Abstract: A turbine engine and method of operating includes an engine core with a compressor, a combustor, and a turbine in axial flow arrangement. A flow path extends through the engine core from the compressor to the turbine to define a flow direction for a working airflow through the engine core.

Abstract: An introduced autonomous aerial vehicle can include multiple cameras for capturing images of a surrounding physical environment that are utilized for motion planning by an autonomous navigation system. In some embodiments, the cameras can be integrated into one or more rotor assemblies that house powered rotors to free up space within the body of the aerial vehicle. In an example embodiment, an aerial vehicle includes multiple upward-facing cameras and multiple downward-facing cameras with overlapping fields of view to enable stereoscopic computer vision in a plurality of directions around the aerial vehicle. Similar camera arrangements can also be implemented in fixed-wing aerial vehicles.

Abstract: A Brayton cycle engine and method for operation. The engine includes an inner wall assembly and an upstream wall assembly each extended from a longitudinal wall into a gas flowpath. An actuator adjusts a depth of the detonation combustion region into the gas flowpath between the inner wall assembly and the upstream wall assembly. The engine flows an oxidizer through the gas flowpath and the inner wall captures a portion of the oxidizer. The engine further adjusts the captured flow of oxidizer via the upstream wall and flows a first flow of fuel to the captured flow of oxidizer to produce rotating detonation gases. The engine flows the detonation gases downstream and to mix with the flow of oxidizer, and flows and burns a second flow of fuel to the detonation gases/oxidizer mixture to produce thrust.

Abstract: A stator blade for a centrifugal compressor having a front portion configured to generate one or more strear-wise vortices in the gas flow around the stator blade in order to avoid and/or delay a detachment of the gas flow from the suction surface of the stator blade especially when the centrifugal compressor is not operating at its operational design speed.

Abstract: A turbine is provided with: a turbine wheel configured to rotate about an axis O1; a turbine housing accommodating the turbine wheel and defining an annular nozzle passage on the outer peripheral side of the turbine wheel; and a plurality of low solidity nozzle vanes 6 arranged in the nozzle passage at an interval in the circumferential direction. Circumferentially adjacent low solidity nozzle vanes 6 are disposed at different radial positions in a connection position of each of the low solidity nozzle vanes 6 with a hub-side wall surface of the hub-side wall surface and a shroud-side wall surface which define the nozzle passage.

Abstract: A method of operating a multi-engine system of a rotorcraft includes, during a cruise flight segment of the rotorcraft, controlling a first engine to provide sufficient power and/or rotor speed demands of the cruise flight segment; and controlling a second engine to by providing a fuel flow to the second engine that is between 70% and 99.5% less than a fuel flow provided to the first engine. A turboshaft engine for a multi-engine system configured to drive a common load is also described.

Abstract: An air guide arrangement, for a ventilation system, has a housing forming a flow channel with a fan arranged in the flow channel to generate an airflow through the housing. A flow guide device is arranged in the flow channels. The flow guide device is axially connected downstream of the fan on the outflow side and directly influences the airflow generated by the fan. The flow guide device has an axis-central through opening delimited by a tubular element extending parallel to the flow direction. Multiple separate flow segments are formed along the tubular element and are evenly distributed in the circumferential direction. The flow segments in the circumferential direction are each separated from one another in terms of flow by flow guide elements extending radially outward from the tubular element.

Abstract: A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

Abstract: A dynamic compressor includes an impeller having impeller vanes disposed around a hub, a shroud, and a diffuser having a shroud surface adjacent to the shroud and a hub surface adjacent to the hub, wherein the diffuser is circumferentially disposed around the impeller. The diffuser includes a plurality of diffuser vanes extending from the hub surface to the shroud surface, each having a vane leading edge and a vane trailing edge. The diffuser includes a centrifugal acceleration stabilizer ring formed in the shroud surface located in a vaneless region defined between an impeller trailing edge and the vane leading edge. The centrifugal acceleration stabilizer ring stabilizes the flow of the fluid by changing the circumferentially-flowing high velocity fluid flow exiting the impeller into a radially-flowing high velocity fluid flow before entering the diffuser, improving the efficiency of the dynamic compressor.

Abstract: A pre-swirler unit can include: a shroud including a bottom surface and a top surface opposite to the bottom surface; a blade disposed on the top surface and including a leading edge and a trailing edge, wherein the blade comprises a concave side surface and a convex side surface that are disposed between the leading edge and the trailing edge, wherein the convex side surface comprises a first plurality of dimples, and wherein the concave side surface comprises a second plurality of dimples.

Abstract: A stator vane segment includes: a plurality of airfoil portions; and one outer band and one inner band joined to a tip portion and a hub portion of each airfoil portion. Each airfoil portion is formed by stacking airfoil profiles in a spanwise direction from the hub portion to the tip portion. Each profile includes a leading edge, a trailing edge, a pressure surface, and a suction surface, in a stacking line connecting respective trailing edges of the profiles at spanwise positions. A portion from the tip portion to a predetermined position is a straight line in a radial direction. A portion from the predetermined position to the hub portion is shifted from the pressure surface toward the suction surface in a circumferential direction from the straight line parallel to the radial direction. An amount of the shift monotonically increases from the predetermined position to the hub portion.

Abstract: A compressor or pump stage is provided. The compressor or pump stage at least comprising a central shaft (8) and one rotor (3), where the axis of rotation of the rotor (3) is the central shaft (8) and where the rotor comprises a number, n, of rows of impellers (5) arranged at an outer perimeter of the rotor with an axial distance between neighbouring rows of impellers (5), where n={2, 3, 4 . . . }.

Abstract: A turbomachine blade including a body that extends mainly in a plane defined by a main axis and a longitudinal direction, which is defined by a lower surface wall, an upper surface wall, a leading edge located at a first longitudinal end of the body and a trailing edge located at a second longitudinal end of the body, wherein the body of the blade includes a plurality of first pipes that extend mainly along the direction of the main axis, for circulation of a gas flow, and a plurality of second pipes that extend mainly along the longitudinal direction, for circulation of a second gas flow.

Abstract: A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

Abstract: A compressor includes a housing, a shaft, and an impeller rotatable relative to the housing by the shaft. The impeller includes a hub, impeller blades, and external blades. The impeller blades extend from a front of the hub. The external blades protrude from a rear surface of the hub or an outer surface of a shroud of the impeller. The external blades are curved along the rear surface or the outer surface. A heat transfer circuit includes a compressor and a working fluid. The compressor includes an impeller having a hub, impeller blades, and external blades.

Abstract: A mixed-flow compressor includes an impeller attached to a shaft and rotatable about a shaft axis. A vaneless diffuser is located axially downstream of the impeller and has a converging portion and a diverging portion. A vaned diffuser is located axially downstream of the vaneless diffuser.

Abstract: A centrifugal compressor includes a casing. An impeller is arranged within the casing. The impeller is rotatable about an axis. A diffuser section is arranged within the casing. The diffuser section is positioned axially downstream from an outlet of the impeller and includes a forward portion fixed relative to the impeller and an aft portion distinct from the forward portion.

Abstract: An exemplary gas turbine engine includes a fan section including a fan rotor and at least one fan blade. A fan pressure ratio across the at least one fan blade is less than 1.45, noninclusive of the pressure across any fan exit guide vane system. The engine further includes a low-pressure compressor having a low-pressure compressor rotor that rotates together with the fan rotor at a common speed in operation, and a geared architecture that drives the low-pressure compressor rotor and the fan rotor. The geared architecture has a gear reduction ratio of greater than 2.5. The engine further includes a high-pressure compressor having a pressure ratio greater than 20, a low-pressure turbine having a pressure ratio greater than 5, and a bypass ratio greater than 10.

Abstract: A compressor includes a housing. An impeller is located within the housing and rotatable about an impeller axis in a first rotational direction. A rotor section is rotatable about the impeller axis in a second rotational direction opposite the first rotational direction.

Abstract: A rechargeable electric fan with a fan portion, a base portion, and a telescoping mast hingedly attached to and connecting the fan portion and the base portion. The fan portion has removable front grille, engageable in a first position in which the grille is retained in the front opening of the housing of the fan portion. The grille is rotatable from the first position to a second position in which the grille is released from the front opening. Slots are disposed on a periphery of the grille, and are configured to engage with corresponding tabs disposed on a periphery of the front opening, to retain the grille in the first position, and to release the grille when it is rotated into the second position. To prevent inadvertent release of the grille, a latch mechanism is configured and disposed to lock the grille in the first position, the latch mechanism comprising a button and a latch configured and disposed to release the lock when the button is depressed.

Abstract: An airfoil of an axial flow machine includes an airfoil part that has a suction surface and a pressure surface, an endwall as a shroud or a platform, and a fillet that is provided along a connecting portion between the airfoil part and the endwall. The fillet includes a first curved surface formed at least on a leading edge side and a trailing edge side in the airfoil part on the suction surface side and a second curved surface formed in a portion other than a portion where the first curved surface is formed. The second curved surface has a smaller curvature radius than the first curved surface.

Abstract: A component for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section extending in a radial direction from a platform, the airfoil section extending in an axial direction between an airfoil leading edge and an airfoil trailing edge, and the airfoil section extending in the circumferential direction between pressure and suction sides. The platform extends in the axial direction between a platform leading edge and a platform trailing edge, and extends in the circumferential direction between a first mate face and a second mate face. The platform has a radially facing surface joined with the airfoil section and has a cold side surface opposed to the radially facing surface. A first thickness is defined between the radially facing surface and the cold side surface adjacent the first mate face, and a second thickness is defined between the radially facing surface and the cold side surface adjacent the second mate face.

Abstract: A compressor of a gas turbine engine includes an impeller having a plurality of impeller blades. The compressor includes a diffuser downstream from the impeller that has a plurality of diffuser blades. Each diffuser blade extends from a hub to a shroud in a spanwise direction, and a leading edge of each diffuser blade is spaced apart from an impeller trailing edge of each of the plurality of impeller blades by a vaneless gap. Each diffuser blade includes a cutback region that extends from proximate the leading edge toward a trailing edge. The cutback region reduces a thickness of each of the diffuser blades such that a throat area defined between adjacent diffuser blades increases in the spanwise direction from the hub to the shroud and the vaneless gap increases in the spanwise direction from the hub to the shroud.

Abstract: A blade of a fan or compressor that reduces loss by enlarging a laminar flow region over a blade surface is provided. The blade is divided into a subsonic region where the relative Mach number of the inlet air flow during rated operation of a turbofan engine is lower than 0.8 and a transonic region where the relative Mach number is equal to or higher than 0.8. A blade surface angle change rate is based on an angle formed by a tangent to the blade surface and the axis of the engine, the leading edge blade surface angle, and the trailing edge blade surface angle at. In each of the subsonic region and the transonic region, values of the blade surface angle change rate on the pressure and suction surfaces are defined at predetermined axial locations along the chord on the pressure and suction surfaces.

Abstract: A turbocharger turbine assembly can include a turbine housing that includes a longitudinal axis, an exhaust inlet, a shroud portion, and an exhaust outlet; and a turbine wheel that includes a hub that includes a rotational axis aligned with the longitudinal axis of the turbine housing, a backdisk and a nose, where the rotational axis defines an axial coordinate (z) in a cylindrical coordinate system; and blades that extend outwardly from the hub, where each of the blades includes a shroud edge, a leading edge, a trailing edge, a pressure side, and a suction side, where the shroud edge includes a minimum axial coordinate position where the shroud edge meets the leading edge, and a maximum axial coordinate position where the shroud edge meets the trailing edge, and where, from the minimum axial coordinate position to the maximum axial coordinate position, the shroud edge has a decreasing radial coordinate position.

Abstract: The present invention provides a steam turbine blade and a steam turbine capable of further improving efficiency. A steam turbine blade includes a proximal end portion T11 which is formed in a blade shape having an intermediate reaction degree, an intermediate portion T12 which is formed in a blade shape having a low reaction degree; and a distal end portion T13 which is formed in a blade shape having a high reaction degree.

Abstract: An apparatus and method of managing negative incidence of an airfoil are provided. The apparatus includes a first row of stationary turbine main vanes and a second row of auxiliary vanes extending radially inwardly from a stationary casing of the turbine proximate the adjacent leading edge of the main vane. The apparatus also includes a flow channel defined between a pressure side of an auxiliary vane of the second row of auxiliary vanes and a suction side of an adjacent main vane of the row of stationary turbine main vanes.

Abstract: A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

Abstract: A centrifugal impeller is disclosed having a non-axisymmetric flowpath surface. The centrifugal compressor may comprise a hub and a plurality of circumferentially spaced vanes. The hub has a flowpath surface and an axis of rotation. The plurality of circumferentially spaced vanes extend from the flowpath surface, with each of the vanes having a pressure-side fillet and a suction-side fillet extending from a leading edge to a trailing edge of the vane. The pressure-side fillet and suction-side fillet intersect the flowpath surface at a runout. The runout of the pressure-side fillet of a first vane is asymmetric to the runout of the suction-side fillet of the first vane.

Abstract: A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

Abstract: A high-Mach engine includes a gas turbine core, an inlet assembly, and a transmission. The inlet assembly including an inlet turbine and the transmission configured to couple the inlet turbine and a core turbine of the gas turbine core cooperate to control air moving through the high-Mach engine.

Abstract: Disclosed is a gas turbine including a housing, a rotor rotatably provided in the housing to transfer a rotary force to a compressor, the compressor receiving the rotary force from the rotor and compressing air, a combustor mixing a fuel with the compressed air supplied from the compressor and igniting the mixture of the fuel and the air to generate combustion gas, and a turbine receiving the rotary force caused by the combustion gas generated by the combustor and rotating the rotor by using the received rotary force.

Abstract: A fan unit that forms part of a gases supply unit used as part of a breathing assistance system for providing heated gases to a user is provided. The fan unit has an impeller surrounded by an upwardly sloped surface to facilitate improved airflow performance under surge conditions. The impeller includes a plurality of blades capped by a lid having a central aperture to provide an inlet to the impeller.

Abstract: An apparatus and method of managing negative incidence of an airfoil are provided. The apparatus includes a row of vane pairs including a first row of main vanes extending radially inwardly from a stationary casing member and spaced circumferentially about a first axial location of the stationary casing member. The apparatus also includes a second row of auxiliary vanes extending radially inwardly from the stationary casing member and spaced circumferentially about a second axial location of the stationary casing member. The apparatus also includes a flow channel defined between a pressure side of each auxiliary vane and a suction side of an adjacent main vane proximate a leading edge of the adjacent main vane.

Abstract: An inlet assembly for a high-mach engine includes a gas turbine core, an inlet turbine, and a core-flow director. When in a closed position, the core-flow director forces air to interact with the inlet turbine before entering the gas turbine core.

Abstract: Turbomachines having one or more flow guiding features designed to increase the performance of the turbomachine (3400, 3700, 4000, 4800). In some examples, flow guiding features are designed and configured to bias a circumferential pressure distribution at a diffuser inlet (2210, 2310, 3410, 4204, 4810, 5208, 808) toward circumferential uniformity, otherwise account for such low-frequency spatial pressure variations, increase the controllability of spatial flow field variations, or modifying flow field variations, etc.

Abstract: A compressor rotor airfoil in a gas turbine engine is presented. Opposed pressure and suction sides are joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. A leading edge dihedral angle is defined at a point on the leading edge between a tangent to the airfoil and a vertical. The leading edge dihedral angle has a span-wise distribution. The distribution has at least one inflection point. A method of reducing a rub angle between a compressor rotor blade and a casing surrounding the blade is also presented.

Abstract: An electric supercharger includes an electric motor, a compressor wheel rotated by the electric motor, and a compressor housing. The compressor housing includes a first passage through which air is introduced, an introducing port connecting to an EGR device, a second passage through which at least one of air and EGR gas is flowed, a bypass passage through which at least one of air and EGR gas is flowed, and a bypass valve to open and close the bypass passage. When the bypass valve is opened, air introduced from the first passage and EGR gas introduced from the introducing port are flowed through the bypass passage to the internal combustion engine. When the bypass valve is closed, air introduced from the first passage and EGR gas introduced from the introducing port are compressed in the compressor wheel and flowed through the second passage to the internal combustion engine.

Abstract: The invention concerns a radial turbo compressor (TCO) including at least one impeller (IP) at least one casing (CS), the impeller (IP) is rotatable about an axis (X), the casing (CS) comprises an inlet (IL) upstream of the impeller (IP), the inlet (IL) comprising an inlet flange (IF) to be mounted to a process gas pipe (PGP), the casing (CS) comprises an outlet (OL) down-stream of the impeller (IP) comprising an outlet flange (OF), the casing (CS) comprises an outlet volute (VL) extending about the axis (X) downstream of the impeller (IP) and upstream of the outlet (OL), and the radial turbo compressor (TCO) comprises a drive unit (DRU) driving the impeller (IP) and being mounted to the casing (CS).

Abstract: A turbine engine part or set of parts including at least first and second obstacles each having a leading edge and a trailing edge, and a platform from which the obstacles extend; wherein the platform has, between the pressure side of the first obstacle and the suction side of the second obstacle a non-axisymmetric surface defining at least one fin with a substantially triangular cross-section, each fin being associated with a leading position and a trailing position on the surface between which the fin extends, such that the leading position is upstream of each of the leading edges; the trailing position is downstream of each of the leading edges.

Abstract: A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

Abstract: A gas turbine engine includes a compressor section and a compressor case with a low pressure compressor (LPC) and a high pressure compressor (HPC). The HPC is aft of the LPC. The compressor case defines a centerline axis. The compressor section also includes a rotor disk defined between the compressor case and the centerline axis. A plurality of stages are defined radially inward relative to the compressor case. The plurality of stages include at least one tandem blade stage. The tandem blade stage includes a plurality of blade pairs. Each blade pair is circumferentially spaced apart from the other blade pairs, and is operatively connected to the rotor disk. Each blade pair includes a forward blade and an aft blade. The aft blade is configured to further condition air flow with respect to the forward blade without an intervening stator vane stage shrouded cavity therebetween.

Abstract: A compressor airfoil of a turbine engine having a geared architecture includes pressure and suction sides that extend in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between a tangential stacking offset and span position that defines a curve that is non-linear.

Abstract: An airfoil for a gas turbine engine includes an airfoil with pressure and suction sides that are joined at leading and trailing edges. The airfoil extends a span from a support to an end in a radial direction. 0% span and 100% span positions respectively correspond to the airfoil at the support and at the end. The leading and trailing edges are spaced apart from one another an axial chord in an axial direction. A cross-section of the airfoil at a span location has a diameter tangent to the pressure and suction sides. The diameter corresponds to the largest circle fitting within the cross-section. A ratio of the diameter to the axial chord is at least 0.4 between 50% and 95% span location.

Abstract: A method for designing a centrifugal pump and a mixed flow pump having a specific speed of 150-1200 comprises a design specification determination step for a turbo pump including an impeller, a specific speed determination step for the impeller, a design variable determination step for the impeller, and a three dimensional shape deriving step for the impeller. According to the present invention, the three dimensional shape of the impeller can be simply designed by converting the design variables related to the specific speed into a function and by putting optimized design variables into a database.

Abstract: A compressor airfoil of a turbine engine having a geared architecture includes pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position. The airfoil has a relationship between leading edge sweep angle and span position defined by a curve in which the leading edge sweep angle is positive at 0% span and crosses to a negative leading edge sweep angle at a span position less than 80% span. A negative sweep angle is in the forward direction, and a positive sweep angle is in the rearward direction.

Abstract: A rotary airfoil in a gas turbine engine is provided. The airfoil includes opposed pressure and suction sides joined together at chordally opposite leading and trailing edges. The pressure and suction sides extend in a span direction from a root to a tip. An axial component of a center of gravity of a cross-section taken chordally toward the tip of the airfoil being upstream relative to an axial component of a center of gravity of a cross-section taken chordally toward the root of the airfoil. A method for forming such blade is also presented.

Abstract: The present application provides a turbine nozzle including an airfoil shape. The airfoil shape may have a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table I. The Cartesian coordinate values of X, Y and Z are non-dimensional values from 0% to 100% convertible to dimensional distances in inches by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil in inches. The X and Y values, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The airfoil profile sections at Z distances may be joined smoothly with one another to form a complete airfoil shape.

Abstract: An airfoil for a turbine engine includes pressure and suction sides that extend in a radial direction from a 0% span position at an inner flow path location to a 100% span position at an airfoil tip. The airfoil has a relationship between a stacking offset and a span position that includes at least one positive and negative slope. The positive slope leans aftward and the negative slope leans forward relative to an engine axis. The positive slope crosses an initial axial stacking offset corresponding to the 0% span position at a zero-crossing position. A first axial stacking offset X1 is provided from the zero-crossing position to a negative-most value on the curve. A second axial stacking offset X2 is provided from the zero-crossing position to a positive-most value on the curve. A ratio of the second to first axial stacking offset X2/X1 is less than 2.0.

Abstract: The present invention relates to a blade row group arrangeable in a main flow path of a fluid-flow machine and including N adjacent member blade rows firmly arranged relative to one another in both the meridional direction (m) and the circumferential direction (u), with the number N of the member blade rows being greater than/equal to 2 and (i) designating the running index with values between 1 and N. Here, a front member blade row with front blades (i) having a leading edge and a trailing edge as well as a rear member blade row with rear blades (i+1) having a leading edge and a trailing edge are provided.