Patent Application
20180356099
The present subject matter relates generally to a system and method of continuous detonation in an engine.
Many propulsion systems, such as gas turbine engines, are based on the Brayton Cycle, where air is compressed adiabatically, heat is added at constant pressure, the resulting hot gas is expanded in a turbine, and heat is rejected at constant pressure. The energy above that required to drive the compression system is then available for propulsion or other work. Such propulsion systems generally rely upon deflagrative combustion to burn a fuel/air mixture and produce combustion gas products which travel at relatively slow rates and constant pressure within a combustion chamber. While engines based on the Brayton Cycle have reached a high level of thermodynamic efficiency by steady improvements in component efficiencies and increases in pressure ratio and peak temperature, further improvements are welcomed nonetheless.
Accordingly, improvements in engine efficiency have been sought by modifying the engine architecture such that the combustion occurs as a detonation in either a continuous or pulsed mode. The pulsed mode design involves one or more detonation tubes, whereas the continuous mode is based on a geometry, typically an annulus, within which single or multiple detonation waves spin. For both types of modes, high energy ignition detonates a fuel/air mixture that transitions into a detonation wave (i.e., a fast moving shock wave closely coupled to the reaction zone). The detonation wave travels in a Mach number range greater than the speed of sound (e.g., Mach 4 to 8) with respect to the speed of sound of the reactants. The products of combustion follow the detonation wave at the speed of sound relative to the detonation wave and at significantly elevated pressure. Such combustion products may then exit through a nozzle to produce thrust or rotate a turbine.
Although detonation combustors may generally provide improved efficiency and performance, there exists a need for propulsion systems further integrating a detonation combustion system that may improve propulsion system efficiency and performance.
Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.
The present disclosure is directed to a propulsion system defining a radial direction extended from a longitudinal centerline extended along a longitudinal direction, and a circumferential direction relative to the longitudinal centerline. The propulsion system includes a rotating detonation combustion (RDC) system defining a plurality of fuel-oxidizer mixing nozzles each defined by a converging-diverging nozzle wall defining a nozzle flowpath. The nozzle wall defines a throat and a lengthwise direction extended between a nozzle inlet and nozzle outlet along the lengthwise direction. The longitudinal centerline of the propulsion system and the radial direction together define a reference plane, and the lengthwise direction of the nozzle intersects the reference plane and defines a nozzle angle greater than zero degrees and approximately 80 degrees or less relative to the reference plane.
In various embodiments, the RDC system further includes an annular outer wall defining at least in part a combustion chamber downstream of the plurality of nozzles. In one embodiment, the RDC system defines the outer wall generally concentric to the longitudinal centerline of the propulsion system. In another embodiment, the propulsion system further includes a turbine nozzle disposed downstream of the combustion chamber. The turbine nozzle includes a plurality of turbine nozzle airfoils defining an exit angle relative to the reference plane.
In one embodiment, the exit angle of the plurality of turbine nozzle airfoils is configured to a desired circumferential direction relative to an exhaust section of the propulsion system. In another embodiment, the exit angle and the nozzle angle are within approximately 20 degrees relative to one another. In still another embodiment, the exit angle and the nozzle angle are approximately equal. In yet another embodiment, the plurality of turbine nozzle airfoils defines a turbine nozzle inlet angle in which the inlet angle is less than or approximately equal to the exit angle. In still yet another embodiment, the plurality of turbine nozzle airfoils defines a turbine nozzle inlet angle in which the turbine nozzle inlet angle is approximately equal to or less than the nozzle angle.
In various embodiments, the RDC system defines an RDC inlet comprising a plurality of RDC inlet airfoils defining an inlet angle relative to the reference plane. In one embodiment, the inlet angle of the RDC inlet airfoils is greater than zero degrees and approximately 80 degrees or less relative to the reference plane. In another embodiment, the inlet angle and the nozzle angle are within approximately 20 degrees relative to one another.
In one embodiment of the propulsion system, each nozzle of the RDC system further defines a fuel injection port disposed approximately at the throat of each nozzle, wherein the fuel injection port is configured to flow a fuel to the nozzle flowpath.
The present disclosure is further directed to a gas turbine engine defining a radial direction extended from a longitudinal centerline extended along a longitudinal direction, and a circumferential direction relative to the longitudinal centerline. The gas turbine engine includes a rotating detonation combustion (RDC) system defining a plurality of fuel-oxidizer mixing nozzles each defined by a converging-diverging nozzle wall defining a nozzle flowpath. The nozzle wall defines a throat and a lengthwise direction extended between a nozzle inlet and nozzle outlet along the lengthwise direction. The longitudinal centerline of the propulsion system and the radial direction together define a reference plane, and the lengthwise direction of the nozzle intersects the reference plane and defines a nozzle angle greater than zero degrees and approximately 80 degrees or less relative to the reference plane. The RDC system further defines an annular outer wall defining at least in part a combustion chamber downstream of the plurality of nozzles, and the combustion chamber defines a combustion inlet proximate to the plurality of nozzles and a combustion outlet downstream thereof. The gas turbine engine further includes a first turbine rotor at the combustion outlet of the RDC system, in which the first turbine rotor is in direct fluid communication with the combustion chamber.
In one embodiment of the gas turbine engine, the nozzle angle is greater than approximately 65 degrees and less than approximately 80 degrees, inclusively.
In another embodiment of the gas turbine engine, each nozzle of the RDC system further defines a fuel injection port disposed approximately at the throat of each nozzle. The fuel injection port is configured to flow a fuel to the nozzle flowpath.
In still another embodiment of the gas turbine engine, the first turbine rotor is configured to rotate co-directional to a direction of bulk swirl of fuel/oxidizer mixture.
In various embodiments of the gas turbine engine, the RDC system defines an RDC inlet comprising a plurality of RDC inlet airfoils defining an inlet angle relative to the reference plane. In one embodiment, the inlet angle of the RDC inlet airfoils is greater than zero degrees and approximately 80 degrees or less relative to the reference plane. In another embodiment, the inlet angle and the nozzle angle are within approximately 20 degrees relative to one another
These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.
A full and enabling disclosure of the present invention, 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:
Reference will now be made in detail to present embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the invention.
As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.
The terms “forward” and “aft” refer to relative positions within a gas turbine engine or vehicle, and refer to the normal operational attitude of the gas turbine engine or vehicle. For example, with regard to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.
The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.
The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.
Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and/or systems. For example, the approximating language may refer to being within a 10 percent margin.
Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.
Embodiments of a propulsion system including a bulk swirl rotating detonation combustion (RDC) system are generally provided herein that may increase a bulk swirl of combustion gases within the combustion chamber of the RDC system, thereby improving propulsion system efficiency and performance. The bulk swirl may reduce a length of the turbine nozzle or altogether eliminate the turbine nozzle, thereby enabling direct fluid communication of the combustion gases from the combustion chamber to a first turbine rotor. Reducing the length of or eliminating the turbine nozzle may improve overall propulsion system efficiency and performance, such as by reducing part counts, length, weight, and improving thermodynamic efficiency by reducing an amount of cooling oxidizer removed from combustion and energy release.
Referring now to the figures,
As will be discussed in further detail below, at least a portion of the flow of oxidizer 195 is mixed with a fuel 163 (shown in
As will be appreciated, in various embodiments of the propulsion system 10 defining a gas turbine engine, rotation of the turbine(s) within the exhaust section 106 generated by the combustion products 138 is transferred through one or more shafts or spools to drive the compressor(s) within the inlet section 104. In various embodiments, the inlet section 104 may further define a fan section, such as for a turbofan engine configuration, such as to propel air across a bypass flowpath outside of the RDC system 100 and exhaust section 106.
It will be appreciated that the propulsion system 10 depicted schematically in
Moreover, it should also be appreciated that the RDC system 100 may further be incorporated into any other suitable aeronautical propulsion system, such as a turboshaft engine, a turboprop engine, a turbojet engine, a ramjet engine, a scramjet engine, etc. Further, in certain embodiments, the RDC system 100 may be incorporated into a non-aeronautical propulsion system, such as a land-based or marine-based power generation system. Further still, in certain embodiments, the RDC system 100 may be incorporated into any other suitable propulsion system, such as a rocket or missile engine. With one or more of the latter embodiments, the propulsion system may not include a compressor in the inlet section 104 or a turbine in the exhaust section 106.
Referring still to
Referring now to
The nozzle assembly 128 is defined at the upstream end of the combustion chamber 122 at the combustion chamber inlet 124. The nozzle assembly 128 generally defines a nozzle inlet 144, a nozzle outlet 146 adjacent to the combustion chamber inlet 124, and a throat 152 between the nozzle inlet 144 and the nozzle outlet 146. A nozzle flowpath 148 is defined from the nozzle inlet 144 through the throat 152 and the nozzle outlet 146.
The nozzle assembly 128 defines a plurality of nozzles 140 each defined by a nozzle wall 150. Each nozzle 140, or more specifically, the nozzle wall 150, generally defines a converging-diverging nozzle, i.e. each nozzle 140 defines a decreasing cross sectional area along a converging area 159 from approximately the nozzle inlet 144 to approximately the throat 152, and further defines an increasing cross sectional area along a diverging area 161 from approximately the throat 152 to approximately the nozzle outlet 146.
Between the nozzle inlet 144 and the nozzle outlet 146, a fuel injection port 162 is defined in fluid communication with the nozzle flowpath 148 through which the oxidizer 195 flows. The fuel injection port 162 introduces a liquid or gaseous fuel 163 (or mixture thereof) to the flow of oxidizer 195 through a fuel port outlet 164 to produce the fuel/oxidizer mixture 132. In various embodiments, the fuel injection port 162 is disposed at approximately the throat 152 of the nozzle assembly 128. Each nozzle 140 may include a plurality of fuel injection ports 162 and fuel port outlets 164 disposed around the throat 152 of each nozzle 140.
Referring briefly to
More particularly, it will be appreciated that the RDC system 100 is of a detonation-type combustor, deriving energy from the continuous detonation wave 130 of detonation. For a detonation combustor, such as the RDC system 100 disclosed herein, the combustion of the fuel/oxidizer mixture 132 is effectively a detonation as compared to a burning, as is typical in the traditional deflagration-type combustors. Accordingly, a main difference between deflagration and detonation is linked to the mechanism of flame propagation. In deflagration, the flame propagation is a function of the heat transfer from a reactive zone to the fresh mixture, generally through conduction. By contrast, with a detonation combustor, the detonation is a shock induced flame, which results in the coupling of a reaction zone and a shockwave. The shockwave compresses and heats the fresh mixture 132, increasing such mixture 132 above a self-ignition point. On the other side, energy released by the combustion contributes to the propagation of the detonation wave 130. Further, with continuous detonation, the detonation wave 130 propagates around the combustion chamber 122 in a continuous manner, operating at a relatively high frequency. Additionally, the detonation wave 130 may be such that an average pressure inside the combustion chamber 122 is higher than an average pressure within typical combustion systems (i.e., deflagration combustion systems). Accordingly, the region 134 behind the detonation wave 130 has very high pressures.
Referring back to
The nozzle 140 defining the nozzle angle 133 generally produces a bulk swirl of the combustion gases 138 at least partially along the circumferential direction C relative to the longitudinal centerline 116. The nozzle angle 133 is disposed co-directional to the detonation wave 130. For example, a schematic reference arrow 127 indicates the direction of the bulk swirl of a fuel/oxidizer mixture 132 egressing the nozzle assembly 128. The nozzle angle 133 is disposed, at least along the circumferential direction C, co-directional to the direction 127 of the bulk swirl of the fuel/oxidizer mixture 132 (further shown in
For example, in one embodiment of the propulsion system 10 such as generally provided in
More particularly, the propulsion system 10 defines a first turbine rotor 131 at the combustion outlet 126 of the RDC system 100. The first turbine rotor 131 is in direct fluid communication with the combustion chamber 122 (shown in
In various embodiments, the first turbine rotor 131 may define a first rotating stage of the high pressure turbine 28 of the turbine section 29. In one embodiment, such as further depicted in
Although generally shown as a turbofan gas turbine engine, the exemplary embodiment of the propulsion system 10 shown in
Referring now to
In one embodiment, the exit angle 139 of the plurality of turbine nozzle airfoils 121 is approximately 80 degrees or less relative to the reference plane 172. In another embodiment, the exit angle 139 is between approximately 65 and approximately 80 degrees relative to the reference plane 172. In yet another embodiment, the exit angle 139 is between approximately 70 and approximately 80 degrees relative to the reference plane 172. In another embodiment, the exit angle 139 and the nozzle angle 133 are within approximately 20 degrees relative to one another. In still another embodiment, the exit angle 139 and the nozzle angle 133 are approximately equal.
The turbine nozzle 125, or more specifically, the plurality of turbine nozzle airfoils 121, may further define a turbine nozzle inlet angle 137 relative to the reference plane 172. In one embodiment, the inlet angle 137 is less than or approximately equal to the exit angle 139. In another embodiment, the inlet angle 137 is approximately equal to or less than the nozzle angle 133. For example, the nozzle assembly 128 defining the nozzle angle 133 may induce a bulk swirl of the fuel/oxidizer mixture 132 through the combustion chamber 122. The combustion gases 138 may at least partially flow at least along the circumferential direction C co-directional to the bulk swirl of the fuel/oxidizer mixture 132. However, losses may incur along the longitudinal direction L such that the combustion gases 138 approach the inlet angle 137 of the turbine nozzle 125 less than the nozzle angle 133. The turbine nozzle 125 may accelerate the flow of combustion gases 138 along the circumferential direction C across the turbine nozzle 125, egressing the turbine nozzle 125 at approximately the exit angle 139. In various embodiments, the inlet angle 137 is approximately equal to or less than the nozzle angle 133, the exit angle 139, or both. In still various embodiments, the exit angle 139 is approximately 80 degrees or less relative to the reference plane 172. As such, the nozzle angle 133 may be approximately 80 degrees or less, and the inlet angle 137 of the turbine nozzle 125 may be approximately equal to a bulk swirl angle at the upstream end of the turbine nozzle 125, such as due to losses as the combustion gases 138 flow along the longitudinal direction L.
The nozzle assembly 128 generally provided in
Referring now to
In various embodiments, the plurality of RDC inlet airfoils 105 defines a pre-diffuser or exit guide vane structure of the RDC system 100. In other embodiments, the plurality of RDC inlet airfoils 105 defines a guide vane structure of the RDC system 100 disposed within the exhaust section 106 defining a turbine section 29, such as generally provided in
In various embodiments, the plurality of RDC inlet airfoils 105 defines an inlet angle 196 relative to the reference plane 172. In one embodiment, the inlet angle 196 is greater than zero degrees and approximately 80 degrees or less relative to the reference plane 172. In another embodiment, the inlet angle 196 and the nozzle angle 133 are within approximately 20 degrees relative to one another. In yet another embodiment, the inlet angle 196 and the nozzle angle 133 are approximately equal.
Referring now to
Referring back to
Referring now to
Referring now to
Embodiments of the propulsion system 10 including the bulk swirl RDC system 100 generally provided herein may increase a bulk swirl of the combustion gases 138 within the combustion chamber 122 of the RDC system 100, thereby reducing a length of the turbine nozzle or altogether eliminating the turbine nozzle, thereby enabling direct fluid communication of the combustion gases 138 from the combustion chamber 122 to the first turbine rotor 131, and reducing a length of the propulsion system 10. Reducing the length of or eliminating the turbine nozzle may improve overall propulsion system efficiency and performance, such as by reducing part counts, length, weight, and improving thermodynamic efficiency by reducing an amount of cooling oxidizer removed from combustion and energy release.
This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.
