EP1731711A1 - Transition de la chambre de combustion à la turbine, écran thermique et aube du distributeur de turbine - Google Patents

Transition de la chambre de combustion à la turbine, écran thermique et aube du distributeur de turbine Download PDF

Info

Publication number
EP1731711A1
EP1731711A1 EP05012555A EP05012555A EP1731711A1 EP 1731711 A1 EP1731711 A1 EP 1731711A1 EP 05012555 A EP05012555 A EP 05012555A EP 05012555 A EP05012555 A EP 05012555A EP 1731711 A1 EP1731711 A1 EP 1731711A1
Authority
EP
European Patent Office
Prior art keywords
turbine
hot gas
edge
transverse
vane
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Withdrawn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
EP05012555A
Other languages
German (de)
English (en)
Inventor
Margarete Herz
Marc Tertilt
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.)
Siemens AG
Original Assignee
Siemens AG
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 Siemens AG filed Critical Siemens AG
Priority to EP05012555A priority Critical patent/EP1731711A1/fr
Publication of EP1731711A1 publication Critical patent/EP1731711A1/fr
Withdrawn legal-status Critical Current

Links

Images

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
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F01D5/14—Form or construction
    • F01D5/141—Shape, i.e. outer, aerodynamic form
    • F01D5/142—Shape, i.e. outer, aerodynamic form of the blades of successive rotor or stator blade-rows
    • F01D5/143—Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour
    • 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
    • F05D2240/80—Platforms for stationary or moving blades
    • 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

Definitions

  • the invention relates to a turbine guide vane for a gas turbine, having a vane profile which can be flowed against a front edge by a hot gas which can flow in a main flow direction, with a platform extending transversely to the vane profile and having a platform surface which is surrounded by a peripheral edge and faces the hot gas the peripheral edge is formed of two longitudinal edges running in the main flow direction as well as of two transversely directed transverse edges, of which a first transverse edge is arranged upstream of the front edge. Furthermore, the invention relates to a method for the calculation and production as well as the use of such a turbine guide vane.
  • the gas turbine has a combustion chamber for a hot gas, which at its downstream end merges into an annular hot gas duct in which the turbine vanes of the first turbine stage attached to guide vanes are arranged. Due to the arrangement of the components of the combustion chamber, the guide vanes and turbine guide vanes and their attachment to each other or to each other at the transition from the combustion chamber to the hot gas duct, an expansion gap is required, which must compensate for their temperature-induced strains.
  • the components can travel such large displacement paths or have such large relative movements to each other that the intended active sealing of the concentric to the axis of rotation of the rotor expansion gap is provided with sealing means, but usually not sufficient and sufficient is reliable.
  • sealing means but usually not sufficient and sufficient is reliable.
  • blocking air is additionally blown out of the rear space of the platforms and the combustion chamber wall through the expansion gap into the hot gas duct.
  • sealing air pressure must be greater than the maximum back pressure along the expansion gap, an excessive amount of sealing air is blown out at the points of the expansion gap with a lower pressure in the hot gas, which unnecessarily reduces the efficiency of the gas turbine.
  • the object of the invention is the calculation, manufacture and creation of a turbine guide vane, with which the penetration of hot gas is made difficult in the expansion gap formed by two hot gas boundaries. It is another object of the invention to provide an improved transition region between a combustion chamber and a turbine unit of a gas turbine to increase the efficiency of the gas turbine. Another object of the invention is to further improve the service lives of turbine vanes. In addition, it is an object of the invention to specify the use of a turbine guide vane.
  • the invention proposes to solve the task directed to the turbine vane, to make an aforementioned turbine vane so that the first transverse edge has a curved course, whereby the platform surface convex, i. extends opposite to the main flow direction.
  • the solution is based on the recognition that due to the blocking of the hot gas channel of an axially flowed, stationary gas turbine through the blade profiles of the turbine blades of the first turbine stage upstream of the blade profiles a near-wall pressure curve in the hot gas sets in the annular hot gas duct along a concentric to the axis of rotation of the rotor Circular path is uneven.
  • stagnation points occur, each with locally increased pressure, in front of the blade profiles.
  • the pressure in the hot gas is lower.
  • the first transverse edge of the peripheral edge of the platform which faces the incoming hot gas, has been partially increased contrary to the main flow direction, ie the platform can counter at least in the region of the leading edge
  • the main flow direction can be extended by a curved course of the first transverse edge is realized.
  • the first transverse edge of the platforms of the turbine guide vanes of the first turbine stage is opposite the exit end of the combustion chamber to form the expansion gap.
  • This end of the combustion chamber is adapted to the wave-shaped course, which is formed by the juxtaposition of a plurality of first transverse edges of adjacent turbine vanes of the first turbine stage of the turbine unit, and thus always runs parallel.
  • the lining or wall of the combustion chamber opposite it has a concave, concave contour corresponding thereto.
  • sealing means may be provided in the expansion gap, which can further complicate the penetration of hot gas into the rear space.
  • the curved, upstream, first transverse edge increases the platform surface convex, ie opposite to the main flow direction.
  • the stagnation points in the hot gas before each leading edge with comparatively highest pressure are thus above or in the region of the platform surface, or in other words: the projections made on the position of the stagnation points parallel to the leading edge of the blade profile (radial direction) lie on the platform surface. This ensures that caused by the local stagnation of the hot gas local pressure increase is in the range of the platform, and not as in the prior art, in the region of the expansion gap.
  • the distance between the stagnation points and the expansion gap can be increased, so that the intermediate pressure gradient is reduced. This causes an improved distribution, ie a comparison of the outflowing blocking air along the expansion gap.
  • a circumferential expansion gap provided in a transitional region between the outlet end of the combustion chamber and an inlet end of an annular hot gas channel of the turbine unit is also proposed for saving blown sealing air, which runs in the circumferential direction in such a wavy manner that it continues upstream in the area in front of each blade profile runs as in the area between two adjacent blade profiles.
  • the first transverse edge runs along a pressure level line which establishes in the hot gas during operation of the gas turbine, along which an identical pressure of the hot gas prevails.
  • a wave-shaped expansion gap curve results, at which an equally large gradient between barrier air pressure and hot gas pressure occurs, so that the barrier air strikes at any point an approximately equal flow resistance or counterpressure and flows out in equal amounts along the expansion gap.
  • the vertex of the convex profile of the first transverse edge is flush with the front edge of the blade profile. It can by such a symmetrical arrangement the convex course of these are particularly easy to adapt to the pressure level line of the hot gas.
  • the platform has an end face in the region of the first transverse edge, which is crowned in the main flow direction.
  • the front side of the combustion chamber facing the end face of the platform is formed almost parallel thereto, so that the expansion gap viewed in the radial direction has a curvature in the main flow direction.
  • the object directed to the use of a turbine vane is solved by the features of claim 5 which is directed to the provision of a transition region by the features of claim 6 and the object directed to the method for calculating and manufacturing a turbine vane by the features of claim 8 solved.
  • the advantages directed at the turbine vane also apply to the transition region of a gas turbine, to the process and to the use of the turbine vane.
  • Fig. 1 shows a gas turbine 1 in a longitudinal partial section. It has inside a rotatably mounted about a rotation axis 2 rotor 3, which is also referred to as a turbine runner. Along the rotor 3 follow one another a suction housing 4, a compressor 5, a toroidal annular combustion chamber 6 with a plurality of coaxially arranged burners 7, a turbine unit 8 and an exhaust housing 9.
  • the annular combustion chamber 6 forms a combustion chamber 17, which communicates with an annular hot gas channel 18.
  • There four successive turbine stages 10 form the turbine unit 8. Each turbine stage 10 is formed of two blade rings.
  • the outlet-side, ie downstream end 19 of the annular combustion chamber 6 opens into the inlet end 29 of the annular hot gas duct 18 which is radially inwardly and radially outwardly bounded respectively by the platforms 20 of the turbine vanes 22 of the first turbine stage 10.
  • the wall 21 of the annular combustion chamber 6 and the platforms 20 of the turbine vanes 22 as two immediately adjacent axially wall elements include a design-related expansion gap 23 (Figure 2), concentric with the axis of rotation 2 of the gas turbine 1 at the radially outer transition and / or at the radial inner transition from annular combustion chamber 6 to the hot gas channel 18 is provided.
  • Fig. 2 shows a detail of the unwound transition of the outer periphery of the hot gas channel 18 in detail.
  • This may be both the outer combustion chamber wall 21a with the outer platforms 20b and the inner combustion chamber wall 21b with the inner platforms 20b (FIG. 3).
  • a vane profile 30 extends in the radial direction R.
  • the circumferential direction is designated by U.
  • the platforms 20 have a circumferential edge 32, which is composed of two along the main flow direction H extending longitudinal edges 36 and two transverse edges 38 extending transversely thereto.
  • the upstream in the flow direction of the hot gas 11 disposed first transverse edge 38 a, with respect to the platform surface 28, a curved, preferably convex contour 39, the apex S, as seen in the main flow direction H, with a front edge 42 of the blade profile 30 is aligned. Since the turbine guide vanes 22 of a turbine stage 10 extend annularly on a ring, a wavy course 37 of the expansion gap 23 results due to the curved first transverse edges 38a in the circumferential direction U.
  • the expansion gap 23 has a constant gap dimension via its extent extending in the circumferential direction U. since the first transverse edge 38a opposite, exit-side end 19 of the annular combustion chamber 6 correspondingly sections concave, ie each provided on the outlet end 19 of the annular combustion chamber 6 ceramic or metallic heat shield 26 has a concave front side contour or transverse edge 27, which the convex transverse edges 38a the platforms 26 is opposite. Alternatively, a circumferential intermediate element could also be provided between the heat shields 26 and the platforms 10, which has a straight course or contour on the hot gas channel side and a wavy and combustion chamber side.
  • hot gas 11 flows from the annular combustion chamber 6 into the hot gas channel 18 of the turbine unit 8. Due to congestion of the hot gas 11 to the arranged in the hot gas channel 18 blade profiles 30 of the turbine vanes 22 of the first turbine stage 10 arise upstream on a concentric to the axis of rotation 2 of the rotor 3 circular path stagnation points 41, their pressures, compared with the also lying on the circular path intermediate points 35th are increased. In the intermediate points 35 there is a lower pressure, since from these locations the hot gas 11 can then flow unhindered downstream between the blade profiles 30.
  • each platform 20 has been extended counter to the main flow direction H of the hot gas 11, in particular in the central region in front of the respective blade profile 30, lie the projections of the stagnation points 41 parallel to the front edge 42 on the platform surfaces 28 and not in the region of an expansion gap 43, which is indicated in FIG. 2 and is known from the prior art.
  • the course 37 of the expansion gap 23 in the circumferential direction U is wave-shaped so that it along a pressure level line of the hot gas 11, along which an equal pressure prevails, so that the air flowing from the expansion gap 23 sealing air 25 in the circumferential direction U always on a smaller, but considered along the expansion gap 23, the same size back pressure can flow. Consequently, approximately the same amount of sealing air 25 flows out along the expansion gap 23 everywhere. The entry of hot gas 11 into the expansion gap 23 is thus equally effectively prevented at each point of the expansion gap 23.
  • Fig. 3 shows a section along the section line III-III of Fig. 2.
  • the transition region 44 between the annular combustion chamber 6 and the hot gas channel 18 is shown in a section.
  • the walls 21a, 21b of the annular combustion chamber 6 and the platforms 20a, 20b of the turbine vanes 22 define the flow path for the hot gas 11.
  • the expansion gap 23 may be formed at the radially inner transition from the annular combustion chamber 6 to the hot gas channel 18.
  • cooled sealing air 25 is blown through a back space 24 through the expansion gap 23 in the direction of the hot gas duct 18.
  • Fig. 4 shows a detail of FIG. 3 in an advantageous embodiment of the invention.
  • the expansion gap 23a initially runs in the radial direction and then curves in an arc range in the axial direction.
  • a second, the expansion gap 23a mitformende end face of the combustion chamber wall runs parallel thereto.
  • the blockage of the hot gas duct 18 caused by the blade profiles 30 leads to stagnation points 41 arranged in front of the leading edge 30 of the blade profile 30. Close to the platform, on the radially outer and inner edges of the hot gas duct 18, swirls 46 also occur in the hot gas flow, passing beyond the expansion gap 23a can flow back.
  • the opposite curved expansion gap 23a the penetration of hot gas 11 is further complicated.
  • each platform 20 has an upstream first transverse edge 38a, the combustion chamber wall facing end face 40 is bent in the main flow direction H.
  • FIG. 5 shows a diagram with a pressure curve 50 of the hot gas 11 along the prior art expansion gap 43 (FIG. 2) upstream of the leading edge 42 of the blade profile 30 of the first turbine stage 10 of the stationary gas turbine 1.
  • the expansion gap 43 extends in a more straight line Transverse edges of the platforms of the turbine vanes along the circumference U on a circular path whose center coincides with the axis of rotation 2.
  • a stagnation point 41 with a local pressure maximum 51 arises during operation of the gas turbine 1, whereas at the same height between the blade profiles 30 the pressure in the hot gas 11 is lower.
  • the pressure curve 52 of the sealing air 25 flowing from the expansion gap 23 extending in the circumferential direction U in the circumferential direction U is likewise shown in the diagram.
EP05012555A 2005-06-10 2005-06-10 Transition de la chambre de combustion à la turbine, écran thermique et aube du distributeur de turbine Withdrawn EP1731711A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP05012555A EP1731711A1 (fr) 2005-06-10 2005-06-10 Transition de la chambre de combustion à la turbine, écran thermique et aube du distributeur de turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP05012555A EP1731711A1 (fr) 2005-06-10 2005-06-10 Transition de la chambre de combustion à la turbine, écran thermique et aube du distributeur de turbine

Publications (1)

Publication Number Publication Date
EP1731711A1 true EP1731711A1 (fr) 2006-12-13

Family

ID=35170059

Family Applications (1)

Application Number Title Priority Date Filing Date
EP05012555A Withdrawn EP1731711A1 (fr) 2005-06-10 2005-06-10 Transition de la chambre de combustion à la turbine, écran thermique et aube du distributeur de turbine

Country Status (1)

Country Link
EP (1) EP1731711A1 (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009019282A2 (fr) 2007-08-06 2009-02-12 Alstom Technology Ltd Installation de turbine à gaz
WO2011138193A1 (fr) * 2010-05-05 2011-11-10 Alstom Technology Ltd. Zone de transition pour une chambre de combustion secondaire d'une turbine à gaz
EP2372102A3 (fr) * 2010-04-02 2014-11-05 United Technologies Corporation Plate-forme des pales de rotor d'une turbine à gaz
EP2937515A1 (fr) * 2010-03-23 2015-10-28 United Technologies Corporation Moteur à turbine à gaz doté d'une plateforme de pale de rotor contourée à surface non axisymétrique

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5647603A (en) * 1979-09-28 1981-04-30 Hitachi Ltd Moving blade of turbine
US4739621A (en) * 1984-10-11 1988-04-26 United Technologies Corporation Cooling scheme for combustor vane interface
EP1067273A1 (fr) * 1999-07-06 2001-01-10 ROLLS-ROYCE plc Configuration d'une bande de recouvrement des aubes de turbine
EP1391582A2 (fr) * 2002-08-22 2004-02-25 Kawasaki Jukogyo Kabushiki Kaisha Structure d'étanchéité pour chemise de chambre de combustion
WO2004057158A1 (fr) * 2002-12-19 2004-07-08 Siemens Aktiengesellschaft Turbine, dispositif de fixation pour aubes fixes et procede de demontage des aubes fixes d'une turbine
US20050100439A1 (en) * 2003-09-09 2005-05-12 Alstom Technology Ltd Turbomachine

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5647603A (en) * 1979-09-28 1981-04-30 Hitachi Ltd Moving blade of turbine
US4739621A (en) * 1984-10-11 1988-04-26 United Technologies Corporation Cooling scheme for combustor vane interface
EP1067273A1 (fr) * 1999-07-06 2001-01-10 ROLLS-ROYCE plc Configuration d'une bande de recouvrement des aubes de turbine
EP1391582A2 (fr) * 2002-08-22 2004-02-25 Kawasaki Jukogyo Kabushiki Kaisha Structure d'étanchéité pour chemise de chambre de combustion
WO2004057158A1 (fr) * 2002-12-19 2004-07-08 Siemens Aktiengesellschaft Turbine, dispositif de fixation pour aubes fixes et procede de demontage des aubes fixes d'une turbine
US20050100439A1 (en) * 2003-09-09 2005-05-12 Alstom Technology Ltd Turbomachine

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
PATENT ABSTRACTS OF JAPAN vol. 005, no. 105 (M - 077) 8 July 1981 (1981-07-08) *

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009019282A2 (fr) 2007-08-06 2009-02-12 Alstom Technology Ltd Installation de turbine à gaz
WO2009019282A3 (fr) * 2007-08-06 2009-05-07 Alstom Technology Ltd Installation de turbine à gaz
US8132417B2 (en) 2007-08-06 2012-03-13 Alstom Technology Ltd. Cooling of a gas turbine engine downstream of combustion chamber
EP2937515A1 (fr) * 2010-03-23 2015-10-28 United Technologies Corporation Moteur à turbine à gaz doté d'une plateforme de pale de rotor contourée à surface non axisymétrique
EP2372102A3 (fr) * 2010-04-02 2014-11-05 United Technologies Corporation Plate-forme des pales de rotor d'une turbine à gaz
US9976433B2 (en) 2010-04-02 2018-05-22 United Technologies Corporation Gas turbine engine with non-axisymmetric surface contoured rotor blade platform
WO2011138193A1 (fr) * 2010-05-05 2011-11-10 Alstom Technology Ltd. Zone de transition pour une chambre de combustion secondaire d'une turbine à gaz
CH703105A1 (de) * 2010-05-05 2011-11-15 Alstom Technology Ltd Gasturbine mit einer sekundärbrennkammer.
CN102884282A (zh) * 2010-05-05 2013-01-16 阿尔斯通技术有限公司 用于燃气涡轮机的二次燃烧室的过渡区域
CN102884282B (zh) * 2010-05-05 2015-07-29 阿尔斯通技术有限公司 用于燃气涡轮机的二次燃烧室的过渡区域
US9097119B2 (en) 2010-05-05 2015-08-04 Alstom Technology Ltd. Transitional region for a secondary combustion chamber of a gas turbine

Similar Documents

Publication Publication Date Title
EP2179143B1 (fr) Refroidissement de fente entre une paroi de chambre de combustion et une paroi de turbine d'une installation de turbine à gaz
EP3176370B1 (fr) Ensemble d'aubes directrices pour turbomachine
DE102015219556A1 (de) Diffusor für Radialverdichter, Radialverdichter und Turbomaschine mit Radialverdichter
DE102012013160A1 (de) Labyrinthdichtungen
CH698041B1 (de) Turbine.
DE102011056638A1 (de) Turbine mit einer Abgashaube
DE102011054586A1 (de) Dichtungsanordnung für eine Turbomaschine
EP3290644B1 (fr) Turbine à gaz
DE102015111843A1 (de) Turbine mit gekühlten Turbinenleitschaufeln
EP1012445B2 (fr) Aube pour une turbomachine
DE102010050185A1 (de) Axialturbomaschine
WO2017140756A1 (fr) Roue de turbine à écoulement mixte d'un turbocompresseur à gaz d'échappement et turbine à gaz d'échappement dotée d'une telle roue
EP1731711A1 (fr) Transition de la chambre de combustion à la turbine, écran thermique et aube du distributeur de turbine
DE102010037692A1 (de) Geformte Wabendichtung für eine Turbomaschine
DE102016107315A1 (de) Rotor mit Überhang an Laufschaufeln für ein Sicherungselement
CH711975A2 (de) Radialturbine, Turbolader und Einsatzstück für ein Turbinengehäuse der Radialturbine.
WO2011088819A2 (fr) Système de boîtier pour une turbomachine axiale
DE102016212649A1 (de) Brennerdichtung einer Gasturbine und Verfahren zu deren Herstellung
EP2860356B1 (fr) Turbomachine
EP2665896B1 (fr) Carter intermédiaire d'une turbine à gaz comprenant une paroi de limite extérieure en amont d'une entretoise avec un contour variable en direction circonférentielle pour minimiser les pertes de flux secondaires
DE102015122238A1 (de) System und Verfahren beinhaltend eine Umfangsdichtungsanordnung zur Erleichterung der Abdichtung in einer Turbine
EP1673519B1 (fr) Dispositif d'etancheite pour une turbine a gaz
EP3056684A1 (fr) Bague intérieure séparée axialement pour une turbomachine, un stator et groupe motopropulseur
EP3536913B1 (fr) Bague intérieure pour une turbomachine et procédé de fabrication de cette bague intérieure
EP3495629B1 (fr) Canal d'écoulement de turbomachine

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): AT BE BG CH CY CZ DE DK EE ES FI FR GB GR HU IE IS IT LI LT LU MC NL PL PT RO SE SI SK TR

AX Request for extension of the european patent

Extension state: AL BA HR LV MK YU

AKX Designation fees paid
STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: THE APPLICATION IS DEEMED TO BE WITHDRAWN

18D Application deemed to be withdrawn

Effective date: 20070614

REG Reference to a national code

Ref country code: DE

Ref legal event code: 8566