Multi-cavity scramjet combustor and control method thereof

Abstract

This invention provides a multi-cavity scramjet engine combustor and its control method. The combustor has a simple structure, can adapt to different equivalence ratio ranges and flight altitudes, exhibits low drag and high combustion performance, low heat load on the combustor wall, is easy to maintain, and facilitates integrated aircraft-engine design. The combustor includes: an isolation section, a first cavity group, a second cavity group, and a third cavity group, an injection module, and a combustor expansion section. Each cavity group has two vertically opposite cavities. The isolation section is a first square tube connected to the first cavity group. The first cavity group is connected to the second cavity group via a second square tube. The second cavity group is connected to the third cavity group via a third third tube. The combustor expansion section is connected to the rear of the third cavity group. The injection unit includes a front injection device and a rear injection device. The front injection device is positioned with its upper and lower walls facing each other on the second square tube. The rear injection device is positioned with its upper and lower walls facing each other on the third third tube.

Description

Technical Field

[0001] This invention relates to a combustion chamber, specifically a multi-cavity scramjet engine combustion chamber, belonging to the technical field of scramjet engine combustion chambers. Background Technology

[0002] Scramjet engines are suitable for hypersonic vehicles and are a hot research topic. However, their combustion chambers have high velocities and low fuel mixing efficiency, necessitating the use of concave cavities, support plates, and other devices to enhance fuel mixing and stabilize the flame. Concave cavities, when used as flame stabilizers, offer simple structure and low drag, and have become a commonly used flame stabilization device for scramjet engines.

[0003] Improving combustion efficiency, widening the combustion boundary, and reducing total pressure loss within a finite-length combustion chamber are key research objectives for scramjet engines. The injection position significantly impacts the combustion performance of scramjet engines. Existing research indicates that changing the fuel injection position in a scramjet engine can alter combustion performance, thereby changing the thrust of the combustion chamber. Studies on the flame stability of scramjet engines with different injection positions reveal that flame stability varies. Research on scramjet engines using ethylene as fuel shows that while multi-orifice injection results in higher combustion efficiency compared to single-orifice injection, it also leads to lower flame stability. Furthermore, existing research indicates that the orifice diameter significantly affects the injection position in scramjet engines. When the orifice diameter is small, the injection position is closer to the leading edge of the cavity, resulting in a lower total kerosene equivalence ratio required for successful ignition; conversely, when the orifice diameter is large, the injection position has little impact on the ignition characteristics of kerosene fuel.

[0004] However, in publicly published literature, including the aforementioned studies, most scramjet engines have only one or a pair (i.e., two concave cavities in parallel) in the flow direction. These scramjet engines struggle to maintain high combustion efficiency and low total pressure loss over a wide stoichiometric range. Furthermore, existing research has found that, for the same length, configuring more concave cavities in the flow direction can significantly improve combustion efficiency without a substantial increase in total pressure loss. Moreover, multi-cavity scramjet engines can adapt to a wide range of flight conditions. Therefore, configuring more concave cavities in the flow direction is suitable for high Mach number scramjet engines.

[0005] However, configuring a greater number of concave cavities in the flow direction makes the flow field more complex. Currently, multi-cavity scramjet engines under research primarily use hydrogen, ethylene, and kerosene as fuel. Hydrogen and ethylene have low densities, making storage difficult, and large-scale practical applications are not yet possible. Aviation kerosene, on the other hand, has high mass per unit volume, is easy to store, and is widely used in various conventional aircraft. Therefore, scramjet engines using aviation kerosene as fuel have engineering advantages and are expected to provide a power source for hypersonic vehicles. However, organizing combustion in scramjet engines using aviation kerosene is challenging, especially at high Mach numbers, requiring further research. For example, existing technology has studied a dual-mode scramjet engine using kerosene at a combustor inlet Mach number of Ma=2.0, finding very low combustion efficiency and high total pressure loss, which require further optimization. Therefore, research on multi-cavity scramjet engines using kerosene at high Mach numbers is needed to improve combustion efficiency, reduce total pressure loss, and thus improve combustion performance.

[0006] Existing scholars have conducted research on multi-cavity scramjet engines using kerosene as fuel at high Mach numbers from various perspectives. For example, in a six-cavity scramjet engine, deploying cavities in front of the injection can enhance combustion mixing and improve combustion efficiency. Studies of a three-cavity scramjet engine at a combustion chamber inlet Mach number of Ma = 2.92 have found that the injection effect through combustion-induced upstream separation zone injection has an excessive impact on fuel injection, potentially increasing penetration depth and improving combustion efficiency. Furthermore, existing research has found that in multi-cavity scramjet engines, the total pressure loss in the combustion chamber is mainly due to friction loss; shortening the combustion chamber length can significantly reduce the total pressure loss. There is also research on four-cavity scramjet engines with a combustion chamber inlet Ma = 3.0, which has revealed significant abrupt changes in combustion intensity and back pressure with variations in stoichiometry, severely affecting flame stability. Furthermore, research on a four-cavity scramjet engine using aviation kerosene as fuel revealed that the front and rear cavities of the series-connected combustion chambers influence each other, causing abrupt changes in thrust with varying equivalence ratios. Studies of the flame position in the series-connected cavities under different total pressures showed that the flame stabilization point is downstream at low equivalence ratios and upstream at high equivalence ratios. Moreover, as the total pressure increases, the flame stabilization point propagates from downstream to upstream.

[0007] Although preliminary studies have been conducted on multi-cavity scramjet engines using aviation kerosene at high Mach numbers, focusing on aspects such as fuel injection, fuel blending, total pressure loss versus combustion chamber length, equivalence ratio versus combustion intensity, equivalence ratio versus thrust, total pressure change versus flame stability, and combustion chamber configuration, aiming to improve combustion efficiency and reduce combustion chamber length, the combustion efficiency within a finite length remains low in high-Mach number scramjet engines using aviation kerosene, necessitating further improvements. Furthermore, the injection position significantly impacts combustion performance; therefore, rationally deploying the injection position and organizing the combustion of scramjet engines using aviation kerosene is crucial for enhancing their combustion performance.

[0008] Meanwhile, the low-velocity recirculating airflow in the concave cavity has a significant impact on fuel mixing, flame stability, and total pressure loss when it enters the mainstream at different angles. When it enters the mainstream at a smaller angle, the total pressure loss is smaller, but the flame stability, fuel mixing, and mass exchange performance between the concave cavity and the mainstream are poorer. When it enters the mainstream at a larger angle, the flame stability, fuel mixing, and mass exchange performance between the concave cavity and the mainstream are higher, but the total pressure loss also increases accordingly.

[0009] Therefore, it is necessary to rationally arrange the combustion chamber configuration of the multi-cavity scramjet engine, deploy appropriate injection methods, and balance the various performance parameters in the combustion chamber so that the combustion chamber can adapt to a wider range of equivalence ratios, thereby obtaining a high-performance combustion chamber configuration.

[0010] Furthermore, changes in intake pressure significantly impact the combustion performance of scramjet engines, especially during high-altitude cruise when the incoming total pressure is lower. For example, existing research suggests that as the incoming total pressure decreases, the combustion chamber pressure decreases, leading to a slower chemical reaction rate. Additionally, fuel atomization becomes less efficient, droplet size increases, fuel-air mixing deteriorates, and flame stabilizer stability declines. Moreover, existing studies have investigated the combustion characteristics of ethylene-fueled scramjet engines with intake total pressures of 1.0, 1.3, and 1.6 MPa, finding that ignition becomes more difficult at lower intake total pressures; however, at higher intake total pressures, combustion becomes more chaotic, with no dominant oscillation frequency in the combustion chamber. A dual-cavity scramjet engine has also been studied, reporting a sudden thrust shift during the uphill phase from an incoming total pressure of Pt0 = 1.5–1.9 MPa due to the transition between ramjet and scramjet engine modes caused by changes in incoming total pressure. The effects of wall injection at incoming total pressures of 1.2, 1.6, and 2.0 MPa were investigated, revealing that fuel penetration depth and mixing efficiency significantly improved with increasing incoming total pressure. Mode conversion experiments were conducted on a scramjet engine under different incoming total pressures. Results showed that when the incoming total pressure was in the range of 0.6–0.9 MPa, the mode conversion equivalence ratio remained constant regardless of pressure variations. Furthermore, the wall cooling effect of a hydrogen-fueled scramjet engine at incoming pressures of 10 bar and 15 bar was studied. It was found that, except for auto-ignition caused by shock-boundary layer interaction, the static wall pressure distribution was largely independent of changes in static wall temperature and cooling efficiency. The combustion performance of the scramjet engine within the inlet total pressure range of 0.6–0.9 MPa was investigated, revealing stable combustion at both higher and lower inlet total pressures, while significant combustion oscillations occurred at moderate inlet total pressures. Furthermore, during the low-altitude climb phase, the total incoming pressure is relatively high. Under high equivalence ratio conditions, when the isolator is short, the shock wave surface of the thermal throat approaches the isolator inlet. It may even cross the isolator inlet and enter the inlet outlet surface, affecting the inlet's starting performance.

[0011] Therefore, when deploying the combustion chamber configuration, the low-altitude performance of the scramjet engine needs to be considered so that the scramjet engine can maintain high combustion performance under different incoming flow pressures, and thus be able to adapt to different flight altitudes.

[0012] Furthermore, many existing combustor configurations do not take into account combustor cooling or vehicle integration. Existing research indicates that traditional separate vehicle-engine configurations introduce significant drag during hypersonic flight, potentially impacting range and maneuverability. Therefore, current hypersonic vehicle shape and overall design require engines to meet vehicle-engine integration requirements, ensuring the engine's shape also functions as part of the vehicle's lifting surface, thus reducing drag losses caused by the engine structure.

[0013] Therefore, existing multi-cavity scramjet engines that use aviation kerosene as fuel need to solve the following problems:

[0014] 1. In the high-altitude low-pressure cycle stage, existing multi-cavity scramjet engines that use aviation kerosene as fuel have relatively difficult combustion organization, low combustion efficiency, and cannot adapt to combustion with a wide equivalence range.

[0015] 2. From the low-altitude climb phase to the high-altitude cruise phase, the total incoming pressure changes significantly from high to low. During the low-altitude climb phase, the existing scramjet engine has a relatively high total incoming pressure, which can easily cause the thermal throat to cross the inlet of the isolation section and enter the outlet surface of the intake, resulting in the engine failing to start. During the cruise phase, due to the lower total incoming pressure, the fuel mixing performance and flame stability of the cavity are lower, which can easily lead to problems with lower combustion performance.

[0016] 3. The design process of a multi-cavity stable combustion scramjet engine needs to take into account fuel mixing performance and total pressure loss, as well as the integration of the scramjet engine and the aircraft. Summary of the Invention

[0017] In view of this, the present invention provides a multi-cavity scramjet engine combustor that is simple in structure, adaptable to different equivalence ratio ranges and different flight altitudes, has low drag, high combustion performance, low heat load on the combustor wall, is easy to maintain, and facilitates the integrated design of the aircraft and engine.

[0018] The technical solution of the present invention is: a multi-cavity scramjet engine combustion chamber, comprising: an isolation section, a first cavity group, a second cavity group and a third cavity group, an injection module and a combustion chamber expansion section;

[0019] Each cavity group has two cavities arranged vertically opposite each other.

[0020] The isolation section is a first square tube, which is connected to the first cavity group; the first cavity group is connected to the second cavity group through a second square tube; the second cavity group is connected to the third cavity group through a third square tube, and the combustion chamber expansion section is connected to the rear of the third cavity group.

[0021] The injection unit includes a front injection device and a rear injection device; the front injection device is arranged opposite to the upper and lower walls of the second square tube; the rear injection device is arranged opposite to the upper and lower walls of the third square tube.

[0022] As a preferred embodiment of the present invention: the injection unit further includes an auxiliary injection device; the auxiliary injection device is respectively arranged opposite to each other on the two cavities of the second cavity group.

[0023] As a preferred embodiment of the present invention: the isolation section is a first square tube with an expansion angle, the upper wall of which is an upwardly inclined slope with an upward inclination angle not exceeding 1°.

[0024] As a preferred embodiment of the present invention: the upper wall surfaces of the second square tube and the third third tube are upwardly inclined slopes, and the upward inclination angle does not exceed 1°.

[0025] As a preferred embodiment of the present invention, the upper wall of the combustion chamber expansion section is inclined upward, and the angle of upward inclination does not exceed 6°.

[0026] As a preferred embodiment of the present invention, the rear wall surfaces of two of the three concave cavities in the first concave cavity group, the second concave cavity group, and the third concave cavity group are inclined surfaces tilted backward at 45°.

[0027] As a preferred embodiment of the present invention, the lower wall surfaces of the isolation section, the first cavity group, the second cavity group, the third cavity group, and the combustion chamber expansion section are horizontal.

[0028] As a preferred embodiment of the present invention: both the front injection device and the rear injection device are provided with a plurality of injection holes that are equally spaced along the lateral direction;

[0029] The diameter of the injection hole is 0.3mm to 0.6mm;

[0030] The distance between the central axis of the front injection device and the front wall of the second cavity group is 55mm to 75mm.

[0031] The distance between the central axis of the rear injection device and the front wall of the third cavity group is 55mm to 75mm.

[0032] Furthermore, the present invention provides a control method for the combustion chamber of the multi-cavity scramjet engine, characterized in that the engine combustion chamber is disposed inside the engine of an aircraft;

[0033] When an aircraft using the scramjet engine combustion chamber is operating in the low-altitude climb phase, the valve connected to the front injection device is closed, and the valve connected to the rear injection device is opened. According to the aircraft control requirements, the opening degree of the valve connected to the rear injection device is adjusted to increase the fuel volume.

[0034] Once the aircraft reaches the set altitude, the opening of the valve connected to the rear injection device is gradually reduced, while the opening of the valve connected to the front injection device is gradually increased, allowing the aircraft to continue climbing to the cruising altitude. After the aircraft reaches the high-altitude cruising altitude, the valve connected to the rear injection device is completely closed, and the valve connected to the front injection device is opened according to the required fuel volume, allowing the aircraft to fully enter the high-altitude cruising phase. Then, the aircraft continues to operate until it completes its journey.

[0035] As a preferred embodiment of the present invention: when auxiliary injection devices are respectively provided opposite to each other on the two cavities of the second cavity group:

[0036] During the low-altitude climb phase, when large-amplitude pressure oscillations occur, the valve connected to the auxiliary injection device is opened, and kerosene with a set equivalent ratio is injected through the auxiliary injection device to suppress the large-amplitude pressure oscillations.

[0037] Beneficial effects:

[0038] (1) The multi-cavity scramjet engine combustor of the present invention, which uses aviation kerosene as fuel, has three cavity groups. The first cavity group is used for mixing, which disrupts the boundary layer of the airflow and increases the turbulence in the airflow without significantly increasing the airflow resistance. The third cavity group is used for secondary auxiliary combustion, which mixes the exhaust gas after combustion in the second cavity group with the mainstream, so that the exhaust gas after combustion in the second cavity group is burned again in the third cavity group, which enhances combustion without significantly increasing the airflow resistance.

[0039] Therefore, during the high-altitude, low-pressure cruise phase: after the airflow enters the first concave cavity group through the isolation section, the first concave cavity group creates entrained vortices in the high-speed airflow. This high-speed airflow with entrained vortices improves the mixing of the high-speed airflow and fuel near the forward injection device. The third concave cavity group further mixes and burns the exhaust gas produced after combustion in the second concave cavity group with the high-speed airflow, improving heat release and combustion efficiency. During the low-altitude, high-pressure climb phase, the first and second concave cavity groups create more entrained vortices in the high-speed incoming gas, ensuring thorough mixing of the fuel injected by the rear injection device with the high-speed airflow in the third concave cavity group. This improves heat release within the limited combustion chamber length and optimizes the climb phase. This combustion chamber maintains high combustion efficiency within an equivalence ratio range of 0–1.1 during both the high-altitude, low-pressure cruise phase and the low-altitude, high-pressure climb phase, adapting to various complex environments, exhibiting high reliability, and improving combustion efficiency while reducing total pressure loss, thereby increasing range and thrust-to-weight ratio.

[0040] (2) In this invention, an auxiliary injection device is provided, which can reduce the large amplitude oscillation during the low-altitude high-pressure climbing stage and reduce the damage to the structure caused by pressure oscillation during the combustion process.

[0041] (3) In this invention, the upper wall of the combustion chamber expansion section is inclined upwards, and the upward inclination angle does not exceed 6°; the combustion expansion section with a small expansion angle can accelerate the high-temperature incoming gas, and at the same time, the small expansion angle makes the high-temperature gas further away from the wall, reducing the negative mass caused by the cooling heat load. And by controlling the expansion angle within a certain range, the backflow caused by the expansion angle being too large is reduced.

[0042] (4) In this invention, the presence of the 45° inclined wall of the first cavity group makes the gas in the vortex backflow zone of the cavity and the incoming gas form a 45° angle, thereby achieving a balance between the total pressure loss of the backflow gas on the incoming flow and the enhanced mixing of the incoming flow; the 45° angle between the rear edge walls of the second and third cavity groups makes the high-temperature gas in the cavity more quickly ejected into the high-speed incoming gas, making the temperature near the pipe wall behind the cavity lower, and also reducing the total pressure loss caused by the excessive inclination angle between the backflow gas and the high-speed incoming gas in the cavity.

[0043] (5) In this invention, the lower part of the combustion chamber expansion section is flush with the horizontal level of the isolation section and the three concave cavity group, which facilitates the integrated design with the shape of the aircraft, and the structure is simple and easy to maintain; and it can reduce the drag of the aircraft in high-speed airflow.

[0044] (6) In this invention, front and rear injection devices are provided; the front injection device is used under high-altitude cruise conditions, enabling complete combustion within a longer combustion chamber length under low-pressure conditions during high-altitude cruise, while the third concave cavity group serves as secondary auxiliary combustion, improving fuel combustion efficiency during high-altitude cruise; the rear injection device injects fuel at low altitudes, ensuring that the upstream shock surface of the thermal throat generated under low-altitude high-pressure conditions is far from the inlet of the isolation section, improving the intake performance of the upstream intake and preventing the upstream intake from failing to start. The combined use of the front and rear injection devices shortens the combustion chamber length. Attached Figure Description

[0045] Figure 1 This is a perspective view of the combustion chamber of a multi-cavity scramjet engine using aviation kerosene as fuel, as described in Example 1.

[0046] Figure 2 This is a structural diagram of the combustion chamber of a multi-cavity scramjet engine using aviation kerosene as fuel, as described in Example 1.

[0047] Figure 3 Combustion flow field diagram of the combustion chamber during cruise phase when the equivalence ratio is 0.1 and the total incoming pressure is 1.0 MPa;

[0048] Figure 4 A comparison chart of combustion performance at different injection positions;

[0049] Figure 5A comparison diagram of the combustion efficiency of the first cavity group under different states of the front and rear injection devices;

[0050] Figure 6 A comparison diagram of wall pressure when only the front injection device is open and when only the rear injection device is open;

[0051] Figure 7 A comparison chart of combustion efficiency at total inflow pressures of 0.8 MPa to 1.2 MPa;

[0052] Figure 8 This is a perspective view of the combustion chamber after adding an auxiliary injection device in Example 3;

[0053] Figure 9 This is a structural diagram of the combustion chamber after adding an auxiliary injection device in Example 3;

[0054] Figure 10 This is a comparison chart showing the combustion oscillation control effect of auxiliary injection during the combustion chamber ramp-up phase in Example 3.

[0055] Wherein: 1-Isolation section; 2-First cavity group; 3-Second cavity group; 4-Third cavity group; 5-Combustion chamber expansion section; 6-Rear injection device; 7-Front injection device; 8-Second square tube; 9-Third square tube; 10-Auxiliary injection device. Detailed Implementation

[0056] The present invention will now be described in further detail with reference to the accompanying drawings and embodiments.

[0057] Example 1:

[0058] This embodiment provides a multi-cavity scramjet engine combustor that is simple in structure, adaptable to different equivalence ratio ranges and different flight altitudes, and is also short in length, has low drag, high combustion performance, low heat load on the combustor wall, is easy to maintain, and facilitates the integrated design of the aircraft and engine.

[0059] like Figure 1 and Figure 2 As shown, the cross-section of the multi-cavity scramjet engine is square, including: an isolation section 1, a cavity group, an injection module, and a combustion chamber expansion section 5; three cavity groups are connected in series between the isolation section 1 and the combustion chamber expansion section 5, namely the first cavity group 2, the second cavity group 3, and the third cavity group 4; each cavity group has two oppositely arranged cavities, that is, the two cavities in each cavity group are connected in parallel along the airflow direction (i.e., including the upper cavity and the lower cavity); thus, the combustion chamber has a total of six cavities, divided into three groups.

[0060] Isolation section 1 is a first square tube with an expansion angle, and isolation section 1 is connected to the first cavity group 2 (e.g., connected together by a fastener); the first cavity group 2 is connected to the second cavity group 3 via a second square tube 8 with an expansion angle (e.g., connected together by a fastener); the second cavity group 3 is connected to the third cavity group 4 via a third square tube 9 with an expansion angle (e.g., connected together by a fastener), and the combustion chamber expansion section 5 is connected to the rear of the third cavity group 4 (e.g., connected together by a fastener). The fasteners can be screws, rivets, or sealing strips. The area between the first cavity group 2 and the third cavity group 4 is the main combustion zone.

[0061] As an example, the lower walls of the isolation section 1, the second square tube 8, and the third square tube 9 are horizontal, while the upper walls are upward-sloping, thus forming an expansion angle; the upward-sloping angle of the upper walls (the angle between the upper walls and the horizontal plane) does not exceed 1°.

[0062] For ease of description, let the end where the isolation section 1 is located be the front end of the combustion chamber, and the end where the combustion chamber expansion section 5 is located be the rear end of the combustion chamber.

[0063] The upper and lower walls of the two concave cavities in the first cavity group 2 are horizontal, and the rear wall has a slope. That is, the surface of the two concave cavities in the first cavity group 2 that is connected to the second square tube 8 is a slope that slopes backward. As an example, the rear wall of the two concave cavities in the first cavity group 2 is a 45° slope (that is, the angle between it and the axis of the combustion chamber is 45°).

[0064] The upper and lower walls of the two concave cavities in the second set of concave cavities 3 are horizontal, and the rear wall has a slope. That is, the surface of the two concave cavities in the second set of concave cavities 3 that are connected to the third third tube 9 is a slope that slopes backward. As an example, the rear wall of the two concave cavities in the second set of concave cavities 3 is a 45° slope (that is, the angle between it and the axis of the combustion chamber is 45°).

[0065] In the third cavity group 4, the upper and lower walls of the two cavities are horizontal, and the rear wall has a slope. That is, the surface of the two cavities in the third cavity group 4 that connects to the combustion chamber expansion section 5 is a rearward inclined slope. As an example, the rear wall of the two cavities in the third cavity group 4 is a 45° slope (that is, the angle between it and the axis of the combustion chamber is 45°).

[0066] The lower wall of the combustion chamber expansion section 5 is horizontal (and thus parallel to the lower wall of the lower cavity in the three cavity groups), and the upper wall has an expansion angle; as an example, the upper wall of the combustion chamber expansion section 5 is an upwardly inclined slope, thereby forming an expansion angle; the upper end face has an upwardly inclined angle (the angle with the horizontal plane) of no more than 6°.

[0067] In the cavity group, the first cavity group 2 serves as a mixing cavity, the second cavity group 3 serves as a main combustion cavity, and the third cavity group 4 serves as a secondary auxiliary combustion cavity.

[0068] The airflow enters through the inlet of isolation section 1, and after passing through isolation section 1, enters the first concave cavity group 2; the rear walls of the two concave cavities in the first concave cavity group 2 are sloped. The high-speed airflow mixes with the vortex-induced backflow gas generated in the first concave cavity group 2, forming a vortex at the center of the high-speed airflow. The high-speed airflow carrying the vortex flows again through the second square pipe 8 with an expansion angle and enters the second concave cavity group 3. Its function is: firstly, to provide an adhesion surface for the upstream shock wave surface of the thermal throat in the scramjet engine, and to separate it from the upstream intake duct, so as to prevent the thermal throat shock wave surface from entering the upstream intake duct and affecting the intake performance of the intake duct. Secondly, the presence of the first concave cavity group 2 allows for the formation of entrained vortices in the high-speed airflow, enhancing its turbulent characteristics and providing a better airflow environment for fuel mixing in the second set of concave cavities. Thirdly, the presence of the second square tube 8 behind the first concave cavity group 2 allows the entrained vortices in the high-speed airflow to fully develop. Fourthly, the presence of the 45° sloping wall of the concave cavity in the first concave cavity group 2 creates a 45° angle between the gas in the vortex recirculation zone and the incoming gas, thus achieving a balance between the total pressure loss of the recirculation zone gas on the incoming flow and the enhanced mixing of the incoming flow. Therefore, the first concave cavity group 2 is used as a mixing cavity in the high-speed airflow, and in conjunction with the second square tube 8 between the first concave cavity group 2 and the second concave cavity group 3, it allows entrained vortices to form and fully develop in the high-speed airflow, enhancing mixing without significantly increasing the resistance in the high-speed airflow.

[0069] During the high-altitude cruise phase, when the equivalent volume is relatively large, a sub-combustion mode forms in the combustion chamber, with combustion mainly concentrated in the central airflow near the second concave cavity group 3. When the equivalent volume is relatively small, a supercombustion mode forms in the combustion chamber, with combustion mainly concentrated within the second concave cavity group 3. At this time, the second concave cavity group 3 exhibits localized fuel richness, resulting in incomplete combustion and a large amount of unburned fuel and waste. Therefore, the third concave cavity group 4 again passes through the vortex recirculation zone within the concave cavity. The vortex in the recirculation zone entrains the high-speed incoming gas into the concave cavity, mixing it with the exhaust gas from the second concave cavity group 3, thus undergoing secondary combustion. The high-temperature gas after combustion in the third concave cavity group 4 is ejected into the high-speed gas center through the 45° inclined wall of the concave cavity in the third concave cavity group 4. The 45° angle between the rear edge walls of the second cavity group 3 and the third cavity group 4 allows the high-temperature gas in the cavity to be ejected into the high-speed incoming gas more quickly, resulting in a lower temperature near the pipe wall behind the cavity, and also reduces the total pressure loss caused by the excessive angle between the return gas and the high-speed incoming gas in the cavity.

[0070] The upper wall of the combustion chamber expansion section 5 has an expansion angle, which accelerates the high-temperature supersonic gas flowing through the third cavity group 4, thereby achieving better momentum performance. Simultaneously, the expansion angle allows the wall surface to be positioned away from the high-temperature, high-speed gas, reducing the thermal load on the expansion section wall and thus mitigating the negative mass caused by high-heat cooling. The lower wall of the combustion chamber expansion section is parallel to the lower walls of the lower cavities in the preceding three cavity groups, allowing it to be inserted into the aircraft and parallel to the aircraft's lower surface. This reduces flight drag caused by the combustion chamber during flight, achieving an integrated aircraft-engine design and improving aircraft performance.

[0071] The spraying unit includes a front spraying device 7 and a rear spraying device 6, both of which use small holes in the wall for spraying. The front spraying device 7 is located on the wall of the second square tube 8, specifically, the front spraying device 7 is located opposite the upper and lower walls of the second square tube 8; the rear spraying device 6 is located on the wall of the third third tube 9, specifically, the rear spraying device 6 is located opposite the upper and lower walls of the third third tube 9.

[0072] As an example, both the front spraying device 7 and the rear spraying device 6 are provided with nine spraying holes evenly distributed along the transverse direction (i.e., the width direction of the second square tube 8 and the third third tube 9). Specifically, nine spraying holes are arranged at equal intervals on each of the upper and lower walls of the second square tube 8; and nine spraying holes are arranged at equal intervals on each of the upper and lower walls of the third third tube 9. The diameter of the spraying holes is 0.3 mm to 0.6 mm, preferably 0.5 mm (0.5 mm droplets have a larger penetrating momentum). The distance between the central axis of the front spraying device 7 and the front wall of the second cavity group 3 is 55 mm to 75 mm, preferably 65 mm, and the distance between the central axis of the rear spraying device 6 and the front wall of the third cavity group 4 is 55 mm to 75 mm, preferably 65 mm.

[0073] The front injection device 7 is used during the circulation phase in a high-altitude, low-pressure environment. Working in conjunction with the concave cavity group, fuel is combusted through the second concave cavity group 3 and the third concave cavity group 4, improving high-altitude, low-pressure combustion performance. The rear injection device 6 is used during the climb phase in a low-altitude, high-pressure environment. The injection position of the rear injection device 6 is far from the inlet of the isolation section 1, ensuring that the upstream shock wave surface of the thermal throat generated at a higher equivalence ratio is far from the inlet of the isolation section. This improves the performance of the upstream intake and prevents intake failure caused by the upstream shock wave surface of the thermal throat being too close to the intake. Simultaneously, during the climb phase, the first concave cavity group 2 and the second concave cavity group 3 repeatedly increase entrainment vortices in the high-speed incoming gas. This increased entrainment vortex allows for more thorough mixing of the fuel and the high-speed incoming gas during the climb phase, thereby improving combustion heat release and combustion efficiency, reducing fuel consumption during the climb phase, and ultimately increasing the range of the scramjet engine.

[0074] Tests have shown that the combustion chamber of this scramjet engine has the following advantages:

[0075] (1) The third cavity group 4 can be used as a secondary auxiliary combustion:

[0076] During the high-altitude cruise phase, the front injection device 7 is activated and the rear injection device 6 is deactivated. At this time, the first cavity group 2 is used for mixing, improving mixing efficiency; the second cavity group 3 is used as the primary combustion cavity; and the third cavity group 4 is used for secondary auxiliary combustion. The exhaust gas produced after combustion in the second cavity group 3 enters the third cavity group 4 for further combustion, improving combustion efficiency. A significant combustion reaction also occurs in the third cavity group 4, thereby improving combustion efficiency. Figure 3 As shown (equivalent ratio 0.1, total inflow pressure 1.0 MPa), it can be clearly observed that heat release is significantly increased at the rear edge of the third cavity group 4. Therefore, the third cavity group 4 improves its combustion heat release, thereby improving combustion performance.

[0077] (2) The combustion chamber of this scramjet engine is capable of efficient combustion over a wide stoichiometric range.

[0078] like Figure 4 As shown, it can be clearly seen that when only the front injection device 7 or only the rear injection device 6 is open, the combustion efficiency is very high, exceeding 90%, within the range of equivalence ratio ER = 0-1.

[0079] (3) The first cavity group 2 can improve fuel blending, thereby improving combustion performance.

[0080] like Figure 5 As shown, it can be clearly seen that when only the front injection device 7 or only the rear injection device 6 is open, the combustion efficiency is much higher when the first cavity group 2 is present than when the first cavity group is not present.

[0081] (4) Capable of adapting to both climbing and cruising phases

[0082] During the climb phase, the total incoming pressure is high, and the equivalent flow rate is large. If the front injection device 7 is opened, the thermal throat shock surface is too close to the inlet of the isolation section. This increases the pressure at the inlet of isolation section 1, affecting the starting performance of the intake duct upstream of isolation section 1, thus preventing the intake duct from starting. However, when the rear injection device 6 is opened, the pressure near isolation section 1 does not change significantly, and the upstream thermal throat shock surface is moved away from the inlet of isolation section 1, thereby improving the starting performance of the intake duct. Figure 6 As shown, when the total incoming pressure is 1.0 MPa and the equivalent ratio ER = 0.6, the front injection device 7 is opened, and the thermal throat shock surface affects the upstream intake.

[0083] Furthermore, combustion performance is also higher under different total incoming pressures during acceleration / deceleration and climb-up phases. For example... Figure 7As shown, the combustion efficiency is calculated when the total inflow pressure is 0.8 MPa, 0.9 MPa, 1.0 MPa, 1.1 MPa, and 1.2 MPa, respectively, within the equivalence ratio range ER = 0 to 1. Figure 7 It can be seen that the combustion efficiency is relatively high.

[0084] (5) While taking into account integration, it also takes into account cooling, combustion-mixing, and integrated design.

[0085] The 45° rear wall slope angle of the three concave cavity groups allows the high-temperature gas to move away from the wall while balancing the total pressure loss and fuel mixing. The square tube with an expansion angle between adjacent concave cavity groups, and the expansion angle of the upper wall of the combustion chamber expansion section 5, keep the high-temperature gas away from the wall, thus also taking into account cooling.

[0086] The lower wall of the combustion chamber expansion section 5 is horizontal, so that when the scramjet engine combustion chamber is inserted into the aircraft, the lower wall is flush with the lower surface of the aircraft, thus realizing the integrated design of the aircraft and engine.

[0087] In summary, this scramjet engine combustor not only achieves high-efficiency combustion over a wide equivalence ratio range during the cruise phase in a high-altitude, low-pressure environment, but also adapts to the climb phase in a low-altitude, high-pressure environment, achieving efficient combustion without affecting intake conditions. Furthermore, the horizontal design of the combustor's lower wall facilitates the integrated design of hypersonic vehicles, thereby reducing drag caused by the combustor. The multi-cavity design further improves combustion performance while shortening the combustor length. Additionally, the small expansion angle of the upper wall of the ducts behind each cavity group reduces wall thickness and the difficulty of thermal protection, thus reducing engine mass. This device effectively solves the current problems of scramjet engines being unable to simultaneously achieve efficient combustion over a wide equivalence ratio range during both high-altitude cruise and low-altitude climb phases, and the significant weight and difficulty of thermal protection associated with existing systems.

[0088] Example 2:

[0089] Based on the above embodiment 1, this embodiment provides an aircraft having the cavity scramjet engine combustion chamber of the above embodiment 1 and its control method.

[0090] The concave scramjet engine combustion chamber of Embodiment 1 described above is located inside the engine of the aircraft.

[0091] When an aircraft using this scramjet engine combustion chamber operates during the low-altitude climb phase, the valve connected to the forward injection device 7 is closed, and the valve connected to the rear injection device 6 is opened. The opening of the valve connected to the rear injection device 6 is adjusted according to the aircraft's control requirements to increase the fuel volume. As the aircraft approaches high-altitude cruise, the valve connected to the rear injection device 6 is gradually closed (i.e., the valve opening of the rear injection device 6 is gradually reduced), and the valve connected to the forward injection device 7 is gradually opened (i.e., the valve opening of the forward injection device 7 is gradually increased), causing the aircraft to begin transitioning to the high-altitude cruise phase. This continues until the valve connected to the rear injection device 6 is completely closed and the valve connected to the forward injection device 7 is opened according to the required fuel volume, indicating that the aircraft has fully transitioned to the high-altitude cruise phase. Operation then continues until the aircraft completes its flight.

[0092] Example 3:

[0093] Based on the above embodiment 1, as follows Figure 8 and Figure 9 As shown, the further injection unit also includes an auxiliary injection device 10; the auxiliary injection device 10 is disposed on the two cavities of the second cavity group 3; that is, the auxiliary injection device 10 is respectively disposed opposite to each other on the wall surface of the two cavities of the second cavity group 3. The auxiliary injection device 10 also adopts wall-mounted small hole injection, including a number of injection holes evenly distributed in the transverse direction.

[0094] As an example, the auxiliary injection device 10 is positioned at the center of the two cavity walls of the second cavity group 3.

[0095] During the low-altitude, high-pressure climb phase, when the rear injection device 6 is opened, due to the high total pressure of the incoming flow, large-amplitude combustion oscillations are likely to occur, even exceeding 100%, causing structural damage. By setting up the auxiliary injection device 10, large-amplitude oscillations can be suppressed, thereby optimizing the oscillation amplitude.

[0096] Specifically: when a large-amplitude pressure oscillation occurs (instantaneous maximum pressure - instantaneous minimum pressure) / average pressure * 100% is greater than 10%, it is considered a large-amplitude oscillation. The valve connected to the auxiliary injection device 10 is opened, diverting a small portion of the flow from the injection device 6 through the auxiliary injection device 10. This ensures that the total equivalence ratio remains constant and the overall thrust change is not too significant, while also allowing the auxiliary injection device 10 to suppress the oscillation in the first cavity group, thereby suppressing large-amplitude combustion oscillations in the engine. Figure 10 As shown ( Figure 10In the diagram, the red line represents the oscillation amplitude at different positions in the combustion chamber when the equivalence ratio of the rear injection device 6 is 1, and the green line represents the oscillation amplitude at different positions in the combustion chamber when the equivalence ratio of the rear injection device 6 is 0.7 and the equivalence ratio of the auxiliary injection device 10 is 0.3. When kerosene with an equivalence ratio of 0.3 is injected through the auxiliary injection device 10, and the equivalence ratio of the rear injection device 6 is reduced (e.g., to 0.7), the oscillation amplitude can be significantly reduced by 50%.

[0097] Although the present invention has been described in detail above with general descriptions and specific embodiments, modifications or improvements can be made to it, which will be obvious to those skilled in the art. Therefore, all such modifications or improvements made without departing from the spirit of the present invention fall within the scope of protection claimed by the present invention.

Claims

1. A control method for the combustion chamber of a multi-cavity scramjet engine, characterized in that, The multi-cavity scramjet engine combustion chamber includes: an isolation section (1), a first cavity group (2), a second cavity group (3) and a third cavity group (4), an injection unit and a combustion chamber expansion section (5); Each cavity group has two cavities arranged vertically opposite each other. The isolation section (1) is a first square tube connected to the first cavity group (2); the first cavity group (2) is connected to the second cavity group (3) through the second square tube (8); the second cavity group (3) is connected to the third cavity group (4) through the third third tube (9); and the combustion chamber expansion section (5) is connected to the rear of the third cavity group (4). The injection unit includes: a front injection device (7) and a rear injection device (6); the front injection device (7) is arranged opposite to the upper and lower walls of the second square tube; the rear injection device (6) is arranged opposite to the upper and lower walls of the third square tube; the front injection device (7) is used in the circulation section in the high-altitude low-pressure environment, and the rear injection device (6) is used in the climbing section in the low-altitude high-pressure environment; The injection unit also includes an auxiliary injection device (10). An auxiliary injection device (10) is respectively provided on the two concave cavities of the second cavity group (3). When an aircraft using the scramjet engine combustion chamber is operating in the low-altitude climb phase, the valve connected to the front injection device (7) is closed, and the valve connected to the rear injection device (6) is opened. According to the aircraft control requirements, the opening degree of the valve connected to the rear injection device (6) is adjusted to increase the fuel volume. Once the aircraft reaches the set altitude, the opening of the valve connected to the rear injection device (6) is gradually reduced, and the opening of the valve connected to the front injection device (7) is gradually increased. The aircraft continues to climb. When the aircraft climbs to the high-altitude cruising altitude, the valve connected to the rear injection device (6) is completely closed, and the valve connected to the front injection device (7) is opened according to the required amount of fuel. The aircraft then fully enters the high-altitude cruising phase. Then it continues to operate until the aircraft completes its journey. When auxiliary injection devices (10) are respectively provided opposite to each other on the two cavities of the second cavity group (3): During the low-altitude climb phase, when a large-amplitude pressure oscillation occurs, the valve connected to the auxiliary injection device (10) is opened, and kerosene with a set equivalent ratio is injected through the auxiliary injection device (10). Under the premise of ensuring that the total equivalent ratio remains unchanged, the large-amplitude pressure oscillation is suppressed. When (instantaneous maximum pressure - instantaneous minimum pressure) / average pressure * 100% is greater than 10%, it is considered that a large-amplitude oscillation has occurred.

2. The control method for the combustion chamber of a multi-cavity scramjet engine as described in claim 1, characterized in that, The isolation section (1) is a first square tube with an expansion angle, and its upper wall is an upwardly inclined slope with an upward inclination angle not exceeding 1°.

3. The control method for the combustion chamber of a multi-cavity scramjet engine as described in claim 1, characterized in that, The upper walls of the second square tube and the third third tube are upwardly inclined slopes, and the upward inclination angle does not exceed 1°.

4. The control method for the combustion chamber of a multi-cavity scramjet engine as described in claim 1, characterized in that, The upper wall of the combustion chamber expansion section (5) is inclined upwards, and the angle of the upward inclination does not exceed 6°.

5. The control method for the combustion chamber of a multi-cavity scramjet engine as described in any one of claims 1-4, characterized in that, The rear wall surfaces of two of the three concave cavities in the first concave cavity group (2), the second concave cavity group (3), and the third concave cavity group (4) are inclined surfaces tilted backward at 45°.

6. The control method for the combustion chamber of a multi-cavity scramjet engine as described in any one of claims 1-4, characterized in that, The lower wall surfaces of the isolation section (1), the first cavity group (2), the second cavity group (3), the third cavity group (4), and the combustion chamber expansion section (5) are horizontal.

7. The control method for the combustion chamber of a multi-cavity scramjet engine as described in any one of claims 1-4, characterized in that, Both the front spraying device (7) and the rear spraying device (6) are provided with a number of spraying holes that are equally spaced along the transverse direction; The diameter of the injection hole is 0.3mm to 0.6mm; The distance between the central axis of the front injection device (7) and the front wall of the second cavity group (3) is 55mm to 75mm; The distance between the central axis of the rear injection device (6) and the front wall of the third cavity group (4) is 55mm to 75mm.

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