Patent Application
20200149496
The present subject matter relates generally to a combustor of an engine, such as a rotating detonation engine.
A rotating detonation engine includes an annulus with an inlet end through which a fuel and air mixture enters and an outlet end from which exhaust exits. A detonation wave travels in a circumferential direction of the annulus and consumes the incoming fuel and air mixture. The burned fuel and air mixture (e.g., combustion gases) exits the annulus and is exhausted with the exhaust flow.
The detonation wave provides a high-pressure region in an expansion region of the combustion system. Rotating detonation pressure gain combustion systems are expected to operate at much higher frequencies than other pressure gain combustion concepts such as pulse detonation combustors.
Maintaining a rotating detonation wave within rotating detonation combustors during low power conditions of the engines, as well as selectively controlling and/or adjusting the operating conditions present technical challenges. For example, when a rotating detonation engine is operating at an idle condition (e.g., not generating enough propulsive force to propel the engine or a vehicle that includes the engine), the detonations rotating within the combustor of the engine may dissipate or be extinguished.
Aspects of the present embodiments are summarized below. These embodiments are not intended to limit the scope of the present claimed embodiments, but rather, these embodiments are intended only to provide a brief summary of possible forms of the embodiments. Furthermore, the embodiments may encompass a variety of forms that may be similar to or different from the embodiments set forth below, commensurate with the scope of the claims.
In one aspect, a combustion system includes an annular tube disposed between an inner wall and an outer wall, the annular tube extending from an inlet end to an outlet end; at least one fluid inlet disposed in the annular tube proximate the inlet end, the fluid inlet providing a conduit through which fluid flows into the annular tube; at least one outlet disposed in the annular tube proximate the outlet end; and at least one inlet fluid plenum disposed upstream of the fluid inlet. The inlet fluid plenum includes at least one reflective surface.
In another aspect, a combustion system includes an annular tube disposed between an inner wall and an outer wall; a diverging section disposed axially forward of the annular tube, the diverging section comprising an inner diverging wall and an outer diverging wall; at least one throat area disposed axially forward of the diverging section; and a fluid inlet plenum disposed axially forward of the throat area, the fluid inlet plenum including at least one reflective surface.
In another aspect, a combustion system includes an annular tube disposed between an inner wall and an outer wall; at least one throat area disposed axially forward of the annular tube; and a fluid inlet plenum disposed axially forward of the throat area, the fluid inlet plenum including at least one reflective surface, wherein at least one rotating detonation wave travels from the reflective surface through the throat area.
These and other features, aspects, and advantages of the present disclosure will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein:
Unless otherwise indicated, the drawings provided herein are meant to illustrate features of embodiments of the disclosure. These features are believed to be applicable in a wide variety of systems comprising one or more embodiments of the disclosure. As such, the drawings are not meant to include all conventional features known by those of ordinary skill in the art to be required for the practice of the embodiments disclosed herein.
In the following specification and the claims, reference will be made to a number of terms, which shall be defined to have the following meanings.
The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.
“Optional” or “optionally” means that the subsequently described event or circumstance may or may not occur, and that the description includes instances where the event occurs and instances where it does not.
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, a value modified by a term or terms, such as “about” 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. Here and throughout the specification and claims, range limitations may be combined and/or interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise.
As used herein, the term “axial” refers to a direction aligned with a central axis or shaft of a gas turbine engine or alternatively the central axis of a propulsion engine, a combustor, and/or internal combustion engine. An axially forward end of the gas turbine engine or combustor is the end proximate the fan, compressor inlet, and/or air inlet where air enters the gas turbine engine and/or the combustor. An axially aft end of the gas turbine engine or combustor is the end of the gas turbine or combustor proximate to the engine or combustor exhaust where combustion gases exit the engine or combustor. In non-turbine engines, axially aft is toward the exhaust and axially forward is toward the inlet.
As used herein, the term “circumferential” refers to a direction or directions around (and tangential to) the circumference of an annulus of a combustor, or for example the circle defined by the swept area of the turbine blades. As used herein, the terms “circumferential” and “tangential” are synonymous.
As used herein, the term “radial” refers to a direction moving outwardly away from the central axis of the gas turbine, or alternatively the central axis of a propulsion engine. A “radially inward” direction is aligned toward the central axis moving toward decreasing radii. A “radially outward” direction is aligned away from the central axis moving toward increasing radii.
The air inlet plenum 21 is defined between an inlet portion 34, a first inner reflective surface 42, a first outer reflective surface 44, an inner sidewall 46, an outer sidewall 52, a second inner reflective surface 54, a second outer reflective surface 56, an inner aft wall 58, and an outer aft wall 64. The inlet portion 34 forms a transition between the axial inlet 50 and the air inlet plenum 21. Each of the axial inlet 50 and the air inlet plenum 21 may be axisymmetric about an inlet centerline 32. The first inner and outer reflective surfaces 42, 44 extend radially inward and radially outward, respectively, from the aft end of the inlet portion 34. Each of the first inner and outer reflective surfaces 42, 44 may be contoured. The first inner and outer reflective surfaces 42, 44 are coupled to the inner and outer sidewalls 46, 52, respectively. Each of the inner and outer sidewalls 46, 52 may be linear and extends radially inward and axially aft toward the throat area 51. The radially inward and aft ends of each of the inner and outer sidewalls 46, 52 are coupled to the second inner and outer reflective surfaces 54, 56 respectively.
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In operation, pressure waves propagate within the annulus area 13 as a result of rotating detonation, and travel primarily circumferentially and toward the aft wall 60. However, portions of pressure waves propagating within the annulus area 13 may also travel back through the throat area 51 toward the axial inlet 50. The portions of the pressure waves that travel back through the throat area 51 may radially, circumferentially, and axially expand within the air inlet plenum 21. Due to the geometry of the air inlet plenum 21, pressure waves that expand therein may reflect off the first inner and outer reflective surfaces 42, 44, as well as the second inner and outer reflective surfaces 54, 56 back toward the throat area 51 (through which they travel back toward the aft wall 60). As a result of portions or one or more pressure waves being reflected from within the air inlet plenum 21 back through the throat area 51 toward the aft wall 60, losses within the combustor 2 associated with propagating pressure and/or detonation waves may be reduced, minimized, and/or eliminated. In some configurations in accordance with the present embodiments, the axial inlet 50 may be angled at least partially in a circumferential and/or radial direction such that pressure waves propagating through the air inlet plenum 21 that travel through the inlet portion 34 may also be reflected off one or more internal walls and/or surfaces of the axial inlet 50 (i.e., in addition to the portions that are reflected off the reflective surfaces 42, 44, 54, 56), back toward the throat area 51.
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The primary fuel injector 26, as well as any other fuel injectors, may disperse fuel through holes and/or orifices that are circular, elliptical, slotted, and/or other suitable shapes. A minimum dimension (i.e., diameter, width, minor axis, etc.) of the holes and/or orifices in each of the primary fuel injector 26 and/or other fuel injectors may be from about 3 to about 30 mils (i.e., thousandths of an inch). In other embodiments, the minimum dimension of the holes and/or orifices may be from about 5 to about 20 mils. In other embodiments, the minimum dimension of the holes and/or orifices may be from about 8 to about 17 mils. In other embodiments, the minimum dimension of the holes and/or orifices may be from about 10 to about 15 mils.
A rotating detonation wave resulting from combustion of a fuel-air mixture from the one or more primary fuel injectors 26 and/or inlet air 18 may travel circumferentially around the combustor 2 as it travels the axial length of the combustor tube (or annular tube) 70, from the inlet end 4 to the outlet end 6. The magnitude of the rotating detonation wave may begin to dissipate as it propagates circumferentially and axially (forward and aft) through the combustor 2. Reflecting pressure and/or detonation waves that travel toward the axial inlet 50 back through the throat area 51 toward the aft end 6 may enhance the performance of the rotating detonation combustor 2, while in operation. As such, the geometry of the air inlet plenum 21, and the reflecting surfaces 42, 44, 54, 56 thereof may allow the combustor 2 to remain in stable operation while simultaneously augmenting the performance of the combustor 2.
In operation, each of the embodiments disclosed herein may include multiple detonation waves simultaneously propagating in a circumferential (and axially aft) direction such that they wrap around the annulus 13 as they move from an inlet end 4 to an outlet end 6. Chemistry and combustor dynamics, as well as other factors, may limit the minimum size of both the combustor 2 as well as the area and/or volume of the annulus 13 due to a minimum amount of time required for the rotating denotation wave to travel around the annulus. As such, the area of the annulus 13, the overall radius of the combustor 2, and/or the overall axial length of the combustor 2 may all be adjusted to ensure the chemistry considerations as well as other factors such as combustor dynamics, aerodynamics, thermal management, and other considerations are all balanced accordingly. In addition, it may be desirable for the combustor 2 to have a non-circular shape in order to increase the distance around the annulus 13 that the rotating detonation wave may travel, while simultaneously allowing the axial length of the combustor 2 to be decreased.
As used herein, “detonation” and “quasi-detonation” may be used interchangeably. Typical embodiments of detonation chambers include a means of igniting a fuel/oxidizer mixture, for example a fuel/air mixture, and a confining chamber, in which pressure wave fronts initiated by the ignition process coalesce to produce a detonation wave. Each detonation or quasi-detonation is initiated either by external ignition, such as spark discharge or laser pulse, or by gas dynamic processes, such as shock focusing, autoignition or by another detonation via cross-firing. The geometry of the detonation chamber is such that the pressure rise of the detonation wave expels combustion products out of the detonation chamber exhaust to produce a thrust force, as well as for other purposes such as flow control actuation. In addition, rotating detonation combustors are designed such that a substantially continuous detonation wave is produced and discharged therefrom. Detonation may be accomplished in a number of types of detonation chambers, including detonation tubes, shock tubes, resonating detonation cavities, and annular detonation chambers.
Each of the embodiments disclosed herein include fuel being combusted in the presence of an oxidizer. Fuel mixes with an oxidizer during or prior to the combustion process. The embodiments disclosed herein include air as one possible oxidizer. However, other oxidizers such as straight oxygen (i.e., pure oxygen) are also possible. In various conditions, oxygen may be a preferred oxidizer over air. In other conditions, air may be the preferred oxidizer. As used herein, the terms “oxygen” and “pure oxygen,” may include gas that is at least about 80% oxygen by mass. In some embodiments, the oxidizer may be at least about 90% oxygen by mass. In other embodiments, the oxidizer may be about 93% to about 99.3% oxygen by mass. In other embodiments, the oxidizer may be greater than about 99.3% oxygen by mass. (By comparison, air is about 21% oxygen, about 78% nitrogen and about 1% other gases). Other oxidizers other than oxygen and air are also possible. In embodiments that use an oxidizer other than air, those embodiments will include the corresponding system components including, for example, an oxidizer inlet, an oxidizer supply line, an oxidizer supply, an oxidizer flow control mechanism, an oxidizer flow modulator, and/or a second oxidizer inlet.
Each of the embodiments disclosed herein include a source of ignition, which may be in the form of a spark igniter and/or via autoignition (i.e., via heated inner and outer walls 10, 8, and/or heated inner and outer diverging walls 28, 30 which have absorbed heat from the combustion process), as well as via volumetric ignition. Some embodiments may include multiple sources of ignition. For example, in some embodiments, at least one spark igniter may be used during some operating conditions and then ignition may transition to autoignition and/or volumetric ignition at other operating conditions.
The present embodiments include an aircraft, an engine, a combustor, and/or systems thereof which include rotating detonation combustion. The embodiments presented herein operate on a kilohertz range (1000 Hz to 1000 kHz), which is faster than the 100 Hz operating frequency of previous pulse detonation actuators (PDA) and/or pulse detonation engines (PDE). As such, the embodiments presented herein may provide a more continuous and less pulsed combustion gas jet discharging from the combustor exhaust 62 compared to previous pulse detonation actuators (PDA).
The present embodiments offer both high operating frequency and significant control authority, which provides benefits in numerous practical applications, such as engine exhaust thrust vectoring for vehicle control or boundary layer separation control for aircraft lift enhancement and drag reduction. The present embodiments may also be used as enhancements or combustion systems for supersonic and/or hypersonic applications, for example, in scramjet engines, as well as in subsonic gas turbine applications. The present embodiments take advantage of a more compact and/or power dense combustion system. The present embodiments may be used as the primary combustion system for engines such as gas turbine engines. The present embodiments may be used as the secondary, tertiary, and/or auxiliary combustion systems for engines such as gas turbine engines, and/or other components of an aircraft or of other applications.
Exemplary applications of the present embodiments may include high-speed aircraft, separation control on airfoils, flame holders, flame stability, augmenters, propulsion, flight stability, flight control, as well as other uses.
Although specific features of various embodiments of the present disclosure may be shown in some drawings and not in others, this is for convenience only. In accordance with the principles of the present disclosure, any feature of a drawing may be referenced and/or claimed in combination with any feature of any other drawing.
This written description uses examples to disclose the embodiments of the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the embodiments described herein 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 have 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 language of the claims.
