Abstract: A transport refrigeration unit for use with a transport trailer includes a structural framework mountable to a wall of the transport trailer, a condenser heat exchanger unit mounted to the structural framework, an evaporator housing separated from the condenser heat exchanger by a distance, and at least one condenser fan assembly positioned aft of the condenser heat exchanger unit and forward of the evaporator housing. The at least one condenser fan assembly includes a fan rotor defining an inlet end. The fan rotor is rotatable about a fan axis. A casing defines a central opening and the fan rotor is arranged within the central opening. The casing includes a plurality of openings spaced about a periphery of the casing. A portion of an airflow moving through the at least one condenser fan assembly is expelled from the at least one condenser fan assembly radially via the plurality of openings.

Abstract: A gearbox assembly for a gas turbofan engine includes a sun gear rotatable about an axis and a plurality of intermediate gears driven by the sun gear. A baffle disposed between at least two of the plurality of intermediate gears includes a first gap distance within a first gap portion and a second gap distance within a second gap portion. The first gap portion is disposed between the baffle and one of the intermediate gears away from the meshed interface with the sun gear and the second gap portion is disposed near the interface with the sun gear. The first gap distance within the first gap portion is different than the second gap distance within the second gap portion to define a desired lubricant flow path.

Abstract: A rotor for a turbofan booster section associated with a fan section of a gas turbine engine includes a rotor blade having an airfoil extending from a root to a tip and having a leading edge and a trailing edge. The airfoil has a plurality of chord lines spaced apart in a spanwise direction. Each chord line of the plurality of chords lines is defined between the leading edge and the trailing edge and has a normalized chord value. From the hub, the normalized chord value decreases to a minimum value between about 20% to about 90% span and increases from the minimum value to the tip. The rotor includes a rotor disk coupled to the rotor blade configured to be coupled to the shaft or the fan to rotate with the shaft or the fan, respectively, at the same speed as the shaft and fan.

Abstract: A ventilator unit has a ventilator housing and at least one ventilator accommodated in the ventilator housing. A first outlet guide apparatus is associated with the ventilator and is arranged in the ventilator housing, wherein the first outlet guide apparatus has first outlet guide vanes distributed circumferentially. A second outlet guide apparatus is correlated with the first outlet guide apparatus and arranged in a flow path of an air stream sucked in by the ventilator. The second outlet guide apparatus is connected to a protective screen. The second outlet guide apparatus has second outlet guide vanes. The second outlet guide apparatus increases the throw of the ventilator unit.

Abstract: This disclosure provides a seal strip assembly for a sealing strip slot to reduce air leakage between gas turbine nozzle segments. The seal strip assembly includes a bimetal element and a backing plate. The bimetal element includes a first layer with a first thermal coefficient of thermal expansion and a second layer with a second coefficient of thermal expansion. The second coefficient of thermal expansion is lower than the first coefficient of thermal expansion allowing the bimetal element to expand with an increase in surrounding temperature.

Abstract: A gas turbine engine for an aircraft including: engine core including a turbine; and fan including a plurality of fan blades extending radially from a hub, each fan blade having a leading and trailing edge. Turbine includes a lowest pressure turbine stage having a row of rotor blades each extending radially and having a leading and trailing edge. A fan-turbine radius difference is measured as radial distance between: a point on a circle swept by a radially outer tip of the trailing edge of each of the rotor blades of the lowest pressure stage of the turbine; and a point on a circle swept by a radially outer tip of the leading edge of each of fan blades; and a fan speed to fan-turbine radius ratio defined as: the ? ? maximum ? ? take ? - ? off ? ? rotational ? ? speed ? ? of ? ? the ? ? fan fan ? - ? turbine ? ? radius ? ? difference ? ? ( 120 ) is in a range between 0.8 rpm/mm to 5 rpm/mm.

Abstract: A turbomachine including a housing having an inlet end opposite and outlet end along a longitudinal axis of the housing, a shaft assembly provided within the housing, the shaft assembly extending from the inlet end to the outlet end, a rotor having at least one rotating impeller extending radially outward from the shaft assembly, and a return channel vane hub extending radially outward from the shaft assembly, the return channel vane hub includes at least one return channel vane extend therefrom, the at least one return channel vane comprising a body having a leading edge and a trailing edge, the leading edge is twisted and extended past an outer edge of the return channel vane hub, and the trailing edge is bowed outwardly.

Abstract: A protective grille for a fan has a web structure which comprises radial webs spaced apart in a circumferential direction and coaxial circumferential webs spaced apart in a radial direction. The protective grille has at least a radial outer region and a central region around a central axis of the protective grille. An envelope surface spanned by the radial outer region and the central region is of convexly curved form in the radial outer region and is of flat form, parallel to a radial plane of the protective grille, in the central region.

Abstract: An example gas turbine engine includes a propulsor assembly including a fan and a fan drive turbine. A weight of the propulsor assembly is less than about 40% of a total weight of the gas turbine engine.

Abstract: A gas turbine engine (10) includes: a compressor system comprising a low pressure compressor (15) and a high pressure compressor (16) coupled to low pressure and high pressure shafts, respectively (23, 24); an inner core casing (34) provided radially inwardly of compressor blades (42), and an outer core casing provided outwardly of compressor blades, the inner core casing and outer core casing defining a core working gas flow path (B) therebetween; a fan (13) coupled to the low pressure shaft via a gearbox (14); wherein the outer core casing comprises a first outer core casing (48) and a second outer core casing (50) spaced radially outwardly from the first outer core casing, and wherein at an axial plane (E) of an inlet to the high pressure compressor, the second outer core casing has an inner radius at least 1.4 times the inner radius of the first outer core casing.

Abstract: A gearbox assembly for a gas turbofan engine includes a sun gear rotatable about an axis and a plurality of intermediate gears driven by the sun gear. A baffle disposed between at least two of the plurality of intermediate gears includes a first gap distance within a first gap portion and a second gap distance within a second gap portion. The first gap portion is disposed between the baffle and one of the intermediate gears away from the meshed interface with the sun gear and the second gap portion is disposed near the interface with the sun gear. The first gap distance within the first gap portion is different than the second gap distance within the second gap portion to define a desired lubricant flow path.

Abstract: A blade and method for producing the blade for a gas turbine engine are described herein. The blade may include a composite airfoil. The airfoil may comprise a ceramic material, and a distal end. A tip may extend from the distal end of the airfoil. The tip of the airfoil may comprise a substantially porous structure and may comprise infiltrated material extending from an airfoil preform to a tip preform to join the airfoil preform and the tip preform.

Abstract: A booster assembly for a gas turbine engine having a first rotor assembly comprising a low pressure turbine drivingly connected to a fan via a first shaft and a second rotor assembly comprising a second turbine drivingly connected to a high pressure compressor via a second shaft. The booster assembly comprises a further compressor arranged to be disposed about said common axis between the fan and high-pressure compressor in a direction of flow and a gearing having first and second input rotors and an output rotor. The first input rotor is arranged to be driven by the first rotor assembly and the second input rotor is arranged to be driven by the second rotor assembly such that the output rotor drives the further compressor in dependence upon the difference in rotational speed between the first and second rotor assemblies.

Abstract: A gas turbine engine comprises a fan includes a plurality of fan blades rotatable about an axis. A compressor section includes at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between about 1000° F. and about 1500° F. A combustor is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.

Abstract: An example gas turbine engine includes a propulsor assembly including at least a fan module and a fan drive turbine module; a gas generator assembly including at least a compressor section, a combustor in fluid communication with the compressor section, and a turbine in fluid communication with the combustor; and a geared architecture driven by the fan drive turbine module for rotating a fan of the fan module. A weight of the fan module and the fan drive turbine module is less than about 40% of a total weight of a gas turbine engine.

Abstract: There is presented a casing arrangement for a gas turbine engine. The casing arrangement includes a plurality of tip treatment bars extending between a pair of spaced-apart annular supports, with each tip treatment bar being elongate and supported at each end by a respective support such that each end of the bar is received within an opening formed in the respective support. At least the ends of each tip treatment bar are tapered in transverse cross-section so as to have a wedge-shaped profile, and the openings in the supports have a complementary tapered shape for receipt and retention of the bar ends therein.

Abstract: An air blower includes a casing having an air inlet portion and an air outlet portion, a fan, and a fan guard. The fan includes a boss portion and a plurality of blades. The fan guard includes a guide portion having a tubular outer shape, which protrudes toward the fan. A center in a distal end shape, which is defined by a contour line of a distal end portion of the guide portion, matches with a rotation axis of the boss portion. A center in a root shape, which is defined by a contour line of a root portion of the guide portion, is shifted with respect to the rotation axis of the boss portion.

Abstract: A fan includes a frame and an impeller. The frame includes a main body, a plurality of guiding elements and a motor base. The main body has an airflow inlet and an airflow outlet. The guiding elements are disposed in the main body and located at the airflow outlet, wherein each of the guiding elements includes an inclined part and an axial extended part, and the inclined par meets the axial extended part at an angle. The motor base connects to the main body. The impeller is disposed on the motor base for providing an airflow, wherein the guiding elements guide the airflow away from the fan.

Abstract: A seal end attachment is provided and includes a vessel through which a working fluid flows, the vessel being formed to define a recess with a mating surface therein, a seal contacting the mating surface and a pressing member being more responsive to a high temperature condition associated with the flow of the working fluid than the vessel and being disposed within the recess to press the seal against the mating surface responsive to the condition being present.

Abstract: A vane assembly for a gas turbine engine includes a first platform, a second platform, and an airfoil that extends radially across an annulus between the first platform and the second platform. The airfoil is centered relative to a centerline axis of the second platform and is offset relative to a centerline axis of the first platform.

Abstract: A gas turbine engine includes a compressor, a combustor section, and a turbine. The turbine includes an integrated case/stator segment that is comprised of a ceramic matrix composite material.

Abstract: A driveshaft vessel assembly for a gas turbine engine is disclosed. The driveshaft vessel assembly comprises a driveshaft vessel and an outlet guide vane. The driveshaft vessel includes a forward face, an aft face, and opposing side faces and is configured to house at least a portion of a radial driveshaft. The outlet guide vane includes a leading edge and a trailing edge. The length of the trailing edge is substantially equal to a length of the forward face of the driveshaft vessel such that the trailing edge of the outlet guide vane is faired into the forward face of the driveshaft vessel.

Abstract: Various embodiments include steam turbine static nozzles having transitional interfaces. In one embodiment a steam turbine static nozzle blade includes: an airfoil; an inner sidewall integral with a first side of the airfoil; and an outer sidewall integral with a second side of the airfoil; the inner sidewall and the outer sidewall each including: a first side having one of: an arcuate concave surface extending substantially an entire length of the sidewall, or an arcuate convex surface extending substantially the entire length of the sidewall; and a second side opposing the first side and having an angled interface extending substantially the entire length of the sidewall.

Abstract: A shell of a stator for a module of an aircraft turbo-engine including a plurality of through openings each configured to house a stator blade, each opening forming a skeleton extending between a first end configured to house the trailing edge of the blade and a second end configured to house the leading edge of the blade. At least one of the openings is associated with a mechanical load transfer slit passing through the shell, and arranged facing and at a distance from the first end of the opening along the direction of the skeleton.

Abstract: A Ceramic Matrix Composites (CMC) airfoil for a gas turbine engine includes at least one CMC ply which defines a suction side, an outer platform, a pressure side and an inner platform with a continuous “I”-shaped fiber geometry.

Abstract: A gas turbine engine includes inner and outer shrouds forming an annular gas path, and a plurality of struts connecting the inner shroud to the outer shroud. Airfoil shaped shields surround the struts, and each of the shields include a main body having an upstream leading edge defining a chordal axis extending in a downstream axial direction from the leading edge toward a downstream end of the shield. A trailing edge flap is located at the downstream end of each shield, the trailing edge flap including first and second span-wise portions. The first span-wise portion is oriented to direct flow at an angle relative to the chordal axis of the main body and the second span-wise portion is oriented to direct flow in a direction that is at a different angle than the angle of first span-wise portion.

Abstract: A turbine of a turbomachine is provided and includes opposing endwalls defining a pathway into which a fluid flow is receivable to flow through the pathway; and a nozzle stage at which adjacent nozzles extend across the pathway between the opposing endwalls to aerodynamically interact with the fluid flow. The adjacent nozzles are configured to define a throat distribution exhibiting endwall throat decambering and pitchline throat overcambering.

Abstract: A CMC turbine stator blade 1 includes a blade 2 which is formed of a ceramic matrix composite material, and a band 11 which is formed of a ceramic matrix composite material or a metallic material and supports the blade 2. The blade 2 includes a first fitting portion 5 to which the band 11 is fitted at the outside thereof. The band 11 includes a second fitting portion 12 to which the first fitting portion 5 is fitted at the inside thereof. The CMC turbine stator blade 1 includes fixing means 6 which sandwiches the second fitting portion 12 from both sides in the height direction of the blade so that the blade 1 and the band 11 are fixed.

Abstract: A rigid mechanical self-centering assembly for a component mounted relative to an opening. Variable radii cylindrical components are accommodated by virtue of variable-length, rigid, structural stays, and flexible mounts with curved faces for flush abuttal to said component, such as a motor or pump. A high efficiency assembly for electrical components is achieved by use of nonferrous materials.

Abstract: A low leakage, integral fixed vane system for a gas turbine engine compression system which can accommodate the latest 3-D aerodynamic airfoil geometries and provide endwall ovalization control. A unitized construction wherein the vanes are integral with the case wall in the compression system.

Abstract: An inner shroud assembly includes an aft core segment mountable to a forward core segment to support a multiple of vanes for rotational movement relative thereto. A shroud backing plate segment engageable with the aft core segment and at least one fastener which passes through the shroud backing plate, the aft core segment and the forward core segment.

Abstract: A turbine exhaust case comprises inner and outer annular shrouds defining therebetween an annular hot gaspath. A circumferential array of exhaust struts extends across the gaspath. The exhaust struts are mounted at one radial end thereof on a flexible strut mounting structure. The flexible strut mounting structure is radially deflectable relative to the outer and inner shrouds to accommodate thermal expansion of the exhaust struts during engine operation.

Abstract: The present invention relates to a turbomachine rectifier comprising a plurality of stator vanes (5) connecting an inner collar (4) to an outer collar (3), each of said vanes (5) comprising a blade (11) and a blade-head platform (10), comprising an intermediate piece (6) arranged between the inner collar (4) and the outer collar (3) and fastened to the outer collar (3), said intermediate piece (6) comprising openings (14) for the passage of the vane blades (11) and said vane platforms (10) resting, on one side, on the outer collar (3) and, on the other side, on the intermediate piece (6).

Abstract: A vane cluster includes a split damped outer shroud and an inner shroud spaced from the split damped outer shroud with a multiple of stator vane airfoils that extend between the split damped outer shroud and the inner shroud.

Abstract: A shroud for a turbine and a method of assembling is provided. The turbine includes a housing, a rotatable shaft and a bucket extending outward therefrom. The shroud includes an alignment member which is coupled to the housing, wherein the alignment member includes a first end, a second end and a body extending between the first and second ends. The second end includes an arcuate portion to facilitate fluid flow downstream from the bucket. The shroud further includes a seal coupled to the body to facilitate sealing a gap defined between the bucket and the body.

Abstract: A turbocharger is disclosed for use with an engine system. The turbocharger may have a housing at least partially defining a compressor shroud and a turbine shroud. The turbine shroud may form a first volute and a second volute, each having an inlet configured to receive exhaust from an exhaust manifold of the engine in a tangential direction, and an axial channel disposed downstream of the inlet. The turbocharger may also have a turbine wheel disposed within the turbine shroud and configured to receive exhaust from the axial channels of the first and second volutes, a compressor wheel disposed within the compressor shroud, and a shaft connecting the turbine wheel to the compressor wheel. The turbocharger may further have a nozzle ring in fluid communication with the axial channels of the first and second volutes at a location upstream of the turbine wheel.

Abstract: A CMC turbine stator blade 1 includes a blade 2 of a ceramic matrix composite material formed or a metallic material, and a band 5 formed of a ceramic matrix composite material and supporting the blade 2. The blade 2 includes a first fitting portion 3 to which the band 5 is fitted outside. The band 5 includes a second fitting portion 7 to which the first fitting portion 3 is fitted inside. A flexible wire 8 is disposed between the fitting portions 3, 7 to fix the blade 2 and the band 5. A first groove 3a in the outer peripheral surface of the first fitting portion 3 extends in the front-to-rear direction of the blade 2. A second groove 7a in the inner peripheral surface of the second fitting portion 7a extends along the first groove 3a. The wire 8 is disposed between the grooves 3a,7a.

Abstract: A vane structure of a gas turbine engine includes a plurality of vanes extending radially between outer and inner shrouds. At least the outer shroud is formed substantially from a single-piece annular skin of sheet metal. A plurality of reinforcing plates are placed against and are affixed to an outer surface of the skin of the outer shroud in respective joining locations where a radial outer end of each of the respective vanes joins the skin.

Abstract: A vane structure for a gas turbine engine according includes a multiple of CMC airfoil sections integrated between a CMC outer ring and a CMC inner ring.

Abstract: A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of radially extending fan blades integral with the hub to form an integrally bladed rotor. Each fan blade defines a leading edge. A hub radius (RHUB) is the radius of the leading edge at the hub relative to a centerline of the fan. A tip radius (RTIP) is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan. The ratio of the hub radius to the tip radius (RHUB/RTIP) is at least less than 0.29. In a particular embodiment, this ratio is between 0.25 and 0.29. In another particular embodiment, this ratio is less than or equal to 0.25.

Abstract: A method is provided for adjusting a position of a blade ring relative to a rotor in a gas turbine engine. An outer casing surrounds the blade ring and the blade ring surrounds the rotor. The method comprises: determining an amount of vertical movement needed to reposition the blade ring relative to the rotor so that the blade ring is at a desired position relative to the rotor; providing at least one torque pin assembly comprising a torque pin and a variable thickness defining structure; determining a change in the thickness of the variable thickness defining structure so as to effect the necessary vertical movement of the blade ring; changing the thickness of the variable thickness defining structure; and coupling the at least one torque pin assembly to the outer casing such that at least one torque pin engages the blade ring. A gas turbine engine is provided having structure for adjusting a position of a blade ring relative to a rotor.

Abstract: A flowpath assembly for a gas turbine engine includes a plurality of flowpath insert ducts arranged in a cascade configuration. Each flowpath duct includes a radially inward wall, a radially outward wall, a first side wall, and a second side wall. A flowpath volume is defined between the inward, outward, first side and second side walls. The first side wall of a given one of the plurality of flowpath insert ducts is positioned adjacent to the second sidewall of an adjacent one of the plurality of flowpath insert ducts in the cascade configuration.

Abstract: A gas turbine engine compressor has a compressor case comprising spaced apart inner and outer walls. An axial rotor is positioned within the outer wall. A bearing structure supports the axial rotor for rotation. A plurality of inlet guide vanes are coupled to the outer wall of the compressor case and radially extend inwardly, wherein each of at least a sub-set of said inlet guide vanes comprises a radial bore. Nested tie rods are received within a respective one of the inlet guide vane radial bores. Each tie rod comprises an outward end attached to the compressor case outer wall and an inward end attached to the compressor case inner wall.

Abstract: A turbine vane ring has radially outer and inner annular shrouds defining therebetween an annular gaspath. Circumferentially spaced-apart airfoil vanes extend radially across the gaspath between the outer and the inner shrouds. The radially outer shroud has a circumferentially continuous cylindrical wall extending axially from a leading edge to a trailing edge. A set of circumferentially distributed stress relieving slots is defined in the leading edge of the cylindrical wall at locations adjacent to the leading edge of at least some of said airfoil vanes. The stress relieving slots extend radially through the cylindrical wall from the radially inner surface to the opposed radially outer surface thereof.

Abstract: In accordance with an exemplary embodiment, a turbine stator component includes a first endwall; a second endwall; a first stator airfoil coupled between the first and second endwalls; and a second stator airfoil adjacent to the first airfoil and coupled between the first and second endwalls. The first stator airfoil has first crystallographic primary and secondary orientations. The second stator airfoil has second crystallographic primary and secondary orientations, the first crystallographic primary and secondary orientations being different from the second crystallographic primary and secondary orientations.

Abstract: A turbine diaphragm has radially inner and outer hollow diaphragm rings (22, 24) comprising box structures that include the radially inner and outer platform portions (261, 262) of static blade units (26) as part of their structures. The box structures are strengthened by the incorporation of ribs or struts (28, 30) which extend radially and axially between the walls of the box structures.

Abstract: A variable vane assembly for a gas turbine engine is disclosed herein. The variable vane assembly includes a vane operable to pivot about a pivot axis. The vane includes a flap portion positionable in a fluid stream for controlling a flow of fluid. The vane also includes a stem portion for mounting the flap portion. The vane also includes a button portion positioned between the flap portion and the stem portion along the pivot axis. The button portion is wider than the stem portion in at least one plane containing the pivot axis. The variable vane assembly also includes a bushing encircling the stem portion and abutting the button portion. The variable vane assembly also includes a labyrinth seal positioned between the bushing and the button portion.

Abstract: An assembling of a guide vane assembly of a turbomachine turbine, formed by an annular row of stationary flow-stabilizing vanes. An outer edge of the guide vane assembly is axially bearing on a line of an external casing of the turbine and an inner edge of the guide vane assembly is in axial sliding connection with the line of the internal casing of the turbine, the axial sliding connection allowing the inner edge to be free along the motor axis with the axial stop being achieved by the axial bearing of the outer edge of the vane.

Abstract: A vane assembly for a gas turbine engine is disclosed and includes a plurality of individual vane segments forming an annular ring. Each of the plurality of vane segments is formed of a ceramic matrix composite including at least one airfoil extending between an inner platform and an outer platform. An outer wrap is disposed about an outer periphery of the annular ring and over an interface between outer platforms of adjacent ones of the plurality of vane segments.

Abstract: A turbomachine component includes a turbine stator nozzle member having an airfoil core shape. The airfoil core shape includes a nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in TABLE 1, and wherein X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z in inches, the profile sections at the Z distances being joined smoothly with one another to form a complete airfoil core shape.