US20140133957A1 - Method of diffusing a gas turbine compression stage, and diffusion stage for implementing same - Google Patents

Method of diffusing a gas turbine compression stage, and diffusion stage for implementing same Download PDF

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Publication number
US20140133957A1
US20140133957A1 US14/126,989 US201214126989A US2014133957A1 US 20140133957 A1 US20140133957 A1 US 20140133957A1 US 201214126989 A US201214126989 A US 201214126989A US 2014133957 A1 US2014133957 A1 US 2014133957A1
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US
United States
Prior art keywords
blades
plates
flow
diffusion
bump
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.)
Abandoned
Application number
US14/126,989
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English (en)
Inventor
Laurent Pierre Tarnowski
Nicolas Bulot
Jerome Yves Felix Gilbert Porodo
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.)
Safran Helicopter Engines SAS
Original Assignee
Turbomeca SA
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 Turbomeca SA filed Critical Turbomeca SA
Assigned to TURBOMECA reassignment TURBOMECA ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: BULOT, Nicolas, PORODO, JEROME YVES, FELIX, GILBERT, TARNOWSKI, LAURENT PIERRE
Publication of US20140133957A1 publication Critical patent/US20140133957A1/en
Assigned to SAFRAN HELICOPTER ENGINES reassignment SAFRAN HELICOPTER ENGINES CHANGE OF NAME (SEE DOCUMENT FOR DETAILS). Assignors: TURBOMECA
Abandoned legal-status Critical Current

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Classifications

    • 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
    • F04D29/42—Casings; Connections of working fluid for radial or helico-centrifugal pumps
    • F04D29/44—Fluid-guiding means, e.g. diffusers
    • F04D29/441—Fluid-guiding means, e.g. diffusers 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
    • F04D19/00—Axial-flow pumps
    • F04D19/002—Axial flow fans
    • 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
    • F04D29/42—Casings; Connections of working fluid for radial or helico-centrifugal pumps
    • F04D29/44—Fluid-guiding means, e.g. diffusers
    • F04D29/441—Fluid-guiding means, e.g. diffusers especially adapted for elastic fluid pumps
    • F04D29/444—Bladed diffusers
    • 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
    • F05D2250/50—Inlet or outlet
    • F05D2250/52—Outlet

Definitions

  • the invention relates to a method for diffusing the flow of air in a compression stage of a gas turbine engine, as well as to a diffusion stage capable of implementing said method.
  • the field of the invention relates to improving the performance levels and the pumping margin of centrifugal and mixed compressors in the diffusion assembly of the relevant stage.
  • the purpose of this diffusion assembly is to convert the kinetic energy of the fluid, obtained at the output of the centrifugal impeller constituting the stage, into static pressure. The operation must occur with a minimum loss of total pressure whilst maintaining a satisfactory level of stability in the compressor in order to maintain a pumping margin that is acceptable for the operation of the turbine engine.
  • a centrifugal compressor has at least one radial compression stage, i.e. which is capable of producing a flow of air perpendicular to the central axis of the compressor.
  • a mixed compressor has at least one compression stage that is inclined relative to said central axis.
  • a diffusion assembly of a compression stage is composed of an impeller formed by two plates, between which the fluid flows in a centrifugal or inclined manner from the centre towards the periphery. Blades are distributed around the impeller between the plates. These blades form a flow cascade between the leading edges of these blades at the centre and the trailing edges on the outside.
  • the plates of the radial and mixed diffusion assemblies are conventionally flat and advantageously the flow cross-sections of the fluid between the blades are tapered.
  • the tapering of the flow cross-sections is defined by the flow cross-section at the neck of the diffuser and by the rate of deceleration between the leading and trailing edges of the cascade.
  • the aim of the invention is to produce such a flow by implementing shape-optimised plates as these plates represent the largest surface area that is “streamed” by the flow.
  • Non-axisymmetric shapes in the direction of the flow and in the tangential direction are thus proposed.
  • the present invention relates to a method for diffusing the flow of air in a compression stage of a gas turbine engine comprising a diffusion assembly composed of an impeller formed by two plates, between which the fluid flows in a centrifugal or inclined manner from the centre towards the periphery. Blades of a cascade are distributed around the impeller between the plates so as to channel the flow of the fluid between the leading edges of these blades at the centre and the trailing edges at the periphery.
  • at least one of the plates has at least one alternation of concave and convex curvatures in at least one of two substantially perpendicular directions, namely in the direction of flow along the blades and in an inter-blade tangential direction.
  • the three-dimensional shape of the stream of the fluid allows its flow in this stream to be redistributed and homogenised: the secondary flows, which generate load losses, are substantially reduced.
  • the position of shocks in the transonic blade assemblies is modified and their intensity is reduced.
  • the aerodynamic locking at the input of the combustion chamber that follows the compression stage is also substantially reduced.
  • the invention further relates to a diffusion stage of a radial or mixed gas turbine engine capable of implementing this method.
  • a diffusion stage of a radial or mixed gas turbine engine capable of implementing this method.
  • Such a stage comprises an impeller formed by two plates, between which the fluid flows in a centrifugal or inclined manner from the centre towards the periphery. Blades of a cascade are distributed around the impeller between the plates and so as to channel the flow of the fluid between the leading edges of these blades at the centre and the trailing edges at the periphery.
  • At least one of the plates has an internal face comprising at least one zone with alternating hollow and bump curvatures between two adjacent blades, in at least one of two substantially perpendicular directions, namely in the direction of flow along the blades and in an inter-blade tangential direction.
  • the diffusion stage has alternating hollow and bump zones between the blades, in particular up to substantially 80% (preferably up to substantially 50%) of a chord line of a blade, at the leading edge of the blades, starting upstream of the leading edge, and/or at the trailing edge, continuing to downstream of the trailing edge.
  • These alternating hollow and bump zones can be applied to one and/or the other of the two centrifugal (radial) and mixed diffusion plates, particularly in a symmetrical manner, relative to a central plane of symmetry of the plates, or in a parallel manner when the two plates of the stage are involved.
  • FIG. 1 is a partial cross-sectional view of a gas turbine engine comprising an air diffuser
  • FIGS. 2 a and 2 b are perspective views of a diffusion stage with blades comprising one and two plates;
  • FIGS. 3 a to 3 c are schematic views of profiles of a plate in the direction of the air flow along a blade, with two zones with curvatures respectively alternating along the blade, at the trailing edge continuing to downstream of the blade, and at the leading edge from upstream of the blade;
  • FIG. 4 is a schematic partial view in the inter-blade tangential direction, with a plate having two zones with alternating curvatures;
  • FIG. 5 is a schematic inter-blade perspective view at the leading edge, with a plate having a zone with alternating curvature;
  • FIG. 6 is a schematic perspective view at the trailing edge of two blades, with a flat plate and a plate having a zone with alternating curvature.
  • downstream and upstream relate to positions in relation to the flow of air.
  • identical reference numerals relate to the passages in the description in which the elements that correspond to these reference numerals are defined.
  • an air flow F is firstly aspirated in a fresh air intake sleeve 2 , then compressed between the vanes 3 of an impeller 4 of a centrifugal compressor 5 and a casing 10 .
  • the turbine has axial symmetry about an axis X′X.
  • the compressor 5 is centrifugal and the compressed flow F then exits radially from the impeller 4 .
  • the flow exits inclined at an angle of between 0° and 90° relative to a radial direction, perpendicular to the axis X′X.
  • the flow F then passes through a diffuser or impeller 6 , disposed at the output of the compressor 4 , in order to be rectified and routed towards intake channels 7 of a combustion chamber 8 .
  • the impeller 6 is composed of a plurality of curved blades 60 that are arranged between two plates 9 at the periphery of the impeller 4 , in a radial manner in this case, and thus rotate about the axis X′X.
  • FIG. 2 b more specifically shows a perspective view of the diffuser 6 with blades 60 rigidly connected to two plates 9 .
  • each blade 60 has, in a known manner, a face 6 e , referred to as an upper face, and a face 6 i , referred to as a lower face.
  • these faces are connected by a tapered leading edge 6 a and a rounded trailing edge 6 f , in the direction of air flow.
  • each blade 60 has flat sides 6 p which are rigidly connected to the plates 9 .
  • the plates 9 of FIGS. 2 a and 2 b are flat. According to the invention, at least one of these plates 9 has, in the space E that is defined between them, at least one zone with alternating curvatures between two blades 60 .
  • such a plate 9 is shown as a profile, in the direction of the air flow F along the blades 60 , from the intake at the leading edge 6 a of the blade to the outlet channel 7 towards the combustion chamber.
  • Two zones Z1 and Z2 with alternating curvatures are produced in the plate 9 , along the blades 60 .
  • Each of these zones has, at the flow F side and in relation to the flat face portion 9 p of the plate 9 , a hollow portion 91 and a bump portion 92 .
  • the zones Z1 and Z2 generally extend, in the non-limiting example shown, over approximately 80% of the length of the chord line 6 c of the blades 60 .
  • the plate 9 has a low thickness in order to simplify the presentation, but in reality it has a certain thickness and the zones with alternating curvatures are formed on the internal face 9 i of the plate where the air flow F flows.
  • the external face 9 e of the plate 9 can remain flat or can also conform to the bump and hollow shapes of the alternating curvatures, which are reversed in this example relative to the hollow and bump shapes of the internal face 9 i .
  • the plate has a variable thickness and, in the second instance, it has a constant thickness.
  • the shape of the plates can depend on the method used to produce the zones with alternating curvatures: by milling, laser, spark erosion, forming, stamping, etc.
  • the two zones Z1 and Z2 with alternating curvatures are respectively formed at the trailing edge 6 f of the blades 60 to downstream of the blades in the intake of the channel 7 , and/or at the leading edge 6 a , from the upstream intake of the blades 60 .
  • FIG. 4 shows a partial schematic view of the diffusion stage in the inter-blade “tangential direction” 6 t , i.e. between two blades 60 .
  • the arrow F indicates the direction of the flow of air.
  • the internal face 9 i of the plate 9 has two zones Z1 and Z2 with alternating curvatures that mainly extend between the lower 6 i and upper 6 e faces of two blades 60 .
  • the perspective inter-blade view of FIG. 5 at the leading edge 6 a more specifically shows the plate 9 with a zone Z1 with an alternating curvature between the two blades 60 .
  • the zone shows the hollow curvature portion 91 , on the lower face 6 i of a blade 60 , and the bump curvature portion 92 , on the upper face 6 e of the other blade 60 .
  • This alternation of curvatures allows the pressures between the lower face and the upper face of each blade 60 to be equalised.
  • the section at the neck 6 s , between the leading edges 6 a is retained.
  • the homogenisation of the air flow F is shown in perspective view in FIG. 6 .
  • a flat plate 90 represented by the hatched lines in the figure
  • aerodynamic locking is created in the zone Z0 with a very small amount of movement.
  • the main air flow (arrow F) has a high Mach number.
  • the aerodynamic locking zone is omitted and the flow of air F is homogenised whilst occupying all of the flow cross-section provided, with a lower Mach number.
  • the curvatures are alternated in the direction of flow of the fluid F but also in the tangential direction 6 t .
  • the zones with alternating curvatures can be juxtaposed so that portions of the surface area with the same type of curvature, either hollow or bump, can be close together.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US14/126,989 2011-06-20 2012-06-19 Method of diffusing a gas turbine compression stage, and diffusion stage for implementing same Abandoned US20140133957A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR1155370A FR2976633B1 (fr) 2011-06-20 2011-06-20 Procede de diffusion d'un etage de compression d'une turbine a gaz et etage de diffusion de mise en oeuvre
FR1155370 2011-06-20
PCT/FR2012/051367 WO2012175855A1 (fr) 2011-06-20 2012-06-19 Procede de diffusion d'un etage de compression d'une turbine a gaz et etage de diffusion de mise en oeuvre

Publications (1)

Publication Number Publication Date
US20140133957A1 true US20140133957A1 (en) 2014-05-15

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US14/126,989 Abandoned US20140133957A1 (en) 2011-06-20 2012-06-19 Method of diffusing a gas turbine compression stage, and diffusion stage for implementing same

Country Status (11)

Country Link
US (1) US20140133957A1 (fr)
EP (1) EP2721305B1 (fr)
JP (1) JP6261498B2 (fr)
KR (1) KR101946084B1 (fr)
CN (1) CN103635698B (fr)
CA (1) CA2838686C (fr)
FR (1) FR2976633B1 (fr)
IN (1) IN2014DN00118A (fr)
PL (1) PL2721305T3 (fr)
RU (1) RU2596691C2 (fr)
WO (1) WO2012175855A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2019063384A1 (fr) * 2017-09-28 2019-04-04 Abb Turbo Systems Ag Diffuseur pour compresseur
US11286952B2 (en) 2020-07-14 2022-03-29 Rolls-Royce Corporation Diffusion system configured for use with centrifugal compressor

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3003908B1 (fr) 2013-03-28 2017-07-07 Turbomeca Diffuseur a ailettes d un compresseur radial ou mixte

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20030235497A1 (en) * 2002-06-20 2003-12-25 The Boeing Company Diffuser having a variable blade height

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB685814A (en) * 1950-03-03 1953-01-14 Escher Wyss Ag Improvements in and relating to radial centrifugal compressors and pumps
US6123506A (en) * 1999-01-20 2000-09-26 Pratt & Whitney Canada Corp. Diffuser pipe assembly
RU2197644C1 (ru) * 2001-07-30 2003-01-27 Открытое акционерное общество "Татнефть" Рабочее колесо центробежного насоса
GB0223756D0 (en) * 2002-10-14 2002-11-20 Holset Engineering Co Compressor
RU2330994C2 (ru) * 2006-05-16 2008-08-10 Открытое акционерное общество "Климов" Центробежный компрессор
US8100643B2 (en) * 2009-04-30 2012-01-24 Pratt & Whitney Canada Corp. Centrifugal compressor vane diffuser wall contouring
CN201461538U (zh) * 2009-07-30 2010-05-12 大同北方天力增压技术有限公司 一种抛物线型叶片式扩压器

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20030235497A1 (en) * 2002-06-20 2003-12-25 The Boeing Company Diffuser having a variable blade height

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2019063384A1 (fr) * 2017-09-28 2019-04-04 Abb Turbo Systems Ag Diffuseur pour compresseur
US11286952B2 (en) 2020-07-14 2022-03-29 Rolls-Royce Corporation Diffusion system configured for use with centrifugal compressor

Also Published As

Publication number Publication date
WO2012175855A1 (fr) 2012-12-27
FR2976633B1 (fr) 2015-01-09
RU2013156047A (ru) 2015-07-27
RU2596691C2 (ru) 2016-09-10
CN103635698A (zh) 2014-03-12
CN103635698B (zh) 2017-06-13
JP6261498B2 (ja) 2018-01-17
KR20140047653A (ko) 2014-04-22
EP2721305B1 (fr) 2021-02-24
IN2014DN00118A (fr) 2015-05-22
FR2976633A1 (fr) 2012-12-21
EP2721305A1 (fr) 2014-04-23
JP2014517217A (ja) 2014-07-17
CA2838686C (fr) 2019-09-17
CA2838686A1 (fr) 2012-12-27
KR101946084B1 (ko) 2019-02-08
PL2721305T3 (pl) 2021-07-12

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AS Assignment

Owner name: TURBOMECA, FRANCE

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:TARNOWSKI, LAURENT PIERRE;BULOT, NICOLAS;PORODO, JEROME YVES, FELIX, GILBERT;REEL/FRAME:031799/0740

Effective date: 20131114

STCB Information on status: application discontinuation

Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION

AS Assignment

Owner name: SAFRAN HELICOPTER ENGINES, FRANCE

Free format text: CHANGE OF NAME;ASSIGNOR:TURBOMECA;REEL/FRAME:046127/0021

Effective date: 20160510