Integrated flow path of wide-range rocket-based ramjet engine coupled with oblique detonation engine flow path

Abstract

The present invention provides an integrated flow channel that couples a wide-range rocket-based ramjet engine with an oblique detonation engine flow channel, and the present invention relates to the technical field of aerospace transportation power systems. The flow channel includes an air inlet duct, a combustion chamber isolation channel, a first-stage concave cavity, a transition channel, a second-stage concave cavity, and a tail nozzle channel that are sequentially connected; a mixing support plate is provided in the combustion chamber isolation channel; an injection hole for injecting fuel into the combustion chamber is provided in the mixing support plate; a rocket support plate is provided in the first-stage concave cavity; the transition channel has an axial cross-section of a rectangle when the Mach number is 0 to 8, and expands to a trapezoidal cross-section when the Mach number is greater than 8. The present invention achieves different expansion ratios by adjusting the angle of the rear transition section of the first-stage concave cavity, thereby realizing the transition from ramjet to detonation mode under high Mach incoming flow.

Description

Technical Field

[0001] The present invention relates to the technical field of aerospace transport power systems, and in particular to an integrated flow channel coupling a wide-range rocket-based ramjet engine with an oblique detonation engine flow channel. Background Art

[0002] Space transport vehicles, characterized by high speeds, possess significant economic value, earning them high attention from aerospace researchers worldwide. Reusable space transport systems are a crucial path to achieving the operationalization of space flight and a key to the integrated development of aviation and aerospace. This requires space vehicles to be able to start from zero speed on the ground and operate efficiently and stably across a wide speed range. This undoubtedly places high demands on their engines.

[0003] Currently, the primary propulsion options for wide-area space transport vehicles are ramjets and scramjets. Because ramjets and scramjets need to decelerate and pressurize high-speed incoming air to meet operating conditions, they require a certain flight speed to function properly. Therefore, achieving zero-speed starting for these engines presents a major challenge.

[0004] The combined propulsion system is a relatively mature solution. The rocket-based ramjet combined cycle engine organically combines a high thrust-to-weight ratio, low specific impulse rocket engine with a low thrust-to-weight ratio, high specific impulse air-breathing ramjet engine, fully leveraging the advantages and characteristics of both propulsion modes. By switching between pure rocket-ejection mode, ramjet mode, and scramjet mode, it successfully solves the scramjet's zero-speed start-up problem, broadens the scramjet's operating speed range, and demonstrates good performance. However, as the operating speed range increases further, typically above Mach 8, the frictional resistance between the scramjet's combustion chamber and the incoming air increases dramatically due to the scramjet's longer combustion chamber, significantly reducing combustion chamber performance. This necessitates the introduction of a new propulsion system.

[0005] Oblique detonation engines have the potential to expand the ramjet's operating speed range to Ma8-Ma16+. ODEs pre-inject fuel into the forebody or inlet, pre-mixing it with air to a certain degree before entering the combustion chamber. Using devices such as oblique splitting, they induce an oblique detonation wave that can settle in the high-speed airflow, completing the mixture combustion and releasing heat over a short distance. The high-temperature, high-pressure combustion products expand through the nozzle, accelerating to generate thrust. This is an engine based on supersonic detonation combustion, with a thermal cycle efficiency far superior to conventional isobaric combustion. Summary of the Invention

[0006] To address the shortcomings of the background technology, the present invention primarily addresses the prior art problem that when the Mach number is greater than 8, due to the length of the scramjet combustion chamber, the frictional resistance between the combustion chamber wall and the incoming air increases dramatically, significantly reducing combustion chamber performance. The present invention provides an integrated flow path that couples the flow paths of a wide-range rocket-based ramjet engine with those of an oblique detonation engine. The present invention couples the flow paths of a rocket-based combined cycle engine (RBCC) with those of an oblique detonation engine (ODE), creating an integrated engine with wide-range rocket-based ramjet and oblique detonation engine flow paths. This engine integrates the flow paths of the RBCC and oblique detonation engines, while simultaneously absorbing the advantages of rockets, ramjet engines, and oblique detonation engines. Through ejection mode, subsonic ramjet mode, scramjet mode, and oblique detonation mode, the engine achieves operation over an ultra-wide speed range (Ma0-Ma8+), thus meeting the requirements of reusability and space-to-air round-trip travel.

[0007] The first object of the present invention is to provide an integrated flow channel for coupling a wide-range rocket-based ramjet engine with an oblique detonation engine flow channel, comprising an air inlet channel, a combustion chamber isolation channel, a first-stage concave cavity, a transition channel, a second-stage concave cavity, and a tail nozzle channel connected in sequence;

[0008] A mixing support plate is provided in the combustion chamber isolation channel; the mixing support plate is provided with injection holes for injecting fuel into the combustion chamber;

[0009] A rocket support plate is provided in the first-stage cavity;

[0010] When the Mach number is between 0 and 8, the axial cross-section of the transition channel is rectangular, and when the Mach number is greater than 8, the axial cross-section expands into a trapezoid.

[0011] Preferably, the first-stage cavity acts as a flame stabilizer and simultaneously generates geometric coupling with the rocket support plate to provide geometric conditions for the generation of oblique detonation waves.

[0012] Preferably, the combustion chamber isolation channel is a channel that isolates the combustion chamber from the air inlet duct, which includes a hollow shell with an air inlet end and an air outlet end. The air inlet end is connected to the air outlet end of the throat section of the equipment to prevent the combustion chamber back pressure from being transmitted forward and causing the air inlet duct to fail to start.

[0013] Preferably, the secondary cavity is a combustion heat release zone when the Mach number is 0-8, or a secondary combustion zone when the Mach number is greater than 8, where the fuel and air are further mixed.

[0014] Preferably, the front section of the rocket support plate is wedge-shaped, with a rounded tip to reduce flow resistance; a rocket nozzle is left at the rear end for ground zero-speed starting of the engine;

[0015] A bulge is provided at the oblique split portion of the front section of the rocket support plate, and the bulge is used to promote the occurrence of oblique blast waves.

[0016] Preferably, the tail jet channel provides space for the incompletely reacted fuel gas to continue reacting, while allowing the high-temperature, high-pressure fuel gas to fully expand under certain geometric constraints.

[0017] Preferably, the injection holes are arranged on the left and right sides of the middle section of the mixing support plate, for injecting fuel into the combustion chamber.

[0018] Preferably, when the Mach number is greater than 8, the axial cross-section of the transition channel expands from a rectangle to a trapezoid, and the engine flow channel as a whole is converted into a configuration working in an oblique detonation mode, and detonation is completed by oblique splitting of the leading edge of the rocket support plate, thereby performing oblique detonation combustion.

[0019] Compared with the prior art, the present invention has the following beneficial effects:

[0020] The present invention provides an integrated flow channel that couples a wide-range rocket-based ramjet engine with an oblique detonation engine flow channel. The present invention achieves different expansion ratios by adjusting the angle of the rear transition section of the first-stage cavity, thereby realizing the transition from ramjet mode to detonation mode under high Mach incoming flow.

[0021] This invention addresses the issue of insufficient ground-based zero-speed startup power for high-altitude space transport vehicles. It utilizes a liquid rocket to provide powerful ground-based propulsion, enabling smooth entry into the low-Mach ramjet mode. Furthermore, the support plate is highly coupled to the liquid rocket, simplifying the engine's internal structure, enhancing overall reliability, reducing production costs, and achieving high economic benefits.

[0022] The modal over-regulation scheme of the present invention is simple and efficient. The present invention can complete the transition from the ramjet mode to the oblique detonation mode by simply adjusting the expansion ratio of the transition section, thereby greatly improving the reliability of the combustion chamber.

[0023] The present invention combines a rocket-based ramjet engine with an oblique detonation engine, significantly broadening the operating speed range and having the ability to adapt to the ultra-high speed range of Ma0-Ma8+. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Figure 1 The cross-sectional view of the integrated flow channel of the wide-area rocket-based ramjet engine coupled with the oblique detonation engine flow channel;

[0025] Figure 2 is the numerical simulation result of the ejection mode;

[0026] Figure 3 is the numerical simulation result of sub-combustion ramjet mode;

[0027] Figure 4 is the numerical simulation result of scramjet mode;

[0028] Figure 5 is the numerical simulation result of oblique detonation mode;

[0029] Figure 6 Flame factor (Fi) cloud chart.

[0030] Among them, 1. Combustion chamber inlet, 2. Combustion chamber isolation section, 3. First-stage cavity, 4. Transition section, 5. Second-stage cavity, 6. Tail nozzle, 7. Combustion chamber outlet, 8. Mixing support plate, 9. Rocket support plate, 10. Bulge, 11. Fuel injection hole, 12. Rocket outlet. DETAILED DESCRIPTION

[0031] Several specific embodiments of the present invention are described in detail below, but it should be understood that the protection scope of the present invention is not limited by the specific embodiments.

[0032] The purpose of the present invention is to provide an integrated flow channel that couples a wide-range rocket-based ramjet engine with an oblique detonation engine flow channel, achieves different expansion ratios by adjusting the angle of the rear transition section of the first-stage cavity, and realizes the transition from ramjet mode to detonation mode under high Mach flow.

[0033] In order to achieve the above-mentioned object, the present invention provides an integrated flow channel for coupling a wide-range rocket-based ramjet engine with an oblique detonation engine flow channel, comprising an air inlet channel, a combustion chamber isolation channel, a first-stage concave cavity, a transition channel, a second-stage concave cavity, and a tail nozzle channel connected in sequence;

[0034] A mixing support plate is provided in the combustion chamber isolation channel; the mixing support plate is provided with injection holes for injecting fuel into the combustion chamber;

[0035] A rocket support plate is provided in the first-stage cavity;

[0036] When the Mach number is between 0 and 8, the axial cross-section of the transition channel is rectangular, and when the Mach number is greater than 8, the axial cross-section expands into a trapezoid.

[0037] The first-stage cavity acts as a flame stabilizer and simultaneously generates geometric coupling with the rocket support plate to provide geometric conditions for the generation of oblique detonation waves.

[0038] The combustion chamber isolation channel is a channel that isolates the combustion chamber from the air inlet duct. It includes a hollow shell with an air inlet end and an air outlet end. The air inlet end is connected to the air outlet end of the throat section of the equipment to prevent the combustion chamber back pressure from being transmitted forward and causing the air inlet duct to fail to start.

[0039] The secondary cavity is a combustion heat release zone when the Mach number is 0-8, or a zone where fuel and air are further mixed and re-combusted when the Mach number is greater than 8.

[0040] The front section of the rocket support plate is in a wedge-shaped shape, and the tip is rounded to reduce flow resistance; a rocket nozzle is reserved at the rear end for the engine to start at zero speed on the ground; a bulge is provided at the oblique wedge of the front section of the rocket support plate, which promotes the occurrence of oblique blast waves.

[0041] The tail jet channel provides space for the incompletely reacted fuel gas to continue reacting, while allowing the high-temperature, high-pressure fuel gas to fully expand under certain geometric constraints.

[0042] The injection holes are arranged on the left and right sides of the middle section of the mixing support plate and are used to inject fuel into the combustion chamber.

[0043] When the Mach number is greater than 8, the axial cross-section of the transition channel expands from a rectangle to a trapezoid, and the engine flow channel as a whole is converted into a configuration working in an oblique detonation mode, and detonation is completed by oblique splitting of the leading edge of the rocket support plate, thereby performing oblique detonation combustion.

[0044] It should be noted that the expansion method in the transition channel of the present invention is: the rear edge of the first-level cavity is used as a fulcrum, and the transition section expands outward by 15 degrees.

[0045] In order to further illustrate the integrated flow channel of the wide-range rocket-based ramjet engine and the oblique detonation engine flow channel provided by the present invention, it is described in conjunction with the accompanying drawings.

[0046] Among them, the air inlet involved in the following embodiments is the combustion chamber inlet, the combustion chamber isolation channel is the combustion chamber isolation section, the transition channel is the transition section, and the tail nozzle channel is the tail nozzle.

[0047] See also Figure 1 The invention relates to a combustion chamber structure integrating a wide-range rocket-based ramjet engine and an oblique detonation engine flow path, which mainly comprises a combustion chamber inlet 1, a combustion chamber isolation section 2, a first-stage concave cavity 3, a transition section 4, a second-stage concave cavity 5, a tail nozzle 6, and a combustion chamber outlet 7, which are connected in sequence.

[0048] Combustion chamber inlet, the cross section where the incoming air enters the combustion chamber.

[0049] The combustion chamber isolation section isolates the combustion chamber from the air inlet duct. It is a hollow shell containing an air inlet end and an air outlet end. Its air inlet end is connected to the air outlet end of the throat section of the equipment to prevent the back pressure of the combustion chamber from being transmitted forward and causing the air inlet duct to fail to start.

[0050] The first-stage cavity acts as a flame stabilizer and simultaneously produces geometric coupling with the rocket support plate, providing geometric conditions for the generation of oblique detonation waves.

[0051] In the transition section, since the oblique detonation mode causes intense combustion in a smaller space, the pressure needs to be released quickly after combustion, so a larger expansion ratio is required. The angle of the transition section is adjusted to adapt to the power requirements of this mode.

[0052] The secondary cavity is the main combustion heat release area under low Mach conditions and the area where fuel and air are further mixed and re-combusted under high Mach conditions.

[0053] The tail nozzle provides space for the incompletely reacted fuel gas to continue reacting while allowing the high-temperature, high-pressure gas to fully expand under certain geometric constraints.

[0054] Combustion chamber outlet, the cross section where high-temperature gas exits the combustion chamber.

[0055] The mixing support plate is located in the isolation section, which generates a pre-combustion shock wave train and injects fuel to mix the fuel with the mainstream.

[0056] The rocket support plate is designed with a wedge-shaped front section and a rounded tip to reduce flow resistance; a rocket nozzle is left at the rear end for the engine to start at 0 speed on the ground; the whole plate runs through the entire combustion chamber.

[0057] The fuel support plate injection holes are arranged on the left and right sides of the middle section of the mixing support plate and are used to inject fuel into the combustion chamber.

[0058] Numerous studies have proven the effectiveness and reliability of the concave cavity as a flame stabilizer. Its operating principle is to improve combustion efficiency and flame stability by increasing the recirculation zone within the combustion chamber. The support plate not only alters the flow field within the combustion chamber but also provides stability, similar to the concave cavity. The combination of the two effectively improves combustion efficiency. This combination promotes fuel mixing and ignition, enhances fuel convection, and ensures more complete combustion, thus shortening the combustion chamber length. Therefore, the present invention utilizes a combination of three flame stabilizers: a mixing support plate, a rocket support plate, and a concave cavity.

[0059] The rocket support plate can effectively provide zero-speed starting power for the aircraft through rocket injection. Designing it in the front position of the combustion chamber can fully expand the rocket gas during the injection process, achieving efficient energy utilization. A bulge 10 is provided at the oblique split of the front section of the rocket support plate to promote the occurrence of oblique blast waves.

[0060] Considering the high flow velocities within the combustion chamber under sub-combustion / scramjet operating conditions, a two-stage concave cavity is designed. This cavity entrains the main flow, creating a recirculation zone within the combustion chamber and extending the residence time of the fuel / air mixture. Furthermore, the required expansion ratio for the combustion chamber varies depending on the operating mode. The isochoric combustion in the oblique detonation mode requires a greater expansion ratio than that in the sub-combustion / scramjet mode. Therefore, the transition section after the first-stage concave cavity is designed to switch between a straight section and a trapezoidal expansion section to accommodate the combustion performance requirements under different operating modes.

[0061] Considering that the conditions for the formation of oblique detonation are relatively harsh, it is necessary to use oblique splitting to generate oblique shock waves under high-speed incoming flow. Therefore, this work couples the rocket support plate with the oblique detonation splitting, and designs a bulge at the support plate splitting to promote the occurrence of oblique detonation waves. In addition, how to inject fuel under high-speed incoming flow is a big problem. Since the incoming flow speed is extremely fast above Ma8, the flame stabilization effect of the cavity and the fuel / air mixing efficiency are greatly reduced. Therefore, the present invention designs the mixing support plate in the isolation section, and completes the mixing of the fuel and the mainstream by injecting kerosene fuel on both sides of the mixing support plate, and completes the detonation and combustion at the oblique shock wave generated by the support plate rocket, thereby solving the problems of difficulty in generating oblique shock waves and difficulty in mixing fuel / air.

[0062] In order to achieve efficient and stable operation under ultra-wide inflow conditions, the present invention adopts multi-mode switching to adapt to the flame stabilization and performance requirements of different inflow conditions. The specific implementation plan is as follows:

[0063] 1. Rocket ejection mode under Ma0-3 inflow conditions. The engine provides power to meet the ground zero-speed takeoff requirement. At this time, the strut rocket activates rocket ejection, utilizing the rocket's powerful acceleration performance to quickly climb.

[0064] 2. Ma3-5: Subsonic ramjet mode. In this mode, the rocket is completely shut down and operates in pure ramjet mode. The combustion chamber pressure and the length of the isolation section are well matched, and the shock wave train in the isolation section is sufficient to decelerate the incoming airflow to subsonic speeds, enabling subsonic ramjet operation.

[0065] 3. Ma5-8: Scramjet mode. The scramjet mode continues to accelerate to Ma5, and then naturally switches to the scramjet mode. In this mode, the total temperature of the downflow is high, the chemical reaction scale is small, and the reaction rate is fast, thus achieving scramjet.

[0066] 4. Mach 8 and above: Oblique detonation mode. The vehicle accelerates to Mach 8, the ramjet engine shuts down, the transition section expands, and the entire engine flow path transforms into a configuration suitable for oblique detonation mode operation. Detonation is achieved by splitting the leading edge of the strut, thus achieving oblique detonation combustion.

[0067] In order to ensure the stability of the engine combustion in an ultra-wide range of Mach numbers, Ma2, Ma4, Ma6, and Ma8 were selected as reference operating conditions, and the transition section expansion scheme is as follows: Figure 1 As shown, the reference operating condition tables are shown in Table 1, Table 3, Table 4, and Table 5.

[0068] The specific working mode of the flow channel provided by the present invention in different working modes is further explained through the calculation results under four reference working conditions: Ma2, Ma4, Ma6 and Ma8.

[0069] Ma2 rocket ejection mode, at this time, the isolation section entrance working conditions are shown in Table 1, the rocket outlet (12) behind the rocket support plate (9) opens and ejects supersonic high-temperature rich-burning gas, the rocket parameters are shown in Table 2, Figure 2 It can be seen that the supersonic, high-temperature, fuel-rich gas and the surrounding subsonic incoming air, under shear forces, form a reactive mixing layer, gradually transitioning from a non-premixed flame to a premixed one, and the combustion chamber pressure increases. As the rocket jet continues downstream, the high-speed rocket jet is subjected to shear forces, its velocity gradually decreasing, while the incoming air velocity gradually increases. The velocity gradient between the jet and the air decreases, the reactive mixing layer gradually thickens, and the jet and air mix more completely, allowing the fuel-rich jet to continue releasing heat. At this point, fuel is injected from the wall of the second-stage cavity. Kerosene droplets break up, atomize, and evaporate within the flow channel, forming gaseous kerosene. This gas mixes thoroughly with the air before entering the main flow channel. The mixed gas enters the rocket jet's reactive mixing layer, where the high temperature and high pressure provide optimal reaction conditions for the kerosene fuel, allowing it to fully release heat. The OH cloud shows that as the reactive mixing layer thickens, the area where the chemical reaction occurs gradually expands, and the heat release within the flow channel becomes more intense. The temperature cloud map shows that the high-temperature zone gradually expands as the flow moves downstream, indicating that the thickening of the reaction mixing layer and the addition of kerosene fuel enhance the chemical reaction and increase heat release. Under these operating conditions, the flame in the combustion chamber remains stable and the engine operates normally, meeting design requirements.

[0070] Ma4 subsonic ramjet mode, the operating conditions of the isolation section entrance are shown in Table 3. In this working mode, the rocket (12) will be completely closed and rely on the pure ramjet mode to work. At this time, the combustion chamber pressure and the length of the isolation section (2) match, and the shock wave string of the isolation section (2) is sufficient to decelerate the incoming flow to subsonic speed, which can achieve subsonic ramjet. Since the total pressure of the Ma4 incoming flow is small and there is a risk of overflow, the starting position of the engine heat release is designed to be inside the second-stage cavity (5). The numerical simulation results under this working condition are shown in Figure 3. Figure 3As shown in the temperature cloud diagram, it can be seen that the high-temperature area is mainly concentrated behind the rocket support plate (9) and in the second-stage concave cavity (5), indicating that the deceleration effect of the rocket support plate (9) generates a low-speed recirculation area behind the rocket support plate (9), and the flow time scale of the chemical reaction is reduced, which can match the time scale of the chemical reaction, so that the fuel can be sucked into it and a chemical reaction occurs. The low-speed airflow continues to flow downstream of the engine and enters the recirculation area of ​​the second-stage concave cavity (5), generating a large-scale low-speed vortex and releasing heat violently. The OH group can characterize the heat release position to a certain extent. It can be seen that the OH content in the rocket support plate (9) and the second-stage concave cavity (5) is the highest, indicating that heat release mainly occurs in this range, and the actual combustion process is in line with the designed heat release range. After calculation, the total temperature of Ma4 is slightly higher than that of Ma2, and only the initial ignition energy is needed to establish stable combustion. The combustion heat release area is behind the rocket support plate and inside the concave cavity, with sufficient space for sufficient chemical reaction, so the combustion efficiency is higher than that of Ma2. Overall, it meets the design requirements.

[0071] Ma6 scramjet mode, the isolation section inlet working conditions are shown in Table 4. Under the condition of Ma6 flight flow, the numerical simulation results are shown in Table 4. Figure 4 As shown, the intensity of the shock wave string of the isolation section (2) is not enough to decelerate the incoming flow to subsonic speed, so the incoming flow velocity in front of the rocket support plate (9) remains at supersonic speed. The density cloud map clearly shows the increase in the flow density after the positive shock wave. The mixing support plate (8) is injected from the side, and the high-speed incoming flow disturbs each other and forms a positive shock wave, which causes the high-speed incoming flow to have a great speed loss and thus can be fully mixed with the kerosene droplets. The fuel begins to release heat at the mixing support plate (8). It can be seen from the temperature cloud map that since the total temperature of the Ma6 incoming flow is relatively high at about 1700K, the static temperature is increased to the ignition temperature of kerosene after the positive shock wave is decelerated, and kerosene can spontaneously combust without forced ignition. The OH-based active group is an intermediate component of kerosene combustion and can generally reflect the location where the reaction occurs. According to the OH cloud map, the chemical reaction and combustion heat release of kerosene are mainly in the shear layer between the combustion flow and the air flow, indicating that the shear layer promotes the mixing of fuel and air, and the chemical reaction is intense, and the combustion heat release is concentrated in this area. Overall, the working mode of the Ma6 scramjet has high combustion efficiency due to its high total temperature, small chemical reaction scale, and fast chemical reaction rate, which can meet the design requirements.

[0072] Ma8 oblique detonation mode, the isolation section inlet working condition is shown in Table 5. The simulation data of this working condition is as follows Figure 5As shown. Analyze the flow field details of the rocket support plate (9) and the first-stage concave cavity (3). By observing the cloud map, it can be seen that the OH base concentration suddenly increases at 1, accompanied by two strong oblique shock waves, indicating that the oblique shock wave generated by the oblique splitting of the leading edge of the rocket support plate (9) successfully detonates. There is a high-temperature reaction area at 2. By observing the OH base cloud map, it is found that the OH base concentration here is extremely high. By analyzing the density map, it can be seen that there are two strong combustion-supporting shock waves formed by the coupling of diffusion shock waves and reflection shock waves. The unreacted fuel further completes the combustion here, and due to the contraction of the flow channel, a large number of OH groups are enriched here. This shows that under the Ma8 inflow condition, the oblique detonation mode of the combustion chamber can not only successfully detonate, but also make the heat release position forward to provide sufficient space for the combustion expansion of the gas, reducing the fuel waste caused by the fuel being blown out of the combustion chamber before it is completely reacted. In order to further analyze the internal mechanism of oblique detonation mode combustion, the flame index analysis method is combined. It is defined as the inner product of the fuel mass fraction gradient and the oxidizer mass fraction gradient. When the flame index is positive, it is premixed combustion, and when the flame index is negative, it is diffusion combustion. The following is obtained: Figure 6 The flame factor cloud diagram shown. Through analysis, it can be found that after the fuel is ejected from the mixing support plate (8), it experiences a pre-combustion shock wave string, which causes a certain diffusion combustion. Combined with the temperature cloud diagram, it can be found that this combustion is not strong, and it is more of a preheating of the premixed gas formed by air / fuel. When the premixed gas of air and fuel flows through the oblique split of the rocket support plate (9), the combustion reaction caused by the oblique shock wave occurs rapidly. From observing the flame factor cloud diagram, it can be found that a violent premixed combustion occurs at the oblique shock wave generated by the rocket support plate (9) and there is no diffusion combustion phenomenon afterwards. This shows that the degree of combustion completion of the fuel and air here is quite high, which is in line with the characteristics of the oblique detonation combustion mode and isochoric combustion. It can complete combustion in a very small shock wave space. At this point, it can be judged that the integrated coupling scheme of the rocket-based ramjet engine and the oblique detonation engine flow channel can indeed achieve a reliable oblique detonation combustion mode.

[0073] In summary, the integrated flow path of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow path innovatively combines the rocket and oblique detonation together, and realizes the integrated flow path design. Its variable internal flow path structure meets the efficient combustion organization under various modes, thereby supporting wide-range Ma0-8+ flight.

[0074] Table 1 Reference operating conditions at the entrance of Ma2 isolation section

[0075]

[0076] Table 2 Rocket exit parameters of Ma2 rocket support plate

[0077]

[0078] Table 3 Reference operating conditions at the entrance of Ma4 isolation section

[0079]

[0080] Table 4 Reference operating conditions at the entrance of Ma6 isolation section

[0081]

[0082] Table 5 Reference operating conditions at the entrance of Ma8 isolation section

[0083]

Claims

1. An integrated flow channel coupling a wide-range rocket-based ramjet engine and an oblique detonation engine flow channel, characterized in that: It includes an air inlet duct, a combustion chamber isolation channel, a first-stage concave cavity, a transition channel, a second-stage concave cavity, and a tail nozzle channel that are connected in sequence; A mixing support plate is provided in the combustion chamber isolation channel; the mixing support plate is provided with injection holes for injecting fuel into the combustion chamber; A rocket support plate is provided in the first-stage cavity; The transition channel has an axial cross-section that is rectangular when the Mach number is 0-8, and expands into a trapezoidal cross-section when the Mach number is greater than 8; The expansion mode of the transition channel is as follows: the rear edge of the first-stage cavity is used as the fulcrum, and the transition section expands outwards by 15 degrees; Through multi-mode switching to adapt to different inflow conditions of flame stabilization and performance requirements, the specific scheme is as follows: Under Ma0-3 inflow conditions, the rocket ejection mode is used. At this time, the support rocket starts the rocket ejection and uses the powerful acceleration performance of the rocket to quickly climb. Ma3-5: Sub-ramjet mode, at which point the rocket is completely shut down and operates in pure ramjet mode; Ma5-8: scramjet mode; Above Ma8: oblique detonation mode, the ramjet stops working. At this time, the transition section expands, and the entire engine flow path is transformed into a configuration suitable for oblique detonation mode operation. Detonation is completed by oblique splitting of the leading edge of the support plate rocket, thereby achieving oblique detonation combustion.

2. The integrated flow channel of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow channel according to claim 1 is characterized in that: The first-stage cavity acts as a flame stabilizer and simultaneously generates geometric coupling with the rocket support plate to provide geometric conditions for the generation of oblique detonation waves.

3. The integrated flow channel of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow channel according to claim 1 is characterized in that: The combustion chamber isolation channel is a channel that isolates the combustion chamber from the air inlet duct. It includes a hollow shell with an air inlet end and an air outlet end. The air inlet end is connected to the air outlet end of the throat section of the equipment to prevent the combustion chamber back pressure from being transmitted forward and causing the air inlet duct to fail to start.

4. The integrated flow channel of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow channel according to claim 1 is characterized in that: The secondary cavity is a combustion heat release zone when the Mach number is 0-8, or a zone where fuel and air are further mixed and re-combusted when the Mach number is greater than 8.

5. The integrated flow channel of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow channel according to claim 1 is characterized in that: The front section of the rocket support plate is in a wedge-shaped shape, with a rounded corner at the tip to reduce flow resistance; a rocket nozzle is left at the rear end for ground zero-speed starting of the engine; A bulge is provided at the oblique split portion of the front section of the rocket support plate, and the bulge is used to promote the occurrence of oblique blast waves.

6. The integrated flow channel of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow channel according to claim 1 is characterized in that: The tail jet channel provides space for the incompletely reacted fuel gas to continue reacting, while allowing the high-temperature, high-pressure fuel gas to fully expand under certain geometric constraints.

7. The integrated flow channel of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow channel according to claim 1 is characterized in that: The injection holes are arranged on the left and right sides of the middle section of the mixing support plate and are used to inject fuel into the combustion chamber.

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