US3864907A - Step cylinder combustor design - Google Patents

Step cylinder combustor design Download PDF

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US3864907A
US3864907A US41285973A US3864907A US 3864907 A US3864907 A US 3864907A US 41285973 A US41285973 A US 41285973A US 3864907 A US3864907 A US 3864907A
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combustor
means
combustor chamber
disposed
heat
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Edward T Curran
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K7/00Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof
    • F02K7/10Plants in which the working fluid is used in a jet only, i.e. the plants not having a turbine or other engine driving a compressor or a ducted fan; Control thereof characterised by having ram-action compression, i.e. aero-thermo-dynamic-ducts or ram-jet engines

Abstract

A supersonic combustion ramjet having a combustor including at least two constant area combustor chamber-sections or stages disposed in open communication between the ramjet inlet and expansion nozzle, and further each incorporating a combustion stabilizing step configuration and a separate fuel injector for respectively and selectively adding heat to the pair of constant area-chambers to control the heat addition to the subsonic or supersonic mode.

Description

e i I United States Patent 1 3,864,907

Curran Feb. 11, 1975 [54] STEP CYLINDER COMBUSTOR DESIGN 3,102,392 9/1963 Bauger et al (M/39.72

[75] lnventor: Edward T. Curran, Dayton, OhlO l 6 I 97 my 6 l 0 R X [73] .Assignee: The United States of America as Primary Examiner-Trygve M. Blix represented by the Secretary of the Assistant Examiner-Paul E. Sauberer United States Air Force, Attorney, Agent, or FirmHarry A. Herbert, .lr.; Washington, DC. Arthur R. Parker 22 F'l d: N 5, 1973 l 1 57 ABSTRACT [2]] App! 412359 A supersonic combustion ramjet having a combustor including at least two constant area combustor cham- [52] U.S. Cl. 60/261, 60/39.72 R, 60/270 R her-sections or stages disposed in open communica- [51] Int. Cl. F02k 7/08 tion between the ramjet inlet and expansion nozzle,

[58] Field of Search..... 60/270 R, 39.72 R, 39.72 P, and further each incorporating a combustion stabiliz- 60/241, 26l; 244/53 R, 53 B ing step configuration and a separate fuel injector for respectively and selectively adding heat to the pair of [56] References Cited constant area-chambers to control the heat addition to UNITED STATES PATENTS the subsonic OI supersonic mode.

2,867,979 l/l959 Mullen 60/39.72 R X 4 Claims, 3 Drawing Figures STEP CYLINDER COMBUSTOR DESIGN- RIGHTS OF THE GOVERNMENT BACKGROUND OF THE INVENTION This invention relates generally and primarily to supersonic combustion ramjet engines and, in particular, to an improved stepcylinder combustor chamber therefor.

In the past, supersonic combustion engines have generally employed conical or, in other words, divergent combustors whose area naturally increases with length. Such divergent combustors are required to prevent the phenomenon known as thermal choking, which limits theamount of heat that may be added to the single constant area-combustor chamber or duct. In this connection, the conventional engine of this type has been previously designed with a step-cylinder-cone geometry, in which a single step is utilized in the combustor chamber immediately downstream from the engine-air inlet. This step configuration acts to stabilize the combustion and is followed both by a single cylindrical or constant area-chamber section or stage in which a limited amount of heat is added, and, finally, by a divergent or conical section or stage in which is added still further heat. However, although this prior art arrangement affords a somewhat increased combustion stability, the aforementioned thermal choking" may still occur in the first, constant area stage if too much heat is attempted to be added and, therefore, this places a definite limit to the actual heat which may be so added therein. Moreover, the further heat that is to be added in the second, conical or divergent stage or chamber section suffers from the inherent disadvantages of the expanding flow in such conical stages that has been previously found to adversely affect the chemical reaction and mixing therein. On the other hand, the new and improved combustor design of the present invention offers considerable improvement over the said step-cylinder-co|1e configuration by a simplified and yet unique means to be set forth hereinafter in the following summary and detailed description.

SUMMARY OF THE INVENTION The present invention consists briefly in improving the conventional single step-cylinder-cone geometry of current combustor chamber configurations, particularly as used in supersonic combustion-ramjet engines, by utilizing a combustor chamber that has at least two stabilizing steps or stages, each being followed by a constant area-cylindrical or quasi-cylindrical duct member or combustor chamber-section or stage in which heat may be selectively or collectively added, as desired. More specifically, the inventive combustor chamber would consist of a first step adjacent to, or immediately downstream of the engine-air inlet, which step would be immediately followed by a first, cylindrical or, in other words, constant area chamber section or stage, and at least one other step forming a second, cylindrical/constant area chamber section or stage, to be followed by the expansion nozzle of the engine.

In its preferred form, the present invention would be applied to, a missile with, of course, the main fuel tank being centrally disposed along; the longitudinal axis thereof and with the various constant area stages being disposed as annular chamber sections around the fuel tank. Actually, it is preferable that a three-step combustor be used in the missile engine in which a total of three constant area chambers of progressively increasing diameter would be used, each being preceeded by a stabilizing step. In this manner, by utilizing additional heat in each of the three constant area stages, the previously-described adverse effects of the divergent flow of more conventional ramjet engines would be avoided, while, at the same time, greater amounts of heat and thus energy could be added in a quite simplified system.

Other objects and advantages of the invention will readily appear hereinafter from the following disclosure, taken in connection with the accompanying draw ings, in which:

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a longitudinally-disposed diagrammatic view of a conventional supersonic combustion ramjet wherein a single step-cylindrical-cone configuration is used for the combustor chamber thereof.

FIG. 2 is a second longitudinally-disposed, diagrammatic view of an improved and basic two step and constant area combustor chamber used to improve the efficiency of the conventional ramjet engine of FIG. 1; and

FIG. 3 is a longitudinal, and partly sectional and schematic view, illustrating details of the further improved ramjet-three step combustor chamber comprising the preferred form of the present invention.

DESCRIPTION OF THE PREFERRED EMBODIMENT Referring initially to FIG. 1 of the drawings, a prior art type of ramjet engine is shown indicated generally at 10 as comprising a main casing or duct member =11, a spike inlet diffuser 12, a main, air inlet 13, and first and second, cylindrical/constant area, and divergent combustor chamber-sections of annular or substantially annular disposition and indicated respectively and generally at 14 and 15. A single combustion stability-step, indicated at 16 as being disposed just downstream of the air inlet 13, leads into, and assists in forming the first, cylindrical/constant areacombustor chamber-section 14. An expansion nozzle 17 completes the principal portions of the engine. As noted hereinbefore, the single step-cylindrical-cone configuration of the combustor chamber of the prior art or conventional ramjet of FIG. 1 affords some improvement over the formerly-used and simple diverging combustor chamber of other engines for the reason that they are operative to prevent or at least drastically alleviate the phenomenon called thermal choking" of the engine that limited the amount of heat which could be added in the constant area-type of duct. However, again, as stated before, although the single step configuration of the aforementioned FIG. I additionally provides a certain amount of improved combustion stability in that limited amounts of heat may be respectively added in both of its constant-area (cylindrical) and divergentcombustor chamber-sections 14 and 15 by means of fuel injection stages, indicated respectively and schematically at 14a and 15a; nevertheless, the divergent configuration of the second combustor chamber-stage 15 makes for a relatively low combustion and therefore engine efficiency because of the previously-mentioned adverse effect, inherent in the expanding flow, on the chemical reaction and mixing in the said divergent section 15.

To specifically provide for a further improved combustion stability and therefore an increased efficiency of operation of the ramjet engine, in accordance with the unique teachings of the present invention, a combustor chamber configuration having two or more stability step-cylinder elements as respectively depicted in FIGS. 2 and 3 may be utilized. In the more basic form of FIG. 2, a supersonic combustion ramjet engine is indicated at 18 as again comprising the main ramjet engine-duct member at 19, a main air inlet indicated generally at 20 and out of which projects the spike inlet diffuser 21, and an expansion nozzle 22.

Disposed between the expansion nozzle 22 and the air inlet 20 is the novel and basic two step-cylinder forming the present invention. This consists of a first step element 23 and a first, cylindrical or constant area combustor chamber-section or stage 24, and a second step element 25 and second, relatively increased diameter and constant area combustor chamber-section or stage 26, each respectively having heat added thereinto by means of the first and second fuel injector stages, indicated schematically and generally at 27 and 28. Thus, by adding heat in at least two constant area stages, increased energy may be produced while, at the same time, the adverse effects of divergent flow may be avoided.

In FIG. 3, a more elaborate form of the invention and one which actually constitutes the preferred embodiment thereof is indicated generally at 29 as comprising the main engine structure or duct member at 30 that houses and incorporates the main air intake 31, a spike inlet diffuser 32 that may also carry the missile payload and which is supported in position by the struts at 33, an expansion nozzle structure 34 supported by struts 35, and intermediately disposed between said spike inlet diffuser 32 and expansion nozzle structure 34 the new and improved multiple step-cylinder-combustor chamber of the preferred form of the invention. The latter comprises a first step 36 and first, cylindrical/- constant area-chamber section or stage 37, a second step 38 and second, cylindrical/constant area-chamber section or stage 39, and a third step 40 and third, cylindrical/constant area-chamber section or stage 41. For each of the said three sections or stages or, in other words, constant area-combustion stages 37, 39 and 41, there is heat added respectively and separately by means of the fuel injectors, indicated schematically at 42, 43 and 44, which fuel injectors are suitably incorporated in a housing structure, indicated generally at 45 as surrounding the outer periphery of the missile fuel tank 46. These fuel injectors are not illustrated in detail since they may be of any well-known conventional design. It is by appropriately varying the outside diameter of the corresponding sections of the said fuel injector-housing structure 45 that the first, second and third, constant area-combustor chamber-sections or stages 37, 39 and 41 may be respectively and progressively increased in diameter or otherwise sized as desired, as is clearly illustrated in FIG. 3.

Thus, a new and improved combustor chamber configuration has been developed primarily for, but not necessarily limited to, supersonic combustion ramjet engines wherein heat may be added to two or more constant area-chamber stages to improve the combustion efficiency thereof and which enables, by preselecting the ratios of each step area, the control of the Mach Number at which combustion occurs. Moreover, the use of discrete steps, as at 36, 38 and 40 (FIG. 3), in the combustor allows closer spatial control of the rate of heat addition to be achieved, by providing predetermined and various stabilizing zones or stages coupled with each fuel injection stage. By well-known fuel control means that may be located generally in the area indicated at 47, the separate combustion stages formed by the present invention may be either placed in operation or kept out of operation as desired, or they may be modulated in such a way as to control the addition of heat either to the supersonic or subsonic mode, and thereby promote and/or facilitate more efficient engine operation under various flight conditions for both supersonic and subsonic combustion ramjets.

I claim:

1. In a ramjet engine having a main duct member, an air inlet and an expansion nozzle-forming structure; combustor chamber means incorporated in said duct member and comprising; a plurality of substantially cylindrically-shaped, combustor chamber-sections disposed in direct and continuous alignment with each other and in open communication between said air inlet and expansion nozzle-forming structure, each combustor chamber-sections being formed into a constant area configuration and respectively increasing in size in the downstream direction to thereby provide for the periodic, non-divergent and relatively even expansion of flow in separate stages and in all directions inherently ensuring better fuel-air mixing and thus the more efficient combustion thereof; separate combustion stability means inherently incorporated with and disposed at or near the upstream end of each of said combustor chamber-sections; and individually actuatable, heat-adding means positioned immediately upstream of each said combustor stability means, and selectively and individually operable to add varying amounts of heat to said combustor chamber means to thereby provide a positive, built-in control and regulation of the speed being inherently produced during the operation of said ramjet engine.

2. In a ramjet engine as in claim 1, wherein said separate combustion stability means each comprise a step formed at the upstream end of each of said combustorchamber sections.

3. In a ramjet engine as in claim 2; and relatively elongated, center body means centrally disposed in said combustor chamber means along the longitudinal axis of said main duct member and incorporating a spike diffuser element at the front end thereof disposed at and through the air inlet, and a center body member extending between said spike diffuser element and the expansion nozzle-forming structure; said center body member comprising a plurality of substantially cylindrical, center body elements respectively decreasing in diameter in the downstream direction to thereby incorporate and form each of said combustion stability steps at the junctures between adjacent center body elethe aforementioned 4. In a ramjet engine as in claim 1, wherein said main duct member comprises a missile having a main fuel tank and combined payload and spike inlet diffuser member, and said combustor chamber means com prises a series of three combustion stability steps and combustor chamber-sections annularily disposed between a supporting structure surrounding said fuel tank and the outside wall of said duct member.

Claims (4)

1. In a ramjet engine having a main duct member, an air inlet and an expansion nozzle-forming structure; combustor chamber means incorporated in said duct member and comprising; a plurality of substantially cylindrically-shaped, combustor chamber-sections disposed in direct and continuous alignment with each other and in open communication between said air inlet and expansion nozzle-forming structure, each combustor chambersections being formed into a constant area configuration and respectively increasing in size in the downstream direction to thereby provide for the periodic, non-divergent and relatively even expansion of flow in separate stages and in all directions inherently ensuring better fuel-air mixing and thus the more efficient combustion thereof; separate combustion stability means inherently incorporated with and disposed at or near the upstream end of each of said combustor chamber-sections; and individually actuatable, heat-adding means positioned immediately upstream of each said combustor stability means, and selectively and individually operable to add varying amounts of heat to said combustor chamber means to thereby provide a positive, built-in control and regulation of the speed being inherently produced during the operation of said ramjet engine.
2. In a ramjet engine as in claim 1, wherein said separate combustion stability means each comprise a step formed at the upstream end of each of said combustor-chamber sections.
3. In a ramjet engine as in claim 2; and relatively elongated, center body means centrally disposed in said combustor chamber means along the longitudinal axis of said main duct member and incorporating a spike diffuser element at the front end thereof disposed at and through the air inlet, and a center body member extending between said spike diffuser element and the expansion nozzle-forming structure; said center body member comprising a plurality of substantially cylindrical, center body elements respectively decreasing in diameter in the downstream direction to thereby incorporate and form each of said combustion stability steps at the junctures between adjacent center body elements, and further respectively define each of the said combustor chambers into a constant area configuration disposed in circumferential relation around the corresponding center body element and progressively increasing in size in the downstream direction to thereby ensure both non-divergent flow coupled with the addition of heat by said individually actuatable, heat-adding means in one or more selected stages to positively promote more efficient combustion and the mode of control to either the supersonic or subsonic range.
4. In a ramjet engine as in claim 1, wherein said main duct member comprises a missile having a main fuel tank and combined payload and spike inlet diffuser member, and said combustor chamber means comprises a series of three combustion stability steps and combustor chamber-sections annularily disposed between a supporting structure surrounding said fuel tank and the outside wall of said duct member.
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Cited By (26)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4091731A (en) * 1976-07-06 1978-05-30 The United States Of America As Represented By The Secretary Of The Navy Fuel injection with flameholding
US4802639A (en) * 1984-09-28 1989-02-07 The Boeing Company Horizontal-takeoff transatmospheric launch system
FR2634005A1 (en) * 1987-05-05 1990-01-12 United Technologies Corp Pilotage fuel injector assembly for a supersonic statoreactor
EP0401107A1 (en) * 1989-05-29 1990-12-05 Societe Europeenne De Propulsion (S.E.P.) S.A. Combustion chamber for ram jet
GB2243409A (en) * 1990-02-28 1991-10-30 Gen Electric Scramjet
US5072581A (en) * 1989-03-23 1991-12-17 General Electric Company Scramjet combustor
US5072582A (en) * 1989-03-23 1991-12-17 General Electric Company Scramjet combustor
US5081831A (en) * 1989-03-23 1992-01-21 General Electric Company Scramjet combustor
US5097663A (en) * 1989-03-23 1992-03-24 General Electric Company Scramjet combustor
US5103633A (en) * 1989-03-23 1992-04-14 General Electric Company Scramjet combustor
US5109670A (en) * 1989-03-23 1992-05-05 General Electric Company Scramjet combustor
US5253474A (en) * 1991-08-30 1993-10-19 General Electric Company Apparatus for supersonic combustion in a restricted length
US5255513A (en) * 1990-02-28 1993-10-26 General Electric Company Method of operating a scramjet including integrated inlet and combustor
US5280705A (en) * 1992-06-22 1994-01-25 General Electric Company Fuel injection system for scramjet engines
US5333445A (en) * 1990-01-29 1994-08-02 General Electric Company Scramjet engine having improved fuel/air mixing
US5349815A (en) * 1991-08-27 1994-09-27 General Electric Company Scramjet combustor having a two-part, aft-facing step
EP0693668A2 (en) * 1994-06-21 1996-01-24 Rockwell International Corporation Gas gun launched scramjet test projectile
US5544586A (en) * 1994-08-30 1996-08-13 The United States Of America As Represented By The Secretary Of The Army Solid fuel ramjet tubular projectile
WO2003010432A1 (en) * 2001-07-23 2003-02-06 Ramgen Power Systems, Inc. Rotary ramjet engine with flameholder extending to running clearance at engine casing interior wall
US20040020211A1 (en) * 2001-07-23 2004-02-05 Ramgen Power Systems, Inc. Trapped vortex combustor
US6694743B2 (en) 2001-07-23 2004-02-24 Ramgen Power Systems, Inc. Rotary ramjet engine with flameholder extending to running clearance at engine casing interior wall
US20090113895A1 (en) * 2001-07-23 2009-05-07 Steele Robert C Vortex combustor for low NOx emissions when burning lean premixed high hydrogen content fuel
US20090158745A1 (en) * 2007-12-21 2009-06-25 Grossi Fabio G Ramjet Superheater
US8484980B1 (en) 2009-11-19 2013-07-16 The United States Of America As Represented By The Administrator Of National Aeronautics And Space Administration Dual-mode combustor
CN104295406A (en) * 2014-05-26 2015-01-21 清华大学 Rocket stamping combination engine with annular injection structure
CN110195881A (en) * 2018-02-26 2019-09-03 通用电气公司 Engine with rotation detonating combustion system

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US2867979A (en) * 1946-04-29 1959-01-13 Experiment Inc Apparatus for igniting fuels
US3102392A (en) * 1959-04-21 1963-09-03 Snecma Combustion equipment for jet propulsion units
US3514956A (en) * 1968-03-11 1970-06-02 William R Bray Injector-ram jet engine

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Publication number Priority date Publication date Assignee Title
US2867979A (en) * 1946-04-29 1959-01-13 Experiment Inc Apparatus for igniting fuels
US3102392A (en) * 1959-04-21 1963-09-03 Snecma Combustion equipment for jet propulsion units
US3514956A (en) * 1968-03-11 1970-06-02 William R Bray Injector-ram jet engine

Cited By (36)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4091731A (en) * 1976-07-06 1978-05-30 The United States Of America As Represented By The Secretary Of The Navy Fuel injection with flameholding
US4802639A (en) * 1984-09-28 1989-02-07 The Boeing Company Horizontal-takeoff transatmospheric launch system
FR2634005A1 (en) * 1987-05-05 1990-01-12 United Technologies Corp Pilotage fuel injector assembly for a supersonic statoreactor
US5109670A (en) * 1989-03-23 1992-05-05 General Electric Company Scramjet combustor
US5103633A (en) * 1989-03-23 1992-04-14 General Electric Company Scramjet combustor
US5072581A (en) * 1989-03-23 1991-12-17 General Electric Company Scramjet combustor
US5072582A (en) * 1989-03-23 1991-12-17 General Electric Company Scramjet combustor
US5081831A (en) * 1989-03-23 1992-01-21 General Electric Company Scramjet combustor
US5097663A (en) * 1989-03-23 1992-03-24 General Electric Company Scramjet combustor
EP0401107A1 (en) * 1989-05-29 1990-12-05 Societe Europeenne De Propulsion (S.E.P.) S.A. Combustion chamber for ram jet
US5333445A (en) * 1990-01-29 1994-08-02 General Electric Company Scramjet engine having improved fuel/air mixing
GB2243409A (en) * 1990-02-28 1991-10-30 Gen Electric Scramjet
US5255513A (en) * 1990-02-28 1993-10-26 General Electric Company Method of operating a scramjet including integrated inlet and combustor
US5085048A (en) * 1990-02-28 1992-02-04 General Electric Company Scramjet including integrated inlet and combustor
US5349815A (en) * 1991-08-27 1994-09-27 General Electric Company Scramjet combustor having a two-part, aft-facing step
US5253474A (en) * 1991-08-30 1993-10-19 General Electric Company Apparatus for supersonic combustion in a restricted length
US5280705A (en) * 1992-06-22 1994-01-25 General Electric Company Fuel injection system for scramjet engines
EP0693668A2 (en) * 1994-06-21 1996-01-24 Rockwell International Corporation Gas gun launched scramjet test projectile
EP0693668A3 (en) * 1994-06-21 1997-01-22 Rockwell International Corp Gas gun launched scramjet test projectile
US5544586A (en) * 1994-08-30 1996-08-13 The United States Of America As Represented By The Secretary Of The Army Solid fuel ramjet tubular projectile
US20100170263A1 (en) * 2001-07-23 2010-07-08 Ramgen Power Systems, Llc Vortex Combustor for Low NOX Emissions when Burning Lean Premixed High Hydrogen Content Fuel
US20040020211A1 (en) * 2001-07-23 2004-02-05 Ramgen Power Systems, Inc. Trapped vortex combustor
US6694743B2 (en) 2001-07-23 2004-02-24 Ramgen Power Systems, Inc. Rotary ramjet engine with flameholder extending to running clearance at engine casing interior wall
US7003961B2 (en) 2001-07-23 2006-02-28 Ramgen Power Systems, Inc. Trapped vortex combustor
US20090113895A1 (en) * 2001-07-23 2009-05-07 Steele Robert C Vortex combustor for low NOx emissions when burning lean premixed high hydrogen content fuel
US8312725B2 (en) 2001-07-23 2012-11-20 Ramgen Power Systems, Llc Vortex combustor for low NOX emissions when burning lean premixed high hydrogen content fuel
US7603841B2 (en) 2001-07-23 2009-10-20 Ramgen Power Systems, Llc Vortex combustor for low NOx emissions when burning lean premixed high hydrogen content fuel
WO2003010432A1 (en) * 2001-07-23 2003-02-06 Ramgen Power Systems, Inc. Rotary ramjet engine with flameholder extending to running clearance at engine casing interior wall
US20090158745A1 (en) * 2007-12-21 2009-06-25 Grossi Fabio G Ramjet Superheater
US8381528B2 (en) * 2007-12-21 2013-02-26 Grossi Aerospace, Inc. Ramjet superheater
US8528341B2 (en) 2007-12-21 2013-09-10 Grossi Aerospace, Inc. Ramjet superheater
US8484980B1 (en) 2009-11-19 2013-07-16 The United States Of America As Represented By The Administrator Of National Aeronautics And Space Administration Dual-mode combustor
US9745921B2 (en) 2009-11-19 2017-08-29 The United States Of America As Represented By The Administrator Of National Aeronautics And Space Administration Process for operating a dual-mode combustor
CN104295406A (en) * 2014-05-26 2015-01-21 清华大学 Rocket stamping combination engine with annular injection structure
CN104295406B (en) * 2014-05-26 2016-08-24 清华大学 A kind of rocket punching press combined engine with ring-like ejection structure
CN110195881A (en) * 2018-02-26 2019-09-03 通用电气公司 Engine with rotation detonating combustion system

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