CN116717813A

CN116717813A - Bimodal combustion system

Info

Publication number: CN116717813A
Application number: CN202310204605.1A
Authority: CN
Inventors: 丹尼尔·路易斯·德佩施米特; 卡皮尔·库马尔·辛; 艾伦·J·格拉泽; 阿林·埃尔斯佩思·拉斯图夫卡·克劳斯; 莎拉·玛丽·莫纳汉; 汉纳·埃琳·鲍尔
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2022-03-07
Filing date: 2023-03-06
Publication date: 2023-09-08
Also published as: US20230280035A1

Abstract

A combustion system may include: a detonation combustor including one or more detonation chamber walls defining a detonation chamber; a deflagration burner comprising one or more deflagration chamber walls defining a deflagration chamber; and one or more conjugate chamber walls defining a conjugate chamber, wherein the conjugate chamber is in fluid communication with the detonation chamber and the deflagration chamber. The detonation chamber includes a detonation zone and a nozzle zone, wherein the nozzle zone provides fluid communication between the detonation zone and the conjugate chamber.

Description

Bimodal combustion system

Technical Field

The present disclosure relates generally to combustion systems for turbine engines, and methods of operating combustion systems for turbine engines.

Background

Combustion systems capable of operating under a wide range of operating conditions and thermal load requirements are of interest in the art, such as combustion systems that exhibit good operating performance (including good combustion efficiency, good fuel consumption, and/or low emissions). While combustion systems that perform deflagration remain an area of interest, the art has shown increased interest in detonation combustion processes. Accordingly, combustion systems that provide improved performance and/or are capable of operating over a wider range of operating conditions and thermal load requirements would be welcomed in the art.

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 including a dual-peak combustion system;

FIG. 1B illustrates a schematic cross-sectional view of an exemplary turbine engine including a dual-peak combustion system;

FIGS. 2A and 2B illustrate schematic cross-sectional views of an exemplary bimodal combustion system, respectively;

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

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

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

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

FIGS. 4A-4C illustrate schematic cross-sectional views of exemplary detonation chambers including a detonation region and a nozzle region, respectively;

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

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

FIGS. 6B and 6C illustrate schematic perspective views of an exemplary detonation fuel manifold;

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

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

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

Detailed Description

Reference will now be made in detail to the present embodiments of the disclosure, one or more examples of which are illustrated in the drawings. The present disclosure uses numerical and letter designations to refer to features in the drawings. Like or similar reference numerals have been used in the drawings and description 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 embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Moreover, all embodiments described herein are to be considered as exemplary unless expressly stated otherwise.

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

The terms "upper", "lower", "right", "left", "vertical", "horizontal", "top", "bottom", "lateral", "longitudinal", and the like should be related to the present disclosure as it is oriented in the drawings. However, it is to be understood that the present 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. Accordingly, 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, where forward refers to a position closer to the engine inlet and aft refers to a position closer to the engine nozzle or exhaust.

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

The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

Unless otherwise indicated herein, the terms "coupled," "fixed," "attached," and the like refer to both direct coupling, fixing, or attaching and indirect coupling, fixing, or attaching through one or more intermediate components or features.

Approximating language, as used herein throughout the specification and claims, may be 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, values modified by terms such as "about," "approximately," and "substantially" are not limited to the precise values specified. In at least some cases, the approximating language may correspond to the precision of an instrument for measuring the value or the precision of a method or machine for constructing or manufacturing a component and/or system. For example, approximating language may refer to being within a margin of 1%, 2%, 4%, 10%, 15%, or 20%.

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.

Furthermore, unless otherwise indicated, the terms "low", "high" or their respective comparison stages (e.g., lower, higher, where applicable) all refer to relative speeds within the engine. For example, a "low pressure turbine" operates at a pressure that is typically lower than a "high pressure turbine". Alternatively, the above terms may be understood at their highest level unless otherwise indicated. For example, a "low pressure turbine" may refer to the lowest maximum pressure turbine within the turbine section, while a "high pressure turbine" may refer to the highest maximum pressure turbine within the turbine section.

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

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

One or more components of the engine described herein may be manufactured or formed using any suitable process (e.g., an additive manufacturing process, such as a 3-D printing process). The use of such a process may allow such components to be integrally formed as a single unitary component, or as any suitable number of sub-components. In particular, the additive manufacturing process may allow such components to be integrally formed and include a variety of features that are not possible using existing manufacturing methods. For example, the additive manufacturing methods described herein may allow for the manufacture of channels, conduits, cavities, openings, housings, manifolds, double-walled, heat exchangers, or other components, or specific positioning and integration of such components, with unique features, configurations, thicknesses, materials, densities, fluid passages, headers, and mounting structures that are not possible or practical using existing manufacturing methods. Some of these features are described herein.

Suitable additive manufacturing techniques according to 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 by, for example, inkjet, laser spraying, and adhesive spraying, stereolithography (SLA), direct Selective Laser Sintering (DSLS), electron Beam Sintering (EBS), electron Beam Melting (EBM), laser Engineering Net Shape (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 making the structures provided herein as unitary, unitary structures include metal alloys, polymers, or ceramic powders. Exemplary metal powder materials are stainless steel alloys, cobalt-chromium alloys, aluminum alloys, titanium alloys, nickel-base superalloys, and cobalt-base superalloys. In addition, suitable alloys may include those designed to have good oxidation resistance, known as "superalloys," which have acceptable strength at elevated operating temperatures in turbine engines, such as Hastelloys,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, CM247 LC, C263, 718, X-850, ECY768, 282, X45, PWA 1483, and CMSX (e.g., CMSX-4) single crystal alloys. The fabricated objects of the present disclosure may be formed with one or more selected crystalline microstructures, such as directional solidification ("DS") or single crystal ("SX").

As used herein, the terms "integral," "unitary," or "integral" as used to describe a structure refer to a structure that is integrally formed from a continuous material or group of materials without seams, connecting joints, or the like. The unitary, single structure described herein may be formed by additive manufacturing to have the structure, or by a casting process, or the like.

The present disclosure generally provides combustion systems configured to perform deflagration combustion and detonation combustion, and engines including such combustion systems. Exemplary engines that may be configured to perform deflagration combustion and detonation combustion include turbine engines, rocket engines, ramjet engines, or combinations thereof, such as a turbine rocket engine, a turbine ramjet engine, or a rocket ramjet engine. Such combustion systems are generally referred to herein as dual-peak combustion systems. The presently disclosed bimodal combustion system can 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 detonation section includes a detonation chamber. The detonation chamber and the explosion chamber are in fluid communication with the conjugate chamber, respectively.

Combustion refers to an exothermic chemical reaction between a fuel and an oxidant that produces combustion products and heat through the conversion of chemicals. Engines may utilize thermal and kinetic energy generated by combustion to provide thrust. In general, combustion may be performed in one or both modes (knocking and knocking). As used herein, the term "deflagration" or "deflagration combustion" refers to combustion that may be thermodynamically described as approximately isobaric. In general, during the deflagration combustion process, the pressure of the combustion products slightly decreases, and the specific volume of the combustion products increases significantly, thereby generating combustion waves having subsonic velocity. For example, the combustion waves generated by the deflagration combustion process may have a velocity of several meters per second (m/s) (e.g., about 10m/s to about 200 m/s). As used herein, the term "detonation" or "detonation combustion" refers to combustion that may be thermodynamically described as approximately isovolumetric. In general, during a detonation combustion process, the pressure and temperature of the combustion products may suddenly rise and the specific volume may slightly drop, thereby generating a supersonic shockwave immediately preceding a combustion wave that also has a supersonic velocity. For example, the combustion waves generated by the detonation combustion process may have a speed of several kilometers per second (km/s) (e.g., about 1km/s to about 6 km/s).

Knock generally provides faster heat release, lower entropy increase, and higher thermal efficiency than knock. An exemplary detonation combustion process may provide a pressure increase of about 5 to about 20 times. In further contrast to knocking, knocking can propagate in lean fuel mixtures that result in relatively low NOx emissions. Detonation combustion has a higher thermodynamic efficiency than deflagration combustion, which translates into significantly improved specific impulse and/or specific fuel consumption. In some embodiments, a gas turbine engine utilizing detonation combustion may have a reduced number of compressor stages and/or reduced compressor pressure requirements due to, for example, the ability of detonation combustion to provide relatively large effective thrust at a relatively low overall compression ratio. Additionally or alternatively, detonation combustion may allow for engines with higher thrust to weight ratios, which may allow for smaller, lighter engines for a given mission requirement.

In an exemplary embodiment, the detonation section of the presently disclosed bimodal combustion system may be configured to perform rotary detonation combustion. The rotary detonation combustion process may separately generate shock waves before combustion waves that annularly propagate through a detonation region of the detonation chamber. The annularly propagating shock wave and combustion wave may be converted to longitudinal waves as the combustion products travel through a nozzle region of the detonation chamber in fluid communication with the conjugate chamber. The detonation occurring in the detonation chamber and/or the conjugate chamber may provide a back pressure that at least partially contributes to initiation and/or stability of the detonation reaction in the detonation chamber. Additionally or alternatively, the nozzle region of the detonation chamber may provide a back pressure that at least partially contributes to initiation and/or stability of the detonation reaction in the detonation chamber.

In some embodiments, a bimodal combustion system may perform detonation combustion during operating conditions that require relatively low thrust and detonation combustion during operating conditions that require relatively high thrust. The presently disclosed bimodal combustion system can detonate and detonate separately or simultaneously. For example, a bimodal combustion system can initiate and sustain detonation combustion consistent with ongoing and sustained detonation combustion. Additionally or alternatively, the dual-peak combustion system may stop detonation combustion while maintaining detonation combustion. Additionally or alternatively, the dual-peak combustion system may transition from knocking to knocking, and/or from knocking to knocking, depending on changing engine operating requirements.

In some embodiments, the presently disclosed bimodal combustion system may be configured to knock when the engine is operating at nominal speed and/or when the engine is operating at cruise speed. Additionally or alternatively, the presently disclosed bimodal combustion system may be configured to detonate when the engine is operating at nominal speed and/or when the engine is operating at cruise speed. Additionally or alternatively, the presently disclosed bimodal combustion system may be configured to initiate knock, for example, when the engine transitions from a nominal operating state to a high power operating state and/or a cruise operating state. Additionally or alternatively, the presently disclosed bimodal combustion system may be configured to stop detonation while maintaining 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 state and/or to a cruise operating state, and/or from a high power operating state to a cruise operating state. Additionally or alternatively, the presently disclosed bimodal combustion system may be configured to stop knocking while maintaining knocking, 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 cruise operating state to a nominal operating state.

As used herein, the term "rated speed" refers to the maximum output that an engine can achieve during normal operation. For turbine engines or other rotating machines, the rated speed refers to the maximum rotational speed that the engine can achieve during normal operation. For engines that do not include rotating machines (e.g., rocket engines), the rated speed refers to the speed at which the engine outputs thrust. Engines (e.g., turbine engines) for providing thrust for aircraft may operate at rated speeds during high power operating conditions (e.g., during takeoff operations and/or during aggressive over-the-air maneuvers).

As used herein, the term "nominal operating state" refers to an engine (e.g., a turbine engine) operating at a speed that is greater than the idle speed of the engine and less than the nominal speed of the engine. For example, the nominal operating state may include an operating speed that is at least 10% greater than the idle speed and at least 10% less than the nominal speed. As an example, the nominal operating state may include a cruise speed.

As used herein, the term "cruise speed" refers to a period of time during which the output of the engine is operating at a relatively high operating speed. For example, turbine engines used to power aircraft may operate at cruising speeds when the aircraft is level after climbing to a particular altitude. In some embodiments, the engine (e.g., turbine engine) may be operated at a cruising speed of about 50% to about 90% of the nominal speed (e.g., about 70% to about 80% of the nominal speed). In some embodiments, cruise speed may be achieved at about 80% of full throttle (e.g., about 50% to about 90% of full throttle, such as about 70% to about 80% of full throttle).

As used herein, the term "low power operating state" refers to an engine (e.g., a turbine engine) operating at a speed that is at least 10% higher than the idle speed of the engine.

As used herein, the term "high power operating state" refers to an engine (e.g., turbine engine) operating at a rotational speed of at least 90% of the rated speed of the engine.

Exemplary embodiments of the present disclosure will now be described in more detail. Referring to fig. 1A and 1B, an exemplary engine 50 including a dual-peak 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 including a bimodal combustion system 200, such as a turbine engine, a rocket engine, a ramjet engine, or a combination thereof, such as a turbine rocket engine, a turbine ramjet engine, or a rocket ramjet engine. As an example, FIG. 1B illustrates an exemplary turbine engine 100 including a dual-peak combustion system 200. Exemplary engine 50 (e.g., turbine engine 100) may be mounted to an aircraft, for example, in an under-wing configuration or a tail mounted configuration. The turbine engine 100 shown in FIG. 1B is provided by way of example and not limitation, and the subject matter of this disclosure may be implemented with other suitable types of engines 50 (including other suitable turbine engines 100).

For example, as shown in FIG. 1A, an exemplary engine 50 may include an inlet section 52, a burner section 54, and an outlet section 56 in serial flow relationship. The engine 50 may include an engine housing 58 that contains and/or defines at least a portion of the inlet section 52, the burner section 54, and/or the outlet section 56. The inlet section 52 may generally direct a flow of an oxidant 60 (such as air or gas) to the combustor section 54. The inlet section 52 may compress the flow of oxidant 60 before the flow of oxidant 60 enters the combustor section 54. For example, the inlet section 52 may define a reduced cross-sectional area downstream to the combustor section 54. At least a portion of the total oxidant 60 flow may be mixed with fuel 62 and may react in a combustion process to produce combustion products 64. The combustor section 54 may include a bimodal combustion system 200 constructed in accordance with the present disclosure. Bimodal combustion system 200 may include a detonation section 202 configured to perform detonation combustion and a detonation section 204 configured to perform detonation 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 before 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 increased cross-sectional area that leads downstream of the combustor section 54 and/or downstream of the turbine section 66. In some embodiments, the turbine section 66 may define a portion of the outlet section 56. Additionally or alternatively, the outlet section 56 may include an outlet nozzle 68 or the like. Expansion of the combustion products 64 generally provides thrust, such as by rotation of the turbine section 66, which may be used as a direct power output in the form of thrust, and/or to generate mechanical energy.

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 of the fan section 102. The fan section 102 may include a fan 106 having any suitable configuration (e.g., 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 in a generally 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 casing 58, with the engine casing 58 surrounding one or more portions of the core engine 104, including the compressor section 114, the combustor section 54, and the turbine section 66. The engine housing 58 may define a core engine inlet 118, an outlet nozzle 68, and a core air flow path 122 therebetween. The core air flow path 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 each include one or more compressor stages. As an example, the compressor section 114 (including the LP compressor 124 and/or the HP compressor 126) may have 1 to 16 compressor stages, such as 1 to 12 stages, such as 1 to 10 stages, such as 1 to 8 stages, such as 1 to 6 stages, or such as 1 to 4 stages, respectively. 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, the combustor section 54, the turbine section 66, and the outlet nozzle 68 may be arranged in serial flow relationship and may each define a portion of a core air flow path 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 wind sector segment 102 may be coupled to a shaft driven by the core engine 104. As an 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.HP shaft 132 may drivingly connect HP turbine 128 to HP compressor 126, and LP shaft 134 may drivingly connect LP turbine 130 to LP compressor 124. In other embodiments, for example in the case of a turbine engine including an intermediate pressure turbine, the turbine engine may have three shafts. The shaft of the core engine 104 along with the rotating portion of the core engine 104 may sometimes be referred to as a "spool". HP shaft 132, the rotating portion of HP compressor 126 coupled to HP shaft 132, and the rotating portion of HP turbine 128 coupled to HP shaft 132 may be collectively referred to as a High Pressure (HP) spool 136. The LP shaft 134, as well as the rotating portion of the LP compressor 124 coupled to the LP shaft 134, the rotating portion of the LP turbine 130 coupled to the LP shaft 134, may be collectively referred to as a Low Pressure (LP) spool 138.

In some embodiments, the fan section may be directly coupled to the shaft of the core engine, such as directly coupled to the LP shaft. Alternatively, as shown in fig. 1B, the fan section 102 and the core engine 104 may be coupled to each other through a power gearbox 140 (e.g., a planetary reduction gearbox, an epicyclic gearbox, etc.). For example, the power gearbox 140 may couple the LP shaft 134 to a fan 106, such as to a fan tray 110 of the fan section 102. The power gearbox 140 may include a plurality of gears for reducing the rotational speed of the LP shaft 134 to a more efficient rotational speed of the fan section 102.

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

During operation of turbine engine 100, inlet airflow 150 enters turbine engine 100 through an inlet 152 defined by nacelle 146 (e.g., a nose mask of nacelle 146). In some embodiments, inlet section 52 (fig. 1A) may include at least a portion of inlet 152, at least a portion of nacelle 146, and/or at least a portion of fan housing 142. Inlet airflow 150 passes through fan blades 108. The inlet airflow 150 splits into a core airflow 154 flowing into and through the core air flow path 122 of the core engine 104 and a bypass airflow 156 flowing through the bypass passage 148. Core gas stream 154 is compressed by 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 flow 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 bypass airflow 156 to flow through bypass passage 148, generating propulsive thrust. The core airflow exiting the outlet nozzle 68 generates additional thrust.

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

As schematically depicted in fig. 1B, in some embodiments, a turbine engine 100 including a bimodal combustion system 200 may be configured with a compressor section 114, the compressor section 114 having a relatively smaller number of compressor stages than other turbine engines 100 that generate a comparable amount of thrust. Additionally or alternatively, the turbine engine 100 including the bimodal combustion system 200 may be configured with a compressor section 114, the compressor section 114 having a relatively short axial length compared to other turbine engines 100 that generate a comparable amount of thrust. In some embodiments, the turbine engine 100 including the bimodal combustion system 200 may be configured without an HP compressor and/or without an LP compressor. As a result, a relatively large proportion of the energy generated by turbine engine 100 may be distributed to thrust forces, e.g., as opposed to the incoming oxidant used in compression combustor section 54. For example, a relatively large proportion of the energy generated by turbine engine 100 may be distributed to thrust generated by rotating fan 106 to move bypass airflow 156 through bypass passage 148 and/or by exhausting core airflow through outlet nozzle 68.

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

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

In some embodiments, the turbine engine 100 including the bimodal combustion system 200 may exhibit a bypass ratio of about 10:1 to about 20:1 (e.g., about 10:1 to about 12:1, such as about 12:1 to about 16:1, such as about 16:1 to about 18:1, or such as about 18:1 to about 20:1), for example, as determined at rated speed and/or cruise speed. As used herein, the term "bypass ratio" refers to the ratio of the mass flow rate through the bypass passage 148 to the mass flow rate of the core air flow path 122 of the core engine 104. In some embodiments, turbine engine 100 including bimodal combustion system 200 and exhibiting the bypass ratio described above may have the total compressor ratio described above at rated speeds (e.g., at takeoff). Additionally or alternatively, turbine engine 100 that generates the amount of thrust described above may exhibit the bypass ratio described above.

In some embodiments, the turbine engine 100 including the bimodal combustion system 200 may have a thrust to weight ratio of about 6.0 to about 9.0 (e.g., about 6.0 to about 7.0, or, for example, about 7.0 to about 8.0, or, for example, about 8.0 to about 9.0). In some embodiments, turbine engine 100 including bimodal combustion system 200 and exhibiting the aforementioned bypass ratio may have the aforementioned thrust-to-weight ratio. Additionally or alternatively, turbine engine 100 having the overall compressor ratio described above may have the thrust-to-weight ratio described above.

In some embodiments, the turbine engine 100 including the bimodal combustion system 200 may have a thrust ratio fuel consumption of about 8 g/kilonewton-seconds (g/kN-s) to about 14 g/kN-s (e.g., about 8 g/kN-s to about 12 g/kN-s, such as about 8 g/kN-s to about 10 g/kN-s, such as about 10 g/kN-s to about 12 g/kN-s, or such as about 12 g/kN-s to about 14 g/kN-s), for example, at rated speed and/or at cruise speed (e.g., at 80% cruise speed at full throttle). In some embodiments, turbine engine 100 including bimodal combustion system 200 and exhibiting the aforementioned bypass ratio, the aforementioned total compressor ratio, and/or the aforementioned thrust-to-weight ratio may have the aforementioned thrust-to-fuel consumption. As an example, a turbine engine 100 generating about 400kN to about 600kN of thrust and/or a turbine engine 100 generating about 10kN to about 300kN of thrust may have the above-described thrust versus fuel consumption, for example, at rated or cruise speeds.

Referring now to fig. 2A and 2B, an exemplary dual-peak combustion system 200 is further described. As shown, bimodal combustion system 200 may include detonation section 202, detonation section 204, and conjugated 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 occur during operation of the detonation section 202 of the bimodal combustion system 200. The detonation combustion may include pulse detonation combustion and/or rotary detonation combustion.

Detonation combustor 206 may include one or more detonation fuel manifolds 212 configured to supply fuel 62 and/or oxidant 60 to detonation chambers 210. The fuel 62 and the oxidant 60 may be mixed within one or more detonation fuel manifolds 212. Additionally or alternatively, fuel 62 and oxidant 60 may be mixed upstream of one or more detonation fuel manifolds 212 and/or within detonation chamber 210. One or more detonation fuel manifolds 212 may define a portion of the detonation chamber walls 208. Additionally or alternatively, detonation fuel manifold 212 may be coupled to one or more detonation chamber walls 208. In some embodiments, detonation fuel manifold 212 may be monolithically integrated with one or more detonation chamber walls 208. One or more detonation chamber walls 208 and detonation fuel manifold 212 may define a single integral component.

Detonation section 202 may include one or more detonation fuel supply lines 214 in fluid communication with detonation fuel manifold 212. One or more detonation fuel supply lines 214 may be coupled to the detonation fuel manifold 212. Additionally or alternatively, one or more detonation fuel supply lines 214 may be defined at least in part by detonation fuel manifold 212 (e.g., by the unitary structure of detonation fuel manifold 212). Detonation section 202 may include one or more detonation oxidant supply lines 216 in fluid communication with detonation fuel manifold 212. One or more detonation oxidant supply lines 216 may be coupled to the detonation fuel manifold 212. Additionally or alternatively, one or more detonation oxidant supply lines 216 may be defined at least in part by the detonation fuel manifold 212 (e.g., by the unitary structure of the detonation fuel manifold 212).

One or more detonation fuel manifolds 212 may include a plurality of detonation fuel orifices 218 to provide fluid communication between detonation fuel supply lines 214 and detonation chambers 210 and/or between detonation oxidant supply lines 216 and detonation chambers 210. A plurality of detonation fuel orifices 218 may provide fuel 62 and/or oxidant 60, respectively, to detonation chamber 210. The plurality of detonation fuel orifices 218 may be defined at least in part by the respective detonation fuel manifold 212 (e.g., by the unitary structure of the respective detonation fuel manifold 212). Additionally or alternatively, a plurality of detonation fuel orifices 218 may be coupled to respective detonation fuel manifolds 212. One or more detonation fuel supply lines 214 and/or one or more detonation oxidant supply lines 216 may be in fluid communication with a corresponding detonation fuel manifold 226 and/or with a corresponding plurality of detonation fuel orifices 218. The fuel 62 and the oxidant 60 may be mixed with each other within the respective detonation fuel manifold 212 (e.g., within the corresponding plurality of detonation fuel orifices 218). Additionally or alternatively, the fuel 62 and the oxidant 60 may be mixed with each other upstream of the detonation fuel manifold 212 and/or within the detonation chamber 210.

Deflagration section 204 includes deflagration burner 220. Deflagration combustor 220 may include one or more deflagration chamber walls 222 defining a deflagration chamber 224 within which deflagration combustion may occur during operation of deflagration section 204 of bimodal combustion system 200. Deflagration combustor 220 may include one or more deflagration chambers 224. Deflagration combustor 220 may include a plurality of deflagration fuel manifolds 226 configured to supply fuel 62 and/or oxidant 60 to one or more deflagration chambers 224. In some embodiments, a plurality of deflagration fuel manifolds 226 may supply fuel 62 and/or oxidant 60 to substantially annular deflagration chamber 224. Additionally or alternatively, a plurality of deflagration fuel manifolds 226 may supply fuel 62 and/or oxidant 60, respectively, to a plurality of generally cylindrical deflagration chambers 224. In some embodiments, explosion chamber 224 may include a plurality of generally cylindrical regions in fluid communication with a generally annular region. Fuel 62 and oxidant 60 may be mixed within respective detonative fuel manifolds 226, upstream of respective detonative fuel manifolds 226, and/or within explosion chambers 224. A plurality of deflagration fuel manifolds 226 may each define a portion of deflagration chamber walls 222. Additionally or alternatively, one or more detonation chamber walls 222 may be coupled to a detonation fuel manifold 226. In some embodiments, one or more of the detonation chamber walls 222 may be monolithically integrated with the detonation fuel manifold 226. For example, deflagration chamber wall 222 and deflagration fuel manifold 226 may define a single integral component.

Deflagration section 204 may include one or more deflagration fuel supply lines 228 in fluid communication with respective ones of a plurality of deflagration fuel manifolds 226. One or more detonation fuel supply lines 228 may be coupled to respective detonation fuel manifolds 226. Additionally or alternatively, one or more of the detonation fuel supply lines 228 may be defined at least in part by the respective detonation fuel manifolds 226 (e.g., by the unitary structure of the respective detonation fuel manifolds 226). The deflagration fuel supply line 228 may provide fuel 62 to a plurality of deflagration fuel manifolds 226. Deflagration section 204 may include one or more deflagration oxidant supply lines 230 in fluid communication with respective ones of the plurality of deflagration fuel manifolds 226. One or more deflagration oxidant supply lines 230 may be coupled to respective deflagration fuel manifolds 226. Additionally or alternatively, one or more deflagration oxidant supply lines 230 may be defined at least in part by respective deflagration fuel manifolds 226 (e.g., by the unitary structure of respective deflagration fuel manifolds 226). Deflagration oxidant supply line 230 may provide oxidant 60 to a 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, the one or more deflagration fuel injectors 232 providing fluid communication between deflagration fuel supply line 228 and deflagration chamber 224 and/or between deflagration oxidant supply line 230 and deflagration chamber 224. Deflagration fuel injector 232 may provide fuel 62 and/or oxidant 60 to explosion chamber 224. One or more detonation fuel injectors 232 are coupled to respective detonation fuel manifolds 226. Additionally or alternatively, one or more of the detonation fuel injectors 232 may be defined at least in part by the respective detonation fuel manifolds 226 (e.g., by the unitary structure of the respective detonation fuel manifolds 226). One or more deflagration fuel supply lines 228 and/or one or more deflagration oxidant supply lines 230 may be in fluid communication with corresponding deflagration fuel manifolds 226 and/or with corresponding deflagration fuel injectors 232. Fuel 62 and oxidant 60 may be mixed with each other within a plurality of deflagration fuel manifolds 226 (e.g., within respective deflagration fuel injectors 232). Additionally or alternatively, fuel 62 and oxidant 60 may be mixed with each other upstream of respective detonative fuel manifolds 226 and/or within detonative chambers 224.

Still referring to fig. 2A and 2B, the oxidant from the inlet section 52 may flow to the burner section 54 through a plurality of fluid paths 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 define an upstream portion of the combustor section 54. The inward combustor casing 238 may define a radially inward portion of the combustor section 54. The plurality of fluid paths 234 may include a plurality of diffusers 240. The plurality of diffusers 240 may be configured to reduce the velocity of the oxidant 60 flowing from the inlet section 52 and/or to increase the static pressure of the oxidant 60 flowing into the combustor section 54. The static pressure of oxidizer 60 may be increased to at least a portion of a pressure level sufficient to support detonation in detonation chamber 210. Additionally or alternatively, 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 paths 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, the combustor inlet plenum 242 being in fluid communication with the plurality of fluid paths 234 and one or more detonation oxidant supply lines 216 corresponding to the respective detonation fuel manifolds 212. Additionally or alternatively, the respective detonation oxidant supply lines 216 may be in direct fluid communication with the plurality of fluid paths 234. Additionally or alternatively to providing fluid communication between the inlet section 52 and one or more detonation fuel manifolds 212, a plurality of fluid paths 234 may provide fluid communication between the inlet section 52 and a plurality of detonation fuel manifolds 226. In some embodiments, combustor inlet plenum 242 may be in fluid communication with a plurality of deflagration fuel manifolds 226. Additionally or alternatively, the plurality of fluid paths 234 may include a plurality of deflagration oxidant supply lines 230 corresponding to respective deflagration fuel manifolds 226.

In some embodiments, additional oxidant 60 and/or additional fuel 62 may be introduced into detonation chamber 210, for example, through one or more detonation chamber dilution paths 207. One or more detonation chamber dilution paths 207 may be defined at least in part by corresponding detonation chamber walls 208. The one or more detonation chamber dilution pathways 207 may also be configured to provide back pressure to the detonation chamber 210 and/or enhance the equivalence ratio within the detonation chamber 210. Additionally or alternatively, in some embodiments, additional oxidant 60 and/or additional fuel 62 may be introduced into explosion chamber 224, for example, through one or more explosion chamber dilution paths 209. One or more explosion chamber dilution paths 209 may be defined at least in part by corresponding explosion chamber walls 222. One or more of chamber dilution paths 209 may also be configured to provide back pressure to chamber 224 and/or enhance the equivalence ratio within chamber 224.

As shown in fig. 2A and 2B, the bimodal combustion system 200 may include a conjugated section 205. Conjugate section 205 may include conjugate burner 211. The conjugate section 205 and/or the conjugate burner 211 may include one or more conjugate chamber walls 244 defining a conjugate chamber 246. Detonation chamber 210 and the explosion chamber may each be in fluid communication with conjugate chamber 246. Detonation chamber 210 and detonation chamber 224 may each have a generally annular configuration. Knock chamber 210 and explosion chamber 224 may have a co-annular arrangement with respect to longitudinal axis 248 of engine 50. The conjugate chamber 246 may be annularly disposed relative to a longitudinal axis 248 of the engine 50. Combustion products 64 from detonation chamber 210 and/or explosion chamber 224 may flow to conjugate chamber 246.

The combustion products 64 generated in the detonation chamber 210 and/or flowing from the detonation chamber 210 may sometimes be referred to as detonation combustion products 64. Combustion products 64 generated in explosion chamber 224 and/or flowing from explosion chamber 224 may sometimes be referred to as explosion combustion products 64. Detonation combustion products 64 and deflagration combustion products 64 may mix with each other in conjugate chamber 246. In some embodiments, further combustion may occur within conjugate chamber 246, such as detonation and/or detonation. The combustion products 64 generated in the conjugate chamber 246 and/or exiting the conjugate chamber 246 may sometimes be referred to as conjugate combustion products 64. Additionally or alternatively, combustion may be substantially completed within detonation chamber 210 and/or detonation chamber 224, respectively. In some embodiments, additional oxidant 60 and/or additional fuel 62 may be introduced into conjugate chamber 246, for example, through one or more conjugate chamber dilution paths 213. One or more conjugate chamber dilution paths 213 may be defined at least in part by corresponding conjugate chamber walls 244. One or more conjugate chamber dilution paths 213 may also be configured to provide back pressure to conjugate chamber 246, detonation chamber 210, and/or detonation chamber 224, and/or to enhance the equivalence ratio within conjugate chamber 246, detonation chamber 210, and/or detonation chamber 224.

In some embodiments, as shown in FIG. 2A, detonation chamber 210 may have a radially outward arrangement relative to detonation chamber 224. Explosion chamber 224 may have a radially inward disposed relative to explosion chamber 210. FIG. 3A illustrates a cross-sectional view of the bimodal combustion system 200 of FIG. 2A at a location along a longitudinal axis 248 denoted by "A". FIG. 3B illustrates a cross-sectional view of the bimodal combustion system 200 of FIG. 2A at a location along a longitudinal axis 248 denoted by "B". As shown in fig. 2A and fig. 3A and 3B, at least a portion of detonation chamber 210 may circumferentially surround at least a portion of detonation chamber 224.

In some embodiments, as shown in FIG. 2B, explosion chamber 224 may have a radially outward arrangement relative to detonation chamber 210. Detonation chamber 210 may have a radially inward arrangement relative to detonation chamber 224. FIG. 3D illustrates a cross-sectional view of the bimodal combustion system 200 of FIG. 2B at a location along a longitudinal axis 248 denoted by "D". FIG. 3E shows a cross-sectional view of the bimodal combustion system 200 of FIG. 2B at a location along a longitudinal axis 248 denoted by "E". As shown in fig. 2B and fig. 3D and 3E, at least a portion of explosion chamber 224 may circumferentially surround at least a portion of explosion chamber 210.

As shown in fig. 2A and 2B and fig. 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 explosion chamber walls 222 defining explosion chamber 224 may include an outer explosion chamber wall 254 and an inner explosion chamber wall 256. At least a portion of the outer detonation chamber wall 254 may circumferentially surround at least a portion of the inner detonation chamber wall 256. In some embodiments, dual-peak combustion system 200 may include a combustor inlet plenum 242 disposed between at least a portion of detonation chamber 210 and at least a portion of detonation chamber 224 relative to a radial axis 258 of engine 50.

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

In some embodiments, outer detonation chamber wall 254 and inner detonation chamber wall 256 may be integrally formed with each other. Alternatively, the outer and inner detonation chamber walls 254, 256 may be coupled to each other, for example, by welding, attachment hardware (e.g., bolts), and the like. In some embodiments, as shown in FIGS. 2A and 2B, one or more of the detonation chamber walls 222 may include a post-detonation chamber wall 262. Rear explosion chamber wall 262 may define a rear region of explosion chamber 224. Post-detonation chamber wall 262 may have a generally annular configuration. For example, post-detonation chamber walls 262 may be annularly disposed between outer detonation chamber walls 254 and inner detonation chamber walls 256. Post-detonation chamber walls 262 may be integrally formed with outer detonation chamber walls 254 and/or with inner detonation chamber walls 256. Additionally or alternatively, post-detonation chamber wall 262 may be coupled to outer detonation chamber wall 254 and/or inner detonation chamber wall 256, such as by welding, attachment hardware (e.g., bolts), and the like.

In some embodiments, the inner detonation chamber wall 252 may adjoin and/or adjoin the outer detonation chamber wall 254. In some embodiments, inner and outer detonation chamber walls 252, 254 may be integrally formed with each other. Alternatively, the inner and outer detonation chamber walls 252, 254 may be coupled to each other, for example, by welding, attachment hardware (e.g., bolts), and the like. In some embodiments, at least a portion of explosion chamber 224 may circumferentially surround at least a portion of combustor inlet plenum 242. Additionally or alternatively, at least a portion of explosion chamber 224 may have a generally cylindrical configuration.

In some embodiments, the inner detonation chamber wall 256 may adjoin and/or adjoin the outer detonation chamber wall 250. In some embodiments, the inner detonation chamber wall 256 and the outer detonation chamber wall 250 may be integrally formed with each other. Alternatively, the inner and outer detonation chamber walls 256, 250 may be coupled to each other, for example, by welding, attachment hardware (e.g., bolts), and the like. In some embodiments, at least a portion of detonation chamber 210 may circumferentially surround a portion of combustor inlet plenum 242. Additionally or alternatively, at least a portion of detonation chamber 210 may have a generally cylindrical configuration.

Still referring to fig. 2A and 2B, and with further reference to fig. 3A-3F, in some embodiments, 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. Knocking of fuel 62 and/or oxidant 60 (e.g., a mixture of fuel 62 and oxidant 60) may occur within a knock region 264 of knock chamber 210. Combustion products 64 resulting from the combustion of fuel 62 and oxidant 60 may flow through nozzle region 266 of detonation chamber 210 to conjugate chamber 246. The 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 then through the outlet nozzle 68 of the engine.

The detonation region 264 and the nozzle region 266 of the detonation chamber 210 may each have a generally annular configuration, e.g., have an annular cross-sectional area. Knock region 264 of knock chamber 210 may include a region of knock chamber 210 in which knocking occurs during operation of knock section 202 of bimodal combustion system 200. During operation of detonation zone 202 within the operating range for which detonation zone 202 may be configured, the detonation reaction within detonation zone 264 may be stable. Detonation section 202 may include detonation nozzles 268. Detonation nozzles 268 may be defined by one or more detonation chamber walls 208. Detonation nozzles 268 may be located at nozzle region 266 of detonation chamber 210. Detonation nozzle 268 may include a detonation throat 270. The detonation throats 270 may define locations of the detonation nozzles 268 that have an annular cross-sectional area with a minimum annular ring width relative to adjacent portions of the detonation nozzles 268. Detonation nozzle 268 may include a converging portion having a decreasing cross-sectional area upstream of detonation throat 270 in a direction from detonation region 264 to detonation throat 270 of detonation nozzle 268. Detonation nozzle 268 may include a diverging portion having an increasing cross-sectional area downstream of detonation throat 270 in a direction from detonation throat 270 toward conjugate chamber 246.

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

Explosion chamber 224 and explosion chamber 210 may be at least partially defined from one another by a conjugate inflection line (inflection line) 272. Additionally or alternatively, detonation chamber wall 208 and detonation chamber wall 222 may be at least partially delineated from each other by conjugate inflection line 272. Conjugate inflection line 272 may define a linear inflection (linear inflection) oriented circumferentially relative to longitudinal axis 248 of engine 50 that represents a tangent to the forefront bevel or forefront curve, delineating detonation chamber wall 208 and detonation chamber wall 222 from each other. For example, as shown in FIG. 2A, the conjugate inflection line 272 may define a linear inflection between the inner and outer detonation chamber walls 252, 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 detonation chamber wall 256.

In some embodiments, detonation section 202 may include conjugate chamber walls 244 disposed between detonation chamber walls 208 and 222. Detonation chamber walls 208 and/or detonation chamber walls 222 may be integrally formed with conjugate chamber walls 244 disposed therebetween. Additionally or alternatively, detonation chamber wall 208 and/or detonation chamber wall 222 may be coupled to conjugate chamber wall 244 disposed therebetween, such as by welding, attachment hardware (e.g., bolts), and the like. In some embodiments, conjugate inflection line 272 may be defined by conjugate chamber wall 244 disposed between such detonation chamber wall 208 and detonation chamber wall 222. For example, the conjugate break line 272 may define a linear break oriented circumferentially relative to the longitudinal axis 248 of the engine 50 that represents a forward-most oblique angle of the conjugate chamber wall 244 or a tangent to the forward-most curve of the conjugate chamber wall 244. Alternatively, the conjugate turn line 272 may define a linear turn oriented circumferentially relative to the longitudinal axis 248 of the engine 50 that represents the final oblique angle of the conjugate chamber wall 244 or a tangent to the final curve of the conjugate chamber wall 244.

Explosion chamber 224 and explosion chamber 210 may each be located on opposite sides of conjugate break line 272. The conjugate turning line 272 may have a generally elliptical or circular shape. The conjugate turning line 272 may circumferentially surround the longitudinal axis 248 of the engine 50. As shown, for example, in fig. 3C and 3F, the conjugate chamber 246 may have a conjugate chamber centerline 274 that defines an annular centerline of the conjugate chamber 246. The conjugate chamber centerline 274 may be determined relative to the volume of the conjugate chamber 246 between the conjugate turning line 272 and the downstream end of the conjugate chamber 246. Conjugate chamber centerline 274 may define the volumetric center of conjugate chamber 246. Conjugate chamber centerline 274 may have a generally elliptical or circular shape. The conjugate chamber centerline 274 may circumferentially surround the longitudinal axis 248 of the engine 50.

A conjugate chamber plane 276 labeled "C" may intersect the conjugate turn line 272 and the conjugate chamber centerline 274 of the conjugate chamber 246. Conjugate chamber plane 276 may circumferentially surround longitudinal axis 248 of engine 50. The conjugate chamber plane 276 may have a generally linear configuration along a conjugate chamber centerline 278, the conjugate chamber centerline 278 intersecting the conjugate break line 272 and the conjugate chamber centerline 274 in an orientation parallel to the conjugate chamber plane 276. As an example, the conjugate chamber plane 276 may have a cylindrical configuration or a frustoconical configuration.

The downstream end of the conjugate chamber 246 may be defined by the structural context of the engine 50. In some embodiments, the downstream end of the conjugate chamber 246 may be defined by a shroud (not shown) (e.g., a combustor exhaust shroud and/or a turbine inlet shroud). Such shrouds may direct the flow of combustion products 64 circumferentially and/or helically into a turbine section 66 (fig. 1A and 1B) of the engine 50. Alternatively, such a shroud may direct the flow of combustion products 64 circumferentially and/or helically into the outlet section 56 (fig. 1A and 1B) of the engine 50. Additionally or alternatively, in some embodiments, the downstream end of the conjugate chamber 246 may be defined by a location of the conjugate chamber 246 having 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, conjugate section 205 may include conjugate nozzle 280. The conjugate nozzle 280 may be defined by one or more conjugate chamber walls 244. In some embodiments, detonation chamber 210 may be at least partially defined with conjugate chamber 246 by conjugate nozzle 280. Conjugate nozzle 280 may include a conjugate throat 282. At least a portion of conjugate nozzle 280 and/or conjugate throat 282 may be adjacent to detonation chamber 210. Conjugate throat 282 may define a location of conjugate nozzle 280 having an annular cross-sectional area perpendicular to conjugate chamber plane 276 that has a minimum annular ring width relative to an adjacent portion of conjugate nozzle 280. The annular cross-sectional area of conjugate nozzle 280 corresponding to conjugate throat 282 may extend from conjugate chamber plane 276 to conjugate chamber wall 244 on a side of conjugate chamber plane 276 radially corresponding to detonation chamber 210. For example, fig. 2A shows conjugate nozzle 280 and conjugate throat 282 located radially outward of conjugate chamber plane 276, and fig. 2B shows conjugate nozzle 280 and conjugate throat 282 located radially inward of conjugate chamber plane 276. Such conjugate chamber walls 244 defining an annular cross-sectional area corresponding to the conjugate throat 282 may be contiguous, and/or integrally formed with the conjugate nozzle 280. Such conjugate chamber walls 244 defining an annular cross-sectional area corresponding to the conjugate throat 282 may be longitudinally adjacent to the detonation chamber walls 208, e.g., contiguous, abutting, and/or integrally integrated with such detonation chamber walls 208. Additionally or alternatively, the conjugate throat 282 may define the conjugate chamber wall 244 and the detonation chamber wall 208. Conjugate nozzle 280 may include a converging portion having a decreasing cross-sectional area upstream of conjugate throat 282 in a direction from detonation throat 270 of detonation nozzle 268 toward conjugate throat 282. In some embodiments, the diverging portion of detonation nozzle 268 may define at least a portion of the converging portion of conjugated nozzle 280. Conjugate nozzle 280 may include a diverging portion having an increasing cross-sectional area downstream of conjugate throat 282 in a direction from conjugate throat 282 toward the downstream end of conjugate chamber 246.

In some embodiments, detonation section 204 may include a detonation nozzle 284. Detonation nozzles 284 may be defined by one or more detonation chamber walls 222. Additionally or alternatively, explosion chamber 224 may be at least partially defined by explosion nozzle 284 and conjugate chamber 246. The detonation nozzle 284 may include a detonation throat 286. At least a portion of knock nozzle 284 and/or knock throat 286 may be adjacent to explosion chamber 224. The detonation throats 286 may define the location of the detonation nozzles 284 with an annular cross-sectional area perpendicular to the conjugate chamber plane 276 having a minimum annular ring width relative to adjacent portions of the detonation nozzles 284. The annular cross-sectional area corresponding to the detonation throat 286 may extend from the conjugate chamber plane 276 to a radial direction of the conjugate chamber plane 276 corresponding to the conjugate chamber wall 244 on one side of the detonation chamber 224. For example, FIG. 2A shows detonation nozzles 284 and detonation throats 286 located radially inward of conjugate chamber plane 276, and FIG. 2B shows detonation nozzles 284 and detonation throats 286 located radially outward of conjugate chamber plane 276. Such conjugate chamber walls 244 defining an annular cross-sectional area corresponding to the detonation throats 286 may be contiguous, abutting, and/or integrally formed with the detonation nozzles 284. Such conjugate chamber walls 244 defining an annular cross-sectional area corresponding to the deflagration throat 286 may be longitudinally adjacent to the deflagration chamber walls 222, e.g., contiguous, abutting, and/or integrally integrated with such deflagration chamber walls 222. Additionally or alternatively, the deflagration throat 286 may define the deflagration chamber wall 222 and the conjugate chamber wall 244. Explosion nozzle 284 may include a converging portion having a decreasing cross-sectional area upstream of explosion throat 286 in a direction from explosion chamber 224 toward explosion throat 286. The detonation nozzle 284 may include a diverging portion having an increasing cross-sectional area downstream of the detonation throat 286 in a direction from the detonation throat 286 toward the downstream end of the conjugate chamber 246.

Referring still to fig. 2A and 2B and 3A-3F, in some embodiments, detonation nozzle 268 and/or detonation throat 270 may be at least partially constructed and arranged to provide a suitable back pressure in detonation chamber 210, for example, to initiate and/or maintain detonation in detonation region 264 of detonation chamber 210. For example, stable detonation may occur within detonation chamber 210 (e.g., in detonation region 264 of detonation chamber 210) at least during operation of detonation section 202 within the operating range for which detonation section 202 may be configured. Additionally or alternatively, detonation nozzle 268 and/or detonation throat 270 may be at least partially constructed and arranged to accelerate combustion products 64 flowing from detonation chamber 210 to conjugate chamber 246. In some embodiments, conjugate nozzle 280 and/or conjugate throat 282 may be at least partially configured to provide a suitable back pressure in detonation chamber 210, for example, to initiate and/or maintain detonation at least during operation of detonation section 202 within an operating range for which detonation section 202 may be configured, such as in detonation region 264 of detonation chamber 210. Additionally or alternatively, conjugate nozzle 280 and/or conjugate throat 282 may be at least partially constructed and arranged to accelerate combustion products 64 flowing from detonation chamber 210 to conjugate chamber 246. In some embodiments, detonation nozzle 284 and/or detonation throat 286 may be configured, at least in part, to provide a suitable back pressure in detonation chamber 210, for example, to initiate and/or maintain detonation at least during operation of detonation section 202 within an operating range for which detonation section 202 may be configured, such as in detonation region 264 of detonation chamber 210. In some embodiments, the detonation within the detonation chamber 224 may provide a backpressure within the detonation chamber 210 that is suitable for initiating and/or maintaining detonation, 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 alternatively, knock nozzle 284 and/or knock throat 286 may be at least partially constructed and arranged to accelerate combustion products 64 flowing from knock chamber 224 to conjugation chamber 246.

In some embodiments, detonation nozzle 268 and/or conjugated nozzle 280 may be configured and arranged in the form of a de Laval (de Laval) type nozzle. For example, detonation nozzle 268 and/or conjugated nozzle 280 may be individually or collectively configured as a de laval nozzle. In some embodiments, detonation nozzle 268 and/or conjugated nozzle 280 may be at least partially constructed and arranged to reduce the static pressure of combustion products 64 flowing from detonation chamber 210 to conjugated chamber 246. The static pressure of conjugate chamber 246 and/or the static pressure of explosion chamber 224 may be less than the static pressure of explosion chamber 210. The static pressure of the combustion products 64 flowing from detonation chamber 210 to conjugate chamber 246 may be reduced by expansion of the combustion products 64 caused by the portion of detonation nozzle 268 downstream of detonation throat 270 and/or by the portion of conjugate nozzle 280 downstream of 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 and the pressure decreases. The velocity of the combustion products 64 traveling from detonation chamber 210 to conjugate chamber 246 may be supersonic in nature. As an example, downstream of detonation nozzle 268, detonation throat 270, and/or conjugate throat 282, the velocity of detonation combustion products 64 may be from about 1,000 meters per second (m/s) to about 5,000m/s (e.g., from about 1,000m/s to about 3,000m/s, such as from about 2,000m/s to about 3,500m/s, such as from about 2,500m/s to about 4,500m/s, or such as from about 3,000m/s to about 5,000 m/s).

As the combustion products 64 flow through the detonation throat 270, the static pressure of the combustion products 64 is generally higher than the static pressure within the conjugate chamber 246. As the combustion products 64 flow through the diverging portion of the detonation nozzle 268, the static pressure of the combustion products 64 is reduced by the expansion. The efficiency of the conversion of kinetic energy of combustion products 64 flowing through detonation nozzle 268 to axial momentum may depend, at least in part, on the ratio of the cross-sectional area of detonation throat 270 to the cross-sectional area of conjugate throat 282. Additionally or alternatively, the efficiency of the conversion of kinetic energy of combustion products 64 flowing through detonation nozzle 268 to axial momentum may depend, at least in part, on a cone half angle (cone-half angle) of the diverging portion of detonation nozzle 268 as determined from a plane perpendicular to detonation throat 270.

In some embodiments, the minimum cross-sectional area of detonation nozzle 268 (e.g., the cross-sectional area defined by detonation throat 270) may be less than the minimum cross-sectional area of conjugated nozzle 280 (e.g., the cross-sectional area defined by conjugated throat 282). For example, the cross-sectional area of detonation throat 270 may be at least 1% smaller than the cross-sectional area of 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% smaller than the cross-sectional area of conjugate throat 282.

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

The pressure drop of the combustion products 64 from detonation chamber 210 to conjugate chamber 246 may be greater than or equal to the pressure drop of the combustion products 64 from detonation chamber 224 to conjugate chamber 246, and the static pressure of the combustion products 64 in conjugate chamber 246, e.g., produced by the combined combustion product 64 flow from detonation chamber 210 and detonation chamber 224, will be approximately equal. The combustion products 64 may exhibit a pressure drop across at least a portion of the conjugate chamber 246 and/or there may be local pressure gradients within the conjugate chamber 246, such as pressure gradients attributable to fluid flow, turbulence, mixing, velocity profiles, shrinkage and/or expansion of cross-sectional areas, introduction of dilution air, and the like, as well as combinations of these. In addition, the pressure at the downstream location is lower than the pressure at the upstream location.

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

As shown in fig. 4A and 4B, detonation nozzle 268 may include a cone half angle 406 determined with reference to detonation chamber plane 402. Additionally or alternatively, as shown in fig. 4C, conjugate nozzle 280 may include a cone half angle 406 determined with reference to detonation chamber plane 402. As used herein with respect to detonation nozzle 268, the term "cone half angle" refers to the angle of detonation nozzle 268 as determined from detonation chamber plane 402. As used herein with respect to conjugate nozzle 280, the term "cone half angle" refers to the angle of conjugate nozzle 280 as determined from detonation chamber plane 402. The cone half angle 406 of detonation nozzle 268 and/or the cone half angle 406 of conjugated nozzle 280, respectively, may be configured to provide suitable combustion conditions within combustor section 54, such as suitable back pressure for stabilizing detonation and/or a pressure drop suitable for steady flow of detonation combustion products 64 from detonation chamber 210 to conjugated chamber 246. Additionally or alternatively, the efficiency of the conversion of kinetic energy of the combustion products 64 flowing through detonation nozzle 268 and/or conjugated nozzle 280 to axial momentum may depend, at least in part, on the respective cone half angle 406. As shown in fig. 4A and 4B, the cone half angle 406 of detonation nozzle 268 may be different between a converging portion 408 of detonation nozzle 268 and a diverging portion 410 of detonation nozzle 268. Additionally or alternatively, the cone half angle 406 of detonation nozzle 268 may differ between annular areas of detonation nozzle 268 on respective opposite sides of detonation chamber plane 402. Respective opposite sides of detonation nozzle 268 relative to detonation chamber plane 402 are sometimes referred to as detonation chamber side 412 of detonation nozzle 268 and detonation chamber side 414 of detonation nozzle 268. Detonation chamber side 412 of detonation nozzle 268 refers to the annular region of detonation nozzle 268 radially adjacent to detonation chamber 210 and radially remote from detonation chamber 224. Explosion chamber side 414 of detonation nozzle 268 refers to the annular area of detonation nozzle 268 radially adjacent to explosion chamber 224 and radially remote from detonation chamber 210.

As shown in FIG. 4A, detonation nozzle 268 may include a divergent cone half angle 416. In some embodiments, divergence cone half angle 416 may be different between detonation chamber side 412 and detonation chamber side 414 of detonation nozzle 268. For example, as shown, diverging cone half-angle 416 may include a first diverging cone half-angle 416a corresponding to an annular region radially proximate to detonation chamber 210 and radially distal from detonation chamber 224, or a second diverging cone half-angle 416b corresponding to an annular region of detonation nozzle 268 radially proximate to detonation chamber 224 and radially distal from detonation chamber 210. The first annular region of detonation nozzle 268 and the second annular region of detonation nozzle 268 may be defined by detonation chamber plane 402 from each other. The first divergent cone half angle 416a may sometimes be referred to as the divergent conjugate cone half angle 416a. The second divergent cone half angle 416b may sometimes be referred to as the 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 alternatively, in some embodiments, divergent conjugate cone half angle 416a may intersect detonation chamber plane 402 upstream of the location where divergent detonation cone half angle 416b intersects detonation chamber plane 402. As shown in fig. 4C, the conjugate nozzle 280 may include a divergent cone half angle 416. The divergence cone half angle 416 of the conjugate nozzle 280 may be greater than the divergence cone half angle 416 of the detonation nozzle 284 (fig. 2A and 2B). Other diverging cone half angles 416 are also contemplated. Divergent cone half-angle 416 (e.g., divergent detonation cone half-angle 416b and/or divergent conjugate cone half-angle 416 a), respectively, may be configured to provide suitable combustion conditions within combustor section 54, such as suitable back pressure for stabilizing detonation and/or a pressure drop suitable for detonation combustion products 64 to stably flow from detonation chamber 210 to conjugate chamber 246.

As shown in FIG. 4B, detonation nozzle 268 may include a converging cone half angle 418. In some embodiments, converging cone half angle 418 may be different between detonation chamber side 412 and detonation chamber side 414 of nozzle 268. For example, as shown, converging cone half angle 418 may include a first converging cone half angle 418a corresponding to an annular region of detonation nozzle 268 radially proximate to detonation chamber 210 and radially distal from detonation chamber 224, and a second converging cone half angle 418b corresponding to an annular region of detonation nozzle 268 radially proximate to detonation chamber 224 and radially distal from detonation chamber 210. The first converging cone half angle 418a may sometimes be referred to as a converging conjugate cone half angle 418a. The second converging cone half angle 418b may sometimes be referred to as converging deflagration cone half angle 418b. In some embodiments, as shown, the converging conjugate cone half angle 418a may be less than the converging deflagration cone half angle 418b. Additionally or alternatively, in some embodiments, converging conjugate cone half angle 418a may intersect detonation chamber plane 402 upstream of the location where converging detonation cone half angle 418b intersects detonation chamber plane 402. Additionally or alternatively, in some embodiments, converging cone half angle 418 (e.g., converging conjugate cone half angle 418a and/or converging deflagration cone half angle 418 b) may be less than diverging cone half angle 416 (e.g., diverging conjugate cone half angle 416a and/or diverging deflagration cone half angle 416 b). Other converging cone half angles 418 are also contemplated. Converging cone half angle 418 (e.g., converging deflagration cone half angle 418b and/or converging conjugate cone half angle 418 a), respectively, may be configured to provide suitable combustion conditions within combustor section 54, such as suitable back pressure for stabilizing detonation and/or a pressure drop suitable for detonation combustion products 64 to stably flow from detonation chamber 210 to conjugate chamber 246.

As an example, cone half angle 406 (e.g., diverging cone half angle 416 and/or converging cone half angle 418) may be about 1 to about 30 degrees, such as about 1 ° to about 10 °, such as about 10 ° to about 20 °, such as about 12 ° to about 18 °, or such as about 20 ° to about 30 °. The divergent conjugate cone half angle 416a may be about 1 to about 30 degrees, such as about 1 ° to about 10 °, such as about 5 ° to about 10 °, such as about 10 ° to about 15 °, such as about 15 ° to about 20 °, such as about 20 ° to about 25 °, or such as about 25 ° to about 30 °. The divergent detonation cone half angle 406 may be about 10 degrees to about 60 degrees, such as about 10 ° to about 20 °, such as about 12 ° to about 18 °, such as about 20 ° to about 30 °, such as about 25 ° to about 40 °, or such as about 40 ° to about 60 °. The divergent deflagration cone half angle 416b may be greater than, less than, or equal to the divergent conjugate cone half angle 416a. In some embodiments, the 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% greater than the divergent conjugate cone half angle 416a, such as from about 50% to about 100% greater than the divergent conjugate cone half angle 416a, such as from about 100% to about 150% greater than the divergent conjugate cone half angle, or such as from about 150% to about 200% greater than the divergent conjugate cone half angle 416a.

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

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

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

Such pressure may be determined during operation of knock section 202 within an operating range for which knock section 202 may be configured (e.g., at steady state conditions at nominal and/or cruise speeds of engine 50). The pressure of the combustion products 64 flowing through and/or exiting detonation nozzle 268 may be determined at a downstream region of the diverging portion of detonation nozzle 268 (e.g., upstream of conjugate throat 282 within 10% of the distance between detonation throat 270 and conjugate throat 282, or upstream of conjugate inflection line 272 within 10% of the distance between detonation throat 270 and conjugate inflection line 272). The pressure of the combustion products 64 in the conjugate chamber 246 may be determined downstream of the conjugate nozzle 280 (e.g., at a longitudinal position along the conjugate chamber centerline 278 that is within 20% of the length of the conjugate chamber 246 from the conjugate chamber centerline 274 of the conjugate chamber 246).

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

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

In some embodiments, detonation nozzle 268 may have an under-inflated configuration, a neutral-inflated configuration, or an over-inflated configuration, as determined with respect to the rated speed and/or the cruise speed of engine 50. In some embodiments, the expanded configuration of detonation nozzle 268 may be different relative to a first portion of detonation nozzle 268 defined between detonation throat 270 and conjugate throat 282 and a second portion of detonation nozzle 268 defined between detonation throat 270 and conjugate inflection line 272. As used with reference to detonation nozzle 268, the term "underflows" or "underflows configuration" refers to a configuration of detonation nozzle 268 that provides at least 5% greater pressure of combustion products 64 exiting the divergent portion of detonation nozzle 268 and/or entering the conjugate throat 282 than the pressure of combustion products 64 in the conjugate chamber 246. When used with reference to detonation nozzle 268, the term "neutral expansion" or "neutral expansion configuration" refers to a configuration of detonation nozzle 268 that provides a pressure of combustion products 64 within 5% of the pressure of combustion products 64 in conjugated chamber 246 exiting the diverging portion of detonation nozzle 268 and/or entering conjugated throat 282. When used with reference to detonation nozzle 268, the term "over-expanded" or "over-expanded configuration" refers to a configuration of detonation nozzle 268 that provides at least 5% less pressure of combustion products 64 exiting the diverging portion of detonation nozzle 268 and/or entering the conjugate throat 282 than the pressure of combustion products 64 in the conjugate chamber 246. In some embodiments, detonation nozzle 268 may have an under-expanded or neutral expanded configuration between detonation throat 270 and conjugate throat 282 and an over-expanded configuration between detonation throat 270 and conjugate inflection line 272.

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

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

In general, neutral expansion nozzles generally have higher efficiency than under-expansion nozzles and over-expansion nozzles. In general, over-expanded nozzles generally have higher efficiency than under-expanded nozzles, but over-expanded nozzles may be less stable. However, in some embodiments, with respect to the rated speed and/or cruising speed of the engine 50, the presently disclosed bimodal combustion system 200 may include a over-expanded detonation nozzle 268, such as the over-expanded portion of the detonation nozzle 268 between the detonation throat 270 and the conjugate transition line 272. Additionally or alternatively, the detonation nozzle 268 (e.g., the portion of the detonation nozzle 268 between the detonation throat 270 and the conjugate inflection line 272) may have an under-expanded configuration relative to the cruising speed of the engine 50 and a neutral expanded configuration relative to the rated speed of the engine 50.

In some embodiments, the flow of combustion products 64 from explosion chamber 224 may facilitate flow separation of combustion products 64 from the portion of detonation nozzle 268 between detonation throat 270 and conjugate transition line 272. The pressure of the combustion products 64 flowing from the detonation 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 alternatively, the stabilization and/or support pressure of combustion products 64 flowing from explosion chamber 224 to conjugation chamber 246 may at least partially facilitate the construction of bimodal combustion system 200 including over-expanded detonation nozzle 268 (e.g., the over-expanded portion of detonation nozzle 268 between detonation throat 270 and conjugated inflection line 272), as determined, for example, with respect to a rated speed and/or a cruising speed of engine 50. In some embodiments, detonation nozzles 268 may exhibit an excessive expansion characteristic at cruising speeds and a neutral expansion characteristic at rated speeds when engine 50 transitions from cruising speeds to rated speeds, for example, due to an increase in pressure in detonation chamber 224 and/or conjugation chamber 246.

In some embodiments, detonation nozzles 268 and conjugate nozzles 280 may define stepped nozzles, double expansion nozzles, and/or double throat nozzles. In some embodiments, under the first operating condition, detonation nozzle 268 may exhibit an under-expansion characteristic and conjugated nozzle 280 may exhibit a neutral expansion characteristic and/or an over-expansion characteristic. Additionally or alternatively, under the second operating condition, detonation nozzle 268 may exhibit neutral expansion characteristics and conjugate nozzle 280 may exhibit excessive expansion characteristics.

In some embodiments, detonation nozzle 268 and/or conjugated nozzle 280 may be configured for choked flow within an operating range for which detonation section 202 may be configured (e.g., at cruise speed and/or rated speed). As used herein, the term "choked flow" refers to a limiting condition in which the mass flow through detonation nozzle 268 and/or conjugated nozzle 280 does not increase with further decreases in downstream pressure for a given upstream pressure and temperature. Choked flow may occur when the pressure ratio of combustion products 64 on opposite sides of detonation nozzle 268 is at least about 1.7 (e.g., at least about 1.7 to at least about 2.1). The detonation nozzle-to-pressure ratio at which choked flow may occur may depend, at least in part, on the configuration of nozzle region 266 and/or other portions of detonation chamber 210, as well as on the composition of combustion products 64. In some embodiments, the engine may be configured to exhibit a detonation nozzle-pressure ratio that is suitable for use in choked flow conditions when a particular operating condition (e.g., a high power operating condition, a nominal operating condition, a cruise speed, and/or a rated speed) is reached.

Referring still to fig. 2A and 2B, and referring to fig. 5A-5D, an exemplary dual-peak combustion system 200 is further described. As shown in fig. 5A, at least a portion of detonation chamber 210 may have a shape that includes a torus shape. For example, detonation region 264 of detonation chamber 210 may include a torus shape. The torus shape of detonation chamber 210 (e.g., detonation region 264 of detonation chamber 210) may include an annular torus, a toroidal torus, or a spindle torus. As shown in fig. 5B, at least a portion of detonation chamber 210 may have a shape that includes a parabolic annular shape. For example, the nozzle region 266 of the detonation chamber 210 may include a parabolic annular shape. The parabolic annular shape of detonation chamber 210 (e.g., nozzle region 266 of detonation chamber 210) may include a converging portion 500, a diverging portion 502, and a saddle region 504 disposed between converging portion 500 and diverging portion 502. At least a portion of the nozzle region 266 including the converging portion 500 and the diverging portion 502 may define a detonation nozzle 268. Saddle region 504 may define detonation throat 270. For example, in addition to defining at least a portion of detonation zone 264 of detonation chamber 210, the torus shape of detonation chamber 210 may define at least a portion of nozzle zone 266 of detonation chamber 210. A portion of the parabolic annular shape of detonation chamber 210 may be defined by a portion of the toroidal shape of detonation chamber 210. For example, the torus shape of the detonation chamber 210 may define at least a portion of the converging portion 500 of the nozzle region 266.

As shown in fig. 5C, explosion chamber 224 may have a generally cylindrical annular shape. Additionally or alternatively, as shown in fig. 5D, the conjugate chamber 246 may have a generally cylindrical annular shape. The cylindrical annular shape of explosion chamber 224 and/or conjugation chamber 246 may have a frustoconical configuration. Additionally or alternatively, the cylindrical annular shape of explosion chamber 224 and/or conjugation chamber 246 may have a contoured configuration with one or more saddle-like or narrowing portions, and/or with one or more node or widening portions.

6A-6C, an exemplary detonation combustor 206 is further described. As shown, detonation combustor 206 may include a detonation fuel manifold 212 having 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 oxidant 60 from the detonation fuel manifold 212 to the detonation chamber 210. For example, a plurality of detonation fuel orifices 218 may supply a mixture of fuel 62 and/or oxidant 60 to detonation chamber 210. Additionally or alternatively, the first plurality of detonation fuel orifices 218a may supply fuel 62 to detonation chamber 210 and the second plurality of detonation fuel orifices 218b may supply oxidant 60 to 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 one or more detonation oxidant supply lines 216.

As shown in FIG. 6A, detonation fuel manifold 212 may include a plurality of detonation fuel orifices 218 circumferentially spaced about an annular face 600 of detonation fuel manifold 212. 6B and 6C, detonation fuel manifold 212 may include a plurality of detonation fuel orifices 218 circumferentially spaced about a circumferential face 604 of detonation fuel manifold 212 (e.g., an inwardly facing circumferential face 604a (FIG. 6B) and/or an outwardly facing circumferential face 604B (FIG. 6C) of detonation fuel manifold 212).

Referring now to fig. 7A-7F, and with further reference to fig. 4A-4C, an exemplary dual-peak combustion system 200 is further described. As described with respect to fig. 4A-4C, and as further shown in fig. 7A-7F, detonation throat centerline 400 may define an annular center of detonation throat 270, and detonation chamber plane 402 may intersect detonation throat centerline 400 tangentially perpendicular to detonation throat centerline 400. As also shown in fig. 7A-7F, chamber 224 may have a chamber centerline 700 defining an annular center of chamber 224. Explosion chamber centerline 700 may be equally spaced between outer explosion chamber wall 254 and inner explosion chamber wall 256 at a location of explosion chamber 224 having a minimum annular ring width defined between outer explosion chamber wall 254 and inner explosion chamber wall 256. Explosion chamber centerline 700 may have a generally elliptical or circular shape. Explosion chamber centerline 700 may circumferentially surround longitudinal axis 248 of engine 50. Explosion chamber plane 702 may intersect explosion chamber centerline 700 tangentially perpendicular to explosion chamber centerline 700. Tangential normal orientation of chamber plane 702 may include a perpendicular orientation relative to an annular plane that spans a minimum annular ring width defined between outer chamber wall 254 and inner chamber wall 256. Explosion chamber planar surface 702 may circumferentially surround longitudinal axis 248 of engine 50. Explosion chamber plane 702 may have a generally linear configuration along an explosion chamber centerline 704, with explosion chamber centerline 704 intersecting a conjugate chamber centerline 278 and explosion chamber centerline 700 in an orientation parallel to explosion chamber plane 702. As an example, chamber planar surface 702 may have a cylindrical configuration or a frustoconical configuration.

Detonation chamber plane 402 may intersect with detonation chamber plane 702 at a conjugate intersection 706 located within conjugate chamber 246. In some embodiments, as shown in FIG. 7A, detonation chamber plane 402 may intersect detonation chamber plane 702 at a conjugate intersection 706 that coincides with conjugate chamber plane 276. Additionally or alternatively, as shown in FIG. 7A, detonation chamber plane 402 may intersect detonation chamber plane 702 at a conjugate intersection 706 that coincides with conjugate chamber centerline 274. In some embodiments, at least one detonation chamber centerline 404 may intersect at least one detonation chamber centerline 704 at a conjugate intersection 706 that coincides with a conjugate chamber plane 276 and/or coincides with a conjugate chamber centerline 274.

In some embodiments, detonation chamber plane 402 may intersect detonation chamber plane 702 at a conjugate intersection 706 located radially outward from conjugate chamber plane 276 or at a location radially inward from conjugate chamber plane 276. Additionally or alternatively, detonation chamber plane 402 may intersect detonation chamber plane 702 at a conjugate intersection 706 located upstream of conjugate chamber centerline 274 or downstream of conjugate chamber centerline 274. In some embodiments, at least one detonation chamber centerline 404 may intersect at least one detonation chamber centerline 704 at a conjugate intersection 706 located radially outward from conjugate chamber plane 276 and/or at a location radially inward from conjugate chamber plane 276. Additionally or alternatively, in some embodiments, at least one detonation chamber centerline 404 may intersect at least one detonation chamber centerline 704 at a conjugate intersection 706 located upstream of conjugate chamber centerline 274 and/or at a location downstream of conjugate chamber centerline 274. Knock chamber plane 402 and knock chamber plane 702 may intersect each other at a normal angle (e.g., right angle) or an oblique angle (such as an acute or obtuse angle). Additionally or alternatively, at least one knock chamber centerline 404 and at least one knock chamber centerline 704 may intersect each other at a normal angle (e.g., right angle) or an oblique angle (such as an acute or obtuse angle). By way of example, FIG. 7A shows detonation chamber plane 402 intersecting detonation chamber plane 702 at, for example, an oblique angle (such as an acute angle) at a conjugate intersection 706 coincident with conjugate chamber centerline 274 of conjugate chamber 246. As shown, at least one detonation chamber centerline 404 may intersect at least one detonation chamber centerline 704, for example, at an oblique or acute angle at a conjugate intersection 706 that coincides with a conjugate chamber centerline 274.

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

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

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

As another example, FIG. 7E shows detonation chamber plane 402 intersecting detonation chamber plane 702 at a normal angle radially outward from conjugate chamber plane 276 and at a location upstream of conjugate chamber centerline 274. As shown, at least one detonation chamber centerline 404 may intersect at a normal angle with at least one detonation chamber centerline 704 at a location radially outward from conjugate chamber plane 276 and upstream of conjugate chamber centerline 274.

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

Referring now to FIG. 8, an exemplary detonation combustor 206 is further described. As shown in fig. 8, detonation waves 800 subsequent to the shock wave 802 may propagate annularly through the detonation chamber 210, such as through the detonation region 264 of the detonation chamber 210. As shown, the shock wave 802 and the corresponding detonation wave 800 may propagate in a counter-clockwise direction, while the combustion products 64 expand in general three dimensions. Alternatively, the shock wave 802 and the corresponding detonation wave 800 may propagate annularly, for example, in a clockwise direction. Although one shock wave 802 and corresponding detonation wave 800 are depicted in fig. 8 for purposes of illustration, in some embodiments, the example detonation combustor 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 propagate simultaneously around the annular volume of the detonation chamber 210, e.g., in a circumferentially spaced relationship.

Although detonation chamber 210 is schematically illustrated in fig. 8 as having a generally cylindrical annular shape, in some embodiments detonation chamber 210 may include any shape that provides a continuous path for shock waves 802 and corresponding detonation waves 800. As an example, detonation chamber 210 may include a torus shape, as shown in fig. 5A. Additionally or alternatively, detonation chamber 210 may include a parabolic annular shape, as shown in fig. 5B. As a further example, detonation chamber 210 may include any annular shape, such as a trapezoidal shape or an oval shape. For detonation chambers 210 that include a torus shape as shown in fig. 5A, the shock wave 802 and corresponding detonation wave 800 may propagate in the circumferential direction 506 (as shown in fig. 5A). The shock wave 802 and the corresponding detonation wave 800 may encompass all or a portion of an annular perimeter defined by the detonation chamber 210, such as all or a portion of an 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 encompass all or a portion of the polar perimeter 508 defined by the detonation region 264 of the detonation chamber 210 (as shown in fig. 5A).

As shown in fig. 8, the area before the shock wave 802 and the 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 oxidant 60 knocks, the shock wave 802 generated by the knocking may temporarily block more fuel 62 and oxidant 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, further depleting the fuel 62 and the oxidant 60. As the shock wave 802 and corresponding detonation wave 800 propagate around the annular volume, additional fuel 62 and oxidizer 60 may generally trail the shock wave 802 and corresponding detonation wave 800 into the detonation chamber 210.

As the combustion products 64 expand as they propagate through the detonation region 264 of the detonation chamber 210, at least a portion of the shock wave 802 may be transferred 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 wave passes from the detonation region 264 through the nozzle region 266 of the detonation chamber 210. The shock wave 802 propagating in the circumferential or hoop direction 506 may sometimes be referred to as a primary shock wave, a circumferential shock wave, or a hoop shock wave. In some embodiments, a longitudinal shock wave 804 may be generated that propagates from detonation chamber 210 into conjugation chamber 246 (fig. 2A and 2B). The longitudinal shock wave 804 may be oriented generally along the at least one detonation chamber centerline 404. As the combustion products 64 enter the conjugate chamber 246 through the nozzle region 266 of the detonation chamber 210, a longitudinal shock wave 804 may be generated. The longitudinal shock wave 804 may sometimes be referred to as a secondary shock wave 804. The longitudinal shock wave 804 may surround at least a portion of the circumference 510 (fig. 5B) of the nozzle region 266 of the detonation chamber 210.

Although one longitudinal shock wave 804 is depicted in FIG. 8 for purposes of illustration, in some embodiments, the exemplary detonation combustor 206 may be configured to continuously generate multiple longitudinal shock waves 804. For example, a plurality of longitudinal shock waves 804 may simultaneously propagate longitudinally from detonation chamber 210 into conjugated chamber 246, e.g., in a circumferentially spaced relationship. Additionally or alternatively, the plurality of longitudinal shock waves 804 may have an annular configuration, such as an annular configuration oriented with respect to the detonation throat centerline 400, the at least one detonation chamber centerline 404, and/or the conjugate chamber centerline 278. Longitudinal shock waves 804 of annular configuration may propagate longitudinally from detonation chamber 210 into conjugate chamber 246, thereby generating thrust. Additionally or alternatively, the longitudinal shock wave 804 may propagate longitudinally from the conjugate chamber 246 into one or more turbine sections 66 (fig. 1A and 1B) of the engine 50 through the outlet section 56 (fig. 1A and 1B) of the engine 50 and/or through the outlet nozzle 68 (fig. 1A and 1B) of the engine 50, thereby generating thrust.

Longitudinal shock wave 804 may emanate from detonation nozzle 268 (e.g., detonation throat 270 of nozzle region 266 of detonation chamber 210). Longitudinal shock wave 804 may propagate from nozzle region 266 of detonation chamber 210 and into conjugation chamber 246. Additionally or alternatively, in some embodiments, the primary shock wave 802 may propagate through the nozzle region 266 of the detonation chamber 210 and into the conjugation chamber 246. In some embodiments, detonation wave 800 may remain within detonation region 264 of detonation chamber 210, e.g., such that the detonation process may be completed within detonation region 264 of detonation chamber 210, while longitudinal shock waves 804 and corresponding combustion products 64 propagate through nozzle region 266 of detonation chamber 210 into 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 of the detonation throat 270. In some embodiments, at least some combustion (e.g., detonation and/or deflagration) may occur within conjugate chamber 246.

In some embodiments, detonation combustor 206 may include a pre-detonation 806, with pre-detonation 806 configured to generate a blast wave 808 suitable for initiating detonation within detonation chamber 210. Additionally or alternatively, in some embodiments, the longitudinal shock wave 804 may be at least partially initiated by the detonation nozzle 268 reflecting the primary shock wave 802. Additionally or alternatively, the longitudinal shock wave 804 may be at least partially induced by backpressure generated by the detonation nozzle 268 and/or the conjugated nozzle 280 (fig. 2A and 2B). Additionally or alternatively, longitudinal shock wave 804 may be at least partially induced by backpressure generated by a deflagration occurring within explosion chamber 224 (fig. 2A and 2B) and/or conjugation chamber 246. In some embodiments, detonation within detonation chamber 210 may be initiated at least in part by backpressure generated by detonation within detonation nozzle 268, conjugated nozzle 280, and/or detonation chamber 224 and/or conjugated chamber 246. In some embodiments, knock may be caused when a choked flow condition is established (e.g., when the knock nozzle pressure ratio increases to a suitable level), for example, as a result of such back pressure. In some embodiments, fuel 62 and oxidant 60 suitable for detonation may be supplied to the detonation chamber when a choked flow condition has been established. In some embodiments, detonation may occur in detonation chamber 210 before detonation is induced within detonation chamber 210. For example, detonation chamber 210 may be used for detonation during specified operating conditions. By providing a fuel 62 and an oxidant 60 suitable for detonation (e.g., a mixture of fuel 62 and oxidant 60 unsuitable for detonation), detonation may be performed within detonation chamber 210. Additionally or alternatively, a detonation may occur within detonation chamber 210 prior to establishing the choked flow condition.

Referring now to fig. 9, an exemplary method according to the present disclosure is further described. As an example, the exemplary method may include a method of generating thrust. Additionally or alternatively, the exemplary method may include a method of combusting a fuel. Additionally or alternatively, the exemplary methods may include methods of operating the engine 50 (e.g., the turbine engine 100, a rocket engine, a ramjet engine, or a combination thereof, such as a turbine rocket engine, a turbine ramjet engine, or a rocket ramjet engine). As shown in FIG. 9, exemplary method 900 may include performing a detonation within a detonation chamber and/or within a conjugate chamber in fluid communication with the detonation chamber at block 902. The combustion products generated by the deflagration in the deflagration chamber may flow through the conjugation chamber, for example from the deflagration chamber. At block 904, the example method 900 may include detonating within a detonation chamber in fluid communication with the conjugate chamber. The combustion products generated by detonation in the detonation chamber may flow from the detonation chamber to the conjugate chamber. The combustion products flowing through the conjugate chamber may generate thrust. The thrust force may be at least partially due to a shock wave emanating from the detonation chamber, such as a longitudinal shock wave emanating from the detonation nozzle as the combustion products enter the conjugate chamber through the nozzle region of the detonation chamber. Additionally or alternatively, the thrust force may be due, at least in part, to detonation occurring within the detonation chamber and/or within the conjugate chamber.

Accordingly, the presently disclosed systems and methods may utilize a bimodal combustion system to provide thrust to an engine while achieving improved performance and/or the ability to operate under a wider range of operating conditions and thermal load requirements. Additionally or alternatively, the presently disclosed systems and methods may provide significantly improved specific wash and/or specific fuel consumption, and/or relatively low NOx emissions. Additionally or alternatively, due to the presently disclosed systems and methods, the exemplary engine for specifying load requirements may be relatively smaller, lighter in weight, and/or may exhibit a higher thrust-to-weight ratio.

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

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

The combustion system of any preceding clause, wherein the nozzle region comprises a detonation nozzle defined by a first of the one or more detonation chamber walls, the detonation nozzle comprising a detonation throat defining a location of the detonation nozzle having 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 includes a diverging portion downstream of the detonation throat, wherein the diverging portion of the detonation nozzle has a cone half angle of 1 to 10 degrees.

The combustion system of any preceding clause, wherein the detonation nozzle includes a converging portion upstream of the detonation throat, wherein the converging portion of the detonation nozzle has a cone half angle of 5 degrees to 30 degrees.

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

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

The combustion system of any preceding clause, wherein at least a portion of the detonation nozzle has an over-expanded configuration, as determined relative 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 neutral expanded configuration, as determined relative to the rated speed and/or the cruise speed of the engine.

The combustion system of any preceding clause, wherein: a detonation throat centerline defines an annular center of the detonation throat, and a detonation chamber plane tangential to the detonation throat centerline intersects the detonation throat centerline; a chamber centerline defines an annular center of the chamber, and a chamber plane tangential to the chamber centerline intersects the chamber centerline; the conjugate inflection line circumferentially about the longitudinal axis defines a linear inflection that delineates the detonation chamber wall and the detonation chamber wall from each other, or a linear inflection that represents a tangent to a forefront bevel or forefront curve of the conjugate chamber wall disposed between the detonation chamber wall and the detonation chamber wall; a conjugate chamber centerline defines a volumetric center of the conjugate chamber as determined relative to a volume of the conjugate chamber located between the conjugate turning line and a downstream end of the conjugate chamber; the conjugate chamber plane intersects the conjugate turning line and the conjugate chamber center line; and the detonation chamber plane intersect each other at a normal or oblique angle.

The combustion system of any preceding clause, wherein the detonation chamber plane and the detonation chamber plane intersect each other at a conjugate intersection comprising a location within the conjugate chamber, the location being at least one of: coincident with the conjugate chamber plane, radially inward from the conjugate chamber plane, or radially outward from the conjugate chamber plane; and coincident with the conjugate chamber centerline, upstream of the conjugate chamber centerline, or downstream of the conjugate chamber centerline.

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

The combustion system of any preceding clause, wherein the detonation chamber and the detonation chamber are each located on opposite sides of a conjugate inflection line circumferentially about the longitudinal axis, the conjugate inflection line defining a linear inflection that delineates the detonation chamber wall and the detonation chamber wall from each other, or the conjugate inflection line defining a linear inflection that represents a tangent to a forefront bevel or forefront curve of a conjugate chamber wall disposed between the detonation chamber wall and the detonation chamber wall.

The combustion system of any preceding clause, wherein the detonation chamber comprises a detonation nozzle defined by a first of the one or more detonation chamber walls, the detonation nozzle comprising a detonation throat defining a location of the detonation nozzle having a first annular cross-sectional area having a first minimum annular width relative to an adjacent portion of the detonation nozzle; wherein a first of the one or more conjugate chamber walls comprises a conjugate nozzle comprising a conjugate throat defining a location of the conjugate nozzle having a second annular cross-sectional area having 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 of the one or more conjugate chamber walls to a conjugate chamber plane intersecting the conjugate turn line and a conjugate chamber centerline defining a volumetric center of the conjugate chamber as determined relative to a volume of the conjugate chamber located between the conjugate turn line and a downstream end of the conjugate chamber, and the first of the one or more conjugate chamber walls is 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 smaller 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 1% to 90% smaller than the second annular cross-sectional area.

The combustion system of any preceding clause, wherein the detonation nozzle includes a first diverging cone half angle corresponding to a first annular region of the detonation nozzle radially proximate to the detonation chamber and radially distal to the detonation chamber, and a second diverging cone half angle corresponding to a second annular region of the detonation nozzle radially proximate to the detonation chamber and radially distal to the detonation chamber, wherein the second diverging cone half angle is greater than the first diverging cone half angle.

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

The combustion system of any preceding clause, wherein the second cone half angle is 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 integrally integrated with the one or more detonation chamber walls, the detonation fuel manifold configured to supply fuel and/or oxidant to the detonation chamber.

The combustion system of any preceding clause, wherein the deflagration burner comprises a plurality of deflagration fuel manifolds each configured to supply fuel and/or oxidant 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 detonation chamber, or wherein at least a portion of the detonation 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 a parabolic torus shape.

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

An engine according to any preceding clause, wherein the engine comprises:

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

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

An engine according to 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 1 to 12 compressor stages.

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

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

The engine of any preceding clause, wherein the turbine engine exhibits a thrust force versus fuel consumption of about 8 g/kn-sec to about 14 g/kn-sec at rated speed and/or cruise speed.

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

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

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

A method of combusting a fuel, the method comprising: deflagrating within a deflagration chamber and/or within a conjugate chamber in fluid communication with the deflagration chamber, producing deflagrated combustion products, wherein the deflagrated combustion products flow through the conjugate chamber, thereby producing thrust forces; and knocking in a detonation chamber in fluid communication with the conjugate chamber to produce detonation combustion products, wherein the detonation combustion products flow through the conjugate chamber to produce thrust.

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

The method of any preceding clause, wherein knocking within the knock chamber comprises: a plurality of shock waves propagating longitudinally through the detonation chamber are generated, 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 downstream of the detonation nozzle of 1,000 meters/second to 5,000 meters/second.

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 example 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. These 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 including one or more detonation chamber walls defining a detonation chamber;

a deflagration burner comprising one or more deflagration chamber walls defining a deflagration chamber; and

one or more conjugate chamber walls defining a conjugate chamber in fluid communication with the detonation chamber and the detonation chamber;

wherein the detonation chamber includes a detonation zone and a nozzle zone providing fluid communication between the detonation zone and the conjugate chamber.

2. The combustion system of claim 1, wherein the nozzle region comprises a detonation nozzle defined by a first of the one or more detonation chamber walls, the detonation nozzle comprising a detonation throat defining a location of the detonation nozzle having 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 includes a diverging portion downstream of the detonation throat, wherein the diverging portion of the detonation nozzle has a cone half angle of 1 to 10 degrees, and/or wherein the detonation nozzle includes a converging portion upstream of the detonation throat, wherein the converging portion of the detonation nozzle has a cone half angle of 5 to 30 degrees.

4. The combustion system of claim 2, wherein the detonation nozzle includes a converging portion and a diverging portion, the converging portion having a decreasing cross-sectional area upstream of the detonation throat in a direction from the detonation zone toward the detonation throat, and the diverging portion having an increasing cross-sectional area downstream of the detonation throat in a direction from the detonation throat toward the conjugate chamber.

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

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

8. The combustion system of claim 2, wherein:

A detonation throat centerline defines an annular center of the detonation throat, and a detonation chamber plane tangential to the detonation throat centerline intersects the detonation throat centerline;

a chamber centerline defines an annular center of the chamber, and a chamber plane tangential to the chamber centerline intersects the chamber centerline;

the conjugate inflection line circumferentially about the longitudinal axis defines a linear inflection that delineates the detonation chamber wall and the detonation chamber wall from each other, or a linear inflection that represents a tangent to a forefront bevel or forefront curve of the conjugate chamber wall disposed between the detonation chamber wall and the detonation chamber wall;

a conjugate chamber centerline defines a volumetric center of the conjugate chamber as determined relative to a volume of the conjugate chamber located between the conjugate turning line and a downstream end of the conjugate chamber;

the conjugate chamber plane intersects the conjugate turning line and the conjugate chamber center line; and is also provided with

The detonation chamber plane and the detonation chamber plane intersect each other at a normal or oblique angle.

9. The combustion system of claim 8, wherein the detonation chamber plane and the detonation chamber plane intersect each other at a conjugate intersection comprising a location within the conjugate chamber, the location being at least one of:

Coincident with the conjugate chamber plane, radially inward from the conjugate chamber plane, or radially outward from the conjugate chamber plane; and

coincident with the conjugate chamber centerline, upstream of the conjugate chamber centerline, or downstream of the conjugate chamber centerline.

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