Patent Application
20190271268
The present subject matter relates generally to gas turbine engines and systems of continuous detonation.
Gas turbine engine designers and manufacturers are generally challenged to improve fuel consumption, increase thrust output, and reduce weight to improve engine efficiency and performance. Known gas turbine engines generally define a Brayton Cycle and include deflagrative combustion systems to burn a fuel/air mixture and produce combustion gas products which travel at relatively slow rates and constant pressure within a combustion chamber. Such combustion systems generally include vane structures at an exit of a compressor section or an inlet of a turbine section (or an exit of a combustion section) to condition a flow of air to the combustion system to improve thermodynamic efficiency and combustor/engine operability and performance (e.g., mitigate lean blow out, reduce combustion hot spots, improve radial flow velocity of gases into and exiting the combustion chamber, etc.). However, these structures increase engine weight, such as via increased lengthwise dimensions and increased part quantities. Still further, these structures may limit engine reliability, such as to require maintenance (e.g., turbine nozzles or vanes). Nonetheless, these structures are generally known as necessary to produce gas turbine engines of a level of performance and operability required in the art and throughout the industry.
As such, there is a need for gas turbine engines that further improve fuel consumption, increase thrust output, and reduce weight to further improve engine 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 turbine engine including a compressor rotor and a rotating detonation combustion (RDC) system. The compressor rotor includes a compressor airfoil defining a trailing edge disposed within a core flowpath of the turbine engine. The core flowpath defines a radial distance between an outer radius and an inner radius at the compressor rotor. The RDC system includes an outer wall and an inner wall each extended along a lengthwise direction and defining a detonation chamber therebetween. The RDC system further includes a strut defining a nozzle assembly and a fuel injection opening providing a flow of fuel to the detonation chamber. The compressor rotor provides a flow of oxidizer in direct fluid communication to the nozzle assembly of the RDC system.
In various embodiments, the fuel injection opening is defined at a first lengthwise distance from the trailing edge of the compressor airfoil equal to or less than approximately nine times the radial distance of the core flowpath. In one embodiment, the strut of the RDC system defines an upstream end defined at a second lengthwise distance less than the first lengthwise distance.
In one embodiment, the radial distance of the core flowpath is defined approximately at the trailing edge or downstream of the compressor airfoil.
In another embodiment, the compressor rotor defines a downstream-most rotor of a compressor section of the turbine engine proximate to the RDC system.
In various embodiments, the strut is disposed at an acute nozzle angle relative to a reference radial plane extended from an axial centerline of the turbine engine. In one embodiment, the nozzle angle is between approximately zero degrees and approximately 85 degrees relative to the axial centerline of the turbine engine. In still various embodiments, the compressor airfoil defines an exit angle relative to the axial centerline of the turbine engine, and wherein a sum of the exit angle and the nozzle angle is between approximately zero degrees and approximately 85 degrees. In one embodiment, the compressor airfoil defines the exit angle within approximately 20 degrees of the nozzle angle.
In various embodiments, the turbine engine further includes a pressure vessel surrounding the outer wall and the inner wall of the RDC system. A cooling passage is defined between the pressure vessel and the RDC system. In one embodiment, the pressure vessel comprises an outer vessel wall and an inner vessel wall. The outer vessel wall and the inner vessel wall are each extended along the lengthwise direction around the outer wall and the inner wall of the RDC system. In another embodiment, the cooling passage is in fluid communication with the core flowpath between the trailing edge of the compressor airfoil and the strut, and a flow of oxidizer is provided from the core flowpath and the cooling passage via an opening in at least one of the outer wall or the inner wall.
In various embodiments, the turbine engine further includes a turbine rotor including a turbine airfoil disposed downstream of the RDC system in direct fluid communication with the detonation chamber. In one embodiment, the core flowpath defines a turbine radial distance at the turbine rotor between an outer turbine radius at the turbine rotor and an inner turbine radius at the turbine rotor. The leading edge of the turbine airfoil defines a turbine lengthwise distance from the detonation chamber exit equal to or less than approximately five times the turbine radial distance of the core flowpath. In another embodiment, the RDC system defines a radial gap between the outer wall and the inner wall, and the strut defines a downstream end defined adjacent to the detonation chamber. The RDC system defines a detonation chamber length between the downstream end and the detonation chamber exit equal to or less than approximately five times the radial gap. In still another embodiment, the trailing edge of the compressor rotor and the leading edge of the turbine rotor define a lengthwise distance equal to or less than approximately nine times the radial gap. In still yet another embodiment, the turbine airfoil of the turbine rotor defines a turbine exit angle relative to a reference radial plane extended from an axial centerline of the turbine engine. The turbine exit angle is between approximately zero and approximately 85 degrees. In another embodiment, the turbine exit angle is within approximately 20 degrees of a nozzle angle relative to the reference radial plane.
In various embodiments, the turbine engine further includes a guide vane disposed between the compressor rotor and the RDC system along the lengthwise direction. The guide vane includes a fin extended at least partially along a circumferential direction to dispose a flow of oxidizer to the RDC system. In one embodiment, one or more of the fins is disposed at an acute angle relative to the lengthwise direction such as to dispose a flow of oxidizer to a cooling passage defined at least partially around the RDC system.
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:
Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention.
Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents.
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 propulsion system or vehicle, and refer to the normal operational attitude of the propulsion system or vehicle. For example, with regard to a propulsion system, forward refers to a position closer to a propulsion system inlet and aft refers to a position closer to a propulsion system 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 gas turbine engine including a rotating detonation combustion (RDC) system are generally provided. The embodiments provided herein may improve engine efficiency and performance via the pressure-gain detonation combustion systems generally shown and described herein relative to a compressor section and a turbine section. The embodiments of the engine generally shown and described herein may improve fuel consumption, increase thrust output, and decrease pressure losses. The embodiments generally provided may further reduce engine complexity, thereby improving maintainability and reduce part quantity and weight, thereby improving cost of operation and reliability. The embodiments of the engine provided herein generally obviate one or more of a diffuser dump region, a pre-diffuser or compressor guide vane structure, or a turbine nozzle or vane assembly generally provided in known gas turbine engines. For example, the embodiments of the engine generally provided herein include the RDC system as providing a fuel injection structure that further conditions, guides, or otherwise orients a flow of oxidizer from a compressor section to a detonation chamber. As another example, the fuel injection structure of the RDC system may further provide a bulk swirl of a flow of detonation gases from the detonation chamber to a turbine section such as to obviate a turbine nozzle or vane assembly upstream of a turbine rotor. As such, a lengthwise dimension of the gas turbine engine may be reduced such as to further reduce weight, increase thrust output, decrease pressure losses, improve fuel consumption, and enable integration of higher thrust gas turbine engines into smaller apparatuses.
Referring now to
The engine 10 may generally define a turbofan, a turboprop, turbojet, or a turboshaft engine configuration, including marine and industrial gas turbine engines and auxiliary power units. The engine 10 includes a compressor rotor 110 and a turbine rotor 210 each coupled together in rotational dependency via a driveshaft 310. The compressor rotor 110, the turbine rotor 210, and the driveshaft 310 may together define a spool of the engine 10. For example, the spool may be a high pressure (HP) spool of the engine 10. The compressor rotor 110 described herein defines a downstream-most rotor of a compressor section 21 of the engine 10. For example, the exemplary compressor rotor 110 defined herein is most proximate to the RDC system 100 relative to one or more other rotors further upstream within the compressor section 21. Still further, the turbine rotor 210 described herein defines an upstream-most rotor of a turbine section 31 of the engine 10. For example, the exemplary turbine rotor 210 defined herein is most proximate to the RDC system 100 relative to one or more other rotors further downstream within the turbine section 31.
The engine 10 generally provided in
Referring now to
The RDC system 100 includes a nozzle assembly 120 disposed downstream of the compressor rotor 110. For example, the nozzle assembly 120 of the RDC system 100 is disposed in direct fluid communication downstream of the compressor rotor 110. As another example, the nozzle assembly 120 of the RDC system 100 is defined downstream of the compressor rotor 110 such as to obviate a diffuser dump region relative to conventional gas turbine engines. In still various embodiments, the nozzle assembly 120 defines a compressor exit guide vane and/or pre-diffuser and a fuel injection system providing a fuel/oxidizer mixture to a downstream detonation chamber 115.
The RDC system 100 includes an outer wall 101 and an inner wall 102 each extended along the lengthwise direction L. The detonation chamber 115 is defined between the outer wall 101 and the inner wall 102. The nozzle assembly 120 includes a nozzle wall 121. In various embodiments the nozzle wall 121 defines a convergent-divergent nozzle, such as defining a decreasing cross sectional area toward the upstream end of the RDC system followed by an increasing cross sectional area toward the downstream end of the RDC system proximate to the detonation chamber 115. In one embodiment, the nozzle wall 121 is defined circumferentially around the axial centerline 12 of the engine 10. In another embodiment, the RDC system 100 defines a plurality of the nozzle assembly 120 including a plurality of the nozzle wall 121 each defining an orifice through which oxidizer or fuel/oxidizer mixture flows into the detonation chamber 115.
The nozzle assembly 120 further includes a strut 105 defined along the radial plane 13 extended from the axial centerline 12. In various embodiments, the strut 105 defines a plurality of vanes disposed in circumferential arrangement through the core flowpath 90. The nozzle assembly 120 further defines a fuel injection opening 108 through which a flow of liquid or gaseous fuel (or combinations thereof) mixes with a flow of oxidizer, such as shown schematically by arrows 82, from the compressor section 21 (
In various embodiments, the strut 105 defines, at least in part, a flowpath through the nozzle assembly 120 at which the flow of oxidizer 82 mixes with a flow of fuel. In one embodiment, the strut 105 defines, at least in part, a contoured flowpath structure such as to guide or orient the flow of oxidizer 82, the flow of fuel, and/or the flow of fuel/oxidizer mixture 232 along the circumferential direction C (
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 230. For a detonation combustor, such as the RDC system 100 disclosed herein, the combustion of the fuel/oxidizer mixture 232 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 and convection. 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 fuel/oxidizer mixture 232, increasing such fuel/oxidizer mixture 232 above a self-ignition point. On the other side, energy released by the combustion contributes to the propagation of the detonation shockwave 230. Further, with continuous detonation, the detonation wave 230 propagates around the combustion chamber 115 in a continuous manner, operating at a relatively high frequency. Additionally, the detonation wave 230 may be such that an average pressure inside the combustion chamber 115 is higher than an average pressure within typical combustion systems (i.e., deflagration combustion systems). Accordingly, the region 234 behind the detonation wave 230 has very high pressures.
Referring back to
It should be appreciated that in one embodiment, the first lengthwise distance 97 is generally defined to a center point of the fuel injection opening 108, such as shown schematically via fuel injection reference plane 103. However, in other embodiments, a difference in distance from a center point of the fuel injection opening 108 to a surrounding wall defining the fuel injection opening 108 is small enough such as to be within the approximation language as used herein. As such, various embodiments of the first lengthwise distance 97 may alternatively be defined to the fuel injection reference plane 103 defined at a wall defining the fuel injection opening 108 (i.e., a radius from a center point of the fuel injection opening 108, or a major axis or a minor axis from a center point of the fuel injection opening 108, etc., to a surrounding wall), at a centerline, center point, or center plane of the fuel injection opening 108, or combinations thereof.
It should further be appreciated that in various embodiments of the engine 10, the RDC system 100 may define one or more fuel injection openings 108 at a plurality of locations along the lengthwise direction L. As such, it should be appreciated that the first lengthwise direction L is generally understood as relative to a forward or upstream-most fuel injection opening 108 along the lengthwise direction L relative to the trailing edge 112 of the compressor airfoil 111.
In still various embodiments of the RDC system 100, the strut 105 of the nozzle assembly 120 defines an upstream end 106 defined at a second lengthwise distance 95 less than the first lengthwise distance 97. The second lengthwise distance 95 is defined from the trailing edge 112 of the compressor rotor 110 to the upstream end 106 of the strut 105 of the RDC system 100.
Referring now to
In one embodiment, a sum of the exit angle 94 from the compressor airfoil 111 and the nozzle angle 107 of the strut 105 is between approximately zero degrees and approximately 85 degrees relative to the radial plane 13. For example, zero degrees relative to the radial plane 13 is along the radial direction from the axial centerline 12. In another embodiment, the compressor airfoil 111 defines the exit angle 94 within approximately 20 degrees of the nozzle angle 107.
Referring still to
Referring now to
Referring still to
In still various embodiments of the engine 10, the trailing edge 112 of the compressor rotor 110 and the leading edge 212 of the turbine rotor 210 defines a lengthwise distance 195 equal to or less than approximately nine times the radial gap 191. In one embodiment, the trailing edge 112 of the compressor rotor 110 and the leading edge 212 of the turbine rotor 210 defines the lengthwise distance 195 equal to or less than approximately five times the radial gap 191.
Referring back to
The strut 105 defined at the acute nozzle angle 107 disposes the flow of fuel/oxidizer mixture 232 (shown in
Referring again to
In still another embodiment, the cooling passage 89 is in fluid communication with the core flowpath 90 between the trailing edge 112 of the compressor airfoil 111 and the strut 105. A flow of oxidizer, shown schematically by arrows 83, is provided from the core flowpath 90 and the cooling passage 89 via an opening 88 in at least one of the outer wall 101 or the inner wall 102 of the RDC system 100.
Referring now to
Referring now to
Referring still to
Referring to
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.
