BIMODAL COMBUSTION SYSTEM

A combustion system may include a detonation combustor comprising one or more detonation chamber walls defining a detonation chamber, a deflagration combustor comprising one or more deflagration chamber walls defining a deflagration chamber, and one or more conjugate chamber walls defining a conjugate chamber, with the conjugate chamber in fluid communication with the detonation chamber and the deflagration chamber. The detonation chamber includes a detonation region and a nozzle region, with the nozzle region providing fluid communication between the detonation region and the conjugate chamber.

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Description
FIELD

The present disclosure generally pertains to combustion systems for a turbine engine, as well as methods of operating a combustion system of a turbine engine.

BACKGROUND

Combustion systems that have an ability to operate over a wide range of operating conditions and thermal load requirements are of interest in the art, as are combustion systems that exhibit good operating performance, including good combustion efficiency, good fuel consumption, and/or low emissions. While combustion systems that perform deflagration continue to be an area of interest, the art has shown an increasing interest in detonation combustion processes. Accordingly, it would be welcomed in the art to provide combustion systems that offer improved performance and/or an ability to operate over a wider range of operating conditions and thermal load requirements.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended Figures, in which:

FIG. 1A shows a schematic cross-sectional view of an engine that includes a bimodal combustion system;

FIG. 1B shows a schematic cross-sectional view of an exemplary turbine engine that includes a bimodal combustion system;

FIGS. 2A and 2B respectively show a schematic cross-sectional view of an exemplary bimodal combustion system;

FIGS. 3A and 3B respectively show schematic cross-sectional views of a detonation chamber circumferentially surrounding a deflagration chamber from the bimodal combustion system shown in FIG. 2A;

FIG. 3C shows a schematic cross-sectional views of a conjugate chamber from the bimodal combustion system shown in FIG. 2A;

FIGS. 3D and 3E respectively show schematic cross-sectional views of a detonation chamber circumferentially surrounding a deflagration chamber from the bimodal combustion system shown in FIG. 2B;

FIG. 3F shows a schematic cross-sectional views of a conjugate chamber from the bimodal combustion system shown in FIG. 2B;

FIGS. 4A-4C respectively shows a schematic cross-sectional view of an exemplary detonation chamber including a detonation region and a nozzle region;

FIGS. 5A-5D respectively show a schematic three-dimensional view of a detonation region of an exemplary detonation chamber, a nozzle region of an exemplary detonation chamber, an exemplary deflagration chamber, and an exemplary conjugate chamber;

FIG. 6A shows a schematic cross-sectional view of an exemplary detonation fuel manifold;

FIGS. 6B and 6C show schematic perspective views of exemplary detonation fuel manifolds;

FIGS. 7A-7F show schematic cross-sectional views of further exemplary bimodal combustion systems;

FIG. 8 schematically depicts an exemplary detonation combustor with a detonation wave propagating through a detonation chamber of the detonation combustor; and

FIG. 9 shows a flow chart depicting an exemplary method of generating thrust.

 DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying figures. The present disclosure uses numerical and letter designations to refer to features in the figures. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure.

The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations. Additionally, unless specifically identified otherwise, all embodiments described herein should be considered exemplary.

As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

The terms “upper”, “lower”, “right”, “left”, “vertical”, “horizontal”, “top”, “bottom”, “lateral”, “longitudinal”, and so forth, shall relate to the disclosure as it is oriented in the drawing figures. However, it is to be understood that the disclosure may assume various alternative orientations, except where expressly specified to the contrary. It is also to be understood that the specific devices illustrated in the attached drawings, and described in the following specification, are simply exemplary embodiments of the disclosure. Hence, specific dimensions and other physical characteristics related to the embodiments disclosed herein are not to be considered as limiting.

The terms “forward” and “aft” refer to relative positions within a turbine engine, with forward referring to a position closer to an engine inlet and aft referring to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.

The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.

The terms “coupled,” “fixed,” “attached to,” and the like, refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein.

Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and/or systems. For example, the approximating language may refer to being within a 1, 2, 4, 10, 15, or 20 percent margin.

Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.

Additionally, the terms “low,” “high,” or their respective comparative degrees (e.g., lower, higher, where applicable) each refer to relative speeds within an engine, unless otherwise specified. For example, a “low-pressure turbine” operates at a pressure generally lower than a “high-pressure turbine.” Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low-pressure turbine” may refer to the lowest maximum pressure turbine within a turbine section, and a “high-pressure turbine” may refer to the highest maximum pressure turbine within the turbine section.

The term “turbomachine” refers to a machine that includes a combustor section and a turbine section with one or more turbines that together generate a thrust output and/or a torque output. In some embodiments, a turbomachine may include a compressor section with one or more compressors that compress air or gases flowing to the combustor section.

As used herein, the term “turbine engine” refers to an engine that may include a turbomachine as all or a portion of its power source. Example turbine engines include gas turbine engines, as well as hybrid-electric turbine engines, such as turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and the like.

One or more components of the engines described herein may be manufactured or formed using any suitable process, such as an additive manufacturing process, such as a 3-D printing process. The use of such a process may allow such component to be formed integrally, as a single monolithic component, or as any suitable number of sub-components. In particular, the additive manufacturing process may allow such component to be integrally formed and include a variety of features not possible when using prior manufacturing methods. For example, the additive manufacturing methods described herein may allow for the manufacture of passages, conduits, cavities, openings, casings, manifolds, double-walls, heat exchangers, or other components, or particular positionings and integrations of such components, having unique features, configurations, thicknesses, materials, densities, fluid passageways, headers, and mounting structures that may not have been possible or practical using prior manufacturing methods. Some of these features are described herein.

Suitable additive manufacturing technologies in accordance with the present disclosure include, for example, Selective Laser Melting (SLM), Direct Metal Laser Melting (DMLM), Fused Deposition Modeling (FDM), Selective Laser Sintering (SLS), 3D printing such as by inkjets, laser jets, and binder jets, Stereolithography (SLA), Direct Selective Laser Sintering (DSLS), Electron Beam Sintering (EBS), Electron Beam Melting (EBM), Laser Engineered Net Shaping (LENS), Laser Net Shape Manufacturing (LNSM), Direct Metal Deposition (DMD), Digital Light Processing (DLP), Direct Selective Laser Melting (DSLM), and other known processes.

Suitable powder materials for the manufacture of the structures provided herein as integral, unitary, structures include metallic alloy, polymer, or ceramic powders. Exemplary metallic powder materials are stainless-steel alloys, cobalt-chrome alloys, aluminum alloys, titanium alloys, nickel-based superalloys, and cobalt-based superalloys. In addition, suitable alloys may include those that have been engineered to have good oxidation resistance, known as “superalloys” which have acceptable strength at the elevated temperatures of operation in a turbine engine, e.g. Hastelloy, Inconel® alloys (e.g., IN 738, IN 792, IN 939), Rene alloys (e.g., Rene N4, Rene N5, Rene 80, Rene 142, Rene 195), Haynes alloys, Mar M, CM 247, CM 247 LC, C263, 718, X-850, ECY 768, 282, X45, PWA 1483 and CMSX (e.g. CMSX-4) single crystal alloys. The manufactured objects of the present disclosure may be formed with one or more selected crystalline microstructures, such as directionally solidified (“DS”) or single-crystal (“SX”).

As used herein, the terms “integral”, “unitary”, or “monolithic” as used to describe a structure refer to the structure being formed integrally of a continuous material or group of materials with no seams, connections joints, or the like. The integral, unitary structures described herein may be formed through additive manufacturing to have the described structure, or alternatively through a casting process, etc.

The present disclosure generally provides combustion systems that are configured to perform both deflagration combustion and detonation combustion, as well as engines that include such a combustion system. Exemplary engines that may be configured to perform both deflagration combustion and detonation combustion include turbine engines, rocket engines, ramjets, or a combination thereof, such as turbo-rocket engines, turbo-ramjets, or rocket-ramjets. Such combustion systems are generally referred to herein as bimodal combustion systems. The presently disclosed bimodal combustion systems may include a detonation section configured to perform detonation combustion, and a deflagration section configured to perform deflagration combustion. The detonation section includes a detonation chamber, and the deflagration section includes a deflagration chamber. The detonation chamber and the deflagration chamber are respectively in fluid communication with a conjugate chamber.

Combustion refers to the occurrence of exothermic chemical reactions between a fuel and an oxidant, producing combustion products and heat by conversion of chemical species. Heat and kinetic energy generated by combustion may be utilized by an engine to provide thrust. Generally, combustion may be performed in one or both of two modes: deflagration and detonation. As used herein, the term “deflagration” or “deflagration combustion” refers to combustion that can be described thermodynamically as approximately isobaric. During a deflagration combustion process, typically, the pressure of the combustion products drops slightly, and the specific volume of the combustion products increases significantly, generating a combustion wave that has a subsonic velocity. For example, a combustion wave generated by a deflagration combustion process may have a velocity on the order of several meters per second (m/s), such as from about 10 m/s to about 200 m/s. As used herein, the term “detonation” or “detonation combustion” refers to combustion that can be described thermodynamically as approximately isochoric. During a detonation combustion process, typically, the pressure and temperature of the combustion products increase abruptly, and the specific volume decreases slightly, generating a supersonic shock wave that closely precedes a combustion wave that also has a supersonic velocity. For example, a combustion wave generated by a detonation combustion process may have a velocity on the order of several kilometers per second (km/s), such as from about 1 km/s to about 6 km/s.

Detonation generally provides a faster heat release, a lower entropy increase, and a greater thermal efficiency, as compared to deflagration. Exemplary detonation combustion processes may provide a pressure increase on the order of a multiple of from about 5 to about 20. In further contrast with deflagration, detonation may propagate in a lean fuel mixture that results in relatively low NOx emissions. Detonation combustion has a higher thermodynamic efficiency than deflagration combustion, which translates to significantly improved specific impulse and/or specific fuel consumption. In some embodiments, a gas turbine engine that utilizes detonation combustion may have a reduced number of compressor stages and/or a reduced compressor pressure demands attributable, for example, to the ability for detonation combustion to provide a relatively large effective thrust at a relatively low overall compression ratio. Additionally, or in the alternative, detonation combustion may allow for engines with a higher thrust-to-weight ratio, which may allow for smaller, lighter-weight engines for given duty requirements.

In exemplary embodiments, the detonation section of the presently disclosed bimodal combustion systems may be configured to perform rotating detonation combustion. A rotating detonation combustion process may generate shock waves respectively preceding a combustion wave that propagates annularly through a detonation region of the detonation chamber. The annularly propagating shock waves and combustion waves may transition to longitudinal waves as combustion products travel through a nozzle region of the detonation chamber in fluid communication with the conjugate chamber. Deflagration taking place within the deflagration chamber and/or the conjugate chamber may provide back pressure that at least partially contributes to the initiation and/or stability of the detonation reaction in the detonation chamber. Additionally, or in the alternative, the nozzle region of the detonation chamber may provide back pressure that that at least partially contributes to the initiation and/or stability of the detonation reaction in the detonation chamber.

In some embodiments, the bimodal combustion system may perform deflagration combustion during operating conditions requiring relatively low thrust, and detonation combustion during operating conditions requiring relatively high thrust. The presently disclosed bimodal combustion systems may perform deflagration and detonation separately or concurrently. For example, a bimodal combustion system may initiate and sustain detonation combustion that coincides with an ongoing and sustained deflagration combustion. Additionally, or in the alternative, a bimodal combustion system may cease detonation combustion while sustaining deflagration combustion. Additionally, or in the alternative, a bimodal combustion system may transition from deflagration to detonation, and/or from detonation to deflagration according to changing operating requirements of the engine.

In some embodiments, the presently disclosed bimodal combustion systems may be configured to perform detonation when the engine is operating at a rated speed and/or when the engine is operating at a cruising speed. Additionally, or in the alternative, the presently disclosed bimodal combustion systems may be configured to perform deflagration when the engine is operating at a rated speed and/or when the engine is operating at a cruising speed. Additionally, or in the alternative, the presently disclosed bimodal combustion systems may be configured to initiate detonation, for example, when the engine transitions from a nominal operating state to a high-power operating and/or to a cruising operating state. Additionally, or in the alternative, the presently disclosed bimodal combustion systems may be configured to cease deflagration while sustaining detonation, for example, when the engine transitions from a low-power operating state to a nominal operating state, from a nominal operating state to a high-power operating and/or to a cruising operating state, and/or from a high-power operating state to a cruising operating state. Additionally, or in the alternative, the presently disclosed bimodal combustion systems may be configured to cease detonation while sustaining deflagration, for example, when the engine transitions from a nominal operating state to a low-power operating state, and/or from a high-power operating state or a cruising operating state to a nominal operating state.

As used herein, the term “rated speed” refers to a maximum output that an engine may achieve when operating properly. For a turbine engine or other rotary machine, a rated speed refers to a maximum rotational speed that the engine may achieve while operating properly. For an engine that does not include a rotary machine, such as a rocket engine, a rated speed refers to a velocity of thrust output by the engine. An engine, such as a turbine engine, utilized to provide thrust for an aircraft may operate at a rated speed during a high-power operating state, such as during takeoff operations and/or during aggressive aerial maneuvers.

As used herein, the “term nominal operating state” refers to operation of an engine, such as a turbine engine, at a speed that is greater than an idle speed and less than a rated speed for the engine. For example, nominal operating state may include an operating speed that is at least 10% greater than an idle speed and at least 10% less than the rated speed. As an example, a nominal operating state may include a cruising speed.

As used herein, the term “cruising speed” refers to operation of an output of an engine at a relatively high operational speed for a sustained period of time. For example, a turbine engine utilized to power an aircraft may operate at a cruising speed when the aircraft levels after climbing to a specified altitude. In some embodiments, an engine such as a turbine engine may operate at a cruising speed that is from about 50% to about 90% of the rated speed, such as from about 70% to about 80% of the rated speed. In some embodiments, a cruising speed may be achieved at about 80% of full throttle, such as from about 50% to about 90% of full throttle, such as from about 70% to about 80% full throttle.

As used herein, the term “low-power operating state” refers to operation of an engine, such as a turbine engine, at a speed that is at least less than 10% greater than an idle speed for the engine.

As used herein, the term “high-power operating state” refers to operation of an engine, such as a turbine engine, at a rotational speed that is at least 90% of a rated speed for the engine.

Exemplary embodiments of the present disclosure will now be described in further detail. Referring to FIGS. 1A and 1B, exemplary engines 50 that include a bimodal combustion system 200 will be described. The engine 50 has a radial axis (R) and a longitudinal axis (L). The engine 50 depicted in FIG. 1A may be any engine 50 that includes a bimodal combustion system 200, such as a turbine engine, a rocket engine, a ramjet, or a combination thereof, such as a turbo-rocket engine, a turbo-ramjet, or a rocket-ramjet. By way of example, FIG. 1B shows an exemplary turbine engine 100 that includes a bimodal combustion system 200. An exemplary engine 50, such as a turbine engine 100 may be mounted to an aircraft, such as in an under-wing configuration or tail-mounted configuration. The turbine engine 100 shown in FIG. 1B is provided by way of example and not to be limiting, and the subject matter of the present disclosure may be implemented with other suitable types of engines 50, including other suitable turbine engines 100.

As shown, for example, in FIG. 1A, an exemplary engine 50 may include, in serial flow relationship, an inlet section 52, a combustor section 54, and an outlet section 56. The engine 50 may include an engine case 58 that contains and/or defines at least a portion of the inlet section 52, the combustor section 54, and/or the outlet section 56. The inlet section 52 may generally guide a flow of oxidizer 60 such as air or gases to the combustor section 54. The inlet section 52 may compress the flow of oxidizer 60 prior to entering the combustor section 54. For example, the inlet section 52 may define a decreasing cross-sectional area leading downstream to the combustor section 54. At least a portion of the overall flow of oxidizer 60 may be mixed with a fuel 62 and may react in a combustion process to generate combustion products 64. The combustor section 54 may include a bimodal combustion system 200 configured according to the present disclosure. The bimodal combustion system 200 may include a detonation section 202 configured to perform detonation combustion, and a deflagration section 204 configured to perform deflagration combustion. Combustion products 64 from the combustor section 54 flow downstream to the outlet section 56. In some embodiments, the combustion products 64 may flow through a turbine section 66 prior to entering the outlet section 56. The turbine section 66 may include one or more turbine stages. In some embodiments, the turbine section 66 may include a high-pressure turbine and/or a low pressure turbine as described herein. The turbine section 66 may be disposed downstream of the combustor section 54. The turbine section 66 may be located between the combustor section 54 and the outlet section 56. The outlet section 56 may generally define an increasing cross sectional area leading downstream from the combustor section 54 and/or downstream from the turbine section 66. In some embodiments, the turbine section 66 may define a portion of the outlet section 56. Additionally, or in the alternative, the outlet section 56 may include an outlet nozzle 68, or the like. Expansion of the combustion products 64 generally provides thrust that may be utilized as a direct power output in the form of thrust, and/or to generate mechanical energy, for example, by rotation of the turbine section 66.

As shown in FIG. 1B, an engine 50 configured as a turbine engine 100 may include a fan section 102 and a core engine 104 disposed downstream from the fan section 102. The fan section 102 may include a fan 106 with any suitable configuration, such as a variable pitch, single stage configuration. The fan 106 may include a plurality of fan blades 108 coupled to a fan disk 110 in a spaced apart manner. The fan blades 108 may extend outwardly from the fan disk 110 generally along a radial direction. The core engine 104 may be coupled directly or indirectly to the fan section 102 to provide torque for driving the fan section 102.

The core engine 104 may include an engine case 58 that encases one or more portions of the core engine 104, including a compressor section 114, a combustor section 54, and a turbine section 66. The engine case 58 may define a core engine-inlet 118, an outlet nozzle 68, and a core air flowpath 122 therebetween. The core air flowpath 122 may pass through the compressor section 114, the combustor section 54, and the turbine section 66, in serial flow relationship. The compressor section 114 may include one or more compressors, such as a first, booster or low pressure (LP) compressor 124 and/or a second, high pressure (HP) compressor 126. The one or more compressors may respectively include one or more compressor stages. By way of example, the compressor section 114, including the LP compressor 124, and/or the HP compressor 126, may respectively have from 1 to 16 compressor stages, such as from 1 to 12 stages, such as from 1 to 10 stages, such as from 1 to 8 stages, such as from 1 to 6 stages, or such as from 1 to 4 stages. The turbine section 66 may include a first, high pressure (HP) turbine 128 and a second, low pressure (LP) turbine 130. The compressor section 114, combustor section 54, turbine section 66, and outlet nozzle 68 may be arranged in serial flow relationship and may respectively define a portion of the core air flowpath 122 through the core engine 104. In some embodiments, the inlet section 52 (FIG. 1A) may include at least a portion of the core engine-inlet 118 and/or at least a portion of the compressor section 114. In some embodiments, the outlet section 56 (FIG. 1A) may include at least a portion of the outlet nozzle 68 and/or at least a portion of the turbine section 66.

The core engine 104 and the fan section 102 may be coupled to a shaft driven by the core engine 104. By way of example, as shown in FIG. 1B, the core engine 104 may include a high pressure (HP) shaft 132 and a low pressure (LP) shaft 134. The HP shaft 132 may drivingly connect the HP turbine 128 to the HP compressor 126, and the LP shaft 134 may drivingly connect the LP turbine 130 to the LP compressor 124. In other embodiments, a turbine engine may have three shafts, such as in the case of a turbine engine that includes an intermediate pressure turbine. A shaft of the core engine 104, together with a rotating portion of the core engine 104, may sometimes be referred to as a “spool.” The HP shaft 132, a rotating portion of the HP compressor 126 coupled to the HP shaft 132, and a rotating portion of the HP turbine 128 coupled to the HP shaft 132, may be collectively referred to as a high pressure (HP) spool 136. The LP shaft 134, and a rotating portion of the LP compressor 124 coupled to the LP shaft 134, a rotating portion of the LP turbine 130 coupled to the LP shaft 134, may be collectively referred to as low pressure (LP) spool 138.

In some embodiments, the fan section may be coupled directly to a shaft of the core engine, such as directly to an LP shaft. Alternatively, as shown in FIG. 1B, the fan section 102 and the core engine 104 may be coupled to one another by way of a power gearbox 140, such as a planetary reduction gearbox, an epicyclical gearbox, or the like. For example, the power gearbox 140 may couple the LP shaft 134 to the fan 106, such as to the fan disk 110 of the fan section 102. The power gearbox 140 may include a plurality of gears for stepping down the rotational speed of the LP shaft 134 to a more efficient rotational speed for the fan section 102.

Still referring to FIG. 1B, the fan section 102 of the turbine engine 100 may include a fan case 142 that at least partially surrounds the fan 106 and/or the plurality of fan blades 108. The fan case 142 may be supported by the core engine 104, for example, by a plurality of outlet guide vanes 144 circumferentially spaced and extending substantially radially therebetween. The turbine engine 100 may include a nacelle 146. The nacelle 146 may be secured to the fan case 142. The nacelle 146 may include one or more sections that at least partially surround the fan section 102, the fan case 142, and/or the core engine 104. For example, the nacelle 146 may include a nose cowl, a fan cowl, an engine cowl, a thrust reverser, and so forth. The fan case 142 and/or an inward portion of the nacelle 146 may circumferentially surround an outer portion of the core engine 104. The fan case 142 and/or the inward portion of the nacelle 146 may define a bypass passage 148. The bypass passage 148 may be disposed annularly between an outer portion of the core engine 104 and the fan case 142 and/or inward portion of the nacelle 146 surrounding the outer portion of the core engine 104.

During operation of the turbine engine 100, an inlet airflow 150 enters the turbine engine 100 through an inlet 152 defined by the nacelle 146, such as a nose cowl of the nacelle 146. In some embodiments, the inlet section 52 (FIG. 1A) may include at least a portion of the inlet 152, at least a portion of the nacelle 146, and/or at least a portion of the fan case 142. The inlet airflow 150 passes across the fan blades 108. The inlet airflow 150 splits into a core airflow 154 that flows into and through the core air flowpath 122 of the core engine 104 and a bypass airflow 156 that flow through the bypass passage 148. The core airflow 154 is compressed by the compressor section 114. Pressurized air from the compressor section 114 flows downstream to the combustor section 54 where fuel is introduced to generate combustion gases 158. The combustion gases 158 exit the combustor section 54 and flows through the turbine section 66, generating torque and/or thrust that rotates the compressor section to support combustion while also rotating the fan section 102. Rotation of the fan section 102 causes the bypass airflow 156 to flow through the bypass passage 148, generating propulsive thrust. Additional thrust is generated by the core airflow exiting the outlet nozzle 68.

In some exemplary embodiments, the turbine engine 100 may be a relatively large power class turbine engine 100 that may generate a relatively large amount of thrust. For example, the turbine engine 100 may be configured to generate from about 300 kilonewtons (kN) of thrust to about 700 kN of thrust, for example, at a rated speed and/or at a cruising speed, such as from about 300 kN to about 500 kN of thrust, such as from about 500 kN to about 620 kN of thrust, or such as from about 620 kN to about 700 kN of thrust. In other embodiments, the turbine engine 100 may be configured to generate from about 10 kN of thrust to about 300 kN of thrust, such as from about 10 kN of thrust to about 50 kN of thrust, such as from about 50 kN of thrust to about 150 kN of thrust, such as from about 100 kN of thrust to about 300 kN of thrust, such as from about 100 kN of thrust to about 200 kN of thrust. However, the various features and attributes of the turbine engine 100 described with reference to FIG. 1B are provided by way of example only and not to be limiting. In fact, the present disclosure may be implemented with respect to any desired turbine engine, including those with attributes or features that differ in one or more respects from the turbine engine 100 described herein.

As schematically depicted in FIG. 1B, in some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 may be configured with a compressor section 114 that has a relatively lower number of compressor stages as compared to other turbine engines 100 that generate a comparable amount of thrust. Additionally, or in the alternative, a turbine engine 100 that includes a bimodal combustion system 200 may be configured with a compressor section 114 that has a relatively shorter axial length as compared to other turbine engines 100 that generate a comparable amount of thrust. In some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 may be configured without an HP compressor and/or without an LP compressor. As a result, a relatively larger proportion of the energy generated by the turbine engine 100 may be allocated to thrust, for example, as opposed to compressing incoming oxidizer utilized in the combustor section 54. For example, a relatively larger proportion of the energy generated by the turbine engine 100 may be allocated to thrust generated by rotating the fan 106 to move bypass airflow 156 through the bypass passage 148, and/or by exhausting core airflow through the outlet nozzle 68.

In some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 may have a compressor section 114 that includes from 1 to 12 stages, such as from 1 to 8 stages, such as from 1 to 6 stages, such as from 1 to 4 stages. Additionally, or in the alternative, the compressor section 114 of a turbine engine 100 that includes a bimodal combustion system 200 may include an LP compressor 124 that has less than 3 stages, such as less than 2 stages, or such as 1 stage. Additionally, or in the alternative, the compressor section 114 of a turbine engine 100 that includes a bimodal combustion system 200 may include an HP compressor 126 that has less than 8 stages, such as less than 4 stages, such as less than 3 stages, or such as 1 stage. Additionally, or in the alternative, the compressor section 114 of a turbine engine 100 that includes a bimodal combustion system 200 may be configured with a spool that does not include a compressor, such as an LP spool 138 that does not include a compressor, or such as an HP spool 136 that does not include a compressor. By way of example, a turbine engine 100 that generates from about 400 kN to about 600 kN of thrust at a rated speed, such as at takeoff, may have an aforementioned compressor section 114. As another example, a turbine engine 100 that generates from about 10 kN to about 300 kN of thrust at a rated speed, such as at takeoff, may have an aforementioned compressor section 114.

In some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 may exhibit an overall compressor ratio at a rated speed, such as at takeoff, of from about 10:1 to about 80:1, such as from about 10:1 to about 20:1, such as from about 20:1 to about 40:1, such as from about 40:1 to about 50:1, such as from about 50:1 to about 70:1, or such as from about 70:1 to about 80:1. By way of example, a turbine engine 100 that generates from about 400 kN to about 600 kN of thrust at a rated speed, such as at takeoff, may have an overall compressor ratio at the rated speed, such as at takeoff, of from about 10:1 to about 55:1, such as from about 20:1 to about 35:1, such as from about 30:1 to about 40:1, or such as from about 40:1 to about 55:1. As another example, a turbine engine 100 that generates from about 10 kN to about 300 kN of thrust at a rated speed, such as at takeoff, may have an overall compressor ratio at the rated speed, such as at takeoff, of from about 10:1 to about 35:1, such as from about 10:1 to about 35:1, such as from about 10:1 to about 20:1, such as from about 15:1 to about 30:1, or such as from about 25:1 to about 35:1.

In some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 may exhibit a bypass ratio of from about 10:1 to about 20:1, such as from about 10:1 to about 12:1, such as from about 12:1 to about 16:1, such as from about 16:1 to about 18:1, or such as from about 18:1 to about 20:1, for example, as determined at a rated speed and/or at a cruising speed. As used herein, the term “bypass ratio” refers to a ratio of the mass flow rate through the bypass passage 148 to the mass flow rate of the core air flowpath 122 of the core engine 104. In some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 and that exhibits an aforementioned bypass ratio may have an aforementioned overall compressor ratio at a rated speed, such as at takeoff. Additionally, or in the alternative, a turbine engine 100 that generates an aforementioned amount of thrust may exhibit an aforementioned bypass ratio.

In some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 may have a thrust to weight ratio of from about 6.0 to about 9.0, such as from about 6.0 to about 7.0, or such as from about 7.0 to about 8.0, or such as from about 8.0 to about 9.0. In some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 and that exhibits an aforementioned bypass ratio may have an aforementioned thrust to weight ratio. Additionally, or in the alternative, a turbine engine 100 that has an aforementioned overall compressor ratio may have an aforementioned thrust to weight ratio.

In some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 may have a thrust specific fuel consumption of from about 8 grams per kilonewton-second (g/kN·s) to about 14 g/kN·s, such as from about 8 g/kN·s to about 12 g/kN·s, such as from about 8 g/kN·s to about 10 g/kN·s, such as from about 10 g/kN·s to about 12 g/kN·s, or such as from about 12 g/kN·s to about 14 g/kN·s, for example, at a rated speed and/or at a cruising speed, such as at cruising speed at 80% of full throttle. In some embodiments, a turbine engine 100 that includes a bimodal combustion system 200 and that exhibits an aforementioned bypass ratio, an aforementioned overall compressor ratio, and/or an aforementioned thrust to weight ratio may have an aforementioned thrust specific fuel consumption. By way of example, a turbine engine 100 that generates from about 400 kN to about 600 kN of thrust and/or a turbine engine 100 that generates from about 10 kN to about 300 kN of thrust, such as at a rated speed or at cruising speed, may have an aforementioned thrust specific fuel consumption.

Now referring to FIGS. 2A and 2B, exemplary bimodal combustion systems 200 are further described. As shown, a bimodal combustion system 200 may include a detonation section 202, a deflagration section 204, and a conjugate section 205. The detonation section 202 may include a detonation combustor 206. The detonation combustor 206 may include one or more detonation chamber walls 208 defining a detonation chamber 210 within which detonation combustion may take place during operation of the detonation section 202 of the bimodal combustion system 200. The detonation combustion may include pulsed detonation combustion and/or rotating detonation combustion.

The detonation combustor 206 may include one or more detonation fuel manifolds 212 configured to supply fuel 62 and/or oxidizer 60 to the detonation chamber 210. The fuel 62 and oxidizer 60 may be mixed within the one or more detonation fuel manifolds 212. Additionally, or in the alternative, the fuel 62 and oxidizer 60 may be mixed upstream from the one or more detonation fuel manifolds 212 and/or within the detonation chamber 210. The one or more detonation fuel manifolds 212 may define a portion of a detonation chamber wall 208. Additionally, or in the alternative, a detonation fuel manifold 212 may be coupled to one or more detonation chamber walls 208. In some embodiments, a detonation fuel manifold 212 may be monolithically integrated with one or more detonation chamber walls 208. The one or more detonation chamber walls 208 and the detonation fuel manifold 212 may define a single monolithic component.

The detonation section 202 may include one or more detonation-fuel supply lines 214 in fluid communication with the detonation fuel manifold 212. The one or more detonation-fuel supply lines 214 may be coupled to the detonation fuel manifold 212. Additionally, or in the alternative, the one or more detonation-fuel supply lines 214 may be defined at least in part by the detonation fuel manifold 212, such as by a monolithic structure of the detonation fuel manifold 212. The detonation section 202 may include one or more detonation-oxidizer supply lines 216 in fluid communication with the detonation fuel manifold 212. The one or more detonation-oxidizer supply lines 216 may be coupled to the detonation fuel manifold 212. Additionally, or in the alternative, the one or more detonation-oxidizer supply lines 216 may be defined at least in part by the detonation fuel manifold 212, such as by a monolithic structure of the detonation fuel manifold 212.

The one or more detonation fuel manifolds 212 may include a plurality of detonation fuel orifices 218 providing fluid communication between the detonation-fuel supply lines 214 and the detonation chamber 210 and/or between the detonation-oxidizer supply lines 216 and the detonation chamber 210. The plurality of detonation fuel orifices 218 may respectively provide fuel 62 and/or oxidizer 60 to the detonation chamber 210. The plurality of detonation fuel orifices 218 may be defined at least in part by the respective detonation fuel manifold 212, such as by a monolithic structure of the respective detonation fuel manifold 212. Additionally, or in the alternative, the plurality of detonation fuel orifices 218 may be coupled to the respective detonation fuel manifold 212. One or more detonation-fuel supply lines 214 and/or one or more detonation-oxidizer supply lines 216 may fluidly communicate with a corresponding deflagration fuel manifold 226 and/or with a corresponding plurality of detonation fuel orifices 218. Fuel 62 and oxidizer 60 may mix with one another within the respective detonation fuel manifold 212, such as within the corresponding plurality of detonation fuel orifices 218. Additionally, or in the alternative, fuel 62 and oxidizer 60 may mix with one another upstream from the detonation fuel manifold 212 and/or within the detonation chamber 210.

The deflagration section 204 includes a deflagration combustor 220. The deflagration combustor 220 may include one or more deflagration chamber walls 222 defining a deflagration chamber 224 within which deflagration combustion may take place during operation of the deflagration section 204 of the bimodal combustion system 200. The deflagration combustor 220 may include one or more deflagration chambers 224. The deflagration combustor 220 may include a plurality of deflagration fuel manifolds 226 configured to supply fuel 62 and/or oxidizer 60 to the one or more deflagration chambers 224. In some embodiments, a plurality of deflagration fuel manifolds 226 may supply fuel 62 and/or oxidizer 60 to a generally annular deflagration chamber 224. Additionally, or in the alternative, the plurality of deflagration fuel manifolds 226 may respectively supply fuel 62 and/or oxidizer 60 to a plurality of generally cylindrical deflagration chambers 224. In some embodiments, a deflagration chamber 224 may include a plurality of generally cylindrical regions in fluid communication with a generally annular region. The fuel 62 and oxidizer 60 may be mixed within the respective deflagration fuel manifolds 226, upstream from the respective deflagration fuel manifolds 226, and/or within the deflagration chamber 224. The plurality of deflagration fuel manifolds 226 may respectively define a portion of a deflagration chamber wall 222. Additionally, or in the alternative, one or more deflagration chamber walls 222 may be coupled to a deflagration fuel manifold 226. In some embodiments, one or more deflagration chamber walls 222 may be monolithically integrated with a deflagration fuel manifold 226. For example, a deflagration chamber wall 222 and a deflagration fuel manifold 226 may define a single monolithic component.

The deflagration section 204 may include one or more deflagration-fuel supply lines 228 in fluid communication with respective ones of the plurality of deflagration fuel manifolds 226. One or more deflagration-fuel supply lines 228 may be coupled to a respective deflagration fuel manifold 226. Additionally, or in the alternative, one or more deflagration-fuel supply lines 228 may be defined at least in part by a respective deflagration fuel manifold 226, such as by a monolithic structure of the respective deflagration fuel manifold 226. The deflagration-fuel supply lines 228 may provide fuel 62 to the plurality of deflagration fuel manifolds 226. The deflagration section 204 may include one or more deflagration-oxidizer supply lines 230 in fluid communication with respective ones of the plurality of deflagration fuel manifolds 226. One or more deflagration-oxidizer supply lines 230 may be coupled to a respective deflagration fuel manifold 226. Additionally, or in the alternative, one or more deflagration-oxidizer supply lines 230 may be defined at least in part by a respective deflagration fuel manifold 226, such as by a monolithic structure of the respective deflagration fuel manifold 226. The deflagration-oxidizer supply lines 230 may provide oxidizer 60 to the plurality of deflagration fuel manifolds 226.

Respective ones of the plurality of deflagration fuel manifolds 226 may include one or more deflagration fuel injectors 232 providing fluid communication between the deflagration-fuel supply lines 228 and the deflagration chamber 224 and/or between the deflagration-oxidizer supply lines 230 and the deflagration chamber 224. The deflagration fuel injectors 232 may provide fuel 62 and/or oxidizer 60 to the deflagration chamber 224. The one or more deflagration fuel injectors 232 be coupled to the respective deflagration fuel manifold 226. Additionally, or in the alternative, the one or more deflagration fuel injectors 232 may be defined at least in part by the respective deflagration fuel manifold 226, such as by a monolithic structure of the respective deflagration fuel manifold 226. One or more deflagration-fuel supply lines 228 and/or one or more deflagration-oxidizer supply lines 230 may fluidly communicate with a corresponding deflagration fuel manifold 226 and/or with a corresponding deflagration fuel injector 232. Fuel 62 and oxidizer 60 may mix with one another within the plurality of deflagration fuel manifolds 226, such as within the respective deflagration fuel injectors 232. Additionally, or in the alternative, fuel 62 and oxidizer 60 may mix with one another upstream from the respective deflagration fuel manifolds 226 and/or within the deflagration chamber 224.

Still referring to FIGS. 2A and 2B, oxidizer from the inlet section 52 may flow to the combustor section 54 through a plurality of fluid pathways 234. The combustor section 54 may be separated from the inlet section 52 by an inlet shroud 236 and/or by an inward combustor casing 238. The inlet shroud 236 may delimit an upstream portion of the combustor section 54. The inward combustor casing 238 may delimit a radially inward portion of the combustor section 54. The plurality of fluid pathways 234 may include a plurality of diffusers 240. The plurality of diffusers 240 may be configured to decrease the velocity of the oxidizer 60 flowing from the inlet section 52 and/or to increase the static pressure of the oxidizer 60 flowing into the combustor section 54. The static pressure of the oxidizer 60 may be increased to at least a portion of a pressure level sufficient to support detonation in the detonation chamber 210. Additionally, or in the alternative, in some embodiments, the plurality of diffusers 240 may be configured to remove swirl from the oxidizer 60 flowing into the combustor section 54.

The plurality of fluid pathways 234 may provide fluid communication between the inlet section 52 and the one or more detonation fuel manifolds 212. In some embodiments, the combustor section 54 may include a combustor inlet plenum 242 in fluid communication with the plurality of fluid pathways 234 and the one or more detonation-oxidizer supply lines 216 corresponding to the respective detonation fuel manifold 212. Additionally, or in the alternative, the respective detonation-oxidizer supply lines 216 may fluidly communicate directly with the plurality of fluid pathways 234. In addition, or in the alternative to providing fluid communication between the inlet section 52 and the one or more detonation fuel manifolds 212, the plurality of fluid pathways 234 may provide fluid communication between the inlet section 52 and the plurality of deflagration fuel manifolds 226. In some embodiments, the combustor inlet plenum 242 may be fluid communication with the plurality of deflagration fuel manifolds 226. Additionally, or in the alternative, the plurality of fluid pathways 234 may include a plurality of deflagration-oxidizer supply lines 230 corresponding to the respective deflagration fuel manifolds 226.

In some embodiments, additional oxidizer 60 and/or additional fuel 62 may be introduced into the detonation chamber 210, such as by way of one or more detonation chamber-dilution pathways 207. The one or more detonation chamber-dilution pathways 207 may be defined at least in part by a corresponding detonation chamber wall 208. The one or more detonation chamber-dilution pathways 207 may also be configured to provide backpressure to the detonation chamber 210 and/or to augment an equivalence ratio within the detonation chamber 210. Additionally, or in the alternative, in some embodiments, additional oxidizer 60 and/or additional fuel 62 may be introduced into the deflagration chamber 224, such as by way of one or more deflagration chamber-dilution pathways 209. The one or more deflagration chamber-dilution pathways 209 may be defined at least in part by a corresponding deflagration chamber wall 222. The one or more deflagration chamber-dilution pathways 209 may also be configured to provide backpressure to the deflagration chamber 224 and/or to augment an equivalence ratio within the deflagration chamber 224.

As shown in FIGS. 2A and 2B, the bimodal combustion system 200 may include a conjugate section 205. The conjugate section 205 may include a conjugate combustor 211. The conjugate section 205 and/or the conjugate combustor 211 may include one or more conjugate chamber walls 244 defining a conjugate chamber 246. The detonation chamber 210 and the deflagration chamber may respectively fluidly communicate with the conjugate chamber 246. The detonation chamber 210 and the deflagration chamber 224 may respectively have a generally annular configuration. The detonation chamber 210 and the deflagration chamber 224 may have a co-annular annularly disposition with respect to a longitudinal axis 248 of the engine 50. The conjugate chamber 246 may be disposed annularly with respect to the longitudinal axis 248 of the engine 50. Combustion products 64 from the detonation chamber 210 and/or the deflagration chamber 224 may flow to the conjugate chamber 246.

Combustion products 64 generated in and/or flowing from the detonation chamber 210 may sometimes be referred to as detonation combustion products 64. Combustion products 64 generated in and/or flowing from the deflagration chamber 224 may sometimes be referred to as deflagration combustion products 64. The detonation combustion products 64 and the deflagration combustion products 64 may mix with one another in the conjugate chamber 246. In some embodiments, further combustion may take place within the conjugate chamber 246, such as deflagration and/or detonation. Combustion products 64 generated in and/or flowing from the conjugate chamber 246 may sometimes be referred to as conjugate combustion products 64. Additionally, or in the alternative, combustion may be substantially completed within the detonation chamber 210 and/or the deflagration chamber 224, respectively. In some embodiments, additional oxidizer 60 and/or additional fuel 62 may be introduced into the conjugate chamber 246, such as by way of one or more conjugate chamber-dilution pathways 213. The one or more conjugate chamber-dilution pathways 213 may be defined at least in part by a corresponding conjugate chamber wall 244. The one or more conjugate chamber-dilution pathways 213 may also be configured to provide backpressure to the conjugate chamber 246, the detonation chamber 210, and/or the deflagration chamber 224, and/or to augment an equivalence ratio within the conjugate chamber 246, the detonation chamber 210, and/or the deflagration chamber 224.

In some embodiments, as shown in FIG. 2A, the detonation chamber 210 may have a radially-outward disposition relative to the deflagration chamber 224. The deflagration chamber 224 may have a radially-inward disposition relative to the detonation chamber 210. FIG. 3A shows a cross-sectional view of the bimodal combustion system 200 of FIG. 2A at a position along the longitudinal axis 248 denoted with an “A”. FIG. 3B shows a cross-sectional view of the bimodal combustion system 200 of FIG. 2A at a location along the longitudinal axis 248 denoted with a “B”. As shown in FIG. 2A and FIGS. 3A and 3B, at least a portion of the detonation chamber 210 may circumferentially surround at least a portion of the deflagration chamber 224.

In some embodiments, as shown in FIG. 2B, the deflagration chamber 224 may have a radially-outward disposition relative to the detonation chamber 210. The detonation chamber 210 may have a radially-inward disposition relative to the deflagration chamber 224. FIG. 3D shows a cross-sectional view of the bimodal combustion system 200 of FIG. 2B at a position along the longitudinal axis 248 denoted with a “D”. FIG. 3E shows a cross-sectional view of the bimodal combustion system 200 of FIG. 2B at a location along the longitudinal axis 248 denoted with a “E”. As shown in FIG. 2B and FIGS. 3D and 3E, at least a portion of the deflagration chamber 224 may circumferentially surround at least a portion of the detonation chamber 210.

As shown in FIGS. 2A and 2B, and FIGS. 3A, 3B, 3D, and 3E, the one or more detonation chamber walls 208 defining the detonation chamber 210 may include an outer detonation chamber wall 250 and an inner detonation chamber wall 252. At least a portion of the outer detonation chamber wall 250 may circumferentially surround at least a portion of the inner detonation chamber wall 252. The one or more deflagration chamber walls 222 defining the deflagration chamber 224 may include an outer deflagration chamber wall 254 and an inner deflagration chamber wall 256. At least a portion of the outer deflagration chamber wall 254 may circumferentially surround at least a portion of the inner deflagration chamber wall 256. In some embodiments, the bimodal combustion system 200 may include a combustor inlet plenum 242 disposed between at least a portion of the detonation chamber 210 and at least a portion of the deflagration chamber 224 with respect to a radial axis 258 of the engine 50.

In some embodiments, the outer detonation chamber wall 250 and the inner detonation chamber wall 252 may be monolithically integrated with one another. Alternatively, the outer detonation chamber wall 250 and the inner detonation chamber wall 252 may be coupled to one another, such as by way of welding, attachment hardware (e.g., bolts), or the like. In some embodiments, as shown in FIGS. 2A and 2B, the one or more detonation chamber walls 208 may include a posterior detonation chamber wall 260. The posterior detonation chamber wall 260 may define a posterior region of the detonation chamber 210. The posterior detonation chamber wall 260 may have a generally annular configuration. For example, the posterior detonation chamber wall 260 may be disposed annularly between the outer detonation chamber wall 250 and the inner detonation chamber wall 252. The posterior detonation chamber wall 260 may be monolithically integrated with the outer detonation chamber wall 250 and/or with the inner detonation chamber wall 252. Additionally, or in the alternative, the posterior detonation chamber wall 260 may be coupled to the outer detonation chamber wall 250 and/or to the inner detonation chamber wall 252, such as by way of welding, attachment hardware (e.g., bolts), or the like.

In some embodiments, the outer deflagration chamber wall 254 and the inner deflagration chamber wall 256 may be monolithically integrated with one another. Alternatively, the outer deflagration chamber wall 254 and the inner deflagration chamber wall 256 may be coupled to one another, such as by way of welding, attachment hardware (e.g., bolts), or the like. In some embodiments, as shown in FIGS. 2A and 2B, the one or more deflagration chamber walls 222 may include a posterior deflagration chamber wall 262. The posterior deflagration chamber wall 262 may define a posterior region of the deflagration chamber 224. The posterior deflagration chamber wall 262 may have a generally annular configuration. For example, the posterior deflagration chamber wall 262 may be disposed annularly between the outer deflagration chamber wall 254 and the inner deflagration chamber wall 256. The posterior deflagration chamber wall 262 may be monolithically integrated with the outer deflagration chamber wall 254 and/or with the inner deflagration chamber wall 256. Additionally, or in the alternative, the posterior deflagration chamber wall 262 may be coupled to the outer deflagration chamber wall 254 and/or to the inner deflagration chamber wall 256, such as by way of welding, attachment hardware (e.g., bolts), or the like.

In some embodiments, the inner detonation chamber wall 252 may adjoin and/or abut the outer deflagration chamber wall 254. In some embodiments, the inner detonation chamber wall 252 and the outer deflagration chamber wall 254 may be monolithically integrated with one another. Alternatively, the inner detonation chamber wall 252 and the outer deflagration chamber wall 254 may be coupled to one another, such as by way of welding, attachment hardware (e.g., bolts), or the like. In some embodiments, at least a portion of the deflagration chamber 224 may circumferentially surround at least a portion of the combustor inlet plenum 242. Additionally, or in the alternative, at least a portion of the deflagration chamber 224 may have a generally cylindrical configuration.

In some embodiments, the inner deflagration chamber wall 256 may adjoin and/or abut the outer detonation chamber wall 250. In some embodiments, the inner deflagration chamber wall 256 and the outer detonation chamber wall 250 may be monolithically integrated with one another. Alternatively, the inner deflagration chamber wall 256 and the outer detonation chamber wall 250 may be coupled to one another, such as by way of welding, attachment hardware (e.g., bolts), or the like. In some embodiments, at least a portion of the detonation chamber 210 may circumferentially surround a portion of the combustor inlet plenum 242. Additionally, or in the alternative, at least a portion of the detonation chamber 210 may have a generally cylindrical configuration.

Still referring to FIGS. 2A and 2B, and with further reference to FIGS. 3A-3F, in some embodiments, the detonation chamber 210 may include a detonation region 264 and a nozzle region 266. The nozzle region 266 may be disposed between the detonation region 264 and the conjugate chamber 246. The nozzle region 266 may provide fluid communication between the detonation region 264 and the conjugate chamber 246. Detonation of fuel 62 and/or oxidizer 60, such as a mixture of fuel 62 and oxidizer 60 may occur within the detonation region 264 of the detonation chamber 210. Combustion products 64 resulting from combustion of fuel 62 and oxidizer 60 may flow through the nozzle region 266 of the detonation chamber 210 to the conjugate chamber 246. Combustion products 64 flowing through the nozzle region 266 to the conjugate chamber 246 may continue through the conjugate chamber 246 to the outlet section 56 of the engine 50 and/or through the turbine section 66, and subsequently through the outlet nozzle 68 of the engine.

The detonation region 264 and the nozzle region 266 of the detonation chamber 210 may respectively have a generally annular configuration, for example, with a cross-sectional area in the shape of an annulus. The detonation region 264 of the detonation chamber 210 may include a region of the detonation chamber 210 within which detonation occurs during operation of the detonation section 202 of the bimodal combustion system 200. The detonation reaction may be stable within the detonation region 264 during operation of the detonation section 202 within an operating range for which the detonation section 202 may be configured. The detonation section 202 may include a detonation nozzle 268. The detonation nozzle 268 may be defined by one or more detonation chamber walls 208. The detonation nozzle 268 may be located at the nozzle region 266 of the detonation chamber 210. The detonation nozzle 268 may include a detonation throat 270. The detonation throat 270 may define a location of the detonation nozzle 268 that has an annular cross-sectional area with a minimum annular ring width relative to an adjacent portion of the detonation nozzle 268. The detonation nozzle 268 may include a convergent portion that has a decreasing cross-sectional area upstream from the detonation throat 270 in a direction from the detonation region 264 to the detonation throat 270 of the detonation nozzle 268. The detonation nozzle 268 may include a divergent portion that has an increasing cross-sectional area downstream from the detonation throat 270 in a direction from the detonation throat 270 towards the conjugate chamber 246.

The deflagration chamber 224 may be in fluid communication with the conjugate chamber 246 and the detonation chamber 210. The deflagration chamber 224 and the detonation chamber 210 may respectively transition to the conjugate chamber 246 along the longitudinal axis 248 of the engine 50. The transition from the deflagration chamber 224 to the conjugate chamber 246 may be at a location of the longitudinal axis 248 that is upstream, downstream, or equidistant from a location of the longitudinal axis 248 corresponding to the transition from the detonation chamber 210 to the conjugate chamber 246.

The deflagration chamber 224 and the detonation chamber 210 may be delineated from one another at least in part by a conjugate inflection line 272. Additionally, or in the alternative, a detonation chamber wall 208 and a deflagration chamber wall 222 may be delineated from one another at least in part by the conjugate inflection line 272. The conjugate inflection line 272 may define a linear inflection oriented circumferentially with respect to the longitudinal axis 248 of the engine 50 representing a forwardmost oblique angle, or a tangent to a forwardmost curve, delineating a detonation chamber wall 208 and a deflagration chamber wall 222 from one another. For example, as shown in FIG. 2A, the conjugate inflection line 272 may define a linear inflection between the inner detonation chamber wall 252 and the outer deflagration chamber wall 254. As shown in FIG. 2B, the conjugate inflection line 272 may define a linear inflection between the outer detonation chamber wall 250 and the inner deflagration chamber wall 256.

In some embodiments, the detonation section 202 may include a conjugate chamber wall 244 disposed between a detonation chamber wall 208 and a deflagration chamber wall 222. The detonation chamber wall 208 and/or the deflagration chamber wall 222 may be monolithically integrated with the conjugate chamber wall 244 disposed therebetween. Additionally, or in the alternative, the detonation chamber wall 208 and/or the deflagration chamber wall 222 may be coupled to the conjugate chamber wall 244 disposed therebetween, such as by welding, attachment hardware (e.g., bolts), or the like. In some embodiments, the conjugate inflection line 272 may be defined by a conjugate chamber wall 244 disposed between such a detonation chamber wall 208 and deflagration chamber wall 222. For example, the conjugate inflection line 272 may define a linear inflection oriented circumferentially with respect to the longitudinal axis 248 of the engine 50 representing a forwardmost oblique angle of the conjugate chamber wall 244 or a tangent to a forwardmost curve of the conjugate chamber wall 244. Alternatively, the conjugate inflection line 272 may define a linear inflection oriented circumferentially with respect to the longitudinal axis 248 of the engine 50 representing an aftmost oblique angle of the conjugate chamber wall 244 or a tangent to an aftmost curve of the conjugate chamber wall 244.

The deflagration chamber 224 and the detonation chamber 210 may be located at respectively opposite sides of the conjugate inflection line 272. The conjugate inflection line 272 may have a generally elliptical or circular shape. The conjugate inflection line 272 may circumferentially surround the longitudinal axis 248 of the engine 50. As shown, for example, in FIGS. 3C and 3F, the conjugate chamber 246 may have a conjugate chamber-center line 274 that defines an annular center line for the conjugate chamber 246. The conjugate chamber-center line 274 may be determined with respect to a volume of the conjugate chamber 246 located between the conjugate inflection line 272 and a downstream end of the conjugate chamber 246. The conjugate chamber-center line 274 may define a volumetric center of the conjugate chamber 246. The conjugate chamber-center line 274 may have a generally elliptical or circular shape. The conjugate chamber-center line 274 may circumferentially surround the longitudinal axis 248 of the engine 50.

A conjugate chamber plane 276, denoted “C”, may intersect the conjugate inflection line 272 and the conjugate chamber-center line 274 of the conjugate chamber 246. The conjugate chamber plane 276 may circumferentially surround the longitudinal axis 248 of the engine 50. The conjugate chamber plane 276 may have a generally linear configuration along a conjugate chamber midline 278 intersecting the conjugate inflection line 272 and the conjugate chamber-center line 274 at an orientation parallel to the conjugate chamber plane 276. By way of example, the conjugate chamber plane 276 may have a cylindrical configuration or a frustoconical configuration.

A downstream end of the conjugate chamber 246 may be determined by the structural context of the engine 50. In some embodiments, a downstream end of the conjugate chamber 246 may be defined by a shroud (not shown), such as a combustor discharge shroud and/or a turbine inlet shroud. Such a shroud may direct a flow of combustion products 64 circumferentially and/or helically into a turbine section 66 (FIGS. 1A and 1B) of the engine 50. Alternatively, such a shroud may direct a flow of combustion products 64 circumferentially and/or helically into the outlet section 56 (FIGS. 1A and 1B) of the engine 50. Additionally, or in the alternative, in some embodiments, a downstream end of the conjugate chamber 246 may be defined by a location of the conjugate chamber 246 that has an annular cross-sectional area with a minimum annular ring width relative to an adjacent upstream portion of the conjugate chamber 246.

In some embodiments, the conjugate section 205 may include a conjugate nozzle 280. The conjugate nozzle 280 may be defined by one or more conjugate chamber walls 244. In some embodiments, the detonation chamber 210 may be delineated from the conjugate chamber 246 at least in part by a conjugate nozzle 280. The conjugate nozzle 280 may include a conjugate throat 282. At least a portion of the conjugate nozzle 280 and/or the conjugate throat 282 may be adjacent to the detonation chamber 210. The conjugate throat 282 may define a location of the conjugate nozzle 280 that has an annular cross-sectional area normal to the conjugate chamber plane 276 with a minimum annular ring width relative to an adjacent portion of the conjugate nozzle 280. The annular cross-sectional area of the conjugate nozzle 280 corresponding to the conjugate throat 282 may extend from the conjugate chamber plane 276 to a conjugate chamber wall 244 on the side of the conjugate chamber plane 276 radially corresponding to the detonation chamber 210. For example, FIG. 2A shows a conjugate nozzle 280 and a conjugate throat 282 located at a radially outward side of the conjugate chamber plane 276, and FIG. 2B shows a conjugate nozzle 280 and a conjugate throat 282 located at a radially inward side of the conjugate chamber plane 276. Such conjugate chamber wall 244 delineating the annular cross-sectional area corresponding to the conjugate throat 282 may be adjoining, abutting, and/or monolithically integrated with the conjugate nozzle 280. Such conjugate chamber wall 244 delineating the annular cross-sectional area corresponding to the conjugate throat 282 may be longitudinally adjacent to a detonation chamber wall 208, for example, adjoining, abutting, and/or monolithically integrated with such detonation chamber wall 208. Additionally, or in the alternative, the conjugate throat 282 may delineate the detonation chamber wall 208 from the conjugate chamber wall 244. The conjugate nozzle 280 may include a convergent portion that has a decreasing cross-sectional area upstream from the conjugate throat 282 in a direction from the detonation throat 270 of the detonation nozzle 268 towards the conjugate throat 282. In some embodiments, the divergent portion of the detonation nozzle 268 may define at least a portion of the convergent portion of the conjugate nozzle 280. The conjugate nozzle 280 may include a divergent portion that has an increasing cross-sectional area downstream from the conjugate throat 282 in a direction from the conjugate throat 282 towards the downstream end of the conjugate chamber 246.

In some embodiments, the deflagration section 204 may include a deflagration nozzle 284. The deflagration nozzle 284 may be defined by one or more deflagration chamber walls 222. Additionally, or in the alternative, the deflagration chamber 224 may be delineated from the conjugate chamber 246 at least in part by a deflagration nozzle 284. The deflagration nozzle 284 may include a deflagration throat 286. At least a portion of the deflagration nozzle 284 and/or the deflagration throat 286 may be adjacent to the deflagration chamber 224. The deflagration throat 286 may define a location of the deflagration nozzle 284 that has an annular cross-sectional area normal to the conjugate chamber plane 276 with a minimum annular ring width relative to an adjacent portion of the deflagration nozzle 284. The annular cross-sectional area corresponding to the deflagration throat 286 may extend from the conjugate chamber plane 276 to a conjugate chamber wall 244 on the side of the conjugate chamber plane 276 radially corresponding to the deflagration chamber 224. For example, FIG. 2A shows a deflagration nozzle 284 and a deflagration throat 286 located at a radially inward side of the conjugate chamber plane 276, and FIG. 2B shows a deflagration nozzle 284 and a deflagration throat 286 located at a radially outward side of the conjugate chamber plane 276. Such conjugate chamber wall 244 delineating the annular cross-sectional area corresponding to the deflagration throat 286 may be adjoining, abutting, and/or monolithically integrated with the deflagration nozzle 284. Such conjugate chamber wall 244 delineating the annular cross-sectional area corresponding to the deflagration throat 286 may be longitudinally adjacent to a deflagration chamber wall 222, for example, adjoining, abutting, and/or monolithically integrated with such deflagration chamber wall 222. Additionally, or in the alternative, the deflagration throat 286 may delineate a deflagration chamber wall 222 from a conjugate chamber wall 244. The deflagration nozzle 284 may include a convergent portion that has a decreasing cross-sectional area upstream from the deflagration throat 286 in a direction from the deflagration chamber 224 towards the deflagration throat 286. The deflagration nozzle 284 may include a divergent portion that has an increasing cross-sectional area downstream from the deflagration throat 286 in a direction from the deflagration throat 286 towards the downstream end of the conjugate chamber 246.

Still referring to FIGS. 2A and 2B, and FIGS. 3A-3F, in some embodiments, the detonation nozzle 268 and/or the detonation throat 270 may be configured and arranged at least in part to provide suitable back pressure in the detonation chamber 210 to initiate and/or sustain detonation, such as in the detonation region 264 of the detonation chamber 210. For example, stable detonation may occur within the detonation chamber 210, such as in the detonation region 264 of the detonation chamber 210, at least during operation of the detonation section 202 within an operating range for which the detonation section 202 may be configured. Additionally, or in the alternative, the detonation nozzle 268 and/or the detonation throat 270 may be configured and arranged at least in part to accelerate combustion products 64 flowing from the detonation chamber 210 to the conjugate chamber 246. In some embodiments, the conjugate nozzle 280 and/or the conjugate throat 282 may be configured at least in part to provide suitable back pressure in the detonation chamber 210 to initiate and/or sustain detonation, such as in the detonation region 264 of the detonation chamber 210, for example, at least during operation of the detonation section 202 within an operating range for which the detonation section 202 may be configured. Additionally, or in the alternative, the conjugate nozzle 280 and/or the conjugate throat 282 may be configured and arranged at least in part to accelerate combustion products 64 flowing from the detonation chamber 210 to the conjugate chamber 246. In some embodiments, the deflagration nozzle 284 and/or the deflagration throat 286 may be configured at least in part to provide suitable back pressure in the detonation chamber 210 to initiate and/or sustain detonation, such as in the detonation region 264 of the detonation chamber 210, for example, at least during operation of the detonation section 202 within an operating range for which the detonation section 202 may be configured. In some embodiments, deflagration within the deflagration chamber 224 may contribute back pressure within the detonation chamber 210 suitable to initiate and/or sustain detonation within the detonation chamber 210, such as in the detonation region 264 of the detonation chamber 210, for example, at least during operation of the detonation section 202 within an operating range for which the detonation section 202 may be configured. Additionally, or in the alternative, the deflagration nozzle 284 and/or the deflagration throat 286 may be configured and arranged at least in part to accelerate combustion products 64 flowing from the deflagration chamber 224 to the conjugate chamber 246.

In some embodiments, the detonation nozzle 268 and/or the conjugate nozzle 280 may be configured and arranged in the form of a de Laval type nozzle. For example, the detonation nozzle 268 and/or the conjugate nozzle 280 may individually or collectively configured as a de Laval type nozzle. In some embodiments, the detonation nozzle 268 and/or the conjugate nozzle 280 may be configured and arranged at least in part to decrease the static pressure of the combustion products 64 flowing from the detonation chamber 210 to the conjugate chamber 246. The static pressure of the conjugate chamber 246 and/or the static pressure of the deflagration chamber 224 may be less than the static pressure of the detonation chamber 210. The static pressure of the combustion products 64 flowing from the detonation chamber 210 to the conjugate chamber 246 may be decreased by way of expansion of the combustion products 64 caused by the portion of the detonation nozzle 268 downstream from the detonation throat 270, and/or by the portion of the conjugate nozzle 280 downstream of the conjugate throat 282.

As the combustion products 64 travel through the detonation nozzle 268 and/or the conjugate nozzle 280, the velocity of the combustion products 64 increases while the pressure decreases. The velocity of the combustion products 64 traveling from the detonation chamber 210 to the conjugate chamber 246 may be supersonic in nature. By way of example, downstream of the detonation nozzle 268, the detonation throat 270, and/or the conjugate throat 282, the detonation combustion products 64 may have a velocity of from about 1,000 meters per second (m/s) to about 5,000 m/s, such as from about 1,000 m/s to about 3,000 m/s, such as from about 2,000 m/s to about 3,500 m/s, such as from about 2,500 m/s to about 4,500 m/s, or such as from about 3,000 m/s to about 5,000 m/s.

As the combustion products 64 flow past the detonation throat 270, the static pressure of the combustion products 64 is generally higher than the static pressure within the conjugate chamber 246. The static pressure of the combustion products 64 is lowered by expansion as the combustion products 64 flow through the divergent portion of the detonation nozzle 268. The efficiency with which the kinetic energy of the combustion products 64 flowing through the detonation nozzle 268 are converted to axial momentum may depend at least in part on a ratio of the cross-sectional area of the detonation throat 270 to the cross-sectional area of the conjugate throat 282. Additionally, or in the alternative, efficiency with which the kinetic energy of the combustion products 64 flowing through the detonation nozzle 268 are converted to axial momentum may depend at least in part on the cone-half angle of the divergent portion of the detonation nozzle 268 determined from a plane normal to the detonation throat 270.

In some embodiments, a minimum cross-sectional area of the detonation nozzle 268, such as a cross-sectional area defined by the detonation throat 270, may be less than a minimum cross-sectional area of the conjugate nozzle 280, such as a cross-sectional area defined by the conjugate throat 282. By way of example, the cross-sectional area of the detonation throat 270 may be at least 1% less than the cross-sectional area of the conjugate throat 282, such as from about 1% to about 90%, such as from about 5% to about 30%, such as from about 5% to about 20%, such as from about 15% to about 30%, such as from about 30% to about 60%, or such as from about 60% to about 90% less than the cross-sectional area of the conjugate throat 282.

In some embodiments, a cross-sectional area of the conjugate throat 282 and/or the detonation throat 270 may respectively be less than a cross-sectional area of the deflagration throat 286. By way of example, the cross-sectional area of the conjugate throat 282 and/or the detonation throat 270 may respectively be at least 1% less than the cross-sectional area of the deflagration throat 286, such as from about 1% to about 90%, such as from about 10% to about 60%, such as from about 20% to about 40%, such as from about 30% to about 60%, or such as from about 60% to about 90% less than the cross-sectional area of the deflagration throat 286.

A pressure drop of combustion products 64 from the detonation chamber 210 to the conjugate chamber 246 may be greater than or equal to a pressure drop of combustion products 64 from the deflagration chamber 224 to the conjugate chamber 246, and the static pressure of the combustion products 64 in the conjugate chamber 246 resulting, for example, from the combined flow of combustion products 64 from the detonation chamber 210 and the deflagration chamber 224 will generally equalize. The combustion products 64 may exhibit a pressure drop across at least a portion of the conjugate chamber 246, and/or that local pressure gradients may exist within the conjugate chamber 246, such as pressure gradients attributable to fluid currents, turbulence, mixing, velocity profiles, contraction and/or expansion in cross-sectional area, introduction of dilution air, and the like, as well as combinations of these. Additionally, a downstream location will have a lower pressure than an upstream location.

Still referring to FIGS. 2A and 2B, and FIGS. 3A-3F, and with reference to FIGS. 4A-4C, exemplary detonation sections 202 of the bimodal combustion system 200 are further described. As shown, the detonation section 202 may include a detonation chamber 210 that has a detonation throat-center line 400 defining an annular center of the detonation throat 270, as seen in FIGS. 4A-4C. The detonation throat-center line 400 may be located equidistantly between opposite sides of the detonation throat 270, such as equidistantly between the outer detonation chamber wall 250 and the inner detonation chamber wall 252 at the detonation throat 270. The detonation throat-center line 400 may have a generally elliptical or circular shape. The detonation throat-center line 400 may circumferentially surround the longitudinal axis 248 of the engine 50. A detonation chamber plane 402 may intersect the detonation throat-center line 400 tangentially normal to the detonation throat-center line 400. The tangentially normal orientation of the detonation chamber plane 402 may include a perpendicular orientation relative to an annular plane spanning the detonation throat 270. The detonation chamber plane 402 may circumferentially surround the longitudinal axis 248 of the engine 50. The detonation chamber plane 402 may have a generally linear configuration along a detonation chamber midline 404 intersecting the conjugate chamber midline 278 and the detonation throat-center line 400 at an orientation parallel to the detonation chamber plane 402. By way of example, the detonation chamber plane 402 may have a cylindrical configuration or a frustoconical configuration.

As shown in FIGS. 4A and 4B, the detonation nozzle 268 may include a cone-half angle 406 determined with reference to the detonation chamber plane 402. Additionally, or in the alternative, as shown in FIG. 4C, the conjugate nozzle 280 may include a cone-half angle 406 determined with reference to the detonation chamber plane 402. The term “cone-half angle,” as used herein in relation to the detonation nozzle 268 refers to an angle of the detonation nozzle 268 determined from the detonation chamber plane 402. As used herein in relation to the conjugate nozzle 280, the term “cone-half angle” refers to an angle of the conjugate nozzle 280 determined from the detonation chamber plane 402. The cone-half angle 406 of the detonation nozzle 268 and/or the cone-half angle 406 of the conjugate nozzle 280 may be respectively configured to provide suitable combustion conditions within the combustor section 54, such as suitable back pressure for stable detonation and/or a pressure drop suitable for stable flow of detonation combustion products 64 from the detonation chamber 210 to the conjugate chamber 246. Additionally, or in the alternative, the efficiency with which the kinetic energy of the combustion products 64 flowing through the detonation nozzle 268 and/or the conjugate nozzle 280 are converted to axial momentum may depend at least in part on the respective cone-half angle 406. As shown in FIGS. 4A and 4B, the cone-half angle 406 of the detonation nozzle 268 may differ as between a convergent portion 408 of the detonation nozzle 268 and a divergent portion 410 of the detonation nozzle 268. Additionally, or in the alternative, the cone-half angle 406 of the detonation nozzle 268 may differ as between annular regions of the detonation nozzle 268 on respectively opposite sides of the detonation chamber plane 402. The respective opposite sides of the detonation nozzle 268 with respect to the detonation chamber plane 402 are sometimes referred to as a detonation chamber-side 412 of the detonation nozzle 268 and a deflagration chamber-side 414 of the detonation nozzle 268. The detonation chamber-side 412 of the detonation nozzle 268 refers to the annular region of the detonation nozzle 268 radially proximal to the detonation chamber 210 and radially distal to the deflagration chamber 224. The deflagration chamber-side 414 of the detonation nozzle 268 refers to the annular region of the detonation nozzle 268 radially proximal to the deflagration chamber 224 and radially distal to the detonation chamber 210.

As shown in FIG. 4A, the detonation nozzle 268 may include a divergent cone-half angle 416. In some embodiments, the divergent cone-half angle 416 may differ as between the detonation chamber-side 412 and the deflagration chamber-side 414 of the detonation nozzle 268. For example, as shown, the divergent cone-half angle 416 may include a first divergent cone-half angle 416a corresponding to the annular region radially proximal to the detonation chamber 210 and radially distal to the deflagration chamber 224, or a second divergent cone-half angle 416b corresponding to the annular region of the detonation nozzle 268 radially proximal to the deflagration chamber 224 and radially distal to the detonation chamber 210. The first annular region of the detonation nozzle 268 and the second annular region of the detonation nozzle 268 may be delineated from one another by the detonation chamber plane 402. The first divergent cone-half angle 416a may sometimes be referred to as a divergent conjugate cone-half angle 416a. The second divergent cone-half angle 416b may sometimes be referred to as a divergent deflagration cone-half angle 416b. In some embodiments, as shown, the divergent conjugate cone-half angle 416a may be less than the divergent deflagration cone-half angle 416b. Additionally, or in the alternative, in some embodiments, the divergent conjugate cone-half angle 416a may intersect the detonation chamber plane 402 upstream from a location at which the divergent deflagration cone-half angle 416b intersects the detonation chamber plane 402. As shown in FIG. 4C, the conjugate nozzle 280 may include a divergent cone-half angle 416. The divergent cone-half angle 416 of the conjugate nozzle 280 may be greater than the divergent cone-half angle 416 of the deflagration nozzle 284 (FIGS. 2A and 2B). Other divergent cone-half angles 416 are also contemplated. A divergent cone-half angle 416, such as the divergent deflagration cone-half angle 416b and/or the divergent conjugate cone-half angle 416a, may be respectively configured to provide suitable combustion conditions within the combustor section 54, such as suitable back pressure for stable detonation and/or a pressure drop suitable for stable flow of detonation combustion products 64 from the detonation chamber 210 to the conjugate chamber 246.

As shown in FIG. 4B, the detonation nozzle 268 may include a convergent cone-half angle 418. In some embodiments, the convergent cone-half angle 418 may differ as between the detonation chamber-side 412 and the deflagration chamber-side 414 of the nozzle 268. For example, as shown, the convergent cone-half angle 418 may include a first convergent cone-half angle 418a corresponding to the annular region of the detonation nozzle 268 radially proximal to the detonation chamber 210 and radially distal to the deflagration chamber 224, and a second convergent cone-half angle 418b corresponding to the annular region of the detonation nozzle 268 radially proximal to the deflagration chamber 224 and radially distal to the detonation chamber 210. The first convergent cone-half angle 418a may sometimes be referred to as a convergent conjugate cone-half angle 418a. The second convergent cone-half angle 418b may sometimes be referred to as a convergent deflagration cone-half angle 418b. In some embodiments, as shown, the convergent conjugate cone-half angle 418a may be less than the convergent deflagration cone-half angle 418b. Additionally, or in the alternative, in some embodiments, the convergent conjugate cone-half angle 418a may intersect the detonation chamber plane 402 upstream from a location at which the convergent deflagration cone-half angle 418b intersects the detonation chamber plane 402. Additionally, or in the alternative, in some embodiments, a convergent cone-half angle 418, such as a convergent conjugate cone-half angle 418a and/or a convergent deflagration cone-half angle 418b, may be less than a divergent cone-half angle 416, such as a divergent conjugate cone-half angle 416a and/or a divergent deflagration cone-half angle 416b. Other convergent cone-half angles 418 are also contemplated. A convergent cone-half angle 418, such as the convergent deflagration cone-half angle 418b and/or the convergent conjugate cone-half angle 418a, may be respectively configured to provide suitable combustion conditions within the combustor section 54, such as suitable back pressure for stable detonation and/or a pressure drop suitable for stable flow of detonation combustion products 64 from the detonation chamber 210 to the conjugate chamber 246.

By way of example, a cone-half angle 406, such as a divergent cone-half angle 416 and/or a convergent cone-half angle 418, may be from about 1 degree to about 30 degrees, such as from about 1° to about 10°, such as from about 10° to about 20°, such as from about 12° to about 18°, or such as from about 20° to about 30°. A divergent conjugate cone-half angle 416a may be from about 1 degree to about 30 degrees, such as from about 1° to about 10°, such as from about 5° to about 10°, such as from about 10° to about 15° , such as from about 15° to about 20°, such as from about 20° to about 25° , or such as from about 25° to about 30°. A divergent deflagration cone-half angle 406 may be from about 10 degrees to about 60 degrees, such as from about 10° to about 20°, such as from about 12° to about 18°, such as from about 20° to about 30°, such as from about 25° to about 40°, or such as from about 40° to about 60°. A divergent deflagration cone-half angle 416b may be greater than, less than, or equal to a divergent conjugate cone-half angle 416a. In some embodiments, a divergent deflagration cone-half angle 416b may be greater than the divergent conjugate cone-half angle 416a, such as from about 10% to about 200% greater than the divergent conjugate cone-half angle 416a, such as from about 10% to about 50%, such as from about 50% to about 100%, such as from about 100% to about 150%, or such as from about 150% to about 200% greater than the divergent conjugate cone-half angle 416a.

A convergent conjugate cone-half angle 418a and/or a convergent deflagration cone-half angle 418b may be from about 1 degree to about 30 degrees, such as from about 1° to about 10°, such as from about 5° to about 10°, such as from about 10° to about 15°, such as from about 15° to about 20°, such as from about 20° to about 25°, or such as from about 25° to about 30°. A convergent deflagration cone-half angle 418b may be greater than, less than, or equal to a convergent conjugate cone-half angle 418a. In some embodiments, a convergent deflagration cone-half angle 418b may be greater than the convergent conjugate cone-half angle 418a, such as from about 10% to about 200% greater than the divergent conjugate cone-half angle 416a, such as from about 10% to about 50%, such as from about 50% to about 100%, such as from about 100% to about 150%, or such as from about 150% to about 200% greater than the convergent conjugate cone-half angle 418a.

In some embodiments, a divergent cone-half angle 416, such as a divergent conjugate cone-half angle 416a and/or a divergent deflagration cone-half angle 416b, may be greater than a convergent cone-half angle 418, such as greater than a convergent conjugate cone-half angle 418a and/or a convergent deflagration cone-half angle 418b. For example, a divergent cone-half angle 416 may be greater than a convergent cone-half angle 418, such as from about 10% to about 200% greater than a convergent cone-half angle 418, such as from about 10% to about 50%, such as from about 50% to about 100%, such as from about 100% to about 150%, or such as from about 150% to about 200% greater than a convergent cone-half angle 418.

In some embodiments, the ratio of the cross-sectional area of the detonation throat 270 to the cross-sectional area of the conjugate throat 282 may be determined based at least in part on a relationship between the operating pressure range of the combustion products 64 in detonation chamber 210 and the pressure of the combustion products 64 in the conjugate chamber 246 and/or the operating pressure range of the combustion products 64 in the deflagration chamber 224. Additionally, or in the alternative, a divergent cone-half angle 416 and/or a convergent cone-half angle 418 may be determined based at least in part on a relationship between the pressure of the combustion products 64 in detonation chamber 210 and the combustion products 64 in the conjugate chamber 246 and/or the combustion products 64 in the deflagration chamber 224 and the combustion products 64 in the conjugate chamber 246. For example, the ratio of the cross-sectional area of the detonation throat 270 to the cross-sectional area of the conjugate throat 282, a divergent cone-half angle 416, and/or a convergent cone-half angle 418, may be determined based at least in part on a relationship between the pressure of the combustion products 64 flowing through and/or exiting the detonation nozzle 268 and the pressure of the combustion products 64 in the conjugate chamber 246. Additionally, or in the alternative, a divergent cone-half angle 416 and/or a convergent cone-half angle 418 may be determined based at least in part on a relationship between the pressure of the combustion products 64 exiting the divergent portion of the detonation nozzle 268 and/or entering the conjugate throat 282, and the pressure of the combustion products 64 in the conjugate chamber 246. Additionally, or in the alternative, a divergent cone-half angle 416 and/or a convergent cone-half angle 418 may be determined based at least in part on a relationship between the pressure of the combustion products 64 approaching conjugate inflection line 272, and the pressure of the combustion products 64 in the conjugate chamber 246.

Such pressures may be determined during operation of the detonation section 202 within an operating range for which the detonation section 202 may be configured, such as under steady state conditions at a rated speed and/or a cruising speed of the engine 50. The pressure of the combustion products 64 flowing through and/or exiting the detonation nozzle 268 may be determined at a downstream region of the diverging portion of the detonation nozzle 268, such as at an upstream side of the conjugate throat 282 within 10% of the distance between the detonation throat 270 and the conjugate throat 282, or such as at an upstream side of the conjugate inflection line 272 within 10% of the distance between the detonation throat 270 and the conjugate inflection line 272. The pressure of the combustion products 64 in the conjugate chamber 246 may be determined downstream from the conjugate nozzle 280, such as at a longitudinal position along the conjugate chamber midline 278 that is within 20% of the length of the conjugate chamber 246 from the conjugate chamber-center line 274 of the conjugate chamber 246.

In some embodiments, the total pressure of the combustion products 64 flowing through and/or exiting the detonation nozzle 268 may be from about 50% greater than the total pressure of the combustion products 64 in the conjugate chamber 246 to about 50% less than the total pressure of the combustion products 64 in the conjugate chamber 246, such as from about 30% greater to about 30% less, such as from about from about 10% greater to about 10% less, such as from about 30% greater to about 5% less, such as from about 5% greater to about 30% less, such as from about 5% less to about 30% less, or such as from about from about 10% to about 20% less, than the total pressure of the combustion products 64 in the conjugate chamber 246, as determined at a rated speed and/or a cruising speed of the engine 50.

In some embodiments and under some operating conditions, the pressure of the combustion products 64 exiting the detonation nozzle 268 as determined between the detonation throat 270 and the conjugate throat 282 may be greater than the pressure of the combustion products 64 exiting the deflagration nozzle 284 as determined between the deflagration throat 282 and the conjugate throat 282. For example, the pressure of the combustion products 64 exiting the detonation nozzle 268 as determined between the detonation throat 270 and the conjugate throat 282 may be from about 1% to about 100% greater, such as from about 10% to about 60% greater, such as from about 10% to about 30% greater, such as from about 50% to about 100% greater, or such as from about 80% to about 100% greater, than the pressure of the combustion products 64 exiting the detonation nozzle 268 as determined between the detonation throat 270 and the conjugate throat 282, as determined at a rated speed and/or a cruising speed of the engine 50.

In some embodiments, the detonation nozzle 268 may have an under-expanded configuration, a neutrally-expanded configuration, or an over-expanded configuration, as determined with respect to a rated speed and/or a cruising speed of the engine 50. In some embodiments, the expansion configuration of the detonation nozzle 268 may differ with respect to a first portion of the detonation nozzle 268 determined between the detonation throat 270 and the conjugate throat 282 and a second portion of the detonation nozzle 268 determined between the detonation throat 270 and the conjugate inflection line 272. The term “under-expanded” or “under-expanded configuration,” when used with reference to the detonation nozzle 268, refers to a configuration of the detonation nozzle 268 that provides for the pressure of combustion products 64 exiting the divergent portion of the detonation nozzle 268 and/or entering the conjugate throat 282 that is at least 5% greater than the pressure of combustion products 64 in the conjugate chamber 246. The term “neutrally-expanded” or “neutrally-expanded configuration,” when used with reference to the detonation nozzle 268, refers to a configuration of the detonation nozzle 268 that provides for the pressure of combustion products 64 exiting the divergent portion of the detonation nozzle 268 and/or entering the conjugate throat 282 that is within 5% of the pressure of combustion products 64 in the conjugate chamber 246. The term “over-expanded” or “over-expanded configuration,” when used with reference to the detonation nozzle 268, refers to a configuration of the detonation nozzle 268 that provides for the pressure of combustion products 64 exiting the divergent portion of the detonation nozzle 268 and/or entering the conjugate throat 282 that is at least 5% less than the pressure of combustion products 64 in the conjugate chamber 246. In some embodiments, the detonation nozzle 268 may have an under-expanded or neutrally-expanded configuration between the detonation throat 270 and the conjugate throat 282 and an over-expanded configuration between the detonation throat 270 and the conjugate inflection line 272.

In some embodiments, a pressure drop of combustion products 64 from the deflagration chamber 224 to the conjugate chamber 246 may be greater than, less than, or equal to a pressure drop of combustion products 64 across the detonation nozzle 268, as determined with respect to a rated speed and/or a cruising speed of the engine 50. In some embodiments, the pressure drop of combustion products 64 from the deflagration chamber 224 to the conjugate chamber 246 may be greater than or equal to the pressure drop of combustion products 64 across the detonation nozzle 268. Additionally, or in the alternative, in some embodiments, the pressure drop of combustion products 64 from the deflagration chamber 224 to the conjugate chamber 246 may be greater than or equal to the pressure drop of combustion products 64 determined between the detonation throat 270 and the conjugate inflection line 272.

By way of example, with respect to a rated speed and/or a cruising speed of the engine 50, the pressure drop of combustion products 64 from the detonation chamber 210 to the conjugate chamber 246 may be from about 100% to about 1% greater, such as from 50% to about 5% greater, such as from 30% to about 5% greater, or such as from 10% to about 1% greater, than the pressure drop of combustion products 64 across the detonation nozzle 268, for example, as determined between the detonation throat 270 and the conjugate inflection line 272.

In general, naturally-expanded nozzles typically have greater efficiency compared to under-expanded nozzles and over-expanded nozzles. In general, over-expanded nozzles typically have greater efficiency than under-expanded nozzles, but that over-expanded nozzles may be less stable. However, in some embodiments, the presently disclosed bimodal combustion systems 200 may include an over-expanded detonation nozzle 268, such as an over-expanded portion of the detonation nozzle 268 between the detonation throat 270 and the conjugate inflection line 272, with respect to a rated speed and/or a cruising speed of the engine 50. Additionally, or in the alternative, the detonation nozzle 268, such as the portion of the detonation nozzle 268 between the detonation throat 270 and the conjugate inflection line 272, may have an under-expanded configuration with respect to a cruising speed of the engine 50 and a neutrally-expanded configuration with respect to a rated speed of the engine 50.

In some embodiments, flow separation of combustion products 64 from the portion of the detonation nozzle 268 between the detonation throat 270 and the conjugate inflection line 272 may be facilitated by a flow of combustion products 64 from the deflagration chamber 224. The pressure of the combustion products 64 flowing from the deflagration chamber 224 to the conjugate chamber 246 may provide support and/or stability to the combustion products 64 exiting the detonation nozzle 268. Additionally, or in the alternative, the stabilizing and/or supporting pressure of the combustion products 64 flowing from the deflagration chamber 224 to the conjugate chamber 246 may at least partially facilitate a configuration of the bimodal combustion system 200 that includes an over-expanded detonation nozzle 268, such as an over-expanded portion of the detonation nozzle 268 between the detonation throat 270 and the conjugate inflection line 272, determined, for example, with respect to a rated speed and/or a cruising speed of the engine 50. In some embodiments, the detonation nozzle 268 may exhibit over-expanded characteristics at a cruising speed and neutrally-expanded characteristics at a rated speed, for example, as a result of increasing pressure in the deflagration chamber 224 and/or in the conjugate chamber 246 as the engine 50 transitions from a cruising speed to a rated speed.

In some embodiments, the detonation nozzle 268 and the conjugate nozzle 280 may define a stepped nozzle, a dual expansion nozzle, and/or a dual throat nozzle. In some embodiments, under a first operating condition, the detonation nozzle 268 may exhibit under-expanded characteristics while the conjugate nozzle 280 may exhibit neutrally-expanded characteristics and/or over-expanded characteristics. Additionally, or in the alternative, under a second operating condition, the detonation nozzle 268 may exhibit neutrally-expanded characteristics while the conjugate nozzle 280 may exhibit over-expanded characteristics.

In some embodiments, the detonation nozzle 268 and/or the conjugate nozzle 280 may be configured for choked flow within an operating range for which the detonation section 202 may be configured, such as at a cruising speed and/or at a rated speed. As used herein, the term “choked flow” refers to a limiting condition where the mass flow through the detonation nozzle 268 and/or the conjugate nozzle 280 will not increase with a further decrease in the downstream pressure for a given upstream pressure and temperature. Choked flow may occur when a ratio of the pressure of the combustion products 64 on opposite sides of the detonation nozzle 268 is at least about 1.7, such as from at least about 1.7 to at least about 2.1. The detonation nozzle-pressure ratio at which choked flow may occur may depend at least in part on the configuration of the nozzle region 266 and/or other portions of the detonation chamber 210, as well as on the composition of the combustion products 64. In some embodiments, an engine may be configured to exhibit a detonation nozzle-pressure ratio suitable for a choked flow condition upon reaching a specified operating state, such as a high-power operating state, a nominal operating state, a cruising speed, and/or a rated speed.

Still referring to FIGS. 2A and 2B, and with reference to FIGS. 5A-5D, exemplary bimodal combustion systems 200 are further described. As shown in FIG. 5A, at least a portion of the detonation chamber 210 may have a shape that includes a torus shape. For example, the detonation region 264 of the detonation chamber 210 may include a torus shape. The toros shape of the detonation chamber 210, such as the detonation region 264 of the detonation chamber 210, may include a ring torus, a horn torus, or a spindle torus. As shown in FIG. 5B, at least a portion of the detonation chamber 210 may have a shape that includes a parabolic annulus shape.

For example, the nozzle region 266 of the detonation chamber 210 may include a parabolic annulus shape. The parabolic annulus shape of the detonation chamber 210, such as the nozzle region 266 of the detonation chamber 210, may include a convergent portion 500, a divergent portion 502, and a saddle region 504 disposed between the convergent portion 500 and the divergent portion 502. At least a portion of the nozzle region 266 that includes the convergent portion 500 and the divergent portion 502 may define the detonation nozzle 268. The saddle region 504 may define the detonation throat 270. The toros shape of the detonation chamber 210 may define at least a portion of the nozzle region 266 of the detonation chamber 210, for example, in addition to defining at least a portion of the detonation region 264 of the detonation chamber 210. A portion of the parabolic annulus shape of the detonation chamber 210 may be defined by a portion of the toros shape of the detonation chamber 210. For example, the toros shape of the detonation chamber 210 may define at least a portion of the convergent portion 500 of the nozzle region 266.

As shown in FIG. 5C, the deflagration chamber 224 may have a generally cylindrical annulus shape. Additionally, or in the alternative, as shown in FIG. 5D, the conjugate chamber 246 may have a generally cylindrical annulus shape. The cylindrical annulus shape of the deflagration chamber 224 and/or of the conjugate chamber 246 may have a frustoconical configuration. Additionally, or in the alternative, the cylindrical annulus shape of the deflagration chamber 224 and/or of the conjugate chamber 246 may have a wavy configuration with one or more saddles or narrowing portions, and/or with one or more nodes or widening portions.

Referring now to FIGS. 6A-6C, exemplary detonation combustors 206 are further described. As shown, a detonation combustor 206 may include a detonation fuel manifold 212 that has a plurality of detonation fuel orifices 218. The fuel orifices may be circumferentially spaced about the detonation fuel manifold 212. The plurality of detonation fuel orifices 218 may supply fuel 62 and/or oxidizer 60 from the detonation fuel manifold 212 to the detonation chamber 210. For example, the plurality of detonation fuel orifices 218 may supply a mixture of fuel 62 and/or oxidizer 60 to the detonation chamber 210. Additionally, or in the alternative, a first plurality of detonation fuel orifices 218a may supply fuel 62 to the detonation chamber 210 and a second plurality of detonation fuel orifices 218b may supply oxidizer 60 to the detonation chamber 210. Respective ones of the plurality of detonation fuel orifices 218 may be in fluid communication with one or more detonation-fuel supply lines 214 and/or with one or more detonation-oxidizer supply lines 216.

As shown in FIG. 6A, the detonation fuel manifold 212 may include a plurality of detonation fuel orifices 218 circumferentially spaced about an annular face 600 of the detonation fuel manifold 212. Additionally, or in the alternative, as shown in FIGS. 6B and 6C, the detonation fuel manifold 212 may include a plurality of detonation fuel orifices 218 circumferentially spaced about a circumferential face 604 of the detonation fuel manifold 212, such as an inward-facing circumferential face 604a (FIG. 6B) and/or an outward-facing circumferential face 604b (FIG. 6C) of the detonation fuel manifold 212.

Referring now to FIGS. 7A-7F, and with further reference to FIGS. 4A-4C, exemplary bimodal combustion systems 200 are further described. As described with respect to FIGS. 4A-4C, and as further shown in FIGS. 7A-7F, a detonation throat-center line 400 may define an annular center of the detonation throat 270, and a detonation chamber plane 402 may intersect the detonation throat-center line 400 tangentially normal to the detonation throat-center line 400. Also as shown in FIGS. 7A-7F, the deflagration chamber 224 may have a deflagration chamber-center line 700 defining an annular center of the deflagration chamber 224. The deflagration chamber-center line 700 may be located equidistantly between the outer deflagration chamber wall 254 and the inner deflagration chamber wall 256 at a location of the deflagration chamber 224 that has a minimum annular ring width determined between the outer deflagration chamber wall 254 and the inner deflagration chamber wall 256. The deflagration chamber-center line 700 may have a generally elliptical or circular shape. The deflagration chamber-center line 700 may circumferentially surround the longitudinal axis 248 of the engine 50. A deflagration chamber plane 702 may intersect the deflagration chamber-center line 700 tangentially normal to the deflagration chamber-center line 700. The tangentially normal orientation of the deflagration chamber plane 702 may include a perpendicular orientation relative to an annular plane spanning the minimum annular ring width determined between the outer deflagration chamber wall 254 and the inner deflagration chamber wall 256.

The deflagration chamber plane 702 may circumferentially surround the longitudinal axis 248 of the engine 50. The deflagration chamber plane 702 may have a generally linear configuration along a deflagration chamber midline 704 intersecting the conjugate chamber midline 278 and the deflagration chamber-center line 700 at an orientation parallel to the deflagration chamber plane 702. By way of example, the deflagration chamber plane 702 may have a cylindrical configuration or a frustoconical configuration.

The detonation chamber plane 402 may intersect the deflagration chamber plane 702 at a conjugate intersection 706 located within the conjugate chamber 246. In some embodiments, as shown in FIG. 7A, the detonation chamber plane 402 may intersect the deflagration chamber plane 702 at a conjugate intersection 706 coinciding with the conjugate chamber plane 276. Additionally, or in the alternative, as shown in FIG. 7A, the detonation chamber plane 402 may intersect the deflagration chamber plane 702 at a conjugate intersection 706 coinciding with the conjugate chamber-center line 274. In some embodiments, at least one detonation chamber midline 404 may intersect at least one deflagration chamber midline 704 at a conjugate intersection 706 coinciding with the conjugate chamber plane 276 and/or coinciding with the conjugate chamber-center line 274.

In some embodiments, the detonation chamber plane 402 may intersect the deflagration chamber plane 702 at a conjugate intersection 706 located radially outward from the conjugate chamber plane 276 or at a location radially inward from the conjugate chamber plane 276. Additionally, or in the alternative, the detonation chamber plane 402 may intersect the deflagration chamber plane 702 at a conjugate intersection 706 located upstream from the conjugate chamber-center line 274 or downstream from the conjugate chamber-center line 274. In some embodiments, at least one detonation chamber midline 404 may intersect at least one deflagration chamber midline 704 at a conjugate intersection 706 located radially outward from the conjugate chamber plane 276 and/or at a location radially inward from the conjugate chamber plane 276. Additionally, or in the alternative, in some embodiments, at least one detonation chamber midline 404 may intersect at least one deflagration chamber midline 704 at a conjugate intersection 706 located upstream from the conjugate chamber-center line 274 and/or at a location downstream from the conjugate chamber-center line 274. The detonation chamber plane 402 and the deflagration chamber plane 702 may intersect one another at a normal angle (e.g., a right angle) or at an oblique angle, such as an acute angle or an obtuse angle. Additionally, or in the alternative, at least one detonation chamber midline 404 and at least one deflagration chamber midline 704 may intersect one another at a normal angle (e.g., a right angle) or at an oblique angle, such as an acute angle or an obtuse angle. By way of example, FIG. 7A shows a detonation chamber plane 402 intersecting a deflagration chamber plane 702 at a conjugate intersection 706 coinciding with the conjugate chamber-center line 274 of the conjugate chamber 246, for example, at an oblique angle, such as an acute angle. As shown, at least one detonation chamber midline 404 may intersect at least one deflagration chamber midline 704 at an oblique or acute angle, for example, at a conjugate intersection 706 coinciding with the conjugate chamber-center line 274.

As another example, FIG. 7B shows a detonation chamber plane 402 intersecting a deflagration chamber plane 702, at an oblique or acute angle, at a location radially inward from the conjugate chamber plane 276 and upstream from the conjugate chamber-center line 274. As shown, at least one detonation chamber midline 404 may intersect at least one deflagration chamber midline 704, at an oblique or acute angle, at a location radially inward from the conjugate chamber plane 276 and upstream from the conjugate chamber-center line 274.

As another example, FIG. 7C shows a detonation chamber plane 402 intersecting a deflagration chamber plane 702, at an oblique or acute angle, at a location radially outward from the conjugate chamber plane 276 and upstream from the conjugate chamber-center line 274. As shown, at least one detonation chamber midline 404 may intersect at least one deflagration chamber midline 704, at an oblique or acute angle, at a location radially outward from the conjugate chamber plane 276 and upstream from the conjugate chamber-center line 274.

As another example, FIG. 7D shows a detonation chamber plane 402 intersecting a deflagration chamber plane 702, at an oblique or acute angle, at a location radially outward from the conjugate chamber plane 276 and downstream from the conjugate chamber-center line 274. As shown, at least one detonation chamber midline 404 may intersect at least one deflagration chamber midline 704, at an oblique or acute angle, at a location radially outward from the conjugate chamber plane 276 and downstream from the conjugate chamber-center line 274.

As another example, FIG. 7E shows a detonation chamber plane 402 intersecting a deflagration chamber plane 702, at normal angle, at a location radially outward from the conjugate chamber plane 276 and upstream from the conjugate chamber-center line 274. As shown, at least one detonation chamber midline 404 may intersect at least one deflagration chamber midline 704, at a normal angle, at a location radially outward from the conjugate chamber plane 276 and upstream from the conjugate chamber-center line 274.

As another example, FIG. 7F shows a detonation chamber plane 402 intersecting a deflagration chamber plane 702, at an oblique angle, such as an obtuse angle, at a location radially outward from the conjugate chamber plane 276 and upstream from the conjugate chamber-center line 274. As shown, at least one detonation chamber midline 404 may intersect at least one deflagration chamber midline 704, at an oblique or obtuse angle, at a location radially outward from the conjugate chamber plane 276 and upstream from the conjugate chamber-center line 274.

Referring now to FIG. 8, exemplary detonation combustors 206 are further described. As shown in FIG. 8, a detonation wave 800 preceded by a shock wave 802 may propagate annularly through the detonation chamber 210, such as through the detonation region 264 of the detonation chamber 210. The shock wave 802 and corresponding detonation wave 800 may propagate in a counter-clockwise direction while combustion products 64 expand in generally three-dimensions, as shown. Alternatively, the shock wave 802 and corresponding detonation wave 800 may propagate annularly, for example, in a clockwise direction. While one shock wave 802 and corresponding detonation wave 800 are depicted in FIG. 8 for illustrative purposes, in some embodiments, exemplary detonation combustors 206 may be configured to continuously generate a plurality of shock waves 802 and corresponding detonation waves 800. For example, a plurality of shock waves 802 and corresponding detonation waves 800 may concurrently propagate around the annular volume of the detonation chamber 210, for example, with a circumferentially spaced relationship.

While the detonation chamber 210 schematically shown in FIG. 8 has a generally cylindrical annulus shape, in some embodiments, the detonation chamber 210 may include any shape that provides a continuous path for the shock wave 802 and corresponding detonation wave 800 to follow. By way of example, a detonation chamber 210 may include a torus shape, as shown in FIG. 5A. Additionally, or in the alternative, a detonation chamber 210 may include a parabolic annulus shape, as shown in FIG. 5B. As further examples, a detonation chamber 210 may include any annular shape, such as a trapezoidal shape or an elliptical shape. For a detonation chamber 210 that includes a torus shape as shown in FIG. 5A, the shock wave 802 and corresponding detonation wave 800 may propagate in a toroidal direction 506 (as shown in FIG. 5A). The shock wave 802 and corresponding detonation wave 800 may envelop all or a portion of the annular perimeter defined by the detonation chamber 210, such as all or a portion of the annular perimeter defined by the detonation region 264 of the detonation chamber 210. As an example, for a detonation chamber 210 that includes a torus shape as shown in FIG. 5A, the shock wave 802 and corresponding detonation wave 800 may envelop all or a portion of the poloidal perimeter 508 (as shown in FIG. 5A) defined by the detonation region 264 of the detonation chamber 210.

As shown in FIG. 8, a region preceding the shock wave 802 and corresponding detonation wave 800 may include a mixture of fuel 62 and oxidizer 60 at a concentration suitable for detonation. As the mixture of fuel 62 and oxidizer 60 detonates, the shock wave 802 generated by the detonation may temporarily inhibit further fuel 62 and oxidizer 60 from entering the detonation chamber 210. The shock wave 802 and corresponding detonation wave 800 may propagate around the annular volume of the detonation chamber 210, consuming further fuel 62 and oxidizer 60. As the shock wave 802 and corresponding detonation wave 800 propagates around the annular volume, additional fuel 62 and oxidizer 60 may flow into the detonation chamber 210 generally trailing the shock wave 802 and corresponding detonation wave 800.

As the combustion products 64 expand while propagating through the detonation region 264 of the detonation chamber 210, at least a portion of the shock wave 802 may pass from the detonation region 264 of the detonation chamber 210 to the nozzle region 266 of the detonation chamber 210. The shock wave 802 may transition from a generally rotational direction of propagation to a helical or longitudinal direction of propagation as the shock waves passes from the detonation region 264 through the nozzle region 266 of the detonation chamber 210. A shock wave 802 propagating in the circumferential or toroidal direction 506 may sometime be referred to as a primary shock wave, a circumferential shock wave, or a toroidal shock wave. In some embodiments, a longitudinal shock wave 804 may be generated that propagates from the detonation chamber 210 into the conjugate chamber 246 (FIGS. 2A and 2B). The longitudinal shock wave 804 may be oriented generally along the at least one detonation chamber midline 404. The longitudinal shock wave 804 may be generated as the combustion products 64 pass through the nozzle region 266 of the detonation chamber 210 into the conjugate chamber 246. The longitudinal shock wave 804 may sometimes be referred to as a secondary shock wave 804. The longitudinal shock wave 804 may envelop at least a portion of the circumference 510 (FIG. 5B) of the nozzle region 266 of the detonation chamber 210.

While one longitudinal shock wave 804 is depicted in FIG. 8 for illustrative purposes, in some embodiments, exemplary detonation combustors 206 may be configured to continuously generate a plurality of longitudinal shock waves 804. For example, a plurality of longitudinal shock waves 804 may concurrently propagate longitudinally from the detonation chamber 210 into the conjugate chamber 246, for example, with a circumferentially spaced relationship. Additionally, or in the alternative, the plurality of longitudinal shock waves 804 may have an annular configuration, such as an annular configuration oriented with respect to the detonation throat-center line 400, the at least one detonation chamber midline 404, and/or the conjugate chamber midline 278. The annularly configured longitudinal shock waves 804 may propagate longitudinally from the detonation chamber 210 into the conjugate chamber 246, generating thrust. Additionally, or in the alternative, the longitudinal shock waves 804 may propagate longitudinally from the conjugate chamber 246 into one or more turbine sections 66 (FIGS. 1A and 1B) of the engine 50, through an outlet section 56 (FIGS. 1A and 1B) of the engine 50, and/or through the outlet nozzle 68 of the engine 50 (FIGS. 1A and 1B), generating thrust.

The longitudinal shock wave 804 may emanate from detonation nozzle 268, such as the detonation throat 270 of the nozzle region 266 of the detonation chamber 210. The longitudinal shock wave 804 may propagate from the nozzle region 266 of the detonation chamber 210 and into the conjugate chamber 246. Additionally, or in the alternative, in some embodiments, the primary shock wave 802 may propagate through the nozzle region 266 of the detonation chamber 210 and into the conjugate chamber 246. In some embodiments, the detonation wave 800 may remain within the detonation region 264 of the detonation chamber 210, for example, such that the detonation process may be completed within the detonation region 264 of the detonation chamber 210, while the longitudinal shock wave 804 and corresponding combustion products 64 propagate through the nozzle region 266 of the detonation chamber 210 into the conjugate chamber 246. Alternatively, in some embodiments, a portion of the detonation may occur within the nozzle region 266 of the detonation chamber 210. For example, in some embodiments, the detonation wave 800 may remain upstream from the detonation throat 270. In some embodiments, at least some combustion, such as detonation and/or deflagration, may occur within the conjugate chamber 246.

In some embodiments, the detonation combustor 206 may include a pre-detonator 806 configured to generate a blast wave 808 suitable to initiate detonation within the detonation chamber 210. Additionally, or in the alternative, in some embodiments, the longitudinal shock wave 804 may be initiated at least in part by reflection of the primary shock wave 802 by the detonation nozzle 268. Additionally, or in the alternative, the longitudinal shock wave 804 may be initiated at least in part by backpressure generated by the detonation nozzle 268 and/or by the conjugate nozzle 280 (FIGS. 2A and 2B). Additionally, or in the alternative, the longitudinal shock wave 804 may be initiated at least in part by backpressure generated from deflagration occurring within the deflagration chamber 224 (FIGS. 2A and 2B) and/or the conjugate chamber 246. In some embodiments, detonation within the detonation chamber 210 may be initiated at least in part by way of back pressure generated by the detonation nozzle 268, by the conjugate nozzle 280, and/or by deflagration occurring within the deflagration chamber 224 and/or the conjugate chamber 246. In some embodiments, detonation may be initiated upon establishing a choked flow condition, for example, at the detonation nozzle-pressure ratio increases to a suitable level, for example, as a result of such backpressure. In some embodiments, fuel 62 and oxidizer 60 suitable for detonation may be supplied to the detonation chamber upon having established a choked flow condition. In some embodiments, deflagration may be performed in the detonation chamber 210 prior to initiating detonation within the detonation chamber 210. For example, the detonation chamber 210 may be utilized for deflagration during specified operating conditions. Deflagration may be performed within the detonation chamber 210, by providing fuel 62 and oxidizer 60 suitable for deflagration, such as a fuel 62 and oxidizer 60 mixture that would be unsuitable for detonation. Additionally, or in the alternative, deflagration may be performed within the detonation chamber 210 prior to establishing a choked flow condition.

Referring now to FIG. 9, exemplary methods in accordance with the present disclosure are further described. By way of example, an exemplary method may include a method of generating thrust. Additionally, or in the alternative, an exemplary method may include a method of combusting fuel. Additionally, or in the alternative, exemplary method may include a method of operating an engine 50, such as a turbine engine 100, a rocket engine, a ramjet, or a combination thereof, such as a turbo-rocket engine, a turbo-ramjet, or a rocket-ramjet. As shown in FIG. 9, an exemplary method 900 may include, at block 902, performing deflagration within a deflagration chamber and/or within a conjugate chamber in fluid communication with the deflagration chamber. Combustion products generated by deflagration in the deflagration chamber may flow through the conjugate chamber, such as from the deflagration chamber through the conjugate chamber. At block 904, the exemplary method 900 may include performing detonation within a detonation chamber in fluid communication with the conjugate chamber. Combustion products generated by detonation in the detonation chamber may flow from the detonation chamber to the conjugate chamber. Combustion products flowing through the conjugate chamber may generate thrust. The thrust may be attributable at least in part to a shock wave emanating from the detonation chamber, such as a longitudinal shock wave emanating from the detonation nozzle as the combustion products pass through the nozzle region of the detonation chamber into the conjugate chamber. Additionally, or in the alternative, the thrust may be attributable at least in part to deflagration performed within the deflagration chamber and/or within the conjugate chamber.

Accordingly, the presently disclosed systems and methods may utilize a bimodal combustion system to provide thrust for an engine while realizing improved performance and/or an ability to operate over a wider range of operating conditions and thermal load requirements. Additionally, or in the alternative, the presently disclosed systems and methods may provide significantly improved specific impulse and/or specific fuel consumption, and/or relatively low NOx emissions. Additionally, or in the alternative, exemplary engines for specified duty requirements may be relatively smaller, lighter-weight, and/or may exhibit a higher thrust-to-weight ratio, as a result of the presently disclosed systems and methods.

Further aspects of the disclosure are provided by the subject matter of the following clauses:

A combustion system, comprising: a detonation combustor comprising one or more detonation chamber walls defining a detonation chamber; a deflagration combustor comprising one or more deflagration chamber walls defining a deflagration chamber; and one or more conjugate chamber walls defining a conjugate chamber, the conjugate chamber in fluid communication with the detonation chamber and the deflagration chamber; wherein the detonation chamber comprises a detonation region and a nozzle region, the nozzle region providing fluid communication between the detonation region and the conjugate chamber.

The combustion system of any preceding clause, wherein the nozzle region comprises a detonation nozzle defined by a first one of the one or more detonation chamber walls, the detonation nozzle comprising a detonation throat defining a location of the detonation nozzle that has an annular cross-sectional area with a minimum annular ring width relative to an adjacent portion of the detonation nozzle.

The combustion system of any preceding clause, wherein the detonation nozzle comprises a divergent portion located downstream of the detonation throat, wherein the divergent portion of the detonation nozzle has a cone-half angle of from 1 degree to 10 degrees.

The combustion system of any preceding clause, wherein the detonation nozzle comprises a convergent portion located upstream of the detonation throat, wherein the convergent portion of the detonation nozzle has a cone-half angle of from 5 degrees to 30 degrees.

The combustion system of any preceding clause, wherein the detonation nozzle comprises a convergent portion and a divergent portion, the convergent portion having a decreasing cross-sectional area upstream from the detonation throat in a direction from the detonation region towards the detonation throat, and the divergent portion having an increasing cross-sectional area downstream from the detonation throat in a direction from the detonation throat towards the conjugate chamber.

The combustion system of any preceding clause, wherein the detonation nozzle is configured as a de Laval type nozzle.

The combustion system of any preceding clause, wherein at least a portion of the detonation nozzle has an over-expanded configuration, as determined with respect to a rated speed and/or a cruising speed of an engine receiving thrust from the combustion system.

The combustion system of any preceding clause, wherein at least a portion of the detonation nozzle has an under-expanded configuration and/or a neutrally-expanded configuration, as determined with respect to the rated speed and/or the cruising speed of the engine.

The combustion system of any preceding clause, wherein: a detonation throat-center line defines an annular center of the detonation throat and a detonation chamber plane intersects the detonation throat-center line tangentially normal to the detonation throat-center line; a deflagration chamber-center line defines an annular center of the deflagration chamber, and a deflagration chamber plane intersects the deflagration chamber-center line tangentially normal to the deflagration chamber-center line; a conjugate inflection line circumferentially surrounding a longitudinal axis defines a linear inflection delineating a detonation chamber wall and a deflagration chamber wall from one another, or a linear inflection representing a forwardmost oblique angle or a tangent to a forwardmost curve of a conjugate chamber wall disposed between a detonation chamber wall and a deflagration chamber wall; a conjugate chamber-center line defines a volumetric center of the conjugate chamber as determined with respect to a volume of the conjugate chamber located between the conjugate inflection line and a downstream end of the conjugate chamber; a conjugate chamber plane intersects the conjugate inflection line and the conjugate chamber-center line; and the detonation chamber plane and the deflagration chamber plane intersect one another at a normal angle or at an oblique angle.

The combustion system of any preceding clause, wherein the detonation chamber plane and the deflagration chamber plane intersect one another at a conjugate intersection comprising a location within the conjugate chamber that is at least one of: coinciding with the conjugate chamber plane, radially inward from the conjugate chamber plane, or radially outward from the conjugate chamber plane; and coinciding with the conjugate chamber-center line, upstream from the conjugate chamber-center line, or downstream from the conjugate chamber-center line.

The combustion system of any preceding clause, wherein the deflagration chamber and the detonation chamber respectively transition to the conjugate chamber along a longitudinal axis.

The combustion system of any preceding clause, wherein the deflagration chamber and the detonation chamber are located at respectively opposite sides of a conjugate inflection line circumferentially surrounding the longitudinal axis, the conjugate inflection line defining a linear inflection delineating a detonation chamber wall and a deflagration chamber wall from one another, or the conjugate inflection line defining a linear inflection representing a forwardmost oblique angle or a tangent to a forwardmost curve of a conjugate chamber wall disposed between a detonation chamber wall and a deflagration chamber wall.

The combustion system of any preceding clause, wherein the detonation chamber comprises a detonation nozzle defined by a first one of the one or more detonation chamber walls, the detonation nozzle comprising a detonation throat defining a location of the detonation nozzle that has a first annular cross-sectional area with a first minimum annular ring width relative to an adjacent portion of the detonation nozzle; wherein a first one of the one or more conjugate chamber walls comprises a conjugate nozzle, the conjugate nozzle comprising a conjugate throat defining a location of the conjugate nozzle that has a second annular cross-sectional area with a second minimum annular ring width relative to an adjacent portion of the conjugate nozzle; wherein the second annular cross-sectional area corresponding to the conjugate throat extends from the first one of the one or more conjugate chamber walls to a conjugate chamber plane, the conjugate chamber plane intersecting the conjugate inflection line and a conjugate chamber-center line, the conjugate chamber-center line defining a volumetric center of the conjugate chamber as determined with respect to a volume of the conjugate chamber located between the conjugate inflection line and a downstream end of the conjugate chamber, and the first one of the one or more conjugate chamber walls being located on a side of the conjugate chamber plane radially corresponding to the detonation chamber; and wherein the first annular cross-sectional area corresponding to the detonation throat is less than the second annular cross-sectional area corresponding to the conjugate throat.

The combustion system of any preceding clause, wherein the first annular cross-sectional area is from 1% to 90% less than the second annular cross-sectional area.

The combustion system of any preceding clause, wherein the detonation nozzle comprises a first divergent cone-half angle corresponding to a first annular region of the detonation nozzle radially proximal to the detonation chamber and radially distal to the deflagration chamber, and a second divergent cone-half angle corresponding to a second annular region of the detonation nozzle radially proximal to the deflagration chamber and radially distal to the detonation chamber, wherein the second divergent cone-half angle is greater than the first divergent cone-half angle.

The combustion system of any preceding clause, wherein the first divergent cone-half angle is from 1 degree to 10 degrees, and/or wherein the second divergent cone-half angle is from 1 degree to 10 degrees.

The combustion system of any preceding clause, wherein the second cone-half angle is from 10% to 200% greater than the first cone-half angle.

The combustion system of any preceding clause, wherein the detonation combustor comprises a detonation fuel manifold coupled to or monolithically integrated with the one or more detonation chamber walls, the detonation fuel manifold configured to supply fuel and/or oxidizer to the detonation chamber.

The combustion system of any preceding clause, wherein the deflagration combustor comprises a plurality of deflagration fuel manifolds respectively configured to supply fuel and/or oxidizer to the deflagration chamber.

The combustion system of any preceding clause, wherein at least a portion of the detonation chamber circumferentially surrounds at least a portion of the deflagration chamber, or wherein at least a portion of the deflagration chamber circumferentially surrounds at least a portion of the detonation chamber.

The combustion system of any preceding clause, wherein at least a portion of the detonation chamber comprises a torus shape, and/or wherein at least a portion of the detonation chamber comprises parabolic annulus shape.

An engine, comprising: an inlet section; a combustor section; and an outlet section; wherein the combustor section comprises a bimodal combustion system, the bimodal combustion system comprising: a detonation combustor comprising one or more detonation chamber walls defining a detonation chamber; a deflagration combustor comprising one or more deflagration chamber walls defining a deflagration chamber; and one or more conjugate chamber walls defining a conjugate chamber, the conjugate chamber in fluid communication with the detonation chamber and the deflagration chamber; wherein the detonation chamber comprises a detonation region and a nozzle region, the nozzle region providing fluid communication between the detonation region and the conjugate chamber.

The engine of any preceding clause, wherein the engine comprises:

a turbine engine, a rocket engine, a ramjet, a turbo-rocket engine, a turbo-ramjet, or a rocket-ramjet.

The engine of any preceding clause, wherein the engine comprises a turbine engine, the turbine engine comprising a turbine section disposed downstream of the combustor section.

The engine of any preceding clause, wherein the turbine engine comprises a compressor section disposed upstream of the combustor section.

The engine of any preceding clause, wherein the compressor section comprises from 1 to 12 compressor stages.

The engine of any preceding clause, wherein the turbine engine exhibits a bypass ratio of from about 10:1 to about 20:1 at a rated speed and/or at a cruising speed.

The engine of any preceding clause, wherein the turbine engine exhibits a thrust to weight ratio of from about 6.0 to about 9.0.

The engine of any preceding clause, wherein the turbine engine exhibits a thrust specific fuel consumption of from about 8 grams per kilonewton-second to about 14 grams per kilonewton-second at a rated speed and/or at a cruising speed.

The engine of any preceding clause, wherein the turbine engine generates from about 300 kilonewtons of thrust to about 700 kilonewtons of thrust at a rated speed and/or at a cruising speed.

The engine of any preceding clause, wherein the turbine engine generates from about 10 kilonewtons of thrust to about 300 kilonewtons of thrust at a rated speed and/or at a cruising speed.

The engine of any preceding clause, wherein the combustion system is configured according to any preceding clause.

A method of combusting fuel, the method comprising: performing deflagration within a deflagration chamber and/or within a conjugate chamber in fluid communication with the deflagration chamber, generating deflagration combustion products, wherein the deflagration combustion products flow through the conjugate chamber, generating thrust; and performing detonation within a detonation chamber in fluid communication with the conjugate chamber, generating detonation combustion products, wherein the detonation combustion products flow through the conjugate chamber, generating thrust.

The method of any preceding clause, wherein performing detonation within the detonation chamber comprises: generating a plurality of primary shock waves that propagate annularly through the detonation chamber.

The method of any preceding clause, wherein performing detonation within the detonation chamber comprises: generating a plurality of shock waves that propagate longitudinally through the detonation chamber, generating thrust.

The method of any preceding clause, wherein the detonation chamber comprises a detonation nozzle, and wherein the detonation combustion products have a velocity of from 1,000 meters per second to 5,000 m/s meters per second downstream of the detonation nozzle.

The method of any preceding clause, wherein the method is performed using the combustion system of any preceding clause or the engine of any preceding clause.

This written description uses exemplary embodiments to describe the presently disclosed subject matter, including the best mode, and also to enable any person skilled in the art to practice such subject matter, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the presently disclosed subject matter is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Claims

1. A combustion system, comprising:

a detonation combustor comprising one or more detonation chamber walls defining a detonation chamber;
a deflagration combustor comprising one or more deflagration chamber walls defining a deflagration chamber; and
one or more conjugate chamber walls defining a conjugate chamber, the conjugate chamber in fluid communication with the detonation chamber and the deflagration chamber;
wherein the detonation chamber comprises a detonation region and a nozzle region, the nozzle region providing fluid communication between the detonation region and the conjugate chamber;
wherein the nozzle region comprises a first divergent cone-half angle and a second divergent cone-half angle, wherein the second divergent cone-half angle is greater than the first divergent cone-half angle.

2. The combustion system of claim 1, wherein the nozzle region comprises a detonation nozzle defined by a first one of the one or more detonation chamber walls, the detonation nozzle comprising a detonation throat defining a location of the detonation nozzle that has an annular cross-sectional area with a minimum annular ring width relative to an adjacent portion of the detonation nozzle.

3. The combustion system of claim 2, wherein the detonation nozzle comprises a divergent portion located downstream of the detonation throat, wherein the divergent portion of the detonation nozzle has a cone-half angle of from 1 degree to 10 degrees, and/or wherein the detonation nozzle comprises a convergent portion located upstream of the detonation throat, wherein the convergent portion of the detonation nozzle has a cone-half angle of from 5 degrees to 30 degrees.

4. The combustion system of claim 2, wherein the detonation nozzle comprises a convergent portion and a divergent portion, the convergent portion having a decreasing cross-sectional area upstream from the detonation throat in a direction from the detonation region towards the detonation throat, and the divergent portion having an increasing cross-sectional area downstream from the detonation throat in a direction from the detonation throat towards the conjugate chamber.

5. The combustion system of claim 2, wherein the detonation nozzle is configured as a de Laval type nozzle.

6. The combustion system of claim 2, wherein at least a portion of the detonation nozzle has an over-expanded configuration, as determined with respect to a rated speed and/or a cruising speed of an engine receiving thrust from the combustion system.

7. The combustion system of claim 6, wherein at least a portion of the detonation nozzle has an under-expanded configuration and/or a neutrally-expanded configuration, as determined with respect to the rated speed and/or the cruising speed of the engine.

8. The combustion system of claim 2, wherein:

a detonation throat-center line defines an annular center of the detonation throat and a detonation chamber plane intersects the detonation throat-center line tangentially normal to the detonation throat-center line;
a deflagration chamber-center line defines an annular center of the deflagration chamber, and a deflagration chamber plane intersects the deflagration chamber-center line tangentially normal to the deflagration chamber-center line;
a conjugate inflection line circumferentially surrounding a longitudinal axis defines a linear inflection delineating a detonation chamber wall and a deflagration chamber wall from one another, or a linear inflection representing a forwardmost oblique angle or a tangent to a forwardmost curve of a conjugate chamber wall disposed between a detonation chamber wall and a deflagration chamber wall;
a conjugate chamber-center line defines a volumetric center of the conjugate chamber as determined with respect to a volume of the conjugate chamber located between the conjugate inflection line and a downstream end of the conjugate chamber;
a conjugate chamber plane intersects the conjugate inflection line and the conjugate chamber-center line; and
the detonation chamber plane and the deflagration chamber plane intersect one another at a normal angle or at an oblique angle.

9. The combustion system of claim 8, wherein the detonation chamber plane and the deflagration chamber plane intersect one another at a conjugate intersection comprising a location within the conjugate chamber that is at least one of:

coinciding with the conjugate chamber plane, radially inward from the conjugate chamber plane, or radially outward from the conjugate chamber plane; and
coinciding with the conjugate chamber-center line, upstream from the conjugate chamber-center line, or downstream from the conjugate chamber-center line.

10. The combustion system of claim 1, wherein the deflagration chamber and the detonation chamber respectively transition to the conjugate chamber along a longitudinal axis.

11. The combustion system of claim 10, wherein the deflagration chamber and the detonation chamber are located at respectively opposite sides of a conjugate inflection line circumferentially surrounding the longitudinal axis, the conjugate inflection line defining a linear inflection delineating a detonation chamber wall and a deflagration chamber wall from one another, or the conjugate inflection line defining a linear inflection representing a forwardmost oblique angle or a tangent to a forwardmost curve of a conjugate chamber wall disposed between a detonation chamber wall and a deflagration chamber wall.

12. The combustion system of claim 11, wherein the detonation chamber comprises a detonation nozzle defined by a first one of the one or more detonation chamber walls, the detonation nozzle comprising a detonation throat defining a location of the detonation nozzle that has a first annular cross-sectional area with a first minimum annular ring width relative to an adjacent portion of the detonation nozzle;

wherein a first one of the one or more conjugate chamber walls comprises a conjugate nozzle, the conjugate nozzle comprising a conjugate throat defining a location of the conjugate nozzle that has a second annular cross-sectional area with a second minimum annular ring width relative to an adjacent portion of the conjugate nozzle;
wherein the second annular cross-sectional area corresponding to the conjugate throat extends from the first one of the one or more conjugate chamber walls to a conjugate chamber plane, the conjugate chamber plane intersecting the conjugate inflection line and a conjugate chamber-center line, the conjugate chamber-center line defining a volumetric center of the conjugate chamber as determined with respect to a volume of the conjugate chamber located between the conjugate inflection line and a downstream end of the conjugate chamber, and the first one of the one or more conjugate chamber walls being located on a side of the conjugate chamber plane radially corresponding to the detonation chamber; and
wherein the first annular cross-sectional area corresponding to the detonation throat is less than the second annular cross-sectional area corresponding to the conjugate throat.

13. The combustion system of claim 12, wherein the first annular cross-sectional area is from 1% to 90% less than the second annular cross-sectional area.

14. The combustion system of claim 12, wherein the first divergent cone-half angle corresponds to a first annular region of the detonation nozzle radially proximal to the detonation chamber and radially distal to the deflagration chamber, and wherein the second divergent cone-half angle corresponds to a second annular region of the detonation nozzle radially proximal to the deflagration chamber and radially distal to the detonation chamber.

15. The combustion system of claim 14, wherein the first divergent cone-half angle is from 1 degree to 10 degrees, and/or wherein the second divergent cone-half angle is from 1 degree to 10 degrees.

16. The combustion system of claim 14, wherein the second divergent cone-half angle is from 10% to 200% greater than the first divergent cone-half angle.

17. The combustion system of claim 1, wherein the detonation combustor comprises a detonation fuel manifold coupled to or monolithically integrated with the one or more detonation chamber walls, the detonation fuel manifold configured to supply fuel and/or oxidizer to the detonation chamber.

18. The combustion system of claim 1, wherein the deflagration combustor comprises a plurality of deflagration fuel manifolds respectively configured to supply fuel and/or oxidizer to the deflagration chamber.

19. The combustion system of claim 1, wherein at least a portion of the detonation chamber circumferentially surrounds at least a portion of the deflagration chamber, or wherein at least a portion of the deflagration chamber circumferentially surrounds at least a portion of the detonation chamber.

20. The combustion system of claim 1, wherein at least a portion of the detonation chamber comprises a torus shape, and/or wherein at least a portion of the detonation chamber comprises parabolic annulus shape.

Patent History
Publication number: 20230280035
Type: Application
Filed: Mar 7, 2022
Publication Date: Sep 7, 2023
Inventors: Daniel Louis Depperschmidt (Saratoga Springs, NY), Kapil Kumar Singh (Rexford, NY), Aaron J. Glaser (Cincinnati, OH), Arin Elspeth Lastufka Cross (Waterford, NY), Sarah Marie Monahan (Latham, NY), Hannah Erin Bower (Niskayuna, NY)
Application Number: 17/687,873
Classifications
International Classification: F23R 3/16 (20060101); F23R 7/00 (20060101); F23R 3/52 (20060101); F23R 3/42 (20060101);