GB2502302A - Gas turbine nozzle guide vane with dilution air exhaust ports - Google Patents

Gas turbine nozzle guide vane with dilution air exhaust ports Download PDF

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Publication number
GB2502302A
GB2502302A GB1209016.3A GB201209016A GB2502302A GB 2502302 A GB2502302 A GB 2502302A GB 201209016 A GB201209016 A GB 201209016A GB 2502302 A GB2502302 A GB 2502302A
Authority
GB
United Kingdom
Prior art keywords
nozzle guide
ports
gas turbine
gas
guide 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.)
Pending
Application number
GB1209016.3A
Other versions
GB201209016D0 (en
Inventor
Bhupendra Khandelwal
Vishal Sethi
Riti Singh
David G Macmanus
Gajanana Hedge
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.)
Cranfield University
Original Assignee
Cranfield University
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 Cranfield University filed Critical Cranfield University
Priority to GB1209016.3A priority Critical patent/GB2502302A/en
Publication of GB201209016D0 publication Critical patent/GB201209016D0/en
Publication of GB2502302A publication Critical patent/GB2502302A/en
Pending legal-status Critical Current

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Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/14Form or construction
    • F01D5/18Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186Film cooling
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00Stators
    • F01D9/02Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • YGENERAL 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
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies, e.g. for aircraft

Abstract

In a gas turbine, a substantial amount of dilution air is introduced into turbine nozzle guide vane 80 from which it leaves via a plurality of ports 100, 101 in the external surfaces 110, 112 which extend between the leading and trailing edges 81, 82 of the vane. The ports 100, 101 are spaced and configured so as to provide a predetermined radial distribution of gas temperature at the nozzle guide vane exit plane (P, fig.3). The ports 100, 101 may be oriented at an angle T to promote tangential flow of dilution air as indicated by arrows 60.

Description

TITLE: GAS TURBINE
DESCRIPTION
TECHNICAL FIELD
The present invention relates to gas turbines.
BACKGROUND ART
Figure 1 illustrates a gas turbine combustor 10 as known, e.g. from Henderson, R.E.
and Blazowski, W.S., "Turbopropulsion Combustion Technology', The Aerothermoa'vnamics qf Aircraft Gas Turbine Engines, AFAPL-TR-78-52 (1978). As indicated by arrow 20, most of the delivery air from the compressor 15 (typically 80% in a modem gas turbine) is supplied to the combustor. A first portion 21 of this air is fed to the primary combustion zone 30 through swirlers 22 where it bums stoichiometrically with fuel introduced through nozzle 35. A second portion of the air is introduced further downstream: as primary cooling flow 41, secondary cooling flow 42 and tertiary cooling flow 43, the latter also being known as "dilution air" and the rearmost region of the combustor into which that air is introduced bcing known as thc "dilution zone" 50. This dilution air 43 typically represents around 30% of the air 20 supplied to the combustor, although this proportion may vary between 20% and 50% depending on the particular design. As indicated at 70, the mixture of combustion products and dilution air leaving the combustion chamber passes through turbine nozzle guide vanes 80 and the turbine rotor 90, where a small amount of air from the compressor (typically 10% in a modem gas turbine) is typically introduced through the turbine nozzle guide vanes for cooling purposes.
As explained in GB774500, published in 1957, dilution air is surplus to that necessary
I
for complete, stoichiometric combustion of the fuel and serves to cool the temperature of the mixture to a value acceptable to the turbine. As illustrated in figure 2A below, GB774500 proposes a gas turbine in which the whole or most of the dilution or cooling air is admifted to the interior 85 of hollow stator guide vanes 80 each having a slot 87 extending along its leading edge 81 for the discharge of air. As indicated by arrows 86 in figure 2B below, the dilution air discharges into the gas stream 70 issuing from the combustor and then reverses in direction to flow and mix with the combustion gases. Air may also exit through outlets (not shown) in the trailing edges of the stator guide vanes. To prevent overheating of the turbine parts, GB774500 suggests that it is usual to attempt to achieve complete mixing.
However, in modem high-performance engines, the desired average radial distribution of temperature at the combustor exit plane is far from flat: as noted in Lefebvre, A.H. Gas Turbine Combustion, Taylor & Francis, Philadelphia, USA. (1998), it instead has a profile that peaks above the mid-height of the turbine blade, having lower temperatures at the turbine blade root, where mechanical stresses are highest, and at the tip of the blade which is the most difficult to cool. Lefebvre indicates that such a distribution may be achieved by appropriate choicc of the numbcr and sizc of thc air admission holes in thc combustor dilution zone as well as the zone length.
DISCLOSURE OF INVENTION
According to the present invention, there is provided a gas turbine having a compressor, a combustor and a turbine; the turbine comprising a nozzle guide vane having leading and trailing edges linked by extemal surfaces; the compressor providing a first gas flow for stoichiometric combustion of fuel in the combustor and a second gas flow for dilution of the products of said stoichiometric combustion, wherein the second gas flow is exhausted through at least one port in at least one of said extemal surfaces of the nozzle guide vane.
Exhaust of the dilution air through an external surface of the nozzle guide vane, rather than through the leading and trailing edges of a nozzle guide vane as disclosed in GB774500, will cause less disturbance of the gas flow and thus lower losses. Moreover, unlike GB774500, the arrangement of the present invention requires no dedicated cooling.
The second gas flow may be exhausted through a plurality of ports in one of the external surfaces of the nozzle guide vane. The plurality of ports may be spaced in a radial direction relative to the axis of rotation of the turbine rotor. Moreover, the plurality of ports may be configured so as to provide a predetermined radial distribution of gas temperature at the nozzle guide vane exit plane. The plurality of ports may also be spaced along the external surface between the leading and trailing edges. At least one of the plurality of ports may be oriented at an angle relative to the surface to promote flow substantially tangential to the surface.
BRIEF DESCRIPTION OF DRAWINGS
An embodiment of the invention will now be described by way of example with reference to the accompanying drawings, in which: Figure 3 is a radial cross-sectional view of a first ernbodirnent of the invention; Figures 4A and B are cross-sectional views of the nozzle guide vane of figure 3; Figure 4C is a detail view of figure 4B.
DETAILED DESCRIPTION OF SPECIFIC EMBODIMENTS
As indicatcd by arrows 60 in figurc 3, a substantial amount of dilution air is introduced into a turbine nozzle guide vane 80 from which it leaves via ports 100, 101 in the external surfaces 110, 112 extending between the leading and trailing edges 81,82 of the nozzle guide vane 80.
Ports 100, 101 are spaced in a radial direction R relative to the axis of rotation A of the turbine rotor. Proper location of these ports will help in maintaining required temperature traverse across the turbine blade, i.e. the holes are configured so as to provide a predetermined radial distribution of gas temperature at the nozzle guide vane exit plane P. Figure 4A is a sectional view through the nozzle guide vane 80 along line BB in figure 3. Ports 100 arc spaccd along thc cxtcrnal surfaccs 110, 112 bctwccn thc lcading and trailing cdgcs 81, 82 and, as shown in thc dctail vicw of figurc 4C, may bc oricntcd at an anglc (T) relative to the surface to prornote flow substantially tangential to the surface as indicated by arrows 60. Such tangential flow of the dilution air will cause less disturbance of the gas flow than in the arrangement of GB774500. Ports (or a single slot) in the leading and trailing edges 8 1,82 of the nozzle guide vane are similarly shaped and sized to not disturb downstream flow.
By feeding a substantial amount of dilution air through the nozzle guide vanes it is possible to partially or indeed completely dispense with a dilution zone in the combustor.
Such an arrangemeilt, shown in figure 3, has a length (and consequently weight) that is lower than that of the conventional combustor design of figure 1. An even shorter combustor may be achieved using hydrogen ifielled micromix combustion which has a S shorter flame length than conventional combustors.
It should be understood that this invention has been descnbed by way of examples only and that a wide variety of modifications can be made without departing from the scope of the invention.

Claims (6)

  1. CLAIMS1. A gas turbine having a compressor, a combustor and a turbine; the turbine comprising a nozzle guide vane having leading and trailing edges linked by external surfaces; the compressor providing a first gas flow for stoichiometric combustion of fuel in the combustor and a second gas flow for dilution of the products of said stoichiometric combustion, wherein the second gas flow is exhausted through at least one port in at least one of 10 said external surfaces of the nozzle guide vanes.
  2. 2. Gas turbine according to claim 1, wherein the second gas is exhausted through a plurality of ports in one of the external surfaces of the nozzle guide vanes.
  3. 3. Gas turbine according to claim 2, wherein the ports are spaced in a radial direction relative to the axis of rotation of the turbine rotor.
  4. 4. Gas turbine according to claim 3, wherein the ports are configured so as to provide a prcdetcrmincd radial distribution of gas temperature at thc nozzle guide vane exit plane.
  5. 5. Gas turbine according to any preceding claim, wherein the ports are spaced along the external surface between the leading and trailing edges.
  6. 6. Gas turbine according to any preceding claim, wherein at least one of the plurality of ports is oriented at an angle relative to the surface to promote flow substantially tangential to the surface.
GB1209016.3A 2012-05-22 2012-05-22 Gas turbine nozzle guide vane with dilution air exhaust ports Pending GB2502302A (en)

Priority Applications (1)

Application Number Priority Date Filing Date Title
GB1209016.3A GB2502302A (en) 2012-05-22 2012-05-22 Gas turbine nozzle guide vane with dilution air exhaust ports

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB1209016.3A GB2502302A (en) 2012-05-22 2012-05-22 Gas turbine nozzle guide vane with dilution air exhaust ports

Publications (2)

Publication Number Publication Date
GB201209016D0 GB201209016D0 (en) 2012-07-04
GB2502302A true GB2502302A (en) 2013-11-27

Family

ID=46546494

Family Applications (1)

Application Number Title Priority Date Filing Date
GB1209016.3A Pending GB2502302A (en) 2012-05-22 2012-05-22 Gas turbine nozzle guide vane with dilution air exhaust ports

Country Status (1)

Country Link
GB (1) GB2502302A (en)

Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5577884A (en) * 1984-03-14 1996-11-26 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Structure for a stationary cooled turbine vane
EP1061236A2 (en) * 1999-06-15 2000-12-20 Mitsubishi Heavy Industries, Ltd. Gas turbine stationary blade
EP1209323A2 (en) * 2000-11-28 2002-05-29 Nuovo Pignone Holding S.P.A. Cooling system for gas turbine stator vanes
EP1262631A2 (en) * 2001-05-21 2002-12-04 United Technologies Corporation Film cooled blade or vane
GB2405451A (en) * 2003-08-23 2005-03-02 Rolls Royce Plc A vane assembly for a gas turbine engine
EP2236752A2 (en) * 2009-04-03 2010-10-06 Rolls-Royce plc Cooled aerofoil for a gas turbine engine
US7862299B1 (en) * 2007-03-21 2011-01-04 Florida Turbine Technologies, Inc. Two piece hollow turbine blade with serpentine cooling circuits
JP2011064207A (en) * 2005-03-30 2011-03-31 Mitsubishi Heavy Ind Ltd High temperature member for gas turbine

Patent Citations (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5577884A (en) * 1984-03-14 1996-11-26 Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. Structure for a stationary cooled turbine vane
EP1061236A2 (en) * 1999-06-15 2000-12-20 Mitsubishi Heavy Industries, Ltd. Gas turbine stationary blade
EP1209323A2 (en) * 2000-11-28 2002-05-29 Nuovo Pignone Holding S.P.A. Cooling system for gas turbine stator vanes
EP1262631A2 (en) * 2001-05-21 2002-12-04 United Technologies Corporation Film cooled blade or vane
GB2405451A (en) * 2003-08-23 2005-03-02 Rolls Royce Plc A vane assembly for a gas turbine engine
JP2011064207A (en) * 2005-03-30 2011-03-31 Mitsubishi Heavy Ind Ltd High temperature member for gas turbine
US7862299B1 (en) * 2007-03-21 2011-01-04 Florida Turbine Technologies, Inc. Two piece hollow turbine blade with serpentine cooling circuits
EP2236752A2 (en) * 2009-04-03 2010-10-06 Rolls-Royce plc Cooled aerofoil for a gas turbine engine

Also Published As

Publication number Publication date
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