EP3109408A1 - Dispositif de stator pour une turbomachine comprenant un dispositif de carter et plusieurs aubes directrices - Google Patents

Dispositif de stator pour une turbomachine comprenant un dispositif de carter et plusieurs aubes directrices Download PDF

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Publication number
EP3109408A1
EP3109408A1 EP16175990.7A EP16175990A EP3109408A1 EP 3109408 A1 EP3109408 A1 EP 3109408A1 EP 16175990 A EP16175990 A EP 16175990A EP 3109408 A1 EP3109408 A1 EP 3109408A1
Authority
EP
European Patent Office
Prior art keywords
platform
region
stator
annular channel
flow
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP16175990.7A
Other languages
German (de)
English (en)
Inventor
Patrick Grothe
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Rolls Royce Deutschland Ltd and Co KG
Original Assignee
Rolls Royce Deutschland Ltd and Co KG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Rolls Royce Deutschland Ltd and Co KG filed Critical Rolls Royce Deutschland Ltd and Co KG
Publication of EP3109408A1 publication Critical patent/EP3109408A1/fr
Withdrawn legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D17/00—Regulating or controlling by varying flow
    • F01D17/10—Final actuators
    • F01D17/12—Final actuators arranged in stator parts
    • F01D17/14—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits
    • F01D17/16—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes
    • F01D17/162—Final actuators arranged in stator parts varying effective cross-sectional area of nozzles or guide conduits by means of nozzle vanes for axial flow, i.e. the vanes turning around axes which are essentially perpendicular to the rotor centre line
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/08—Sealings
    • F04D29/083—Sealings especially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/541—Specially adapted for elastic fluid pumps
    • F04D29/545—Ducts
    • F04D29/547—Ducts having a special shape in order to influence fluid flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/40—Casings; Connections of working fluid
    • F04D29/52—Casings; Connections of working fluid for axial pumps
    • F04D29/54—Fluid-guiding means, e.g. diffusers
    • F04D29/56—Fluid-guiding means, e.g. diffusers adjustable
    • F04D29/563—Fluid-guiding means, e.g. diffusers adjustable specially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines

Definitions

  • the invention relates to a stator for a turbomachine with a housing device and a plurality of guide vanes according to the type defined in more detail in claim 1 and a Schaufelradvorutter according to the closer defined in claim 15.
  • Stator devices of compressors for aircraft engines are well known in practice.
  • Such stator devices are designed with adjustable running guide vanes, which are arranged circumferentially distributed in a housing device and each having an airfoil and in the radial direction of the stator laterally adjoining, also referred to as penny platform.
  • the platforms together with the housing means define a core flow channel of the aircraft engine in the radial direction of the stator device.
  • a spindle-shaped region adjoins the platforms, via which the guide vanes are rotatably mounted about a central axis of the spindle-shaped region relative to the housing device.
  • the platform which has a circular cross-section with respect to the central axis of the spindle-shaped area, has a larger cross section than the spindle-shaped area with respect to the central axis of the spindle-shaped area.
  • the platforms are each mounted in a recess of the housing device which is concentric with the central axis of the spindle-shaped region, wherein there is a circumferential gap between the housing device and the platforms of the guide vanes.
  • a surface facing away from the core flow channel of the platforms relative to the housing means in the radial direction is spaced.
  • a turbomachine designed with such a stator device disadvantageously has an undesirably low efficiency.
  • stator of a compressor or a turbine for a turbomachine in particular a stationary gas turbine or aircraft engine, proposed with a housing device and a plurality of vanes, which are arranged circumferentially distributed on the housing means, wherein the vanes each with an airfoil and in each case at least a platform are executed.
  • the platforms form, at least in some areas, a surface of an annular channel through which working fluid flows during operation of the stator device and limit this, at least in regions, in the radial direction of the stator device.
  • the platforms are each adjustable relative to the housing device, in particular rotatably mounted about a central axis of the platform.
  • At least one platform in the axial direction of the stator device is arranged between two reference points of the annular channel, wherein a first reference point represents an edge point of the annular channel which is 10% of an axial extent of the platform upstream of a front end of the platform with respect to a central longitudinal axis of the platform Platform is arranged, and wherein a second reference point represents an edge point of the annular channel, which is arranged with respect to the central longitudinal axis of the platform by 10% of the axial extent of the platform downstream of a rear end of the platform, wherein at least one edge region of the platform in the radial direction of the stator opposite a rectilinear connection of the two reference points in the annular channel protrudes.
  • the solution according to the invention is based on the recognition that a flow region, in particular at least partially through a recess in the radial direction of the guide vane between the platform and the housing device and a recess in the radial direction of the stator between a side facing away from the airfoil surface of the platform and the housing device is formed, a part of the guided through the annular channel working fluid is guided as a leakage flow during operation of a running machine according to the invention with the stator device.
  • the leakage flow is guided in operation due to a pressure difference between the pressure side and the suction side of the airfoil and an increasing pressure gradient in the flow direction of the working fluid in the annular channel through the flow region.
  • part of a main flow passing through the annular passage is undesirably guided from the downstream pressure side of the airfoil to an upstream suction side of the airfoil via a side of the platform remote from the annular passage the area of which the pressure is lower than the pressure in the area of the pressure side.
  • the leakage flow in the region of the suction side of the blade leaves the flow region, the leakage flow interacts with the main flow of the working fluid in the annular channel, whereby a so-called blocking region with a flow velocity reduced in relation to the surrounding regions of the main flow occurs in the main flow. This effect causes the leakage flow has a significant negative impact on the efficiency of the turbomachine.
  • the pressure conditions of the main flow in the region of the edge region of the platform of the guide vanes are advantageously influenced in such a way that, during operation of the stator device, through the flow region flowing mass flow opposite an embodiment of the guide vane is reduced with not projecting into the annular channel platform.
  • a lower mass flow in the region of the suction side of the blade leaves the flow region compared to conventionally designed platforms, as a result of which a lossy interaction of the leakage flow with the main flow is reduced.
  • a turbomachine embodied with the stator device according to the invention advantageously has an improved efficiency and thus also a reduced specific fuel consumption.
  • the pressure conditions in the region of a platform adjacent in the circumferential direction of the stator device are advantageously influenced by the edge region of the platform projecting into the annular channel.
  • a leakage flow guided through the flow region during operation of the stator device is particularly low if the edge region of the platform projecting into the annular channel in relation to the straight-line connection of the two reference points is located in a front region of the platform in the axial direction of the stator device. This results from the fact that the main flow in the region of the edge region projecting into the annular channel is deflected and dammed thereby, whereby a static pressure in the region of an exit of the leakage flow from the flow region is increased. A pressure difference between the downstream pressure side and the upstream suction side of the airfoil is thus reduced, which in turn results in a reduced mass flow passing through the flow region.
  • the edge region of the platform projecting into the annular channel in relation to the straight-line connection of the two reference points is located in a rear region of the platform in the axial direction of the stator device.
  • the platform of the guide vane is arranged in an inner and / or outer edge region of the vane blade with respect to the radial direction of the stator device. Regardless of which in which in the radial direction facing edge region of the annular channel, the platform is arranged, can be reduced by the projecting into the annular channel edge region of the platform flowing through the flow area mass flow.
  • the pressure conditions in the region of the edge region of the platform extending into the annular channel are improved in a particularly advantageous manner if the edge region of the platform projecting into the annular channel is at least 0.3%, in particular between 0.5% and .mu., Relative to the straight-line connection of the reference points 2.5% to 4%, preferably between 0.7% and 1.5% of an extension of the annular channel in the radial direction of the stator in the range of the edge region, ie a perpendicular to the axial direction of the stator arranged width of the annular channel, extending into the annular channel.
  • the edge region of the platform extending into the annular channel extends in an axial direction of the stator device .
  • rear area is executed with a rounding.
  • the leakage flow becomes in one arrangement of the overhang in an axial direction of the stator device rear edge region of the platform deflected after a discharge from the flow area in the region of the overhang and accelerated around it, so that the main flow is advantageously influenced by the leakage flow discharged from the flow area leakage in a reduced extent.
  • the overhang By arranging the overhang in a front edge region of the platform in the axial direction of the stator device, a static pressure in the region of the front edge region of the platform during operation of the stator device is advantageously further increased.
  • the overhang can be executed in particular nose-shaped.
  • the overhang in the axial direction of the stator device overlaps the housing device adjoining the platform at least in regions.
  • a pressure in the region of the suction side of the guide vane becomes advantageously strong due to a large blocking effect by the overhang increased, whereby a funded by the flow area mass flow is advantageously low.
  • the edge region of the platform projecting beyond the linear connection of the two reference points in the annular channel extends over an angular range of, for example, greater than 20 °, in particular greater than 30 °, with respect to a circumferential direction of the guide vane.
  • a transition from the edge region projecting into the annular channel to regions of the platform which do not protrude into the annular channel is preferably flowing, ie. H. without step, executed. It when the projecting into the annular channel edge region of the platform completely surrounds the platform is particularly advantageous.
  • a flow region is provided, over which, during operation of the stator device, a working fluid flows at least partially in the radial direction of the stator device on a side of the platform facing away from the annular channel from a pressure side of the blade to a suction side of the blade, at least a suction device, which adjoins the flow region and is formed by a recess, via which working fluid can be diverted from the flow region during operation of the stator device.
  • a mass flow of the leakage flow which enters the main flow of the annular channel in the region of the suction side of the blade from the flow region, is reduced or flow of leakage flow into the annular channel in the region of the suction side of the blade is completely prevented.
  • the suction device is connected on a side facing away from the flow region with a space in which there is a static pressure which is less than a static pressure in the flow region.
  • the main flow in the region of the suction side of the blade is significantly less affected compared to embodiments without a suction device.
  • a lossy interaction of the leakage flow with the main flow is thereby reduced, which advantageously results in an improved efficiency and thus also a reduced specific fuel consumption of a turbomachine designed with the stator device according to the invention.
  • the reduction of the leakage flow flowing into the main flow in the region of the stator device by the provision of the suction device according to the invention also has, in addition, a positive effect on vane devices arranged downstream in the annular channel of the stator device.
  • the platforms may form an inner part of the surface of the annular channel that is external in the radial direction of the stator device and / or a radial direction of the stator device, wherein a suction device may be provided in the region of the inner and / or outer platforms in the radial direction of the stator device.
  • the suction device is preferably designed as a material recess in the housing device and may be formed, for example, channel-shaped or as a bore. Alternatively, the recess may be designed with a separate component.
  • the suction device directly adjoins the surface of the annular channel.
  • the suction device adjoins the flow region at a distance from the surface of the annular channel.
  • the suction device adjoins, in the case of a platform, which adjoins the blade of the guide blade radially outward, preferably in the radial direction of the stator device outside of the annular channel to the flow region, whereas the suction device is adjacent to the flow region in the radial direction of the stator device within the annular channel in the case of a platform which connects radially inwards to the blade of the guide blade.
  • a stator device characterized by small losses in the region of the suction device extends essentially in the radial direction of the stator device.
  • the suction device designed, for example, as a bore can also be angled relative to the radial direction of the stator device or have a curved course, wherein an embodiment of the suction device is selected in particular such that a flow in the region of the suction device does not come off during operation of the stator device.
  • the suction device may be connected to the flow region in one of the pressure side of the blade and / or in an area facing the suction side of the blade of the guide blade, wherein a suction effect of the suction device is not significantly influenced by its position.
  • the suction device extends in an advantageous embodiment of the stator device in the circumferential direction of the guide blade over an angular range which is in particular greater than 20 °, preferably greater than 30 °. wherein the respectively selected angular range is adapted to the maximum adjustment angle of the guide vane and may for example also be 180 °.
  • the suction device integrated into the housing device with the flow region in each position of the guide vane.
  • a mass flow that enters the main flow again in the area of the suction side of the blade can be reduced.
  • the suction device is designed such that the suction device is connected to the flow region only at certain adjustment positions of the guide vanes and is not connected to the flow region at other adjustment positions. In this way, for example, it can be achieved in a simple manner that leakage flow is extracted via the suction device from the flow region only during a partial load operation of the aircraft engine and not during a nominal operation.
  • the suction device extends with respect to a central axis of the stator device circumferentially substantially circumferentially in the housing device.
  • Such an embodiment of the suction device is easy to manufacture. Leakage flows in the flow regions of all the stator blades of the stator device can in this case be sucked off from the respective flow regions in a simple manner and supplied, for example, to a common space.
  • webs can be provided distributed circumferentially relative to a central axis of the stator device, via which the housing device is reinforced in the region of the peripheral suction device.
  • the housing device may have, in the region of the guide vane, a recess adjoining the annular channel, via which a mass flow can be specifically removed in a conventional manner during operation of the stator device of the main flow. Together with the extracted via the suction device from the leakage flow mass flow taken in the region of the recess mass flow can be used as bleed air in a known manner.
  • An efficiency of a turbomachine designed with such a Schaufelradvoriques is advantageously high, since in addition to an improvement in efficiency by the reduction of the introduced from the flow area in the main flow mass flow in the range described in more detail above the extracted during operation of the stator of the leakage flow mass flow itself to improve the efficiency of the turbomachine is used.
  • the mass flow of the stator device removed by the suction device of the leakage flow is preferably fed to the rotor device, which is directly upstream in the axial direction of the impeller device of the stator device. There is an optimum of aspirated mass flow in which a maximum efficiency improvement is achieved.
  • the solution according to the invention also increases the surge limit of a paddle wheel device designed as a compressor, as a result of which, for example, a number of blades of the compressor can be reduced or a step pressure ratio can be increased.
  • the line region has at least one nozzle, via which working fluid of the rotor device can be supplied during operation of the bucket wheel device.
  • a plurality of nozzles arranged distributed in the circumferential direction of the blade wheel device or one or more nozzles extending over a larger angular range of, for example, greater than 45 ° or a completely circumferential nozzle may be provided.
  • the mass flow taken in the region of the stator device via the suction device of the leakage flow is used for other applications.
  • the mass flow taken from the leakage flow is introduced into a bypass duct of an engine.
  • Fig. 1 shows a section of a turbomachine, which is designed in the present case as a jet engine 1, but in an alternative embodiment can also represent a stationary gas turbine.
  • an annular channel or core flow channel 3 of the jet engine 1 is shown in the region of a high-pressure compressor 2 Schaufelradvoriques, wherein various stages 6A, 6B, 6C, 6D of the high-pressure compressor 2 can be seen, each consisting of a rotor device 4 and one in the axial direction A of the jet engine 1 downstream of the rotor device 4 arranged stator 5 exist.
  • the rotor device 4 has a multiplicity of blade devices 9, which are designed to be circumferentially distributed with a disk wheel 11 and rotate during operation of the jet engine 1 about a central axis of the jet propulsion system 1.
  • the stator 5 is, however, with a plurality of likewise each an airfoil 13 having guide vanes 12, wherein the respective identical executed guide vanes 12 circumferentially distributed in the radial direction R of the jet engine 1 are arranged on the outside of a housing device 8.
  • the platforms 14 are each connected to a spindle-shaped region 15 and embodied in the present case, wherein the platforms 14 with respect to a central axis 18 of the spindle-shaped portion 15 have a larger cross-section than the spindle-shaped portion 15.
  • the guide vanes 12 are arranged with the platforms 14 and the spindle-shaped regions 15 in recesses 16 of the housing device 8, wherein the spindle-shaped regions 15 are mounted in the recesses 16 via bushes 17.
  • the guide vanes 12 are rotatably arranged in the recesses 16 of the housing device 8 in a known manner about the central axis 18 of the spindle-shaped region 15, wherein the guide vanes 12, for example via the spindle-shaped regions 15 by an angle between 18 ° and 45 ° relative to the housing means 8 are rotatable ,
  • a platform 19 is provided, which is carried out in an analogous manner as the platform 14 with a spindle-shaped portion 20 and the core flow channel 3 at least partially in the radial Limited R direction of the jet engine 1.
  • the guide vane 12 is in turn mounted on a socket 21 in a housing part 22, a so-called shroud, wherein the guide vane 12 to the Center axis 18 relative to the housing part 22 is rotatably mounted.
  • the housing part 22 is arranged overall in a recess 24 which is formed by two rotor devices 4 which are adjacent to one another in the axial direction A of the jet engine 1 or the stator device 5.
  • the area of the rotor device 4 facing the housing part 22 rotates about the engine axis, whereas the housing part 22 is stationary with respect to the engine axis.
  • Fig. 2a shows the platform 14 and the spindle-shaped portion 14 of a vane 12, wherein it can be seen that the executed with a circular cross-section platform 14 in the likewise circular and concentric with the central axis 18 recess 16 is mounted.
  • a gap 28 surrounding the central axis 18 in the radial direction r which extends outward from the surface 27 of the core flow channel 3 in the axial direction a of the center axis 18 to the outside extends.
  • the housing device 8 has a recess 36 in a mutually facing region of guide vanes 12 adjacent to the central axis in the circumferential direction U, so that the gap 28 is formed in this area by the platforms 14 of the guide vanes 12.
  • the in the Fig. 2a shown embodiment in the following representative of the in the Fig. 2b shown embodiment described.
  • Fig. 2a It can also be seen that a surface 30 of the platform 14 facing away from the core flow channel 3 is spaced apart from the housing device 8 in the radial direction R of the jet engine 1. By this distance and the gap 28, a flow region 31 is formed.
  • a pressure of a working fluid, in this case air, in the region of the high-pressure compressor 2 in the core flow channel 3 in the axial direction A of the jet engine 1 increases in the flow direction, so that a pressure of flowing through the core flow channel 3 main flow on a downstream pressure side
  • a portion of the main flow flows as leakage flow from the pressure side 33 of the blade 13 through the flow area 31 to the suction side 34 of the airfoil 13.
  • the leakage flow is here in the region of the pressure side 33 through the gap 28 via the core flow channel 3 facing away from the surface 30 to the gap 28 in the region of the suction side 34.
  • the occurring during operation leakage flow is in the Fig. 2a and Fig. 2b shown by flow lines 38, wherein in the present case only the flow lines 38 emerging from the gap 28 through the region 39 are shown.
  • a suction device 40 is provided which directly adjoins the flow region 31 in the region of a side facing the pressure side 33 of the airfoil 13.
  • the suction device 40 opens in a transition region in the flow region 31, in which the gap 28 is connected to the surface 30 of the platform 14.
  • the suction device 40 in the present case forms a channel which has an angle of approximately 45 ° with respect to the radial direction R and the axial direction A of the jet engine 1.
  • the suction device 40 opens with a flow region 31 remote end into a space 42 and a plenum, wherein the space 42 is separated by the housing means 8 of the core flow channel 3.
  • the space 42 is arranged in the radial direction R of the jet engine 1 outside the core flow channel 3 and in the axial direction A of the jet engine 1 downstream of the guide vanes 12.
  • a cross section of the channel-shaped suction device 40 in the present case increases continuously in the direction of the space 42.
  • the suction device 40 extends completely circumferentially in the circumferential direction U of the jet engine 1, so that the flow regions 31 of all guide vanes 12 of the stator device 5 are connected to one another and to the space 42 via the suction device 40.
  • this is also apparent from another perspective, in which representation the guide vanes 12 are not shown.
  • the suction device 40 is connected to the flow region 31 over an angular range of approximately 45 °, so that a connection of the suction device 40 to the flow region 31 occurs even if the guide blade 12 is in the respective end position is ensured.
  • the execution according to Fig. 6 differs from the design according to FIG. 3 to FIG. 5 in that, downstream of the gap 28, the housing device 8 has a recess 43 in the area of the pressure side 33 of the blade leaf 13, via which bleed air is additionally taken from the main flow. The bleed air taken over the recess 43 is also supplied to the space 42.
  • FIG. 7 An alternative to the suction device 40 design is in the Fig. 7 shown with the suction device 44.
  • the suction device 44 differs from the suction device 40 in that the suction device 44 directly adjoins the core flow channel 3 in the region of the gap 28 on the pressure side 33 of the blade 13, wherein the suction device 44, starting from the core flow channel 3, again essentially has an angle extends from 45 ° both relative to the radial direction R and the axial direction A of the jet engine 1 in the housing means and similar to the suction device 40 opens into the space 42.
  • a cross section of the channel-shaped suction device 44 expands continuously, starting from the flow region 31.
  • Fig. 9 extends the suction device 48, starting from the gap 28 in the region of the pressure side 33 of the blade 13 in the radial direction R of the jet engine 1 to the outside and opens directly into the space 42.
  • the Extraction means 48 not circumferentially U of the jet engine 1 circulating, but is arranged substantially concentric with the central axis 18 of the spindle-shaped portion 15 and extends over an angular range of for example 45 ° about the central axis 18.
  • each stator blade 12 of the stator device 5 associated with a separate suction device 48, which are each connected to the space 42.
  • suction devices 50, 52, 54 In the 10 to FIG. 12 Further embodiments of suction devices 50, 52, 54 are shown, with the suction devices 50, 52, 54 described in greater detail below, in contrast to the suction devices 40, 44, 46, 48, being connected to the flow region 31 in the area of the suction side 34 of the blade 13 , Otherwise, the suction devices 50, 52, 54 may be designed essentially comparable to the suction devices 40, 44, 46, 48.
  • the suction device 50 is adjacent to the flow region 31 in a region facing away from the core flow channel 3 and, starting therefrom, extends with a curvature in the radial direction R of the jet engine 1 to the outside to a space 51 which is separated from the core flow channel 3 by the housing device 8 analogously to the space 42 , However, in the axial direction A of the jet engine 1 upstream of the guide vane 12 is arranged.
  • the channel-shaped suction device 50 extends circumferentially with respect to the central axis of the jet engine 1 in the circumferential direction U, wherein a cross section of the suction device 50, starting from the flow region 31 to a mouth region in the space 51 is substantially constant.
  • suction device 52 is again channel-shaped and extends the gap 28 in the radial direction R of the jet engine 1 to the outside, wherein the suction device 52 opens via a curvature with respect to the axial direction A of the jet engine 1 in the flow direction of the main flow in the space 51.
  • the suction device 52 opens via a curvature with respect to the axial direction A of the jet engine 1 in the flow direction of the main flow in the space 51.
  • the execution according to Fig. 11 are distributed in the circumferential direction U of the jet engine 1 more webs 55 are provided, via which parts of the housing means 8 are connected to each other for stability reasons, wherein in the present case in each case in a region between two vanes 12, a web 55 is arranged.
  • the suction device 54 of the FIGS. 12 to 14 is performed substantially similar to the suction device 52.
  • Fig. 13 are In this case, webs 56 are again provided, which have a greater extent in the circumferential direction U of the jet engine 1 than the webs 55.
  • the webs 55, 56 depending on the application in the circumferential direction U of the jet engine 1 can also extend over a larger or a smaller area.
  • the suction device 54 is connected to the flow region 31 with respect to the central axis 18 over an angular range of about 30 °.
  • the mass flow taken in the operation of the jet engine 1 via the respective suction device 40, 44, 46, 48, 50, 52, 54 of the leakage flow can in principle be used for various applications, the mass flow being used analogously to the bleed air taken from the main flow in a conventional manner can.
  • Fig. 15 and Fig. 16 the mass flows taken from the leakage flow are discharged via a suction device 52 Fig. 11 a line region 57 which extends substantially in the axial direction A of the jet engine 1 upstream of a rotor device 4 directly upstream of the stator 5.
  • the line region 57 is formed by a part 60 of the housing device 8 forming a surface 27 of the core current channel 3 and a part 61 of the housing device 8, which in the present case has a substantially plate-shaped configuration and is here connected to support elements 62, 63 of the housing device 8.
  • the line region 57 runs circumferentially in the circumferential direction U of the jet engine 1.
  • the mass flow routed via the line region 57 is conducted into the mainstream via a nozzle 58 extending circumferentially in the circumferential direction U of the jet engine 1 in the region of rotor tips 59 of the rotor blade devices 9.
  • a nozzle 58 extending circumferentially in the circumferential direction U of the jet engine 1 in the region of rotor tips 59 of the rotor blade devices 9.
  • Fig. 16 a section of the stage 6C of the high-pressure compressor 2 is shown, that of the Fig. 15 equivalent.
  • Fig. 16 differs from Fig. 15 merely by the embodiment of the line region 65, which essentially corresponds to the line region 57 of FIG Fig. 15 corresponds, in contrast to the line region 57, however, in a rotor tip 59 facing end portion having a plurality of nozzles 66 which are arranged distributed over the circumference of the line portion 65.
  • FIG. 17 and Fig. 18 Another alternative is in Fig. 17 and Fig. 18 shown, wherein the housing device 8 is not shown here.
  • a plurality of structurally identical line regions 68 are provided, via which a mass flow drawn off from the flow region 31 can be supplied to the rotor tips 57 of the rotor blade devices 9.
  • Fig. 19 a highly simplified section of the high-pressure compressor 2 of the jet engine 1 is shown, wherein it is apparent that in the axial direction A of the jet engine 1 present two spaces 70, 71 are provided, which are interconnected via a line region 72.
  • the spaces 70, 71 can be supplied via the suction devices 40, 44, 46, 48, 50, 52, 54 in each case one of the leakage flow removed mass flow in the manner described in more detail above.
  • any spaces 70, 71 can be connected to one another in such a way and a mass flow from one of these spaces 70 can be supplied to a desired place of use.
  • the leakage flow can be conveyed upstream via a suction device 40, 44, 46, 48, 50, 52, 54 in the axial direction A of the jet engine 1 and, for example, fed to a rotor device 4 in a manner described above.
  • FIG. 20 to FIG. 28 shown further possibilities, wherein the embodiments shown alone or in addition to the suction devices 40, 44, 46, 48, 50, 52, 54 may be provided.
  • the vane 12 through the central axis 18 is both the platform 14 with the spindle-shaped portion 15 in a radial direction R of the jet engine 1 outer region of the core flow channel 3 and the platform 19 and the spindle-shaped portion 20 in a radial direction R of the jet engine.
  • 1 inner region of the core flow channel 3 and the platforms 14 and 15 connecting blade 13 can be seen.
  • a leading edge 74 of the airfoil 13 facing upstream in the axial direction A of the jet engine 1 and a trailing edge 75 of the airfoil 13 facing downstream in the axial direction A of the jet engine 1 are shown.
  • the platforms 14, 19 have a front edge region 77 or 79 facing upstream in the axial direction A of the jet engine 1 with a front end 81 or 83 and a rear edge region 78 or 80 arranged downstream in the axial direction A of the jet engine 1 with a rear end region End 82 and 84, wherein the flow area 31 formed by the gap 28 and the distance of the surface 30 of the platform 14 and 19 in the radial direction R of the jet engine 1 of the housing device 8 and the housing part 22 can be seen.
  • arrows 94 the flow direction of the leakage flow is shown both in the region of the platform 14 and in the region of the platform 19.
  • reference points 86 and 88 are respectively upstream of the platform 14 and 19 and reference points 87 and 89 downstream of the platform 14 and 19, respectively, the reference points 86 and 88 being at a distance upstream of the front end 81 and 83, respectively % of an extension of the platform 14 and 19 in the axial direction A of the jet engine 1 amounts to.
  • the reference points 87 and 89 have a downstream distance from the rear end 82 and 84, respectively Platform 14 and 19, which corresponds to about 10% of the extension of the platform 14 and 19 in the axial direction A of the jet engine 1.
  • the reference points 86 to 89 are respectively arranged on the surface 27 of the core flow channel 3.
  • the reference numeral 91 denotes a straight-line connection of the reference points 86 and 87 of the platform 14 and the reference numeral 92 denotes a straight-line connection of the reference points 88 and 89 of the platform 19.
  • a detail representing the front edge region 77 of the platform 14 is provided with the reference symbol I, whereas a detail encompassing the rear edge region 78 of the platform 14 is designated by the reference symbol II.
  • a cutout comprising the front edge region 79 of the platform 19 is designated by III and a cutout comprising the rear edge region 80 of the platform 19 is designated by IV.
  • Fig. 20 to Fig. 23 extends the front edge portion 77 of the platform 14 in the radial direction R of the jet engine 1 with respect to the connection 91 by an extension 93 in the core flow channel 3, whereas the rear edge portion 78 of the platform 14 in the radial direction R of the jet engine 1 does not protrude into the core flow channel 3, but is arranged substantially in alignment with the housing device 8.
  • the edge region 77 extending into the core flow channel 3 may extend in the circumferential direction u of the central axis 18 over an angular range of, for example, 20 ° to approximately 180 °, in particular a flowing transition between the front edge region 77 which extends into the core flow channel 3 and the rear edge region 78 of the platform 14, which does not extend into the core flow channel 3, is provided.
  • a sharp edge 96 is provided in the region of the front end 81 of the platform 14 in a transition region between a substantially extending in the radial direction R of the jet engine 1 side surface 97 and the core flow channel 3 facing surface 98. From a manufacturing point of view, the edge 96 can be provided with a small radius.
  • FIG. 21 is one of the execution of Fig. 20 corresponding variant shown, which only of the areas I and III of Fig. 20 different.
  • a larger radius 99 is provided in the transition region between the side surface 97 extending essentially in the radial direction R of the jet engine 1 and the surface 98 facing the core flow channel 3.
  • the section III of the Fig. 20 corresponding area of the platform 19 is substantially horizontally mirrored to the cutout I 'formed.
  • the section I is alternative to the embodiment according to Fig. 20 designed, wherein in the transition region between the substantially radially in the radial direction R of the jet engine 1 extending side surface 97 and the core flow channel 3 facing surface 98 as a Nose 100 executed overhang is arranged.
  • the platform 14 in the region of the surface 98 has a greater extent in the axial direction A of the jet engine 1 than in the region of the side surface 97.
  • a front end 81 of the platform 14 in the region of the nose 100 is in this case with respect to the axial direction A of the jet engine. 1 approximately at the level of a gap 28 limiting side wall 101 of the housing device 8 is arranged.
  • the embodiment of the lug 100 described above ensures that the leakage flow emerging from the gap 28 in the region of the suction side 33 of the blade 13 does not enter the main flow directly in the radial direction R of the jet engine 1, but is previously dammed up in the region of the nose 100.
  • a static pressure in this area is further increased with the advantageous effects for the leakage flow described in more detail above.
  • the leaking from the gap 28 leakage flow is deflected before being introduced into the main flow around the nose 100 around and accelerated, so that the leakage flow interacts only to an advantageously small extent with the main flow.
  • the section III of the Fig. 20 corresponding area of the platform 19 is also in the design according to Fig. 22 essentially horizontally mirrored to the cutout I "formed.
  • the platform 14 in the front edge region 77 again has an overhang designed as a nose 102, as an alternative to the region I of FIG Fig. 20 designed cutout I '"in Fig. 23 shows.
  • the nose 102 is executed in principle comparable to the nose 100.
  • the nose 102 has a greater extent with respect to the axial direction A of the jet engine 1, so that the nose 102 engages over the housing device 8 with respect to the side surface 101 in the axial direction A of the jet engine 1 by a length 107.
  • a greater increase in the static pressure is achieved in the region of the gap 28 than with the nose 100.
  • the section III of the Fig. 20 corresponding area of the platform 19 substantially horizontally mirrored to the cutout I '"formed.
  • Fig. 24 is an alternative embodiment of section I with I “" Fig. 20 shown, in which the front edge region 77 of the platform 14 in the radial direction R of the jet engine 1 has a substantially corresponding to the connection 91 in this area extent.
  • the execution I "" according to Fig. 24 is with the in the FIGS. 25 to 28 shown variants II 'to II "" of the rear edge portion 78 of the platform 14 combined.
  • the front edge region 79 of the platform 19 corresponding to the cutout III in Fig. 20 horizontally mirrored to the front edge region 77 of the platform 14 executed.
  • the rear edge region 80 of the platform 19 corresponding to the section IV in Fig. 20 horizontally mirrored to the rear edge region 78 of the platform 14.
  • sections II 'to II "" of the FIGS. 25 to 28 extends the rear edge portion 78 of the platform 14 in the radial direction R of the jet engine 1 relative to the connection 91 by an extension 110 in the core flow channel 3.
  • the extending into the core flow channel 3 edge region can in the circumferential direction u of the central axis 18 in turn over an angular range of, for example 20 ° to about 180 °, wherein in particular a smooth transition between the rear edge region 78 comprehensive edge region extending into the core flow channel 3, and a front edge region 77 comprising the edge region of the platform 14, which does not extend into the core flow channel 3 , is provided.
  • section II 'according to Fig. 25 is in a transition region between the substantially radially in the radial direction R of the jet engine 1 extending side surface 97 and the core flow channel 3 facing surface 98, a sharp edge 103 is provided, which may be provided in particular from a manufacturing point of view with a small radius.
  • the respective nose 105 and 106 substantially mirror-symmetrical or vertically mirrored to the nose 100 and the nose 102 is executed.
  • the platforms 14, 19 in the cutouts I, II, III and IV can form any combination of the respective embodiments described in each case.
  • the transition region from the side surface 97 of the platform 14 to the surface 98 of the platform 14 is designed in particular in the front edge region 77 or 79, comparable to the rear edge region 78 or 80.
  • the edge region 77, 78 of the platform 14 projecting into the core flow channel 3 or the edge region 79, 80 of the platform 19 projecting into the core flow channel 3 is preferably executed completely circumferentially in the circumferential direction u of the central axis 18.
  • the transition region from the side surface 97 to the surface 98 of the platform but also in the front edge region 77 or 79 may be designed differently than in the rear edge region 78 or 80.
  • the extension 93 in the front edge region 77 or 79 and the extension 110 in the rear edge region 78 or 80 can both have a mutually corresponding value.
  • one of the extensions 93 and 110 may be larger than the other extension 110 and 93, respectively.
  • the extension 93 or the extension 110 of the edge region 77, 78, 79, 80 projecting into the core flow channel 3 amounts to approximately 0.8% of a width B in the present case of the core flow channel 3 perpendicular to the axial direction A of the jet engine 1 in the region of the edge region 77, 78, 79, 80.
  • a suction device 108 may adjoin the flow region 31 in the region of the pressure side 33 of the blade 13 and / or in the region of the suction side 34 of the blade.
  • suction devices 40, 44, 46, 48, 50, 52 or 54 can be combined.
  • Fig. 3 to Fig. 18 shown platforms 14 and 19 according to the embodiments according to FIGS. 20 to 30 protrude into the core flow channel 3.
  • a suction device 40, 44, 46, 48, 50, 52, 54, 108 By combining the projecting into the core flow channel 3 platform 14, with a suction device 40, 44, 46, 48, 50, 52, 54, 108, an efficiency of the jet engine 1 is advantageously further increased, since the effect of the suction device 40, 44, 46, 48, 50, 52, 54, 108 with the effect of projecting into the core flow channel platform 14, 19 added.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP16175990.7A 2015-06-25 2016-06-23 Dispositif de stator pour une turbomachine comprenant un dispositif de carter et plusieurs aubes directrices Withdrawn EP3109408A1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102015110250.1A DE102015110250A1 (de) 2015-06-25 2015-06-25 Statorvorrichtung für eine Strömungsmaschine mit einer Gehäuseeinrichtung und mehreren Leitschaufeln

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EP3109408A1 true EP3109408A1 (fr) 2016-12-28

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Publication number Priority date Publication date Assignee Title
US10227930B2 (en) * 2016-03-28 2019-03-12 General Electric Company Compressor bleed systems in turbomachines and methods of extracting compressor airflow
DE102019217394A1 (de) * 2019-11-11 2021-05-12 MTU Aero Engines AG Leitschaufelanordnung für eine strömungsmaschine
FR3109959B1 (fr) * 2020-05-06 2022-04-22 Safran Helicopter Engines Compresseur de turbomachine comportant une paroi fixe pourvue d’un traitement de forme

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GB2234299A (en) * 1989-07-06 1991-01-30 Rolls Royce Plc Support arrangement for nozzle guide vane
EP1010862A2 (fr) * 1998-12-16 2000-06-21 General Electric Company Rondelle et joint pour une aube variable
US6210106B1 (en) * 1999-04-30 2001-04-03 General Electric Company Seal apparatus for gas turbine engine variable vane
EP1528226A2 (fr) * 2003-10-29 2005-05-04 United Technologies Corporation Disque de palier en électro-graphite pour une aube variable
EP2113637A2 (fr) * 2008-04-30 2009-11-04 Rolls-Royce Deutschland Ltd & Co KG Unité rotative pour un compresseur axial

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US4711084A (en) * 1981-11-05 1987-12-08 Avco Corporation Ejector assisted compressor bleed
GB2210935B (en) * 1987-10-10 1992-05-27 Rolls Royce Plc Variable stator vane assembly
US6283705B1 (en) * 1999-02-26 2001-09-04 Allison Advanced Development Company Variable vane with winglet
FR2899637B1 (fr) * 2006-04-06 2010-10-08 Snecma Aube de stator a calage variable de turbomachine
US7806652B2 (en) * 2007-04-10 2010-10-05 United Technologies Corporation Turbine engine variable stator vane
FR2916815B1 (fr) * 2007-05-30 2017-02-24 Snecma Compresseur a reinjection d'air
EP2058524A1 (fr) * 2007-11-12 2009-05-13 Siemens Aktiengesellschaft Compresseur à purge d'air doté de conduits dans les aubes variables
DE102008015207A1 (de) * 2008-03-20 2009-09-24 Rolls-Royce Deutschland Ltd & Co Kg Fluid-Injektor-Düse
DE102008019603A1 (de) * 2008-04-18 2009-10-22 Rolls-Royce Deutschland Ltd & Co Kg Strömungsmaschine mit schaufelreiheninterner Fluid-Rückführung
FR2933148B1 (fr) * 2008-06-25 2010-08-20 Snecma Compresseur de turbomachine
FR2933149B1 (fr) * 2008-06-25 2010-08-20 Snecma Injection d'air dans la veine d'un compresseur de turbomachine

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Publication number Priority date Publication date Assignee Title
GB743782A (en) * 1952-11-07 1956-01-25 Power Jets Res & Dev Ltd Nozzle assemblies for bladed fluid flow machines such as turbines and compressors
GB2234299A (en) * 1989-07-06 1991-01-30 Rolls Royce Plc Support arrangement for nozzle guide vane
EP1010862A2 (fr) * 1998-12-16 2000-06-21 General Electric Company Rondelle et joint pour une aube variable
US6210106B1 (en) * 1999-04-30 2001-04-03 General Electric Company Seal apparatus for gas turbine engine variable vane
EP1528226A2 (fr) * 2003-10-29 2005-05-04 United Technologies Corporation Disque de palier en électro-graphite pour une aube variable
EP2113637A2 (fr) * 2008-04-30 2009-11-04 Rolls-Royce Deutschland Ltd & Co KG Unité rotative pour un compresseur axial

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DE102015110250A1 (de) 2016-12-29

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