Abstract: A housing structure for a turbo-engine, especially for a gas turbine or an aircraft engine, includes an outer housing wall and an inner housing wall, the inner and outer housing walls annularly enclosing a flow channel for the turbo-engine and being spaced apart radially from the flow channel. At least one heat shield is arranged between the inner and outer housing walls, and a bar or a fixture projects at least somewhat radially from the inner housing wall. The bar or the fixture has on at least one side a broadening element that includes a sealing face against which the heat shield is positioned in a sealing manner.

Abstract: An augmentor assembly for a gas turbine engine includes an alignment assembly disposed along an axis having a base and a plate. The base includes a plurality of grooves. The plate includes a plurality of flanges in mating engagement to the plurality of grooves. A first component and a second component are attached to the alignment assembly. The mating engagement of the plurality of grooves and the plurality of flanges is arranged to provide incremental adjustment to define a position of the first component relative to the second component.

Abstract: A structural support system is provided in a can annular gas turbine engine having an arrangement including a plurality of integrated exit pieces (IEPs) forming an annular chamber for delivering gases from a plurality of combustors to a first row of turbine blades. A bracket structure is connected between an IEP and an inner support structure on the engine. The bracket structure includes an axial bracket member attached to an IEP and extending axially in a forward direction. A transverse bracket member has an end attached to the inner support structure and extends circumferentially to a connection with a forward end of the axial bracket member. The transverse bracket member provides a fixed radial position for the forward end of the axial bracket member and is flexible in the axial direction to permit axial movement of the axial bracket member.

Abstract: A fuel staging system is described which comprises a staging valve 28 operable to divide a metered fuel supply between a pilot delivery line 18 and a mains delivery line 16, and a recirculation valve 54 connected to the mains delivery line 16 and operable to permit a recirculation fuel flow to be supplied to the mains delivery line 16 when the staging valve 28 occupies a position in which substantially no fuel is delivered to the mains delivery line 16 through the staging valve 28.

Abstract: A pump sharing valve for a gas turbine engine fuel system includes a housing having multiple ports. A sleeve is arranged in the housing and includes an inlet window and first, second and third windows axially spaced apart from one another. A spool is slidably received in the sleeve and includes a diameter and first and second cylindrical portions. The first cylindrical portion selectively connects the inlet window to the second and third windows. The second cylindrical portion selectively connects the inlet window to the first window. The spool is movable between first and second positions relative to the first, second and third windows to regulate flow. The first and second cylindrical portions respectively includes first and second widths. A first ratio corresponds to the first width to the spool diameter and a second ratio corresponds to the second width to the spool diameters. The first ratio is between 1.83 and 1.93 and the second ratio is between 1.22 and 1.32.

Abstract: A buffer air cooler system for gas turbine engines disposed in a bypass duct of the engine, includes a housing for containing the buffer air cooler therein and an inlet portion attached to the housing. In one embodiment, the inlet portion has a double-skin configuration in at least one region of a top, bottom and sides of the inlet portion.

Abstract: A gas turbine combustion chamber has a housing (1), a flame tube (2) received therein, and a perforated plate (3) which at least partially surrounds the latter and which is fastened at both of its end faces (3.1, 3.2) to the housing.

Abstract: A propulsion system of an aircraft includes an inner fixed structure (IFS) and the outer sleeve that may be separately coupled to the pylon. For instance, the inner fixed structure and the outer sleeve may move independently with respect to each other. A single latch of the latching system disclosed herein, may be configured to latch all 4 panels, (e.g. both halves of the IFS and both doors of the outer sleeve) together and/or in a closed and retained position.

Abstract: A heat shield device for a yoke suspending an engine from an aircraft is provided. The device includes panels able to be fixed to the yoke, at least one of the panels extending radially in the direction of the axis of the engine, and a component configured to be fixed to the engine and including a groove shaped in such a way that the lower end of the panel can be inserted into it.

Abstract: A system includes a gas turbine system that has a turbine, a combustor, and a compressor having an air intake. The air intake is configured to supply an air flow to the compressor. The system also includes a fluid injection system that has a fluid wash system configured to inject a wash fluid into the gas turbine system via the air intake. The fluid wash system includes an injector configured to inject the wash fluid into the gas turbine system and an anti-icing system configured to inject an anti-icing fluid into the gas turbine system via the air intake. The anti-icing system is configured to inject the anti-icing fluid into the gas turbine system via the injector.

Abstract: An apparatus for supporting an aft portion of a transition duct in a gas turbine engine includes a transition aft frame that engages with an annular shaped stator component disposed in a turbine section of the gas turbine engine. The transition aft frame includes radially inner and outer panels and circumferentially spaced first and second side panels connecting the inner and outer panels. A forward face of the stator component includes first and second connection points circumferentially spaced apart. The transition aft frame includes first and second attachment structures that respectively engage with the first and second connection points when the transition duct is aligned axially with the stator component. The first and second attachment structures are spaced apart in a manner effective to transfer moment load from the first and second attachment structures to the first and second side panels respectively.

Abstract: The present invention relates to a system and method of LO airflow modulation for use with a DTSA engine. A DTSA engine is positioned within an aircraft fuselage and the second DTSA turbine fan includes it own dedicated decoupled air inlet duct that is formed co-centrically about the housing. An airflow modulator member is positioned in duct to form the duct wall when the modulator is in a closed position. The Modulator member is pivotally connected to the duct wall, and is movable by an actuator to a second open position that allows airflow to escape the third stream duct, and provide airflow to both the DTSA fan blades as well as engine bay for cooling. The method of the present invention provides airflow modulation to an aircraft employing a DTSA engine, said method including the receiving of airflow from a supersonic aircraft intake; sensing a number of parameters, including, but not limited to aircraft speed, temperature, engine load and/or altitude.

Abstract: A wall structure for restricting a hot-gas path in a gas turbine or in a combustion chamber including a first wall segment having a first face, and a second wall segment having a second face. Each wall segment disposed adjacent to each other so as to define an inside of the wall segments and an outside of the wall segments. The first and second faces being disposed so as to form a gap leading from the inside to the outside. Each of the faces having a receiving groove. The wall structure further including a sealing element having a thickness and disposed in the receiving grooves so as to bridge the gap. The first face having a recess and the second wall segment having a projection that extends from the second face and protrudes into the recess in the first face. The projection has a tapered cross section.

Abstract: The present disclosure relates to a beam, in particular for a cascade thrust reverser, including a skin made of a composite material and defining a closed section, wherein said beam is filled with a core material.

Abstract: The present disclosure relates to methods and systems for reducing silica recession of silicon-containing ceramics or silicon-containing ceramic composites, particularly those exposed to a combustion gas or to combustion gas environments, including those exposed to high temperature combustion gas environments. The methods and systems involve silicon-doping of compressed air and/or removal of moisture from compressed air prior to co-mingling the treated compressed air with the combustion gas to which the silicon-containing ceramics or silicon-containing ceramic composites are exposed.

Abstract: The present invention relates to a low signature jet engine outlet comprising movable border segments forming at least one circular arc and being movable radially for controlling the form and area of the outlet, at least one stationary border element for defining the form and area of the outlet, and movable coupling means for connecting the at least one circular arc to the at least one stationary border element. The coupling means are movable along a connection part of the at least one stationary border element, and the connection part of the at least one stationary border element is substantially flat and extending in a substantially radial direction of the at least one circular arc.

Abstract: A wake manipulating structure for a turbine system includes a combustor liner defining a combustor chamber. Also included is an airflow path located along an outer surface of the combustor liner. Further included is a wake generating component disposed in the airflow path and proximate the combustor liner, wherein the wake generating component generates a wake region located downstream of the wake generating component. Yet further included is a venturi structure or section disposed in the airflow path configured to reduce the wake region.

Abstract: A nacelle for a turbofan engine includes a variable area fan nozzle (VAFN) and a proximity sensor that is attached to a nacelle forward portion to sense the presence of the VAFN when stowed. The proximity sensor is not attached to the VAFN, thereby enabling the use of a proximity sensor such as linear variable displacement transformer having a relatively compact operational stroke.

Abstract: A cooling system is provided for cooling a turbine of a gas turbine engine. The system has first and second flow paths for guiding cooling air received from the compressor of the engine. The routes of both flow paths bypass the combustor of the engine. The system also has a preswirler for receiving the cooling air at the ends of the two flow paths, swirling the cooling air tangentially to the engine axis, and delivering the swirled cooling air to a rotor of the turbine. The first flow path is routed through a heat exchanger which cools the cooling air guided by the first flow path relative to the cooling air guided by the second flow path.

Abstract: A turbomachine combustor portion includes a combustion chamber. A center injection nozzle is arranged within the combustion chamber and includes a center nozzle inlet and a center nozzle outlet. An outer premixed injection nozzle is positioned radially outward of the center injection nozzle and includes an outer nozzle inlet and an outer nozzle outlet that is arranged upstream of the center nozzle outlet. A late lean injector is positioned downstream of the center nozzle and the outer premixed nozzle. The combustor portion includes a first combustion zone arranged downstream of the outer nozzle outlet, a second combustion zone arranged downstream of the center nozzle outlet, and a third combustion zone arranged further downstream of the center nozzle outlet. The center injection nozzle, outer premixed injection nozzle, and late lean injector are selectively operated to establish a combustion flame front in the first, second, and third combustion zones.