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US20070122280A1 - Method and apparatus for reducing axial compressor blade tip flow - Google Patents

Method and apparatus for reducing axial compressor blade tip flow Download PDF

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Publication number
US20070122280A1
US20070122280A1 US11/164,636 US16463605A US2007122280A1 US 20070122280 A1 US20070122280 A1 US 20070122280A1 US 16463605 A US16463605 A US 16463605A US 2007122280 A1 US2007122280 A1 US 2007122280A1
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US
United States
Prior art keywords
airfoil
blade
tip
inlet opening
channel
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
US11/164,636
Inventor
Zhifeng Dong
John Rahaim
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.)
General Electric Co
Original Assignee
General Electric Co
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 General Electric Co filed Critical General Electric Co
Priority to US11/164,636 priority Critical patent/US20070122280A1/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: DONG, ZHIFENG, RAHAIM, JOHN JOSEPH
Priority to CA002569177A priority patent/CA2569177A1/en
Priority to CNA2006101729989A priority patent/CN101008402A/en
Priority to JP2006323501A priority patent/JP2007154887A/en
Priority to EP06125091A priority patent/EP1793089A3/en
Publication of US20070122280A1 publication Critical patent/US20070122280A1/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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/10—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using sealing fluid, e.g. steam
    • 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
    • F04D29/08—Sealings
    • F04D29/16—Sealings between pressure and suction sides
    • F04D29/161—Sealings between pressure and suction sides especially adapted for elastic fluid pumps
    • F04D29/164—Sealings between pressure and suction sides especially adapted for elastic fluid pumps of an axial flow wheel
    • 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
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers 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
    • F04D29/00—Details, component parts, or accessories
    • F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
    • F04D29/68—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers
    • F04D29/681—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps
    • F04D29/684—Combating cavitation, whirls, noise, vibration or the like; Balancing by influencing boundary layers especially adapted for elastic fluid pumps by fluid injection
    • 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
    • F05D2240/307—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 tip of a rotor blade

Definitions

  • This invention relates to a method and apparatus for reducing axial compressor tip flow in airfoils, such as blades and vanes.
  • Blade tip flow in the compressor area of a turbine engine results in loss of compressor efficiency and stall margin.
  • flow recirculation in seal cavities along the inner flow path between the vanes and blades also degrades compressor performance.
  • One prior art solution for reducing tip flow is to reduce the blade tip clearance. This is done by a variety of means, including control of the casing and vane interface using mechanical and/or thermal methods. These methods can cause tip rubbing, excess wear and loss of engine efficiency.
  • a method of reducing air flow between a tip of a turbine airfoil rotating in a closely-spaced apart casing comprises the step of providing in the airfoil a radially-extending channel having an inlet opening proximate a base of the airfoil and an exit opening on the airfoil tip. Air is extracted and pressurized from a region proximate the base of the airfoil, introduced into the channel and conveyed through the channel to the airfoil tip. The air exits the channel through the exit openings in the airfoil tip into an area between the airfoil tip and casing under sufficient pressure to resist axial air flow from a pressure side to a suction side of the airfoil.
  • Another aspect of the invention provides a method of reducing air flow between a tip of a turbine airfoil rotating in a closely-spaced apart casing, comprising the steps of providing a first radially-extending channel having an inlet opening proximate a base of the airfoil on a leading edge side thereof, and an exit opening on the airfoil tip, and providing a second radially-extending channel having an inlet opening proximate the base of the airfoil on a trailing edge side thereof, and an exit opening on the airfoil tip.
  • the air is extracted from a region proximate the base of the airfoil into the channel and pumped through the channel to the airfoil tip.
  • the air exits the channel through the exit openings in the airfoil tip into an area between the airfoil tip and casing under sufficient pressure to resist axial air flow from a pressure side to a suction side of the airfoil.
  • a turbine machine compressor airfoil comprising a airfoil base, a airfoil tip, and an air flow channel extending radially from an air inlet opening in the airfoil proximate the airfoil base to an exit opening in the airfoil tip for providing a air blockage against an axial flow of air from the pressure side of the airfoil to the suction side of the airfoil to thereby reduce compressor airfoil tip flow.
  • FIG. 1 is a fragmentary cross-section of an axial flow compressor section of a turbine engine illustrating one embodiment of the invention
  • FIG. 2 is an enlarged fragmentary cross-section of a portion of the compressor shown in FIG. 1 ;
  • FIG. 3 is a fragmentary cross-section of the compressor section of a turbine engine illustrating another embodiment of the invention
  • FIG. 4 is an enlarged fragmentary cross-section of a portion of the compressor shown in FIG. 3 ;
  • FIG. 5 is a fragmentary cross-section of the compressor section of a turbine engine illustrating yet another embodiment of the invention.
  • FIG. 6 is an enlarged fragmentary cross-section of a portion of the compressor shown in FIG. 5 .
  • FIG. 1 a partial section of the axial compressor section of a turbine engine T 1 illustrating a method and apparatus for controlling axial compressor blade tip flow according to the present invention is illustrated in FIG. 1 .
  • the turbine engine “T 1 ” includes compressor blades 10 - 14 and intermediately-positioned stator vanes 15 - 19 in a casing C 1 .
  • the compressor blades 10 - 14 include respective leading edges 10 A- 14 A.
  • blade 10 is shown in enlarged detail for clarity, and is also exemplary of blades 11 - 14 .
  • Air is extracted and pressurized from the area of the leading edge side 10 A of the blade 10 through holes 10 B in a disk 20 .
  • the holes 10 B communicate with a channel 10 C that extends radially outwardly through the blade 10 to the tip where it exits through holes 10 D.
  • the channel 10 C may branch out before exiting the tip of the blade 10 .
  • the size of the channel 10 C and the location and number of the branches is determined empirically based on blade size, shape and volume, and engine performance, rating, tip clearance and similar factors. Note in the drawings that the tip clearance is sufficiently small in relation to the scale of the drawings that actual representation of the tip clearance cannot be shown.
  • a turbine engine “T 2 ” includes compressor blades 30 - 34 and intermediately-positioned stator vanes 35 - 39 in a casing C 2 .
  • the compressor blades 30 - 34 include respective trailing edges 30 A- 34 A.
  • blade 31 is shown in enlarged detail for clarity, and is exemplary of blades 30 and 32 - 34 .
  • Air is extracted from the area of the trailing edge side 31 A of the blade 31 through holes 31 B in the disk rim 40 .
  • the holes 31 B communicate with a channel 31 C that extends radially outwardly through the blade 31 to the tip where it preferably branches before exiting through holes 31 D.
  • a turbine engine “T 3 ” includes compressor blades 50 - 54 and intermediately-positioned stator vanes 55 - 59 in a casing C 3 .
  • the compressor blades 50 - 54 include respective leading edges 50 A- 54 A and respective trailing edges 50 B- 54 B.
  • FIG. 6 illustrates a blade 52 that is shown in enlarged detail for clarity, and is exemplary of blades 51 and 52 - 54 .
  • Air is extracted from both the areas of the leading edge side 52 A and trailing edge side 52 B of blade 52 through holes 52 C and 52 D in the disk rim 60 .
  • the holes 52 C and 52 D communicate with channels 52 E and 52 F, respectively, that extend radially outwardly through the blade 52 to the tip, where they preferably branch before exiting through holes 52 G.
  • the air discharged at the blade tip reduces or prevents blade tip flow by aerodynamically blocking air flow in the region of the tip clearance between the blade tip and the casing. Air from the inner flow path is brought to the tip clearance, as described above, to form this air block. The pressure of the extracted air increases due to the compressor rotor pumping and, when exiting the blade at the tip, resists air flow across the blade tip from the pressure side to the suction side.
  • the methods described above can be applied to both low pressure compressors (boosters) and high pressure compressors. There is no chargeable flow loss when these methods are utilized. Furthermore, by reducing air flow by aerodynamic air blockage rather than by a tight running clearance between the blade tips and the casing, a larger assembly clearance between the blade tips and the casing can be established and maintained. Blade tip rubs are thus reduced, as is recirculation in the inner flow path between the vane and the blade. The extracted air is continuously pumped from the inner flow path to the blade tip, thus providing a continuous air blockage to the blade tip at all times during engine operation.
  • the methods described in this application also have application in blisk (blade integrated disk), skewed or circumferential dovetailed blades.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

Abstract

A turbine machine of the type having a high pressure compressor positioned in a casing is provided with a plurality of rotating compressor blades with at least one air channel formed in the blades for air flow communication from the base to the tip for extracting and pressurizing air from an inner area proximate the disk and conveying the extracted and pressurized air into an area between the blade tip and casing for blocking air flow across the tips of the blades.

Description

    TECHNICAL FIELD AND BACKGROUND OF THE INVENTION
  • This invention relates to a method and apparatus for reducing axial compressor tip flow in airfoils, such as blades and vanes. Blade tip flow in the compressor area of a turbine engine results in loss of compressor efficiency and stall margin. In addition, flow recirculation in seal cavities along the inner flow path between the vanes and blades also degrades compressor performance. One prior art solution for reducing tip flow is to reduce the blade tip clearance. This is done by a variety of means, including control of the casing and vane interface using mechanical and/or thermal methods. These methods can cause tip rubbing, excess wear and loss of engine efficiency.
    • BRIEF DESCRIPTION OF THE INVENTION
  • According to one aspect of the invention, a method of reducing air flow between a tip of a turbine airfoil rotating in a closely-spaced apart casing is provided, and comprises the step of providing in the airfoil a radially-extending channel having an inlet opening proximate a base of the airfoil and an exit opening on the airfoil tip. Air is extracted and pressurized from a region proximate the base of the airfoil, introduced into the channel and conveyed through the channel to the airfoil tip. The air exits the channel through the exit openings in the airfoil tip into an area between the airfoil tip and casing under sufficient pressure to resist axial air flow from a pressure side to a suction side of the airfoil.
  • Another aspect of the invention provides a method of reducing air flow between a tip of a turbine airfoil rotating in a closely-spaced apart casing, comprising the steps of providing a first radially-extending channel having an inlet opening proximate a base of the airfoil on a leading edge side thereof, and an exit opening on the airfoil tip, and providing a second radially-extending channel having an inlet opening proximate the base of the airfoil on a trailing edge side thereof, and an exit opening on the airfoil tip. The air is extracted from a region proximate the base of the airfoil into the channel and pumped through the channel to the airfoil tip. The air exits the channel through the exit openings in the airfoil tip into an area between the airfoil tip and casing under sufficient pressure to resist axial air flow from a pressure side to a suction side of the airfoil.
  • In another aspect of the invention a turbine machine compressor airfoil is provided, comprising a airfoil base, a airfoil tip, and an air flow channel extending radially from an air inlet opening in the airfoil proximate the airfoil base to an exit opening in the airfoil tip for providing a air blockage against an axial flow of air from the pressure side of the airfoil to the suction side of the airfoil to thereby reduce compressor airfoil tip flow.
    • BRIEF DESCRIPTION OF THE DRAWINGS
  • The invention is described below in conjunction with the following drawings, in which:
  • FIG. 1 is a fragmentary cross-section of an axial flow compressor section of a turbine engine illustrating one embodiment of the invention;
  • FIG. 2 is an enlarged fragmentary cross-section of a portion of the compressor shown in FIG. 1;
  • FIG. 3 is a fragmentary cross-section of the compressor section of a turbine engine illustrating another embodiment of the invention;
  • FIG. 4 is an enlarged fragmentary cross-section of a portion of the compressor shown in FIG. 3;
  • FIG. 5 is a fragmentary cross-section of the compressor section of a turbine engine illustrating yet another embodiment of the invention; and
  • FIG. 6 is an enlarged fragmentary cross-section of a portion of the compressor shown in FIG. 5.
    • DESCRIPTION OF THE PREFERRED EMBODIMENT AND BEST MODE
  • Referring now specifically to the drawings, a partial section of the axial compressor section of a turbine engine T1 illustrating a method and apparatus for controlling axial compressor blade tip flow according to the present invention is illustrated in FIG. 1. The turbine engine “T1” includes compressor blades 10-14 and intermediately-positioned stator vanes 15-19 in a casing C1. The compressor blades 10-14 include respective leading edges 10A-14A.
  • As is shown in FIG. 2, blade 10 is shown in enlarged detail for clarity, and is also exemplary of blades 11-14. Air is extracted and pressurized from the area of the leading edge side 10A of the blade 10 through holes 10B in a disk 20. The holes 10B communicate with a channel 10C that extends radially outwardly through the blade 10 to the tip where it exits through holes 10D. Note that the channel 10C may branch out before exiting the tip of the blade 10. The size of the channel 10C and the location and number of the branches is determined empirically based on blade size, shape and volume, and engine performance, rating, tip clearance and similar factors. Note in the drawings that the tip clearance is sufficiently small in relation to the scale of the drawings that actual representation of the tip clearance cannot be shown.
  • Referring now to FIG. 3, a turbine engine “T2” includes compressor blades 30-34 and intermediately-positioned stator vanes 35-39 in a casing C2. The compressor blades 30-34 include respective trailing edges 30A-34A.
  • In FIG. 4, blade 31 is shown in enlarged detail for clarity, and is exemplary of blades 30 and 32-34. Air is extracted from the area of the trailing edge side 31A of the blade 31 through holes 31B in the disk rim 40. The holes 31B communicate with a channel 31C that extends radially outwardly through the blade 31 to the tip where it preferably branches before exiting through holes 31D.
  • Referring now to FIG. 5, a turbine engine “T3” includes compressor blades 50-54 and intermediately-positioned stator vanes 55-59 in a casing C3. The compressor blades 50-54 include respective leading edges 50A-54A and respective trailing edges 50B-54B.
  • FIG. 6 illustrates a blade 52 that is shown in enlarged detail for clarity, and is exemplary of blades 51 and 52-54. Air is extracted from both the areas of the leading edge side 52A and trailing edge side 52B of blade 52 through holes 52C and 52D in the disk rim 60. The holes 52C and 52D communicate with channels 52E and 52F, respectively, that extend radially outwardly through the blade 52 to the tip, where they preferably branch before exiting through holes 52G.
  • In each of the embodiments described above, the air discharged at the blade tip reduces or prevents blade tip flow by aerodynamically blocking air flow in the region of the tip clearance between the blade tip and the casing. Air from the inner flow path is brought to the tip clearance, as described above, to form this air block. The pressure of the extracted air increases due to the compressor rotor pumping and, when exiting the blade at the tip, resists air flow across the blade tip from the pressure side to the suction side.
  • The methods described above can be applied to both low pressure compressors (boosters) and high pressure compressors. There is no chargeable flow loss when these methods are utilized. Furthermore, by reducing air flow by aerodynamic air blockage rather than by a tight running clearance between the blade tips and the casing, a larger assembly clearance between the blade tips and the casing can be established and maintained. Blade tip rubs are thus reduced, as is recirculation in the inner flow path between the vane and the blade. The extracted air is continuously pumped from the inner flow path to the blade tip, thus providing a continuous air blockage to the blade tip at all times during engine operation.
  • The methods described in this application also have application in blisk (blade integrated disk), skewed or circumferential dovetailed blades.
  • A method and apparatus for controlling axial compressor blade tip flow is described above. Various details of the invention may be changed without departing from its scope. Furthermore, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation—the invention being defined by the claims.

Claims (18)

1. A turbine machine of the type having a high pressure compressor positioned in a casing, a plurality of rotating compressor blades having respective blade tips, respective blade bases affixed to a central disk, and a plurality of stationary vanes positioned between respective ones of the blades, comprising at least one air channel formed in respective ones of the blades for air flow communication from the base to the tip for extracting and pressurizing air from an inner area proximate the disk and conveying the extracted and pressurized air through the channel into an area between the blade tip and casing for blocking air flow across a tip of the blades.
2. A turbine machine according to claim 1, wherein the channel exits the blade tip through a respective plurality of exit holes.
3. A turbine machine according to claim 1, wherein the channel includes an air inlet opening in on leading edge side of the blade.
4. A turbine machine according to claim 1, wherein the channel includes an air inlet opening on a trailing edge side of the blade.
5. A turbine machine according to claim 1, wherein the channel includes an air inlet opening in on a leading edge side and an air inlet opening on a trailing edge side of the blade.
6. A turbine machine according to claim 1, and including a first channel having an air inlet opening on a leading edge side of the blade and a second channel having an air inlet opening on a trailing edge side of the blade.
7. A turbine machine according to claims 3, wherein the channel includes a plurality of exit holes in the blade tip.
8. A method of reducing air flow between a tip of a turbine airfoil rotating in a closely-spaced apart casing, comprising the steps of:
(a) providing in the airfoil a radially-extending channel having an inlet opening proximate a base of the airfoil and an exit opening on the airfoil tip;
(b) extracting and pressurizing air from a region proximate the base of the airfoil into the channel;
(c) conveying the air through the channel to the airfoil tip; and
(d) discharging the air through the exit openings in the airfoil tip into an area between the airfoil tip and casing under sufficient pressure to resist axial air flow across the tip of the airfoil.
9. A method according to claim 8, wherein the step of providing a radially-extending channel having an inlet opening proximate a base of the airfoil and an exit opening on the airfoil tip includes the step of forming the inlet opening on a leading edge side of the airfoil.
10. A method according to claim 8, wherein the step of providing a radially-extending channel having an inlet opening proximate a base of the airfoil and an exit opening on the airfoil tip includes the step of forming the inlet opening on a trailing edge side of the blade.
11. A method according to claim 8, wherein the step of providing a radially-extending channel having an inlet opening proximate a base of the airfoil and an exit opening on the airfoil tip includes the steps of forming an air inlet opening on a leading edge side of the airfoil, and an air inlet opening on the trailing edge side of the airfoil.
12. A method according to claim 8, including the steps of providing:
(a) a first radially-extending channel having an inlet opening proximate the base of the airfoil on the leading edge side thereof and an exit opening at the airfoil tip; and
(b) a second radially-extending channel having an inlet opening proximate the base of the airfoil on the trailing edge side thereof and an exit opening at the airfoil tip.
13. A method of reducing air flow between a tip of a turbine airfoil rotating in a closely-spaced apart casing, comprising the steps of:
(a) providing a first radially-extending channel having an inlet opening proximate a base of the airfoil on a leading edge side thereof, and an exit opening on the airfoil tip;
(b) providing a second radially-extending channel having an inlet opening proximate the base of the tip on a trailing edge side thereof, and an exit opening on the airfoil tip;
(c) extracting high pressure air from a region proximate the base of the airfoil into the channels:
(d) conveying the air through the channels to the airfoil tip; and
(e) exiting the air through the exit openings in the airfoil tip into an area between the airfoil tip and casing under sufficient pressure to resist axial air flow across the tip of the airfoil.
14. A turbine machine compressor blade, comprising:
(a) a blade base;
(b) a blade tip;
(c) an air flow channel extending radially from an air inlet opening in the blade proximate the blade base to an exit opening in the blade tip.
15. A turbine machine compressor blade according to claim 14, wherein the air inlet opening is formed on a leading edge side of the blade.
16. A turbine machine compressor blade according to claim 14, wherein the air inlet opening is formed on a trailing edge side of the blade.
17. A turbine machine compressor blade according to claim 14, wherein the channel includes an air inlet opening on a leading edge side of the blade and an air inlet opening on a trailing edge side of the blade.
18. A turbine machine compressor blade according to claim 14, and including a first channel having an air inlet opening on a leading edge side of the blade and a second channel having an air inlet opening on a trailing edge side of the blade.
US11/164,636 2005-11-30 2005-11-30 Method and apparatus for reducing axial compressor blade tip flow Abandoned US20070122280A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
US11/164,636 US20070122280A1 (en) 2005-11-30 2005-11-30 Method and apparatus for reducing axial compressor blade tip flow
CA002569177A CA2569177A1 (en) 2005-11-30 2006-11-29 Method and apparatus for reducing axial compressor blade tip flow
CNA2006101729989A CN101008402A (en) 2005-11-30 2006-11-30 Method and apparatus for reducing axial compressor blade tip flow
JP2006323501A JP2007154887A (en) 2005-11-30 2006-11-30 Method for reducing axial compressor blade tipflow and turbine machine
EP06125091A EP1793089A3 (en) 2005-11-30 2006-11-30 Method and apparatus for reducing axial compressor blade tip flow

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US11/164,636 US20070122280A1 (en) 2005-11-30 2005-11-30 Method and apparatus for reducing axial compressor blade tip flow

Publications (1)

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US20070122280A1 true US20070122280A1 (en) 2007-05-31

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US (1) US20070122280A1 (en)
EP (1) EP1793089A3 (en)
JP (1) JP2007154887A (en)
CN (1) CN101008402A (en)
CA (1) CA2569177A1 (en)

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EP1793089A2 (en) 2007-06-06

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