US20050201859A1 - Gas turbine ventilation circuitry - Google Patents

Gas turbine ventilation circuitry Download PDF

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Publication number
US20050201859A1
US20050201859A1 US10/517,613 US51761304A US2005201859A1 US 20050201859 A1 US20050201859 A1 US 20050201859A1 US 51761304 A US51761304 A US 51761304A US 2005201859 A1 US2005201859 A1 US 2005201859A1
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US
United States
Prior art keywords
flange
upstream
labyrinth
labyrinths
air
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
US10/517,613
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English (en)
Inventor
Sylvie Coulon
Jean-Claude Taillant
Jean-Philippe Maffre
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 Aircraft Engines SAS
Original Assignee
SNECMA Moteurs 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 SNECMA Moteurs SA filed Critical SNECMA Moteurs SA
Assigned to SNECMA MOTEURS reassignment SNECMA MOTEURS ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: COULON, SYLVIE, MAFFRE, JEAN-PHILIPPE, TAILLAN, JEAN-CLAUDE
Publication of US20050201859A1 publication Critical patent/US20050201859A1/en
Abandoned 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/08—Heating, heat-insulating or cooling means
    • F01D5/081—Cooling fluid being directed on the side of the rotor disc or at the roots of the 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • 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
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • F01D5/3015—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type with side plates
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/28—Arrangement of seals
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the invention relates to the field of ventilating a high pressure turbine in an aircraft turbomachine.
  • the invention relates to a turbomachine having a sealing device between the turbine rotor and the inner casing of the combustion chamber, said turbine rotor comprising firstly, a turbine disk presenting an upstream clamping annulus for fastening it to the downstream cone of a compressor and, secondly, a flange that is disposed upstream from said disk and spaced apart from the disk by a cavity, said flange having an inside bore that is traversed by the upstream clamping annulus of said disk and an upstream clamping annulus so it can be fastened onto said downstream cone, a first air circuit secured to said inner casing in order to deliver a first flow of cooling air into said cavity via main injectors and holes made in said flange, said sealing device comprising a discharge labyrinth between the downstream cone and said inner casing, a main under-injector labyrinth disposed between the flange and the inside wall of the first air circuit, and at least one over-injector labyrinth disposed between the
  • FIG. 1 shows one such high pressure turbine rotor 1 , disposed downstream from a combustion chamber 2 , and comprising a turbine disk 3 fitted with blades 4 , and a flange 5 disposed upstream from the disk 3 .
  • the disk 3 and the flange 5 both include respective upstream clamping annuluses, referenced 3 a for the disk 3 and 5 a for the flange 5 , to enable them to be fastened to the downstream end 6 of the downstream cone 7 of the high pressure compressor driven by the rotor 1 .
  • the disk 3 includes an inside bore 8 passing the shaft 9 of a low pressure turbine, and the flange 5 presents an inside bore 10 that surrounds the clamping annulus 3 a of the disk 3 , and ventilation holes 11 through which a first flow C 1 of cooling air taken from the bottom of the combustion chamber is delivered into the cavity 12 that separates the downstream face of the flange 5 from the upstream face of the disk 3 .
  • the flow C 1 of cooling air flows radially outwards and penetrates into the recesses 4 a that contain the roots of the blades 4 in order to cool said roots.
  • the flow of air comes from the bottom of the combustion chamber, flows into a duct 13 disposed inside the enclosure 14 that separates the flange 5 from the bottom of the combustion chamber, and is drawn into rotation by injectors 15 in order to lower the temperature of the air delivered into the cavity 12 .
  • a second flow C 2 of cooling air taken from the bottom of the combustion chamber flows downstream inside the enclosure 16 that separates the downstream cone 7 in the high pressure compressor from the inner casing 17 of the combustion chamber 2 .
  • the flow C 2 of air flows through a discharge labyrinth 18 and penetrates into the enclosure 14 from whence one portion C 2 a flows through the orifices 19 made in the upstream clamping annulus 5 a of the flange 5 , and then passes through the bore 10 of the flange 5 in order to cool the radially inner portion of the flange and joins flow C 1 of air cooling the blades 4 .
  • Another portion C 2 b of the second flow C 2 of air cools the upstream face of the flange 5 , flows around the injectors 15 and is evacuated into the upstream venting cavity 20 of the turbine rotor 1 .
  • a third portion C 2 c of the second flow C 2 of air serves to ventilate the upstream top face 21 of the flange 5 through a second labyrinth 22 that is situated under the injectors 15 .
  • the third portion C 2 c penetrates into the enclosure 23 that is situated downstream of the second labyrinth 22 , between the flange 5 and the injectors 15 , and is evacuated into the upstream venting cavity of the turbine rotor 1 through a third labyrinth 24 that is situated above the injectors 15 , or else it is mixed with the first flow C 1 of air.
  • the second flow C 2 of air serves to cool the downstream cone 7 , the connection drum connecting the high pressure compressor to the high pressure turbine, and the flange 5 .
  • the second air flow that flows axially in an annular space delimited by the stationary walls secured to the chamber and the adjustable rotating walls secured to the rotor is heated by the power dissipated between the rotor and the stator.
  • the flow C 2 c of air that serves to cool the flange downstream from the second labyrinth 22 , said labyrinth being situated under the injectors 15 cannot be easily controlled since it is subject to the variations that occur, while the engine is in operation and over the lifetime of said engine, in the clearances in the discharge labyrinth 18 , in the second labyrinth 22 , and in the third labyrinth 24 , said third labyrinth being situated above the injectors 15 .
  • said third labyrinth comprises three successive wipers that are formed on an angled portion 25 of the flange 5 , said wipers cooperating with sealing elements 26 secured to an annular structure 27 inserted between the outside wall 28 of the duct 13 and the upstream portion 29 of the inside casing 27 .
  • This type of three-wiper labyrinth is of a considerable weight, and, because of centrifugal forces, it requires the flange 5 to be fastened onto the upstream face of the turbine disk 3 by means of a claw coupling 30 .
  • the first object of the invention is to modify the sealing device upstream from the main injectors, in order to lighten the upstream flange.
  • a second object of the invention is to allow for a decrease in the venting flow rate upstream from the rotor, thereby achieving a saving in specific consumption.
  • a third object of the invention is to raise the pressure levels in the cooling air supply circuit of the turbine wheel, which is favorable for cooling the blades.
  • the invention achieves the first object by the fact that, downstream from the main injectors in the flow direction of the second flow of cooling air, the sealing device comprises at least three labyrinths that are radially spaced apart, being disposed between the flange and the annular structure.
  • each of said three labyrinths comprises a single wiper.
  • Each of the labyrinths is thus light in structure, which makes it possible to omit the claw coupling.
  • the invention achieves the second and third objects by the fact that one of the annular cavities lying between two consecutive labyrinths out of said three labyrinths is fed by cooling air coming from the second circuit upstream from the under-injector labyrinth.
  • this third flow of air is drawn into rotation in the same direction as the rotation of the rotor by the secondary injectors.
  • the secondary injectors are made in the form of sloping holes formed in the annular structure.
  • FIG. 1 is an axial half-section of a high pressure turbine rotor in a turbojet, showing the cooling air circuits and the different sealing labyrinths of the prior art
  • FIG. 2 is an axial half-section of a turbojet turbine rotor showing the disposition of the flange and the labyrinths of the invention, upstream from the main injectors.
  • FIG. 1 The prior art shown in FIG. 1 is described in the introduction and needs no further explanation.
  • FIG. 2 shows a high pressure turbine rotor referenced 1 that is disposed downstream from a combustion chamber 2 , which comprises a turbine disk 3 fitted with blades 4 on its periphery, and a flange 5 that is disposed upstream from the disk 3 .
  • the disk 3 and the flange 5 define between them a cavity 12 that is fed with cooling air via main injectors 15 and via holes 11 made in the flange 5 and opposite the main injectors 15 .
  • the main injectors 15 slope relative to the axis of rotation of the turbine so as to direct the air they supply in the direction of rotation of the turbine rotor 1 .
  • the main injectors 15 are fed with air taken from the bottom of the combustion chamber by means of an annular duct 13 which comprises a radially inner wall 13 a and a radially outer wall 28 .
  • a second labyrinth is disposed under the main injectors, between the radially inner wall 13 a and the flange 5 .
  • An annular structure 27 is inserted between the radially outer wall 28 of the duct 13 and the upstream portion 29 of the inner casing of the combustion chamber 2 .
  • the invention provides for three radially spaced apart labyrinths 31 , 32 , and 33 , replacing the third labyrinth 24 of the prior art, said three labyrinths lying between the cavity 23 that is situated upstream from the second labyrinth and the upstream venting cavity 20 in the turbine rotor 1 and above the main injectors 15 .
  • Each of the three labyrinths 31 , 32 , and 33 comprises a single wiper, and together they define two intermediate cavities 34 and 35 between the enclosure 23 , into which the main injectors 15 and the upstream venting cavity 2 emerge.
  • the labyrinths 31 , 32 , and 33 could be replaced by other rotor/stator sealing systems, such as brush gaskets, and there could also be a combination of labyrinths and of brush gaskets.
  • Branch holes 36 made through the wall of the annular duct 13 serve to put the enclosure 14 , which is under the combustion chamber and is disposed downstream from the second labyrinth that is situated under the main injectors, into communication with the enclosure 37 that is situated radially outer the annular duct 13 .
  • Bore holes 38 that slope relative to the axis of rotation of the turbine rotor 1 are made in the annular structure 27 between the enclosure 37 and the cavity 35 that is situated immediately upstream from the venting cavity 20 .
  • the bore holes 38 slope in the direction of rotation of the turbine rotor 1 in order to reduce the temperature of the cooling air of the radially outer wall of the flange 5 .
  • the invention replaces one over-injector labyrinth 24 of the prior art, comprising three wipers, with three radially spaced apart labyrinths 31 , 32 and 33 , each having a single wiper, makes it possible to simplify the structure of the radially outer portion of the flange 5 .
  • This portion comes in the form of a web having a radially inner end bearing on the roots of the blades 4 and on the teeth of the disk.
  • Such a disposition makes it possible to reduce the weight of the flange 5 and eliminates the claw coupling of the flange 5 onto the disk 3 , which increases the life times of the flange 5 and of the disk 3 .
  • the bore holes 38 are graded in order to reduce the amount of air escaping from the venting cavity 20 , which makes it possible to reduce specific consumption by about 0.1%.
  • the bore holes 38 constitute a system of secondary injectors which make it possible, via the branch holes 36 to make use of most of the air in the under-chamber cavity to cool the top of the upstream flange. This flow of air meets the air that cools the blades, which is why it is commonly referred to as a “shunt” flow.
  • the sloping bore holes 38 may be replaced with vaned injectors or with sloping tubes, assembled in the wall of the annular structure 27 .

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Sealing Using Fluids, Sealing Without Contact, And Removal Of Oil (AREA)
US10/517,613 2002-06-27 2003-06-25 Gas turbine ventilation circuitry Abandoned US20050201859A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
FR02/07979 2002-06-27
FR0207979A FR2841591B1 (fr) 2002-06-27 2002-06-27 Circuits de ventilation de la turbine d'une turbomachine
PCT/FR2003/001958 WO2004003347A1 (fr) 2002-06-27 2003-06-25 Circuits de ventilation de la turbine d'une turbomachine

Publications (1)

Publication Number Publication Date
US20050201859A1 true US20050201859A1 (en) 2005-09-15

Family

ID=29724922

Family Applications (1)

Application Number Title Priority Date Filing Date
US10/517,613 Abandoned US20050201859A1 (en) 2002-06-27 2003-06-25 Gas turbine ventilation circuitry

Country Status (9)

Country Link
US (1) US20050201859A1 (fr)
EP (1) EP1552111A1 (fr)
JP (1) JP2005530956A (fr)
AU (1) AU2003253082A1 (fr)
CA (1) CA2490619A1 (fr)
FR (1) FR2841591B1 (fr)
MA (1) MA27255A1 (fr)
RU (1) RU2005101887A (fr)
WO (1) WO2004003347A1 (fr)

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20050169749A1 (en) * 2003-10-21 2005-08-04 Snecma Moteurs Labyrinth seal device for gas turbine engine
US20170030196A1 (en) * 2015-07-28 2017-02-02 MTU Aero Engines AG Gas turbine
US20180209290A1 (en) * 2017-01-26 2018-07-26 United Technologies Corporation Gas turbine seal
CN108716423A (zh) * 2018-05-08 2018-10-30 中国科学院工程热物理研究所 一种燃气轮机涡轮转静子间鱼嘴封严结构
US20180355743A1 (en) * 2015-12-09 2018-12-13 Mitsubishi Hitachi Power Systems, Ltd. Seal fin, seal structure, turbo machine, and method for manufacturing seal fin
EP3819463A1 (fr) * 2019-11-11 2021-05-12 Rolls-Royce plc Ensemble de turbine avec composants composites à matrice céramique et moyens d'étanchéité inter-étages
US20220243657A1 (en) * 2021-02-03 2022-08-04 Pratt & Whitney Canada Corp. Tangential on-board injector
US11421597B2 (en) 2019-10-18 2022-08-23 Pratt & Whitney Canada Corp. Tangential on-board injector (TOBI) assembly
CN116537895A (zh) * 2023-07-04 2023-08-04 中国航发四川燃气涡轮研究院 一种带有篦齿间隙控制的预旋供气系统

Families Citing this family (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1508672A1 (fr) 2003-08-21 2005-02-23 Siemens Aktiengesellschaft Anneau de fixation segmenté pour une turbine
FR2869094B1 (fr) * 2004-04-15 2006-07-21 Snecma Moteurs Sa Chambre de combustion annulaire de turbomachine a bride interne de fixation amelioree
GB0412476D0 (en) * 2004-06-04 2004-07-07 Rolls Royce Plc Seal system
US20090238683A1 (en) * 2008-03-24 2009-09-24 United Technologies Corporation Vane with integral inner air seal
EP2963377A3 (fr) 2014-06-30 2016-04-13 Marc Hartmann Appareil de libération de fluide dans l'atmosphère
CN107131009B (zh) * 2017-05-16 2019-02-15 中国科学院工程热物理研究所 一种叶轮机械自锁封严结构及具有其的发动机
FR3101670B1 (fr) 2019-10-08 2021-10-08 Safran Aircraft Engines Injecteur pour une turbine haute pression
FR3115562A1 (fr) 2020-10-26 2022-04-29 Safran Aircraft Engines Injecteur d’air de refroidissement pour turbine de turbomachine
FR3118891B1 (fr) 2021-01-15 2023-03-24 Safran Aircraft Engines Fabrication d’un injecteur de turbine par fusion laser sur lit de poudre
FR3126140B1 (fr) 2021-08-11 2024-04-26 Safran Aircraft Engines Flasque d’étanchéité pour turbine de turbomachine
FR3127521B1 (fr) 2021-09-24 2023-12-15 Safran Aircraft Engines Carter d’injection d’air de refroidissement pour turbine de turbomachine
FR3127518A1 (fr) 2021-09-28 2023-03-31 Safran Helicopter Engines Étage de turbomachine comprenant au moins un anneau d’étanchéité
FR3127979B1 (fr) * 2021-10-11 2024-05-31 Safran Aircraft Engines Joint axial à léchettes pour turbomachine
FR3129426A1 (fr) 2021-11-19 2023-05-26 Safran Aircraft Engines Turbomachine à régulation passive du débit de ventilation des injecteurs de turbine

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5143512A (en) * 1991-02-28 1992-09-01 General Electric Company Turbine rotor disk with integral blade cooling air slots and pumping vanes
US5402636A (en) * 1993-12-06 1995-04-04 United Technologies Corporation Anti-contamination thrust balancing system for gas turbine engines
US5984630A (en) * 1997-12-24 1999-11-16 General Electric Company Reduced windage high pressure turbine forward outer seal
US6776573B2 (en) * 2000-11-30 2004-08-17 Snecma Moteurs Bladed rotor disc side-plate and corresponding arrangement
US7048497B2 (en) * 2001-11-08 2006-05-23 Snecma Moteurs Gas turbine stator

Family Cites Families (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4466239A (en) * 1983-02-22 1984-08-21 General Electric Company Gas turbine engine with improved air cooling circuit
FR2744761B1 (fr) * 1996-02-08 1998-03-13 Snecma Disque labyrinthe avec raidisseur incorpore pour rotor de turbomachine

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5143512A (en) * 1991-02-28 1992-09-01 General Electric Company Turbine rotor disk with integral blade cooling air slots and pumping vanes
US5402636A (en) * 1993-12-06 1995-04-04 United Technologies Corporation Anti-contamination thrust balancing system for gas turbine engines
US5984630A (en) * 1997-12-24 1999-11-16 General Electric Company Reduced windage high pressure turbine forward outer seal
US6776573B2 (en) * 2000-11-30 2004-08-17 Snecma Moteurs Bladed rotor disc side-plate and corresponding arrangement
US7048497B2 (en) * 2001-11-08 2006-05-23 Snecma Moteurs Gas turbine stator

Cited By (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7296415B2 (en) * 2003-10-21 2007-11-20 Snecma Moteurs Labyrinth seal device for gas turbine engine
US20050169749A1 (en) * 2003-10-21 2005-08-04 Snecma Moteurs Labyrinth seal device for gas turbine engine
US20170030196A1 (en) * 2015-07-28 2017-02-02 MTU Aero Engines AG Gas turbine
US10428656B2 (en) * 2015-07-28 2019-10-01 MTU Aero Engines AG Gas turbine
US11105213B2 (en) * 2015-12-09 2021-08-31 Mitsubishi Power, Ltd. Seal fin, seal structure, turbo machine, and method for manufacturing seal fin
US20180355743A1 (en) * 2015-12-09 2018-12-13 Mitsubishi Hitachi Power Systems, Ltd. Seal fin, seal structure, turbo machine, and method for manufacturing seal fin
US20180209290A1 (en) * 2017-01-26 2018-07-26 United Technologies Corporation Gas turbine seal
US10408077B2 (en) * 2017-01-26 2019-09-10 United Tehnologies Corporation Gas turbine seal
CN108716423A (zh) * 2018-05-08 2018-10-30 中国科学院工程热物理研究所 一种燃气轮机涡轮转静子间鱼嘴封严结构
US11421597B2 (en) 2019-10-18 2022-08-23 Pratt & Whitney Canada Corp. Tangential on-board injector (TOBI) assembly
US11815020B2 (en) 2019-10-18 2023-11-14 Pratt & Whitney Canada Corp. Tangential on-board injector (TOBI) assembly
US11415016B2 (en) 2019-11-11 2022-08-16 Rolls-Royce Plc Turbine section assembly with ceramic matrix composite components and interstage sealing features
EP3819463A1 (fr) * 2019-11-11 2021-05-12 Rolls-Royce plc Ensemble de turbine avec composants composites à matrice céramique et moyens d'étanchéité inter-étages
US20220243657A1 (en) * 2021-02-03 2022-08-04 Pratt & Whitney Canada Corp. Tangential on-board injector
US11598265B2 (en) * 2021-02-03 2023-03-07 Pratt & Whitney Canada Corp. Tangential on-board injector
CN116537895A (zh) * 2023-07-04 2023-08-04 中国航发四川燃气涡轮研究院 一种带有篦齿间隙控制的预旋供气系统

Also Published As

Publication number Publication date
FR2841591A1 (fr) 2004-01-02
RU2005101887A (ru) 2005-06-27
AU2003253082A1 (en) 2004-01-19
JP2005530956A (ja) 2005-10-13
CA2490619A1 (fr) 2004-01-08
FR2841591B1 (fr) 2006-01-13
EP1552111A1 (fr) 2005-07-13
MA27255A1 (fr) 2005-03-01
WO2004003347A1 (fr) 2004-01-08

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AS Assignment

Owner name: SNECMA MOTEURS, FRANCE

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:COULON, SYLVIE;TAILLAN, JEAN-CLAUDE;MAFFRE, JEAN-PHILIPPE;REEL/FRAME:016100/0511

Effective date: 20041210

STCB Information on status: application discontinuation

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