US20180010800A1 - Shock compression based supersonic combustor - Google Patents
Shock compression based supersonic combustor Download PDFInfo
- Publication number
- US20180010800A1 US20180010800A1 US15/182,447 US201615182447A US2018010800A1 US 20180010800 A1 US20180010800 A1 US 20180010800A1 US 201615182447 A US201615182447 A US 201615182447A US 2018010800 A1 US2018010800 A1 US 2018010800A1
- Authority
- US
- United States
- Prior art keywords
- combustor
- oxidizer
- supersonic
- fuel
- core
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
Images
Classifications
-
- 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
-
- 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
- F23R3/26—Controlling the air flow
-
- 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
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
-
- 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
- F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
- F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
-
- 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
- F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
- F23R3/58—Cyclone or vortex type combustion chambers
-
- 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
- F23R7/00—Intermittent or explosive combustion chambers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
- F23C—METHODS OR APPARATUS FOR COMBUSTION USING FLUID FUEL OR SOLID FUEL SUSPENDED IN A CARRIER GAS OR AIR
- F23C2205/00—Pulsating combustion
-
- 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
- Embodiments of the invention relates to the field of propulsion. More specifically, the invention relates to combustors based on detonation combustion for gas turbine or other engine based application.
- a combustor of an engine is a component that houses the burning process of fuel-oxidizer (F/O) mixture or some combination thereof.
- the combustion process in majority of the engines in operation is subsonic i.e. the rate at which the F/O mixture burns is slower than the local speed of sound. This is a constant pressure combustion process also known as deflagration.
- a detonation combustor houses a similar burning process, however the rate at which the F/O mixture is burnt is faster than the local speed of sound. This a constant volume process also known as detonation. A detonation process is thermodynamically superior to the deflagration process.
- the challenge with existing detonation combustors and detonation based engines are the valves and ignition system required to maintain the pulsed regime of the detonation waves and its unsteady flow characteristics.
- the present invention aims to address this issue with a unique new technique.
- the present invention utilizes shock compression and/or shock reflection to carry out the ignition process.
- the present invention also allows for comparably more uniform and steady flow at the end of the combustion chamber in order to reduce the fatiguing from the pulsed or unsteady nature of other concepts and technologies.
- One or more embodiments of the invention comprises of an injector module supporting both fuel and oxidizer, a combustion core with grooves wherein the detonation occurs and propagates towards the exit and an outer shell that envelopes the injector module and the combustion core.
- a select mass flow of fuel is injected into to the corresponding groove of the combustion core from the fuel nozzle located at the injection module.
- the said fuel nozzle would be an atomizer.
- the oxidizer line ejects a micro shock wave into the same groove of the combustion core.
- the shock wave compresses the injected fuel against a wall of the groove until the critical pressure, dictated by the Chapman-Jouguet detonation theory, is achieved upon which deflagration to detonation transition (DDT) or direct detonation combustion is initiated.
- DDT deflagration to detonation transition
- the detonation waves and the reflected Mach waves then propagate out of the combustion section of the combustor core to the mixing section where it then becomes periodic with respect to the reactive flow from the other grooves before exiting the combustor completely.
- FIG. 1 is an exploded isometric view of the supersonic combustor in accordance with one or more embodiments of the present invention.
- FIG. 2 is a downstream exploded view of the supersonic combustor illustrating the main components comprising the injector module, combustor core and the outer shell, in accordance with one or more embodiments of the present invention.
- FIG. 3A is an isometric view of the injector module in accordance with one or more embodiments of the present invention.
- FIG. 3B is an isometric view of the injector module internals detailing the oxidizer and fuel lines therein in accordance with one or more embodiments of the present invention.
- FIG. 4A is a front cross-sectional view of the injector module in accordance with one or more embodiments of the present invention.
- FIG. 4B is a close-up view of a set of fuel and oxidizer injector ports in accordance with one or more embodiments of the present invention.
- FIG. 5 is an upstream cross-sectional view of the combustor core in accordance with one or more embodiments of the present invention.
- FIG. 6 is a side profile view of the combustor core in accordance with one or more embodiments of the present invention.
- FIG. 7 is an upstream cross-sectional view of the combustor core in accordance with one or more embodiments of the present invention.
- FIG. 8 is an isometric view of the outer shell in accordance with one or more embodiments of the present invention.
- FIG. 9 is a side profile view of the injector module geometrically mated to the combustor core and partially depicting the combustion process in accordance with one or more embodiments of the present invention.
- first”, “second” and the like, herein do not denote any order, quantity or importance, but rather are used to distinguish one element from another, and the terms “a” and “an” herein do not denote a limitation of quantity, but rather denote the presence of at least one of the referenced item.
- FIGS. 1-9 One or more embodiments of the present invention will now be described with references to FIGS. 1-9 .
- FIG. 1 Illustrated in FIG. 1 is an exploded isometric view of an exemplary embodiment of the supersonic combustor 100 of the present invention.
- the supersonic combustor 100 comprises injector module 110 , combustor core 120 and the outer shell 130 .
- One or more embodiments of the present invention may include alignment keys, e.g. 114 on the injector module 110 and keys 124 on combustor core 120 .
- the down-stream front-side face 113 of the injector module 110 is coupled (coincident) with the upstream back-side face 123 of the combustor core 120 , and the keys 114 of the injector module 110 are aligned with the keys 124 of the combustor core.
- the alignment of the injector module 110 and combustor core 120 ensures proper placement of the oxidizer line injector port 118 and fuel line injector port 119 within the entrance 125 of the corresponding grooves 121 of the combustor core 120 .
- the aligned keys 114 and 124 slide into the corresponding slots 134 of the outer shell 130 , locking the various components into position.
- the sets of fuel lines 116 and oxidizer lines 115 are also illustrated in FIG. 3A .
- the pressurized oxidizer from the compressor (usually air) moves into a plenum which then channels the required amount into the oxidizer lines 115 through oxidizer inlet ports 111 of the injector module 110 as well as any cooling systems utilized by the device and/or engine.
- FIG. 2 Illustrated in FIG. 2 , is a downstream exploded isometric view of the exemplary embodiment of supersonic combustor 100 detailing the assembly of the injector module 110 , combustor core 120 and the outer shell 130 .
- This illustration provides an example of the positioning of the line injector ports 119 and 118 of the fuel line 116 and oxidizer line 115 , respectively, at their corresponding entrances 125 (see FIG. 6 ) to grooves 121 of the combustor core 120 .
- the direction of the fuel line 116 is in the z-axis of the supersonic combustor 100 and is coupled through fuel inlet ports 112 , whereas the downstream injector port 118 of the oxidizer line 115 and in turn any valve mechanism that would be attached is at an angle ⁇ to the opposite sidewall 122 of the corresponding groove 121 belonging to the combustor core 120 .
- FIGS. 3A and 3B Illustrated in FIGS. 3A and 3B are exemplary illustrations of the downstream isometric views of the injector module 110 and its internals in accordance with one or more embodiments of the present invention. These illustrations detail the keys 114 and the front-side face 113 of the injector module 110 that mate with the rear-side face 123 and the keys 124 of the combustor core 120 . Each set of the fuel lines 116 and oxidizer lines 115 of the injector module 110 correspond to a groove 121 of the combustor core.
- the fuel system (storage, pump and pipes outside of the present invention) delivers the fuel to the fuel line 116 through the fuel inlet ports 112 .
- the fuel line 116 may be configured to transport the fluid (e.g.
- the oxidizer line 115 Delivered from the compressor and plenum of the engine, the oxidizer line 115 is filled with oxidizer.
- a valve located at the downstream injector port 118 of the oxidizer line 115 forms a quick circular opening, analogous to that of a diaphragm burst in a conventional shock tube. This results in production of a micro-shock wave of a specific strength.
- the length, internal diameter, and cross-sectional area of the oxidizer lines 115 and fuel lines 116 may vary based on the scale, performance, efficiency, flow characteristics and/or any other desired parameter.
- FIGS. 4A and 4B Illustrated in FIGS. 4A and 4B are a front cross-sectional view and a close-up of a set of fuel injector ports 119 and oxidizer injector ports 118 in accordance with one or more embodiments of the injector module of the present invention.
- the number of sets of fuel lines 116 and oxidizer lines 115 correspond to the number of grooves 121 , e.g. eight (8), in the combustor core 120 .
- the number of fuel lines 116 and oxidizer lines 115 as well as the corresponding grooves 121 of the combustor core 120 can be varied based on the scale of the present invention and its application.
- the injector port 118 of the oxidizer line 115 is at an acute angle ⁇ with respect to the front-side cross-section face 113 of the injector module 110 and the axial direction (z-axis) of the combustor 100 . This angle may vary based on the combustion requirements and/or any other parameters that might contribute to the operation of the device.
- the linear and circumferential spacing between the fuel lines 116 and oxidizer lines 115 , and the thickness of the injector module 110 may also vary due to the device scale, flow characteristics and any other optimization parameters.
- FIGS. 5, 6 and 7 Illustrated in FIGS. 5, 6 and 7 are exemplary embodiments of the combustor core 120 of the present invention. These drawings illustrate the upstream cross-section, side profile and the upstream isometric views of the combustor core 120 respectively.
- FIG. 5 an exemplary configuration of the keys 124 of the combustor core 120 are illustrated. In the illustrative embodiment present herein, keys 124 are configured to align with the keys 114 of the injector module 110 .
- the location of the keys 114 and 124 may change based on the device scale and other design criteria, however, alignment of the keys 114 , 124 are used to ensure the correct placement of the fuel injector ports 119 and oxidizer injector ports 118 of the injector module 110 , within the corresponding entrance 125 of the grooves 121 .
- the rear face 123 of the combustor core 120 is shown, which mates with the front face 113 of the injector module 110 allowing the keys 114 and 124 to align.
- a center hole 127 is present along the entire axial length of the combustor core 120 .
- the center hole 127 may house the shaft and bearings of an engine to which the present invention is fitted.
- FIG. 5 also shows the entrance of the grooves 125 where the corresponding fuel injector ports 119 and oxidizer injector ports 118 are positioned to enable the device operation.
- the combustor core comprises a combustion section 128 and a mixing section 129 .
- a configuration for grooves 121 along the axial and circumferential direction of combustion section 128 of the combustor core 120 are illustrated.
- the path of the grooves 121 turns along with the circumference of the combustor core 120 , i.e. spiral, to add a component of rotation to the reactive flow.
- the configuration and path of the grooves 121 may be different, e.g. straight, linear and/or follow a complex curve based on the device optimization requirements.
- the dimension of the walls 122 can vary based on the shock compression and detonation requirements.
- the axial and circumferential lengths of the grooves 121 and its walls 122 may vary in order to condition the reactive flow for the entrance of the turbine section downstream of the present invention.
- the mixing section 129 of the combustor core 120 downstream of the grooves 121 , are illustrated in FIGS. 6 and 7 .
- the mixing section 129 allows the periodic reactive flow from each groove 121 to merge and create a more uniform flow downstream.
- the axial length of the mixing section 129 can vary in order to ensure proper mixing of the reactive flows.
- FIG. 8 Illustrated in FIG. 8 is an illustration of the outer shell 130 of the exemplary embodiment of the present invention. This drawing details the general configuration of the outer shell 130 of the supersonic combustor 100 from an upstream isometric view.
- the slots 134 can be seen, into which the keys 114 and 124 of injector module 110 and combustor core 120 respectively, slide in. This ensures that the mated and aligned injector module 110 and combustor core 120 are locked into their appropriate positions, allowing the supersonic combustor 100 to be installed and operated with an engine.
- the design of the outer shell 130 is illustrated as a simple hollow cylinder, those of skill in the art would appreciate that it could take other configurations to accommodate other systems, such as cooling, mounting or control systems based on the device and/or engine operation.
- the inner configuration of the outer shell 130 may be of different design, for example, dictated by combustion and flow regime requirements as well as the compatibility to an engine's turbine section.
- Alignment, locking and mounting techniques for the present invention and its components may change as required. Additional components may be added to the present invention in order to refine its design and function based on its application. For example, those of skill in the arts would appreciate that cooling, mounting and control systems may change the general design of the components of the supersonic combustor detailed herein.
- the actuation of the valves at the downstream oxidizer injector ports 118 of the oxidizer lines 115 and the fuel injection from the fuel lines 116 will be governed by a control system.
- the key parameters of the present invention responsible for the throttling of the device are mass flow of fuel injected from each fuel line 116 into its corresponding grooves 121 ; strength and mass flow of the micro shock wave generated by the oxidizer line 115 and its valve system; and the frequency of fuel 116 and oxidizer 115 fired within their corresponding grooves 121 .
- FIG. 9 is a side profile view of the injector module geometrically mated to the combustor core and partially depicting the combustion process in accordance with one or more embodiments of the present invention.
- the mated injector module 110 and combustor core 120 are configured to fit inside the outer shell 130 .
- a valve e.g. an iris valve, or any other suitable mechanism placed at the downstream injector port 118 of the oxidizer line 115 forms a circular opening to mimic a diaphragm burst of a conventional shock-tube, thereby creating a micro shock wave.
- the oxidizer line injector port aperture, timing and other operational parameters may be governed by a control system.
- the generated micro shock wave enters the grove 121 , it compresses the injected fuel mass against the sidewall 122 of the groove 121 opposite to the location of the oxidizer line injector port 118 .
- Detonation combustion occurs as the critical Chapman-Jouguet (CJ) conditions are achieved during the compression process.
- the detonation waves and reflected Mach waves then propagate down the groove 121 .
- the downstream cross section 126 of the grooves 121 can be tailored to change the regime of the propagating reactive flow between subsonic and supersonic based on the application of the present invention (e.g. supersonic regime for thrust and subsonic regime for power generation applications).
- the reactive flow exits the combustion section 128 and enters the mixing section 129 of the combustor core 120 , it merges with the periodic reactive flow from other grooves 121 , resulting in a uniform flow at the exit 139 of supersonic combustor 100 .
Abstract
A supersonic combustor containing an injector module, a combustor core and an outer shell. The injector module houses both fuel and oxidizer nozzles. The combustor core contains grooves within which the combustion process takes place. The outer shell holds both the injector module and the combustor core and allows for other cooling, mounting and structural mechanisms required for operation.
Description
- Embodiments of the invention relates to the field of propulsion. More specifically, the invention relates to combustors based on detonation combustion for gas turbine or other engine based application.
- A combustor of an engine is a component that houses the burning process of fuel-oxidizer (F/O) mixture or some combination thereof. The combustion process in majority of the engines in operation is subsonic i.e. the rate at which the F/O mixture burns is slower than the local speed of sound. This is a constant pressure combustion process also known as deflagration.
- A detonation combustor houses a similar burning process, however the rate at which the F/O mixture is burnt is faster than the local speed of sound. This a constant volume process also known as detonation. A detonation process is thermodynamically superior to the deflagration process.
- The challenge with existing detonation combustors and detonation based engines are the valves and ignition system required to maintain the pulsed regime of the detonation waves and its unsteady flow characteristics. The present invention aims to address this issue with a unique new technique.
- Where other detonation combustor technologies rely on spark or flame as the source of ignition, the present invention utilizes shock compression and/or shock reflection to carry out the ignition process. The present invention also allows for comparably more uniform and steady flow at the end of the combustion chamber in order to reduce the fatiguing from the pulsed or unsteady nature of other concepts and technologies.
- One or more embodiments of the invention comprises of an injector module supporting both fuel and oxidizer, a combustion core with grooves wherein the detonation occurs and propagates towards the exit and an outer shell that envelopes the injector module and the combustion core.
- A select mass flow of fuel is injected into to the corresponding groove of the combustion core from the fuel nozzle located at the injection module. In the case of a liquid fuel, the said fuel nozzle would be an atomizer. Accurately timed, the oxidizer line ejects a micro shock wave into the same groove of the combustion core. The shock wave compresses the injected fuel against a wall of the groove until the critical pressure, dictated by the Chapman-Jouguet detonation theory, is achieved upon which deflagration to detonation transition (DDT) or direct detonation combustion is initiated. The detonation waves and the reflected Mach waves then propagate out of the combustion section of the combustor core to the mixing section where it then becomes periodic with respect to the reactive flow from the other grooves before exiting the combustor completely.
- The above and other aspects, features and advantages of the invention will be more apparent from the following more particular description thereof, presented in conjunction with the following drawings wherein:
-
FIG. 1 is an exploded isometric view of the supersonic combustor in accordance with one or more embodiments of the present invention. -
FIG. 2 is a downstream exploded view of the supersonic combustor illustrating the main components comprising the injector module, combustor core and the outer shell, in accordance with one or more embodiments of the present invention. -
FIG. 3A is an isometric view of the injector module in accordance with one or more embodiments of the present invention. -
FIG. 3B is an isometric view of the injector module internals detailing the oxidizer and fuel lines therein in accordance with one or more embodiments of the present invention. -
FIG. 4A is a front cross-sectional view of the injector module in accordance with one or more embodiments of the present invention. -
FIG. 4B is a close-up view of a set of fuel and oxidizer injector ports in accordance with one or more embodiments of the present invention. -
FIG. 5 is an upstream cross-sectional view of the combustor core in accordance with one or more embodiments of the present invention. -
FIG. 6 is a side profile view of the combustor core in accordance with one or more embodiments of the present invention. -
FIG. 7 is an upstream cross-sectional view of the combustor core in accordance with one or more embodiments of the present invention. -
FIG. 8 is an isometric view of the outer shell in accordance with one or more embodiments of the present invention. -
FIG. 9 is a side profile view of the injector module geometrically mated to the combustor core and partially depicting the combustion process in accordance with one or more embodiments of the present invention. - The present invention comprising shock compression based supersonic combustor will now be described. In the following exemplary description numerous specific details are set forth in order to provide a more thorough understanding of embodiments of the invention. It will be apparent, however, to an artisan of ordinary skill that the present invention may be practiced without incorporating all aspects of the specific details described herein. Furthermore, although steps or processes are set forth in an exemplary order to provide an understanding of one or more systems and methods, the exemplary order is not meant to be limiting. One of ordinary skill in the art would recognize that the steps or processes may be performed in a different order, and that one or more steps or processes may be performed simultaneously or in multiple process flows without departing from the spirit or the scope of the invention. In other instances, specific features, quantities, or measurements well known to those of ordinary skill in the art have not been described in detail so as not to obscure the invention. It should be noted that although examples of the invention are set forth herein, the claims, and the full scope of any equivalents, are what define the metes and bounds of the invention.
- For a better understanding of the disclosed embodiment, its operating advantages, and the specified object attained by its uses, reference should be made to the accompanying drawings and descriptive matter in which there are illustrated exemplary disclosed embodiments. The disclosed embodiments are not intended to be limited to the specific forms set forth herein. It is understood that various omissions and substitutions of equivalents are contemplated as circumstances may suggest or render expedient, but these are intended to cover the application or implementation.
- Those of skill in the art would appreciate that the dimensions, geometric parameters and the components detailed in the drawings of the present invention are subjected to change based on the device application, scale, and/or related flow characteristics in order to ensure optimal efficiency and performance.
- The term “first”, “second” and the like, herein do not denote any order, quantity or importance, but rather are used to distinguish one element from another, and the terms “a” and “an” herein do not denote a limitation of quantity, but rather denote the presence of at least one of the referenced item.
- One or more embodiments of the present invention will now be described with references to
FIGS. 1-9 . - Illustrated in
FIG. 1 is an exploded isometric view of an exemplary embodiment of thesupersonic combustor 100 of the present invention. As illustrated, thesupersonic combustor 100 comprisesinjector module 110,combustor core 120 and theouter shell 130. One or more embodiments of the present invention may include alignment keys, e.g. 114 on theinjector module 110 andkeys 124 oncombustor core 120. As illustrated, the down-stream front-side face 113 of theinjector module 110 is coupled (coincident) with the upstream back-side face 123 of thecombustor core 120, and thekeys 114 of theinjector module 110 are aligned with thekeys 124 of the combustor core. The alignment of theinjector module 110 andcombustor core 120 ensures proper placement of the oxidizerline injector port 118 and fuelline injector port 119 within theentrance 125 of thecorresponding grooves 121 of thecombustor core 120. Thealigned keys corresponding slots 134 of theouter shell 130, locking the various components into position. The sets offuel lines 116 andoxidizer lines 115 are also illustrated inFIG. 3A . The pressurized oxidizer from the compressor (usually air) moves into a plenum which then channels the required amount into theoxidizer lines 115 throughoxidizer inlet ports 111 of theinjector module 110 as well as any cooling systems utilized by the device and/or engine. - Illustrated in
FIG. 2 , is a downstream exploded isometric view of the exemplary embodiment ofsupersonic combustor 100 detailing the assembly of theinjector module 110,combustor core 120 and theouter shell 130. This illustration provides an example of the positioning of theline injector ports fuel line 116 andoxidizer line 115, respectively, at their corresponding entrances 125 (seeFIG. 6 ) to grooves 121 of thecombustor core 120. The direction of thefuel line 116 is in the z-axis of thesupersonic combustor 100 and is coupled throughfuel inlet ports 112, whereas thedownstream injector port 118 of theoxidizer line 115 and in turn any valve mechanism that would be attached is at an angle θ to theopposite sidewall 122 of thecorresponding groove 121 belonging to thecombustor core 120. - Illustrated in
FIGS. 3A and 3B are exemplary illustrations of the downstream isometric views of theinjector module 110 and its internals in accordance with one or more embodiments of the present invention. These illustrations detail thekeys 114 and the front-side face 113 of theinjector module 110 that mate with the rear-side face 123 and thekeys 124 of thecombustor core 120. Each set of thefuel lines 116 andoxidizer lines 115 of theinjector module 110 correspond to agroove 121 of the combustor core. The fuel system (storage, pump and pipes outside of the present invention) delivers the fuel to thefuel line 116 through thefuel inlet ports 112. Thefuel line 116 may be configured to transport the fluid (e.g. fuel or fuel/air mixture) to thecorresponding groove 121 of thecombustor core 120 or it can act as a retainer for a different pipe and nozzle/injector depending on the type of fluid and the application. Delivered from the compressor and plenum of the engine, theoxidizer line 115 is filled with oxidizer. Upon achieving critical mass and pressure, a valve located at thedownstream injector port 118 of theoxidizer line 115 forms a quick circular opening, analogous to that of a diaphragm burst in a conventional shock tube. This results in production of a micro-shock wave of a specific strength. The length, internal diameter, and cross-sectional area of theoxidizer lines 115 andfuel lines 116 may vary based on the scale, performance, efficiency, flow characteristics and/or any other desired parameter. - Illustrated in
FIGS. 4A and 4B are a front cross-sectional view and a close-up of a set offuel injector ports 119 andoxidizer injector ports 118 in accordance with one or more embodiments of the injector module of the present invention. As illustrated, there are one or more, preferably a plurality (e.g. eight (8)), sets offuel lines 116 andoxidizer lines 115 on theinjector module 110. The number of sets offuel lines 116 andoxidizer lines 115 correspond to the number ofgrooves 121, e.g. eight (8), in thecombustor core 120. The number offuel lines 116 andoxidizer lines 115 as well as the correspondinggrooves 121 of thecombustor core 120 can be varied based on the scale of the present invention and its application. Theinjector port 118 of theoxidizer line 115 is at an acute angle θ with respect to the front-side cross-section face 113 of theinjector module 110 and the axial direction (z-axis) of thecombustor 100. This angle may vary based on the combustion requirements and/or any other parameters that might contribute to the operation of the device. The linear and circumferential spacing between thefuel lines 116 andoxidizer lines 115, and the thickness of theinjector module 110 may also vary due to the device scale, flow characteristics and any other optimization parameters. - Illustrated in
FIGS. 5, 6 and 7 are exemplary embodiments of thecombustor core 120 of the present invention. These drawings illustrate the upstream cross-section, side profile and the upstream isometric views of thecombustor core 120 respectively. InFIG. 5 , an exemplary configuration of thekeys 124 of thecombustor core 120 are illustrated. In the illustrative embodiment present herein,keys 124 are configured to align with thekeys 114 of theinjector module 110. The location of thekeys keys fuel injector ports 119 andoxidizer injector ports 118 of theinjector module 110, within thecorresponding entrance 125 of thegrooves 121. Therear face 123 of thecombustor core 120 is shown, which mates with thefront face 113 of theinjector module 110 allowing thekeys center hole 127 is present along the entire axial length of thecombustor core 120. Thecenter hole 127 may house the shaft and bearings of an engine to which the present invention is fitted.FIG. 5 also shows the entrance of thegrooves 125 where the correspondingfuel injector ports 119 andoxidizer injector ports 118 are positioned to enable the device operation. - As illustrated in
FIG. 6 , the combustor core comprises acombustion section 128 and amixing section 129. A configuration forgrooves 121 along the axial and circumferential direction ofcombustion section 128 of thecombustor core 120 are illustrated. In the present exemplary embodiment of thecombustor core 120, the path of thegrooves 121 turns along with the circumference of thecombustor core 120, i.e. spiral, to add a component of rotation to the reactive flow. Those of skill in the art would appreciate that the configuration and path of thegrooves 121 may be different, e.g. straight, linear and/or follow a complex curve based on the device optimization requirements. The dimension of thewalls 122 can vary based on the shock compression and detonation requirements. The axial and circumferential lengths of thegrooves 121 and itswalls 122 may vary in order to condition the reactive flow for the entrance of the turbine section downstream of the present invention. - The
mixing section 129 of thecombustor core 120, downstream of thegrooves 121, are illustrated inFIGS. 6 and 7 . Themixing section 129 allows the periodic reactive flow from eachgroove 121 to merge and create a more uniform flow downstream. The axial length of themixing section 129 can vary in order to ensure proper mixing of the reactive flows. - Illustrated in
FIG. 8 is an illustration of theouter shell 130 of the exemplary embodiment of the present invention. This drawing details the general configuration of theouter shell 130 of thesupersonic combustor 100 from an upstream isometric view. Theslots 134 can be seen, into which thekeys injector module 110 andcombustor core 120 respectively, slide in. This ensures that the mated and alignedinjector module 110 andcombustor core 120 are locked into their appropriate positions, allowing thesupersonic combustor 100 to be installed and operated with an engine. Although the design of theouter shell 130 is illustrated as a simple hollow cylinder, those of skill in the art would appreciate that it could take other configurations to accommodate other systems, such as cooling, mounting or control systems based on the device and/or engine operation. Similarly, the inner configuration of theouter shell 130, may be of different design, for example, dictated by combustion and flow regime requirements as well as the compatibility to an engine's turbine section. - Alignment, locking and mounting techniques for the present invention and its components may change as required. Additional components may be added to the present invention in order to refine its design and function based on its application. For example, those of skill in the arts would appreciate that cooling, mounting and control systems may change the general design of the components of the supersonic combustor detailed herein. The actuation of the valves at the downstream
oxidizer injector ports 118 of theoxidizer lines 115 and the fuel injection from thefuel lines 116 will be governed by a control system. The key parameters of the present invention responsible for the throttling of the device are mass flow of fuel injected from eachfuel line 116 into itscorresponding grooves 121; strength and mass flow of the micro shock wave generated by theoxidizer line 115 and its valve system; and the frequency offuel 116 andoxidizer 115 fired within theircorresponding grooves 121. -
FIG. 9 is a side profile view of the injector module geometrically mated to the combustor core and partially depicting the combustion process in accordance with one or more embodiments of the present invention. The matedinjector module 110 andcombustor core 120 are configured to fit inside theouter shell 130. - During the combustion process, as the fuel is injected into the
groove 121 via thefuel line 116, a valve, e.g. an iris valve, or any other suitable mechanism placed at thedownstream injector port 118 of theoxidizer line 115 forms a circular opening to mimic a diaphragm burst of a conventional shock-tube, thereby creating a micro shock wave. The oxidizer line injector port aperture, timing and other operational parameters may be governed by a control system. As the generated micro shock wave enters thegrove 121, it compresses the injected fuel mass against thesidewall 122 of thegroove 121 opposite to the location of the oxidizerline injector port 118. Detonation combustion occurs as the critical Chapman-Jouguet (CJ) conditions are achieved during the compression process. The detonation waves and reflected Mach waves then propagate down thegroove 121. Thedownstream cross section 126 of thegrooves 121 can be tailored to change the regime of the propagating reactive flow between subsonic and supersonic based on the application of the present invention (e.g. supersonic regime for thrust and subsonic regime for power generation applications). As the reactive flow exits thecombustion section 128 and enters themixing section 129 of thecombustor core 120, it merges with the periodic reactive flow fromother grooves 121, resulting in a uniform flow at theexit 139 ofsupersonic combustor 100. - While the invention herein disclosed has been described by means of specific embodiments and applications thereof, numerous modifications and variations could be made thereto by those skilled in the art without departing from the scope of the invention set forth in the claims.
Claims (15)
1. A supersonic combustor comprising:
an injector module comprising at least one set of fuel and oxidizer lines, wherein each of said fuel lines includes a fuel line injector port and each of said oxidizer lines includes an oxidizer line injector port;
a combustor core coupled to said injector module, wherein said combustor core comprises a combustion section towards its proximal end and a mixing section towards its distal end, wherein the combustion section comprises a corresponding groove for each set of fuel and oxidizer lines, wherein each groove is positioned axially and circumferentially along the combustor core and has an entrance at said proximal end of said combustor core, wherein said oxidizer line injector port and said fuel line injector port are coupled to said entrance of said corresponding groove; and
an outer shell, wherein said injector module and said combustor core are coupled inside said outer shell.
2. The supersonic combustor of claim 1 , wherein each groove comprises a sidewall configured to provide a desired shock compression and detonation.
3. The supersonic combustor of claim 1 , wherein the oxidizer injector port further comprises a valve attached at an angle θ to an opposite sidewall.
4. The supersonic combustor of claim 1 , wherein said oxidizer line is configured for generating micro shock waves.
5. The supersonic combustor of claim 1 , further comprising a cooling system coupled to said outer shell.
6. The supersonic combustor of claim 1 , wherein said outer shell further comprises mounting components.
7. The supersonic combustor of claim 3 , wherein actuation of the valve at the oxidizer injector port and fuel injection into the groove is governed by a control system.
8. The supersonic combustor of claim 1 , wherein each of said groove is spiral.
9. The supersonic combustor of claim 1 , wherein a downstream cross-section of the groove is configured to change a propagating reactive flow in said groove between subsonic and supersonic based on the application.
10. A supersonic combustor comprising:
an injector module comprising a plurality of fuel and oxidizer lines, wherein each of said fuel lines includes a fuel line injector port and each of said oxidizer lines includes an oxidizer line injector port; and
a combustor core coupled to said injector module, wherein said combustor core comprises a combustion section towards its proximal end and a mixing section towards its distal end, wherein the combustion section comprises a plurality of grooves positioned axially and circumferentially along the combustor core, wherein at least one fuel line injector port and one oxidizer line injector port are coupled to an entrance of one of said plurality of grooves of said combustor core.
11. The supersonic combustor of claim 10 , further comprising an outer shell, wherein said injector module and said combustor core are coupled inside said outer shell.
12. The supersonic combustor of claim 10 , wherein each one of said plurality of grooves is spiral.
13. The supersonic combustor of claim 10 , wherein each one of said plurality of grooves comprises a sidewall configured to provide a desired shock compression and detonation.
14. The supersonic combustor of claim 10 , wherein the oxidizer injector port further comprises a valve attached at an angle θ to an opposite sidewall.
15. The supersonic combustor of claim 10 , wherein each of said plurality of oxidizer lines is configured for generating micro shock waves.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US15/182,447 US20180010800A1 (en) | 2016-06-14 | 2016-06-14 | Shock compression based supersonic combustor |
Applications Claiming Priority (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US15/182,447 US20180010800A1 (en) | 2016-06-14 | 2016-06-14 | Shock compression based supersonic combustor |
Publications (1)
Publication Number | Publication Date |
---|---|
US20180010800A1 true US20180010800A1 (en) | 2018-01-11 |
Family
ID=60892597
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US15/182,447 Abandoned US20180010800A1 (en) | 2016-06-14 | 2016-06-14 | Shock compression based supersonic combustor |
Country Status (1)
Country | Link |
---|---|
US (1) | US20180010800A1 (en) |
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
CN111207009A (en) * | 2019-12-26 | 2020-05-29 | 中国空气动力研究与发展中心 | Method for initiating oblique detonation wave in supersonic velocity airflow by using external instantaneous energy source |
US11105511B2 (en) | 2018-12-14 | 2021-08-31 | General Electric Company | Rotating detonation propulsion system |
US11236908B2 (en) | 2018-10-24 | 2022-02-01 | General Electric Company | Fuel staging for rotating detonation combustor |
US11371711B2 (en) | 2018-11-28 | 2022-06-28 | General Electric Company | Rotating detonation combustor with offset inlet |
Citations (14)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6062018A (en) * | 1993-04-14 | 2000-05-16 | Adroit Systems, Inc. | Pulse detonation electrical power generation apparatus with water injection |
US6460342B1 (en) * | 1999-04-26 | 2002-10-08 | Advanced Research & Technology Institute | Wave rotor detonation engine |
US6845620B2 (en) * | 2001-07-06 | 2005-01-25 | Mohamed Razi Nalim | Rotary ejector enhanced pulsed detonation system and method |
US7100360B2 (en) * | 2002-12-30 | 2006-09-05 | United Technologies Corporation | Pulsed combustion engine |
US20060254252A1 (en) * | 2005-05-13 | 2006-11-16 | General Electric Company | Pulse detonation assembly and hybrid engine |
US20070180810A1 (en) * | 2006-02-03 | 2007-08-09 | General Electric Company | Pulse detonation combustor with folded flow path |
US20090102203A1 (en) * | 2007-10-23 | 2009-04-23 | Lu Frank K | System and method for power production using a hybrid helical detonation device |
US20090165438A1 (en) * | 2007-12-26 | 2009-07-02 | Occhipinti Anthony C | Pulse detonation engine |
US20090193786A1 (en) * | 2008-02-01 | 2009-08-06 | General Electric Company | System And Method Of Continuous Detonation In A Gas Turbine Engine |
US20090196733A1 (en) * | 2008-02-01 | 2009-08-06 | General Electric Company | Rotary Pressure Rise Combustor For A Gas Turbine Engine |
US20100242435A1 (en) * | 2009-03-30 | 2010-09-30 | Alliant Techsystems Inc. | Helical cross flow (hcf) pulse detonation engine |
US20150184860A1 (en) * | 2013-12-27 | 2015-07-02 | United Technologies Corporation | Spiral pulse detonation tube configuration |
US20150300630A1 (en) * | 2012-11-07 | 2015-10-22 | Exponential Technologies, Inc. | Pressure-gain combustion apparatus and method |
US20180363555A1 (en) * | 2017-06-15 | 2018-12-20 | General Electric Company | Combustion Section Heat Transfer System for a Propulsion System |
-
2016
- 2016-06-14 US US15/182,447 patent/US20180010800A1/en not_active Abandoned
Patent Citations (15)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6062018A (en) * | 1993-04-14 | 2000-05-16 | Adroit Systems, Inc. | Pulse detonation electrical power generation apparatus with water injection |
US6460342B1 (en) * | 1999-04-26 | 2002-10-08 | Advanced Research & Technology Institute | Wave rotor detonation engine |
US6845620B2 (en) * | 2001-07-06 | 2005-01-25 | Mohamed Razi Nalim | Rotary ejector enhanced pulsed detonation system and method |
US7100360B2 (en) * | 2002-12-30 | 2006-09-05 | United Technologies Corporation | Pulsed combustion engine |
US20060254252A1 (en) * | 2005-05-13 | 2006-11-16 | General Electric Company | Pulse detonation assembly and hybrid engine |
US20070180810A1 (en) * | 2006-02-03 | 2007-08-09 | General Electric Company | Pulse detonation combustor with folded flow path |
US20090102203A1 (en) * | 2007-10-23 | 2009-04-23 | Lu Frank K | System and method for power production using a hybrid helical detonation device |
US20090165438A1 (en) * | 2007-12-26 | 2009-07-02 | Occhipinti Anthony C | Pulse detonation engine |
US20090193786A1 (en) * | 2008-02-01 | 2009-08-06 | General Electric Company | System And Method Of Continuous Detonation In A Gas Turbine Engine |
US20090196733A1 (en) * | 2008-02-01 | 2009-08-06 | General Electric Company | Rotary Pressure Rise Combustor For A Gas Turbine Engine |
US8082728B2 (en) * | 2008-02-01 | 2011-12-27 | General Electric Company | System and method of continuous detonation in a gas turbine engine |
US20100242435A1 (en) * | 2009-03-30 | 2010-09-30 | Alliant Techsystems Inc. | Helical cross flow (hcf) pulse detonation engine |
US20150300630A1 (en) * | 2012-11-07 | 2015-10-22 | Exponential Technologies, Inc. | Pressure-gain combustion apparatus and method |
US20150184860A1 (en) * | 2013-12-27 | 2015-07-02 | United Technologies Corporation | Spiral pulse detonation tube configuration |
US20180363555A1 (en) * | 2017-06-15 | 2018-12-20 | General Electric Company | Combustion Section Heat Transfer System for a Propulsion System |
Cited By (5)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US11236908B2 (en) | 2018-10-24 | 2022-02-01 | General Electric Company | Fuel staging for rotating detonation combustor |
US11371711B2 (en) | 2018-11-28 | 2022-06-28 | General Electric Company | Rotating detonation combustor with offset inlet |
US11105511B2 (en) | 2018-12-14 | 2021-08-31 | General Electric Company | Rotating detonation propulsion system |
US11898757B2 (en) | 2018-12-14 | 2024-02-13 | General Electric Company | Rotating detonation propulsion system |
CN111207009A (en) * | 2019-12-26 | 2020-05-29 | 中国空气动力研究与发展中心 | Method for initiating oblique detonation wave in supersonic velocity airflow by using external instantaneous energy source |
Similar Documents
Publication | Publication Date | Title |
---|---|---|
RU2430307C2 (en) | Air-fuel mix injector, combustion chamber and gas turbine engine with said injector | |
US20210003285A1 (en) | Systems, Apparatuses And Methods For Improved Rotation Detonation Engines | |
US11674476B2 (en) | Multiple chamber rotating detonation combustor | |
US20190093553A1 (en) | Reverse-flow core gas turbine engine with a pulse detonation system | |
US10641169B2 (en) | Hybrid combustor assembly and method of operation | |
JP4555654B2 (en) | Two-stage pulse detonation system | |
CN109028142B (en) | Propulsion system and method of operating the same | |
US20180010800A1 (en) | Shock compression based supersonic combustor | |
US7980056B2 (en) | Methods and apparatus for controlling air flow within a pulse detonation engine | |
US20180231256A1 (en) | Rotating Detonation Combustor | |
US9027324B2 (en) | Engine and combustion system | |
Duvall et al. | Study of the effects of various injection geometries on the operation of a rotating detonation engine | |
US20180355792A1 (en) | Annular throats rotating detonation combustor | |
US20100242435A1 (en) | Helical cross flow (hcf) pulse detonation engine | |
US20220235727A1 (en) | Rotating detonation engine | |
US11131461B2 (en) | Effervescent atomizing structure and method of operation for rotating detonation propulsion system | |
US20210190320A1 (en) | Turbine engine assembly including a rotating detonation combustor | |
US11236908B2 (en) | Fuel staging for rotating detonation combustor | |
US7131260B2 (en) | Multiple detonation initiator for frequency multiplied pulsed detonation combustion | |
US20200332744A1 (en) | Liquid Combustion Concentric Injector and Ignitor | |
US11378040B2 (en) | Swirl preburner system and method | |
US20220252004A1 (en) | Radial pre-detonator | |
CN114777162A (en) | Continuous rotation knocking ramjet engine with radial oil supply and air supply | |
US10101033B2 (en) | Spiral pulse detonation tube configuration | |
CN114877377B (en) | Outer ring detonation combustor |
Legal Events
Date | Code | Title | Description |
---|---|---|---|
STPP | Information on status: patent application and granting procedure in general |
Free format text: DOCKETED NEW CASE - READY FOR EXAMINATION |
|
STPP | Information on status: patent application and granting procedure in general |
Free format text: NON FINAL ACTION MAILED |
|
STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |