US3232054A - Liquid fuel combustion apparatus, for jet-propulsion engines, gas turbines, or other prime movers - Google Patents

Liquid fuel combustion apparatus, for jet-propulsion engines, gas turbines, or other prime movers Download PDF

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Publication number
US3232054A
US3232054A US346278A US34627864A US3232054A US 3232054 A US3232054 A US 3232054A US 346278 A US346278 A US 346278A US 34627864 A US34627864 A US 34627864A US 3232054 A US3232054 A US 3232054A
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United States
Prior art keywords
liquid fuel
jet
nose piece
gas turbines
prime movers
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Expired - Lifetime
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US346278A
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Squire R Jackson
George S Cooper
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ZF International UK Ltd
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Lucas Industries Ltd
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Publication date
Priority to FR962434A priority Critical patent/FR1381149A/en
Priority to DE1964L0046955 priority patent/DE1235671B/en
Application filed by Lucas Industries Ltd filed Critical Lucas Industries Ltd
Priority to US346278A priority patent/US3232054A/en
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Publication of US3232054A publication Critical patent/US3232054A/en
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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
    • F23R3/10—Air inlet arrangements for primary air

Definitions

  • LIQUID FUEL COMBUSTION APPARATUS FOR JET-PROPULSION ENGINES, GAS TURBINES, OR OTHER PRIME MOVERS Filed Feb. 20, 1964 5 Sheets-Sheet 1 Feb. 1, 1966 s.
  • R. JACKSON ETAL 3,232,054 STION APPARATUS FOR JET-PROPULS GAS TURBINES, OR OTHER PRIME MOVERS ION 3 Sheets-Sheet 2 LIQUID FUEL COMBU ENGINES Filed Feb. 20, 1964 1966 s.
  • JACKSON ETAL 3 LIQUID FUEL COMBUSTION APPARATUS, FOR JET-PROPULSION ENGINES, GAS TURBINES, OR OTHER PRIME MOVERS Filed Feb. 20, 1964 5 Sheets-Sheet 1 Feb. 1, 1966 s.
  • R. JACKSON ETAL 3,232,054 STION APPARATUS FOR JET-PROPULS GAS TURBINES, OR OTHER PRIME MOVERS ION
  • LIQUID FUEL COMBUSTION APPARATUS FOR JET-PROPULSION ENGINES, GAS TURBINES, OR OTHER PRIME MOVERS Filed Feb. 20, 1964 3 Sheets-Sheet 5 United States Patent M 3,232,054 LIQUID FUEL COMBUSTION APPARATUS, FOR JET-PROPULSION ENGINES, GAS TURBINES, OR OTHER PRIME MOVERS Squire R. Jackson, Burnley, and George S. Cooper, Rawtenstall, England, assignors to Joseph Lucas (Industries) Limited, Birmingham, England Filed Feb. 20, 1964, Ser. No. 346,278 1 Claim. (Cl. Gil-39.74)
  • This invention relates to an improved liquid fuel combustion apparatus, for jet-propulsion engines, gas turbines, or other prime movers and of the kind comprising a combustion chamber of substantially circular cross-section, a liquid fuel burner at the inlet end of the combustion chamber, a swirler surrounding the burner, an air jacket surrounding the combustion chamber and having an air inlet at one end, and a nose piece extending from the combustion chamber towards said inlet.
  • the nose piece is connected to the combustion chamber through the intermediary of a hollow tapering part of generally circular cross-section the nose piece being shaped to define between it and said part a pair of substantially segmental openings through which air can flow to the swirler.
  • FIGURE 1 is a sectional side view of a part of .an apparatus in accordance with one embodiment of the invention
  • FIGURE 2 is a part sectional end view of a pair of adjacent combustion chambers of the apparatus shown in FIGURE 1, and
  • FIGURE 3 is a cross-sectional view on the line 3-3 in FIGURE 1.
  • combustion chambers 3 of circular cross-section which are located within an annular air jacket 4.
  • the combustion chambers are interconnected by pressure balancing pipes 5 and are spaced from the walls of the air jacket 4 which converge at one end to define an annular air inlet 6.
  • each combustion chamber Centrally disposed within the inlet end of each combustion chamber is a liquid fuel burner 7 which is surrounded by a swirler 8 located in a substantially conical end plate 9 of the combustion chamber. Moreover, secured to the inlet end of the combustion chamber is a hollow tapering part 10 of generally circular cross-section throughout its length, whilst to the narrower end of said part 10 is connected a nose piece 11.
  • the combined part and nose piece is of a symmetrical form with its surface presented to the outer annular wall of the .air jacket 4, curving progressively inwards (in an upstream direction) towards the centre of the annular row of combustion chambers, more sharply than the surface, presented to the inner annular wall of the air jacket, which curves progressively outwards.
  • FIGURES 2 and 3 it will be seen that the combined nose piece 11 and part 10 is symmetrical in a plane at right angles to the plane of the longitudinal cross-section of FIGURE 1.
  • the maximum width of the nose piece 11 in this plane is less than the narrower end of the part 10 so as to define, at opposite sides of the nose piece 11 a pair of substantially segmental openings 12 through which air can enter the part 10 and flow to the swirler.
  • the sides of the nose piece adjacent the part 10 are parallel and only its outer end tapers.
  • Ribs 18 span the openings 12 between the side Wall of the nose piece 11 and the edge of the part 10. Inside the nose piece are a pair of brackets 19, which are joined together, as shown in FIGURE 3, in the plane of the section of FIGURE 1. The edges of the nose piece 11 have stiffening strips 21.
  • an aperture .13 for accommodating a fuel supply pipe 14 for the burner 7, this pipe also extending through the outer Wall of the air jacket.
  • the width of the nose piece 11 as viewed in FIGURE 3 is approximately equal to the diameter of the swirler 8 so as to protect the latter from the direct impingement of air entering the part 10.
  • a proportion of the air entering the part 10 flows through the swirler 8 to the combustion chamber. Another part of this air flows over the conical end plate 9 of the combustion chamber and enters the combustion chamber through flared openings therein, .and around the outer edge thereof.
  • the combined combustion chamber 3, part 10 and nose piece 11 is supported in spaced relationship within the air jacket 4 by means of a forked bracket 15 secured to the jacket and embracing a reduced portion 10 of the part 10 adjacent the nose piece 11.
  • the reduced portion of the nose piece is for-med as a result of the formation of a pair of recesses 16 in the outer surface of the part 10, and a pin 17 passes transversely through the portion 10 .and the two limbs of the bracket 15.
  • Liquid fuel combustion apparatus for gas turbine engines comprising a combustion chamber of substantially circular cross-section having an inlet at an upstream end thereof and an outlet at a downstream end thereof, a liquid fuel burner at the upstream end of the combustion chamber, a swirler surrounding the burner, an air jacket surrounding the combustion chamber having an air inlet at an upstream end of the air jacket adjacent to the upstream end of the combustion chamber, .

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Spray-Type Burners (AREA)

Description

1966 s. R. JACKSON ETAL 3,232,054
LIQUID FUEL COMBUSTION APPARATUS, FOR JET-PROPULSION ENGINES, GAS TURBINES, OR OTHER PRIME MOVERS Filed Feb. 20, 1964 5 Sheets-Sheet 1 Feb. 1, 1966 s. R. JACKSON ETAL 3,232,054 STION APPARATUS, FOR JET-PROPULS GAS TURBINES, OR OTHER PRIME MOVERS ION 3 Sheets-Sheet 2 LIQUID FUEL COMBU ENGINES Filed Feb. 20, 1964 1966 s. R. JACKSON ETAL 3,
LIQUID FUEL COMBUSTION APPARATUS, FOR JET-PROPULSION ENGINES, GAS TURBINES, OR OTHER PRIME MOVERS Filed Feb. 20, 1964 3 Sheets-Sheet 5 United States Patent M 3,232,054 LIQUID FUEL COMBUSTION APPARATUS, FOR JET-PROPULSION ENGINES, GAS TURBINES, OR OTHER PRIME MOVERS Squire R. Jackson, Burnley, and George S. Cooper, Rawtenstall, England, assignors to Joseph Lucas (Industries) Limited, Birmingham, England Filed Feb. 20, 1964, Ser. No. 346,278 1 Claim. (Cl. Gil-39.74)
This invention relates to an improved liquid fuel combustion apparatus, for jet-propulsion engines, gas turbines, or other prime movers and of the kind comprising a combustion chamber of substantially circular cross-section, a liquid fuel burner at the inlet end of the combustion chamber, a swirler surrounding the burner, an air jacket surrounding the combustion chamber and having an air inlet at one end, and a nose piece extending from the combustion chamber towards said inlet.
According to the invention in a liquid fuel combustion apparatus of the kind specified the nose piece is connected to the combustion chamber through the intermediary of a hollow tapering part of generally circular cross-section the nose piece being shaped to define between it and said part a pair of substantially segmental openings through which air can flow to the swirler.
An example of the invention will now be described with particular reference to the accompanying drawings in which:
FIGURE 1 is a sectional side view of a part of .an apparatus in accordance with one embodiment of the invention,
FIGURE 2 is a part sectional end view of a pair of adjacent combustion chambers of the apparatus shown in FIGURE 1, and
FIGURE 3 is a cross-sectional view on the line 3-3 in FIGURE 1.
Referring to the drawings there is provided a plurality of annularly arranged combustion chambers 3 of circular cross-section which are located within an annular air jacket 4. The combustion chambers are interconnected by pressure balancing pipes 5 and are spaced from the walls of the air jacket 4 which converge at one end to define an annular air inlet 6.
Centrally disposed within the inlet end of each combustion chamber is a liquid fuel burner 7 which is surrounded by a swirler 8 located in a substantially conical end plate 9 of the combustion chamber. Moreover, secured to the inlet end of the combustion chamber is a hollow tapering part 10 of generally circular cross-section throughout its length, whilst to the narrower end of said part 10 is connected a nose piece 11. In longitudinal sectional side elevation (as shown in FIGURE 1) the combined part and nose piece is of a symmetrical form with its surface presented to the outer annular wall of the .air jacket 4, curving progressively inwards (in an upstream direction) towards the centre of the annular row of combustion chambers, more sharply than the surface, presented to the inner annular wall of the air jacket, which curves progressively outwards. From FIGURES 2 and 3, it will be seen that the combined nose piece 11 and part 10 is symmetrical in a plane at right angles to the plane of the longitudinal cross-section of FIGURE 1. However, the maximum width of the nose piece 11 in this plane is less than the narrower end of the part 10 so as to define, at opposite sides of the nose piece 11 a pair of substantially segmental openings 12 through which air can enter the part 10 and flow to the swirler. Moreover Patented Feb. 1, 1966 as viewed in this direction the sides of the nose piece adjacent the part 10 are parallel and only its outer end tapers.
Ribs 18 span the openings 12 between the side Wall of the nose piece 11 and the edge of the part 10. Inside the nose piece are a pair of brackets 19, which are joined together, as shown in FIGURE 3, in the plane of the section of FIGURE 1. The edges of the nose piece 11 have stiffening strips 21.
In the outer surface of the nose piece 11, presented to the outer annular wall of the air jacket is formed an aperture .13 for accommodating a fuel supply pipe 14 for the burner 7, this pipe also extending through the outer Wall of the air jacket. Moreover, the width of the nose piece 11 as viewed in FIGURE 3 is approximately equal to the diameter of the swirler 8 so as to protect the latter from the direct impingement of air entering the part 10.
A proportion of the air entering the part 10 flows through the swirler 8 to the combustion chamber. Another part of this air flows over the conical end plate 9 of the combustion chamber and enters the combustion chamber through flared openings therein, .and around the outer edge thereof.
The combined combustion chamber 3, part 10 and nose piece 11 is supported in spaced relationship within the air jacket 4 by means of a forked bracket 15 secured to the jacket and embracing a reduced portion 10 of the part 10 adjacent the nose piece 11. The reduced portion of the nose piece is for-med as a result of the formation of a pair of recesses 16 in the outer surface of the part 10, and a pin 17 passes transversely through the portion 10 .and the two limbs of the bracket 15.
Having thus described our invention what we claim as new and desire to secure by Letters Patent is:
Liquid fuel combustion apparatus for gas turbine engines comprising a combustion chamber of substantially circular cross-section having an inlet at an upstream end thereof and an outlet at a downstream end thereof, a liquid fuel burner at the upstream end of the combustion chamber, a swirler surrounding the burner, an air jacket surrounding the combustion chamber having an air inlet at an upstream end of the air jacket adjacent to the upstream end of the combustion chamber, .a hollow tapering part of generally circular cross-section connected to the combustion chamber near its upstream end, a nose piece connected to the hollow tapering part, the nose piece being of non-circular cross-section to define between it and the hollow tapering part, a pair of substantially segmental openings, the position of the nose piece being arranged to prevent direct impingement of air entering the openings upon the swirler, said openings being disposed on opposite sides of a plane passing through a longitudinal axis of the combustion chamber and swirler, and the edges of the openings nearest to said plane being spaced apart by a distance at least equal to the width of the swirler.
References Cited by the Examiner UNITED STATES PATENTS 2,872,971 2/1959 Clarke 60-39.65 X 3,119,234 l/1964 Murray 60-39.37
FOREIGN PATENTS 216,743 8/ 1958 Australia.
MARK NEWMAN, Primary Examiner.
RALPH D. BLAKESLEE, Examiner.
US346278A 1964-02-03 1964-02-20 Liquid fuel combustion apparatus, for jet-propulsion engines, gas turbines, or other prime movers Expired - Lifetime US3232054A (en)

Priority Applications (3)

Application Number Priority Date Filing Date Title
FR962434A FR1381149A (en) 1964-02-03 1964-02-03 Liquid fuel combustion apparatus for jet propulsion engine, gas turbine or other prime mover
DE1964L0046955 DE1235671B (en) 1964-02-04 1964-02-04 Combustion chamber for liquid fuels for gas turbine and jet engines
US346278A US3232054A (en) 1964-02-20 1964-02-20 Liquid fuel combustion apparatus, for jet-propulsion engines, gas turbines, or other prime movers

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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3927835A (en) * 1973-11-05 1975-12-23 Lucas Aerospace Ltd Liquid atomising devices
US4214925A (en) * 1977-10-25 1980-07-29 Kobe Steel, Limited Method for fabricating brazed aluminum fin heat exchangers
FR2500905A1 (en) * 1981-03-02 1982-09-03 Gen Electric REAR MOUNTING SYSTEM FOR TRANSITION CHANNEL ELEMENTS

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2872971A (en) * 1959-02-10 Combustion chambers for jet propulsion
US3119234A (en) * 1960-09-13 1964-01-28 Rolls Royce Combustion chamber for a gas turbine engine

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2872971A (en) * 1959-02-10 Combustion chambers for jet propulsion
US3119234A (en) * 1960-09-13 1964-01-28 Rolls Royce Combustion chamber for a gas turbine engine

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3927835A (en) * 1973-11-05 1975-12-23 Lucas Aerospace Ltd Liquid atomising devices
US4214925A (en) * 1977-10-25 1980-07-29 Kobe Steel, Limited Method for fabricating brazed aluminum fin heat exchangers
FR2500905A1 (en) * 1981-03-02 1982-09-03 Gen Electric REAR MOUNTING SYSTEM FOR TRANSITION CHANNEL ELEMENTS
US4422288A (en) * 1981-03-02 1983-12-27 General Electric Company Aft mounting system for combustion transition duct members

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