Two-stage pulse detonation system

The present invention relates to a pulse detonation system, and more specifically to a two-stage pulse detonation system.

A variable cycle turbofan ramjet engine can be used to obtain aircraft flight speeds from low subsonic Mach numbers to high supersonic Mach numbers of approximately Mach 6. Known engines include a core engine system and a dual mode augmenter. The dual-mode augmenter adds additional heat to the exhaust airflow from the core engine system to increase engine thrust. The core engine system provides power to drive the fan assembly and includes compressors, combustors, high pressure turbines and low pressure turbines, typically in series axial flow relationships. The dual mode augmenter is located downstream of the core engine and receives air from the core engine and the bypass duct surrounding the core engine.

Known engines can operate over a wide range of flight speed operations when utilizing several different combustion systems. During flight speed operation from takeoff to approximately Mach 3, the core engine and engine fan system provide the pressure and amount of airflow used by the augmenter to generate engine thrust. To maintain flight speed operation between Mach 3 and Mach 6, the ram airflow is dual-mode by either shutting down the core engine system and idling the fan system or utilizing an auxiliary ram duct. Introduce in the augmenter. To maintain flight speed record above Mach 6, either a separate supersonic combustion system or scramjet is used or a separate rocket-based thrust generation system is used. A rocket-based thrust generation system is used to achieve flight speed operation in outer space. As a result, several different combustion systems will be used in engines that operate efficiently over a wide range of operating flight speeds.

Pulse detonation helps increase the available flight speed range while reducing the need for combustion system combinations by efforts to solve the drawback of requiring a combination of combustion systems to obtain a wide range of flight speeds. (Intermittent detonation) Engines have been developed.

One design example proposed for an engine with a pulse detonation system includes at least one tube pulse detonation engine (PDE). The PDE can be placed as an augmenter and / or as a main combustor. However, the tube pulse detonation system has some drawbacks. These systems use mechanical valves to help control detonation, which increases complexity and cost and limits the frequency of detonation that can be achieved. Low detonation frequencies can also adversely affect the components within the engine system due to the shock and vibration generated by the detonation. In addition, tube PDE systems do not work efficiently with commonly used aviation fuels.<patcit num="1"><text>U.S. Pat. No. 6,349,538</text></patcit><patcit num="2"><text>U.S. Pat. No. 6,477,829</text></patcit>

<p> The two-stage PDE described in the present invention meets the requirements for aviation fuel use and high frequency operation. Produce thrust using detonation, which has proven to be the most efficient way to extract energy from any fuel-air mixture that can be detonated.</p>

<p> In embodiments of the present invention, the turbofan engine comprises a two-stage pulse detonation system that allows the engine to operate efficiently and effectively over a wide range of flight speeds. The two-stage pulse detonation engine increases the frequency of operating detonation of the engine without the physical constraints caused by mechanical control valves or other turbomachinery. This increases the total engine thrust and extends the expected life of the components. The two-stage pulse detonation system includes a pre-combustor and one or more shape resonators that provide additional thrust to the engine. The pre-combustor can be any system that makes it easier to detonate the fuel / oxidizing material mixture. Examples of pre-combustors include constant pressure combustors, pyrolysis systems and catalytic systems.</p><p> During operation, the rich fuel-air mixture is burned under constant pressure in the pre-combustor to produce CO and H.₂Produces a mixture of detonable ingredients such as. Following the pre-combustor, the resulting air-fuel mixture is then guided into either the two-dimensional or three-dimensional shape resonator with additional air to create detonation conditions and detonate the air-fuel mixture within the shape resonator. cause. In some cases, an ignition source is also added to the air-fuel mixture in the resonator to facilitate detonation. This ignition source can be a spark, an external shock, or the like. Detonation of the fuel / air mixture in the shape resonator produces a hot combustion gas with high temperature and pressure. This high-temperature combustion gas is directed outward from the engine to generate thrust.</p>

The advantages, properties and various additional features of the invention will become more apparent by considering exemplary embodiments of the invention outlined in the drawings.

The present invention will be described in more detail with reference to the accompanying drawings, but this description does not limit the technical scope of the invention.

FIG. 1 is a side sectional view of a turbofan engine 10 including a two-stage pulse detonation system 12 with at least one shape resonator. In this case, the PDE system is located as an augmenter, i.e. behind the engine turbine. FIG. 2 is a cross-sectional view of the two-stage pulse detonation augmenter 13 shown in FIG. 1 taken along line 2-2. In one embodiment, the turbofan engine 10 is an F110 / 129 engine available from General Electric, Inc., located in Cincinnati, Ohio. The engine 10 has an axis or centerline 14 that extends approximately longitudinally, which centerlines extend in the anterior direction 16 and the posterior direction 18. The engine 10 includes a core engine 30 with a high pressure compressor 34, a combustor 36, a high pressure turbine 38 and an output turbine or low pressure turbine 39, all arranged in an axial flow relationship in series. In another embodiment, the engine 10 includes a core fan assembly (not shown).

In yet another embodiment, the two-stage system can also be placed between the compressor and the turbine. In this case, the two-stage PDE 12 replaces the combustor 36 in FIG. Yet another embodiment includes an engine with a two-stage PDE system that acts as the main combustor 36 and also as an augmenter.

The two-stage pulse detonation system 12 is located downstream of both the core engine 30 and the engine fan system to receive bypass air from the engine fan system. The two-stage pulse detonation system 12 generates high-frequency detonation of the fuel / air mixture to generate high-temperature and high-pressure gas, and ejects this gas from the engine to generate thrust. The operation of the two-stage pulse detonation system will be described in more detail later. The two-stage pulse detonation system 12 includes an inlet side 70 with a precombustor 60, a group of at least one shape resonator 62, an outlet side 72 and a shock focusing subsystem 74. In one embodiment, one pre-combustor 60 supplies a detonable air-fuel mixture to one or more shape resonators 62. The inlet side 70 is located upstream of the exit side 72 and surrounds the engine central body 76 in the circumferential direction.

As used herein, "pulse detonation engine" is understood to mean any device or system that causes both pressure and speed increases due to a series of repeated or pseudo-detonations within the device. I want to be. "Pseudo-detonation" is a supersonic turbulent combustion process that produces a pressure increase and velocity increase greater than the pressure increase and velocity increase caused by the defragmentation wave. A typical embodiment of a pulse detonation engine is a detonation (detonation) in which a means of igniting a fuel / oxidizing material mixture, such as a fuel / air mixture, and a pressure wave front that begins in the ignition process are combined. ) Includes a detonation chamber that produces waves. Each detonation or pseudo-detonation is initiated either by external ignition such as spark discharge or laser pulse, by gas dynamic processes such as shock focusing, self-ignition, or by another detonation (crossfire) ( Detonate). The geometry of the detonation chamber is such that the pressure rise of the detonation wave causes combustion products to be released from the pulse detonation engine exhaust to generate thrust. As is known to those of skill in the art, pulse detonation can be achieved in many types of detonation chambers, including detonation tubes, shock tubes, resonant detonation cavities and annular detonation chambers.

The shock focusing subsystem 74 includes at least one nozzle 78 extending from the shape resonator 62 and a system outlet side 72. Nozzle 78 allows the fuel and air mixture detonated within the shape resonator 62 to flow to the outlet side 72 to generate additional thrust. In one embodiment, each nozzle 78 has a circular cross section and the shock focusing subsystem 74 has a circular cross section contour. In another embodiment, the shock focusing subsystem 74 has a non-circular cross-sectional contour. In another embodiment, the nozzle 78 has a non-circular cross section with, for example, a polygonal or rectangular cross section. The nozzle 78 extends from the shape resonator 62 to the exhaust nozzle 84. The exhaust nozzle 84 is arranged downstream of the two-stage pulse detonation system 12 and the nozzle 78.

During operation, airflow flows into engine 10 and fuel is introduced into core engine 30. The air and fuel are mixed and ignited in the core engine 30 to generate hot combustion gas. Specifically, the pressurized air from the high-pressure compressor 34 is mixed with the fuel in the combustor 36 and ignited, whereby combustion gas is generated. This combustion gas drives the high-pressure turbine 38, and the high-pressure turbine 38 drives the high-pressure compressor 34. The combustion gas is discharged from the high-pressure turbine 38 into the low-pressure turbine 39. The core airflow is discharged from the low pressure turbine 39.

A small portion of the bypass airflow is flushed into the precombustor 60, where additional fuel is added to the airflow. The fuel used throughout the engine is any known hydrocarbon fuel that can be used for aircraft engine applications, including, for example, JP-4, Jet-A and JP-8. The amount of fuel injected into the pre-combustor 60 is such that a rich fuel-air ratio can be obtained. In one embodiment, the fuel-to-air ratio of this rich mixture is in the range of 2-3. In another embodiment, an additional gas, such as oxygen, is used in the air-fuel mixture of the pre-combustor 60.

After mixing the rich fuel / air mixture, the resulting mixture is burned in the pre-combustor 60 by the constant pressure combustion method. During the constant pressure combustion process, some of the hydrocarbon fuels are burned, but some are partially oxidized, and this partial oxidation chemically alters some of the fuels, resulting in CO and H.₂Produces additional detonable components such as.

In one embodiment, the pressure in the pre-combustor that burns the fuel / air mixture is approximately 3 atm. A rich fuel-air mixture is obtained by forming a premixed turbulent jet flame using at least one dual premixed swirler.

The burned fuel mixture and the detonable component are guided from the premixer 60 into the shape resonator 62 through the divergent nozzle 64 at or near the speed of sound. Prior to flowing into the shape resonator 62, the burned fuel mixture is further mixed with the secondary gas injected into the stream flowing through the nozzle 64. The newly formed air-fuel mixture exits the convergence-divergence nozzle 64 and flows into the shape resonator 62 at high pressure. In one embodiment, this secondary gas is air.

In one embodiment, the amount of secondary air added to the burned pre-combustor stream is such that the resulting fuel-air ratio is 1. By setting the final fuel / air ratio of the fuel / air mixture flowing into the shape resonator to 1, it becomes possible to achieve the maximum fuel consumption when the fuel / air mixture is detonated in the shape resonator 62. The amount of secondary air added to the burned pre-combustor flow is such that a fuel-to-air ratio of 1 can be obtained with respect to the total amount of air used in the pre-combustor and resonator.

As the new fuel / air mixture exits the convergent-divergence nozzle 64, this outflow is radially oriented towards the axis of symmetry of the shape resonator 62. The flow from the nozzle 64 collides along the axis of symmetry of the shape resonator 62 to form a stagnation region with high pressure and temperature. The resulting pressure rise in the shape resonator 62 creates a pressure wave that blocks the flow exiting the nozzle 64. In one embodiment, the formed high pressure and high temperature region acts as a starting (detonation) part for detonating the fuel / air mixture. The generated detonation propagates downstream through the resonator outlet nozzle 78 and flows out of the nozzle 78 at the outlet side 72 to generate additional thrust.

The frequency and intensity of detonation of the fuel / air mixture in the shape resonator 62 is determined by the geometry, composition, temperature and pressure as the fuel / air mixture exits the nozzle 64 and flows into the shape resonator 62. All of these parameters are optimized for high frequency detonation, efficient fuel consumption within the shape resonator 62 and sufficient ignition wait time for detonation.

Before the flow is detonated in the resonator, the flow discharged from the nozzle 64 is blocked due to the pressure ratio before and after the nozzle. This limits the flow into the shape resonator 62. In addition, this flow blockage helps allow detonation of the fuel / air mixture within the shape resonator 62. When detonation occurs, the detonation creates a high pressure downstream of the nozzles 64, which can significantly reduce the flow through these nozzles or even reverse the direction of the flow. When the fuel / air mixture is detonated in the shape resonator 62, the detonated fuel / air mixture flows out of the shape resonator 62 and travels along the resonator outlet nozzle 78. When the detonated fuel / air mixture flows out of the shape resonator, a pressure drop occurs in the shape resonator 62. This pressure drop allows the nozzle flow to flow into and re-occlude the shape resonator 62, allowing cycle iterations.

In one embodiment, some of the pressure waves generated by the inflow of the fuel / air mixture from the nozzle 64 into the shape resonator 62 are also radiated toward the reflective surface 68 of the shape resonator 62. The pressure wave directed to the reflecting surface 68 is reflected by the shape of the reflecting surface 68 and directed to a high-pressure and high-temperature focus. The high pressure and high temperature created at the focal point helps detonate the fuel / air mixture emitted from the nozzle 64. This detonation propagates along the resonator outlet nozzle 78 and exits at the outlet side 72.

In one embodiment, the focus of the pressure wave reflected from the reflection surface 68 coincides with the stagnation region formed by the flow of fuel / air mixture from the convergence-divergence nozzle 64. The convergence of the reflected pressure wave and the stagnation region of the flow flowing in from the nozzle 64 create a high-pressure high-temperature region at the convergence point, causing detonation of the fuel / air mixture. The detonation is then guided along the resonator outlet nozzle 78 and exits the nozzle 72 at the outlet side 72 to generate additional thrust.

It should be noted that the engine 10 shown in FIGS. 1 and 2 has a plurality of shape resonators 62 and resonator outlet nozzles 78. However, in one embodiment, a combination of a single shape resonator 62 and a resonator outlet nozzle 78 is used.

FIG. 3 is a cross-sectional view of a part of the two-stage pulse detonation system 100. As mentioned above, the hydrocarbon fuel is mixed with air to form a fuel-rich air-fuel mixture, which is burned at a constant pressure in the pre-combustor 102. The partially burned fuel mixture and the components remaining after combustion flow through the convergence-divergence nozzle 104 and are further mixed with the secondary air exiting port 112 at the convergence-divergence nozzle 104. Create an equal fuel-air ratio. The resulting air-fuel mixture flows into the shape resonator 106 having a cylindrical pressure wave reflecting surface 108. The two-stage pulse detonation engine further includes a resonator outlet nozzle 110 for sending the detonated fuel / air mixture to the outlet of the engine (not shown). The shape of the pressure wave reflecting surface 108 has a two-dimensional curved surface having a focal point F in which the reflected pressure waves in the shape resonator 106 are focused. This shape is the shape of a part of a cylinder with its long axis passing through the focal point F.

In one embodiment, the curvature of the pressure wave reflecting surface 108 is three-dimensional. In this embodiment, the surface shape is part of a sphere with a focal point F.

The shape resonator 106, reflective surface 108 and resonator outlet nozzle 110 are made of heat resistant aerospace material that can withstand high and high temperature conditions as a result of repeated detonation of the fuel / air mixture within the shape resonator 106.

The convergence-divergence nozzle 104 is configured as an annular space formed in the radial direction surrounding the shape resonator 106. The fuel / air air-fuel mixture flows out radially from the nozzle 104 toward the center of the shape resonator 106, forming a high-pressure, high-temperature stagnation region where detonation and detonation occur. In another embodiment, the nozzle 104 is configured as a plurality of individual channel openings, in which case at least one channel opening is placed on top of the shape resonator 106 and at least one other channel opening is on the shape resonator 106. Installed on the bottom.

In one embodiment, a plurality of individual individual nozzles 104 are radially distributed around the perimeter of the shape resonator 106. Each of the nozzles directs the fuel / air mixture from the pre-combustor 102 towards the central region of the shape resonator 106. In another embodiment, a number of individual nozzles 104 direct the fuel / air mixture from the pre-combustor 102 toward the center of the shape resonator 106, while the remaining nozzles 104 direct the secondary air to the center of the shape resonator 106. Turn to.

FIG. 4 is a cross-sectional view of the convergence-divergence nozzle 200 surrounding the shape resonator 202. The convergence-divergence nozzle 200 is configured to allow supersonic or near-supersonic flow from the pre-combustor 204 to the shape resonator 202. The nozzle 200 is made of heat resistant aerospace material. In one embodiment, the inner surface of the nozzle 200 is coated with a heat insulating film.

In one embodiment, the dimensions of the nozzle inlet 206, throat 208 and outlet 210 and the dimensional relationships between them limit the amount of pressure drop that occurs in the flow from the pre-combustor 204 into the shape resonator 202. Optimized as Typical geometric parameters of such a radial inward convergence-divergence nozzle include the shape contour of the wall, which includes the wall angle, convergence-divergence length, throat width. And its radial position.

By optimizing these geometric dimensions, the pressure loss between the inlet 206 and the outlet 210 of the nozzle 200 is reduced. The reduction in pressure drop allows the two-stage pulse detonation system to operate at a lower total pressure and increases the overall performance of the two-stage PDE system. The lower pressure drop at nozzle 200 allows the use of lower working pressure within the precombustor 204, resulting in improved cycle efficiency. In addition, the reduction in pressure drop at the nozzle reduces the ignition delay prior to the initiation of detonation within the shape resonator 202.

As mentioned above, the detonation frequency and pre-detonation time of the fuel / air mixture in the shape resonator 202 is related to the geometry of the resonator and the composition, temperature and pressure of the fuel / air mixture. As the rate of pressure rise within the shape resonator 202, i.e. the rate of reaching a given level, increases, the delay between sequential detonations decreases. Therefore, the frequency with which the two-stage pulse detonation system can operate increases as the pressure loss in the convergence-divergence nozzle 200 decreases. In addition, the pressure in the premixer 204 is maintained at a lower total pressure without adversely affecting detonation in the resonator 202, thus increasing overall system efficiency.

A plurality of secondary air ports 212 installed in the convergence-divergence nozzle 200 are arranged to inject secondary air into the fuel / air mixture flowing from the pre-combustor 204 to the shape resonator 202. In one embodiment, a fuel-to-air ratio of 1 within the shape resonator 202 is used to achieve optimum detonation efficiency and frequency. This is achieved by injecting an appropriate amount of secondary air from the air port 212 into the burned fuel-rich air-fuel mixture that exits the pre-combustor 204 and enters the shape resonator 202. Normal atmosphere is ejected from air port 212. In one embodiment, the high pressure airflow from the core engine 30 is used as secondary air. In another embodiment, at least one of these ports can be fueled instead of air.

The air port 212 is located on the convergent side of the convergent-divergence nozzle 200 to inject secondary air upstream into the fuel / air mixture. The air ports 212 are installed along the nozzle 200 at equal intervals from each other. The pressure of the secondary airflow from the air port prevents backflow into the air port 212 of the fuel / air mixture and prevents the formation of a recirculation zone in the nozzle 200 near the air port 212. Adjusted to.

In addition, the air port 212 tilts upstream into the fuel / air mixture flow to ensure optimized secondary air mixing from the precombustor 204 into the fuel / air mixture. are doing. In one embodiment, the angle A for injecting secondary air into the nozzle stream is in the range 0 ° to 45 ° with respect to the perpendicular N to the surface 214 of the nozzle 200.

In one embodiment, secondary air is injected into the nozzle stream upstream at 45 ° with respect to the perpendicular N to the surface 214 of the nozzle 200.

The spacing, size, and number of air ports 212 along the nozzle 200 optimizes the mixing of secondary air within the nozzle 200 with the fuel / air mixture from the precombustor 204. In one embodiment, the spacing between the edges of the air port 212 is equal to the diameter of the air port 212.

Further, in one embodiment, instead of providing the air port 212 on the convergent side of the nozzle 200, the air port 212 is located on the divergent surface 216 on the divergent side of the nozzle 200. In this embodiment, the pressure required for the secondary airflow is lower than the pressure required in the configuration where the air port 212 is on the convergent side of the nozzle 200. Convergence-When placed on the divergence side of the divergence nozzle 200, the secondary airflow exits the air port 212 at an angle and pressure that maximizes the mixture of the secondary airflow with the fuel / air mixture. In one embodiment, the secondary airflow exits the divergent air port of the nozzle 200 at an angle of 0 ° to 90 ° with respect to the perpendicular to the surface of the divergent portion of the nozzle and is the flow of the fuel / air mixture. Go inside. Bringing secondary air into the fuel / air mixture on the divergent side of nozzle 200 allows the secondary air system (not shown) to operate at a lower total working pressure.

In one embodiment, the secondary air is preheated prior to flowing into the nozzle 200. Preheating the secondary air helps the fuel / air mixture from the precombustor 204 maintain a high temperature or become hot, which helps initiate detonation within the shape resonator 202.

In another embodiment, the secondary air port 212 is installed on one or more surfaces of the convergence-divergence nozzle 200. A series of air ports 212 are provided on the convergent surface 214 of the nozzle and a second series of air ports are provided on the divergent surface 216 of the nozzle. In yet another embodiment, the air port arranged on the converging side of the nozzle 200 is arranged on the side opposite to the side on the diverging side where the air port 212 is arranged. This helps to achieve the optimum mixing of secondary air into the fuel / air mixture.

Further, in one embodiment, the air port 212 is located on the opposing sides of the convergent-divergent nozzle 200 on the convergent side, or on the opposite sides of the nozzle 200 on the divergent side, or a combination thereof. Placed on both.

FIG. 5 is a cross-sectional view of one embodiment of the shape resonator 300 used in a two-stage pulse detonation engine. The figure also shows the pre-combustor 302 and the convergent-divergent nozzle 304. The pressure wave reflecting surfaces 306, 308 of the shape resonator 300 form a wedge shape with an angle α in order to achieve the best pressure recovery of the pressure waves reflected from these surfaces. Achieving optimal pressure recovery of the reflected pressure wave helps maximize the detonation of the fuel / air mixture within the shape resonator 300. In one embodiment, the angle α between the surfaces 306, 308 is in the range of 45 ° to 90 °. In another embodiment, the angle α between the reflecting surfaces 306, 308 is 55 °.

In order to align the focus of the pressure wave reflected from the surfaces 306, 308 with the hot and high pressure stagnation region formed by the fuel / air mixture exiting the nozzle 304, the reflecting surfaces 306, 308 are the openings of the convergent-divergent nozzle 304. Placed at a distance D from. In one embodiment, the reflective surfaces 306, 308 start at the opening of the nozzle 304 and have a distance D of zero.

In another embodiment, the shape resonator 300 has a three-dimensional conical shape.

FIG. 6 is a cross-sectional view of another shape resonator 400 having a pressure wave reflecting surface 402 having a parabolic shape. In one embodiment, the reflective surface 402 has a three-dimensional configuration in which the curved reflective parabolic surface 402 is rotated about the radial direction of the center line of the shape resonator 400. The shape of the reflective surface and the distance D match the fuel / air mixture stagnation region in the shape resonator 400 so as to achieve the best pressure recovery of the pressure wave reflected from the reflective surface 402. So that it is optimized.

FIG. 7 is a cross-sectional view of another shape resonator 500 having a pressure wave reflecting surface 502 having another parabolic shape. In one embodiment, the reflective surface 502 has a three-dimensional configuration in which the curved reflective surface 502 is rotated around the center line of the shape resonator 500 in the radial direction. The shape of the reflective surface and the distance D match the fuel / air mixture stagnation region in the shape resonator 500 so as to achieve the best pressure recovery of the pressure wave reflected from the reflective surface 502. So that it is optimized.

FIG. 8 is a cross-sectional view of another shape resonator 600 having a pressure wave reflecting surface 602 having a flat surface. Distance D is optimized to achieve the best pressure recovery of the pressure wave reflected from the reflective surface 602 and to match the reflected pressure wave with the fuel / air mixture stagnation region in the shape resonator 600. To.

FIG. 9 is a cross-sectional view of another shape resonator 700 having a pressure wave reflecting surface 702 having a plurality of facets. In one embodiment, the reflective surface 702 has a three-dimensional configuration in which the faceted reflective surface 702 is rotated about the radial direction of the center line of the shape resonator 700. The shape and distance D of the reflective surface 700 aligns the reflected pressure wave with the fuel / air mixture stagnation region in the shape resonator 700 so as to achieve the best pressure recovery of the pressure wave reflected from the reflective surface 702. Optimized to let you.

FIG. 10 is a cross-sectional view of another shape resonator 800 having a plurality of pressure wave reflecting surfaces 802 having a cylindrical shape. In one embodiment, the reflective surfaces 802, 804 have a three-dimensional configuration such that the reflective surfaces 802, 804 are spherically formed and radially arranged around the centerline of the shape resonator 800. The shape and distance D of the reflective surfaces 802, 804 are such that the best pressure recovery of the pressure waves reflected from the reflective surfaces 802, 804 is achieved, and the reflected pressure waves are shaped. The fuel / air mixture in the resonator 800 Optimized to match the stagnation area.

In another embodiment, the reflective surfaces 802, 804 have a parabolic shape. In yet another embodiment, these reflective surfaces are flat or have faceted surfaces.

In the embodiment of the resonator configuration (FIGS. 5 to 10) described above, the secondary port is added in the convergent nozzle portion, in the divergent nozzle portion, in both portions of the nozzle, and in any side of the nozzle. ..

The two-stage pulse detonation system described above includes at least one pre-combustor and a shape resonator that produces additional engine thrust without the need for mechanical valves or other turbomachinery. As a result, engines using this two-stage pulse detonation system achieve higher detonation frequencies while using standard aviation hydrocarbon fuels, thus requiring different fuels without the adverse effects caused by lower operating frequencies. It is possible to obtain a higher thrust without using. The result is an overall engine system that allows the engine to operate with high efficiency and performance over a wide range of driving flight speeds. In addition, the two-stage pulse detonation engine system described above can be used in turbofan, turbojet and ramjet engine structures.

Although the present invention has been described with respect to various specific embodiments, it will be apparent to those skilled in the art that the invention can be practiced with the technical ideas of the claims and modifications within the technical scope. It should be noted that the reference numerals described in the claims are for the sake of comprehension and do not limit the technical scope of the invention to the examples.

<figref num="1">Side sectional view of a turbofan engine including a two-stage pulse detonation system.</figref><figref num="2">Sectional view of the pulse detonation augmenter used in the two-stage pulse detonation system shown in Figure 1 taken along line 2-2.</figref><figref num="3">Sectional cross-sectional view of a two-stage pulse detonation system.</figref><figref num="4">Sectional view of the convergence-divergence nozzle used within a two-stage pulse detonation system.</figref><figref num="5">Sectional view of another shape resonator used within a two-stage pulse detonation system.</figref><figref num="6">Sectional view of yet another shape resonator used within a two-stage pulse detonation system.</figref><figref num="7">Sectional view of yet another shape resonator used within a two-stage pulse detonation system.</figref><figref num="8">Sectional view of yet another shape resonator used within a two-stage pulse detonation system.</figref><figref num="9">Sectional view of yet another shape resonator used within a two-stage pulse detonation system.</figref><figref num="10">Sectional view of yet another shape resonator used within a two-stage pulse detonation system.</figref>

Code description

10 Turbofan engine 12 Two-stage pulse detonation system 13 Two-stage pulse detonation augmenter 30 Core engine 34 High-pressure compressor 36 Combustor 38 High-pressure turbine 39 Low-pressure turbine 60 Pre-combustor 62 Shape resonator 64 Convergence-divergence nozzle 68 Pressure wave reflection Surface 74 Shock Focus Subsystem 76 Engine Central Body 78 Resonator Outlet Nozzle 84 Exhaust Nozzle