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EP2846000A2 - Roue statorique d'une turbine à gaz - Google Patents

Roue statorique d'une turbine à gaz Download PDF

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Publication number
EP2846000A2
EP2846000A2 EP14184094.2A EP14184094A EP2846000A2 EP 2846000 A2 EP2846000 A2 EP 2846000A2 EP 14184094 A EP14184094 A EP 14184094A EP 2846000 A2 EP2846000 A2 EP 2846000A2
Authority
EP
European Patent Office
Prior art keywords
vane
pressure side
section
turbine
passage
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.)
Granted
Application number
EP14184094.2A
Other languages
German (de)
English (en)
Other versions
EP2846000A3 (fr
EP2846000B1 (fr
Inventor
Lars Willer
Knut Lehmann
Philipp Amtsfeld
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 EP2846000A2 publication Critical patent/EP2846000A2/fr
Publication of EP2846000A3 publication Critical patent/EP2846000A3/fr
Application granted granted Critical
Publication of EP2846000B1 publication Critical patent/EP2846000B1/fr
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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
    • 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
    • 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
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/122—Fluid guiding means, e.g. vanes related to the trailing edge of a stator vane
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/12—Fluid guiding means, e.g. vanes
    • F05D2240/123—Fluid guiding means, e.g. vanes related to the pressure side of a stator vane
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/304—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the trailing edge of a rotor blade
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/305—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the pressure side of a rotor blade
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F05D2250/711—Shape curved convex
    • 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
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F05D2250/713—Shape curved inflexed

Definitions

  • the invention relates to a Turbinenleitrad, in particular a Hochdruckturbinenleitrad a gas turbine, in particular for use in a gas turbine engine.
  • the stator is to be understood in this context that During operation of the gas turbine, in particular the trailing edge of the first stator of the high-pressure turbine can burn off under the extreme thermal loads that occur. This means that the vane, starting from the blade trailing edge, is shortened by the burn-back. Since the first stator of a high-pressure turbine determines the flow through the entire turbomachine, maintaining the flow (capacity) of the first stator is crucial so that the entire turbomachine and all individual components can continue to operate at a nominal mass flow at the design point. It is thus required that the flow rate (capacity) of the turbine does not change significantly due to the burn-back.
  • the passage cross-section within the stator must remain approximately constant upstream of the throat cross-section (ie in the direction of the progressive burn-back of the blade trailing edge), so that the then effective passage cross-section also remains approximately constant even in the case of burn-back of the thermally highly loaded trailing edge. This ensures that the flow remains similar even in the event of burnback.
  • Such an embodiment is for example from the FIG. 4 of the DE 10 2005 025 213 A1 known.
  • the invention has for its object to provide a Turbinenleitrad of the type mentioned, which has a high efficiency with a simple structure and simple design, while at the same time the mentioned burn-back criterion is met. Especially in the case of a burn-back, the turbine capacity should remain largely unchanged, so that the entire engine with its individual components can continue to be operated at the design point. According to the invention the object is achieved by the feature combination of claim 1.
  • the subclaims show advantageous embodiments of the invention.
  • the solution according to the invention then considers a turbine nozzle in which two adjacent guide vanes each form a passage which comprises a constant passage section.
  • the constant passage section is characterized in that it has a substantially constant passage cross-section.
  • the constant passage portion has an entrance area in the constant passage portion and an exit area. The exit area is located at the blade trailing edge and is typically identical to the narrowest section (throat section) of the passage.
  • Each vane forms a rearward pressure side extending from the blade trailing edge adjacent to the constant passage portion to the entry portion of the passage portion, and a forward portion extending upstream of the trailing portion. The rear region is thus the region of the pressure side of the guide vane which delimits the constant passage section.
  • the guide vanes on the pressure side have a convex pressure side contour, which produces a transition from the rear region of the guide vane to the front region of the vane.
  • the inventive solution provides a convex pressure side contour on the pressure side of the vane such that a transition from a rear portion of the vane in which there is a constant passage portion to the front portion of the vane is established by the convex pressure side contour.
  • the rear portion of the vane is thus connected to the front portion of the vane via the convex pressure side contour.
  • the convex pressure side contour or the convex curvature of the pressure side provided by this, it is possible to make the passage between two guide vanes constant over a certain length, even if the adjacent guide vane is provided on the suction side with a considerable convex curvature for realizing a loss-optimized turbine nozzle. without compensation the convex pressure side contour - would lead to a significant passage widening.
  • the invention thus ensures a burn-back capability even in the event that a loss-optimized turbine nozzle is provided with guide vanes with considerable convex curvature of the suction side in the area of the throat cross-section.
  • the solution according to the invention thus provides that the wall of the pressure side of the Guide vane convex pressure side contour, that is, forms a convex curvature, which forms the transition between the rear portion of the vane, which is adjacent to the constant passage portion, and the front area extending upstream thereto.
  • the invention produces by a convex contouring of the pressure side of the guide vane of the stator, the remindbrandrati without the aerodynamic design of the suction side of the guide vane is affected. According to the invention, it is therefore possible to freely define the suction side of the guide vane of the guide wheel and to optimally shape the loss while realizing guide vanes with considerable convex curvature of the suction side in the area of the throat cross section or adjacent to the throat cross section.
  • the configuration of the pressure-side contour of the guide vane according to the invention ensures that during a burn-back the cross section of the passage between adjacent guide vanes remains substantially constant, so that the flow (capacity) of the turbine and thus the efficiency of the entire engine by a burn-back not or only slightly to be influenced.
  • the profile thickness of the guide vanes in the direction of the blade trailing edge in front of the rear region of the guide vanes increases or is constant or decreases to a lesser extent than in the rear region of the guide vane.
  • this embodiment provides that the profile thickness increases in the direction of the blade trailing edge in front of the entry region into the passage or is constant or decreases to a lesser extent than in the region of the contant passage section.
  • the convex pressure-side contour forms a maximum at or upstream of the entry region into the constant passage cross-section. Furthermore, it can be provided that the convex pressure-side contour forms a maximum of the curvature at or upstream of the entry region into the constant passage cross-section. The maximum of the curvature is close to the locally furthest from the pressure side point or near the locally furthest from the pressure side protruding line of the pressure side contour.
  • the maximum and / or the maximum of the curvature are thus not in the rear region of the guide vane, but in the front region of the guide vane, but preferably at a short distance from the rear region (eg at a distance corresponding to a maximum of 10% of the length of the skeleton line ) or directly at the junction of the two areas.
  • a further embodiment of the invention provides that the convex pressure side contour on the pressure side of the guide vanes is formed predominantly or completely in the front region of the guide vane. It can be provided that a part of the convex pressure side contour is additionally formed in the rear region of the guide vane. Basically, in the rear region of the vane, which limits the constant passage portion, but also a straight or even concave curvature may be provided, which merges into the convex printing side contour.
  • a substantially constant passage cross section is present, for example, if the passage cross section does not deviate more than 20% from the throat cross section in the region of the blade trailing edge. Preferably, this deviation from the narrow section is less and is less than 10%, 5% or 2% of the throat section.
  • the passage cross section in the constant passage section is exactly constant. It can further be provided that the constant passage section extends over a chord length which, for example, in the region between 5% and 40% of the total chord, for example, about 5%, 10%, 15%, 20%, 25%, 30%, 35% or 40% of the total chord.
  • the convex pressure side contour extends over the entire height of the guide blade. Furthermore, it can be provided that the pressure-side contour extends at least over a partial region of the blade height (for example over at least 50% or at least 70% of the blade height). It is also possible that the configuration of the curvature varies over the blade height.
  • the guide blade starting from the blade trailing edge, is subsequently provided with a concave region on the convex region.
  • This embodiment leads in particular to an optimal surface pressure distribution on the blade surface.
  • the cooling air consumption turbine turbine according to the invention has considerable advantages. Since the blade contour in the trailing edge region has a greater thickness, it is possible to expand the internal cooling geometry further in the direction of the blade trailing edge. This can be done, for example, by pedestal banks located further back. This gives the possibility Save cooling air, since the difficult to cool and thermally highly loaded rear edge overhang can be reduced in length.
  • Engine development costs can be reduced because of the secure capacity prediction, since the need for subsequent capacity change is reduced.
  • the engine development time can also be shortened.
  • the gas turbine engine 10 is a generalized example of a turbomachine, in which the invention can be applied.
  • the engine 10 is formed in a conventional manner and comprises in succession an air inlet 11, a fan 12 circulating in a housing, a medium pressure compressor 13, a high pressure compressor 14, a combustion chamber 15, a high pressure turbine 16, a medium pressure turbine 17 and a low pressure turbine 18 and a Exhaust nozzle 19, which are all arranged around a central engine axis 1.
  • the medium pressure compressor 13 and the high pressure compressor 14 each include a plurality of stages, each of which includes a circumferentially extending array of fixed stationary vanes 20, commonly referred to as stator vanes, extending radially inward from the core engine casing 21 in an annular flow passage through the compressors 13, 14 protrude.
  • the compressors further include an array of compressor blades 22 projecting radially outwardly from a rotatable drum or disc 26 coupled to hubs 27 of high pressure turbine 16 and mid pressure turbine 17, respectively.
  • the turbine sections 16, 17, 18 have similar stages including an array of fixed vanes 23 projecting radially inward from the housing 21 into the annular flow passage through the turbines 16, 17, 18 and a downstream array of turbine blades 24 projecting outwardly from a rotatable hub 27.
  • the compressor drum or compressor disk 26 and the blades 22 arranged thereon and the turbine rotor hub 27 and the Turbine blades 24 disposed thereon rotate about the engine axis 1 during operation.
  • the Fig. 2 shows a view of a known from the prior art Turbinenleitrads frontal view of adjacent vanes 23. These each have a pressure side 30 and a suction side 31 and form a passage 29 through which the exiting the combustion chamber hot gases flow. From the representation of Fig. 2 shows that in the region of a blade trailing edge 32, the passage 29 has a narrowest cross section (narrow section 36). This is formed with regard to the desired profile shape of the guide vanes 23. Due to the thermal load during operation, the area of the blade trailing edge 32 is burned off, so that a burn-back 35 results. This means that the hatched area of the blade profile burns off. This results in an effective passage cross-section 37, which is significantly widened compared to the narrow section 36 and consequently leads to a significant reduction in the efficiency. With the widening of the passage cross-section, the flow and the capacity change.
  • FIG. 3 shows a view of an embodiment of the invention.
  • the guide vanes 23 in turn have a pressure side 30 and a suction side 31, with two adjacent vanes 23 forming a passage 29 between the suction side 31 of one vane and the pressure side 30 of the other vane starting from the vane trailing edge 32 through the hot gases leaving the combustion chamber stream.
  • the passage 29 comprises a constant passage portion 29a, in which the passage 29 has a substantially constant passage cross-section 37.
  • the constant passage portion 29a has an entrance portion 38 and an exit portion 36 having substantially the same passage cross section.
  • the outlet region 36 is delimited by the blade trailing edge 32, so that the outlet region 36 corresponds to the narrow cross section of the passage 29.
  • passage cross section 37 in the constant passage section 29a is substantially constant means that the deviation of the passage cross section 37 from the throat section in this constant passage section 29a is below a defined value which is defined, for example, as 20% of the throat cross section.
  • a constant passage cross-section 29a can be defined, for example, in that the deviation from the throat section is below 15%, 10% or 5% of the throat cross-section.
  • the vane 23 further defines on the pressure side a rear portion 320 extending from the blade trailing edge 32 adjacent to the constant passage portion 29a to the entrance portion 38 of the constant passage portion 29a.
  • the pressure-side rear region 320 of the guide vane is thus that region which delimits the constant passage section 29 a on the pressure side. Upstream of the rear region 320 extends a front region 310, which in principle extends to the blade leading edge, but for the purposes of the present invention only the region adjacent to the rear region 320 is considered in detail.
  • the vane 23 also has on the pressure side 30 a convex pressure side contour 33, which creates a transition from the rear region 320 to the front region 310.
  • This means that the convex pressure side contour 33 is formed in the transition region between the two regions 310 and 320, wherein it may extend exclusively in the front region 310 or alternatively over both regions 310, 320.
  • the convex pressure side contour 33 has a maximum M, which in the cross-sectional view of FIG. 3 indicates the point at which the curvature provided by the convex pressure side contour 33 protrudes locally most strongly from the pressure side 30.
  • the profile thickness in front of the rear region 320 does not increase or is constant, but decreases to a lesser extent (ie by a smaller value per unit length) than then in the rear region 320.
  • This course of the profile thickness d corresponds with the realization of a maximum M of the curvature provided by the convex pressure side contour 33 in front of or at the entrance area into the constant passage section 29a.
  • a convex pressure-side contour 33 leads to an increase in the wedge angle between the surfaces of the pressure side 30 and the suction side 31 in the adjacent region to the blade trailing edge 32 and in particular to avoid widening of the passage cross-section in the event of burn-back.
  • Such a widening is avoided precisely by providing a constant passage section 29a through the solution according to the invention, so that the throat section does not change in the case of burn-back 35 in the region of this constant passage section 29a.
  • a burn-back 35 is in the FIG. 3 marked greatly exaggerated in order to better explain the effectiveness of the invention can. It follows that the passage cross-section 37 in the constant passage cross-section 29a remains substantially the same in the event of burn-back, since the throat cross-section 36 in this section is substantially equal to the passage cross-section 37.
  • FIG. 4 shows a further inventive embodiment of two vanes 23 of a Turbinenleitrads.
  • the embodiment corresponds to the embodiment of FIG. 3 which is referred to with reference to the reference numerals used.
  • the rear region 320 of the vane 23 is given a shape that allows a constant passage portion 29a with a substantially constant passage cross section 29a between an entry region 38 and to provide an exit region 36 of this constant passage portion 29a.
  • the corresponding curvature of the convex pressure side contour 33 causes the profile thickness d of the guide vane 23 to rise or remain substantially constant in front of the rear region 320 and to decrease sharply only in the rear region 320 of the vane (see profile thickness d1, d2 and d3 of FIG FIG. 4 ).
  • a difference in the design of the FIG. 4 opposite to the embodiment of FIG. 3 consists in the curvature of the pressure side 30 of the vane in the rear region 320. While this curvature in the FIG. 3 is at least approximately concave, it is in the embodiment of FIG. 4 formed convex, so that the rear portion 320 forms a portion of the convex pressure side contour 33 and contributes to the latter. However, the maximum M of the convex printing side contour 33 is located in front of the constant passage portion 29a in the front portion 310.
  • the convex pressure side contour 33 makes the transition from the rear portion 320 of the vane to the front portion 310 of the vane.
  • line 40 indicates the profile of the pressure side of a prior art vane, with the wall of the vane 23 adjacent the blade trailing edge 32 being substantially straight or of slight uniform curvature.
  • the line 40 thus illustrates the pressure side contour of a conventional vane.
  • a thickening 50 is also in the embodiment of FIG. 3 in front. This has in the embodiment of FIG. 3 but a different shape and does not run is generally convex, but also has a convex portion in the transition from the rear portion 320 to the front portion 310. In the embodiment of FIG. 4 the thickening 50 is completely formed by the convex printing side contour 33.
  • FIG. 4 Another special feature of the design of the FIG. 4 is that a concave portion 34 is provided adjacent to the convex portion 33 on the pressure side 30 of the vane 23. This leads to a further optimization of the surface pressure distribution on the blade surface.
  • FIG. 5 shows a comparison of the configuration according to the prior art, as this in the FIG. 2 is shown (left half of FIG. 5 ) and an embodiment of the invention according to the FIG. 4 ,
  • inventively provided contouring of the pressure side 30 the advantages described above arise. This is in particular also from the comparative representation of the static surface pressures according to FIG. 6 seen.
  • the standardized chord length of 0.0 corresponds to the position of the blade leading edge
  • the normal chord length of 1.0 corresponds to the position of the blade trailing edge.
  • Fig. 6 shows the geometric configuration according to the prior art ( Fig. 5 left) associated surface pressure distribution.
  • the lower half of Fig. 6 shows the embodiment according to the invention ( Fig. 5 right) associated surface distribution. Visible is the inventively resulting advantageous pressure curve on the suction side ( Fig. 6 below), which can be implemented without infringing the burn-back criterion.
  • the S-beat of the pressure curve on the pressure side in the area of the blade trailing edge at the chord length 0.7 to 1.0 results from the contouring of the pressure side according to the invention for compliance with the burnback criterion.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP14184094.2A 2013-09-09 2014-09-09 Roue statorique d'une turbine à gaz Active EP2846000B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE102013217997 2013-09-09

Publications (3)

Publication Number Publication Date
EP2846000A2 true EP2846000A2 (fr) 2015-03-11
EP2846000A3 EP2846000A3 (fr) 2015-04-29
EP2846000B1 EP2846000B1 (fr) 2020-04-08

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EP14184094.2A Active EP2846000B1 (fr) 2013-09-09 2014-09-09 Roue statorique d'une turbine à gaz

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US (1) US9896950B2 (fr)
EP (1) EP2846000B1 (fr)

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EP2987956A1 (fr) * 2014-08-18 2016-02-24 Siemens Aktiengesellschaft Aube de compresseur
CN108757508A (zh) * 2018-05-03 2018-11-06 西北工业大学 一种带有减震凸台导流叶片的压气机
BE1026579B1 (fr) * 2018-08-31 2020-03-30 Safran Aero Boosters Sa Aube a protuberance pour compresseur de turbomachine
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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN109763928A (zh) * 2017-11-09 2019-05-17 株式会社东芝 导流叶片以及流体机械

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