Rocket supporting plate and concave cavity combined oblique detonation engine

Abstract

The invention relates to the technical field of rocket engines, in particular to a rocket support plate and concave cavity combined oblique detonation engine which comprises a mixing section, a combustion chamber and an exhaust nozzle which are sequentially connected, the mixing section is of a hollow structure, a fuel-rocket support plate assembly is arranged in the mixing section, and a detonation slope-concave cavity combined support plate assembly is arranged in the combustion chamber; a middle flow channel for constant-pressure combustion is formed in the middle of the detonation slope-concave cavity combined supporting plate assembly, and an upper flow channel and a lower flow channel for constant-volume combustion are formed by the upper surface and the lower surface of the detonation slope-concave cavity combined supporting plate assembly and the inner wall of the combustion chamber. According to the device, an oblique detonation combustion mode, a rocket mode, a subsonic combustion ramjet mode and a scramjet mode are integrated into the same engine, a split flow channel structure is formed, full-speed domain coverage from low Mach to ultrahigh Mach (more than 8 Mach) is achieved, and combustion organizations in different flight stages are optimized through a mode switching mechanism.

Description

Technical Field

[0001] The invention relates to the technical field of rocket engines, and in particular to a rocket support plate cavity combined oblique detonation engine. Background Art

[0002] With the development of science and technology, aircraft are also developing towards higher altitudes and higher speeds. The power performance of aerospace vehicles determines its key capabilities such as flight envelope, range, and maneuverability, and also supports the realization of the vision of free aerospace flight in the future. From turbine engines, subsonic ramjets to scramjets, and then to various forms of new combined power, air-breathing aerospace power has been developing towards higher speed, more compactness, wider range, and more economical directions.

[0003] Turbine engines have the problem of "thrust trap". At around Mach 3, most of the energy will be lost in the form of shock waves on the impeller. The principle of the subsonic engine is to compress the supersonic incoming flow through the inlet and the isolation section, and then reduce the airflow speed to subsonic speed through some flame stabilization devices such as the structure of the support plate cavity to organize combustion. However, at higher Mach numbers, the total temperature of the incoming flow is very high. When the supersonic incoming flow is reduced to subsonic speed, the static temperature of the airflow is very high. When the static temperature of the airflow is close to or even higher than the constant pressure combustion temperature of the fuel, the thrust gain generated by the fuel will be greatly reduced or even no thrust gain. For this reason, the scramjet engine has emerged. Its principle is to reduce the speed of the supersonic incoming flow, but still keep it in the supersonic range, and then organize combustion. At this time, the static temperature of the airflow will be much lower than the subsonic airflow, and the energy released by the fuel will also be more. However, when the Mach number of the scramjet engine is higher, due to the high airflow speed in the combustion chamber, the fuel needs a long distance to complete mixing and constant pressure combustion, which will cause the engine's combustion chamber to become very long, which is not conducive to engine performance and flight engine integrated design.

[0004] In order to solve the defects of a single power solution, various combined power types have emerged. The combined cycle power system (referred to as the combined power system) organically combines two or more power types to achieve complementary advantages, maximize the advantages of different powers under different working environments, and achieve power performance improvement within the full working envelope, thereby improving the wide envelope working capability and full envelope working performance of the aircraft. It is an ideal power for future earth-to-earth transportation systems, near-space hypersonic aircraft, etc. However, under high Mach number conditions, it will face the same problems as subsonic ramjet engines or scramjet engines. The flight upper limit of existing combined powers is restricted by the flight upper limit of subsonic ramjet engines or scramjet engines. In addition, traditional air-breathing aerospace powers all adopt isobaric combustion based on slow combustion, and the cycle thermal efficiency is close to its theoretical upper limit, and it is very difficult to further improve it.

[0005] like Figure 11As shown in the figure, the existing rocket-based combined cycle (RBCC) engine is a combined power system formed by organically combining a rocket engine with a high thrust-to-weight ratio and a ramjet engine with a high specific impulse. It can generally operate in the entire airspace and speed range, has strong acceleration and maneuvering orbit-changing capabilities, and is more suitable for use in the power of hypersonic flight platforms, the first and second stages of two-stage-to-orbit space shuttles, or the power of single-stage-to-orbit space shuttles. It consists of Figure 11 It can be seen that RBCC has four working modes. In the subsonic or low Mach number stage, the rocket ejector mode is adopted. The high-temperature rich-burning gas generated by the rocket in the flow channel is mixed with the fuel injected later and burned again for a certain distance, and then expands and does work through the nozzle; when the incoming flow Mach number further increases, the subsonic combustion ramjet mode is adopted. At this time, the rocket is still running, and the high-temperature rich-burning gas generated is mixed with the incoming flow compressed by the inlet and isolator, and then a thermal choke occurs in the thermal throat and its speed is reduced to subsonic speed through a certain flow channel, and then it burns with the subsequent injected fuel and expands and does work through the nozzle; when the incoming flow Mach number further increases, reducing the air flow to subsonic speed to organize combustion will lose most of the energy. Therefore, the supersonic combustion ramjet mode will be adopted. At this time, the air flow will be decelerated through a series of shock wave trains after passing through the rocket, but still remains in a supersonic state, and mixes with the fuel and organizes combustion in the supersonic state; when the air flow Mach number further increases, at this time the supersonic combustion ramjet mode will face the problems of high static temperature or a long combustion distance. Therefore, the pure rocket mode will be adopted. At this time, the high-temperature gas generated by the rocket further burns in the flow channel and expands and does work through the nozzle.

[0006] In summary, the existing combustion mode of the rocket-based combined power cycle is still constant-pressure combustion. When in the scramjet power mode at the highest Mach number during the air-breathing power stage, due to the problems of the scramjet engine itself, when the oncoming flow Mach number further increases, if combustion is to be completed within a certain distance, it is necessary to increase the compression intensity of the inlet to reduce the air flow velocity. However, increasing the compression intensity of the inlet will not only increase the static temperature of the oncoming flow in the combustion chamber but also increase the total pressure loss of the inlet, which is not conducive to the heat release of the engine combustion and will significantly reduce the engine performance. If the compression intensity of the inlet is not changed, the length of the combustion chamber needs to be increased. At this time, the overall length of the engine increases, and increasing the length of the combustion chamber will increase the thermal choking in the subsonic combustion mode of the RBCC engine, which may cause the engine to overflow and the inlet to unstart in this mode. That is, increasing the length of the combustion chamber to increase the performance of the scramjet mode will weaken the performance of the engine in the subsonic combustion mode, and it may even be counterproductive. Currently, other combined power methods that adopt constant-pressure combustion also have this problem. For combined power engines that adopt constant-volume combustion, such as engines combined with rotating detonation technology, the upper limit of the oncoming flow Mach number for constant-volume combustion is lower than the working upper limit of the scramjet, and it is also impossible to broaden the working upper limit of the combined power engine.

[0007] Therefore, a rocket strut cavity combined oblique detonation engine is needed to solve the above problems. Summary of the Invention

[0008] To solve the problems in the prior art that increasing the length of the combustion chamber will increase the thermal choking in the subsonic combustion mode of the RBCC engine, which may cause the engine to overflow and the inlet to unstart in this mode. That is, increasing the length of the combustion chamber to increase the performance of the scramjet mode will weaken the performance of the engine in the subsonic combustion mode, and for combined power engines that adopt constant-volume combustion, such as engines combined with rotating detonation technology, the upper limit of the oncoming flow Mach number for constant-volume combustion is lower than the working upper limit of the scramjet, and it is also impossible to broaden the working upper limit of the combined power engine, the present invention provides a rocket strut cavity combined oblique detonation engine to solve the existing problems.

[0009] A rocket strut cavity combined oblique detonation engine of the present invention adopts the following technical solutions, including: A mixing section, a combustion chamber, and a tail nozzle connected in sequence. The mixing section is a hollow structure, and a fuel-rocket strut assembly is arranged therein. An initiating ramp-cavity combined strut assembly is arranged in the combustion chamber. The middle of the initiating ramp-cavity combined strut assembly forms an intermediate flow path for constant-pressure combustion, and its upper surface and lower surface form an upper flow path and a lower flow path for constant-volume combustion with the inner wall of the combustion chamber; Among them, the fuel-rocket strut assembly includes: a first strut, which is horizontally arranged inside the mixing section, has a rhomboid-like cross-section along the axial section of the engine, is provided with a first fuel supply channel inside, and second fuel supply channels are arranged on the upper and lower sides of the first fuel supply channel. Both the first fuel supply channel and the second fuel supply channel are communicated with the injection pipeline connected to the side wall of the mixing section. A combustion nozzle cavity is arranged on the end face of the first strut facing away from the oncoming flow direction, and the middle section of the combustion nozzle cavity is communicated with the first fuel supply channel. The combustion nozzle cavity is used for burning the fuel supplied by the first fuel supply channel and spraying the gas; a plurality of injection holes communicating the inner cavity of the mixing section and the second fuel supply channel are arranged on the first strut, and nozzles are arranged on the injection holes.

[0010] Preferably, the combustion nozzle cavity includes: a combustion chamber cavity and a nozzle cavity. The inlet of the combustion chamber cavity is communicated with the middle section of the first fuel supply channel, the outlet end of the combustion chamber cavity is communicated with the inlet of the nozzle cavity, and the outlet of the nozzle cavity extends outside the end face of the first strut.

[0011] Preferably, the detonation ramp-recess combined strut assembly includes: two second struts oppositely arranged inside the combustion chamber. The inner wall surfaces of the second struts facing the combustion chamber have the same shape. The opposite surfaces of the two second struts are parallel, and grooves are provided on the opposite surfaces. The two grooves form a recess relatively. Among them, an inlet section of the intermediate flow channel is formed between the opposite surfaces of the second struts close to the oncoming flow end, and the recess of the second struts facing away from the oncoming flow end forms an outlet section of the intermediate flow channel.

[0012] Preferably, one end of the second fuel supply channel passes through the first strut and is communicated with the fuel injection pipeline of the engine connected to the side wall of the mixing section, and the other end is closed inside the first strut.

[0013] Preferably, the first fuel supply channel penetrates through the two side end faces of the first strut, and the ports of the first fuel supply channel are communicated with the injection pipelines of the fuel and oxidant of the rocket.

[0014] Preferably, the nozzle is a 90° vertically injecting straight-through nozzle or a straight-through nozzle that is inclined 45° backward in the direction away from the oncoming flow.

[0015] Preferably, the diameter of the nozzle is 0.2 mm to 0.3 mm.

[0016] Preferably, the cross-sectional area of the second fuel supply channel is 1.2 to 1.5 times the total area of all the injection holes on the second fuel supply channel.

[0017] Preferably, the blockage ratio of the first strut is 15% to 30%.

[0018] Preferably, the leading edge angles of the first strut are all 10° to 30°.

[0019] The beneficial effects of the present invention are: 1. By arranging a fuel-rocket strut assembly in the mixing section and an ignition ramp-recess combined strut assembly in the combustion chamber, an intermediate flow channel is formed in the middle of the ignition ramp-recess combined strut assembly, and upper and lower flow channels are formed between the upper and lower surfaces of the ignition ramp-recess combined strut assembly and the inner wall of the combustion chamber. Thus, the oblique detonation combustion mode is integrated with the rocket, subsonic ramjet, and supersonic ramjet modes in the same engine, forming a separated flow channel structure to achieve full-speed coverage from low Mach to ultra-high Mach (above Mach 8), and optimizing the combustion organization in different flight stages through a mode switching mechanism. Combining the rocket combustion chamber, fuel injection function with the aerodynamic design of the fuel-rocket strut assembly, that is, by arranging a combustion nozzle cavity communicating with the fuel supply channel on the first strut, the mixing of rocket gas and air inflow is achieved, and the efficient distribution of fuel is realized by using the injection holes and nozzles on both sides of the first strut, reducing the complexity and weight of the engine.

[0020] 2. By combining two second struts of the ignition ramp-recess combined strut assembly to form a recess structure, the coupling of shock waves and combustion waves is induced in the oblique detonation mode to stabilize the standing oblique detonation wave, shorten the combustion distance and improve the combustion efficiency. A 90° straight nozzle or 45° oblique straight nozzle design is adopted to balance the fuel penetration depth and the influence of the low-speed zone. Combining the parameter settings of the injection hole diameter of 0.2 - 0.4 mm and the interval of 10 mm, the mixing effect is improved and the risk of early combustion is reduced. Based on the blockage ratio (15% - 30%), the key dimensions of the height of the first strut and the width of the flow channel in the mixing section are determined to optimize the mechanical properties of the first strut and the aerodynamic characteristics of the combustion chamber, avoid boundary layer separation and the formation of high-temperature low-speed zones, and ensure stable operation under multiple modes. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the description of the embodiments or the prior art. Obviously, the following drawings are only some embodiments of the present invention. For those of ordinary skill in the art, other drawings can be obtained based on these drawings without creative efforts.

[0022] Figure 1 It is a schematic diagram of the overall structure of a rocket strut-recess combined oblique detonation engine of the present invention; Figure 2 For Figure 1 a schematic diagram of the structure of the mixing section in Figure 3 For Figure 1 a schematic diagram of the structure of the combustion chamber in Figure 4 For Figure 1 a schematic diagram of the structure of the tail nozzle in Figure 5 ForFigure 1 Structural schematic diagram of the middle fuel-rocket strut assembly; Figure 6 is Figure 5 Side end face schematic diagram of; Figure 7 is Figure 6 Cross-sectional view taken along A-A in; Figure 8 is Figure 6 Cross-sectional view taken along B-B in; Figure 9 is Figure 5 Top view of; Figure 10 is Figure 9 Cross-sectional view taken along C-C in; Figure 11 Structural schematic diagram of an existing rocket-based combined cycle (RBCC) engine; Figure 12 Temperature cloud map in the Ma10 oblique detonation mode in the embodiments of the present invention; Figure 13 Density cloud map in the Ma10 oblique detonation mode in the embodiments of the present invention; Figure 14 OH cloud map in the Ma10 oblique detonation mode in the embodiments of the present invention; Figure 15 Density gradient cloud map in the Ma10 oblique detonation mode in the embodiments of the present invention.

[0023] In the figure: 1, mixing section; 2, combustion chamber; 3, tail nozzle; 4, first strut; 5, detonation ramp-recess combined strut assembly; 11, installation groove; 41, injection hole; 42, first fuel supply channel; 43, combustion chamber cavity; 44, nozzle cavity; 45, second fuel supply channel. Detailed implementation manners

[0024] Next, the technical solutions in the embodiments of the present invention will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present invention. Obviously, the described embodiments are only a part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those of ordinary skill in the art without creative efforts shall fall within the protection scope of the present invention.

[0025] Due to the characteristics of short-distance heat release, supersonic combustion, and constant-volume combustion of the oblique detonation engine, it can achieve combustion organization at higher Mach numbers, complete combustion in a shorter combustion chamber, reduce the total mass of the engine, and the constant-volume combustion has higher efficiency than the constant-pressure combustion of subsonic combustion and supersonic combustion ramjet engines, which is determined by the thermodynamic cycle. Therefore, integrating the oblique detonation engine technology into the combined power form is expected to increase the upper limit of the flight speed of the combined power engine. Thus, a rocket strut cavity combined oblique detonation engine of the present invention is provided.

[0026] An embodiment of the rocket strut cavity combined oblique detonation engine of the present invention is as follows Figure 1 shown, including: a mixing section 1, a combustion chamber 2, and a nozzle 3 connected in sequence. The nozzle 3 is as Figure 4 shown, as Figure 2 shown, the mixing section 1 is a hollow structure, and a fuel-rocket strut assembly is arranged inside the mixing section 1, as Figure 3 shown, a detonation ramp-cavity combined strut assembly 5 is arranged inside the combustion chamber 2. An intermediate flow passage is formed in the middle of the detonation ramp-cavity combined strut assembly 5, and upper and lower flow passages are respectively formed between its upper and lower surfaces and the inner wall of the combustion chamber 2. Similar to the multi-duct engine of an aeroengine, it has a rocket ejector mode, a subsonic combustion ramjet mode, a supersonic combustion ramjet mode, an oblique detonation mode, and a pure rocket mode.

[0027] It should be noted that detonation combustion only occurs in the upper and lower flow passages, and this part is constant-volume combustion; of course, in the detonation working mode, the intermediate flow passage is always deflagration, that is, constant-pressure combustion.

[0028] Among them, the fuel-rocket strut assembly includes: a first strut 4, the first strut 4 is horizontally arranged inside the mixing section 1, as Figure 1 and Figure 2 shown, the corresponding end sides of the first strut 4 are installed in the installation grooves 11 opened on the shell of the mixing section 1, as Figure 9 and Figure 10 shown, the axial section of the first strut 4 along the engine is rhomboid-like, as Figure 5 , Figure 6 , Figure 7 and Figure 8As shown, a first fuel supply channel 42 is provided in the first support plate 4. Second fuel supply channels 45 are provided on the upper and lower sides of the first fuel supply channel 42. Both the first fuel supply channel 42 and the second fuel supply channels 45 communicate with a jet pipe connected to the side wall of the mixing section 1. The end face of the first support plate 4 facing away from the oncoming flow direction is provided with a combustion nozzle cavity, and the middle section of the combustion nozzle cavity communicates with the first fuel supply channel 42. The combustion nozzle cavity is used to burn the fuel supplied by the first fuel supply channel 42 and eject the combustion gas; A plurality of injection holes 41 communicating the inner cavity of the mixing section 1 and the second fuel supply channels 45 are provided on the first support plate 4, and nozzles are provided on the injection holes 41.

[0029] It should be noted that the first support plate 4 can be used as a rocket support plate or a support plate for fuel mixing. Fuel is injected into the channels on the upper and lower sides of the first support plate 4 through the second fuel supply channels 45 of the first support plate 4 via the injection holes 41 and the nozzles, so that the fuel is distributed in the upper and lower parts of the combustion chamber when entering the combustion chamber. Among them, the fuel and oxidant in the second fuel supply channels 45 are used as the fuel and oxidant sources of the rocket, and corresponding igniters are provided to ensure that the rocket can be stably ignited. When the fuel is aviation kerosene, the structural schematic diagram of the fuel-rocket support plate assembly of the present invention is Figure 5 as shown.

[0030] Exemplarily, as Figure 7 shown, in a specific embodiment, the combustion nozzle cavity includes: a combustion chamber cavity 43 and a nozzle cavity 44. The inlet of the combustion chamber cavity 43 communicates with the middle section of the first fuel supply channel 42. The outlet end of the combustion chamber cavity 43 communicates with the inlet of the nozzle cavity 44. The outlet of the nozzle cavity 44 extends outside the end face of the first support plate 4.

[0031] Exemplarily, as Figure 1 and Figure 3 shown, in a specific embodiment, the initiating ramp-recess combination support plate assembly 5 includes: two second support plates oppositely arranged in the combustion chamber 2. The inner wall surfaces of the second support plates facing the combustion chamber 2 have the same shape. The opposite surfaces of the two second support plates are parallel, and grooves are provided on the opposite surfaces. The two grooves form a recess relative to each other. Among them, the inlet section of the middle flow channel is formed between the opposite surfaces of the second support plates near the oncoming flow end, and the recess of the second support plate facing away from the oncoming flow end forms the outlet section of the middle flow channel.

[0032] Exemplarily, as Figure 1 、 Figure 6 and Figure 8As shown, in a specific embodiment, one end of the second fuel supply channel 45 is connected to the fuel injection pipeline of the engine connected to the side wall of the mixing section 1, and the other end of the second fuel supply channel 45 is closed inside the first support plate 4. Among them, the first fuel supply channel 42 penetrates through the two side end faces of the first support plate 4, and the ports of the first fuel supply channel 42 are connected to the injection pipelines of the fuel and oxidizer of the rocket. That is, the fuel and oxidizer enter the second fuel supply channel 45 and the first fuel supply pipeline 42 respectively from both ends of the combustion chamber 2. The fuel and oxidizer in the first fuel supply pipeline 42 enter the combustion chamber cavity 43, and preliminary combustion is organized in the combustion chamber cavity 43. The high-temperature gas generated by combustion is discharged into the axial air flow channel composed of the mixing section 1 and the fuel-rocket support plate assembly after expanding through the nozzle cavity 44.

[0033] Exemplarily, in a specific embodiment, the nozzle is a 90° vertically injecting direct current nozzle or a direct current nozzle that sprays obliquely 45° away from the oncoming flow direction. The diameter of the nozzle is 0.2 mm to 0.3 mm. Among them, the 90° vertically injecting direct current nozzle is the injection scheme with the highest fuel penetration depth and can obtain the optimal mixing effect. The direct current nozzle that sprays obliquely 45° away from the oncoming flow direction sacrifices a certain penetration ability and mixing effect. The low-speed area generated by its injection is smaller than that of the 90° vertical injection, reducing the possibility of premature combustion. The two injection schemes have their own advantages and disadvantages. The purpose of the second fuel supply channel 45 and the first fuel supply pipeline 42 of the fuel-rocket support plate assembly is to spray fuel (kerosene) as the main fuel in the oblique detonation mode and other non-pure rocket modes. The diameter of the nozzle being 0.2 mm to 0.3 mm can take into account both the atomization and penetration of the fuel. If the processing technology is not good, this standard can be relaxed to 0.4 mm. A nozzle with too large a diameter will cause the ejected liquid droplets to require a long distance to break and evaporate in the mixing section, which is not conducive to combustion. A too small fuel nozzle has a greater processing difficulty itself, is easily affected by impurities in the fuel or distortion due to thermal expansion and contraction of the material, resulting in nozzle blockage. Moreover, the momentum of small liquid droplets is relatively smaller, and the penetration depth is insufficient, which will lead to a reduction in mixing efficiency and is not conducive to organizing combustion. In this embodiment, 0.2 mm is adopted.

[0034] Exemplarily, in a specific embodiment, the cross-sectional area of the second fuel supply channel 45 is 1.2 to 1.5 times the total area of all the injection holes 41 on the second fuel supply channel 45. Since the cross-sectional area of the injection holes 41 is too small, it may cause the second fuel supply channel 45 to not reach the full pipe flow state, resulting in less fuel injection or even no fuel injection at some positions of the injection holes 41, which is disadvantageous for fuel mixing; if the cross-sectional area is too large, it will cause the mechanical properties of the first splitter plate 4 to decline, thereby causing partial damage to the first splitter plate 4 and even causing engine damage. Therefore, in this embodiment, the cross-sectional area of the second fuel supply channel 45 is 1.3 times the total area of the injection holes 41 thereon. Secondly, in this embodiment, the interval between adjacent injection holes 41 is 10 mm. If the interval between the injection holes 41 is too close, more injection holes 41 will be arranged under the same distribution length, resulting in a smaller area for each injection hole 41. Similarly, if the interval between the injection holes 41 is too large, the area of each injection hole 41 will become larger.

[0035] Exemplarily, in a specific embodiment, the blockage ratio of the first splitter plate 4 is 15% to 30%, and the leading edge angles of the first splitter plate 4 are all 10° to 30°. Among them, in this embodiment, the blockage ratio of the first splitter plate 4 is taken as 20%. Based on the blockage ratio of the first splitter plate 4, the width of the first splitter plate 4 in the mixing section 1 is determined to be 100 mm, and the height of the first splitter plate 4 is 20 mm. In this embodiment, the leading edge angle of the first splitter plate 4 is 20°. Although a too small leading edge angle can enhance the performance of the combined oblique detonation engine in the oblique detonation mode, since a fuel supply channel needs to be designed inside the first splitter plate 4, the leading edge angle of the first splitter plate 4 cannot be designed too low, which will result in an overly long first splitter plate 4, being not conducive to mechanical properties and possibly causing a long high-temperature and low-speed region formed by the comprehensive action of the boundary layer of the first splitter plate 4 and the recirculation region at the trailing edge of the first splitter plate 4 in the central flow field, increasing the risk of pre-ignition. To ensure mechanical properties, a chamfer not exceeding 1 mm can be set at the leading edge of the first splitter plate 4. A too small chamfer has little improvement on mechanical properties, while an overly large chamfer will cause flow separation at the trailing edge of the chamfer of the first splitter plate 4 in the oblique detonation combustion mode, thereby thickening the boundary layer and increasing the risk of pre-ignition.

[0036] The present invention will be described below with reference to specific simulation data: Ma10 oblique detonation mode, at this time, the combustion chamber inlet parameters are shown in Table 1, and the simulation data for this working condition are as Figures 12 to 15 shown.

[0037] Table 1

[0038] Among them, as Figure 15As shown in the figure, when the high-speed mixed air flow enters the combustion chamber 3, it is geometrically deflected (with a 30° deflection angle) by the second splitter plate at the inlet of the combustion chamber 4 and compressed by the wall surface, forming an initial oblique shock wave (OSW). This shock wave compresses the air flow to a high-temperature and high-pressure state, triggering the spontaneous combustion of the fuel. In the short combustion chamber, the initial oblique shock wave (OSW) reflects off the upper wall surface, forming a reflected shock wave and a boundary layer separation zone. The reattachment shock wave downstream of the separation zone further compresses the air flow, causing a sudden increase in local pressure and forcing the combustion wave to propagate at supersonic speed, forming an overdriven detonation wave (ODW). The detonation wave is nearly perpendicular to the flow channel, and the shock wave is fully coupled with the flame front, with the heat release of combustion concentrated in a narrow area behind the shock wave.

[0039] The initial oblique shock wave (OSW) forms near the inlet of the combustion chamber 3. As shown in Figure 12 the temperature contour map, a narrow high-temperature zone appears at the inlet of the combustion chamber 3, corresponding to the adiabatic heating caused by shock wave compression; Figure 12 simultaneously shows a steep density gradient in Figure 14 identifying the shock wave front. At this time, as shown in Figure 14 the OH concentration is relatively low, indicating that the high-temperature environment has triggered local spontaneous combustion of the fuel but has not formed stable combustion. As the air flow penetrates deeper into the combustion chamber, the interaction between shock wave reflection and the boundary layer generates a local high-pressure zone, and an overdriven detonation wave (ODW) is formed accordingly. Figure 12 A high-temperature zone expands downstream of the shock wave in Figure 13 As shown in the density contour map, the density first increases and then decreases behind the shock wave, reflecting the competitive effects of compression and expansion; as shown in Figure 14 the OH contour map, a sharp jump in the OH distribution is shown, forming a continuous high-concentration zone, indicating that combustion is completed instantaneously behind the shock wave, the shock wave is fully coupled with the flame front, and there is no decoupling phenomenon.

[0040] From Figure 12 、 Figure 13 and Figure 14 the spatial alignment (the coincidence of the temperature peak, density abrupt change, and high OH concentration zone) further verifies the strong coupling characteristics of the overdriven detonation wave, reveals the dynamic processes of shock wave reflection, boundary layer separation, and detonation wave stabilization, and can be used as a basis for the design and optimization of an oblique detonation engine at extremely high speeds.

[0041] Working principle In the rocket ejector mode, rocket gas and oxidizer are respectively injected into the left and right of the combustion chamber cavity 43 and the nozzle cavity 44 from the second fuel supply channel 45 of the fuel-rocket splitter plate assembly, forming a high-temperature fuel-rich jet to eject air; the first fuel supply channels 42 on the upper and lower sides of the second fuel supply channel 45 inject fuel through the injection holes 41 to the upper and lower sides of the mixing section 1, mixing with the rocket gas for pre-combustion, providing an initial mixing condition for subsequent ram combustion. It provides ground zero-speed thrust for the whole.

[0042] In the subsonic combustion ramjet mode, the fuel injection and oxidizer injection into the combustion chamber cavity 43 and the nozzle cavity 44 are stopped, the rocket combustion is shut down, and fuel is injected into the combustion chamber 3 only through the first fuel supply channels 42 on the upper and lower sides of the second fuel supply channel 45. The fuel is mixed with the low-speed recirculation zone behind the fuel-rocket strut assembly and the subsonic airflow. The middle flow channel of the detonation ramp-recess combined strut assembly 5 maintains constant-pressure combustion, the fuel burns stably in the recess recirculation zone, the strut rockets are completely shut down, and the main combustion is maintained only by the fuel strut injection.

[0043] In the supersonic combustion ramjet mode, the first fuel supply channels 42 on the upper and lower sides of the second fuel supply channel 45 inject fuel through transverse jets, and self-ignition is achieved by using the supersonic airflow shear. The middle flow channel of the detonation ramp-recess combined strut assembly 5 continues constant-pressure combustion, and the upper and lower flow channels formed by the second strut of the detonation ramp-recess combined strut assembly 5 and the combustion chamber 3 enhance the mixing efficiency through shock wave matching. The strut rockets remain closed, and the rhomboid cross-section structure of the struts aids supersonic mixing through aerodynamic disturbances.

[0044] In the oblique detonation mode, the first fuel supply channels 42 on the upper and lower sides of the second fuel supply channel 45 inject fuel in advance to form a premixed gas. When flowing through the ramp and recess of the detonation ramp-recess combined strut assembly 5, an oblique shock wave is induced to trigger detonation, and the locally high-temperature gas synergistically stabilizes the detonation wave. The upper and lower flow channels formed by the second strut of the detonation ramp-recess combined strut assembly 5 and the combustion chamber 3 are converted to constant-volume combustion, the middle flow channel still maintains constant-pressure combustion, and the transition channel expands to a trapezoid to match the detonation expansion requirement, realizing short-distance and high-efficiency heat release.

[0045] The above are only the preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent replacements, improvements, etc. made within the spirit and principles of the present invention shall be included in the protection scope of the present invention.

Claims

1. A rocket strut cavity combined oblique detonation engine, comprising: A mixing section, a combustion chamber, and a tail nozzle connected in sequence, characterized in that the mixing section is a hollow structure, inside which a fuel-rocket strut assembly is arranged, and an ignition ramp-recess combined strut assembly is arranged in the combustion chamber. A middle flow path for constant-pressure combustion is formed in the middle of the ignition ramp-recess combined strut assembly, and an upper flow path and a lower flow path for constant-volume combustion are formed between its upper surface, lower surface and the inner wall of the combustion chamber: Among them, the fuel-rocket strut assembly includes: a first strut, which is horizontally arranged inside the mixing section. Its axial cross-section along the engine is rhomboid-like. A first fuel supply channel is arranged inside it. Second fuel supply channels are arranged on both the upper and lower sides of the first fuel supply channel. Both the first fuel supply channel and the second fuel supply channel are connected to the injection pipeline connected to the side wall of the mixing section. A combustion nozzle cavity is arranged on the end face of the first strut facing away from the oncoming flow direction, and the middle section of the combustion nozzle cavity is communicated with the first fuel supply channel. The combustion nozzle cavity is used for burning the fuel supplied by the first fuel supply channel and ejecting the gas; A plurality of injection holes communicating the inner cavity of the mixing section and the second fuel supply channel are arranged on the first strut, and nozzles are arranged on the injection holes.

2. The combined inclined detonation engine with rocket strut and cavity according to claim 1, characterized in that, The combustion nozzle cavity includes: a combustion chamber cavity and a nozzle cavity. The inlet of the combustion chamber cavity is communicated with the middle section of the first fuel supply channel. The outlet end of the combustion chamber cavity is communicated with the inlet of the nozzle cavity. The outlet of the nozzle cavity extends outside the end face of the first strut.

3. The combined inclined detonation rocket strut cavity engine according to claim 1, characterized in that, The ignition ramp-recess combined strut assembly includes: two second struts oppositely arranged inside the combustion chamber. The inner wall surfaces of the second struts facing the combustion chamber have the same shape. The opposite surfaces of the two second struts are parallel, and grooves are opened on the opposite surfaces. The two grooves form a recess relative to each other. Among them, an inlet section of the middle flow path is formed between the opposite surfaces of the second struts near the oncoming flow end, and an outlet section of the middle flow path is formed by the recess of the second struts facing away from the oncoming flow end.

4. A combined rocket strut cavity oblique detonation engine according to claim 1, characterized in that, One end of the second fuel supply channel passes through the first strut and is connected to the fuel injection pipeline of the engine connected to the side wall of the mixing section, and the other end is closed inside the first strut.

5. A combined inclined detonation rocket strut cavity engine according to claim 1, wherein The first fuel supply channel penetrates through the two side end faces of the first strut, and the ports of the first fuel supply channel are connected to the injection pipelines of the fuel and oxidant of the rocket.

6. The combined inclined detonation rocket strut cavity engine according to claim 1, characterized in that, The nozzle is a 90° vertically injecting straight-through nozzle or a rear-injecting straight-through nozzle inclined 45° away from the oncoming flow direction.

7. A combined rocket strut and cavity oblique detonation engine according to claim 1, characterized in that The diameter of the nozzle is 0.2 mm to 0.3 mm.

8. A combined inclined detonation rocket strut cavity engine according to claim 1, characterized in that, The cross-sectional area of the second fuel supply channel is 1.2 to 1.5 times the total area of all the injection holes on the second fuel supply channel.

9. The combined inclined detonation rocket strut cavity engine according to claim 1, characterized in that The blockage ratio of the first strut is 15% to 30%.

10. The combined inclined detonation rocket strut cavity engine according to claim 1, characterized in that The leading edge angle of the first strut is 10° to 30°.

Images