Combined cycle propulsion system for hypersonic flight

Abstract

A combined cycle propulsion system for an aircraft includes a compressor-fed internal combustion engine and a multi-mode supersonic engine. The multi-mode supersonic engine includes an adjustable inlet section, a combustion section disposed downstream of the adjustable inlet section and including a first combustor portion having at least one rotating detonation combustor and a second combustor portion having a supersonic combustion-type combustor, and an adjustable exhaust nozzle section disposed downstream of the combustion section. The at least one rotating detonation combustor functions as an igniter for the supersonic combustion-type combustor.

Description

Technical Field

[0001] This disclosure relates to a combined cycle propulsion system for an aircraft. Background Technology

[0002] Some conventional supersonic aircraft are known to include ramjet engines for achieving supersonic speeds, while other types of aircraft may include scramjet engines for achieving even higher supersonic or hypersonic speeds. Both ramjet and scramjet engines rely on the aircraft's forward momentum to deliver supersonic airflow into an inlet duct, which uses shock waves to compress the supersonic airflow. Ramjet engines typically include an isolator downstream of the inlet duct, through which the compressed airflow from the inlet passes before being supplied to the ramjet engine's combustor. The isolator reduces the velocity of the inlet airflow to subsonic speeds, causing combustion in the ramjet engine's combustion chamber to occur at subsonic speeds. Combustion products from the ramjet engine's combustor are then expelled through nozzles, thereby increasing the aircraft's speed. Scramjet engines similarly have an inlet duct into which a supersonic airflow is input, and where compressed airflow passes through an isolator before being fed into the scramjet combustor. Combustion in a scramjet engine subsequently occurs in the supersonic flow of inlet air within the scramjet engine's combustor. Attached Figure Description

[0003] The features and advantages of this disclosure will become apparent from the following description of various exemplary embodiments, as illustrated in the accompanying drawings, wherein the same reference numerals generally denote the same, functionally similar and / or structurally similar elements.

[0004] Figure 1 This is a schematic diagram of an exemplary combined cycle propulsion system according to an embodiment of the present disclosure.

[0005] Figure 2 This is a cross-sectional side view of an exemplary turbojet engine according to aspects of this disclosure.

[0006] Figure 3 Based on aspects of this disclosure Figure 1 A magnified view of the adjustable entry section cropped at 125 in the detail view.

[0007] Figure 4 Based on aspects of this disclosure Figure 1 A magnified view of the first burner section of the combustion zone, taken at point 126 in the detail view.

[0008] Figure 5 Based on aspects of this disclosure Figure 1A detailed view of the enlarged view of the second burner section and the adjustable exhaust nozzle section of the combustion zone, taken at point 154.

[0009] Figure 6 Based on aspects of this disclosure Figure 5 A schematic cross-sectional view taken at point 6-6 on the plane.

[0010] Figure 7 This is an enlarged view of an alternative second burner section according to another aspect of this disclosure.

[0011] Figure 8 According to another aspect of this disclosure, a combined cycle propulsion system is in Figure 1 A schematic cross-sectional view taken at point 8-8 on the plane.

[0012] Figure 9 This is a schematic partial cross-sectional side view of a combined cycle propulsion system according to another aspect of this disclosure.

[0013] Figure 10 It is based on another aspect of this disclosure. Figure 9 A partial cross-sectional rear view taken at point 10-10 on the plane.

[0014] Figure 11 It is based on another aspect of this disclosure. Figure 10 The top view of the cross section taken at plane 11-11 depicts the multi-mode supersonic engine.

[0015] Figure 12 Depicting according to this disclosure Figure 11 Another arrangement in terms of aspects.

[0016] Figure 13 This is a flowchart of the process steps of a method for operating a combined cycle propulsion system according to aspects of this disclosure.

[0017] Figure 14 An example of an aircraft in which a combined cycle propulsion system can be implemented according to aspects of this disclosure is depicted. Detailed Implementation

[0018] Various embodiments are discussed in detail below. Although specific embodiments are discussed, they are for illustrative purposes only. Those skilled in the art will recognize that other components and configurations can be used without departing from the spirit and scope of this disclosure.

[0019] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0020] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows out, and "downstream" refers to the direction in which the fluid flows in.

[0021] The features, advantages, and embodiments of this disclosure will be apparent or obvious upon consideration of the following detailed description, accompanying drawings, and claims. Furthermore, the following detailed description is exemplary and intended to provide further explanation without limiting the scope of the claimed disclosure.

[0022] Some conventional supersonic aircraft are known to include ramjet engines for achieving supersonic speeds, while other types of aircraft may include scramjet engines for achieving even higher supersonic or hypersonic speeds. Both ramjet and scramjet engines rely on the aircraft's forward momentum to deliver supersonic airflow into an inlet duct. The inlet duct compresses the airflow, and the compressed airflow then passes through an isolator downstream of the inlet before being supplied to the ramjet engine's combustor. The isolator also reduces the velocity of the inlet airflow to subsonic speeds before combustion in the combustion chamber. Combustion products from the ramjet engine's combustor are then expelled through nozzles, thereby increasing the aircraft's speed. Scramjet engines similarly have an inlet duct in which supersonic airflow is input and compressed via a shock wave generated within the isolator, then flows through the isolator before being supplied to the ramjet engine's combustor. Combustion in a scramjet engine occurs in the supersonic flow of inlet air leading to the scramjet combustor. One problem with scramjet engines is the no-start condition, in which the supersonic airflow breaks down, putting it outside the design range of the scramjet combustor. The no-start condition is likely to occur more frequently at lower Mach numbers.

[0023] This disclosure addresses the aforementioned problems by providing a combined cycle propulsion system comprising a compressor-fed internal combustion engine and a multi-mode supersonic engine having a first combustor section with a rotating detonation combustor (RDC) and a second combustor section, the second combustor section including a supersonic combustion type combustor, such as a scramjet combustor. Combustion products from the RDC in the first combustor section can be used as an ignition source for the supersonic combustion type combustor. The inclusion of an RDC in the first combustor section of the multi-mode supersonic engine more or less allows for ramjet operation of the multi-mode supersonic engine. Once the aircraft reaches a specific flight speed, the scramjet combustor can operate, and combustion products from the RDC can be used to ignite the scramjet combustor. Since the scramjet combustor does not operate at lower Mach numbers, the likelihood of a failure to ignite in the scramjet combustor is reduced.

[0024] Now refer to the attached diagram, Figure 1 This is a schematic diagram of an exemplary combined cycle propulsion system according to aspects of this disclosure. Figure 1 As can be seen in the diagram, the combined cycle propulsion system 2 includes a compressor-fed internal combustion engine 4 and a multi-mode supersonic engine 6. The compressor-fed internal combustion engine 4 can be either a gas turbine engine or a compressor-fed rotating detonation combustion (RDC) engine. As will be described below, the multi-mode supersonic engine 6 may include a combustion section 8 having a first combustor portion 62 and a second combustor portion 66. The first combustor portion 62 may include, for example, an RDC combustor for ramjet operation, and the second combustor portion 66 is a supersonic combustion type combustor 85, such as a scramjet combustor for scramjet operation. In the following description, the compressor-fed internal combustion engine 4 will be described as constituting a gas turbine engine comprising a gas turbine inlet section 12, a gas turbine compressor section 14, a gas turbine combustion section 16, a gas turbine turbine section 18, and a gas turbine exhaust nozzle section 20 that may include an afterburner 21. Each of the aforementioned components of the compressor-fed internal combustion engine 4 will be described in more detail below. However, typically, the components of the compressor supplying the internal combustion engine 4 can be defined around the centerline axis 10, and the compressor supplying the internal combustion engine 4 can extend longitudinally along the centerline axis 10.

[0025] Extending from the front end 36 of the gas turbine compressor section 14 is a translational spike 22. The translational spike 22 can be longitudinally translated along the centerline axis 10 in the upstream direction 30 and the downstream direction 32 by a spike actuator 23. When the translational spike 22 is translated in the upstream direction 30, it is translated to a first (extended) position 24, such that the circumferential apex 34 of the translational spike 22 disengages from the inner surface 28 of the gas turbine inlet section 12. A gap 38 is thus formed between the circumferential apex 34 of the translational spike 22 and the inner surface 28 of the gas turbine inlet section 12 to allow the inlet airflow 40 to enter the gas turbine inlet section 12 and reach the compressor inlet 29 of the gas turbine compressor section 14. On the other hand, when the translatable peak 22 translates in the downstream direction 32, the translatable peak 22 translates to a second (retracted) position 26, such that the circumferential apex 34 of the translatable peak 22 engages with the inner surface 28 of the gas turbine inlet section 12 to restrict or prevent the inlet airflow 40 from entering the gas turbine inlet section 12. Of course, in addition to the translatable peak 22, other technologies and structures (such as doors, slots, flaps, etc.) can also be implemented to control (i.e., allow or restrict) the inlet airflow 40 allowed to enter the gas turbine inlet section 12.

[0026] Still referencing Figure 1The compressor-supplied internal combustion engine 4 may be surrounded or enclosed by an outer casing 42 extending circumferentially around the centerline axis 10. The casing 42 may include a cooling flow passage 44 extending from an upstream end 48 of the gas turbine inlet section 12 to a downstream end 50 of the gas turbine exhaust nozzle section 20. The cooling flow passage 44 allows cooling airflow 46 to flow through it to provide cooling to the outer surface 140 of the multi-mode supersonic engine 6. The cooling flow passage 44 may also include a plurality of air exhaust openings 52 that allow a portion of the cooling airflow 46 to enter various components of the multi-mode supersonic engine 6 from the cooling flow passage 44. The air exhaust openings 52 may also be arranged to allow some of the inlet airflow 40, compressed airflow 54, and / or combustion gases 56 supplied from the compressor to the internal combustion engine 4 into the multi-mode supersonic engine 6.

[0027] exist Figure 1 In this respect, the multi-mode supersonic engine 6 is shown to include a first combustor section 62 connected in series with a second combustor section 66, wherein the second combustor section 66 is arranged downstream of the first combustor section 62. Additionally, in Figure 1 In terms of the first burner section 62 and the second burner section 66, they are depicted as arranged circumferentially around the compressor supplying the internal combustion engine 4. The multi-mode supersonic engine 6 includes an adjustable inlet section 58, which will... Figure 3 This is described in more detail below. However, in brief, the adjustable inlet section 58 may be an adjustable converging inlet that compresses the supersonic inlet airflow 74 as it passes through, producing a compressed inlet airflow 76 that may typically include a shock wave. Downstream of the adjustable inlet section 58, the isolator section 60 organizes the shock wave to produce an isolator airflow 78, which may then be reduced to a subsonic airflow 78 (a) or may be a supersonic flow 78 (b), depending on the operating state of the multi-mode supersonic engine 6. When the isolator airflow 78 is reduced to a subsonic airflow, the now subsonic isolator airflow 78 enters the inlet end 70 of the first combustor section 62. A first portion 81 of the isolator airflow 78 enters the combustor 79 of the first combustor section 62, while a second portion 80 of the isolator airflow 78 enters the combustor bypass passage 84 as a bypass airflow 80 to bypass the combustor 79 of the first combustor section 62. When the isolator airflow 78 is a subsonic airflow 78(a), the bypass airflow 80 can also be a subsonic airflow. On the other hand, when the isolator airflow 78 is a supersonic airflow 78(b), the bypass airflow 80 can also be a supersonic airflow leading to the second burner section 66.

[0028] As burner 79, the first burner section 62 is considered to include a rotating detonation burner (RDC) 82, to which a burner inlet airflow 81 is supplied, and fuel 147 (as will be described in more detail below) Figure 4 ) through fuel injection port 148 ( Figure 4 The gas is injected into the RDC 82 and detonates, thereby generating a rotating detonation combustion wave. The burner bypass passage 84 is arranged radially outside the rotating detonation burner 82 to allow the burner bypass gas flow 80 to flow through it to the inlet section 64 of the second burner section 66.

[0029] Inlet section 64 draws in combustor bypass airflow 80 from combustor bypass passage 84 and combustion gas 86 from rotary detonator combustor 82. Inlet section 64 can also function as an isolator. Inlet section 64 is in fluid communication with outlet end 65 of rotary detonator combustor 82, allowing combustor bypass airflow 80 and combustion gas 86 to flow downstream into second combustor section 66, where fuel is injected by fuel injector 88 to generate a fuel-air mixture within second combustor section 66. In operation, the fuel-air mixture can be ignited by utilizing combustion gas 86 from rotary detonator combustor 82, or, depending on the fuel used in second combustor section 66, the fuel-air mixture can be automatically ignited. Therefore, rotary detonator combustor 82 is configured to act as an igniter for second combustor section 66 to initiate ignition when switching between operation of first combustor section 62 and second combustor section 66 of multi-mode supersonic engine 6. Once second combustor section 66 is successfully ignited, rotary detonator combustor 82 can be shut off or can continue operating until the desired space speed is reached. When the second burner section 66 is ignited, the use of the rotary detonation burner 82 as an igniter can further reduce the magnitude and suddenness of the pressure drop, thereby reducing the possibility of a failure to start. Therefore, the second burner section 66 can operate at a higher Mach number, making a failure to start less likely. The combustion product gas 90 then expands and is discharged through the adjustable exhaust nozzle section 68.

[0030] Figure 2 This is a schematic partial cross-sectional side view of an exemplary gas turbine engine according to aspects of this disclosure. As described above, the compressor supplying the internal combustion engine 4 can be a turbojet engine, or any similar engine including a mechanically driven compressor to compress air for combustion. Figure 2 An example is depicted where a turbojet engine acts as a compressor-supplying internal combustion engine 4, which can be implemented in the combined cycle propulsion system 2 of this disclosure. For example... Figure 2As shown, the compressor supplying the internal combustion engine 4 includes a gas turbine compressor section 14, a gas turbine combustion section 16, a gas turbine turbine section 18, and a gas turbine exhaust nozzle section 20. The gas turbine compressor section 14 may include a low-pressure compressor section 92, an intermediate-pressure compressor section 94, and a high-pressure compressor section 96. The low-pressure compressor section 92 includes a plurality of low-pressure compressor rotors 98 and a plurality of low-pressure compressor stators, each low-pressure compressor stator including a plurality of low-pressure compressor stator blades 100 circumferentially arranged between each low-pressure compressor rotor 98. Similarly, the intermediate-pressure compressor section 94 includes a plurality of intermediate-pressure compressor rotors 102 and a plurality of intermediate-pressure compressor stators, each intermediate-pressure compressor stator having a plurality of intermediate-pressure compressor stator blades 104 circumferentially arranged between each intermediate-pressure compressor rotor 102. Furthermore, the high-pressure compressor section 96 includes multiple high-pressure compressor rotors 106 and multiple high-pressure compressor stators, each stator having multiple high-pressure compressor stator blades 108 arranged circumferentially between each high-pressure compressor rotor 106. Inlet gas flow 40 enters the low-pressure compressor section 92, where initial compression occurs, and then enters the intermediate-pressure compressor section 94, where the inlet gas flow is further compressed before entering the high-pressure compressor section 96 to compress the inlet gas flow 40, resulting in a compressed gas flow 54. The compressed gas flow 54 enters the diffuser 110 before entering the gas turbine combustion section 16.

[0031] The gas turbine combustion section 16 includes a gas turbine combustor 112, which can be any type of combustor known to be implemented in a gas turbine engine, including a knock-type combustor or a detonation combustor. For example, the gas turbine combustor 112 can be a knock-type combustor, which can be a canister combustor or an annular knock-type combustor. Alternatively, the gas turbine combustion section 16 may include a rotating detonation combustor 121. Figure 9 and Figure 13 ).like Figure 2The annular detonation combustor shown may include multiple vortex cups or fuel-air mixer assemblies 113 connected to fuel nozzles 115 to inject a fuel-air mixture into a combustion chamber 117, where the fuel-air mixture is ignited and burned to produce combustion gases 56. The combustion gases 56 then flow from a gas turbine combustor 112 through a turbine nozzle, which may include multiple blades 114 arranged circumferentially around the turbine nozzle. The combustion gases 56 then flow to a gas turbine turbine section 18. The gas turbine turbine section 18 may include a high-pressure turbine 116, an intermediate-pressure turbine 118, and a low-pressure turbine 120. Multiple turbine stator blades 122 may be arranged circumferentially around the gas turbine turbine section 18 between the high-pressure turbine 116 and the intermediate-pressure turbine 118, and between the intermediate-pressure turbine 118 and the low-pressure turbine 120. The combustion gases 56 passing through the gas turbine turbine section 18 cause rotation of each of the high-pressure turbine 116, the intermediate-pressure turbine 118, and the low-pressure turbine 120.

[0032] Drive shaft assembly 124 connects the gas turbine turbine section 18 to the gas turbine compressor section 14. Drive shaft assembly 124 typically includes a plurality of concentrically arranged drive shafts (not shown) within the drive shaft assembly. High-pressure turbine 116 is driven to high-pressure compressor section 96 via a high-pressure drive shaft (not shown, but included in drive shaft assembly 124) to drive each high-pressure compressor rotor 106. Similarly, intermediate-pressure turbine 118 is driven to drive each intermediate-pressure compressor rotor 102 via an intermediate-pressure drive shaft (not shown, but included in drive shaft assembly 124). Additionally, low-pressure turbine 120 is driven to drive each low-pressure compressor rotor 98 via a low-pressure drive shaft (not shown, but included in drive shaft assembly 124).

[0033] In the operation of an aircraft including a combined cycle propulsion system 2 with the aforementioned compressor-supplied internal combustion engine 4, the compressor-supplied internal combustion engine 4 is generally operable for all flight operations of the aircraft up to a specific airspeed. For example, an aircraft with the compressor-supplied internal combustion engine 4 may be able to reach airspeeds exceeding Mach 1, particularly when the translational spike 22 can be used as a convergent-diffusion nozzle to compress the inlet airflow 40 entering the gas turbine inlet section 12, such that the inlet airflow 40 entering the compressor inlet 29 is below Mach 1. As the launch vehicle approaches the operating airspeed limit of the compressor-supplied internal combustion engine 4, the multi-mode supersonic engine 6 can be engaged to become operable. After successful ignition of the multi-mode supersonic engine 6, the compressor-supplied internal combustion engine 4 can be deactivated by, for example, shutting off the fuel flow to the compressor-supplied internal combustion engine 4 and retracting the translational spike 22 to limit or prevent the inlet airflow 40 from entering the gas turbine inlet section 12 of the compressor-supplied internal combustion engine 4. The combined cycle propulsion system 2 can then operate in conjunction with the multi-mode supersonic engine 6. As will be explained in more detail below, the first operating mode of the multi-mode supersonic engine 6 can be operated with the first combustor section 62 engaged, and then later, after approaching another airspeed limit of the first combustor section 62, the second combustor section 66 of the multi-mode supersonic engine 6 can be operated, and the first combustor section 62 can be stopped from operating by cutting off the fuel to the first combustor section 62.

[0034] According to this disclosure, Figure 3 Is Figure 1 A magnified view of the adjustable inlet section 58, cropped at point 125 in the detailed view. (See attached image.) Figure 3 As shown, the adjustable inlet section 58 may include an outer wall portion 59 connected to the actuator 61. The actuator 61 may adjust an angle 63 of the outer wall portion 59 to adjust the size of the inlet 67 of the adjustable inlet section 58. For example, the actuator 61 may adjust the angle 63 to translate the outer wall portion 59 to a position (as shown in outer wall portion 59(a)) to increase the size of the inlet 67. Additionally, or alternatively, the adjustable inlet section 58 may include a convergent-diffusion nozzle portion 69 defining a throat 83 and may be connected to the actuator 61 and / or actuator 71. The actuator 61 and / or actuator 71 may be operable to translate the convergent-diffusion nozzle portion 69 in the longitudinal direction (L) or to adjust the height of the throat 83. As will be described below, the adjustable inlet section 58 can be adjusted via actuator 61 and / or actuator 71 to regulate the intake airflow 74 of the multi-mode supersonic engine based on any of the various operating modes of the multi-mode supersonic engine 6.

[0035] Figure 4 Based on aspects of this disclosure Figure 1The enlarged view of the first burner section 62, taken at detail view 126, is shown. The first burner section 62 is generally referred to as a ramjet burner, wherein the first burner section 62 is typically operated using a subsonic inlet airflow entering the burner, although the first burner section 62 can also be operated using a supersonic inlet airflow. Therefore, the first burner section 62 can be referred to as a ramjet burner. The first burner section 62 is defined between an outer wall 128 and an inner wall 130, wherein the inner wall 130 defines the outer wall of the cooling flow passage 44. Figure 3 In this section, both the outer wall 128 and the inner wall 130 extend circumferentially around the centerline axis 10. The first burner section 62 includes a rotary detonation burner 82. The rotary detonation burner 82 includes a housing 132 which may define a generally triangular shape in the longitudinal direction and extends circumferentially around the centerline axis 10. Of course, the shape of the housing 132 may be other shapes and may be based on shock waves that can be reflected and / or induced in the bypass airflow 80 passing through the burner bypass passage 84. The triangular shape of the housing 132 may at least partially form, together with the outer wall 128, a convergent-diffusion nozzle 134 within the burner bypass passage 84 through which the bypass airflow 80 passes. The housing 132 includes a downstream inner surface 136 forming the outer wall of the detonation chamber 138, while the outer surface 140 of the inner wall 130 defines the inner wall of the detonation chamber 138. The detonation chamber 138 extends circumferentially around the centerline axis 10.

[0036] An inlet 142 leading to the detonation chamber 138 is defined by the upstream inner surface 144 of the housing 132 and the outer surface 140 of the inner wall 130. Burner inlet airflow 81 enters the detonation chamber 138 from the isolator section 60 through inlet 142. The housing 132 also includes a fuel manifold 146 that supplies fuel 147 to a plurality of fuel injection ports 148 circumferentially spaced around a centerline axis 10. One or more igniters 150 extend from the housing 132 into the detonation chamber 138, wherein the igniters 150 may be circumferentially spaced around the centerline axis 10. In operation, burner inlet airflow 81 from the isolator section 60 enters the detonation chamber 138 through inlet 142. Fuel 147 is injected into the detonation chamber 138 through multiple fuel injection ports 148 to generate a fuel-air mixture within the detonation chamber 138, and the fuel-air mixture is ignited by an igniter 150 to generate a rotating detonation wave within the detonation chamber 138. The rotating detonation wave propagates circumferentially around the detonation chamber 138 and extends longitudinally through the outlet 152 of the detonation chamber 138 to exhaust combustion gases 86. When the first burner section 62 operates while the second burner section 66 is not in operation, the combustion gases 86 exiting the detonation wave at outlet 152 and the bypass airflow 80 travel through the inlet section 64 of the second burner section 66, and then, as Figure 1As shown, propulsion is provided through the second burner section 66 and the adjustable exhaust nozzle section 68. When the multi-mode supersonic engine 6 operates without the second burner section 66 in operation, the second burner section 66 is idle and only serves as a passive flow passage to the adjustable exhaust nozzle section 68.

[0037] Figure 5 Based on aspects of this disclosure Figure 1 A magnified view of the second burner section 66 and the adjustable exhaust nozzle section 68, taken at detail view 154. The second burner section 66 may be a supersonic burner, also known as a scramjet burner. The second burner section 66 is defined between an outer wall 128 and an inner wall 130, wherein the inner wall 130 defines the outer wall of the cooling flow passage 44. Figure 5 In the middle, both the outer wall 128 and the inner wall 130 extend circumferentially around the central axis 10 and define the combustion chamber 166 therebetween. Also seen is the second burner section 66, which includes a plurality of circumferentially spaced inclined fuel injectors 156. (Brief reference) Figure 6 , Figure 6 for Figure 5 A schematic cross-sectional view taken at plane 6-6 shows multiple tilting fuel injectors 156 circumferentially spaced apart, with gaps 158 between them. (Return to reference) Figure 5 Each tilting fuel injector 156 can be seen to include a fuel injector fuel passage 160, which can be connected to a fuel manifold (not shown) and a fuel injection port 162. When the second burner section 66 is in operation, fuel 164 is injected from the fuel injection port 162 into the combustion chamber 166. The fuel 164 is mixed with the bypass airflow 80 and the combustion gas 86 from the rotary detonation burner 82 of the first burner section 62. The combustion gas 86 from the rotary detonation burner 82 of the first burner section 62 can be used as an igniter (i.e., as an ignition source) for the second burner section 66. Therefore, the second burner section 66 can be operated in conjunction with the first burner section 62, and more specifically, with the rotary detonation burner 82 used as an igniter, in order to reduce the possibility of an inactive state in the second burner section 66.

[0038] Figure 5An adjustable exhaust nozzle section 68 is also depicted. It can be seen that the adjustable exhaust nozzle section 68 includes an outer exhaust nozzle portion 159 and an inner exhaust nozzle portion 161, each of which is connected to an exhaust nozzle actuator 167. The exhaust nozzle actuator 167 can actuate the outer exhaust nozzle portion 159 and the inner exhaust nozzle portion 161 to define either a diverging exhaust nozzle 163 or a converging exhaust nozzle 165 (the outer exhaust nozzle portion 159(a) and the inner exhaust nozzle portion 161(a) are shown). Therefore, in an operating state where the first combustor section 62 is operating and the second combustor section 66 is not operating, causing the multi-mode supersonic engine 6 to operate in ramjet mode, the adjustable exhaust nozzle section 68 can form a converging exhaust nozzle 165. On the other hand, in an operating state where the second combustor section 66 is operating, the adjustable exhaust nozzle section 68 is actuated to form a diverging exhaust nozzle 163. Of course, when transitioning from operating the first burner section 62 as the igniter for the second burner section 66, the adjustable exhaust nozzle section 68 transitions between the converging exhaust nozzle 165 and the diverging exhaust nozzle 163. For example, after igniting the second burner section 66 using the rotary detonation burner 82 as the igniter, the operation of the rotary detonation burner 82 of the first burner section 62 can be terminated and only the second burner section 66 can be operated, and the adjustable exhaust nozzle section 68 is actuated accordingly during this transition.

[0039] Figure 7 This is an enlarged view of an alternative second burner section 66 according to another aspect of this disclosure. Figure 7 In this aspect, multiple fuel injector blades 168 are implemented instead of the tilting fuel injector 156. The multiple fuel injector blades 168 can be coupled with... Figure 6 The tilting fuel injectors 156 shown are circumferentially spaced from each other in the same manner. Each fuel injector blade 168 may extend between the inner wall 130 and the outer wall 128 and may be aligned at an angle 174. Furthermore, each fuel injector blade 168 includes a blade fuel injector fuel passage 170 that provides a fuel flow 164 from a fuel manifold (not shown) to a fuel injector port 172 on the downstream side of the fuel injector blade 168. Each fuel injector blade 168 functions similarly to the tilting fuel injector 156 to provide fuel for supersonic combustion.

[0040] Figure 8 According to another aspect of this disclosure, a combined cycle propulsion system is in Figure 1 A schematic cross-sectional view taken at point 8-8 on the plane. Figures 1 to 7 In this respect, the first combustor section 62 of the multi-mode supersonic engine 6 is described as having a single rotating detonation combustor 82 extending circumferentially around a central axis 10. Figure 8Alternatively, the first burner section 62 includes a plurality of rotary detonation burners 176. Each of the plurality of rotary detonation burners 176 may be similar to rotary detonation burner 82, but on a smaller (individual) scale. The plurality of rotary detonation burners 176 are circumferentially spaced from each other about a central axis 10. A plurality of burner bypass flow channels 178 (i.e., gaps) are provided between the respective rotary detonation burners 176. The function of each burner bypass flow channel 178 is similar to Figure 4 The burner bypass passage 84 provides a bypass airflow 80 that bypasses the rotating detonation burner 176 and flows to the inlet section 64 of the second burner section 66 (see...). Figure 1 ). Figure 8 The aspect also describes, such as Figure 7 The fuel injector blades 168 are shown arranged circumferentially around the centerline axis 10. The fuel injector blades 168 are shown circumferentially aligned with the corresponding bypass flow passages 178. Each fuel injector blade 168 may be arranged in a second burner section 66 downstream of the rotary detonator 176, or, as will be described in more detail below, within a burner bypass flow passage 178 circumferentially adjacent to the rotary detonator 176.

[0041] refer to Figure 9 and Figure 10 , Figure 9 This is a schematic partial cross-sectional side view of a combined cycle propulsion system according to another aspect of this disclosure, and Figure 10 yes Figure 9 A partial cross-sectional view taken at point 10-10 on the plane. Figures 1 to 8 In each of the aforementioned aspects, the multi-mode supersonic engine 6 is described as extending circumferentially around the centerline axis 10. However, in Figure 9 and Figure 10 In this respect, the multi-mode supersonic engine 6 is arranged radially outward on one side of the compressor-supplied internal combustion engine 4. In this respect, the compressor-supplied internal combustion engine 4 defines a longitudinal direction (L) along the centerline axis 10 of the compressor-supplied internal combustion engine 4, a circumferential direction (C) around the centerline axis 10, and a radial direction (R) extending from the centerline axis 10, and the multi-mode supersonic engine 6 extends along the length of the compressor-supplied internal combustion engine 4. However, the multi-mode supersonic engine 6 may be shorter or longer than the length of the compressor-supplied internal combustion engine 4.

[0042] like Figure 10As shown, the multi-mode supersonic engine 6 can define a stadium-shaped profile 180, wherein the height 182 of the stadium-shaped profile 180 extends in the radial direction (R) and the length 184 of the stadium-shaped profile 180 extends perpendicular to both the longitudinal direction (L) and the radial direction (R). The multi-mode supersonic engine 6 includes a first combustor section 62 having a plurality of rotating detonation combustors 176. Figure 9 Multiple rotating detonation burners are spaced apart from each other along a length 184 of the stadium's 180-degree outline. Similar to... Figure 8 In this respect, multiple burner bypass flow channels 178 are arranged between corresponding rotary detonation burners in multiple rotary detonation burners 176. This is also similar to... Figure 8 In this respect, it can be seen that multiple fuel injector blades 168 are aligned with corresponding bypass flow channels in the burner bypass flow channel 178, and as... Figure 9 As shown, the fuel injector blades 168 can be arranged downstream of the rotary detonation burner 176. Figure 9 and Figure 10 The multi-mode supersonic engine 6 can be operated similarly to Figures 1 to 8 Regarding the operation of the adjustable inlet section 58, the first burner section 62, the second burner section 66, and the adjustable exhaust nozzle section 68. Therefore, the plurality of rotating detonation burners 176 of the first burner section 62 can be used as igniters for the second burner section 66.

[0043] Figure 11 It is based on another aspect of this disclosure. Figure 10 The top view of the cross section taken at plane 11-11 depicts the multi-mode supersonic engine 6. Figure 11 Similar to Figure 9 and Figure 10 The aspect is Figure 11 Aspects include such as Figure 10 The stadium shape outline shown is 180, and is related to... Figure 10 In the same manner, multiple rotary detonation burners 176 are spaced apart along a length 184, and multiple burner bypass flow channels 178 are included between corresponding rotary detonation burners in the multiple rotary detonation burners 176. Figure 10 As described in the text, multiple rotary detonation burners 176 share a common adjustable inlet section 58 and a common isolator section 60 downstream of the adjustable inlet section 58. However, with Figure 9 The difference between the first burner section 62 and the second burner section 66 upstream is that, Figure 11In this configuration, the first burner section 62 and the second burner section 66 are arranged in a shared combustion zone 186. Therefore, the shared combustion zone 186 includes each of a plurality of rotating detonation burners 176, a plurality of burner bypass flow channels 178 between each of the plurality of rotating detonation burners 176, and fuel injector blades 168 arranged within corresponding bypass flow channels of the plurality of burner bypass flow channels 178. Regarding... Figure 11 In terms of operation, when the first burner section 62 is not operated while the second burner section 66 is not operated, the rotary detonation burner 176 is operable to receive the burner inlet airflow 81 and the bypass airflow 80 passes through the burner bypass flow passage 178. The combustion gas 86 from the rotary detonation burner 176 mixes with the bypass airflow 80 to exit the adjustable exhaust nozzle section 68. When the second burner section 66 begins operation, the rotary detonation burner 176 acts as an igniter for the second burner section 66, causing the combustion gas 86 from the rotary detonation burner 176 to ignite the fuel 164 injected from the fuel injector port 172 of the fuel injector impeller 168. Therefore, in operation, Figure 11 The multi-mode supersonic engine 6 is similar to Figure 10 Operations in this area.

[0044] Figure 12 This disclosure describes Figure 11 Another arrangement in this regard. In Figure 11 In this arrangement, each of the multiple rotating detonation burners 176 receives an isolator airflow 78 from a common (single) adjustable inlet section 58 and a common (single) isolator section 60. Similarly, Figure 11 Each of the multiple burner bypass flow channels 178 receives isolator airflow 78 from a common (single) adjustable inlet section 58 and a common (single) isolator section 60. However, in Figure 12 In this arrangement, the multi-mode supersonic engine 6 includes multiple first combustor section inlet sections 188 and multiple corresponding isolator sections 190. Each of the multiple inlet sections 188 and each of the multiple isolator sections 190 corresponds to a corresponding rotating detonation combustor 176. Furthermore, each combustor bypass flow passage 178 includes a corresponding bypass inlet section 192 and isolator section 194 to provide airflow 196 to a second combustor section 66, which includes fuel injector blades 168 sharing the combustion section 186.

[0045] Each of the aforementioned engine types can use different types of fuel. For example, when the compressor supplies internal combustion engine 4 as a gas turbine engine, any of a variety of kerosene-based fuels, such as Jet A, Jet A-1, Jet B, JP4, JP8, etc., can be used. On the other hand, when the compressor supplies internal combustion engine including a rotary detonation burner instead of a knock-off burner, various other fuel types, such as propane, ethane, ethylene, and hydrogen, can be used in combination with air or oxygen as a fuel-air mixture. Similarly, the first burner section 62 can also use any of the aforementioned fuel types, while the second burner section 66 can use hydrogen as fuel. Therefore, the compressor supplies internal combustion engine 4 with the aforementioned first fuel type, the first burner section 62 can use a second fuel type different from the first fuel type, and the second burner section 66 can use a third fuel type different from both the first and second fuel types.

[0046] Figure 13 This is a flowchart of the process steps of a method for operating a combined cycle propulsion system according to aspects of this disclosure. Figure 14 An example of an aircraft 3 is depicted, which could be a hypersonic aircraft in which a combined cycle propulsion system 2 can be implemented. The method of operating the combined cycle propulsion system 2 can be implemented in the aircraft 3 in which the combined cycle propulsion system 2 is provided, wherein the combined cycle propulsion system 2 can correspond to... Figures 1 to 12 Any of the aforementioned aspects. In the first step 1300, the compressor is operated to supply power to the internal combustion engine 4 to obtain a first operational flight state of the aircraft. The first operational flight state may correspond to a specific airspeed of the aircraft 3, and the operation of the compressor supplying power to the internal combustion engine 4 may correspond to a flight airspeed between approximately Mach 0.9 and approximately Mach 2. For airspeeds exceeding Mach 1, in order to maintain the operation of the compressor supplying power to the internal combustion engine 4, the spike 22 may be shifted ( Figure 1 The inlet section 40 can be shifted to control the inlet airflow 40 through the gas turbine inlet section 12, such that the inlet airflow 40 at the compressor inlet 29 is a subsonic airflow. During the operation of the compressor supplying the internal combustion engine 4 in step 1300, the multi-mode supersonic engine 6 can passively allow the inlet airflow 74 to pass only through the multi-mode supersonic engine 6. Alternatively, the adjustable inlet section 58 can be actuated to close or restrict the inlet airflow 74 through the multi-mode supersonic engine 6.

[0047] In step 1301, once the first operational flight state of the aircraft 3 is reached, or as it approaches the first operational flight state, a first transition operation is initiated to activate the first supersonic operating mode of the multi-mode supersonic engine 6. The first transition operation may include actuating the adjustable inlet section 58 ( Figure 3), so as to compress the inlet airflow 74 to generate a compressed inlet airflow 76, and organize the shock wave from the compressed inlet airflow 76 via the isolator section 60, thereby obtaining the inlet air 81 entering the first burner section 62 ( Figure 1 and Figure 4 The subsonic inlet flow. Figure 1 In terms of the first burner section 62, it includes at least one rotating detonation burner 82 through which the burner inlet airflow 81 passes and a burner bypass passage 84 through which the bypass airflow 80 passes. In the first transition operation of step 1301, the operation of the first burner section (i.e., at least one rotating detonation burner 82) is initiated, as follows: Figure 4 As shown, fuel 147 is injected from fuel injection port 148 into detonation chamber 138, and the fuel-air mixture is ignited via igniter 150. As... Figure 4 As discussed above, this generates a detonation wave, and the combustion gases 86, along with the bypass flow 80, flow downstream to the second combustor section 66. The second combustor section 66 is inactive during this phase, and the combustion gases 86 and bypass flow 80 flow through the adjustable exhaust nozzle section 68. The adjustable exhaust nozzle section 68, during this operational phase of the multi-mode supersonic engine 6, limits the convergent exhaust nozzle 165 (…). Figure 5 This provides thrust. During transitional operation, the multi-mode supersonic engine 6 typically operates as a ramjet engine.

[0048] In step 1302, once the first combustor section 62 has been successfully started, a gradual winddown operation for the compressor supply to the internal combustion engine 4 can begin. The winddown operation may include terminating the operation of the gas turbine combustor 112 and compressor sections 92, 94, and 96. Furthermore, the translational spike 22 can be actuated to a second retracted position 26 to close or limit the inlet airflow 40 entering the gas turbine inlet section 12 supplied to the internal combustion engine 4.

[0049] In step 1303, a first supersonic operating mode is obtained, wherein the first burner section 62, the adjustable inlet section 58, and the adjustable exhaust nozzle section 68 are operated to obtain a second operating state of the aircraft 3. For example, as the airspeed of the vehicle increases due to the operation of the first burner section 62, the adjustable inlet section 58 can be adjusted based on the inlet airflow 74, and the adjustable exhaust nozzle section 68 can also be adjusted to increase the airspeed of the vehicle 3. The second operating state can be, for example, an airspeed of the aircraft 3 of approximately Mach 3 to approximately Mach 4. Within this aircraft airspeed range, the burner inlet airflow 81 and the bypass airflow 80 can begin to become supersonic within the first burner section 62 and can reach, for example, inlet airflow velocities greater than Mach 1 and up to approximately Mach 2 within the first burner section 62. The rotating detonation burner 82 can be operated with a supersonic inlet airflow, but as the burner inlet airflow 81 approaches, for example, Mach 2, the operation of the rotating detonation burner 82 may become less efficient and may approach its operating limits.

[0050] In step 1304, during the second operational flight state, or as the aircraft 3 approaches the second operational flight state, a second transitional supersonic operating mode is initiated to activate the second supersonic operating mode of the multi-mode supersonic engine 6. The second transitional supersonic operating mode may include actuating an adjustable inlet section 58 for scramjet engine operation, thus allowing further adjustment of the adjustable inlet section 58 to obtain a supersonic inlet flow of compressed inlet airflow 76 leading to the second combustor section 66 and a bypass airflow 80. Furthermore, the adjustable exhaust nozzle section 68 may be adjusted to transition from a converging exhaust nozzle 165 to a diverging exhaust nozzle 163. Figure 5 ), and initiate the operation of the second burner section 66. When initiating the operation of the second burner section 66 (as described above, the second burner section 66 may include a scramjet jet burner, which includes a tilting fuel injector 156 ( Figure 5 ) or fuel injector blade 168 ( Figure 7 Fuel 164 is injected into the second burner section 66. To reduce the possibility of the second burner section 66 not starting, combustion gas 86 from the rotating detonation burner 82 of the first burner section 62 is used as an igniter to start the second burner section 66. Therefore, during the second transition supersonic operating mode, both the first burner section 62 and the second burner section 66 are operating.

[0051] Once the second burner section 66 has been successfully started, in step 1305, the operation of the first burner section is gradually stopped to gradually shut down the first burner section 62. In other words, supplying fuel 147 to the rotating detonation burner 82 can terminate the operation of the rotating detonation burner 82. In step 1306, the second burner section 66 operates as a scramjet burner, and the adjustable inlet section and adjustable exhaust nozzle section operate as a scramjet engine to obtain a third operational flight state of the aircraft. The third operational flight state can be, for example, the airspeed of the aircraft, such as Mach 5 or greater. In scramjet engine operation, the inlet airflow 76 and bypass airflow 80 to the second burner section 66 can be, for example, about Mach 3, but in any case greater than Mach 1. Therefore, the combined cycle propulsion system 2 can provide a transitional flight mode from subsonic to supersonic to hypersonic flight.

[0052] While the foregoing description generally pertains to gas turbine engines, it is readily understood that gas turbine engines can be implemented in a variety of environments. For example, the engine can be implemented in aircraft, but it can also be implemented in non-aircraft applications, such as power plants, marine applications, or oil and gas production applications. Therefore, this disclosure is not limited to use in aircraft.

[0053] Further aspects of this disclosure are provided by the subject matter of the following clauses.

[0054] A combined cycle propulsion system for an aircraft includes: a compressor-fed internal combustion engine; and a multi-mode supersonic engine, the multi-mode supersonic engine including: an adjustable inlet section; a combustion section disposed downstream of the adjustable inlet section and including a first combustor portion having at least one rotating detonation combustor and a second combustor portion having a supersonic combustion type combustor; and an adjustable exhaust nozzle section disposed downstream of the combustion section, wherein the at least one rotating detonation combustor serves as an igniter for the supersonic combustion type combustor.

[0055] According to the combined cycle propulsion system described in the foregoing clause, the compressor supplies the internal combustion engine, the first burner section and the second burner section with a first fuel type, a second fuel type different from the first fuel type, or a third fuel type different from both the first fuel type and the second fuel type.

[0056] According to any of the preceding clauses, the combined cycle propulsion system includes a compressor-supply internal combustion engine comprising an outer housing extending circumferentially around a centerline axis of the compressor-supply internal combustion engine, and the multi-mode supersonic engine extending circumferentially around the centerline axis and disposed radially outside the outer housing.

[0057] According to any of the preceding clauses, in a combined cycle propulsion system, the first burner section and the second burner section are arranged in series, with the second burner section arranged downstream of the first burner section.

[0058] According to any of the preceding clauses, in a combined cycle propulsion system, the at least one rotary detonator is a single rotary detonator extending circumferentially around the centerline axis.

[0059] According to any of the preceding clauses, in a combined cycle propulsion system, the first burner portion includes a burner bypass flow passage at the rotating detonation burner, the burner bypass flow passage providing a bypass airflow from upstream of the at least one rotating detonation burner to downstream of the at least one rotating detonation burner to the second burner portion.

[0060] According to any of the preceding clauses, in a combined cycle propulsion system, the first burner portion includes a plurality of rotary detonation burners circumferentially spaced around the centerline axis supplied by the compressor to the internal combustion engine.

[0061] According to any of the preceding clauses, in a combined cycle propulsion system, a plurality of bypass flow channels are provided between respective rotary detonation burners in a plurality of rotary detonation burners, and provide bypass airflow through which to the second burner section.

[0062] According to any of the preceding clauses, in a combined cycle propulsion system, a cooling flow channel is provided between the compressor-supplied internal combustion engine and the multi-mode supersonic engine to provide cooling to the inner walls of the multi-mode supersonic engine.

[0063] According to any of the preceding clauses, in a combined cycle propulsion system, a plurality of air exhaust openings are arranged between the compressor-supply internal combustion engine and the multi-mode supersonic engine to provide airflow and / or combustion gas flow from the compressor-supply internal combustion engine to the multi-mode supersonic engine.

[0064] According to any of the preceding clauses, in a combined cycle propulsion system, the compressor-supply internal combustion engine defines a longitudinal direction along the centerline axis of the compressor-supply internal combustion engine, a circumferential direction around the centerline axis, and a radial direction extending from the centerline axis; the multi-mode supersonic engine is arranged radially outside the compressor-supply internal combustion engine and defines a stadium-shaped profile, the height of which extends in the radial direction and the length of which extends orthogonally to the longitudinal and radial directions; and the multi-mode supersonic engine includes a plurality of rotating detonation burners spaced apart from each other along the length of the stadium-shaped profile.

[0065] According to any of the preceding clauses, in a combined cycle propulsion system, the second burner portion includes a plurality of fuel injectors disposed between respective rotary detonation burners in the plurality of rotary detonation burners, the plurality of fuel injectors being arranged to provide a fuel flow downstream of the outlet end of each rotary detonation burner in the plurality of rotary detonation burners.

[0066] According to any of the preceding clauses, the combined cycle propulsion system includes a plurality of first burner section inlet sections corresponding to a respective rotary detonation burner among the plurality of rotary detonation burners, and a plurality of bypass inlet sections arranged between the respective rotary detonation burners among the plurality of rotary detonation burners.

[0067] According to any of the preceding clauses of the combined cycle propulsion system, wherein the adjustable inlet section includes at least one actuator, the adjustable inlet section being adjustable via the at least one actuator to regulate the inlet flow of air to the multi-mode supersonic engine based on (a) a first supersonic operating mode in which the first burner portion operates and the second burner portion does not operate, (b) a second supersonic operating mode in which the second burner portion operates and the first burner portion does not operate, and (c) a transitional supersonic operating mode between the first and second supersonic operating modes in which both the first and second burner portions operate.

[0068] According to any of the preceding clauses, in the combined cycle propulsion system, the second burner portion includes a scramjet burner, wherein in the first supersonic operating mode, the at least one rotating detonation burner operates and the scramjet burner does not operate; in the second supersonic operating mode, the scramjet burner operates and the at least one rotating detonation burner does not operate; and in the transition supersonic operating mode, the at least one rotating detonation burner operates as the igniter of the scramjet burner.

[0069] According to any of the preceding clauses, the combined cycle propulsion system wherein the adjustable exhaust nozzle section includes at least one exhaust nozzle actuator, the adjustable exhaust nozzle section being actuated in the first supersonic operating mode to define a convergent exhaust outlet, b) being actuated in the second supersonic operating mode to define a divergent exhaust outlet, and c) being actuated in the transition supersonic operating mode to transition between the convergent exhaust outlet and the divergent exhaust outlet.

[0070] A method of operating a combined cycle propulsion system for an aircraft, the combined cycle propulsion system comprising a compressor-supplied internal combustion engine and a multi-mode supersonic engine, the multi-mode supersonic engine having an adjustable inlet section, a combustion section, and an adjustable exhaust nozzle section, the combustion section being disposed downstream of the adjustable inlet section and including a first combustor section having at least one rotating detonation combustor and a second combustor section having a supersonic combustion type combustor, the adjustable exhaust nozzle section being disposed downstream of the combustion section, the method comprising: operating the compressor-supplied internal combustion engine to obtain the first operational flight state of the aircraft; in the first operational flight state of the aircraft, initiating the first transition operation to actuate the adjustable inlet section to obtain a subsonic inlet flow of air to the first combustor section. The aircraft then initiates the operation of the first burner section; initiates a compressor-supply-to-internal-combustion-engine-gradual-stop operation to gradually stop the compressor-supply-to-internal-combustion-engine operation; operates the first burner section and actuates the adjustable inlet section and the adjustable exhaust nozzle section to obtain a second operational flight state of the aircraft; in the second operational flight state, initiates a second transition operation to actuate the adjustable inlet section to obtain a supersonic inlet airflow to the second burner section, and uses the first burner section as an igniter to initiate the operation of the second burner section; initiates a first burner section-gradual-stop operation to gradually stop the operation of the first burner section; and operates the second burner section, the adjustable inlet section, and the adjustable exhaust nozzle section to obtain a third operational flight state of the aircraft.

[0071] According to the method described in the foregoing clause, the first burner section includes at least one rotary detonation burner, and the combustion product gas generated by the at least one rotary detonation burner is used as the igniter of the second burner section.

[0072] According to the method described in any of the preceding clauses, the second burner portion includes a scramjet jet burner.

[0073] According to the method of any of the preceding clauses, the first operational flight state includes an airspeed between Mach 0.9 and Mach 2, the second operational flight state of the aircraft includes an airspeed of approximately Mach 3, and the third operational flight state of the aircraft includes an airspeed of at least Mach 5.

[0074] Although the foregoing description is directed to preferred embodiments, it should be noted that other variations and modifications will be apparent to those skilled in the art and can be made without departing from the spirit or scope of this disclosure. Furthermore, features described in connection with one embodiment may be used in conjunction with other embodiments, even if not explicitly stated above.

Claims

1. A combined cycle propulsion system for an aircraft, characterized in that, The combined cycle propulsion system includes: The compressor supplies power to the internal combustion engine; and A multi-mode supersonic engine, wherein the gas flow path of the supersonic engine is defined by a series flow relationship, the multi-mode supersonic engine comprising: An adjustable inlet section, wherein the adjustable inlet section has an inlet at its upstream end; A combustion section, the combustion section being arranged downstream of and in fluid communication with the adjustable inlet section and comprising a first burner section having at least one rotary detonation burner and a second burner section having a supersonic combustion burner; and An adjustable exhaust nozzle section is provided, which is located downstream of and in fluid communication with the combustion section. The at least one rotary detonation burner serves as the ignition source for the supersonic combustion burner. The compressor-supply internal combustion engine includes an outer casing extending circumferentially around the centerline axis of the compressor-supply internal combustion engine, and the multi-mode supersonic engine extends circumferentially around the centerline axis and is arranged radially outward from the outer casing. The housing extends from the inlet of the adjustable inlet section to the adjustable exhaust nozzle section, and has an inner housing wall and an outer housing wall, the outer housing wall being the inner wall of the multi-mode supersonic engine, extending from the inlet of the adjustable inlet section to the adjustable exhaust nozzle section, and... A cooling airflow passage is defined between the inner wall of the housing and the outer wall of the housing, from the inlet of the adjustable inlet section to the adjustable exhaust nozzle section, to provide cooling airflow through it, thereby providing cooling to the inner wall of the multi-mode supersonic engine.

2. The combined cycle propulsion system of claim 1, wherein, in, The compressor supplies the internal combustion engine, the first burner section, and the second burner section with a first fuel type, a second fuel type different from the first fuel type, or a third fuel type different from both the first and second fuel types.

3. The combined cycle propulsion system of claim 1, wherein, in, The compressor-supplied internal combustion engine includes an outer housing extending circumferentially around the centerline axis of the compressor-supplied internal combustion engine, and the multi-mode supersonic engine extending circumferentially around the centerline axis and arranged radially outside the outer housing.

4. The combined cycle propulsion system of claim 3, wherein, in, The first burner section and the second burner section are arranged in series, with the second burner section located downstream of the first burner section.

5. The combined cycle propulsion system of claim 4, wherein, in, The at least one rotary detonation burner is a single rotary detonation burner that extends circumferentially around the central axis.

6. The combined cycle propulsion system of claim 5, wherein, in, The first burner section includes a burner bypass flow passage at the rotary detonation burner, the burner bypass flow passage providing a bypass air flow from upstream of the at least one rotary detonation burner to downstream of the at least one rotary detonation burner to the second burner section.

7. The combined cycle propulsion system of claim 3, wherein, in, The first burner section includes a plurality of rotary detonation burners, which are circumferentially spaced around the centerline axis of the internal combustion engine supplied by the compressor.

8. The combined cycle propulsion system of claim 7, wherein, in, Multiple bypass flow channels are provided between corresponding rotary detonation burners in a plurality of rotary detonation burners, and provide bypass airflow through them to the second burner section.

9. The combined cycle propulsion system of claim 1, wherein, in, A cooling flow channel is provided between the compressor-supplied internal combustion engine and the multi-mode supersonic engine to provide cooling to the inner wall of the multi-mode supersonic engine.

10. The combined cycle propulsion system of claim 1, wherein, in, Multiple air exhaust openings are arranged between the compressor-supply internal combustion engine and the multi-mode supersonic engine to provide airflow and / or combustion gas flow from the compressor-supply internal combustion engine to the multi-mode supersonic engine.

11. The combined cycle propulsion system according to claim 1, characterized in that, in, The compressor supplies power to the internal combustion engine in a longitudinal direction along the centerline axis of the compressor-supply-internal combustion engine, a circumferential direction around the centerline axis, and a radial direction extending from the centerline axis. The multi-mode supersonic engine is arranged radially outside the compressor-supplied internal combustion engine and defines a stadium-shaped profile, the height of which extends in the radial direction and the length of which extends orthogonally to both the longitudinal and radial directions. The multi-mode supersonic engine includes multiple rotating detonation burners spaced apart from each other along the length of the stadium's shaped profile.

12. The combined cycle propulsion system according to claim 11, characterized in that, in, The second burner section includes a plurality of fuel injectors arranged between respective rotary detonation burners in the plurality of rotary detonation burners, the plurality of fuel injectors being arranged to provide a fuel flow downstream of the outlet end of each rotary detonation burner in the plurality of rotary detonation burners.

13. The combined cycle propulsion system according to claim 12, characterized in that, in, The first burner section includes a plurality of first burner section inlet sections corresponding to a respective rotary detonation burner among the plurality of rotary detonation burners, and a plurality of bypass inlet sections arranged between the respective rotary detonation burners among the plurality of rotary detonation burners.

14. The combined cycle propulsion system according to claim 1, characterized in that, in, The second burner section includes a scramjet burner, wherein in the first supersonic operating mode, at least one rotating detonation burner operates and the scramjet burner does not operate; in the second supersonic operating mode, the scramjet burner operates and the at least one rotating detonation burner does not operate; and in the transitional supersonic operating mode, the at least one rotating detonation burner operates as the igniter of the scramjet burner.

15. The combined cycle propulsion system according to claim 14, characterized in that, in, The adjustable exhaust nozzle section includes at least one exhaust nozzle actuator, the adjustable exhaust nozzle section being actuated in a first supersonic operating mode to define a convergent exhaust outlet, b) being actuated in a second supersonic operating mode to define a divergent exhaust outlet, and c) being actuated in the transition supersonic operating mode to transition between the convergent exhaust outlet and the divergent exhaust outlet.

16. A method for operating a combined cycle propulsion system for an aircraft, characterized in that, The combined cycle propulsion system includes a compressor-fed internal combustion engine and a multi-mode supersonic engine. The multi-mode supersonic engine has an adjustable inlet section, a combustion section, and an adjustable exhaust nozzle section. The combustion section is located downstream of the adjustable inlet section and includes a first burner section having at least one rotating detonation burner and a second burner section having a supersonic combustion burner. The adjustable exhaust nozzle section is located downstream of the combustion section. The method includes: The compressor is operated to supply power to the internal combustion engine to obtain the first operational flight state of the aircraft; In the first operational flight state of the aircraft, a first transition operation is initiated to actuate the adjustable inlet section to obtain a subsonic inlet flow of air to the first burner section, and the operation of the first burner section is initiated. The operation of gradually stopping the compressor's supply to the internal combustion engine is initiated and then gradually stopped. Operate the first burner section and actuate the adjustable inlet section and the adjustable exhaust nozzle section to obtain the second operational flight state of the aircraft; In the second operational flight state, a second transition operation is initiated to actuate the adjustable inlet section to obtain a supersonic inlet flow of air to the second burner section, and the operation of the second burner section is initiated using the first burner section as an igniter; The operation of the first burner section is gradually stopped to gradually cease operation of the first burner section; and The second burner section, the adjustable inlet section, and the adjustable exhaust nozzle section are operated to obtain the third operational flight state of the aircraft.

17. The method according to claim 16, characterized in that, in, The first burner section includes at least one rotary detonation burner, and the combustion product gas generated by the at least one rotary detonation burner is used as the igniter of the second burner section.

18. The method according to claim 17, characterized in that, in, The second burner section includes a scramjet jet burner.

19. The method according to claim 16, characterized in that, in, The first operational flight state includes an airspeed between Mach 0.9 and Mach 2, the second operational flight state includes an airspeed between Mach 3 and Mach 4, and the third operational flight state includes an airspeed of at least Mach 5.

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