Integrated flow channel for coupling flow channels of wide-area rocket-based ramjet engine and oblique detonation engine

Abstract

The invention provides a wide-area rocket-based ramjet engine and oblique detonation engine flow channel coupled integrated flow channel, and relates to the technical field of aerospace transportation power systems. The flow channel comprises 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 spraying channel which are communicated in sequence. A mixing support plate is arranged in the combustion chamber isolation channel; the mixing support plate is provided with an injection hole used for injecting fuel into the combustion chamber. A rocket supporting plate is arranged in the first-stage concave cavity; when the Mach number of the transition channel is 0-8, the axial section of the transition channel is rectangular, and when the Mach number is larger than 8, the axial section of the transition channel expands to be trapezoidal. Different expansion ratios are achieved by adjusting the angle of the rear transition section of the first-stage cavity, and conversion from a stamping mode to a knocking mode is achieved 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 for coupling a wide-range rocket-based ramjet engine with an oblique detonation engine flow channel. Background Art

[0002] Space transport vehicles have the characteristics of high flight speed, and their economic value has been highly valued by aerospace researchers around the world. Reusable space transportation systems are an important way for humans to realize space flight flights and are the key to the integrated development of aerospace. This requires space vehicles to have the ability to start at zero speed on the ground and to work efficiently and stably in a wide speed range. This undoubtedly places high demands on the engines of space vehicles.

[0003] The main power schemes for wide-area space transport aircraft are ramjet and scramjet. Since ramjet and scramjet engines need to decelerate and pressurize high-speed incoming air to meet working conditions, they need to reach a certain flight speed to work properly. How to achieve zero-speed starting of ramjet and scramjet engines has become a major problem.

[0004] The combined power scheme is a relatively mature solution at present. Among them, the rocket-based ramjet combined cycle engine organically combines a rocket engine with a high thrust-to-weight ratio and low specific impulse with an air-breathing ramjet engine with a low thrust-to-weight ratio and high specific impulse, giving full play to the advantages and characteristics of the two propulsion methods. Through the conversion between the pure rocket launch mode, the ramjet mode, and the scramjet mode, the problem of zero-speed start of the scramjet engine is successfully solved, and the working speed range of the scramjet engine is broadened and good performance is shown. However, when the working speed range is further increased, usually when the Mach number is greater than 8, due to the long combustion chamber of the scramjet engine, the friction resistance between the inner wall of the combustion chamber and the incoming air will increase sharply, and the performance of the combustion chamber will be significantly reduced, which requires the introduction of a new power scheme.

[0005] The oblique detonation engine has the potential to expand the operating speed range of ramjet engines to Ma8-Ma16+. ODE injects fuel in advance in the forebody or inlet duct, premixes it with air to a certain degree before entering the combustion chamber, and induces oblique detonation waves that can stay in the high-speed airflow through devices such as oblique splitting, completes the mixed gas combustion and heat release in a short distance, and the high-temperature and high-pressure combustion products are accelerated through the nozzle expansion to generate thrust. It is an engine based on supersonic detonation combustion with a thermal cycle efficiency far higher than conventional isobaric combustion. Summary of the invention

[0006] In order to solve the deficiencies in the background technology, the present invention mainly solves the problem in the prior art that when the Mach number is greater than 8, due to the long scramjet combustion chamber, the friction resistance between the combustion chamber inner wall and the incoming air will increase sharply, and the combustion chamber performance will be significantly reduced. 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 couples the flow channels of a rocket-based combined cycle engine (RBCC) and an oblique detonation engine (ODE), that is, a wide-range rocket-based ramjet engine and an oblique detonation engine flow channel integrated combined engine, while absorbing the advantages of rockets, ramjet engines, and oblique detonation engines, and realizing the engine's ultra-wide speed range (Ma0-Ma8+) operation through the ejection mode, subsonic ramjet mode, scramjet mode, and oblique detonation mode to meet the needs of reusability and round-trip between space and air.

[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, including an 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; A mixing support plate is provided in the combustion chamber isolation channel; and an injection hole for injecting fuel into the combustion chamber is provided in the mixing support plate; A rocket support plate is arranged in the first-stage cavity; 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.

[0008] 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.

[0009] Preferably, the combustion chamber isolation channel is a channel that isolates the combustion chamber from the air inlet duct, and comprises a hollow shell with an air inlet end and an air outlet end, wherein 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.

[0010] Preferably, the secondary cavity is a combustion heat release zone when the Mach number is 0-8, or a secondary combustion zone where the fuel and air are further mixed when the Mach number is greater than 8.

[0011] Preferably, 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 starting the engine at zero speed on the ground; A bulge is provided at the oblique split of the front section of the rocket support plate, and the bulge is used to promote the occurrence of oblique blast waves.

[0012] Preferably, the tail injection channel provides space for the incompletely reacted fuel gas to continue to react, while allowing the high-temperature, high-pressure fuel gas to fully expand under certain geometric constraints.

[0013] Preferably, 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.

[0014] 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 operating 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.

[0015] Compared with the prior art, the present invention has the following beneficial effects: 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 a rear transition section of a first-stage cavity, thereby realizing the transition from ramjet mode to detonation mode under high Mach inflow.

[0016] The present invention solves the problem of insufficient power for high-altitude space transport aircraft to start at zero speed on the ground, and uses a liquid rocket on the ground to provide strong power, so that it can smoothly enter the low Mach ramjet mode. At the same time, the support plate is highly coupled with the liquid rocket, the internal structure of the engine becomes simple, the overall reliability is stronger, the production cost is low, and the economic benefit is high.

[0017] The modal overregulation 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.

[0018] The present invention combines a rocket-based ramjet engine with an oblique detonation engine, which significantly broadens the operating speed range and has the ability to adapt to the ultra-high speed range of Ma0-Ma8+. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] Figure 1 It is a cross-sectional view of the integrated flow passage of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow passage; Figure 2 is the numerical simulation result of the ejection mode; Figure 3 is the numerical simulation result of sub-combustion ramjet mode; Figure 4 This is the numerical simulation result of scramjet mode; Figure 5 is the numerical simulation result of oblique detonation mode; Figure 6 Flame factor (Fi) cloud chart.

[0020] 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

[0021] 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.

[0022] 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, achieve different expansion ratios by adjusting the angle of the rear transition section of the first-stage cavity, and realize the transition from ramjet mode to detonation mode under high Mach flow.

[0023] 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 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; A mixing support plate is provided in the combustion chamber isolation channel; and an injection hole for injecting fuel into the combustion chamber is provided in the mixing support plate; A rocket support plate is arranged in the first-stage cavity; 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.

[0024] The first-stage cavity acts as a flame stabilizer and simultaneously produces geometric coupling with the rocket support plate to provide geometric conditions for the generation of oblique detonation waves.

[0025] 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.

[0026] The secondary cavity is a combustion heat release zone when the Mach number is 0-8, or a secondary combustion zone where the fuel and air are further mixed when the Mach number is greater than 8.

[0027] 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 starting the engine at zero speed on the ground; a bulge is provided at the oblique wedge of the front section of the rocket support plate, and the bulge is used to promote the occurrence of oblique blast waves.

[0028] The tail jet channel provides space for the incompletely reacted fuel gas to continue to react, while enabling the high-temperature, high-pressure fuel gas to fully expand under certain geometric constraints.

[0029] 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.

[0030] 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 transformed into a configuration operating 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.

[0031] 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.

[0032] In order to further illustrate the integrated flow passage of a wide-range rocket-based ramjet engine coupled with an oblique detonation engine flow passage provided by the present invention, it is described in conjunction with the accompanying drawings.

[0033] 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.

[0034] See also Figure 1 The invention discloses a combustion chamber configuration integrating flow passages of a wide-range rocket-based ramjet engine and an oblique detonation engine, 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; Combustion chamber inlet, the cross section where the incoming air enters the combustion chamber.

[0035] The combustion chamber isolation section isolates the combustion chamber from the air inlet duct. It is a hollow shell including 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.

[0036] The first-stage cavity acts as a flame stabilizer and simultaneously produces geometric coupling with the rocket support plate to provide geometric conditions for the generation of oblique detonation waves.

[0037] Transition section: Since the oblique detonation mode will cause violent 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.

[0038] The secondary cavity is the main combustion heat release area under low Mach conditions and the supplementary combustion area where fuel and air are further mixed under high Mach conditions.

[0039] The tail nozzle provides space for the incompletely reacted fuel gas to continue reacting while allowing the high-temperature, high-pressure combustion gas to fully expand under certain geometric constraints.

[0040] The combustion chamber outlet is the cross section where the high-temperature combustion gas exits the combustion chamber.

[0041] The mixing support plate is located in the isolation section, generates a pre-combustion shock wave train and injects fuel to mix the fuel with the mainstream.

[0042] The rocket support plate is designed with a pointed front section and a rounded tip to reduce flow resistance; a rocket nozzle is left at the rear end for zero-speed starting of the engine on the ground; the whole plate runs through the entire combustion chamber.

[0043] 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.

[0044] The concave cavity is a flame stabilizer that has been proven to be highly efficient and reliable by a large number of studies. Its working principle is to improve the combustion efficiency and flame stabilization ability by increasing the recirculation area in the combustion chamber. The support plate not only changes the flow field in the combustion chamber but also has a stabilizing effect, and has similar functions to the concave cavity. The combination of the two can effectively improve the combustion efficiency. The combination of the two has many advantages, such as promoting fuel mixing and ignition, enhancing fuel convection, making combustion more complete, and facilitating shortening the length of the combustion chamber. Therefore, the present invention uses a combination of three flame stabilizers: a mixing support plate, a rocket support plate, and a concave cavity.

[0045] 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 make the rocket gas fully expand during the injection process, so as to achieve efficient use of energy. A bulge 10 is provided at the oblique split of the front section of the rocket support plate, and the bulge is used to promote the occurrence of oblique explosion waves.

[0046] Considering that the flow velocity in the combustion chamber is high under the subcombustion / scramjet mode, a two-stage concave cavity is designed to form a recirculation zone in the combustion chamber by using the concave cavity to entrain the mainstream to extend the residence time of the fuel / air mixture in the combustion chamber. In addition, the expansion ratio required for the combustion chamber in different working modes is different. The expansion ratio required for the isochoric combustion in the oblique detonation mode is greater than that in the subcombustion / scramjet mode. Therefore, the transition section after the first-stage concave cavity can be structurally transformed in the equal straight section and the trapezoidal expansion section to meet the requirements of combustion performance under different working modes.

[0047] 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 stabilizing effect of the cavity and the mixing efficiency of fuel / air are greatly reduced. Therefore, the present invention designs the mixing support plate in the isolation section, and completes the mixing of fuel and mainstream by spraying 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.

[0048] In order to achieve efficient and stable operation under ultra-wide flow conditions, the present invention uses multi-mode switching to adapt to the flame stabilization and performance requirements of different flow conditions. The specific implementation plan is as follows: 1. Rocket ejection mode under Ma0-3 inflow conditions. The engine provides power to meet the ground zero-speed takeoff requirements. At this time, the support rocket starts the rocket ejection and uses the rocket's powerful acceleration performance to climb quickly.

[0049] 2. Ma3-5: Subsonic ramjet mode. At this time, the rocket is completely shut down and works in pure ramjet mode. The combustion chamber pressure and the length of the isolation section match, and the isolation section shock wave train is sufficient to decelerate the incoming flow to subsonic speed, which can achieve subsonic ramjet.

[0050] 3.Ma5-8: Scramjet mode. The scramjet mode continues to accelerate to Ma5, and 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, so scramjet can be achieved.

[0051] 4. Above Ma8: oblique detonation mode. The vehicle accelerates to Ma8, and the ramjet stops working. At this time, the transition section expands, and the engine flow path is transformed into a configuration suitable for oblique detonation mode. Detonation is completed by oblique splitting of the leading edge of the support plate rocket, thereby achieving oblique detonation combustion.

[0052] 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 conditions. 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.

[0053] 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.

[0054] 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 fuel gas. The rocket parameters are shown in Table 2. Figure 2It can be seen that the supersonic high-temperature rich-burning gas and the surrounding subsonic incoming air form a reaction mixing layer under the action of shear force, gradually transitioning from a non-premixed flame to a premixed flame, and the pressure in the combustion chamber rises. The rocket jet continues to develop downstream, and the rocket high-speed jet is subjected to shear force, the speed gradually decreases, the incoming flow speed gradually increases, the velocity gradient between the jet and the air becomes smaller, the reaction mixing layer gradually thickens, the jet and the air are mixed more completely, and the rich-burning jet continues to release heat. At this time, the fuel is injected from the wall of the second-stage cavity, and the kerosene droplets are broken, atomized, and evaporated in the flow channel to form gaseous kerosene, which is fully mixed with air and enters the mainstream of the flow channel. The mixed gas directly enters the reaction mixing layer of the rocket jet, where the high temperature and high pressure environment provides good reaction conditions for the kerosene fuel, and the kerosene can fully release heat. The OH cloud map can show that as the reaction mixing layer gradually thickens, the area where the chemical reaction occurs gradually increases, and the heat release in the flow channel is more intense. It can be seen from the temperature cloud map that the range of the high temperature area gradually expands as it flows downstream, indicating that the thickening of the reaction mixing layer and the addition of kerosene fuel have jointly enhanced the chemical reaction and released more heat. Under this condition, the flame in the combustion chamber can be stable and the engine works normally, meeting the design requirements.

[0055] Ma4 subsonic ramjet mode, the inlet operating conditions of the isolation section are shown in Table 3. In this working mode, the rocket (12) will be completely closed and work in a pure ramjet mode. At this time, the combustion chamber pressure and the length of the isolation section (2) match, and the shock wave train of the isolation section (2) is sufficient to decelerate the incoming flow to subsonic speed, so that subsonic ramjet can be achieved. 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 condition are shown in Figure 3 As shown, it can be seen from the temperature cloud map 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 chemical reaction time scale, 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, enters the recirculation area of ​​the second-stage concave cavity (5), generates a large range of low-speed vortices and releases 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 meets 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 required to establish stable combustion. The combustion heat release area is behind the rocket support plate and inside the concave cavity, and there is sufficient space for sufficient chemical reaction, so the combustion efficiency is higher than that of Ma2. Overall, it meets the design requirements.

[0056] Ma6 scramjet mode, the isolation section inlet operating conditions are shown in Table 4. Under the condition of Ma6 flight flow, the numerical simulation results are shown in 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 speed in front of the rocket support plate (9) remains at supersonic speed. The density cloud map can clearly show the increase in flow density after the positive shock wave. The mixing support plate (8) is injected from the side, and it disturbs the high-speed incoming flow and forms a positive shock wave, so that the high-speed incoming flow has a great speed loss and 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 incoming flow of Ma6 is relatively high, 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, it can be seen that 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.

[0057] Ma8 oblique detonation mode, the isolation section inlet operating conditions are shown in Table 5. The simulation data of this operating condition is as follows Figure 5 As 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 group 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 group cloud map, it is found that the OH group concentration here is extremely high. Analysis of the density map shows that there are two strong combustion-supporting shock waves formed by the coupling of diffusion shock waves and reflection shock waves. The incompletely reacted fuel further completes 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 fuel gas, thereby 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, which is defined as the inner product of the fuel mass fraction gradient and the oxidant mass fraction gradient. When the flame index is positive, it is premixed combustion, and when the flame index is negative, it is diffusion combustion. Figure 6The flame factor cloud diagram is shown. Through analysis, it can be found that after the fuel is ejected from the mixing support plate (8), it undergoes a pre-combustion shock wave train, which causes a certain diffusion combustion. Combined with the temperature cloud diagram, it can be found that this combustion is not intense, 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 combustion completion degree of the fuel and air here is quite high, which conforms to the characteristics of the oblique detonation combustion mode and isochoric combustion, and can complete the combustion in a very small shock wave space. 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.

[0058] 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 the oblique detonation together, and realizes the integrated flow path design. Its variable internal flow path structure satisfies the efficient combustion organization under various modes, thereby supporting wide-range Ma0-8+ flight.

[0059] Table 1 Reference operating conditions at the entrance of Ma2 isolation section

[0060] Table 2 Rocket exit parameters of Ma2 rocket support plate

[0061] Table 3 Reference operating conditions at the entrance of Ma4 isolation section

[0062] Table 4 Reference operating conditions at the entrance of Ma6 isolation section

[0063] Table 5 Reference operating conditions for the entrance of Ma8 isolation section

Claims

1. An integrated flow channel for coupling a wide-range rocket-based ramjet engine with 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 which are connected in sequence; A mixing support plate is provided in the combustion chamber isolation channel; and an injection hole for injecting fuel into the combustion chamber is provided in the mixing support plate; A rocket support plate is arranged in the first-stage cavity; 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.

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 produces 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 passage of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow passage 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 secondary combustion zone where the fuel and air are further mixed when the Mach number is greater than 8.

5. The integrated flow passage of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow passage according to claim 1 is characterized in that: 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 ground zero-speed starting of the engine; A bulge is provided at the oblique split 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 passage of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow passage according to claim 1 is characterized in that: The tail jet channel provides space for the incompletely reacted fuel gas to continue to react, while enabling the high-temperature, high-pressure fuel gas to fully expand under certain geometric constraints.

7. The integrated flow passage of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow passage 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.

8. The integrated flow passage of the wide-range rocket-based ramjet engine coupled with the oblique detonation engine flow passage according to claim 1 is characterized in that: 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 transformed into a configuration operating 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.

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