Abstract: An exemplary gas turbine engine includes a turbine section operative to impart rotational energy to a compressor section. The turbine section includes at least a low-pressure turbine and a high-pressure turbine, and a number of stages in the low pressure turbine is from three to five.

Abstract: A steam turbine 10 according to an embodiment includes: rotor blade cascades each made up at a turbine rotor 22; an inner casing 21 where the turbine rotor 22 is provided to penetrate; an outer casing 20 surrounding the inner casing 21; stationary blade cascades each made up at an inner side of the inner casing 21; and an annular diffuser 60 provided at a downstream side of a final turbine stage, formed by a steam guide 40 and a bearing cone 50, and discharging steam toward outside in a radial direction. An enlarged inclination angle ?1 of an inner surface 70 of a diaphragm outer ring 26a at the final turbine stage relative to a turbine rotor axial direction is an enlarged inclination angle ?2 of an inner surface 80 at an inlet of the steam guide 40 relative to the turbine rotor axial direction or more.

Abstract: A pump for a fluid which can be blood has a stator housing and a rotor hub with leading and trailing portions and an intermediate portion disposed therebetween. At least one impeller blade at the leading portion drives circumferential and axial components of a flow into a pump annulus or intermediate pathway portion. At least one stator blade extends radially inward from the stator housing within the intermediate pathway portion and is configured to reduce a circumferential component of the flow.

Abstract: An interstage seal is provided for installation between forward and aft stage discs of a turbine section of a gas turbine engine. The seal has a forward annular portion, which extends forward to contribute to an air seal at the outer periphery of the forward stage disc, and an aft annular portion, which extends rearward to contribute to an air seal at the outer periphery of the aft stage disc. The seal has a peripheral slot into which static vane inner rails are loadable. The seal includes a first component, which provides the forward annular portion, and a second component, which provides the aft annular portion. The second component is separably connectable to the first component such that the first and second components, when connected, cooperate to form the peripheral slot. Loading the static vane inner rails into the slot prevents the second component separating from the first component.

Abstract: A turbofan engine is disclosed and includes a fan and a compressor in communication with the fan section, a combustor, a turbine and a speed reduction mechanism coupled to the fan and rotatable by the turbine. The turbine includes a first turbine section that includes three or more stages and a second turbine section that includes at least two stages. A ratio of airfoils in the first turbine section to a bypass area is less than about 170.

Abstract: A turbofan engine has an engine case and a gaspath through the engine case. A fan has a circumferential array of fan blades. The engine further has a compressor, a combustor, a gas generating turbine, and a low pressure turbine section. A speed reduction mechanism couples the low pressure turbine section to the fan. A bypass area ratio is greater than about 6.0. The low pressure turbine section airfoil count to bypass area ratio is below about 170.

Abstract: A turbine section of a gas turbine engine is provided. The turbine section is annular about a longitudinal axis and includes first turbine with a first inlet and a first outlet; a second turbine with a second inlet and a second outlet; an inter-turbine duct extending from the first outlet to the second inlet and configured to direct an air flow from the first turbine to the second turbine; and a first guide vane disposed within the inter-turbine duct.

Abstract: A vacuum pump comprises an exhaust portion having a plurality of rotating blade portions arranged in multiple stages, each of the rotating blade portions having a plurality of rotor blades, and a plurality of stationary blade portions arranged between the rotating blade portions, in which outer circumferential rims are supported via spacers, each of the stationary blade portions having a plurality of stator blades. In at least one stationary blade portion among the plurality of stationary blade portions, a blade height on the inner circumferential side of the plurality of stator blades provided in the stationary blade portion is set to be smaller than a blade height on the outer circumferential side, and the stationary blade portion is supported by the spacers in such a manner that the inner circumferential side of the stationary blade portion is floated up toward an intake port side.

Abstract: An aerofoil, such as a vane, for an axial flow machine, the aerofoil having first and second ends and opposing pressure and suction surfaces extending between said first and second ends in a span-wise direction. A first portion of the span of the aerofoil towards the first end the aerofoil has a location of greatest camber which is closer to a trailing edge than a leading edge of the aerofoil. In a further portion of the aerofoil span the camber is either uniform between the leading and trailing edges or else a location of greatest camber is closer to the leading edge than the trailing edge. In the first portion of the span of the aerofoil the pressure and/or suction surface may turn to a greater extent towards the trailing edge than towards the leading edge of the aerofoil.

Abstract: A counter-rotating axial flow fan with reduced noise at the target operating point achieved without modifying a front impeller, a rear impeller, or a middle stationary portion is provided. An annular rib including a projecting surface for generating turbulent flow is formed on an inner wall portion of a casing at a position off from the middle stationary portion to a side of the rear impeller, the projecting surface extending radially inwardly of the inner wall portion and extending continuously in the circumferential direction of the inner wall portion. A fluid striking the projecting surface for generating turbulent flow is partially disturbed to form a turbulent flow before entering an area in which the rear impeller is provided. The turbulent flow suppresses flow separation of a fluid flowing along the surfaces of rear blades of the rear impeller from the surfaces of the rear blades.

Abstract: A mid-turbine frame for a gas turbine engine includes a vane assembly supported within a case. The vane assembly directs airflow between a first turbine and a second turbine. A support rod extends through the vane assembly and is secured to the case by a retaining nut. The retaining nut extends through a guide attached to the vane assembly for aligning the vane assembly relative to the case portion.

Abstract: In a turbomolecular pump, in connection with a dimensionless number X that is the ratio of an inter-vane distance S to a chord length C for moving vane blades of rotor impeller (4B) and stationary vane blades of stator impeller (2B), with dimensionless numbers at the outer circumferential portion and the inner circumferential portion of a first vane stage being termed Xo(R) and Xi(R) and dimensionless numbers at the outer circumferential portion and the inner circumferential portion of a second vane stage being termed Xo(S) and Xi(S), and with respect to vane stages that are adjacent along the direction of the rotational shaft, at least one vane stage is provided that satisfies a first relational equation “Xo(R)>Xo(S)” and a second relational equation “Xi(R)<Xi(S)”.

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 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: 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 turbine stage for a turbomachine, the stage comprising a rotor wheel mounted inside a sectorized ring carried by an outer casing, and a sectorized nozzle fastened to the casing downstream from the wheel, each ring sector having a downstream end that is held bearing radially outwards against a cylindrical rail of the casing by means of the nozzle, the casing rail including an annular row of through radial notches reducing the areas for heat exchange by conduction both between the casing and the ring sectors and between the casing and the nozzle.

Abstract: A serial fan module includes a first fan, a second fan and a plurality of guide elements. The first fan has a first frame body and a first impeller disposed within the first frame body. The second fan has a second frame body and a second impeller disposed within the second frame body. The guide elements are disposed between the first impeller and the second impeller, wherein the guide elements guide airflow from the first fan to the second fan. Each of the guide elements has an axial extended part and an inclined part and the inclined part meets the axial extended part at a camber angle. Each of the inclined part has an inner edge toward the first impeller and each of the axial extended part has an outer edge toward the second impeller.

Abstract: A turbine rotor (1) for a gas turbine with a one-piece compressor housing (11) or a compressor housing (11) which is divided in a first plane, a one-piece turbine housing (12) connected to the compressor housing (11) or a turbine housing (12) which is divided in a second plane different from the first plane comprises a compressor area which includes at least two compressor disks (2.1-2.5) and a turbine area which includes at least one turbine disk (13.1, 13.2). The compressor disks and turbine disks are clamped together by a tie rod arrangement which includes at least one tie rod (6) which has a first fastening area, a second fastening area for clamping at least two compressor disks between the first fastening area and the second fastening area, and a third fastening area for clamping at least one turbine disk.

Abstract: A counter-rotating axial flow fan in which the shape of stationary blades of a middle stationary portion is optimized to reduce noise is provided. Defining the maximum axial chord length of front blades as Lf, the maximum axial chord length of rear blades as Lr, and the maximum axial chord length of stationary blades as Lm, a relationship of Lm/(Lf+Lr)<0.14 is satisfied. Defining the maximum dimension between the blade chord for lower surfaces of the stationary blades and the lower surfaces as K1, the maximum axial chord length Lm of the stationary blades satisfies a relationship of Lm/K1>5.8.

Abstract: A low pressure cooling system for a turbine engine for directing cooling fluids at low pressure, such as at ambient pressure, through at least one cooling fluid supply channel and into a cooling fluid mixing chamber positioned immediately downstream from a row of turbine blades extending radially outward from a rotor assembly to prevent ingestion of hot gases into internal aspects of the rotor assembly. The low pressure cooling system may also include at least one bleed channel that may extend through the rotor assembly and exhaust cooling fluids into the cooling fluid mixing chamber to seal a gap between rotational turbine blades and a downstream, stationary turbine component. Use of ambient pressure cooling fluids by the low pressure cooling system results in tremendous efficiencies by eliminating the need for pressurized cooling fluids for sealing this gap.

Abstract: A steam turbine is provided having a relief groove which is arranged in the region of the equalizing piston and extends in the circumferential direction of the rotor. The relief groove, with regard to an inlet passage, is arranged in the axial upstream direction so that it is arranged on the rotor outside a region in which the steam flow enters the bladed flow path via the inlet passage. The relief groove, with regard to the first blade row, is arranged in a region in which the greatest thermal stresses can arise in the rotor. As an option, the relief groove has a cover for reducing vortex flows, and also devices for reducing heating of the groove or devices for active cooling. The steam turbine allows an increased number of risk-free running up and running down operations of the steam turbine with minimum detriment to the turbine performance.

Abstract: A gas turbine engine compressor spacer (230) includes a body (229) with a hollow cylinder shape. The body (229) includes an outer surface (232) with an axially forward end. The body (229) also includes a forward face (233) extending radially inward from the axially forward end. A forward lip (237) extends axially from the body (229). The forward lip (237) includes a forward lip surface (239) located radially inward from the forward face (233). The forward lip surface (239) is the radially outer circumferential surface of the forward lip (237). The spacer (230) also includes a forward undercut (235). The forward undercut (235) is located between the forward face (233) and the forward lip surface (239).

Abstract: This wet compression invention with a vaporizable fluid mist demonstrates major performance improvements over the relevant art in achieving a high degree of saturation, providing sensible cooling, strongly reducing the temperature increase due to compression work, reducing excess diluent air flow for downstream combustion, reducing compression noise, and increasing the achievable compressor pressure ratio.

Abstract: A stage of a steam turbine, having a rotor extending along a longitudinal axis, has a fixed stage and a movable stage arranged successively along a flow channel for feeding steam in a direction substantially parallel to the longitudinal axis; the fixed stage has an inner ring coaxial with the longitudinal axis, and a number of stator blades arranged radially about the inner ring; each stator blade is fixed to an annular top portion of the inner ring, which has a top surface facing the flow channel and crosswise to the longitudinal axis; and the movable stage has a number of rotor blades arranged radially about the rotor and fixed to the rotor.

Abstract: The invention is directed to a turbine with a housing mountable in a drill string and having an inlet opening and an outlet opening for a fluid conveyed through the drill string, and with an impeller arranged in the housing and adapted to be impinged upon by the fluid flowing into the housing, wherein the impeller is axially movably mounted and yieldably supported against the afflux direction in such a way that the supporting force of the yielding support increases as the axial movement of the impeller increases. To reduce the dependence of the turbine rotational frequency and turbine performance on the feed rate and to avoid overloading of the generator, the housing includes in the range of movement of the impeller a frontal first housing section of smaller inner diameter as seen in the afflux direction and a rear second housing section of larger inner diameter as seen in the afflux direction.

Abstract: A multi-stage pump rotor (10) for a turbomolecular pump has at least two separate blade disk rings (17), each having a motor ring (12) and at least one blade disk (14). A cylindrical reinforcement pipe (18), which surrounds the rotor rings (12) of the blade disk rings (17) on the outside without clearance, is provided between the blade disks (14) of adjacent blade disk rings (17). The reinforcement pipe (18) absorbs a large part of the tangential forces occurring during operation such that the pump rotor (10) has improved stability at high rotor speeds.

Abstract: A micro gas turbine engine in which the turbine rotor blades are formed as an integral bladed rotor with cooling air passages formed within the blades and the rotor disk by an EDM process. an adjacent stator vane includes an air riding seal with an air cushion supplied through the vanes to provide cooling, and where the air cushion is then passed into the turbine blades and rotor disk to provide cooling for the turbine blades. With cooling of the turbine blades, higher turbine inlet temperatures for micro gas turbine engines can be produced.

Abstract: A longitudinal stack of gas turbine engine rotor disks each include an annular spacer which transmits compressive preloading of the stack to an adjacent disk, the spacer and an annular shoulder on the disk rim defining an annular slot which accommodates the base of a segmented annular blade cluster which shields the rim from some of the heat associated with the flow of working fluid around the disks.

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 rotor (1) of a gas turbine compressor comprises a multiplicity of welded-together rotor disks (3a, 3b, 4, 5), of which two or more rotor disks (3a, 3b), in a radially outer region (9?), are welded together, and, in a radially inner central region (9), are abutment-joined together. Via the abutment-joining of two rotor disks (3a, 3b), a heat flow (8) radially outward from the center of the rotor (1) is achieved so that the material temperature of the rotor (1) can be kept below a predetermined level during operation. As a result, the service life of the rotor (1) can be increased. In one embodiment, one rotor disk (3a), on its surface, additionally has a recess (7) which can be cooled from outside. The rotor disks (3a, 3b) according to the invention, which are welded and abutment-joined together, can especially be used at the last point in the flow direction of the compressor.

Abstract: Gas turbine engine systems and related methods involving vane-blade count ratios greater than unity are provided. In this regard, a representative turbine stage for a gas turbine engine includes a first set of rotatable blades operative to be positioned downstream of and adjacent to a first set of vanes, a number of blades of the first set of blades being less than a number of vanes of the first set of vanes.

Abstract: A compression stage having a plurality of stator vanes and rotor blades coaxial with a longitudinal centerline axis, each stator vane having an exit swirl angle distribution such that the exit swirl angle has a maximum value at an intermediate radius location and each rotor blade having a blade leading edge adapted to receive the flow from the stator vanes with the exit swirl angle distribution profile.

Abstract: In accordance with one embodiment, a system includes a multi-stage turbine including a first turbine stage including a first wheel having multiple first slots spaced circumferentially about the first wheel, and multiple first blade segments each coupled to at least one of the multiple first slots. The multi-stage turbine also includes a second turbine stage including a second wheel having multiple second slots spaced circumferentially about the second wheel, and multiple second blade segments each coupled to at least one of the multiple second slots. The multi-stage turbine further includes a one-piece interstage seal extending axially between the first and second turbine stages, wherein the one-piece interstage seal comprises a first radial support coupled to the first wheel and a second radial support coupled to the second wheel, and the one-piece interstage seal extends circumferentially across at least two of the multiple first slots or at least two of the multiple second slots.

Abstract: A steam turbine rotor shaft and method of manufacturing the same are provided wherein the sliding characteristics of a journal are improved, and the journal is free from welding cracks and does not need a post heat treatment. The low alloy steel coating layer having better sliding characteristics than 9 to 13% Cr heat resisting steel and an area rate of defects including pores and oxides in a range of 3 to 15% is formed by a high velocity flame spray coating method on a sliding surface of the journal.

Abstract: An exemplary gas turbine engine includes a turbine section operative to impart rotational energy to a compressor section. The turbine section includes at least a low-pressure turbine and a high-pressure turbine, and a ratio of a number of stages in the low-pressure turbine to a number of stages in the high-pressure turbine is 2.

Abstract: A turbo-molecular pump, includes a blade pumping section having a plurality of rotor blades and a plurality of stator blades which are alternately arranged in plural stages along an axial direction of the pump, and a thread groove pumping section having a cylindrical-shaped screw stator, and a rotor cylinder adapted to be rotated inside the screw stator. The rotor cylinder has a lower edge surface located on an upstream side relative to a downstream edge of the screw stator with respect to the axial direction. The turbo-molecular pump of the present invention can minimize flying-out of broken pieces of the rotor cylinder from a discharge port.

Abstract: A gas turbine includes a compressor, a turbine downstream of the compressor having stators, and a rotor connecting the compressor to the turbine. The rotor includes rotor cavities. A supply plenum containing a process by-product gas and a control valve is connected to at least one of the rotor cavities to supply the process by-product gas to at least one of the rotor cavities. A process for operating a gas turbine having a compressor and a turbine includes connecting the compressor to the turbine by a rotor, creating cavities in the rotor, and injecting a process by-product gas into at least one of the cavities. The process further includes regulating the flow of the process by-product gas into at least one of the cavities according to a pre-programmed parameter.

Abstract: A cooling path is defined through a double-flow steam turbine. The steam turbine includes a generator end and a turbine end and a double-flow tub diaphragm disposed axially between the generator end and the turbine end. A first stage of the steam turbine includes a plurality of first stage nozzles and a plurality of first stage buckets, and a steam path is defined from an inlet region, through the turbine end, to a steam outlet. The cooling path includes cooling holes in the tub diaphragm and root radial seals permitting steam flow beneath the tub diaphragm and around the first stage nozzles and back into the steam path before the first stage buckets.

Abstract: A compressor drum for an axial turbo-machine includes a hollow body generally symmetrical in revolution about its axis of rotation and forming a shell at the periphery of which are arranged a plurality of parallel rows of vanes, each row corresponding to one stage of the compressor. Each vane includes at its root an attachment platform housed in a corresponding opening made in the shell of the hollow body. The outer surface of each platform is flush with the outer surface of the shell at the edge of the opening and each platform is attached to the hollow body by welding between the outer edge of the platform and the inner edge of the corresponding opening. The platform has the general shape of a parallelogram, a bisector of which corresponds approximately to the main axis of the profile of the vane at the platform.

Abstract: A rim seal assembly includes an annular seal element circumscribing an engine centerline or axis and mounted on an annular platform including a radially inwardly extending annular platform flange disposed between and connected to forward and aft flanges at distal ends of forward and aft annular elements respectively. Annular forward and aft outer rim cavities radially disposed between the forward and aft annular elements and the platform are axially separated by the platform flange. Cooling slots extend radially across axially facing forward and aft surfaces of the forward and aft flanges. The platform flange and the forward and aft flanges are bolted together. The annular seal element may include seal teeth in sealing relationship with an annular seal land such as in a labyrinth seal. The rim seal assembly may be incorporated in a low pressure turbine use a compressor as a source of cooling air.

Abstract: A steam turbine includes a rotor having a plurality of early stages and a stator portion surrounding a portion of the rotor and arranged such that a leakage region exists between the stator portion and the rotor and having a cooling steam channel that passes cooling steam from one portion of the stator portion to the leakage region. The turbine also includes at least one cooling steam transmission channel axially displaced about the rotor that receives the cooling steam from the leakage region and provides it to at least a portion of the early stages.

Abstract: A device for detecting a shaft rupture in a rotor of a turbine, in particular a medium-pressure turbine, of a gas turbine, in particular of an aircraft engine, at least one stator-side sensor element being positioned downstream from the turbine, in particular in the area of a stator-side blade ring of another turbine, in particular a low-pressure turbine, and, in case of a shaft rupture of the rotor of the turbine, a radially externally situated segment of a last (seen in the direction of flow) rotor-side blade ring of the turbine working together with the, or with each, sensor element to generate an electrical signal corresponding to the shaft rupture.

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: A compressor includes: a compressor spool rotatable about an axis carrying axially-spaced-apart blade rows of compressor blades; a casing surrounding the compressor blades, the casing carrying a liner assembly defining a boundary of a primary compressor flowpath; and a plurality of axially-spaced-apart stator rows of stator vanes carried by the liner assembly, the stator rows alternating axially with the blade rows. At least some of the stator rows are variable stator rows, the stator vanes of which are mounted on trunnions passing through the casing, and are pivotable relative to the casing. An actuator arm is coupled to each trunnions, outside the casing. At least one first bleed slot passes through the liner structure between axially adjacent first and second ones of the variable stator rows; and a first flow path defined by the casing communicates with the first bleed slot and with the exterior of the casing.

Abstract: A compression stage having a plurality of stator vanes and rotor blades coaxial with a longitudinal centerline axis, each stator vane having an exit swirl angle distribution such that the exit swirl angle has a maximum value at an intermediate radius location and each rotor blade having a blade leading edge adapted to receive the flow from the stator vanes with the exit swirl angle distribution profile.

Abstract: A method is disclosed for assembling a multi-stage compressor or a multi-stage turbine for use in a gas-turbine engine. The method comprises the steps of assembling a rotor drum so as to comprise a plurality of rotor discs 17, 18, and then releasably connecting a plurality of static components 38 to the assembled rotor drum 19, thus forming an intermediate structure. The intermediate structure is then inserted within an outer casing 50, preferably by lowering the outer casing 50 over the intermediate structure, whereafter the static components 38 are fixed to the outer casing 50. The static components 38 are then released from their connection to the rotor drum 19 in order to permit rotation of the drum 19 relative to the static components 38 and the outer casing 50.

Abstract: The distance h between a base end face 12ta (12tb) of a stator-blade-wheel outer ring part 12a and the corresponding rotor-blade end face 4aa (4ab), is set at a value larger than the maximum deformation of the running rotor blade wheel 4a during the pump operation. The larger distance h keeps perfectly stator-rotor separation, while the outer ring part 12a of a stator blade wheel 12 goes, in assembling, into the rotor area on account of large shift of any half of stator blade wheels owing to the wide cutting width. It makes also easy stator blade wheel assembling of the pump, as it is allowable to arrange half stator blade wheels with some eccentricity.

Abstract: A turbomachine includes an annular flow path section between a plurality of radially extending stator blades and a plurality of radially extending rotor blades. At least a portion of the flow path section has a circumferentially varying outer periphery.

Abstract: A fluid flow machine includes at least one rotor blade row designed for high work coefficients, which—in its center section and/or averaged over the blade height—provides for an essentially swirl-free relative outflow. The respective rotor, in its trailing edge area, is essentially orientated in meridional direction, which essentially corresponds to the meridional flow direction.

Abstract: A hybrid gas compressor has at least one rotor and shroud which define a compressor gas path extending from an inlet to an outlet. The compressor includes at least two compression stages and one diffusion stage between the inlet and outlet, the compression stages including respective circumferential arrays of blades extending from the rotor and the diffusion stage including a circumferential array of vanes between the compression stages. Blade aerodynamic loadings may be controlled, particularly in the last stage, to provide desired compression characteristics across the compressor. A bleed outlet is optionally located between the inlet and outlet for bleeding from the gas path.