Abstract: A serial axial fan includes a first axial fan, a second axial fan, and a rectifying portion. The second axial fan is on one side in the axial direction with respect to the first axial fan and is connected in series to the first axial fan with the rectifying portion interposed therebetween. The rectifying portion includes a chassis with a cylindrical shape surrounding the central axis, a rectifying blade portion extending to a radially inner side from a radially inner surface of the chassis, and an air feeding space on the radially inner side of the rectifying blade portion. The rectifying blade portion extends at least in the axial direction and is inclined relative to another side in the circumferential direction from another side in the axial direction toward the one side in the axial direction.

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: The present invention relates to a method for designing a flow channel for a turbomachine, in particular a gas turbine that comprises a guide vane cascade having a plurality of guide vanes, which are distributed in the peripheral direction, and flow passages, each of which is bounded by two successive guide vanes, and a support rib arrangement having at least one support rib, wherein a design of one of the flow passages is adapted to this support rib, that it is situated downstream of, in order to reduce a pressure loss and/or a vibrational stimulation.

Abstract: A turbomachine type chemical reactor for processing a process fluid is presented. The turbomachine type chemical reactor includes at least one impeller section and a stationary diffuser section arranged downstream. The impeller section accelerates the process fluid to a supersonic flow. A shock wave is generated in the stationary diffuser section that instantaneously increases static temperature of the process fluid downstream the shock wave for processing the process fluid. which allows thermally cracking a chemical compound, such as hydrocarbon, in the process fluid. Static pressure of the process fluid is simultaneously increased across the shock wave. The turbomachine type chemical reactor significantly reduces residence time of the process fluid in the chemical reactor and improves efficiency of the chemical reactor.

Abstract: A tandem rotor disk apparatus may include a rotor disk body concentric about an axis. The tandem rotor disk apparatus may also include a first blade extending radially outward of the rotor disk body and a second blade extending radially outward of the rotor disk body. The first blade may be offset from the second blade in a direction parallel to the axis. The tandem rotor disk apparatus may be implemented in a gas turbine engine with no intervening stator vane stages disposed between the first blade and the second blade. The tandem rotor disk apparatus may include two separate rotor disk bodies.

Abstract: A turbomachine type chemical reactor for processing a process fluid is presented. The turbomachine type chemical reactor includes at least one impeller section and a stationary diffuser section arranged downstream. The impeller section accelerates the process fluid to a supersonic flow. A shock wave is generated in the stationary diffuser section that instantaneously increases static temperature of the process fluid downstream the shock wave for processing the process fluid. Static pressure of the process fluid is simultaneously increased across the shock wave. The turbomachine type chemical reactor significantly reduces residence time of the process fluid in the chemical reactor and improves efficiency of the chemical reactor.

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 supply duct of a compressor of a turbine engine, formed from internal and external walls of revolution around an axis and opposite one another to define a circulation stream of a fluid, is provided. The stream allows the fluid to be routed from the inlet of the duct to the inlet of the compressor. The radius of the external wall at the inlet of the duct is greater than the radius of the duct at the inlet of the compressor. The duct includes a portion for which the radius of the external wall along the portion is less than the radius of the external wall at the inlet of the compressor, and the radius of the internal wall along the portion of the duct is less than the radius of the internal wall at the inlet of the compressor.

Abstract: A turbine nozzle segment includes a radially-inner endwall, a radially-outer endwall, a pair of airfoil-shaped vanes extending between the radially-inner endwall and the radially-outer endwall, and respective reinforcing ribs extending between the pressure and suction sidewalls of the vanes. The back face of the radially-inner endwall and/or the back face of the radially-outer endwall has a pocket formed therein in an area between the pressure sidewall of the first vane and the suction sidewall of the second vane to enhance stiffness distribution between the second vane and the radially-inner endwall and/or radially-outer endwall.

Abstract: A series fan structure includes at least one first combining part and a series fan assembly, which has a first fan and a second fan mated with the first fan. The first fan has a first base and a lateral side, which together internally define a first passage. The first base is outwardly extended to form a plurality of a first supporting portions connected to the first lateral side. A first connecting space is defined among the first supporting portions and communicated with the first passage. The second fan has a second base and a lateral side, which together internally define a second passage, which is communicated with the first passage and the first connecting space. The first combining part is located in the first connecting space and connected to the first supporting portions and has a plurality of first holes communicated with the first and the second passage.

Abstract: A diffuser of a thermal energy machine, in particular of a gas turbine, has a diffuser inlet, a diffuser outlet, and a plurality of air-guiding elements, wherein an air mass flow enters the diffuser through the diffuser inlet, and wherein the air mass flow that has entered the diffuser exits the diffuser through the diffuser outlet and flows off as a plurality of partial air mass flows by the air-guiding elements. At least two immediately adjacent air-guiding elements of the plurality of air-guiding elements are designed in such a way that the flow-off angles thereof with respect to the circumferential surface formed by the outlet opening of the diffuser outlet extending circumferentially in the circumferential direction differ from each other.

Abstract: Disclosed is a method for operating a compressor of a turbomachine, in which, when viewed in the direction of a main flow, a slowly dropping degree of reaction being raised in one compressor area more quickly than previously dropped and the raised degree of reaction then being reduced again in a subsequent compressor area more slowly that it was previously raised, a compressor of a turbomachine including at least one degree of reaction rising more quickly in one compressor area in comparison to adjacent compressor areas with respectively a dropping degree of reaction, whereby a sawtooth-shaped curve of the degree of reaction results, and a turbomachine including a compressor of this type.

Abstract: The present invention relates to a compressor stage for a gas turbine, in particular, an aircraft engine, having a row of rotating blades (3) and a row of guide vanes (4), which is adjacent downstream, wherein the choke point ? and the aspect ratio ARax, which is defined by the quotient between average channel height (h) and average chord length (lax), satisfy the condition ?>?1.33·ARax+5.16.

Abstract: A series fan structure includes a series fan assembly and an assembling member. The series fan assembly has a first fan and a second fan. The first and second fans are correspondingly serially connected with each other. The first fan has a first fan frame defining a first receiving space. The second fan has a second fan frame defining a second receiving space. The assembling member is disposed between the first and second fan frames. The assembling member is formed with multiple straight-through perforations in communication with the first and second receiving spaces. The series fan structure improves the vibration and noise problem of the conventional series fan structure and is able to increase airflow volume.

Abstract: An exhaust diffuser for a gas turbine includes an annular duct. A row of struts is arranged in the duct. In a region downstream of the trailing edges of the struts, the cross-sectional area of the duct decreases to a local minimum and then increases again towards the outlet end of the duct. Thereby the gas flow is locally accelerated downstream of the struts. This stabilizes the boundary layer of the flow in this region and leads to a marked increase in pressure recovery for a wide range of operating conditions.

Abstract: An accessory gear box for a gas turbine engine having a drive shaft with a rotational axis and a tower shaft coupled to the drive shaft is provided. The accessory gear box includes a first plurality of gears arranged, which extend along a first axis substantially parallel to the rotational axis of the drive shaft. The accessory gear box includes a second plurality of gears, which extend along a second axis. The accessory gear box includes a first shaft, with one of the first plurality of gears coupled to the first shaft, and one of the second plurality of gears coupled to a second shaft. The one of the second plurality of gears coupled to the first shaft includes a first engagement surface and a second engagement surface, and the second engagement surface is coupled to another one of the second plurality of gears to drive the second shaft.

Abstract: Architecture that harnesses energy from natural atmospheric wind and water currents and self-generated wind and water currents from moving vehicles and natural fluid flow found in nature for moving or stationary applications. The power generation system harnesses energy from natural atmospheric sources utilizing pneumatic and/or hydraulic turbines with compound nozzles, meteorological sensors, computer controlled harmonic resonance valves, a control system, and other components.

Abstract: A system, in certain embodiments, includes a plurality of detachable, three-dimensional diffuser vanes attached to a diffuser plate of a centrifugal compressor. In certain embodiments, the detachable, three-dimensional diffuser vanes may be attached to the diffuser plate using threaded fasteners. In addition, dowel pins may be used to align the detachable, three-dimensional diffuser vanes with respect to the diffuser plate. However, in other embodiments, the detachable, three-dimensional diffuser vanes may include a tab configured to fit securely within a groove in the diffuser plate. In addition, the tabs of the detachable, three-dimensional diffuser vanes may include indentions that mate with extensions extending from the diffuser plate, wherein the tabs may slide into slots between the extensions and the grooves of the diffuser plate.

Abstract: A process is provided for replacing a first nozzle block coupled to a nozzle chamber with a second nozzle block. The process may comprise removing the first nozzle block from the nozzle chamber. The nozzle chamber may comprise a main body having at least one inlet, at least one passage and at least one exit. The process may further comprise coupling inner and outer retaining rings to the nozzle chamber main body; engaging a second nozzle block with the inner and outer retaining rings; forming a bore so as to extend partly in one of the inner and outer retaining rings and the second nozzle block; and locating an anti-rotation pin in the bore.

Abstract: A turbomachine including at least one blade-row group that is arranged in the main flow path and at least two rows of blades that are adjacent to each other in the main flow direction, each row having a plurality of blades, whereby the trailing edges of the blades of the upstream row of blades and the leading edges of the blades of the downstream row of blades are arranged at an axial edge distance that decreases from the center of the main flow path in the direction of at least one main flow limiter.

Abstract: There is provided an axial flow compressor that improves reliability on an increase in a blade loading on a last-stage stator vane of the axial flow compressor due to a partial load operation of a gas turbine. An annular flow passage is formed by a rotor having multiple rotor blades fitted thereto and a casing having multiple stator vanes fitted thereto, two or more of the stator vanes are disposed downstream of a last-stage rotor blade that is the rotor blade disposed at the most downstream side in a flow direction of the annular flow passage, a blade loading on a first stator vane disposed at the most upstream side is set to be smaller than a blade loading of a second stator vane disposed downstream of the first stator vane by one row.

Abstract: In an engine disk and blade combination, the metallic disk has a plurality of first blade attachment slots and a plurality of second blade attachment slots circumferentially interspersed with each other. There is a circumferential array of a plurality of first blades. Each first blade has an airfoil and an attachment root. The attachment roots are respectively received in associated said first attachment slots. There is a circumferential array of second blades. Each second blade has an airfoil and an attachment root. The attachment roots are respectively received in associated said second slots. The first blades and second blades are non-metallic. The first blades are radially longer than the second blades. The first slots are radially deeper than the second slots.

Abstract: A turbine includes a turbine housing having two gas passages of substantially the same flow area. A nozzle ring is disposed in the housing and around the turbine wheel. The nozzle ring includes first and second outer rings, and an inner ring disposed between the first and second outer rings. First and second pluralities of vanes are disposed between the rings. The second outer ring has a thicker cross section than the first outer ring such that a larger flow area is created between the first outer ring and the inner ring than a flow area created between the second outer ring and the inner ring.

Abstract: The present application relates to the compressor stator of an axial turbomachine. The stator comprises an annular row of main stator blades and auxiliary blades each of which are associated with a main blade. The auxiliary blades are located at the trailing edges of the main blades and are in the vicinity of the pressure faces of the main blades. The auxiliary blades are aligned to generate a low-pressure area at the trailing edges of the main blades. Thus, a flow bypassing a main blade by its suction face is sucked in by the low-pressure area when it approaches the trailing edge of the main blade. Stalling is thus avoided and the efficiency of the machine is improved.

Abstract: A turbomachine includes a plurality of rotating blades adjacent to a plurality of stationary vanes attached to a stationary casing. Stationary vanes are spaced apart circumferentially with equal spacing around an inner perimeter of the casing. Optionally, stationary vanes are offset radially and/or circumferentially. Stationary vanes may alternate with leading and/or trailing edges at different distances from the rotating blades. In one embodiment, stationary vanes have tapered leading and/or trailing edge angles to homogenize flow and reduce stator wake excitation, flow excitation, and acoustic excitation due to interaction with spinning modes of acoustic pressure pulsations at rotating blade passing frequency.

Abstract: The nozzles include inlet ports circumferentially arranged into a group and an opening and closing valve disposed so as to be commonly functioned. This arrangement realizes a gas turbine with no deterioration of efficiency against fluctuation of gas inflow rate by only one heat-resistant moving component additionally installed. The variable nozzle number turbine according to this invention also realizes a turbine compressor with no surge limit when used as a regenerator. These turbines further realizes a small-sized gas turbine with improved fuel economy under partial load conditions and a turbocharger for regenerating electric power as well as improving efficiency of a displacement engine.

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 diaphragm for a steam turbine is disclosed that has at least one arc of admission. The arc of admission has a plurality of nozzles arranged about the circumference of the diaphragm and are configured to eject a working fluid at succeeding rotor blades axially-spaced from the diaphragm. The flow area of the first few nozzle vanes in the arc of admission is gradually increased along the arcuate length of the diaphragm, thereby mitigating the load impulse absorbed by each rotor blade as it enters the arc of admission. The flow area of the last few nozzle vanes in the arc of admission is gradually decreased so that each rotor blade does not suddenly go from full load impulse to zero and thereby contribute to the fatigue of the rotor blade and create unwanted noise.

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 system includes a rotary machine including, a stator, a rotor configured to rotate relative to the stator, a plurality of axial grooves disposed along a circumference of the stator or the rotor, a plurality of blade segments disposed along the circumference, wherein each blade segment of the plurality of blade segments comprises a blade coupled to a mounting base supported in a respective axial groove of the plurality of axial grooves, and the plurality of blades has a non-uniform blade spacing about the circumference.

Abstract: A system, including a rotary machine including: a stator, a rotor configured to rotate relative to the stator, wherein the rotor comprises a plurality of blades having a non-uniform spacing about a circumference of the rotor.

Abstract: A system, in certain embodiments, includes a centrifugal compressor diffuser including a flow path having a first surface and a second surface defining opposite axial sides of the flow path. The centrifugal compressor diffuser also includes multiple vanes extending from the first surface to the second surface of the flow path. A first profile of each vane varies along an axial direction. The centrifugal compressor diffuser further includes multiple vanelets extending from the first surface toward the second surface in the axial direction. A first axial extent of each vanelet is less than a second axial extent of the flow path. In addition, a second profile of each vanelet varies along the axial direction and/or the vanelets form a non-periodic pattern around a circumference of the flow path.

Abstract: An exemplary control stage of a steam turbine is disclosed, which includes plural staging valves circumferentially distributed around the turbine for regulating steam admission flow so as to control the loading of the turbine. Nozzle chambers are connected to a downstream end of each staging valve. An arc of admission forms the downstream portion of each nozzle chamber, and control stage nozzles in the arcs of admission define the downstream end of the nozzle chamber. Each nozzle chamber has at least two arcs of admission, each with a different circumferential dimension.

Abstract: A stator for a turbomachine. The stator includes a plurality of nozzle vanes, with the distance between each of the plurality of nozzle vanes varying circumferentially around the stator in a repeating pattern configured to distribute a vibration energy over a plurality of side band vibration frequencies around a nominal passing frequency in a rotor.

Abstract: A guide vane for a gas turbine with a vane base body which is of single-piece design and comprises a profiled vane blade extending between a vane root and a cover plate and also the vane root formed integrally with the vane blade and the cover plate formed integrally with the vane blade, is intended, in a relatively simple way to be able to be matched to the individual conditions of use with especially little outlay on apparatus and logistics. For this purpose, according to the invention, a flow-routing body with an advance guide blade that is connected upstream of the vane blade as seen in the direction of flow of the working medium of the gas turbine is joined to the vane base body.

Abstract: In a blade row of an axial flow type compressor in which a rotor blade row and a stator blade row are alternately arranged in an axial direction, the stator blade row 10 is formed by plural main stator blades 12 and plural sub-stator blades 14. Each main stator blade 12 is formed by a basic blade portion 12a which has the same shape as that of each sub-stator blade and a forward blade portion 12b which extends to the upstream side of the basic blade portion. The basic blade portion 12a are located at the same position in an axial direction. The forward blade portion 12b forms a forward stator blade row which has a circumferential interval larger than that of the basic stator blade row in the vicinity of at least a radial inner end.

Abstract: A method of designing a multistage turbine for a turbomachine in which each turbine stage comprises a stator row and a rotor row each made up of an annular row of airfoils, wherein, for all of the stator or rotor rows, it consists in simultaneously modifying the shapes of the airfoils of said rows to straighten out the wakes from the trailing edges of said airfoils, then in angularly positioning the rows in such a manner that the wakes from the airfoils of the stator (or rotor) airfoils impact against the leading edges of the stator (or rotor, respectively) airfoils of the rows situated downstream, in order to achieve multistage aerodynamic coupling simultaneously over the turbine as a whole.

Abstract: An assembly of airfoils in a compressor of a turbine engine that includes at least three axially stacked rows of airfoils: a middle airfoil row, a first upstream airfoil row, and a first downstream airfoil row. The middle airfoil row may be bordered on each side by the first upstream airfoil row, which comprises the first row of airfoils in the upstream direction from the middle airfoil row, and the first downstream airfoil row, which comprises the first row of airfoils in the downstream direction from the middle airfoil row. The first upstream airfoil row and the first downstream airfoil row may have substantially the same number of similarly shaped airfoils. The first upstream airfoil row and the first downstream airfoil row each may comprise a row of rotor blades, which rotate at substantially the same speed during operation. The middle airfoil row may comprise a row of stator blades, which remains substantially stationary during operation.

Abstract: A gas turbine engine turbine comprises an annular array of nozzle guide vanes and an annular array of turbine blades mounted within its annular casing. An array of radially extending protrusions are positioned axially upstream of said array of said nozzle guide vanes and protruding inwardly from the inner casing wall so as to mix the tangential momentum component of the overtip leakage fluid flow before it reaches the array of nozzle guide vanes.

Abstract: A flow structure of a gas turbine, in particular for an aircraft engine, in a transitional channel between two compressors or in a transitional channel between two turbines or in a transitional channel of a turbine outlet housing downstream from a low-pressure turbine is disclosed. Supporting ribs are positioned in the transitional channel and spaced a distance apart in the circumferential direction of the transitional channel. At least one guide vane and/or guide rib is positioned between two supporting ribs spaced a distance apart from one another. The flow outlet edge of the guide rib or each guide rib runs upstream from the flow outlet edges of the supporting ribs.

Abstract: A method of designing a multistage turbine for a turbomachine in which each turbine stage comprises a stator row and a rotor row each made up of an annular row of airfoils, wherein, for all of the stator or rotor rows, it consists in simultaneously modifying the shapes of the airfoils of said rows to straighten out the wakes from the trailing edges of said airfoils, then in angularly positioning the rows in such a manner that the wakes from the airfoils of the stator (or rotor) airfoils impact against the leading edges of the stator (or rotor, respectively) airfoils of the rows situated downstream, in order to achieve multistage aerodynamic coupling simultaneously over the turbine as a whole.

Abstract: A return section in a multistage centrifugal compressor includes a plurality of circumferentially positioned flow vanes, and a plurality of splitter vanes disposed between the flow vanes. The splitter vanes serve to minimize or eliminate regions of reverse or separated flow, resulting in improved compressor performance.

Abstract: A compressor includes a stator having a plurality of vane sectors each with an outer shroud having a forward and a rearward guide vane extending there form, the vane sector being mounted in an annular slot of the casing, the guide vanes forming a gap between the tips and an inner shroud, and a spring member located on the outer shroud in the rearward portion to bias the rearward shroud in a radial direction, where a space is formed in the slot to allow for the outer shroud to move in the radial direction when the rearward guide vane tip makes contact with the inner shroud due to thermal growth of the stator components. This arrangement allows for the rearward guide vane tip to maintain contact with the inner shroud without inducing high compressive stresses in the stator components due to the thermal growth.

Abstract: A nozzle box includes a torus, bridge ring and nozzle ring portions wherein steam flowing generally circumferentially in the torus portion is redirected for flow generally axially by the bridge ring portion. The steam exiting the bridge ring portion flows into nozzles formed by adjacent partitions of the nozzle ring portion. Bridges in the bridge ring portion are tangentially leaned to exactly match the angle of inclination of the leading edges of the axially adjacent partitions at like circumferential locations. For strength purposes, there are two bridges for each partition. Every other bridge axially registers with the leading edge of the axially downstream partition matching its angle while remaining bridges are equally spaced between the aligned bridges. In this manner, the steam flow is straightened by the bridges to match the angle of the leading edge of the nozzle partitions with consequent reduction of passage area loss.

Abstract: A method of assembling a gas turbine engine includes providing a plurality of stator vane sectors that each include an equal number of stator vanes that are circumferentially-spaced such that a first circumferential spacing is defined between each pair of adjacent stator vanes within the sector, and coupling the plurality of stator vane sectors together to form a stator vane assembly such that a second circumferential spacing is defined between each pair of adjacent stator vanes coupled to adjacent sectors, wherein the second circumferential spacing is different from the first circumferential spacing.

Abstract: A gas turbine engine turbine stage has an annular row of turbine airfoils including a first plurality of first airfoils having a first chord length and a second plurality of second airfoils having a second chord length shorter than the first chord length. At least one of the second airfoils is circumferentially disposed between each adjacent pair of the first airfoils. A second leading edges of the second airfoils are located downstream of first leading edges of the first airfoils. The first and second airfoils have different first and second temperature capabilities respectively and the first temperature capability is greater than the second temperature capability. The first and second airfoils may be made from different alloys and/or the second airfoils may be constructed to use lower amounts of cooling airflow than the first airfoils. More than one of the second airfoils may be disposed between each pair of the first airfoils.

Abstract: A method of assembling a gas turbine engine includes coupling an inner engine casing to an outer engine casing such that an annular flow path is defined therebetween, coupling a plurality of circumferentially-spaced support struts between the inner and outer casings, and coupling a plurality of circumferentially-spaced outlet guide vanes within the flow path upstream from the support struts, such that a first circumferential spacing is defined between a first guide vane and a second guide vane, and a second circumferential spacing is defined between the second guide vane and a third guide vane. The guide vanes are arranged such that the second guide vane is between the first and third guide vanes and the second circumferential spacing is different from the first circumferential spacing.

Abstract: A gas turbine engine turbine stage has an annular row of turbine airfoils including a first plurality of first airfoils having a first chord length and a second plurality of second airfoils having a second chord length shorter than the first chord length. At least one of the second airfoils is circumferentially disposed between each adjacent pair of the first airfoils. A second leading edges of the second airfoils are located downstream of first leading edges of the first airfoils. The first and second airfoils have different first and second temperature capabilities respectively and the first temperature capability is greater than the second temperature capability. The first and second airfoils may be made from different alloys and/or the second airfoils may be constructed to use lower amounts of cooling airflow than the first airfoils. More than one of the second airfoils may be disposed between each pair of the first airfoils.

Abstract: A blade row arrangement for turbo-engines has an axial construction with two guide blade rows fixedly positioned relative to one another and having a different number of blades while the blade pitch is constant in each case, and having a moving blade row arranged between the two guide blade rows. The blades of the first guide blade row, in a first partial area of the row, successively have an identical axial offset; the axial offset being selected as a function of the blade number ratio of the two guide blade rows such that it increases the effective flow-off cross-section when the first guide blade row has more guide blades than the second guide blade row and reduces the effective flow-off cross-section when the first guide blade row has less guide blades than the second guide blade row. The blades of the first guide blade row, in a second partial area of the row, successively have an axial offset which is opposite in relation to the blades in the first partial area.

Abstract: Two stages of relatively rotatable guide vanes (22,24) are arranged in tandem. Shroud portions (42,46) on the vanes overlap in loose engagement, until an associated gas turbine engine starts operating. Gas loads then act on the downstream vane stage (24) and move it in a downstream direction until the opposing flanks of radially opposing lands (44,48) on respective shrouds (42,46) engage, whereupon, gas leakage via the overlap onto the turbine casing (40) is at least substantially reduced.