idle thrust.

Abstract: A power plant for burning a fuel in a low pressure combustion chamber to produce electrical power. A first compressor supplies compressed air through a first heat exchanger to add heat to the compressed air. The heated compressed air is passed through a first turbine to drive a first electric generator. The first turbine outlet is passed through a second heat exchanger in series with the first heat exchanger to further heat the compressed air. The compressed air is then passed through a second turbine to drive a second electric generator and produce electric power. The outlet from the second turbine is passed through a first combustor to produce the hot gas flow through the second heat exchanger. The outlet from the second heat exchanger is passed through a second combustor before passing through the first heat exchanger. The outlet from the first heat exchanger is passed through a heat recovery steam generator to generate steam to drive another turbine and another generator.

Abstract: A burner (27) of a gas turbine engine (10) includes a cylindrical basket (60) comprising an air flow reversal region (86). The flow reversal region ends at an air inlet plane (84) of the basket. The burner also includes a flow conditioner (90) disposed in the flow reversal region transecting an air flow (80) flowing non-uniformly through the flow reversal region, the flow conditioner being effective to mitigate variation of the air flow entering the basket across the inlet plane.

Abstract: A sensor assembly for a gas turbine engine includes a telemetry module mounted at a rotor bearing compartment for sensing engine operational parameters and a cooling system for cooling the telemetry module separate from a rotor bearing lubricant flow.