EP1936125A1 - Turbomaschinenverdichter - Google Patents

Turbomaschinenverdichter Download PDF

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Publication number
EP1936125A1
EP1936125A1 EP06292027A EP06292027A EP1936125A1 EP 1936125 A1 EP1936125 A1 EP 1936125A1 EP 06292027 A EP06292027 A EP 06292027A EP 06292027 A EP06292027 A EP 06292027A EP 1936125 A1 EP1936125 A1 EP 1936125A1
Authority
EP
European Patent Office
Prior art keywords
ferrules
annular
blades
flanges
compressor
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
EP06292027A
Other languages
English (en)
French (fr)
Other versions
EP1936125B1 (de
Inventor
Enrique Penalver Castro
Frédéric Gebbia
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 Aero Boosters SA
Original Assignee
Techspace Aero 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 Techspace Aero SA filed Critical Techspace Aero SA
Priority to AT06292027T priority Critical patent/ATE425346T1/de
Priority to ES06292027T priority patent/ES2323283T3/es
Priority to EP06292027A priority patent/EP1936125B1/de
Priority to DE602006005693T priority patent/DE602006005693D1/de
Publication of EP1936125A1 publication Critical patent/EP1936125A1/de
Application granted granted Critical
Publication of EP1936125B1 publication Critical patent/EP1936125B1/de
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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
    • F01D25/243—Flange connections; Bolting arrangements

Definitions

  • the present invention relates to a turbomachine compressor, in particular a low-pressure compressor, and a turbomachine such as a turbojet engine or an airplane turboprop engine equipped with this compressor.
  • a low-pressure turbomachine compressor comprises a plurality of compression stages each comprising an upstream annular row of rotor blades rotating inside a fixed casing and driven by a shaft of the turbomachine, and a downstream annular row of fixed blades for rectifying the flow of air, which are carried by the casing at their radially outer ends.
  • the casing has a substantially cylindrical or frustoconical shape and is formed of annular rings arranged coaxially end to end, and each carrying an annular row of fixed vanes rectification.
  • the ferrules comprise at their ends cylindrical flanges each terminating in an annular flange.
  • the cylindrical upstream rim of greater axial dimension of each ferrule surrounds the annular row of upstream rotor blades, and comprises an upstream annular flange for attachment on a corresponding flange provided on the downstream cylindrical rim of smaller axial dimension of another shell located upstream.
  • the joining plane of the fastening flanges of two adjacent ferrules extends substantially upstream of the leading edges of the rotor blades located between the vanes carried by these ferrules.
  • the junction plane of these flanges may extend downstream of the trailing edges of the rotor blades located between the vanes carried by the ferrules.
  • An annular element of abradable material is fixed on the inner cylindrical surface of the axial large upstream flange of each ferrule, and cooperates with the tops of the blades to limit the radial clearances between the blades and the casing.
  • the purpose of the invention is in particular to provide a simple solution, effective and economical to all of the aforementioned problems of the prior art.
  • turbomachine compressor comprising a substantially cylindrical housing formed of annular ferrules arranged coaxially end to end and each carrying an annular row of straightening vanes, the ferrules having at their axial ends cylindrical flanges carrying annular flanges. fastening ferrules together, characterized in that the flanges for fixing two adjacent ferrules are in the axial direction substantially in the middle of an annular row of blades located between these ferrules.
  • the flanges for fixing the flanges of two adjacent ferrules surround an annular array of moving blades, the flanges of these ferrules having substantially the same axial dimension.
  • the upstream and downstream cylindrical flanges of a ferrule may have the same axial dimension or different axial dimensions.
  • the fastening flanges of two adjacent ferrules surround rotor blades and are substantially equidistant from the ferrules, which makes it possible to stiffen the crankcase, which can thus withstand significant shocks such as those produced by parts of blades when rupture of the compressor rotor.
  • the axial dimensions of the upstream and downstream flanges of each shell can be optimized so as to reduce the axial size of the compressor, which was not possible in the current technique.
  • the cylindrical edges of the ends of the adjacent ferrules extend over at least the entire axial length of the blades situated between these ferrules and form an annular groove around these moving blades, this annular groove having a internal diameter greater than the outer diameter of the theoretical vein in this part of the housing.
  • An element of abradable material may be mounted in the annular groove and advantageously has an axial dimension at least equal to that of the top of the blades which it surrounds, and an internal diameter equal to or greater than the external diameter of the theoretical vein in this part crankcase.
  • the outer diameter of the blades of each annular row can thus be, cold, equal to or greater than the outer diameter of the theoretical vein in the corresponding part of the housing.
  • the abradable element can be in two parts mounted end to end, each part of the abradable element being fixed on a cylindrical end rim of a ferrule.
  • the abradable element extends on either side of the junction plane between the ferrules, which limits the formation of projections within the housing.
  • the tops of the moving blades which come into operation use the abradable element by friction make it possible to make up for any unevenness or protrusion between the two parts of the element. abradable.
  • the invention also relates to a turbomachine such as a turbojet engine or a turboprop engine, characterized in that it comprises a compressor as described above.
  • the compressor 10 of the figure 1 comprises a number of compression stages of which only one has been shown, each stage comprising an annular row of rotor blades 12, whose radially inner ends are fixed on a disk carried by a shaft of the turbomachine, and a annular row of fixed vanes 14 of rectification, arranged downstream of the annular row of moving blades 12 and whose radially outer ends are carried by a substantially cylindrical housing 16 of axis 18.
  • the housing 16 is formed of annular ferrules 20, 22 arranged coaxially end to end, and each carrying a row of fixed blades 14.
  • the ferrules 20, 22 each comprise an upstream cylindrical rim 26 and a downstream cylindrical rim 28 ending in annular flanges 30, 30 'for fixing ferrules together.
  • the flange 30 carried by the upstream flange of the shell 22 is applied and tightened by appropriate means of the screw-nut type on the corresponding flange 30 'carried by the downstream flange of the ferrule 20 located upstream.
  • the upstream flange 26 of the shell 22 further comprises at its free end a cylindrical surface 31 in which is engaged the downstream end of the downstream flange 28 of the shell 20.
  • junction plane J of the fastening flanges 30, 30 'of two adjacent ferrules extends substantially upstream of the leading edges (or downstream of the trailing edges) of the blades 12 situated between these ferrules. ( figure 2 ).
  • this technique has several disadvantages detailed in the foregoing.
  • the invention provides a simple and effective solution to all of these problems by moving in the axial direction of the junction plane J 'flanges 20, 20' substantially in the middle of the blades 12 and the formation of a ring groove ("Trench") in the edges of the ferrules around the blades.
  • the upstream cylindrical flange 26 of each ferrule 20, 22 has an axial dimension substantially equal to that of the downstream cylindrical flange 28 to which it is attached.
  • the flanges 26, 28 connecting two ferrules adjacent ones surround a row of moving blades 12 and define an annular groove 32 opening radially inwards.
  • each flange 26, 28 has an end portion 36 of small axial dimension and small internal diameter which is aligned axially with a shell 20.
  • Each flange 26, 28 further comprises an end portion 40 of larger axial dimension and larger internal diameter which extends around the blades 12 and internally delimits a portion of the annular groove 32 above.
  • the inside diameter of the end portion 40 of the flange is greater than the outside diameter D of the theoretical vein in the corresponding part of the casing, and the inside diameter of the end portion 36 of this flange is substantially equal to the outside diameter D of this theoretical vein.
  • a cylindrical element 34 of abradable material is fixed in the annular groove 32 and extends around the blades 12, over substantially their entire axial dimension. This element 34 is intended to cooperate with the tops of the blades to prevent the passage of air in axial direction between the tops of the blades 12 and the casing 16, as will be described in more detail in the following.
  • the element 34 is formed of two parts mounted coaxially end to end. Each of these parts is fixed by appropriate means on an end portion 40 of larger internal diameter of a flange 26, 28.
  • the internal diameter of each portion of the abradable element is substantially equal to one of its axial ends, with that of the end portion 36 of the flange to which it is attached, and at the other of its ends, with that of the other complementary part of the element.
  • the outer diameter of the blades is, in the cold, substantially equal to the external diameter D of the theoretical vein in the corresponding part of the casing, that is to say that the tops of the blades 12 are aligned axially with the cylindrical surface internal part of the housing.
  • the tops of the blades 12 define with the abradable element 34 which surrounds them a small radial clearance so as to limit the passage of air between the blades 12 and the casing 16.
  • the blades 12 expand more or less in the radial direction and this radial clearance can disappear.
  • the tops of the blades then come into contact with the abradable element 34 by friction, which makes it possible to make up any jump or difference between the two parts of the abradable element 34.
  • the invention has many other advantages detailed above.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP06292027A 2006-12-22 2006-12-22 Turbomaschinenverdichter Active EP1936125B1 (de)

Priority Applications (4)

Application Number Priority Date Filing Date Title
AT06292027T ATE425346T1 (de) 2006-12-22 2006-12-22 Turbomaschinenverdichter
ES06292027T ES2323283T3 (es) 2006-12-22 2006-12-22 Compresor de turbomaquina.
EP06292027A EP1936125B1 (de) 2006-12-22 2006-12-22 Turbomaschinenverdichter
DE602006005693T DE602006005693D1 (de) 2006-12-22 2006-12-22 Turbomaschinenverdichter

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP06292027A EP1936125B1 (de) 2006-12-22 2006-12-22 Turbomaschinenverdichter

Publications (2)

Publication Number Publication Date
EP1936125A1 true EP1936125A1 (de) 2008-06-25
EP1936125B1 EP1936125B1 (de) 2009-03-11

Family

ID=37913167

Family Applications (1)

Application Number Title Priority Date Filing Date
EP06292027A Active EP1936125B1 (de) 2006-12-22 2006-12-22 Turbomaschinenverdichter

Country Status (4)

Country Link
EP (1) EP1936125B1 (de)
AT (1) ATE425346T1 (de)
DE (1) DE602006005693D1 (de)
ES (1) ES2323283T3 (de)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP2733314A1 (de) * 2012-11-15 2014-05-21 Techspace Aero S.A. Radiale Befestigungs- und Positionierflansche für Gehäuseabschnitte einer axialen Turbomaschine
US8926277B2 (en) 2011-03-03 2015-01-06 Rolls-Royce Plc Fan casing for a turbofan engine
US10100674B2 (en) 2012-11-15 2018-10-16 Safran Aero Boosters Sa Radial fixing and positioning flanges for shells of axial turbine compressor housings
FR3098241A1 (fr) * 2019-07-04 2021-01-08 Safran Aircraft Engines Tambour pour une turbomachine d’aeronef

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB960812A (en) 1963-04-08 1964-06-17 Rolls Royce Compressor for a gas turbine engine
WO1995025879A1 (en) 1994-03-21 1995-09-28 United Technologies Corporation Compressor stator vane assembly

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB960812A (en) 1963-04-08 1964-06-17 Rolls Royce Compressor for a gas turbine engine
WO1995025879A1 (en) 1994-03-21 1995-09-28 United Technologies Corporation Compressor stator vane assembly

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8926277B2 (en) 2011-03-03 2015-01-06 Rolls-Royce Plc Fan casing for a turbofan engine
EP2733314A1 (de) * 2012-11-15 2014-05-21 Techspace Aero S.A. Radiale Befestigungs- und Positionierflansche für Gehäuseabschnitte einer axialen Turbomaschine
US10100674B2 (en) 2012-11-15 2018-10-16 Safran Aero Boosters Sa Radial fixing and positioning flanges for shells of axial turbine compressor housings
FR3098241A1 (fr) * 2019-07-04 2021-01-08 Safran Aircraft Engines Tambour pour une turbomachine d’aeronef

Also Published As

Publication number Publication date
ATE425346T1 (de) 2009-03-15
DE602006005693D1 (de) 2009-04-23
EP1936125B1 (de) 2009-03-11
ES2323283T3 (es) 2009-07-10

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