Abstract: A guide vane within an annular inlet duct of a gas-turbine engine provides for generating swirl within an annular inlet duct so as to provide for reducing the rate of deceleration of the inlet air flow within the annular inlet duct while providing for diffusion of the meridional component of velocity thereof.

Abstract: A vacuum pump comprises: a pump rotor portion configured such that multiple rotor blade stages including stacking portions formed in a ring shape and turbine blades radially formed on an outer peripheral side of the stacking portions are stacked on each other; a rotor shaft to which inner peripheral surfaces of the stacking portions of the multiple rotor blade stages stacked on each other are fixed; and a communication path allowing communication between a clearance space between the rotor shaft and each stacking portion and a pump exhaust path in which the turbine blades are arranged and discharging gas in the clearance space through the pump exhaust path.

Abstract: An airfoil profile for a turbine nozzle of a gas turbine is provided. The turbine nozzle may include an airfoil portion having an uncoated nominal profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, wherein the X, Y, and Z coordinates are distances in inches measured in a Cartesian coordinate system, the corresponding X and Y coordinates, when connected by a smooth continuous arc, define one of a plurality of airfoil profile sections at each Z distance, and the plurality of airfoil profile sections, when joined together by smooth continuous arcs, define an airfoil shape.

Abstract: An axial flow fan device 1 includes a first axial flow fan 10 and a second axial flow fan 20. The first axial flow fan 10 of a first casing 14 includes first engagement portions 152. Two sets of first engagement portions 152 are provided at positions which protrude outward in the direction of the axis x from a first base portion 16, and which are symmetrical with respect to the axis x, each first engagement portion 152 having a stepped portion in the radial direction. The second axial flow fan 20 of a second casing 24 includes second engagement portions 252. Two sets of second engagement portions 252 are provided on a second peripheral wall 241 at positions which are symmetrical with respect to the axis x, each second engagement portion 252 having a stepped portion in the radial direction.

Abstract: A fan cover is configured to be mounted on a housing of a fan module. The fan cover includes a frame having an opening formed therein, a central hub positioned within the opening of the frame, and a plurality of spiral-shaped air guidance members that extend from the central hub to the frame. Gaps between the spiral-shaped air guidance members of the plurality of spiral-shaped air guidance members enable air to flow from the fan module through the fan cover. Each spiral-shaped air guidance member is configured to extend perpendicularly from the central hub and curve towards the frame at an angle with respect to the opening of the frame. Each spiral-shaped air guidance member further has a plurality of openings formed therein to facilitate air flow.

Abstract: A rotor disk assembly for a gas turbine engine. A multiple of blade attachments each received within a slot in a rotor disk. The rotor blade attachment manufactured of a metal alloy and a rotor blade retained between each two of the multiple of blade attachments, the rotor blade manufactured of non-metal alloy.

Abstract: A turbomachinery apparatus includes: a turbine, including: a turbine component defining an arcuate flowpath surface; an array of axial-flow turbine airfoils extending from the flowpath surface, the turbine airfoils defining spaces therebetween; and a plurality of fences extending from the flowpath surface, in the spaces between the turbine airfoils, each fence having opposed concave and convex sides extending between a leading edge and a trailing edge, wherein the fences have a nonzero camber and a constant thickness, are axially located near the leading edges of adjacent turbine airfoils, and wherein at least one of a chord dimension of the fences and a span dimension of the fences is less than the corresponding dimension of the turbine airfoils.

Abstract: In a vacuum pump, the proportion of hole portions at a distal end (outer peripheral side/outer peripheral portion) of a stator disc is locally increased. More specifically, hole portions of substantially the same size are provided in parallel from the inner peripheral side to the outer peripheral side of the stator disc, and the number of hole portions provided in the outermost row is set to be larger than the number of hole portions provided in an inner row. Hole portions of substantially the same size are provided in parallel from the inner peripheral side to the outer peripheral side of the stator disc, and some of hole portions provided in the outermost row are integrated to be a single hole portion. T shaped hole portions are provided in the stator disc. L shaped hole portions are provided in the stator disc.

Abstract: Circumferentially-split diffuser assemblies utilized within compression systems, such as centrifugal and mixed-flow compression systems employed within gas turbine engines, are provided. In embodiments, the diffuser assembly includes flow passages, which extend through the diffuser assembly and which include diffuser flow passage sections. Diffuser airfoils are interspersed with the diffuser flow passage sections. The diffuser airfoils include inboard and outboard airfoil segments distributed around a diffuser assembly centerline. The inboard and outboard airfoil segments are contained in and, thus, defined by inner and outer annular diffuser structures, respectively. The outer annular diffuser structure circumscribes the inner annular diffuser structure.

Abstract: A turbine and a turbine-generator device for use in electricity generation. The turbine has a universal design and so may be relatively easily modified for use in connection with generators having a rated power output in the range of 50 KW to 5 MW. Such modifications are achieved, in part, through use of a modular turbine cartridge built up of discrete rotor and stator plates sized for the desired application with turbine brush seals chosen to accommodate radial rotor movements from the supported generator. The cartridge may be installed and removed from the turbine relatively easily for maintenance or rebuilding. The rotor housing is designed to be relatively easily machined to dimensions that meet desired operating parameters.

Abstract: A serial axial flow fan in which an end portion of a first axial flow fan on an exhaust side, and an end portion of a second axial flow fan on an intake side are connected to each other. At least either of a plurality of first blades of a first impeller of the first axial flow fan, and a plurality of second blades of a second impeller of the second axial flow fan are provided with auxiliary blade portions.

Abstract: A compressor for a turbine engine includes an annular row of blades including a first blade with an extrados surface and a second blade with an intrados surface; and a connecting surface joining the intrados surface to the extrados surface, the connecting surface having a bulge of volume VB and a recess of volume VC, the bulge and recess forming a contouring 3D. The volume of the bulge VB is greater than four times the volume VC of the recess, the connecting surface includes a flat or axisymmetric band circumferentially separating the bulge from the recess, and the bottom of the recess being formed against the intrados surface.

Abstract: A propulsor for a gas turbine engine may comprise a case including a duct disposed along an axis to define a flow path, a rotor including a row of propulsor blades extending in a generally radial direction outwardly from a hub, the hub rotatable about the axis such that the propulsor blades deliver airflow into the flow path, and a row of guide vanes situated in the flow path. A first guide vane may comprise a radially inner length (LI), a radially outer length (LO), and a midspan length (LM). The first guide vane may have a first dimensional relationship defined as LO/LM. The first guide vane may have a second dimensional relationship defined as LI/LM. The first dimensional relationship and/or the second dimensional relationship may be greater than 1.05.

Abstract: A gas turbine engine has a fan tip air angle and/or a fan blade tip air angle in a defined range to achieve improved over all performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of fan tip air angle and/or a fan blade tip air angle may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

Abstract: A compressor wheel (1) of a radial compressor (2) of an exhaust-gas turbocharger (3), having a hub (8); and having two blade rows (9, 10) which are arranged adjacent to one another on the hub (8), a first blade row (9) forming the leading blade row, and the adjacent, second blade row (10) forming the trailing blade row, wherein the number of blades (11) of the first blade row (9) is equal to the number of blades (12) of the second blade row (10).

Abstract: A axial flow fan includes an impeller having a plurality of blades, a motor arranged to rotate the impeller, and a casing accommodating the impeller and the motor. The casing includes a cylindrical housing, a motor base configured to hold the motor, and a plurality of spokes connecting the housing with the motor base. The plurality of spokes include a spoke having a straight shape and a spoke having a curved shape that is convex toward a rotational direction of the impeller. The axial flow fan further includes a lead wire for supplying electric power to the motor. When a radial length of the spoke having the straight shape is defined as L and an amount of the curvature of the spoke having the curved shape is defined as X, a displacement of the curved shape (X/L) is less than 0.2.

Abstract: A stator airfoil of a gas turbine engine according to one embodiment includes at least two winglets projecting transversely from opposed sides of a stator airfoil, respectively. Each winglet includes a leading edge axially and outwardly extending from one of the opposed sides of the stator airfoil and a trailing edge extending from the stator airfoil to join with the leading edge of the winglet. The winglets therefore provide a wide trailing edge thereof to aerodynamically manage boundary layers and micro-shocks of a working fluid.

Abstract: A fan exit guide vane assembly for a gas turbine engine includes a plurality of guide vanes having a pressure side and a suction side extending between a leading edge and a trailing edge. The vanes further include a span extending between a root and tip with a stagger angle defined as an angle between a longitudinal axis parallel to an engine axis of rotation and a line connecting the leading edge and the trailing edge that is less than about 15°.

Abstract: A series fan frame body structure made of different materials includes at least two fans, which are serially connected with each other. The fan frames of the two fans are made of different plastic materials. Accordingly, the excited frequency of the fan is prevented from being close to the natural frequency of the fan frames. Therefore, the vibration frequencies of the fan frames are different from each other so that the co-vibration is reduced to lower the noise. By means of the series fan frame body structure made of different materials, the problem of vibration of the fan is effectively improved.

Abstract: Described is a gas turbine diffuser (300) comprising a strut (302) with a leading edge (304) extending between a first wall portion (308a) and a second wall portion (308b), wherein a first edge portion (304a) of the leading edge (304) is inclined towards a diffuser section outlet (326), i.e. in flow direction (328) of an exhaust stream, with regard to a normal direction (319a) perpendicular to the first wall portion (308a) at a first leading end point (320a) at which the leading edge (304) meets the first wall portion (308). Hence the leading edge (304) is partially inclined towards a diffuser section outlet (326).

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 stator blade diaphragm ring includes a plurality of stator blades each having a root and a radial pin receiver. The diaphragm ring further includes a first ring having a plurality of slots, wherein each slot is engaged by the root of one of the plurality of stator blades and a second ring having a plurality of radial through passageways, wherein each radial through passageway is aligned with the radial pin receiver of one of the plurality of stator blades. The diaphragm also includes a plurality of pins, wherein each pin extends from one of the radial through passageways into the radial pin receiver of one of the plurality of stator blades.

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: 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 includes an impeller and a frame. The frame is used for accommodating the impeller. The frame includes a plurality of static blade groups. Each of the static blade groups has a plurality of static blades. Moreover, at least one first static blade of a first static blade group and at least one first static blade of a second static blade group are symmetric with respect to a central axis of the frame.

Abstract: A rotor blade is disclosed herein. The rotor blade comprises a nominal surface profile substantially in accordance with Cartesian coordinates X, Y and Z as set forth in TABLE 1. Wherein X and Y are distances in millimeters which, when connected by smooth, continuing arcs, define airfoil profile sections at each distance Z in millimeters. The airfoil profile sections at the Z distances being joined smoothly with one another to form a complete airfoil shape.

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 1. 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 are distances in inches which, 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: Compressor stators (22) for gas turbine engines (10) are disclosed. An exemplary compressor stator (22) comprises a circumferential array of stator vanes (28) and a shroud (24, 26) for supporting the stator vanes (28). The shroud (24, 26) has a circumferentially extending inner endwall (30, 32) exposed to an annular gas path (20) of the compressor (14). The endwall (30, 32) has a circumferentially uniform axial cross-sectional profile (42). The axial cross-sectional profile (42) comprises at least one deviation (48, 50) from a nominal axial cross-sectional profile (44) defining an overall shape of the annular gas path (20). The at least one deviation comprises a concave deviation (48) and a convex deviation (50).

Abstract: A method of modifying a steam turbine to change from a first maximum thermal power of the steam generator to a second maximum thermal power of the steam generator. The method can include replacing, in the high-pressure module, at least one set of fixed blades sized for the first maximum thermal power by at least one set of fixed blades sized for the second maximum thermal power. The at least one set of moving blades are sized to operate at the first and second maximum thermal powers. The rotor and the at least one set of moving blades of the high-pressure module remain unchanged on changing from the first maximum thermal power to the second maximum thermal power.

Abstract: A gas turbine compression system (1) comprising a gas channel (5), a low pressure compressor section (8) and a high pressure compressor section (9) for compression of the gas in the channel and a compressor structure (14) arranged between the low pressure compressor section (8) and the high pressure compressor section (9). The compressor structure (14) is designed to conduct a gas flow in the gas channel and comprises a plurality of radial struts (15,16,21,24,25) for transmission of load, wherein at least one of the struts (15,16,21,24,25) is hollow for housing service components. The compressor structure (14) is arranged directly downstream of a last rotor (10) in the low pressure compressor section (8) and designed for substantially turning a swirling gas flow from the rotor (10) by a plurality of the struts (15,16,21,24,25) having a cambered shape.

Abstract: When cold and in the non-coated state, the aerodynamic profile is substantially identical to a nominal profile determined by the Cartesian coordinates X,Y,Z? given in Table 1, in which the coordinate Z? is the quotient D/H where D is the distance of the point under consideration from a first reference plane P0 situated at the base of the nominal profile, and H is the height of the profile measured from the first reference plane to a second reference plane P1. The measurements D and H are taken radially relative to the axis of the turbine, while the X coordinate is measured in the axial direction of the turbine.

Abstract: A turbine bucket is provided including a bucket airfoil having an airfoil shape, the bucket airfoil having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in Table 1 wherein 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, and wherein X and Y are distances in inches which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z, the airfoil profile sections at Z distances being joined smoothly with one another to form a complete airfoil shape.

Abstract: A gas turbine that can effectively cool an axial cavity between a stator blade and a rotor blade and a blade-end surface is provided. In a gas turbine in which a blade-count difference between stator-blade segments of a second-stage stator blade and a third-stage stator blade is set to zero, a relative position between the second-stage stator blade and the third-stage stator blade in the circumferential direction is set by clocking so that a wake flow produced by the stator-blade segments of the second-stage stator blade is guided to the leading edges of the stator-blade segments of the third-stage stator blade within a range of 0% to 15% of the length in a span direction from a hub side of the stator-blade segments of the third-stage stator blade.

Abstract: A ring diffuser for an axial turbomachine is provided. The ring diffuser includes an outer wall and an inner wall coaxial thereto, between which a diffuser channel extends in a ring shape along an axial extension from an inflow-side end diverging to an outflow-side end, wherein the inner wall and the outer wall each include a wall surface bounding the diffuser channel. In order to specify a diffuser channel that is tailored to an inflow that is uneven along the circumference and that facilitates a particularly efficient conversion of kinetic energy into static energy, a rotationally asymmetric wall surface bounding the diffuser channel on the inlet side is provided.

Abstract: An arrangement with a ring diffuser for an axial turbomachine, for example for a gas turbine, is provided. The ring diffuser includes an outer wall and with an inner wall coaxial thereto, between which a diffuser duct extends in ring form along an axial extent so as to diverge from an inflow-side end to an outflow-side end, the inner wall and the outer wall in each case including a wall surface delimiting the diffuser duct. In order to specify a diffuser duct which is adapted to an inflow uneven along the circumference and by means of which a specially efficient conversion of kinetic energy into static energy is possible, the wall surface which delimits the diffuser duct on the inlet side is rotationally asymmetrical.

Abstract: A fan assembly includes an axial flow fan between an inlet stator and an outlet stator. The inlet stator has inlet stator blades which extend outwardly from an inner ring. Each inlet stator blade has a downstream edge which has a tangent which is oriented at a first variable angle with respect to a plane which is perpendicular to the fan axis. The first angle increases with increasing distance from the inner support ring. The outlet stator has a plurality of outlet stator blades which extend outwardly from a second inner ring. Each outlet stator blade has an upstream edge which has a tangent which is oriented at a second variable angle with respect to a plane which is perpendicular to the fan axis. The second angle decreases with increasing distance from the inner ring.

Abstract: A turbine apparatus includes: A first nozzle comprising an array of first vanes each including a concave pressure side, a convex suction side, and leading and trailing edges; A rotor downstream from the first nozzle comprising a plurality of blades carried by a rotatable disk; and a second nozzle disposed downstream from the rotor comprising an array of second vanes each including a concave pressure side, a convex suction side, and leading and trailing edges; wherein the first and second vanes of the first and second nozzles are circumferentially clocked relative to each other such that, in a predetermined operating condition, wakes discharged from the first vanes are aligned in a circumferential direction with the leading edges of the second vanes, wherein a stacking axis of the first vanes is nonlinear. An inner band of the first nozzle is contoured in a non-axisymmetric shape.

Abstract: A stator blade carrier, particularly for a gas turbine, is provided. The stator blade carrier includes a number of axial segments and is intended to attain a particularly high degree of operational reliability and long service life. To this end, two neighboring axial segments are connected to a number of tie rods, each enclosed by a support tube, wherein a spherical disk is arranged on at least one end of the respective support tube and mounted in a conical socket supported on the respective axial segment.

Abstract: A tandem blade design for an axial turbomachine, comprising a front blade and a rear blade disposed with an offset thereto in the circumferential direction and in the axial direction. The rear blade is profiled and positioned with respect to the front blade such that it raises the speed level at the trailing edge of the front blade in a predetermined working range in interaction with the front blade.

Abstract: A main flow path of a fluid flow machine includes N adjacent member blade rows firmly arranged relative to each other in both a meridional direction m and a circumferential direction u. A number of the member blade rows, N, is greater than/equal to 2 and (i) designates a running index with values between 1 and N. A trailing edge HK (i) of a blade of the member blade row (i) is spaced from a leading edge VK(i+1) of a blade of the adjacent, downstream member blade row (i+1) by a meridional edge distance D in a meridional plane established by axial direction x and radial direction r. A value of D along a height of the main flow path increases towards the main flow path confinement at least along a part of the area between the main flow path center and the main flow path confinement.

Abstract: A mechanical joint for a gas turbine engine includes:(a) an annular first component having an annular, radially-extending first flange; (b) an annular second component having an annular, radially-extending second flange abutting the first flange; (c) a plurality of generally radially-extending radial channels passing through at least one of the first and second flanges; (d) a plurality of generally axially-extending channels extending through the first flange and communicating with respective ones of the radial channels; and (e) a plurality of fasteners clamping the first and second flanges together.

Abstract: A turbine airfoil assembly for installation in a gas turbine engine. The airfoil assembly includes an endwall and an airfoil extending radially outwardly from the endwall. The airfoil includes pressure and suction sidewalls defining chordally spaced apart leading and trailing edges of the airfoil. An airfoil mean line is defined located centrally between the pressure and suction sidewalls. An angle between the mean line and a line parallel to the engine axis at the leading and trailing edges defines gas flow entry angles, ?, and exit angles, ?. Airfoil inlet and exit angles are substantially in accordance with pairs of inlet angle values, ?, and exit angle values, ?, set forth in one of Tables 1, 3, 5 and 7.

Abstract: A fan module and server equipment are provided that can achieve a balance between increased airflow and noise reduction when an axial flow fan is mounted in the server equipment. The fan module for taking in and discharging air includes a stator located on an upstream side with respect to airflow and an axial flow fan located on the downstream side. When the fan module is viewed from the rotational-axial direction of the axial flow fan, if a leading edge of a rotor vane constituting part of the axial flow fan passes a trailing edge of a stator vane constituting part of the rotor, a skew is formed in which the leading edge of the rotor vane constantly intersects the leading edge of the rotor vane at a single point.

Abstract: To reduce secondary flow loss and to improved turbine efficiency, a section located radially outward of a border section 28 of a stationary blade 21 is bent in the rotational direction of a rotor. Thus, even if combustion gas leaks from a tip clearance between an end wall of a casing and a tip portion of a rotor blade, and a stagnation line 35 near a tip portion 22 is situated in the side of a back surface 24, because a section located radially outward of the border section 28 is bent in the rotational direction of the rotor, the stagnation line 35 is also situated toward the rotational direction of the rotor. Therefore, the stagnation lines 35 formed at various heights in the heightwise direction of the stationary blade 21 are generally aligned in the rotational direction of the rotor. Thus, fluctuation of pressure distribution in the heightwise direction of the stationary blade 21, of the combustion gas flowing into the stationary blade 21 can be reduced.

Abstract: An assembly of airfoils in a turbine engine, the assembly comprising: at least three successive axially stacked rows of airfoils in one of a compressor and a turbine: a first airfoil row, a second airfoil row, and a third airfoil row; wherein the first airfoil row and the third airfoil row each comprise one of a row of rotor blades and a row of stator blades, and the second airfoil row comprises the other; and wherein at least a majority of the mid-channel points of the airfoils in the third airfoil row are positioned circumferentially within +/?25% pitch of the third airfoil row with respect to the location at which a wake flow during a selected operating condition from the first airfoil row is determined to enter the third airfoil row.

Abstract: An inlet guide vane is installed at a suction port into which a fluid is sucked by the rotation of an impeller so as to be rotatable about an axis, and is comprised of a rod-like shaft supported by an inner peripheral surface of the suction port so as to be rotatable about the axis, and a plate-like vane body connected to the shaft and provided so as to protrude from the inner peripheral surface of the suction port. The vane body includes a tapered portion formed such that both side surfaces approach each other toward an outer edge of the vane body in its width direction and an outer edge thereof at a tip in the direction of the axis, and a parallel portion arranged on the axis and formed such that both the side surfaces are parallel to each other along the direction of the axis.

Abstract: An article of manufacture having a nominal profile substantially in accordance with Cartesian coordinate values of X, Y and Z set forth in TABLE A. 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 can be joined smoothly with one another to form a complete airfoil shape.

Abstract: A steam turbine is provided having a rotor which rotates around a machine axis, and a stator which concentrically encompasses the rotor with clearance, between which an annular passage is formed which is exposed to throughflow by steam in the axial direction and in which a multiplicity of rotor blades, which are fastened on the rotor, and fixed stator blades are arranged in a plurality of stages, wherein the stator blades of the last stage have a sweep with a sweep angle which changes in sign over the relative blade height.

Abstract: A two-stage high pressure turbine includes a second stage blade having an airfoil with a profile substantially in accordance with at least an intermediate portion of the Cartesian coordinate values of X, Y and Z set forth in Table 2. The X and Y values are distances, which when smoothly connected by an appropriate continuing curve, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form a complete airfoil shape.

Abstract: A turbomachine includes a housing, and a plurality of turbomachine nozzles. At least one of the plurality of turbomachine nozzles includes a body having a first surface and an opposing second surface that extend between a first end and a second end. At least one of the first and second surfaces includes a profile having an angle of between about 3° and about 8° between a first line passing through the first end and a first position located between about 28% and about 38% from the first end and a second line passing through the first position and a second position located between about 53% and about 63% from the first end, and an angle of between about 10° and about 20° between the second line and a third line passing through the second position and a third position located between about 74% and about 84% from the first end.