US3877221A - Combustion apparatus air supply - Google Patents

Combustion apparatus air supply Download PDF

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Publication number
US3877221A
US3877221A US391888A US39188873A US3877221A US 3877221 A US3877221 A US 3877221A US 391888 A US391888 A US 391888A US 39188873 A US39188873 A US 39188873A US 3877221 A US3877221 A US 3877221A
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US
United States
Prior art keywords
wall
liner
diffuser
air
walls
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.)
Expired - Lifetime
Application number
US391888A
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English (en)
Inventor
Arthur H Lefebvre
Harold L Stocker
Samuel B Reider
Jerry G Tomlinson
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.)
Motors Liquidation Co
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Motors Liquidation 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 Motors Liquidation Co filed Critical Motors Liquidation Co
Priority to US391888A priority Critical patent/US3877221A/en
Priority to CA196,154A priority patent/CA993668A/en
Priority to GB1814274A priority patent/GB1425438A/en
Priority to DE2422362A priority patent/DE2422362B2/de
Priority to IT50887/74A priority patent/IT1011411B/it
Priority to FR7418267A priority patent/FR2245855B1/fr
Priority to JP49058891A priority patent/JPS5045111A/ja
Application granted granted Critical
Publication of US3877221A publication Critical patent/US3877221A/en
Priority to JP1977132889U priority patent/JPS5395004U/ja
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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    • 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
    • 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

  • ABSTRACT An improvement in diffusers to guide air flow from the compressor of a gas turbine engine into the combustion apparatus, splitting the air between front, outer, and inner walls of the combustion liner. A slot in a wall of the diffuser transfers air from compressor mid-radius to the passage leading to the liner inner wall.
  • Our invention is directed to combustion apparatus for gas turbine engines. and particularly to an inlet diffuser structure adapted to divide the air flowing to the apparatus between several flow paths. More specifically. our invention resides in diffusing structure which is placed ahead of the combustion zone of an annular gas turbine combustor to divide the air between a portion flowing to the forward wall of the liner. a portion flowing to the radially outer wall. and a portion flowing to the radially inner wall.
  • Our invention is particularly adapted to maintain the desired division of flow between the several paths leading into the combustion apparatus notwithstanding changes in operating conditions of the engine which alter the pressure distribution profile at the outlet of a compressor which supplies the combustion apparatus with air.
  • Our apparatus is adapted to be employed at relatively high inlet velocity of the air with a broad operating range and a high pressure recovery for low pressure drop through the combustion apparatus. It is characterized by prevention of localized flow separation and by control of the air so as to maintain a good burner outlet pattern.
  • the principal objects ofour invention are to improve the operation of gas turbine combustion apparatus. particularly those of the annular type: to improve the inlet diffusing arrangements of such combustion apparatus; and to provide a combustion apparatus which is relatively insensitive to variations in the pressure and velocity profiles of air delivered to the combustion apparatus by a compressor which supplies the air for combustion.
  • FIG. 1 is a sectional view of the combustion apparatus of a gas turbine engine taken in a plane containing the axis of the engine and illustrating the environment of our invention.
  • FIG. 2 is an enlarged view corresponding to FIG. 1 of the diffuser and upstream end of the combustion apparatus.
  • a typical axial-flow gas turbine engine 2 which is only partly shown. includes. in flow series, an axial-flow compressor 3, combustion apparatus 4. and a turbine 6. Only the discharge end of the compressor is illustrated. and only a portion of the turbine nozzle 7 through which the combustion products flow is illustrated.
  • the turbine is connected by a shaft to drive the compressor to force compressed air into the combustion apparatus. Fuel is burned in the air so supplied and the resulting combustion products are fed to the turbine to drive the compressor. Power may be taken off as shaft power or as a pressurized exhaust stream for jet propulsion.
  • the general structure of such engines is well known. and there is no need to describe such in greater detail to explain our invention.
  • the combustion apparatus 4 which is shown by way of example comprises an outer casing or wall 8 and an inner casing or wall 9. these defining between them an annular space extending from the outlet of the compressor to the inlet of the turbine. Combustion takes place in an annular combustion liner 10 disposed between the outer and inner walls.
  • the combustion liner comprises an outer wall ll and an inner wall 12. These being approximately cylindrical and slightly tapered.
  • An air passage 13 is defined by walls 8 and 11, and an air passage 14 by walls 9 and 12.
  • the liner 10 also comprises a ring-shaped forward wall 15 fixed to and joining the outer and inner walls of the liner.
  • Fuel nozzles 16 are mounted on struts 18 extending through the outer wall 8. through which fuel is supplied to the nozzle. In a particular example. there are sixteen such nozzles disposed equally aroung the axis of the engine.
  • the multistage compressor so far as illustrated. comprises a final rotor stage 22 and two rows of outlet guide vanes 23.
  • the compressor delivers that air in an axial direction into the combustion apparatus. specifically delivering air through the forward. outer. and inner walls of the liner.
  • the diffuser comprises an outer diffuser wall 26 of sheet metal the rear edge of which overlaps and is welded or brazed to the forward portion of the liner outer wall 11.
  • the forward end of outer diffuser wall 26 abuts and is welded to a ring 27 forming the outer boundary of an annular central air entrance 28.
  • the inner boundary of this air entrance is defined by a ring 30.
  • the inner diffuser wall 34 comprises a forward section 32 which is attached to the rear edge of the ring 30 and an aft section the rear end of which overlaps and is fixed to the forward end of the liner inner wall 12. Sections 32 and 35 of the inner diffuser wall 34 are overlapped and spaced to define a communicating passage 36, the parts being connected and maintained in proper relation by spacers 38 of any suitable structure distributed around the axis of the diffuser. As will be seen, the walls 26 and 34 define between them a rather large cavity or air space 39 to which air is introduced through a diverging diffuser passage 40 defined between rings 27 and 30.
  • the assembly of rings 27 and 30 may be characterized as a snout 42.
  • this snout projects very close to the outlet of the compressor; that is. the outlet guide vanes 23. and that it serves to split the air discharged by the compressor into three portions, one flowing through the diffusing passage 40, a second flowing through an outer diffusing passage 43 between walls 8 and 26 into passage 13, and a third flowing through an inner diffusing passage 44 between walls 30. 32. and 35 and the wall 9 into passage 14.
  • the means by which the diffuser 24 and the forward end of liner 10 are supported are not material to the invention. but it may be mentioned that they are attached to circumferentially spaced struts (not illustrated) extending between walls 8 and 9 abreast of snout 42. the forward edge of which is recessed to clear these struts.
  • the outer and inner walls 11 and 12 of the combustion liner are similar. Each consists of a number of overlapping sections which define narrow gaps at the over-laps between them for flow of air which performs the function of convection cooling of the liner and then flows over the heated surface of the wall for film cooling.
  • Primary combustion air is admitted through holes near the forward end of the liner wall as indicated by the arrows 45. Secondary or dilution air flows through holes in the walls farther downstream as indicated by the arrows 46. Some primary combustion air is admitted through the fuel nozzle 16 to atomize the fuel. Additional air is admitted through the forward wall and flows between the margins of baffles 47 and the outer and inner walls to serve as film cooling air for the up stream end of the liner.
  • the film cooling air so intro quizted may become primary combustion air and. so far as any of this film cooling air is not combined with the fuel. it ultimately becomes dilution air. So far as our invention is concerned. any suitable structure of the forward. outer. and inner walls of the liner may be employed.
  • the strut 18 has a circumferential flange 50 which is disposed with some clearance in a hole 51 in wall 26.
  • a hat-section sheet metal ring or ferrule 52 is assembled around flange 50. The outer margin of the ferrule is slidably retained between the inner surface of wall 26 and a flanged retain ing ring 54 welded to wall 26. This structure minimizes air leakage around strut 18.
  • our invention is directed to the diffuser structure 24 which diffuses and divides the air going to the three liner walls.
  • the air may be divided into approximately 44% flowing into passage 13. including about 4% flowing past barrier 19 for turbine cooling;
  • passage 14 including some 6% flowing past barrier 20 for turbine cooling; and about l6% through the front liner wall 15.
  • the amount entering through the central air inlet 28 and flowing through diffusing passage 40 is. in this case. approximately 23% of the total air and about 7% of the air is redirected under normal operating-conditions through the air communicating passage 36 into the inner passage 14.
  • the exact proportion of air which is diverted from cavity 39 through the communicating passage to the inner passage 14 varies with the change in outlet velocity and total pressure profiles of the-air entering the combustion apparatus through vanes 23-of the compressor. If the pressure near the inner boundary of the compressor discharge passage decreases the effect of this is to decrease both the quantity and pressure of air flowing through the inner passage 14. while. at the same time it creates an increase in pressure differential between cavity 39 and the exit of the inner diffusing passage 44. as a result of which an increased amount of air flows through the communicating passage 36 from cavity 39 to inner passage 14 thereby increasing both pressure and air flow quantity in passage l4.'ln this manner the communicating passage compensates for variations in compressor discharge velocity and pressure profiles.
  • the diffusing passage 40 is directed much more toward the inner margin of the diffuser'assembly 24 than its'outer margin and thus the air flows more directly toward the communicating passage 36 into inner passage 14. Because of the great enlargement of the passage as it leaves passage 40 and enters cavity 39, the air in cavity 39 is rapidly diffused and thus there are no significant velocity effects to upset distribution of flow through the nozzles 16 and the cooling air inlets between the baffles 47 and the liner walls. Thus. there is a fast-moving stream of air flowing over the outer surface of both walls of the liner and a relatively slower moving stream of air within the cavity 39.
  • a gas turbine engine comprising a compressor with an annular axial outlet. outer and inner casing walls defining an annular combustion space connected to the outlet. and an annular combustion liner disposed between the said casing walls. the casing walls having forward portions diverging to define a diffusing zone beginning at the compressor outlet. the combustion liner having an upstream wall and outer and inner walls with openings for entrance of'air into the liner in each of the said liner walls. the improvement comprising an airflow dividing diffuser disposed in the diffusing zone and connected to the upstream end of the liner. the diffuser having a snout providing an annular entrance adjacent the compressor outlet and at a mid-radius of the outlet.
  • the outer diffuser wall and outer casing wall defining a diffusing flow path to the outer liner wall.
  • the outer diffuser wall blocking flow between the said air space and the outer liner wall. and the inner casing wall defining a diffusing flow path to the inner liner wall; the said diffusing flow paths taking air from the outer. central. and inner annular zones of the compressor outlet. respectively; and the diffuser defining means for diverting a portion of the air from the path leading to the upstream liner wall into the path leading to the inner liner wall responsive to relative pressure conditions in the said paths.
  • a gas turbine engine comprising a compressor with an annular axial outlet. outer and inner casing walls defining an annular combustion space connected to the outlet. and an annular combustion liner disposed between the said casing walls. the casing walls having forward portions diverging to define a diffusing zone beginning at the compressor outlet. the combustion liner having an upstream wall and outer and inner walls with openings for entrance of air into the liner in each of the said liner walls. the improvement comprising an airflow dividing diffuser disposed in the diffusing zone and connected to the upstream end of the liner. the diffuser having a snout providing an annular entrance adjacent the compressor outlet and at a mid-radius of the outlet.
  • the outer diffuser wall and outer casing wall defining a diffusing flow path to the outer liner wall.
  • the outer diffuser wall blocking flow between the said air space and the outer liner wall. and the inner casing wall defining a diffusing flow path to the inner liner wall; the said diffusing flow paths taking air from the outer. central. and inner annular zones of the compressor outlet. respectively; and the inner diffuser wall defining circumferentially extending slot means for diverting a portion of the air from the path leading to the upstream liner wall into the path leading to the inner liner wall in response to relative pressure conditions in the said paths.
  • a gas turbine engine comprising a compressor with an annular axial outlet. outer and inner casing walls defining an annular combustion space connected to the outlet. and an annular combustion liner disposed between the said casing walls. the casing walls having forward portions diverging to define a diffusing zone beginning at the compressor outlet. the combustion liner having an upstream wall and outer and inner walls with openings for entrance of air into the liner in each of the said liner walls. the improvement comprising an airflow dividing diffuser disposed in the diffusing zone and connected to the upstream end of the liner. the diffuser having a snout providing an annular entrance adjacent the compressor outlet and at a mid-radius of the outlet.
  • the outer diffuser wall and outer casing wall defining a diffusing flow path to the outer liner wall.
  • the outer diffuser wall blocking flow between the said air space and the outer liner wall. and the inner casing wall defining a diffusing flow path to the inner liner wall; the said diffusing flow paths taking .air from the outer. central. and inner annular zones ofthe compressor outlet.
  • the inner diffuser wall defining circumferentially extending slot.means for diverting a portion of the air from the path leading to the upstream liner wall into the path leading to the inner liner wall in response to relative pressure conditions in the said paths; and the diffusing flow path to the upstream liner wall being disposed to discharge the air into the said air space in a direction more predominantly toward the inner diffuser wall than toward the outer diffuser wall.
US391888A 1973-08-27 1973-08-27 Combustion apparatus air supply Expired - Lifetime US3877221A (en)

Priority Applications (8)

Application Number Priority Date Filing Date Title
US391888A US3877221A (en) 1973-08-27 1973-08-27 Combustion apparatus air supply
CA196,154A CA993668A (en) 1973-08-27 1974-03-27 Combustion apparatus air supply for gas turbine engine
GB1814274A GB1425438A (en) 1973-08-27 1974-04-25 Combustion apparatus air supply of gas turbine engines
DE2422362A DE2422362B2 (de) 1973-08-27 1974-05-06 Ringbrennkammer für ein Gasturbinentriebwerk
IT50887/74A IT1011411B (it) 1973-08-27 1974-05-09 Perfezionamento nei dispositivi di alimentazione dell aria per combustori
FR7418267A FR2245855B1 (ja) 1973-08-27 1974-05-27
JP49058891A JPS5045111A (ja) 1973-08-27 1974-05-27
JP1977132889U JPS5395004U (ja) 1973-08-27 1977-10-04

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US391888A US3877221A (en) 1973-08-27 1973-08-27 Combustion apparatus air supply

Publications (1)

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US3877221A true US3877221A (en) 1975-04-15

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ID=23548373

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US391888A Expired - Lifetime US3877221A (en) 1973-08-27 1973-08-27 Combustion apparatus air supply

Country Status (7)

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US (1) US3877221A (ja)
JP (2) JPS5045111A (ja)
CA (1) CA993668A (ja)
DE (1) DE2422362B2 (ja)
FR (1) FR2245855B1 (ja)
GB (1) GB1425438A (ja)
IT (1) IT1011411B (ja)

Cited By (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4098074A (en) * 1976-06-01 1978-07-04 United Technologies Corporation Combustor diffuser for turbine type power plant and construction thereof
US4458479A (en) * 1981-10-13 1984-07-10 General Motors Corporation Diffuser for gas turbine engine
US4870826A (en) * 1987-06-18 1989-10-03 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma) Casing for a turbojet engine combustion chamber
US5077967A (en) * 1990-11-09 1992-01-07 General Electric Company Profile matched diffuser
US5134855A (en) * 1989-12-15 1992-08-04 Rolls-Royce Plc Air flow diffuser with path splitter to control fluid flow
US5279126A (en) * 1992-12-18 1994-01-18 United Technologies Corporation Diffuser-combustor
US5339622A (en) * 1992-08-19 1994-08-23 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) Gas turbine engine with improved water ingestion prevention
US5737915A (en) * 1996-02-09 1998-04-14 General Electric Co. Tri-passage diffuser for a gas turbine
US6401447B1 (en) 2000-11-08 2002-06-11 Allison Advanced Development Company Combustor apparatus for a gas turbine engine
GB2390890A (en) * 2002-07-17 2004-01-21 Rolls Royce Plc Diffuser for gas turbine engine
EP1426688A1 (en) * 2002-11-19 2004-06-09 General Electric Company Combustor inlet diffuser with boundary layer blowing
US20180266690A1 (en) * 2017-03-14 2018-09-20 Safran Aircraft Engines Combustion chamber of a turbine engine
US20220373181A1 (en) * 2021-05-20 2022-11-24 General Electric Company Active boundary layer control in diffuser

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS5932863U (ja) * 1982-08-27 1984-02-29 三菱重工業株式会社 ガスタ−ビン燃焼器
DE59504264D1 (de) * 1994-01-24 1998-12-24 Siemens Ag Brennkammer für eine gasturbine

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3403510A (en) * 1966-11-23 1968-10-01 United Aircraft Corp Removable and replaceable fuel nozzle holder assembly for an annular combustion burner
US3589127A (en) * 1969-02-04 1971-06-29 Gen Electric Combustion apparatus
US3631675A (en) * 1969-09-11 1972-01-04 Gen Electric Combustor primary air control

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB902511A (en) * 1960-04-19 1962-08-01 Bristol Siddeley Engines Ltd Improvements in or relating to combustion chambers

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3403510A (en) * 1966-11-23 1968-10-01 United Aircraft Corp Removable and replaceable fuel nozzle holder assembly for an annular combustion burner
US3589127A (en) * 1969-02-04 1971-06-29 Gen Electric Combustion apparatus
US3631675A (en) * 1969-09-11 1972-01-04 Gen Electric Combustor primary air control

Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4098074A (en) * 1976-06-01 1978-07-04 United Technologies Corporation Combustor diffuser for turbine type power plant and construction thereof
US4458479A (en) * 1981-10-13 1984-07-10 General Motors Corporation Diffuser for gas turbine engine
US4870826A (en) * 1987-06-18 1989-10-03 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma) Casing for a turbojet engine combustion chamber
US5134855A (en) * 1989-12-15 1992-08-04 Rolls-Royce Plc Air flow diffuser with path splitter to control fluid flow
US5077967A (en) * 1990-11-09 1992-01-07 General Electric Company Profile matched diffuser
US5339622A (en) * 1992-08-19 1994-08-23 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (S.N.E.C.M.A.) Gas turbine engine with improved water ingestion prevention
US5279126A (en) * 1992-12-18 1994-01-18 United Technologies Corporation Diffuser-combustor
US5737915A (en) * 1996-02-09 1998-04-14 General Electric Co. Tri-passage diffuser for a gas turbine
KR100476353B1 (ko) * 1996-02-09 2005-06-16 제너럴 일렉트릭 캄파니 압축기배출디퓨저및가스터빈
US6401447B1 (en) 2000-11-08 2002-06-11 Allison Advanced Development Company Combustor apparatus for a gas turbine engine
US20040011043A1 (en) * 2002-07-17 2004-01-22 Anthony Pidcock Diffuser for gas turbine engine
GB2390890A (en) * 2002-07-17 2004-01-21 Rolls Royce Plc Diffuser for gas turbine engine
GB2390890B (en) * 2002-07-17 2005-07-06 Rolls Royce Plc Diffuser for gas turbine engine
US7181914B2 (en) 2002-07-17 2007-02-27 Rolls-Royce Plc Diffuser for gas turbine engine
EP1426688A1 (en) * 2002-11-19 2004-06-09 General Electric Company Combustor inlet diffuser with boundary layer blowing
CN100416062C (zh) * 2002-11-19 2008-09-03 通用电气公司 具有附面层吹除的燃烧室进口扩压器
US20180266690A1 (en) * 2017-03-14 2018-09-20 Safran Aircraft Engines Combustion chamber of a turbine engine
US10684018B2 (en) * 2017-03-14 2020-06-16 Safran Aircraft Engines Combustion chamber of a turbine engine
US20220373181A1 (en) * 2021-05-20 2022-11-24 General Electric Company Active boundary layer control in diffuser
US11578869B2 (en) * 2021-05-20 2023-02-14 General Electric Company Active boundary layer control in diffuser

Also Published As

Publication number Publication date
DE2422362A1 (de) 1975-03-27
DE2422362B2 (de) 1979-08-23
JPS5045111A (ja) 1975-04-23
IT1011411B (it) 1977-01-20
FR2245855B1 (ja) 1978-08-04
FR2245855A1 (ja) 1975-04-25
DE2422362C3 (ja) 1980-05-08
GB1425438A (en) 1976-02-18
CA993668A (en) 1976-07-27
JPS5395004U (ja) 1978-08-02

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