US3899884A - Combustor systems - Google Patents
Combustor systems Download PDFInfo
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- US3899884A US3899884A US094289A US9428970A US3899884A US 3899884 A US3899884 A US 3899884A US 094289 A US094289 A US 094289A US 9428970 A US9428970 A US 9428970A US 3899884 A US3899884 A US 3899884A
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- Prior art keywords
- swirler
- air
- mixing chamber
- combustion chamber
- venturi
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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
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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
- F23R3/12—Air inlet arrangements for primary air inducing a vortex
- F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes
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- 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 improvements in combustor systems, particularly in the generation of a hot gas stream as in gas turbine engines.
- Combustor systems as employed in gas turbine engines, comprise a spray nozzle which injects fuel into a combustion zone. Once ignited, a continuous flame front is maintained in the combustion chamber as fuel and pressurized air flow therein and a high energy, hot gas stream is discharged therefrom.
- the object of the invention is to reduce the temperature levels of fuel nozzles and, in so doing, minimize the buildup of carbon thereon and further, to maintain minimal smoke generation in the combustion process.
- venturi into which the fuel spray is discharged.
- the acceleration provided by the initial portion of the venturi tube accelerates mixing air introduced around the nozzle to an extent sufficient to prevent combustion for a predetermined distance from the nozzle. By thus displacing the flame front from the nozzle, its temperature may be reduced sufficiently to prevent undesirable carbon buildup thereon.
- the divergent portion of the venturi tube re-expands the mixing air to discharge it into the combustion chamber at a relatively wide angle consistent with the smoke minimizing dispersion process referenced above.
- venturi tube to the means for introducing mixing air as well as in the amount of mixing air introduced.
- FIG. 1 is a schematic representation of a gas turbine engine employing a combustor of the type herein referenced;
- FIG. 2 is an enlarged longitudinal section illustrating the details of the combustor referenced in FIG. 1 as they embody the present invention
- FIG. 3 is a section taken on line IIl-IlI in FIG. 2;
- FIG. 4 is a longitudinal section, similar to FIG. 2, illustrating another embodiment of the invention.
- the illustrated engine comprises an axial flow compressor 10 which pressurizes air.
- This pressurized air flows through an annular passageway 12 to an annular combustor 14 where fuel is introduced.
- the pressurized air supports combustion of fuel within the combustor to generate a high energy level, hot gas stream.
- This hot gas stream drives a turbine 16 which in turn powers the rotor of the compressor 10.
- the hot gas stream is then converted to a useful output as by being discharged from a nozzle 17 to provide propulsive. thrust for an aircraft.
- the combustor 14 as illustrated in FIG. 2, comprises outer and inner liners 18 and 20 which define an annular combustion chamber 21. These liners are joined by compositely formed dome portion 22 at their upstream ends. Cylindrical conduits or mixing chambers 24 which open into the dome portion 22 and the combustion chamber 21. Transition segments 26 blend the conduit openings into the dome portion. An axial flow swirler 28 is mounted at the upstream end of each conduit 24.
- the swirler has a central opening which receives the discharge end 29 of a fuel spray nozzle 30.
- the nozzle 30 may take many forms but is preferably characterized by at least one orifice outlet which produces a conical spray discharge having a relatively large included angle relative to an axis a which extends lengthwise of the mixing chamber.
- the swirler 28 includes a row of passageways 32 annularly surrounding the axis of the nozzle discharge end 29. The passageways 32 are angled relative to the nozzle axis to produce a vortical flow field.
- the swirler 28 may be mounted on the upstream end of the conduit 24 in the manner taught in copending US. application Ser. No. 796,391, filed Feb. 4, 1969 and of common assignment.
- the referenced application also describes in further detail, the relationship of the swirler passageways 32 in obtaining a high degree of fuel dispersion for low smoke combustion.
- the passageway 12 is defined by outer and inner, generally cylindrical casings 34 and 36 which extend along the lengths of the liners 18 and 20 are respectively spaced therefrom to define annular passageways 38 and 40.
- An annular snout assembly 42 is secured to and projects upstream from the liners l8 and 20.
- the snout assembly has a central passageway 43 having an entrance facing the discharge passageway 12 and discharging into an annular chamber 44 surrounding the entrances to the mixing chambers, i.e., swirler passageways 32.
- the pressurized discharge from the compressor then is split into three annular flow-paths along passageways 38, 40 and 43.
- Air from the passageways 38 and 40 may enter the combustion chamber 21 to serve three functions. First, it may pass through relatively small holes 46 which are oriented to cool the liners 18 and 20. Second, it may enter relatively large holes 48 to penetrate the combustion chamber and supply requisite, primary air for the combustion process. Third, it may enter holes (not shown) further downstream as dilution air to reduce the temperature of the hot gas stream to a temperature compatible with the capabilities of the materials forming the turbine. The air entering holes 48 may also serve a dilution function.
- Air entering the snout passageway 43 and chamber 44 is then metered by and injected into the mixing chamber as discrete jets by the swirler passageways 32.
- the vortical flow of mixing air is highly effective in dispersing or mixing the fuel from the spray cone into fine droplets which support a combustion process at a flame front generally identified by the broken line in FIG. 2.
- the flame front is primarily controlled by two factors, once ignition is had. These are the presence and degree of a combustible mixture and the velocity of that mixture. In accordance with the present invention, these factors are taken into account in maintaining a desired distance between the flame front and the discharge end 29 of the fuel nozzle to minimize the temperature of the latter.
- a venturi tube 50 is disposed concentrically of the nozzle axis a in a surrounding relationship with the spray cone discharged by the nozzle.
- the inlet diameter of the venturi tube approximates the mean diameter of the annular row of swirler passageways 32.
- approximately one-third of the swirler mixing air is captured by and axially accelerated through the convergent portion of the venturi tube 50.
- the divergent portion of the venturi tube then re-expands the mixture to approximately the original cone shape of the nozzle discharge.
- the divergent portion of the venturi terminates at a generally tangent relationship with the spray cone.
- venturi tube 50 While maintaining the desired low smoke characteristics, the venturi tube 50 is highly effective in maintaining the flame front at a desired downstream distance. This distance can be controlled as desired by the length of venturi tube and its contraction ratio as well as the amount of mixing air flow therethrough, all of these factors being balanced so that the spray cone angle is relatively divergent as it discharges from the venturi tube into the combustion chamber.
- venturi tube may be expressed in another fashion in that it increases the total pressure in the core of the vortical flow field created by the axial flow swirler. Without the venturi, the low pressure of the vortical core would enable recirculating primary air to flow forwardly, as indicated by the broken arrows, and create a combustible mixture and flame front closely adjacent the nozzle discharge. With the venturi such recirculation occurs, as indicated by the solid arrows, but is more limited in the magnitude of its upstream movement.
- venturi tube is welded or otherwise bonded to the swirler 28 intermediate the lengths of the passageways 32. In so doing, added strength and rigidity is obtained for the swirler.
- FIG. 4 illustrates an embodiment of the invention wherein all of the mixing air flows through a venturi tube 52 which also serves as a mixing chamber for introducing the dispersed fuel mixture into the combustion zone.
- the action of this venturi in obtaining a desired downstream displacement of the flame front and a consequent reduction in nozzle temperature is essentially the same as previously described except that the velocity factor of the venturi is more predominant in displacing the flame front since a greater amount of air enters the venturi.
- the benefits of the broader aspects of the present invention can be attained where the amount of mixing air flow is sufficient to produce a combustible mixture in the mixing chamber.
- a combustor system comprising:
- a spray nozzle for discharging fuel as an atomized cone into said mixing chamber, the axis of said spray cone extending lengthwise of said mixing chamber toward said combustion chamber,
- venturi means surrounding said spray cone and through which at least a portion of the axially flowing air passes, and wherein the air introducing means additionally provide a vortical component to the air flow field,
- venturi means comprise a venturi tube mounted within said mixing chamber
- said air introducing means includes an annular row of generally radially on'ented passageways angled through said swirler,
- said axial swirler has a central opening through which the discharge end of the spray nozzle projects
- the inlet end of the venturi tube has a diameter intermediate the minimum and maximum diameters of the annular row of swirler passageways and is bonded to said swirler.
- a combustor as in claim 2 wherein: a pressurized plenum is provided, in combination with the upstream face of said axial swirler, and Produce a Combustible mixture thereinthe swirler passageways meter air into the mixing means are provided for introducing further pressurized air into said combustion chamber sufficient to
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Pressure-Spray And Ultrasonic-Wave- Spray Burners (AREA)
- Application Of Or Painting With Fluid Materials (AREA)
- Combustion Of Fluid Fuel (AREA)
- Nozzles (AREA)
Abstract
The disclosure shows two versions of providing a venturi around a fuel spray cone in a mixing chamber having an axial, vortical flow of pressurized air therethrough. This mixture is discharged into a combustion chamber. The venturi maintains desirable low smoke formation while spacing the flame front of the ignited mixture from the spray nozzle to reduce its temperature and thus prevent undesired carbon formation thereon.
Description
United States Patent Ekstedt Aug. 19, 1975 1 COMBUSTOR SYSTEMS 3,430,443 3/1969 Richardson 60/39.65 3,570,242 3/1971 Leonardi 431/183 [75] hvemo 3 3 Eksted" Cmcmnau, 3,589.127 6/1971 Kenworthy 60/39.65
[73] Assignee: General Electric Company, Primary ExaminerDoug1as Hart Cincinnati, Ohio Attorney, Agent, or Firm-Lee H. Sachs; Derek P. 22 Filed: Dec. 2, 1970 Lawrence The disclosure shows two versions of providing a ven- [52] U.S. Cl. 60/39.74 R; 431/183 i around a f l Spray Cone in a mixing Chamber [5 Clhaving an axial vortical flow of pressurized air there- Field of Search 60/3965 39-74; 431/183 through. This mixture is discharged into a combustion chamber. The venturi maintains desirable low smoke [56] References Clted formation while spacing the flame front of the ignited UNITED STATES PATENTS mixture from the spray nozzle to reduce its tempera- 1,290,607 1/1919 Lovekin 431/183 ture and thus Prevent undesired carbon formation 1,322,999 11/1919 Bester 60/39.65 thereon- 2,398,654 4/1946 Lubbock... 60/39.65 3 Cl 4 D 3,285,007 11/1966 Carlisle 60/39.74 R raw'ng gums III! III/III/lfl PATENTEUAUG-I ems 3, 899 884 SWEET 1 [1F 2 INVENTOR. EDWARD E. EKSTEDT nrromiEY- PATENTEU W75 3, 899 884 SZIET 2 UF 2 INVENTOR. EDWARD E. EKSTEDT ATTORNEY- COMBUSTOR SYSTEMS The invention herein described was made in the course of or under a contract, or a subcontract thereunder, with the U.S. Department of the Air Force.
The present invention relates to improvements in combustor systems, particularly in the generation of a hot gas stream as in gas turbine engines.
Combustor systems, as employed in gas turbine engines, comprise a spray nozzle which injects fuel into a combustion zone. Once ignited, a continuous flame front is maintained in the combustion chamber as fuel and pressurized air flow therein and a high energy, hot gas stream is discharged therefrom.
Recent emphasis had been placed on the problem of minimizing smoke generated in the combustion process and discharged to the atmosphere from gas turbine engines. One highly successful approach to this problem has been to discharge the fuel into a conduit or mixing chamber which opens into the combustion chamber. Pressurized air is introduced axially and swirled vorti cally about the discharge axis of the spray nozzle. In so doing, the turbulent flow field created produces a highly dispersed, over-stoichiometric mixture of fuel and air in the conduit. This mixture is discharged in a generally conical fashion into the combustion chamber where it is further dispersed by additional air entering other openings in the combustion chamber. This addi tional air is sufficient to sustain combustion and a flame front is produced in this mixing zone at the mixing chamber discharge. Because of the high degree of dispersion obtained, over-rich fuel zones are essentially eliminated and smoke is minimized to the point of being virtually undetectable.
While smoke is thus effectively reduced, the frequency of maintenance of the fuel nozzles has increased. This is due to the coking of fuel on the nozzles which adversley affects the pattern of the fuel spray cone discharged therefrom. Various forms of spray nozzles and air shrouds therefor have had limited degrees of success in overcoming this problem and still maintaining minimized smoke levels. This lack of success is attributed to the fine dispersion of fuel which recirculates, in localized regions, and contacts the discharge end of the nozzle. The fuel then cokes or carbonizes when it contacts the nozzle due to the high temperature thereof.
Accordingly, the object of the invention is to reduce the temperature levels of fuel nozzles and, in so doing, minimize the buildup of carbon thereon and further, to maintain minimal smoke generation in the combustion process.
These ends are broadly attained by providing a venturi means into which the fuel spray is discharged. The acceleration provided by the initial portion of the venturi tube accelerates mixing air introduced around the nozzle to an extent sufficient to prevent combustion for a predetermined distance from the nozzle. By thus displacing the flame front from the nozzle, its temperature may be reduced sufficiently to prevent undesirable carbon buildup thereon. The divergent portion of the venturi tube re-expands the mixing air to discharge it into the combustion chamber at a relatively wide angle consistent with the smoke minimizing dispersion process referenced above.
Other features are found in the relationship of the venturi tube to the means for introducing mixing air as well as in the amount of mixing air introduced.
The above and other related objects and features of the invention will be apparent from a reading of the following description of the disclosure found in the accompanying drawings and the novelty thereof pointed out in the appended claims.
In the drawings:
FIG. 1 is a schematic representation of a gas turbine engine employing a combustor of the type herein referenced;
FIG. 2 is an enlarged longitudinal section illustrating the details of the combustor referenced in FIG. 1 as they embody the present invention;
FIG. 3 is a section taken on line IIl-IlI in FIG. 2; and
FIG. 4 is a longitudinal section, similar to FIG. 2, illustrating another embodiment of the invention.
Referencing FIG. 1, the illustrated engine comprises an axial flow compressor 10 which pressurizes air. This pressurized air flows through an annular passageway 12 to an annular combustor 14 where fuel is introduced. The pressurized air supports combustion of fuel within the combustor to generate a high energy level, hot gas stream. This hot gas stream drives a turbine 16 which in turn powers the rotor of the compressor 10. The hot gas stream is then converted to a useful output as by being discharged from a nozzle 17 to provide propulsive. thrust for an aircraft.
The combustor 14, as illustrated in FIG. 2, comprises outer and inner liners 18 and 20 which define an annular combustion chamber 21. These liners are joined by compositely formed dome portion 22 at their upstream ends. Cylindrical conduits or mixing chambers 24 which open into the dome portion 22 and the combustion chamber 21. Transition segments 26 blend the conduit openings into the dome portion. An axial flow swirler 28 is mounted at the upstream end of each conduit 24.
The swirler has a central opening which receives the discharge end 29 of a fuel spray nozzle 30. The nozzle 30 may take many forms but is preferably characterized by at least one orifice outlet which produces a conical spray discharge having a relatively large included angle relative to an axis a which extends lengthwise of the mixing chamber. The swirler 28 includes a row of passageways 32 annularly surrounding the axis of the nozzle discharge end 29. The passageways 32 are angled relative to the nozzle axis to produce a vortical flow field.
The swirler 28 may be mounted on the upstream end of the conduit 24 in the manner taught in copending US. application Ser. No. 796,391, filed Feb. 4, 1969 and of common assignment. The referenced application also describes in further detail, the relationship of the swirler passageways 32 in obtaining a high degree of fuel dispersion for low smoke combustion.
Introduction of pressurized air from the annular compressor discharge passageway 12 into the combustion chamber 21 will now be described. The passageway 12 is defined by outer and inner, generally cylindrical casings 34 and 36 which extend along the lengths of the liners 18 and 20 are respectively spaced therefrom to define annular passageways 38 and 40. An annular snout assembly 42 is secured to and projects upstream from the liners l8 and 20. The snout assembly has a central passageway 43 having an entrance facing the discharge passageway 12 and discharging into an annular chamber 44 surrounding the entrances to the mixing chambers, i.e., swirler passageways 32. The pressurized discharge from the compressor then is split into three annular flow-paths along passageways 38, 40 and 43.
Air from the passageways 38 and 40 may enter the combustion chamber 21 to serve three functions. First, it may pass through relatively small holes 46 which are oriented to cool the liners 18 and 20. Second, it may enter relatively large holes 48 to penetrate the combustion chamber and supply requisite, primary air for the combustion process. Third, it may enter holes (not shown) further downstream as dilution air to reduce the temperature of the hot gas stream to a temperature compatible with the capabilities of the materials forming the turbine. The air entering holes 48 may also serve a dilution function.
Air entering the snout passageway 43 and chamber 44 is then metered by and injected into the mixing chamber as discrete jets by the swirler passageways 32. The vortical flow of mixing air is highly effective in dispersing or mixing the fuel from the spray cone into fine droplets which support a combustion process at a flame front generally identified by the broken line in FIG. 2.
The flame front is primarily controlled by two factors, once ignition is had. These are the presence and degree of a combustible mixture and the velocity of that mixture. In accordance with the present invention, these factors are taken into account in maintaining a desired distance between the flame front and the discharge end 29 of the fuel nozzle to minimize the temperature of the latter.
To this end, a venturi tube 50 is disposed concentrically of the nozzle axis a in a surrounding relationship with the spray cone discharged by the nozzle. The inlet diameter of the venturi tube approximates the mean diameter of the annular row of swirler passageways 32. Thus approximately one-third of the swirler mixing air is captured by and axially accelerated through the convergent portion of the venturi tube 50. This creates a condition of where the mixture, at the throat of the venturi, is sufficiently over-stoichiometric as not to support combustion and the mixture velocity is increased to further deter combustion. The divergent portion of the venturi tube then re-expands the mixture to approximately the original cone shape of the nozzle discharge. In this connection, it will be noted that the divergent portion of the venturi terminates at a generally tangent relationship with the spray cone.
While maintaining the desired low smoke characteristics, the venturi tube 50 is highly effective in maintaining the flame front at a desired downstream distance. This distance can be controlled as desired by the length of venturi tube and its contraction ratio as well as the amount of mixing air flow therethrough, all of these factors being balanced so that the spray cone angle is relatively divergent as it discharges from the venturi tube into the combustion chamber.
The effect of the venturi tube may be expressed in another fashion in that it increases the total pressure in the core of the vortical flow field created by the axial flow swirler. Without the venturi, the low pressure of the vortical core would enable recirculating primary air to flow forwardly, as indicated by the broken arrows, and create a combustible mixture and flame front closely adjacent the nozzle discharge. With the venturi such recirculation occurs, as indicated by the solid arrows, but is more limited in the magnitude of its upstream movement.
It will be noted that the venturi tube is welded or otherwise bonded to the swirler 28 intermediate the lengths of the passageways 32. In so doing, added strength and rigidity is obtained for the swirler.
FIG. 4 illustrates an embodiment of the invention wherein all of the mixing air flows through a venturi tube 52 which also serves as a mixing chamber for introducing the dispersed fuel mixture into the combustion zone. The action of this venturi in obtaining a desired downstream displacement of the flame front and a consequent reduction in nozzle temperature is essentially the same as previously described except that the velocity factor of the venturi is more predominant in displacing the flame front since a greater amount of air enters the venturi. In fact, the benefits of the broader aspects of the present invention can be attained where the amount of mixing air flow is sufficient to produce a combustible mixture in the mixing chamber.
It has been demonstrated that the described use of a venturi is highly efiective in preventing carbon buildup on fuel nozzles and, thus, assuring long maintenancefree operation. At the same time, low smoke levels continue to be maintained. 7
While the invention has been described with reference to an annular combustion system, it is equally applicable to a cannular system. Further, in the broader aspects of the invention, the mixing air could flow through the mixing chamber with little or no vortical component. These and other variations in the described embodiments will occur to those skilled in the art within the spirit and scope of the present inventive concepts which are to be derived solely from the appended claims.
Having thus described the invention, what is claimed as novel and desired to be secured by Letters Patent of the United States is:
1. A combustor system comprising:
a combustion chamber,
a mixing chamber opening into the upstream end of the combustion chamber,
a spray nozzle for discharging fuel as an atomized cone into said mixing chamber, the axis of said spray cone extending lengthwise of said mixing chamber toward said combustion chamber,
means for introducing pressurized air into the upstream end of said mixing chamber and producing an axial flow field for dispersing the fuel in fine droplets,
venturi means surrounding said spray cone and through which at least a portion of the axially flowing air passes, and wherein the air introducing means additionally provide a vortical component to the air flow field,
the venturi means comprise a venturi tube mounted within said mixing chamber,
a cylindrical conduit and an axial swirler at the upstream end of the conduit define said mixing chamber,
said air introducing means includes an annular row of generally radially on'ented passageways angled through said swirler,
said axial swirler has a central opening through which the discharge end of the spray nozzle projects, and
the inlet end of the venturi tube has a diameter intermediate the minimum and maximum diameters of the annular row of swirler passageways and is bonded to said swirler.
6 2. A combustion chamber as in claim I wherein the chamber at a rate insufficient to produce a comdivergent discharge end of the venturi tube terminates bustibie mixture with the f l introduced by said generally as a tangent to the spray cone discharge from Spray noflle and the nozzle.
3. A combustor as in claim 2 wherein: a pressurized plenum is provided, in combination with the upstream face of said axial swirler, and Produce a Combustible mixture thereinthe swirler passageways meter air into the mixing means are provided for introducing further pressurized air into said combustion chamber sufficient to
Claims (3)
1. A combustor system comprising: a combustion chamber, a mixing chamber opening into the upstream end of the combustion chamber, a spray nozzle for discharging fuel as an atomized cone into said mixing chamber, the axis of said spray cone extending lengthwise of said mixing chamber toward said combustion chamber, means for introducing pressurized air into the upstream end of said mixing chamber and producing an axial flow field for dispersing the fuel in fine droplets, venturi means surrounding said spray cone and through which at least a portion of the axially flowing air passes, and wherein the air introducing means additionally provide a vortical component to the air flow field, the venturi means comprise a venturi tube mounted within said mixing chamber, a cylindrical conduit and an axial swirler at the upstream end of the conduit define said mixing chamber, said air introducing means includes an annular row of generally radially oriented passageways angled through said swirler, said axial swirler has a central opening through which the discharge end of the spray nozzle projects, and the inlet end of the venturi tube has a diameter intermediate the minimum and maximum diameters of the annular row of swirler passageways and is bonded to said swirler.
2. A combustion chamber as in claim 1 wherein the divergent discharge end of the venturi tube terminates generally as a tangent to the spray cone discharge from the nozzle.
3. A combustor as in claim 2 wherein: a pressurized plenum is provided, in combination with the upstream face of said axial swirler, and the swirler passageways meter air into the mixing chamber at a rate insufficient to produce a combustible mixture with the fuel introduced by said spray nozzle, and means are provided for introducing further pressurized air into said combustion chamber sufficient to produce a combustible mixture therein.
Priority Applications (12)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US094289A US3899884A (en) | 1970-12-02 | 1970-12-02 | Combustor systems |
CA121,071A CA943352A (en) | 1970-12-02 | 1971-08-23 | Combustor systems |
AU32634/71A AU452434B2 (en) | 1970-12-02 | 1971-08-23 | Combustor systems |
CH1235971A CH537519A (en) | 1970-12-02 | 1971-08-24 | Combustion device, in particular for gas turbine engines |
NL7111784A NL7111784A (en) | 1970-12-02 | 1971-08-26 | |
GB4034571A GB1353335A (en) | 1970-12-02 | 1971-08-27 | Combustion equipment |
DE2143012A DE2143012C3 (en) | 1970-12-02 | 1971-08-27 | Burner arrangement in a gas turbine combustor |
BE771990A BE771990A (en) | 1970-12-02 | 1971-08-31 | IMPROVEMENTS TO COMBUSTION SYSTEMS |
NO3228/71A NO134997C (en) | 1970-12-02 | 1971-08-31 | |
FR7131637A FR2116363B1 (en) | 1970-12-02 | 1971-09-01 | |
JP6704971A JPS5521250B1 (en) | 1970-12-02 | 1971-09-02 | |
DK431671A DK132539C (en) | 1970-12-02 | 1971-09-02 | COMBUSTION SYSTEM |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
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US094289A US3899884A (en) | 1970-12-02 | 1970-12-02 | Combustor systems |
Publications (1)
Publication Number | Publication Date |
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US3899884A true US3899884A (en) | 1975-08-19 |
Family
ID=22244287
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
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US094289A Expired - Lifetime US3899884A (en) | 1970-12-02 | 1970-12-02 | Combustor systems |
Country Status (12)
Country | Link |
---|---|
US (1) | US3899884A (en) |
JP (1) | JPS5521250B1 (en) |
AU (1) | AU452434B2 (en) |
BE (1) | BE771990A (en) |
CA (1) | CA943352A (en) |
CH (1) | CH537519A (en) |
DE (1) | DE2143012C3 (en) |
DK (1) | DK132539C (en) |
FR (1) | FR2116363B1 (en) |
GB (1) | GB1353335A (en) |
NL (1) | NL7111784A (en) |
NO (1) | NO134997C (en) |
Cited By (55)
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US4215535A (en) * | 1978-01-19 | 1980-08-05 | United Technologies Corporation | Method and apparatus for reducing nitrous oxide emissions from combustors |
FR2529954A1 (en) * | 1982-07-06 | 1984-01-13 | Gen Electric | CARBURATION ASSEMBLY AND METHOD FOR CONTROLLING THE ASSEMBLY |
US4445338A (en) * | 1981-10-23 | 1984-05-01 | The United States Of America As Represented By The Secretary Of The Navy | Swirler assembly for a vorbix augmentor |
US4458479A (en) * | 1981-10-13 | 1984-07-10 | General Motors Corporation | Diffuser for gas turbine engine |
US4693074A (en) * | 1983-11-26 | 1987-09-15 | Rolls-Royce Plc | Combustion apparatus for a gas turbine engine |
WO1988006231A1 (en) * | 1987-02-19 | 1988-08-25 | Hi-Tech International Laboratory Company Limited | Combustion system for internal combustion engine and combustor used therefor |
US4934145A (en) * | 1988-10-12 | 1990-06-19 | United Technologies Corporation | Combustor bulkhead heat shield assembly |
US4974416A (en) * | 1987-04-27 | 1990-12-04 | General Electric Company | Low coke fuel injector for a gas turbine engine |
US4982564A (en) * | 1988-12-14 | 1991-01-08 | General Electric Company | Turbine engine with air and steam cooling |
US5117637A (en) * | 1990-08-02 | 1992-06-02 | General Electric Company | Combustor dome assembly |
US5123248A (en) * | 1990-03-28 | 1992-06-23 | General Electric Company | Low emissions combustor |
US5209067A (en) * | 1990-10-17 | 1993-05-11 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. | Gas turbine combustion chamber wall structure for minimizing cooling film disturbances |
US5261239A (en) * | 1991-02-28 | 1993-11-16 | Societe Nationale D'etude Et De Construction De Motors D'aviation | Lean premixture combustion-chamber comprising a counterflow enclosure to stabilize the premixture flame |
US5274995A (en) * | 1992-04-27 | 1994-01-04 | General Electric Company | Apparatus and method for atomizing water in a combustor dome assembly |
US5477671A (en) * | 1993-07-07 | 1995-12-26 | Mowill; R. Jan | Single stage premixed constant fuel/air ratio combustor |
US5490388A (en) * | 1992-09-28 | 1996-02-13 | Asea Brown Boveri Ltd. | Gas turbine combustion chamber having a diffuser |
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US5613357A (en) * | 1993-07-07 | 1997-03-25 | Mowill; R. Jan | Star-shaped single stage low emission combustor system |
US5628182A (en) * | 1993-07-07 | 1997-05-13 | Mowill; R. Jan | Star combustor with dilution ports in can portions |
US5638674A (en) * | 1993-07-07 | 1997-06-17 | Mowill; R. Jan | Convectively cooled, single stage, fully premixed controllable fuel/air combustor with tangential admission |
JP2641551B2 (en) * | 1987-02-19 | 1997-08-13 | 有限会社ハイ・テク・インターナショナル研究所 | Combustion system for internal combustion engine and combustion device |
US5924276A (en) * | 1996-07-17 | 1999-07-20 | Mowill; R. Jan | Premixer with dilution air bypass valve assembly |
US6021635A (en) * | 1996-12-23 | 2000-02-08 | Parker-Hannifin Corporation | Dual orifice liquid fuel and aqueous flow atomizing nozzle having an internal mixing chamber |
US6220034B1 (en) | 1993-07-07 | 2001-04-24 | R. Jan Mowill | Convectively cooled, single stage, fully premixed controllable fuel/air combustor |
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US6363726B1 (en) | 2000-09-29 | 2002-04-02 | General Electric Company | Mixer having multiple swirlers |
US6367262B1 (en) | 2000-09-29 | 2002-04-09 | General Electric Company | Multiple annular swirler |
US6381964B1 (en) | 2000-09-29 | 2002-05-07 | General Electric Company | Multiple annular combustion chamber swirler having atomizing pilot |
US6418726B1 (en) | 2001-05-31 | 2002-07-16 | General Electric Company | Method and apparatus for controlling combustor emissions |
US6460344B1 (en) | 1999-05-07 | 2002-10-08 | Parker-Hannifin Corporation | Fuel atomization method for turbine combustion engines having aerodynamic turning vanes |
US6474071B1 (en) | 2000-09-29 | 2002-11-05 | General Electric Company | Multiple injector combustor |
WO2002088602A1 (en) * | 2001-04-25 | 2002-11-07 | Pratt & Whitney Canada Corp. | Turbine premixing combustor |
US6484489B1 (en) | 2001-05-31 | 2002-11-26 | General Electric Company | Method and apparatus for mixing fuel to decrease combustor emissions |
US6543233B2 (en) * | 2001-02-09 | 2003-04-08 | General Electric Company | Slot cooled combustor liner |
US6550251B1 (en) | 1997-12-18 | 2003-04-22 | General Electric Company | Venturiless swirl cup |
US6564555B2 (en) * | 2001-05-24 | 2003-05-20 | Allison Advanced Development Company | Apparatus for forming a combustion mixture in a gas turbine engine |
US20030196440A1 (en) * | 1999-05-07 | 2003-10-23 | Erlendur Steinthorsson | Fuel nozzle for turbine combustion engines having aerodynamic turning vanes |
US6925809B2 (en) | 1999-02-26 | 2005-08-09 | R. Jan Mowill | Gas turbine engine fuel/air premixers with variable geometry exit and method for controlling exit velocities |
US20070119177A1 (en) * | 2005-11-30 | 2007-05-31 | General Electric Company | Turbine engine fuel nozzles and methods of assembling the same |
US20070227150A1 (en) * | 2006-03-31 | 2007-10-04 | Pratt & Whitney Canada Corp. | Combustor |
US20090212139A1 (en) * | 2008-02-21 | 2009-08-27 | Delavan Inc | Radially outward flowing air-blast fuel injector for gas turbine engine |
US20100162713A1 (en) * | 2008-12-31 | 2010-07-01 | Shui-Chi Li | Cooled flameholder swirl cup |
US20100199684A1 (en) * | 2008-12-31 | 2010-08-12 | Edward Claude Rice | Combustion liner assembly support |
US7779636B2 (en) | 2005-05-04 | 2010-08-24 | Delavan Inc | Lean direct injection atomizer for gas turbine engines |
US20110067403A1 (en) * | 2009-09-18 | 2011-03-24 | Delavan Inc | Lean burn injectors having multiple pilot circuits |
EP2592351A1 (en) | 2011-11-09 | 2013-05-15 | Delavan, Inc. | Staged pilots in pure airblast injectors for gas turbine engines |
CN103836647A (en) * | 2014-02-27 | 2014-06-04 | 中国科学院工程热物理研究所 | Venturi tube flow channel wall face structure |
US8893500B2 (en) | 2011-05-18 | 2014-11-25 | Solar Turbines Inc. | Lean direct fuel injector |
US8919132B2 (en) | 2011-05-18 | 2014-12-30 | Solar Turbines Inc. | Method of operating a gas turbine engine |
US9046039B2 (en) | 2008-05-06 | 2015-06-02 | Rolls-Royce Plc | Staged pilots in pure airblast injectors for gas turbine engines |
US9182124B2 (en) | 2011-12-15 | 2015-11-10 | Solar Turbines Incorporated | Gas turbine and fuel injector for the same |
US20160209038A1 (en) * | 2013-08-30 | 2016-07-21 | United Technologies Corporation | Dual fuel nozzle with swirling axial gas injection for a gas turbine engine |
US11371711B2 (en) * | 2018-11-28 | 2022-06-28 | General Electric Company | Rotating detonation combustor with offset inlet |
FR3142533A1 (en) * | 2022-11-28 | 2024-05-31 | Safran Aircraft Engines | Combustion chamber for turbomachine |
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CH577627A5 (en) * | 1974-04-03 | 1976-07-15 | Bbc Sulzer Turbomaschinen | |
DE2641605C2 (en) * | 1975-12-24 | 1986-06-19 | General Electric Co., Schenectady, N.Y. | Device for supplying air and fuel |
AT351131B (en) * | 1976-07-05 | 1979-07-10 | Henkel Kgaa | TEXTILE DETERGENTS |
US4180974A (en) * | 1977-10-31 | 1980-01-01 | General Electric Company | Combustor dome sleeve |
JPS5634550A (en) * | 1979-08-28 | 1981-04-06 | Fuji Heavy Ind Ltd | Fluid-pressure holding device for car brake |
JPS58448A (en) * | 1981-06-22 | 1983-01-05 | Kazuichi Tsukamoto | Two-system type brake hydraulic pressure keeping apparatus |
JPS58100160U (en) * | 1981-12-28 | 1983-07-07 | 富士重工業株式会社 | Brake hydraulic pressure holding device for automobile brakes |
EP0153842B1 (en) * | 1984-02-29 | 1988-07-27 | LUCAS INDUSTRIES public limited company | Combustion equipment |
FR2673454B1 (en) * | 1991-02-28 | 1995-01-13 | Snecma | COMBUSTION CHAMBER COMPRISING A BOTTOM WALL COMPRISING A PLURALITY OF PARTIAL CONE TRUNKS. |
GB2398375A (en) * | 2003-02-14 | 2004-08-18 | Alstom | A mixer for two fluids having a venturi shape |
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Cited By (76)
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US4215535A (en) * | 1978-01-19 | 1980-08-05 | United Technologies Corporation | Method and apparatus for reducing nitrous oxide emissions from combustors |
US4458479A (en) * | 1981-10-13 | 1984-07-10 | General Motors Corporation | Diffuser for gas turbine engine |
US4445338A (en) * | 1981-10-23 | 1984-05-01 | The United States Of America As Represented By The Secretary Of The Navy | Swirler assembly for a vorbix augmentor |
FR2529954A1 (en) * | 1982-07-06 | 1984-01-13 | Gen Electric | CARBURATION ASSEMBLY AND METHOD FOR CONTROLLING THE ASSEMBLY |
US4584834A (en) * | 1982-07-06 | 1986-04-29 | General Electric Company | Gas turbine engine carburetor |
US4693074A (en) * | 1983-11-26 | 1987-09-15 | Rolls-Royce Plc | Combustion apparatus for a gas turbine engine |
WO1988006231A1 (en) * | 1987-02-19 | 1988-08-25 | Hi-Tech International Laboratory Company Limited | Combustion system for internal combustion engine and combustor used therefor |
JP2641551B2 (en) * | 1987-02-19 | 1997-08-13 | 有限会社ハイ・テク・インターナショナル研究所 | Combustion system for internal combustion engine and combustion device |
US4974416A (en) * | 1987-04-27 | 1990-12-04 | General Electric Company | Low coke fuel injector for a gas turbine engine |
US4934145A (en) * | 1988-10-12 | 1990-06-19 | United Technologies Corporation | Combustor bulkhead heat shield assembly |
US4982564A (en) * | 1988-12-14 | 1991-01-08 | General Electric Company | Turbine engine with air and steam cooling |
US5123248A (en) * | 1990-03-28 | 1992-06-23 | General Electric Company | Low emissions combustor |
US5117637A (en) * | 1990-08-02 | 1992-06-02 | General Electric Company | Combustor dome assembly |
US5209067A (en) * | 1990-10-17 | 1993-05-11 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. | Gas turbine combustion chamber wall structure for minimizing cooling film disturbances |
US5261239A (en) * | 1991-02-28 | 1993-11-16 | Societe Nationale D'etude Et De Construction De Motors D'aviation | Lean premixture combustion-chamber comprising a counterflow enclosure to stabilize the premixture flame |
US5274995A (en) * | 1992-04-27 | 1994-01-04 | General Electric Company | Apparatus and method for atomizing water in a combustor dome assembly |
US5490388A (en) * | 1992-09-28 | 1996-02-13 | Asea Brown Boveri Ltd. | Gas turbine combustion chamber having a diffuser |
US5477671A (en) * | 1993-07-07 | 1995-12-26 | Mowill; R. Jan | Single stage premixed constant fuel/air ratio combustor |
US5613357A (en) * | 1993-07-07 | 1997-03-25 | Mowill; R. Jan | Star-shaped single stage low emission combustor system |
US5628182A (en) * | 1993-07-07 | 1997-05-13 | Mowill; R. Jan | Star combustor with dilution ports in can portions |
US5638674A (en) * | 1993-07-07 | 1997-06-17 | Mowill; R. Jan | Convectively cooled, single stage, fully premixed controllable fuel/air combustor with tangential admission |
US5572862A (en) * | 1993-07-07 | 1996-11-12 | Mowill Rolf Jan | Convectively cooled, single stage, fully premixed fuel/air combustor for gas turbine engine modules |
US5765363A (en) * | 1993-07-07 | 1998-06-16 | Mowill; R. Jan | Convectively cooled, single stage, fully premixed controllable fuel/air combustor with tangential admission |
US6220034B1 (en) | 1993-07-07 | 2001-04-24 | R. Jan Mowill | Convectively cooled, single stage, fully premixed controllable fuel/air combustor |
US5592819A (en) * | 1994-03-10 | 1997-01-14 | Societe Nationale D'etude Et De Construction De Moteurs D'aviation S.N.E.C.M.A. | Pre-mixing injection system for a turbojet engine |
US5924276A (en) * | 1996-07-17 | 1999-07-20 | Mowill; R. Jan | Premixer with dilution air bypass valve assembly |
US6021635A (en) * | 1996-12-23 | 2000-02-08 | Parker-Hannifin Corporation | Dual orifice liquid fuel and aqueous flow atomizing nozzle having an internal mixing chamber |
US6708498B2 (en) | 1997-12-18 | 2004-03-23 | General Electric Company | Venturiless swirl cup |
US6550251B1 (en) | 1997-12-18 | 2003-04-22 | General Electric Company | Venturiless swirl cup |
US6925809B2 (en) | 1999-02-26 | 2005-08-09 | R. Jan Mowill | Gas turbine engine fuel/air premixers with variable geometry exit and method for controlling exit velocities |
US6560964B2 (en) | 1999-05-07 | 2003-05-13 | Parker-Hannifin Corporation | Fuel nozzle for turbine combustion engines having aerodynamic turning vanes |
US20030196440A1 (en) * | 1999-05-07 | 2003-10-23 | Erlendur Steinthorsson | Fuel nozzle for turbine combustion engines having aerodynamic turning vanes |
US6460344B1 (en) | 1999-05-07 | 2002-10-08 | Parker-Hannifin Corporation | Fuel atomization method for turbine combustion engines having aerodynamic turning vanes |
US6883332B2 (en) | 1999-05-07 | 2005-04-26 | Parker-Hannifin Corporation | Fuel nozzle for turbine combustion engines having aerodynamic turning vanes |
EP1106919A1 (en) * | 1999-12-10 | 2001-06-13 | General Electric Company | Methods and apparatus for decreasing combustor emissions |
US6609377B2 (en) | 2000-09-29 | 2003-08-26 | General Electric Company | Multiple injector combustor |
US6381964B1 (en) | 2000-09-29 | 2002-05-07 | General Electric Company | Multiple annular combustion chamber swirler having atomizing pilot |
US6367262B1 (en) | 2000-09-29 | 2002-04-09 | General Electric Company | Multiple annular swirler |
US6363726B1 (en) | 2000-09-29 | 2002-04-02 | General Electric Company | Mixer having multiple swirlers |
US6474071B1 (en) | 2000-09-29 | 2002-11-05 | General Electric Company | Multiple injector combustor |
US6543233B2 (en) * | 2001-02-09 | 2003-04-08 | General Electric Company | Slot cooled combustor liner |
WO2002088602A1 (en) * | 2001-04-25 | 2002-11-07 | Pratt & Whitney Canada Corp. | Turbine premixing combustor |
US6564555B2 (en) * | 2001-05-24 | 2003-05-20 | Allison Advanced Development Company | Apparatus for forming a combustion mixture in a gas turbine engine |
US6418726B1 (en) | 2001-05-31 | 2002-07-16 | General Electric Company | Method and apparatus for controlling combustor emissions |
US6484489B1 (en) | 2001-05-31 | 2002-11-26 | General Electric Company | Method and apparatus for mixing fuel to decrease combustor emissions |
US20100287946A1 (en) * | 2005-05-04 | 2010-11-18 | Delavan Inc | Lean direct injection atomizer for gas turbine engines |
US8156746B2 (en) | 2005-05-04 | 2012-04-17 | Delavan Inc | Lean direct injection atomizer for gas turbine engines |
US7779636B2 (en) | 2005-05-04 | 2010-08-24 | Delavan Inc | Lean direct injection atomizer for gas turbine engines |
US20070119177A1 (en) * | 2005-11-30 | 2007-05-31 | General Electric Company | Turbine engine fuel nozzles and methods of assembling the same |
JP2007155318A (en) * | 2005-11-30 | 2007-06-21 | General Electric Co <Ge> | Turbine engine fuel nozzle and turbine engine |
US7788927B2 (en) | 2005-11-30 | 2010-09-07 | General Electric Company | Turbine engine fuel nozzles and methods of assembling the same |
US20070227150A1 (en) * | 2006-03-31 | 2007-10-04 | Pratt & Whitney Canada Corp. | Combustor |
US7950233B2 (en) | 2006-03-31 | 2011-05-31 | Pratt & Whitney Canada Corp. | Combustor |
US8146837B2 (en) | 2008-02-21 | 2012-04-03 | Delavan Inc | Radially outward flowing air-blast fuel injection for gas turbine engine |
US8128007B2 (en) | 2008-02-21 | 2012-03-06 | Delavan Inc | Radially outward flowing air-blast fuel injector for gas turbine engine |
US7926744B2 (en) | 2008-02-21 | 2011-04-19 | Delavan Inc | Radially outward flowing air-blast fuel injector for gas turbine engine |
US20110089264A1 (en) * | 2008-02-21 | 2011-04-21 | Delavan Inc. | Radially outward flowing air-blast fuel injection for gas turbine engine |
US20110089262A1 (en) * | 2008-02-21 | 2011-04-21 | Delavan Inc | Radially outward flowing air-blast fuel injector for gas turbine engine |
US20090212139A1 (en) * | 2008-02-21 | 2009-08-27 | Delavan Inc | Radially outward flowing air-blast fuel injector for gas turbine engine |
US9046039B2 (en) | 2008-05-06 | 2015-06-02 | Rolls-Royce Plc | Staged pilots in pure airblast injectors for gas turbine engines |
US20100162713A1 (en) * | 2008-12-31 | 2010-07-01 | Shui-Chi Li | Cooled flameholder swirl cup |
US20100199684A1 (en) * | 2008-12-31 | 2010-08-12 | Edward Claude Rice | Combustion liner assembly support |
US8281597B2 (en) | 2008-12-31 | 2012-10-09 | General Electric Company | Cooled flameholder swirl cup |
US9046272B2 (en) * | 2008-12-31 | 2015-06-02 | Rolls-Royce Corporation | Combustion liner assembly having a mount stake coupled to an upstream support |
US8607571B2 (en) | 2009-09-18 | 2013-12-17 | Delavan Inc | Lean burn injectors having a main fuel circuit and one of multiple pilot fuel circuits with prefiliming air-blast atomizers |
US20110067403A1 (en) * | 2009-09-18 | 2011-03-24 | Delavan Inc | Lean burn injectors having multiple pilot circuits |
US9239167B2 (en) | 2009-09-18 | 2016-01-19 | Rolls-Royce Plc | Lean burn injectors having multiple pilot circuits |
US8893500B2 (en) | 2011-05-18 | 2014-11-25 | Solar Turbines Inc. | Lean direct fuel injector |
US8919132B2 (en) | 2011-05-18 | 2014-12-30 | Solar Turbines Inc. | Method of operating a gas turbine engine |
EP2592351A1 (en) | 2011-11-09 | 2013-05-15 | Delavan, Inc. | Staged pilots in pure airblast injectors for gas turbine engines |
US9182124B2 (en) | 2011-12-15 | 2015-11-10 | Solar Turbines Incorporated | Gas turbine and fuel injector for the same |
US20160209038A1 (en) * | 2013-08-30 | 2016-07-21 | United Technologies Corporation | Dual fuel nozzle with swirling axial gas injection for a gas turbine engine |
US10228137B2 (en) * | 2013-08-30 | 2019-03-12 | United Technologies Corporation | Dual fuel nozzle with swirling axial gas injection for a gas turbine engine |
CN103836647A (en) * | 2014-02-27 | 2014-06-04 | 中国科学院工程热物理研究所 | Venturi tube flow channel wall face structure |
US11371711B2 (en) * | 2018-11-28 | 2022-06-28 | General Electric Company | Rotating detonation combustor with offset inlet |
FR3142533A1 (en) * | 2022-11-28 | 2024-05-31 | Safran Aircraft Engines | Combustion chamber for turbomachine |
Also Published As
Publication number | Publication date |
---|---|
DE2143012B2 (en) | 1974-10-31 |
DE2143012C3 (en) | 1975-06-12 |
DK132539C (en) | 1976-05-24 |
BE771990A (en) | 1971-12-31 |
DE2143012A1 (en) | 1972-06-15 |
NL7111784A (en) | 1972-06-06 |
CH537519A (en) | 1973-05-31 |
AU3263471A (en) | 1973-03-01 |
GB1353335A (en) | 1974-05-15 |
NO134997B (en) | 1976-10-11 |
DK132539B (en) | 1975-12-22 |
CA943352A (en) | 1974-03-12 |
JPS5521250B1 (en) | 1980-06-09 |
FR2116363A1 (en) | 1972-07-13 |
AU452434B2 (en) | 1974-09-05 |
FR2116363B1 (en) | 1974-12-20 |
NO134997C (en) | 1977-01-19 |
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