US3737152A - Cooling of hot fluid ducts - Google Patents

Cooling of hot fluid ducts Download PDF

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Publication number
US3737152A
US3737152A US00218286A US3737152DA US3737152A US 3737152 A US3737152 A US 3737152A US 00218286 A US00218286 A US 00218286A US 3737152D A US3737152D A US 3737152DA US 3737152 A US3737152 A US 3737152A
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United States
Prior art keywords
holes
cooling
duct
ins
section
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Expired - Lifetime
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US00218286A
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English (en)
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J Wilson
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UK Secretary of State for Defence
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Secr Defence
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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
    • F23R3/06—Arrangement of apertures along the flame tube
    • F23R3/08—Arrangement of apertures along the flame tube between annular flame tube sections, e.g. flame tubes with telescopic sections
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present invention relates to ducts, one surface of which is exposed to a hot fluid, and relates particularly to means for providing film cooling for said surfaces.
  • the invention has application, for example, in combustion chambers of gas turbine engines.
  • the ducts In combustion chambers the ducts, known as flame tubes, are provided with machined rings which separate upstream and downstream wall sections of the duct, and which are provided with a row of drilled holes circumferentially spaced around the ring.
  • the solid rings are made from high temperature resistant materials and are difficult to machine and drill, which makes them expensive to manufacture.
  • the cooling air flow through the holes which are usually about /8th inch diameter, takes time to coalesce into a continuous film and hot combustion gases can be induced into the spaces between the cooling air holes and overheat the ring.
  • a duct comprises coaxial upstream and downstream wall sections which lie at different distances from the duct axis, and a cooling section through which a cooling fluid can be passed to form a film over one of the surfaces of the duct, the cooling section having end portions connected to said wall sections and an intermediate portion which is provided with at least three spaced rows of holes therethrough, each hole in any one row being staggered with respect to the holes in an adjacent row, the diameter of the holes lying in the range from 0.005 to 0.030 ins.
  • the duct may have a circular cross-section and the cooling section is then in the form of a ring connecting two wall sections of the duct having different diameters.
  • the intermediate portion of the cooling section is preferably conical and may be convergent or divergent in an upstream direction to produce a cooling film on the exterior surface of a downstream wall section of reduced diameter, or to produce a cooling film on the interior surface of a downstream wall section of increased diameter respectively.
  • the lower practical limit of diameter of the holes is likely to be between 0.005 and 0.010 in order to maintain an adequate porosity and to reduce the possibility of blockage of the holes by dirt carried in the fluid stream.
  • the upper limit of the range of diameters is set inter alia by the power of the machine on which the holes are produced. A further consideration is the heat released during drilling on an electron beam or laser pulse drilling machine.
  • a preferred range of diameters of the holes is between 0.0l ins and 0.020.ins.
  • a second advantage comes from an improved performance in cooling the duct walls, because the air flow through a large number of tiny holes coalesces into a continuous film more quickly than in the conventional cooling rings.
  • FIG. 1 is a diagrammatic view of a gas turbine engine, the combustion chamber of which includes a cooling ring of the present invention
  • FIG. 2 shows a cooling ring of FIG. 1 in greater detail
  • FIG. 3 is a view on the arrow A of FIG. 2.
  • FIG. 1 a gas turbine jet propulsion engine having a compressor l, combustion equipment 2, and a turbine 3 all in flow series, the exhaust from the turbine being passed to atmosphere through an exhaust nozzle 4.
  • the combustion equipment comprises outer and inner flame tubes 5 and 6, radially spaced respectively from each other and from outer and inner casings 7 and 8 to form an annular flame chamber surrounded by annular cooling air passages.
  • Each flame tube is constructed in a plurality of axially extending sections joined to each other by means of cooling rings 9.
  • Air from the compressor 1 flows into the passages between the casings 7 and 8 and the flame tubes 5 and 6, and some of said air passes into the flame tubes through the cooling rings 9, each of which is adapted to direct said air in a substantially continuous circumferential film over the internal surface of the flame tube wall section downstream thereof.
  • the cooling ring comprises axially extending end portions 10 and 11 which are dimensioned to be connected to the downstream end of an upstream wall section, and the upstream end of the adjacent downstream wall section respectively.
  • the directions of cooling air and hot gas flows are shown by arrow A.
  • the respective ends of the two adjacent wall sections are disposed at different radii from the axis of the engine so that an intermediate portion 12 of the cooling ring 9, which connects the end portions 10 and 11, lies at an angle to the engine axis.
  • the angle may be as large as but in the example shown it is 40.
  • the intermediate portion 12 is drilled, by a high powered electron beam drilling machine, with a very large number of tiny holes 13 which pass completely through the portion to allow a flow of cooling air therethrough. Each hole is drilled using a single pulsed beam of electrons which makes the process very rapid.
  • the example shown has seven rows of holes in which the holes are pitched 0.027 ins. apart and each hole in any one row is staggered with respect to the holes in an adjacent row, to cover the whole of the area of the intermediate portion 12.
  • the radial pitching and staggering of the holes is such that the centres of holes in any two rows lie on the apices of equilateral triangles.
  • the diameter of the holes is 0.016 ins. and the total area of the holes is approximately two fifths of the area of the surface of the intermediate portion.
  • Each of the holes was drilled at an angle of 30 to the I normal to the surface of the intermediate portion, so that the axis of each hole was-at 20to the axis of the engine.
  • the angle of the intermediate portion, and of the holes, to the axis of the engine is a compromise between the ideal condition of a radially extending intermediate portion (i.e., at 90 to the engine axis) with holes whose axes are parallel to the engine axis, and the need to have an adequate number of rows of holes without making the intermediate portion of the cooling ring too deep radially, or without the axes of the holes departing too far from being parallel to the axis of the engine.
  • the radial depth) between the radially inner surface of the cooling ring and the first cooling air hole must be kept to a minimum. This may be achieved by varying the angles of the axes of the holes so that the radially inner row of cooling holes lie at a steeper angle to the axis of the engine than the radially outer row.
  • the holes are preferably as small as they can practicably be made, i.e., down to 0.005 ins. in diameter. Very small holes, however, can get blocked by dirt and debris in the cooling air so that 0.010 ins. should be considered as a minimum diameter for the holes in dirty environments such as gas turbine engines. ln cleaner environments holes down to 0.005 ins. diameter may be used.
  • the upper limit of diameter is reached by considerations of the increasing cost of the technique of drilling, due to increased power requirements, and the fall-off in cooling performance as the hole diameter increases. With single pulse electron beam drilling, which is the recommended technique because of the high speed and low cost, the rate of increase of cost with diameter increases steeply for diameters of holes above 0.020 ins.
  • Cooling rings made as described above are of basically simple shape, and using a high power, a single pulse electron beam drilling process, can be manufactured at a greatly reduced cost compared with a conventional drilled cooling ring, particularly when made from high temperature resisting alloys such as PK.24.
  • the invention has a further advantage in that areas of the combustion chamber wall, which for a given engine design are known to be hotter than others, can be cooled by varying the numbers of holes in the hot regions. In an automatic machine for drilling the holes using an electron beam, such variations can easily be made by altering the instructions to the machine on the punched tape or other controlled programme. Thus the most efficient use of cooling air can be obtained through cooling rings of the present invention.
  • the invention has been described with reference to an annular combustion chamber flame tube of a gas turbine engine, but clearly is applicable to the walls of other hot fluid ducts, for example, liners for reheat pipes, or to the walls of the separate annular array of flame tubes or cans, which are sometimes used in combustion equipment in gas turbine engines.
  • a duct comprising coaxial upstream and downstream wall sections which lie at different distances from the longitudinal axis of the duct and a cooling section through which a cooling fluid can be passed to form a film over one of the surfaces of the duct, the cooling section having end portions connected to said wall sections and an intermediate portion which is pro vided with means defining at least three spaced rows of holes therethrough, each hole in any one row being staggered with respect to the holes in an adjacent row, the diameter of the holes lying in the range from 0.005 ins. to 0.030 ins.
  • a duct according to claim 1 having a circular cross section and wherein the cooling section is in the form of a ring connecting the upstream and downstream wall sections of the duct which have different diameters.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US00218286A 1971-01-25 1972-01-17 Cooling of hot fluid ducts Expired - Lifetime US3737152A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB305371 1971-01-25

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US3737152A true US3737152A (en) 1973-06-05

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US00218286A Expired - Lifetime US3737152A (en) 1971-01-25 1972-01-17 Cooling of hot fluid ducts

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US (1) US3737152A (enrdf_load_stackoverflow)
FR (1) FR2123386B1 (enrdf_load_stackoverflow)
GB (1) GB1320482A (enrdf_load_stackoverflow)
IT (1) IT949676B (enrdf_load_stackoverflow)

Cited By (23)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3952503A (en) * 1973-03-20 1976-04-27 Rolls-Royce (1971) Limited Gas turbine engine combustion equipment
US4232527A (en) * 1979-04-13 1980-11-11 General Motors Corporation Combustor liner joints
US4233123A (en) * 1978-12-18 1980-11-11 General Motors Corporation Method for making an air cooled combustor
US4242871A (en) * 1979-09-18 1981-01-06 United Technologies Corporation Louver burner liner
US4329848A (en) * 1979-03-01 1982-05-18 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Cooling of combustion chamber walls using a film of air
US4380905A (en) * 1979-03-22 1983-04-26 Rolls-Royce Limited Gas turbine engine combustion chambers
US4923371A (en) * 1988-04-01 1990-05-08 General Electric Company Wall having cooling passage
US5000005A (en) * 1988-08-17 1991-03-19 Rolls-Royce, Plc Combustion chamber for a gas turbine engine
US5181379A (en) * 1990-11-15 1993-01-26 General Electric Company Gas turbine engine multi-hole film cooled combustor liner and method of manufacture
US5261223A (en) * 1992-10-07 1993-11-16 General Electric Company Multi-hole film cooled combustor liner with rectangular film restarting holes
US5307637A (en) * 1992-07-09 1994-05-03 General Electric Company Angled multi-hole film cooled single wall combustor dome plate
US5329773A (en) * 1989-08-31 1994-07-19 Alliedsignal Inc. Turbine combustor cooling system
WO1995025932A1 (en) * 1989-08-31 1995-09-28 Alliedsignal Inc. Turbine combustor cooling system
US6021570A (en) * 1997-11-20 2000-02-08 Caterpillar Inc. Annular one piece combustor liner
US20100139324A1 (en) * 2007-04-12 2010-06-10 Saint- Gobain Isover Internal combustion burner
US9494081B2 (en) 2013-05-09 2016-11-15 Siemens Aktiengesellschaft Turbine engine shutdown temperature control system with an elongated ejector
US20220299206A1 (en) * 2021-03-19 2022-09-22 Raytheon Technologies Corporation Cmc stepped combustor liner
US11859823B2 (en) 2022-05-13 2024-01-02 General Electric Company Combustor chamber mesh structure
US11859824B2 (en) 2022-05-13 2024-01-02 General Electric Company Combustor with a dilution hole structure
US11867398B2 (en) 2022-05-13 2024-01-09 General Electric Company Hollow plank design and construction for combustor liner
US11994294B2 (en) 2022-05-13 2024-05-28 General Electric Company Combustor liner
US12066187B2 (en) 2022-05-13 2024-08-20 General Electric Company Plank hanger structure for durable combustor liner
US20250244014A1 (en) * 2024-01-30 2025-07-31 Honda Motor Co., Ltd. Combustor for gas turbine engine

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR2406075A1 (fr) * 1977-10-11 1979-05-11 Snecma Appareil de combustion et son procede de realisation

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2548485A (en) * 1946-01-09 1951-04-10 Shell Dev Combustion chamber lining
US3623711A (en) * 1970-07-13 1971-11-30 Avco Corp Combustor liner cooling arrangement

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2548485A (en) * 1946-01-09 1951-04-10 Shell Dev Combustion chamber lining
US3623711A (en) * 1970-07-13 1971-11-30 Avco Corp Combustor liner cooling arrangement

Cited By (27)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3952503A (en) * 1973-03-20 1976-04-27 Rolls-Royce (1971) Limited Gas turbine engine combustion equipment
US4233123A (en) * 1978-12-18 1980-11-11 General Motors Corporation Method for making an air cooled combustor
US4329848A (en) * 1979-03-01 1982-05-18 Societe Nationale D'etude Et De Construction De Moteurs D'aviation Cooling of combustion chamber walls using a film of air
US4380905A (en) * 1979-03-22 1983-04-26 Rolls-Royce Limited Gas turbine engine combustion chambers
US4232527A (en) * 1979-04-13 1980-11-11 General Motors Corporation Combustor liner joints
US4242871A (en) * 1979-09-18 1981-01-06 United Technologies Corporation Louver burner liner
US4923371A (en) * 1988-04-01 1990-05-08 General Electric Company Wall having cooling passage
US5000005A (en) * 1988-08-17 1991-03-19 Rolls-Royce, Plc Combustion chamber for a gas turbine engine
US5329773A (en) * 1989-08-31 1994-07-19 Alliedsignal Inc. Turbine combustor cooling system
WO1995025932A1 (en) * 1989-08-31 1995-09-28 Alliedsignal Inc. Turbine combustor cooling system
AU638331B2 (en) * 1990-11-15 1993-06-24 General Electric Company Gas turbine engine multi-hole film cooled combustor liner and method of manufacture
US5181379A (en) * 1990-11-15 1993-01-26 General Electric Company Gas turbine engine multi-hole film cooled combustor liner and method of manufacture
US5307637A (en) * 1992-07-09 1994-05-03 General Electric Company Angled multi-hole film cooled single wall combustor dome plate
US5261223A (en) * 1992-10-07 1993-11-16 General Electric Company Multi-hole film cooled combustor liner with rectangular film restarting holes
US6021570A (en) * 1997-11-20 2000-02-08 Caterpillar Inc. Annular one piece combustor liner
US9587822B2 (en) 2007-04-12 2017-03-07 Saint-Gobain Isover Internal combustion burner
US20100139324A1 (en) * 2007-04-12 2010-06-10 Saint- Gobain Isover Internal combustion burner
US9494081B2 (en) 2013-05-09 2016-11-15 Siemens Aktiengesellschaft Turbine engine shutdown temperature control system with an elongated ejector
US20220299206A1 (en) * 2021-03-19 2022-09-22 Raytheon Technologies Corporation Cmc stepped combustor liner
US11867402B2 (en) * 2021-03-19 2024-01-09 Rtx Corporation CMC stepped combustor liner
US12359816B2 (en) 2021-03-19 2025-07-15 Rtx Corporation CMC stepped combustor liner
US11859823B2 (en) 2022-05-13 2024-01-02 General Electric Company Combustor chamber mesh structure
US11859824B2 (en) 2022-05-13 2024-01-02 General Electric Company Combustor with a dilution hole structure
US11867398B2 (en) 2022-05-13 2024-01-09 General Electric Company Hollow plank design and construction for combustor liner
US11994294B2 (en) 2022-05-13 2024-05-28 General Electric Company Combustor liner
US12066187B2 (en) 2022-05-13 2024-08-20 General Electric Company Plank hanger structure for durable combustor liner
US20250244014A1 (en) * 2024-01-30 2025-07-31 Honda Motor Co., Ltd. Combustor for gas turbine engine

Also Published As

Publication number Publication date
FR2123386B1 (enrdf_load_stackoverflow) 1976-07-23
IT949676B (it) 1973-06-11
FR2123386A1 (enrdf_load_stackoverflow) 1972-09-08
DE2202356B2 (de) 1976-02-19
DE2202356A1 (de) 1972-08-10
GB1320482A (en) 1973-06-13

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