USH903H - Cool tip combustor - Google Patents

Cool tip combustor Download PDF

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Publication number
USH903H
USH903H US06/374,355 US37435582A USH903H US H903 H USH903 H US H903H US 37435582 A US37435582 A US 37435582A US H903 H USH903 H US H903H
Authority
US
United States
Prior art keywords
turbine
combustor
wall
section
cooling
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
US06/374,355
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English (en)
Inventor
Barry Weinstein
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
Assigned to GENERAL ELECTRIC COMPANY, A NY CORP. reassignment GENERAL ELECTRIC COMPANY, A NY CORP. ASSIGNMENT OF ASSIGNORS INTEREST. Assignors: WEINSTEIN, BARRY
Priority to US06/374,355 priority Critical patent/USH903H/en
Priority to CA000414954A priority patent/CA1194803A/fr
Priority to IL67317A priority patent/IL67317A/xx
Priority to GB08234168A priority patent/GB2119861B/en
Priority to IT24980/82A priority patent/IT1153896B/it
Priority to JP57227827A priority patent/JPS58195027A/ja
Priority to FR8221913A priority patent/FR2526083B1/fr
Priority to DE19823248439 priority patent/DE3248439A1/de
Publication of USH903H publication Critical patent/USH903H/en
Application granted granted Critical
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
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on 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
    • 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
    • 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
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/06—Arrangement of apertures along the flame tube
    • 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

  • This invention relates to means for directing cooling air to critical parts of hot section turbine blades in gas turbine engines.
  • blade cooling configurations are generally restricted to fairly simple designs because of small dimensions and limitations of current manufacturing technologies. The implication is a typical smaller engine turbine blade or vane cannot be provided with the highly complex, internal air cooling passage configuration typically used today in larger gas turbine engines.
  • an object of the present invention to provide a means for cooling tips of turbine blades in turbine sections of gas turbine engines with a system that can be utilized in relatively small engine configurations.
  • Another object of the present invention is to provide a source of cooling air that can be directed specifically to turbine blade tips in a turbine section of a small gas turbine engine.
  • means are provided for introducing cooling air into a turbine section of a gas turbine engine in the region of tip sections of turbine blades.
  • the source of this cooling air is compressor discharge air that has bypassed the combustor.
  • This compressor discharge air is introduced at an aft section of the combustor through inlet air holes just upstream of the turbine section.
  • the air is introduced along a radially outer section of the combustor only.
  • the cooling air flows initially into annulus regions within the combustor that are protected from hot combustion gases. From these annulus regions, the air flows downstream in the combustor and forms a thick film that blankets the combustor wall.
  • FIG. 1 is a schematic cross-sectional illustration of a central section of a gas turbine engine.
  • FIG. 2 is a schematic cross-sectional illustration of a combustor and high pressure turbine section of a gas turbine engine with the present invention embodied therein.
  • FIG. 3 is a cross-sectional illustration of a downstream portion of a combustor wall with one embodiment of part of the present invention incorporated therein.
  • FIG. 4 is a graphical representation of test results of turbine blade temperatures.
  • FIG. 1 a central section of a typical gas turbine engine 10 is shown that involves substantial turbomachinery that rotates about an engine centerline 12.
  • Components of this turbomachinery include, in serial flow relationship, a compressor 14, a combustor 16, a high-pressure turbine section 18, and a low-pressure turbine section 20.
  • inlet air is directed into and pressurized by the compressor 14 from which the air is discharge through a diffuser 22.
  • a major portion of this compressor discharge air is then passed into the combustor 16 where it is mixed with fuel and vaporized to form high-pressure, high-temperature combustion gases which flow downstream into the high-pressure turbine 18.
  • the high-pressure gases cause turbine blades 24 in the high-pressure turbine 18 to rotate at high velocities thereby providing mechanical power. These high-temperature, high-pressure gases then continue to flow downstream into the low-pressure turbine 20 where they cause low-pressure turbine blades 26 to rotate thereby providing additional mechanical power. From the low-pressure turbine 20, the gases are discharged downstream so as to pass out of the engine 10.
  • a portion of the air discharge from the compressor 14 that passes through the diffuser 22 is circulated to cool a variety of hot pairs of the engine 10. Some of that air used for cooling flows to the region of the combustor 16 and surrounds the combustor walls. In some engines, small cooling holes are provided in the combustor walls so that cooling air can enter the combustor to cool interior combustor surfaces. Other portions of the cooling air are directed internally to hot temperature parts inside the high-pressure turbine 18. A part of this air used to cool the high-pressure turbine is directed into the interior of a high-pressure turbine nozzle 28 so as to provide an internal cooling function by impingement and diffusion processes. Another part of the compressor discharge air is directed along other paths to cool interior regions of the turbine blades 24 of the high-pressure turbine 18. These cooling flowpaths are generally represented by the dark arrows in FIG. 1.
  • a prior art solution to the problem of cooling turbine blade tips is a ducting of a small portion of the compressor discharge cooling air into the high-pressure turbine 18 at an inlet location 32 just downstream of the turbine nozzle 28. Cooling air ducted in this manner would bypass the combustor and flow into the high-pressure turbine 18 just upstream of the turbine blade 24.
  • This proposal reduces turbine tip temperatures, but this approach also has a negative impact on engine performance, both in terms of thrust and fuel consumption.
  • the detrimental effect on engine performance is caused because the cooling air enters the gas flow stream behind the first-stage turbine nozzle 28 and is, therefore, chargeable to the engine's thermodynamic cycle. As a result, the amount of air burned compared to the allowable turbine rotor inlet temperature level is reduced and engine performance decreases.
  • FIG. 2 a portion of a gas turbine engine 11 is shown that is generally similar to part of the engine shown in FIG. 1 but, this time, incorporating an embodiment of the present invention.
  • a portion of the cooling air discharged from the compressor does not enter the combustor 16, but, instead, flows downstream around the combustor as indicated by dark arrows in FIG. 2.
  • This cooling air does not undergo the mixing and combustion processes occurring during engine operation inside the combustor 16. Because the air does not undergo combustion, it remains relatively cool and serves as a source of high-pressure cooling air that can be utilized in the high-pressure turbine sections of the engine.
  • any cooling air used in the high-pressure turbine section must be at high pressure because the internal gases flowing through the high-pressure turbine area, as the name suggests, are at very high pressure.
  • the cooling air introduced into the high-pressure turbine must be even higher in pressure than those gases flowing through the turbine so that the cooling air will be caused by its own pressure forces to flow into the turbine blades and vanes and from there into the combustion gas flow passage of the turbine section. If the cooling air that was used for cooling in this region were lower in pressure than the combustion gases flowing through the turbine section, pressure forces would not permit the cooling air to flow from interior regions of the turbine blades and vanes out into the combustion gas flow passage.
  • cooling air is introduced at an inlet location 32 immediately downstream of the turbine nozzle 28, the air will tend to cool the turbine blade tips 20. However, because the air has not been expanded and directed by the turbine nozzle 28, it will not be useful for providing appropriate gas forces for causing the turbine blades 24 to rotate.
  • the present invention comprises means for introducing cooling air forward or upstream of the first-stage turbine nozzle 28 so that there is no associated engine performance penalty.
  • This means is shown in FIG. 2, and a part of the invention is shown in larger scale in FIG. 3.
  • a portion of the compressor discharge air that is flowing outside of the combustor 16 is directed into combustor inlet air holes 36 at a location just upstream of the turbine nozzle 28.
  • the air is introduced at a location just upstream of the turbine nozzle 28, partly to prevent that cooling air from undergoing the normal combustion processes inside the turbine 16, and also to lessen heating of the cooling air from prolonged exposure to the hot combustion gases. If this cooling air were to undergo combustion, it would climb dramatically in temperature and be rendered relatively useless for the purpose of cooling tips 30 of turbine blades.
  • FIG. 3 the inlet air holes 36 through which the cooling air is directed into a downstream section of the combustor 16 are shown in greater detail.
  • a portion of a radially outer wall section 38 of the combustor 16 is shown in FIG. 3. This portion of the combustor wall section 38 is located just upstream of the turbine nozzle 28 (not shown). In the cross-sectional view shown, three inlet air holes 36 can be seen and their relative configuration can be appreciated. It should first be noted that the downstream portion of the combustor wall section 38 is actually double-walled. An outer wall section 40 connects to the turbine nozzle in a standard manner as would be the normal practice in many gas turbine engines.
  • An inner combustor wall section 42 is provided and is protected from hot combustion gases at its upstream end by a flange 44. At its downstream end, the inner wall section 42 extends almost to the turbine nozzle inlet. Cooling air from the compressor discharge is bled into annulus regions 46 that are open in a downstream direction and are generally protected from the combustion occurring inside the combustor 16. Because the cooling air is bled into these protected annulus regions 46, the cooling air does not undergo combustion, and the air enters the turbine nozzle at substantially compressor discharge temperature, thereby forming a thick, low-temperature film along a radially outer wall of the turbine flow path.
  • each of the holes 36 represents one of a row of holes that extend around the entire circumference of the radially outer wall section 38 of the combustor 16.
  • the total number of inlet air holes 36 could vary widely as could their general configuration.
  • a row of upstream inlet air holes 48 is provided to bleed cooling air into the annulus region between the flange 44 and the inner wall 42.
  • a row of intermediate inlet air holes 50 is provided to bleed additional cooling air into the annulus region between the inner wall 42 and the outer wall 40.
  • a row of downstream inlet air holes 52 is provided to direct additional cooling air into the annulus between the inner wall 42 and the outer wall 40.
  • FIG. 4 a comparison of test results is shown that graphically represents turbine blade temperatures in a typical gas turbine engine and, additionally, represents turbine blade temperatures in a second gas turbine engine incorporating the present invention.
  • the X (horizontal) coordinate in FIG. 4 is marked off in degrees Fahrenheit.
  • the Y (vertical) coordinate in FIG. 4 is a dimensionless representation of turbine blade height, beginning at a root of the turbine blade and ending at the tip of the turbine blade.
  • the lines shown on the graph of FIG. 4 designated 54 represent turbine blade temperatures in two typical gas turbine engines, generally having an engine configuration similar to that shown in FIG. 2 but without incorporating the present invention.
  • the line designated 56 in FIG. 4 represents turbine blade temperature, again within an engine having generally the same configuration as shown in FIG.

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)
US06/374,355 1982-05-03 1982-05-03 Cool tip combustor Abandoned USH903H (en)

Priority Applications (8)

Application Number Priority Date Filing Date Title
US06/374,355 USH903H (en) 1982-05-03 1982-05-03 Cool tip combustor
CA000414954A CA1194803A (fr) 1982-05-03 1982-11-05 Turbine a extremites d'aubes refroidies
IL67317A IL67317A (en) 1982-05-03 1982-11-21 Gas turbine engine with means for cooling turbine blade tips
GB08234168A GB2119861B (en) 1982-05-03 1982-12-01 Cooling high pressure turbine blade tips in a gas turbine engine
IT24980/82A IT1153896B (it) 1982-05-03 1982-12-24 Sistema di raffreddamento di palette di turbina a gas
JP57227827A JPS58195027A (ja) 1982-05-03 1982-12-28 ガスタービンエンジン
FR8221913A FR2526083B1 (fr) 1982-05-03 1982-12-28 Moteur a bouts d'aubes de turbine refroidis
DE19823248439 DE3248439A1 (de) 1982-05-03 1982-12-29 Gasturbinentriebwerk mit gekuehlten schaufelspitzen

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US06/374,355 USH903H (en) 1982-05-03 1982-05-03 Cool tip combustor

Publications (1)

Publication Number Publication Date
USH903H true USH903H (en) 1991-04-02

Family

ID=23476435

Family Applications (1)

Application Number Title Priority Date Filing Date
US06/374,355 Abandoned USH903H (en) 1982-05-03 1982-05-03 Cool tip combustor

Country Status (8)

Country Link
US (1) USH903H (fr)
JP (1) JPS58195027A (fr)
CA (1) CA1194803A (fr)
DE (1) DE3248439A1 (fr)
FR (1) FR2526083B1 (fr)
GB (1) GB2119861B (fr)
IL (1) IL67317A (fr)
IT (1) IT1153896B (fr)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5398496A (en) * 1993-03-11 1995-03-21 Rolls-Royce, Plc Gas turbine engines
US5749218A (en) * 1993-12-17 1998-05-12 General Electric Co. Wear reduction kit for gas turbine combustors
US6494678B1 (en) 2001-05-31 2002-12-17 General Electric Company Film cooled blade tip
US20090060723A1 (en) * 2007-08-31 2009-03-05 Snecma separator for feeding cooling air to a turbine
US20110000218A1 (en) * 2008-02-27 2011-01-06 Mitsubishi Heavy Industries, Ltd. Gas turbine and method of opening chamber of gas turbine
US20110129336A1 (en) * 2008-05-29 2011-06-02 Snecma Assembly including a turbine disk for a gas turbine engine and a bearing-supporting journal, and cooling circuit for the turbine disk of such an assembly
US10954796B2 (en) * 2018-08-13 2021-03-23 Raytheon Technologies Corporation Rotor bore conditioning for a gas turbine engine
US20220018540A1 (en) * 2020-07-15 2022-01-20 Raytheon Technologies Corporation Deflector for conduit inlet within a combustor section plenum

Families Citing this family (6)

* 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
US4739621A (en) * 1984-10-11 1988-04-26 United Technologies Corporation Cooling scheme for combustor vane interface
US4785623A (en) * 1987-12-09 1988-11-22 United Technologies Corporation Combustor seal and support
US5188510A (en) * 1990-11-21 1993-02-23 Thomas R. Norris Method and apparatus for enhancing gas turbo machinery flow
US5178660A (en) * 1991-06-26 1993-01-12 Libbey-Owens-Ford Co. Apparatus for bending glass sheets
EP2837769B1 (fr) 2013-08-13 2016-06-29 Alstom Technology Ltd Arbre de rotor pour turbomachine

Family Cites Families (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2541171A (en) * 1947-01-25 1951-02-13 Kellogg M W Co Air inlet structure for combustion chambers
CH323720A (de) * 1954-05-20 1957-08-15 Sulzer Ag Gekühlte Gasturbine
DE1270889B (de) * 1964-05-13 1968-06-20 Rolls Royce Kuehlvorrichtung fuer Brennkammern von Gasturbinentriebwerken
US3418808A (en) * 1966-07-05 1968-12-31 Rich David Gas turbine engines
GB1193587A (en) * 1968-04-09 1970-06-03 Rolls Royce Nozzle Guide Vanes for Gas Turbine Engines.
CA982828A (en) * 1972-06-01 1976-02-03 General Electric Company Combustor casing cooling structure
US3965066A (en) * 1974-03-15 1976-06-22 General Electric Company Combustor-turbine nozzle interconnection
GB1550368A (en) * 1975-07-16 1979-08-15 Rolls Royce Laminated materials
DE2643049A1 (de) * 1975-10-14 1977-04-21 United Technologies Corp Schaufel mit gekuehlter plattform fuer eine stroemungsmaschine
GB1524956A (en) * 1975-10-30 1978-09-13 Rolls Royce Gas tubine engine
DE2810240C2 (de) * 1978-03-09 1985-09-26 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Verstelleitgitter für axial durchströmte Turbinen, insbesondere Hochdruckturbinen von Gasturbinentriebwerken
FR2490728A1 (fr) * 1980-09-25 1982-03-26 Snecma Dispositif de refroidissement par film d'air pour tube a flamme de moteur a turbine a gaz

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5398496A (en) * 1993-03-11 1995-03-21 Rolls-Royce, Plc Gas turbine engines
US5749218A (en) * 1993-12-17 1998-05-12 General Electric Co. Wear reduction kit for gas turbine combustors
US6494678B1 (en) 2001-05-31 2002-12-17 General Electric Company Film cooled blade tip
US20090060723A1 (en) * 2007-08-31 2009-03-05 Snecma separator for feeding cooling air to a turbine
US8069669B2 (en) * 2007-08-31 2011-12-06 Snecma Separator for feeding cooling air to a turbine
US20110000218A1 (en) * 2008-02-27 2011-01-06 Mitsubishi Heavy Industries, Ltd. Gas turbine and method of opening chamber of gas turbine
US9080464B2 (en) * 2008-02-27 2015-07-14 Mitsubishi Hitachi Power Systems, Ltd. Gas turbine and method of opening chamber of gas turbine
US20110129336A1 (en) * 2008-05-29 2011-06-02 Snecma Assembly including a turbine disk for a gas turbine engine and a bearing-supporting journal, and cooling circuit for the turbine disk of such an assembly
US8899913B2 (en) 2008-05-29 2014-12-02 Snecma Assembly including a turbine disk for a gas turbine engine and a bearing-supporting journal, and cooling circuit for the turbine disk of such an assembly
US10954796B2 (en) * 2018-08-13 2021-03-23 Raytheon Technologies Corporation Rotor bore conditioning for a gas turbine engine
US20220018540A1 (en) * 2020-07-15 2022-01-20 Raytheon Technologies Corporation Deflector for conduit inlet within a combustor section plenum
US11371700B2 (en) * 2020-07-15 2022-06-28 Raytheon Technologies Corporation Deflector for conduit inlet within a combustor section plenum

Also Published As

Publication number Publication date
IT8224980A0 (it) 1982-12-24
DE3248439A1 (de) 1983-11-03
FR2526083B1 (fr) 1987-11-27
JPS58195027A (ja) 1983-11-14
IT1153896B (it) 1987-01-21
GB2119861A (en) 1983-11-23
IL67317A (en) 1989-10-31
FR2526083A1 (fr) 1983-11-04
IL67317A0 (en) 1983-03-31
GB2119861B (en) 1985-07-31
IT8224980A1 (it) 1984-06-24
JPH0415378B2 (fr) 1992-03-17
CA1194803A (fr) 1985-10-08

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Owner name: GENERAL ELECTRIC COMPANY, A NY CORP.

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST.;ASSIGNOR:WEINSTEIN, BARRY;REEL/FRAME:003989/0787

Effective date: 19820429

STCF Information on status: patent grant

Free format text: PATENTED CASE