Multi-cavity scramjet engine combustion chamber and control method thereof

Abstract

The invention provides a multi-cavity scramjet engine combustion chamber and a control method thereof. The combustion chamber is simple in structure, capable of adapting to different equivalence ratio ranges and different flight heights, low in resistance, high in combustion performance, low in thermal load of the wall face of the combustion chamber, convenient to maintain and capable of facilitating aircraft-engine integrated design. The combustion chamber comprises an isolation section, a first concave cavity group, a second concave cavity group, a third concave cavity group, an injection module and a combustion chamber expansion section, each concave cavity group is provided with two concave cavities which are oppositely arranged up and down; the isolation section is a first square tube and is connected with the first concave cavity group; the first concave cavity group is connected with the second group of concave cavities through a second square tube; the second group of concave cavities is connected with a third concave cavity group through a third square pipe; a combustion chamber expansion section is connected behind the third concave cavity group; the injection unit comprises a front injection device and a rear injection device; front injection devices are oppositely arranged on the upper wall surface and the lower wall surface of the second square pipe; rear injection devices are oppositely arranged on the upper and lower wall surfaces of the third square pipe.

Description

Technical Field

[0001] The invention relates to a combustion chamber, in particular to a multi-cavity scramjet engine combustion chamber, and belongs to the technical field of scramjet engine combustion chambers. Background Art

[0002] Scramjet engines are suitable for use in hypersonic vehicles and are a hot topic of research. Their combustion chamber speed is high and the fuel mixing efficiency is low, so they need to use concave cavities and support plates as devices to enhance fuel mixing and stabilize the flame. When the concave cavity is used as a flame stabilizer, it has a simple structure and low resistance, and has become a commonly used flame stabilization device for scramjet engines.

[0003] In a combustion chamber of limited length, improving combustion efficiency, widening the combustion boundary and reducing total pressure loss are one of the important research goals of scramjet engines. The injection position has a significant impact on the combustion performance of scramjet engines. Existing studies have shown that in scramjet engines, changing the injection position of the fuel can change the combustion performance, thereby changing the thrust of the combustion chamber. By studying the flame stability of scramjet engines at different injection positions, it was found that the stability of the flame at different injection positions is not the same. A study of scramjet engines fueled by ethylene found that compared with single-hole injection, although the combustion efficiency is higher during multi-hole injection, the flame stability is lower. Existing studies have found that in scramjet engines, the nozzle diameter has a greater impact on the injection position. When the nozzle diameter is small, the injection position is close to the front edge of the cavity, and the total kerosene equivalence ratio required for successful ignition is low; when the nozzle diameter is large, the injection position has little effect on the ignition characteristics of kerosene fuel.

[0004] However, in the publicly published literature, including the above-mentioned research, most scramjet engines have only one or a pair of cavities in the flow direction (i.e., two cavities in parallel). These scramjet engines are difficult to maintain high combustion efficiency and low total pressure loss in a wide equivalence ratio range. In addition, existing studies have found that under the same length, the configuration of more cavities in the flow direction can greatly improve the combustion efficiency, while the total pressure loss does not increase much. And multi-cavity scramjet engines can adapt to a wide range of flight conditions. Therefore, configuring more cavities in the flow direction is suitable for high Mach number scramjet engines.

[0005] However, when more cavities are configured in the flow direction, the flow field will become more complicated. The main fuels of the multi-cavity scramjet engines currently studied are hydrogen fuel, ethylene, and kerosene. Among them, hydrogen fuel and ethylene have low density and are difficult to store, and large-scale practical applications have not yet been realized. However, aviation kerosene has high mass per unit volume and is easy to store. It is widely used in various existing conventional aircraft. Therefore, the scramjet engine using aviation kerosene as fuel has engineering application advantages and is expected to provide a power source for hypersonic aircraft. However, it is difficult to organize the combustion of the scramjet engine using aviation kerosene as fuel, especially the combustion at high Mach numbers, which needs further research. For example, in the prior art, a dual-mode scramjet engine fueled by kerosene with a combustion chamber inlet Mach number Ma=2.0 was studied, and it was found that the combustion efficiency was very low and the total pressure loss was high, which requires further optimization. Therefore, it is necessary to carry out research on multi-cavity scramjet engines fueled by kerosene at high Mach numbers to improve the combustion efficiency, reduce the total pressure loss, and thus improve the combustion performance.

[0006] Existing scholars have conducted research on multi-cavity scramjet engines fueled by kerosene at high Mach numbers from different directions. For example, in a six-cavity scramjet engine, the cavity is deployed in front of the injection, which can enhance combustion mixing and improve combustion efficiency. The injection effect of a three-cavity scramjet engine at the Mach number Ma=2.92 at the combustion chamber inlet was studied, and it was found that the injection of the upstream separation zone induced by combustion had too much influence on the fuel injection, which may increase the penetration depth and improve the combustion efficiency. In addition, existing studies have also found that in a multi-cavity scramjet engine, the total pressure loss of the combustion chamber is mainly the loss along the combustion chamber. Shortening the length of the combustion chamber can greatly reduce the total pressure loss. There are also studies on four-cavity scramjet engines in the prior art, and the Mach number of the combustion chamber inlet Ma=3.0. It was found that with the change of the equivalence ratio, there is a large mutation in the combustion intensity and back pressure, which seriously affects the stability of the flame. In addition, the study of the four-cavity scramjet engine fueled by aviation kerosene found that the front and rear cavities of the cavity-series cavity combustion chamber affect each other, causing a sudden change in thrust when the equivalence ratio changes. By conducting a study on the flame position of the cavities in series under different total pressures, it was found that the flame stability point is downstream under low equivalence ratios and upstream under high equivalence ratios, and the flame stability point propagates from downstream to upstream as the total pressure increases.

[0007] Although preliminary research has been conducted on the multi-cavity scramjet engine fueled by aviation kerosene at high Mach numbers from the aspects of fuel injection, fuel blending, total pressure loss-combustion chamber length, equivalence ratio-combustion intensity, equivalence ratio-thrust, total pressure change-flame stability performance, and combustion chamber configuration, in an attempt to improve combustion efficiency and reduce combustion chamber length, in the multi-cavity scramjet engine fueled by aviation kerosene at high Mach numbers, the combustion efficiency within a limited length is still low, and its combustion efficiency still needs to be further improved. The injection position has a great influence on its combustion performance. Therefore, the injection position is reasonably deployed to organize the combustion of the scramjet engine fueled by aviation kerosene in order to improve its combustion performance.

[0008] At the same time, the low-speed backflow airflow in the cavity has a great influence on the mixing of fuel, flame stability, and total pressure loss when entering the mainstream at different angles. When entering the mainstream at a smaller angle, the total pressure loss is small, but the flame stability performance and the mixing of fuel, as well as the fluid exchange performance between the cavity and the mainstream are poor. When entering the mainstream at a larger angle, the flame stability performance and the mixing of fuel, as well as the fluid exchange performance between the cavity and the mainstream are higher, but the total pressure loss also increases.

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

[0010] In addition, the change of intake pressure has a great influence on the combustion performance of scramjet, especially in the high-altitude cruising stage, the total pressure of the incoming flow is low, which has a great influence on the combustion performance. For example, existing studies believe that with the decrease of the total pressure of the incoming flow, on the one hand, the pressure in the combustion chamber decreases, resulting in a decrease in the chemical reaction rate; on the other hand, the fuel atomization quality is not ideal, the average droplet size is enlarged, the mixing effect of fuel and air becomes worse, and the stability of the flame stabilizer decreases. In addition, the prior art studies the combustion characteristics of ethylene-fueled scramjet engines when the total intake pressure is 1.0, 1.3 and 1.6MPa, and finds that when the total intake pressure is low, ignition becomes more difficult; however, when the total intake pressure is high, the combustion becomes more chaotic, and there is no dominant oscillation frequency in the combustion chamber. A double-cavity scramjet engine was also studied, and it was reported that in the upward stage from the total intake pressure Pt0 = 1.5-1.9MPa, the conversion between the ramjet engine and the scramjet engine mode caused by the change of the total intake pressure leads to a sudden change in thrust. The influence of wall injection when the total pressure of the incoming flow is 1.2, 1.6 and 2.0 MPa was studied, and it was found that with the increase of the total pressure of the incoming flow, the penetration depth and mixing efficiency of the fuel were significantly improved. The modal conversion experiment was carried out on the scramjet engine under different total pressures of the incoming flow. The results show that when the total pressure of the incoming flow is in the range of 0.6-0.9 MPa, the modal conversion equivalence ratio remains unchanged regardless of how the total pressure changes. In addition, the wall cooling effect of the hydrogen-fueled scramjet engine under two inlet pressures of 10 bar and 15 bar was studied. It was found that except for the autoignition caused by the shock wave-boundary layer-interaction, the static wall pressure distribution is basically unrelated to the changes in the static wall temperature and cooling efficiency. The combustion performance of the scramjet engine in the range of 0.6-0.9 MPa total inlet pressure was studied, and it was found that stable combustion can be achieved at higher and lower inlet total pressures, while obvious combustion oscillations will occur at medium inlet total pressures. In addition, during the low-altitude climb phase, the total pressure of the incoming flow is high. In the case of a large equivalence ratio, when the isolation section is short, the shock wave surface of the thermal throat is close to the isolation section entrance. It may even pass through the isolation section entrance and enter the inlet outlet surface, affecting the starting performance of the inlet.

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

[0012] In addition, many existing combustion chamber configurations do not take into account combustion chamber cooling and aircraft integration. Existing studies have shown that the traditional aircraft-engine separation configuration will bring great resistance to the aircraft during hypersonic flight, and even affect the aircraft's range and operational performance. Therefore, in the existing hypersonic aircraft shape design and overall design, the engine is required to meet the aircraft-engine integration design requirements, so that the engine shape can also serve as part of the aircraft's lift surface, reducing the drag loss caused by the engine structure.

[0013] Therefore, the existing multi-cavity scramjet engine using aviation kerosene as fuel needs to solve the following problems:

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

[0015] 2. From the low-altitude climbing stage to the high-altitude cruising stage, the total pressure of the incoming flow changes greatly from high to low. In the low-altitude climbing stage, the total pressure of the incoming flow of the existing scramjet engine is relatively high, and it is easy for the thermal throat to cross the entrance of the isolation section and enter the inlet outlet surface, resulting in the intake failure. In the cruising stage, due to the low total pressure of the incoming flow, the fuel mixing performance and flame stability performance of the cavity are low, and the problem of low combustion performance is easy to occur.

[0016] 3. The process of designing a multi-cavity stable combustion scramjet engine needs to take into account the 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 combustion chamber with a simple structure, which can adapt to different equivalence ratio ranges and different flight altitudes, has low resistance, high combustion performance, low heat load on the combustion chamber wall, is easy to maintain, and is convenient for aircraft-engine integrated design.

[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 opposite to each other up and down;

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

[0021] The injection unit comprises: a front injection device and a rear injection device; the front injection device is disposed oppositely on the upper and lower walls of the second square tube; and the rear injection device is disposed oppositely on the upper and lower walls of the third square tube.

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

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

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

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

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

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

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

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

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

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

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

[0033] When the aircraft using the scramjet engine combustion chamber is working in the low-altitude climbing stage, the valve connected to the front injection device is closed, the valve connected to the rear injection device is opened, and the opening of the valve connected to the rear injection device is adjusted according to the aircraft control requirements to increase the oil volume;

[0034] When the aircraft reaches the set altitude, the opening of the valve connected to the rear injection device is gradually reduced, and the opening of the valve connected to the front injection device is gradually increased, and the aircraft continues to climb to the cruising altitude; when 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 oil volume, and the aircraft completely enters the high-altitude cruising stage; then it continues to operate until the aircraft completes the voyage.

[0035] As a preferred embodiment of the present invention, when auxiliary injection devices are respectively arranged on two concave cavities of the second concave cavity group:

[0036] During the low-altitude climbing stage, when large-amplitude pressure oscillation occurs, the valve connected to the auxiliary injection device is opened, and kerosene with a set equivalence ratio is injected through the auxiliary injection device to suppress the large-amplitude pressure oscillation.

[0037] Beneficial effects:

[0038] (1) The multi-cavity scramjet engine combustion chamber using aviation kerosene as fuel of the present invention has three cavity groups. The first cavity group is used for mixing, destroying the boundary layer of the airflow, increasing the turbulence in the airflow, and not significantly increasing the resistance in the airflow. The third cavity group is used for secondary auxiliary combustion, mixing the combustion 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, enhancing combustion, and not significantly increasing the resistance of the airflow.

[0039] Therefore, in the high-altitude and low-pressure cruise stage: after the airflow enters the first cavity group through the isolation section, the first cavity group forms an entrained vortex in the high-speed airflow. The high-speed airflow with the entrained vortex improves the mixing of the high-speed airflow and the fuel near the front injection device. The third cavity group mixes and burns the tail gas generated by the combustion of the second cavity group with the high-speed airflow again, thereby improving the combustion heat release and combustion efficiency. In the low-altitude and high-pressure climbing stage, the first cavity group and the second cavity group allow the high-speed incoming gas to have more entrained vortices, so that in the third cavity group, the fuel injected by the rear injection device is fully mixed with the high-speed airflow, thereby improving the combustion heat release within the limited combustion chamber length and reducing the optimization of the climbing stage. The combustion chamber can maintain a high combustion efficiency within the equivalent ratio range of 0 to 1.1 in the high-altitude and low-pressure cruise stage and the low-altitude and high-pressure climbing stage, adapt to different complex environments, have high reliability, and improve combustion efficiency, reduce total pressure loss, and thus improve range and thrust-to-weight ratio.

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

[0041] (3) In the present invention, the upper wall surface of the combustion chamber expansion section is inclined upward, 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 farther away from the wall surface, reducing the negative mass caused by the cooling heat load. The expansion angle is controlled within a certain range to reduce the backflow caused by a too large expansion angle.

[0042] (4) In the present invention, the existence of the 45° sloped wall surface of the concave cavity in the first concave cavity group makes the gas in the trapped vortex recirculation zone in the concave cavity and the incoming gas form a 45° angle, thereby achieving a balance between the total pressure loss of the gas in the recirculation zone to the incoming flow and the enhanced mixing of the incoming flow; the 45° angle of the trailing edge wall surface of the second and third concave cavity groups makes the high-temperature gas in the concave cavity be ejected into the high-speed incoming gas more quickly, making the temperature near the pipe wall surface behind the concave cavity lower, and reducing the total pressure loss caused by the excessively large inclination angle between the recirculation gas in the concave cavity and the high-speed incoming gas.

[0043] (5) In the present invention, the bottom of the combustion chamber expansion section is flush with the isolation section and the level of the three cavity groups, which is convenient for integrated design with the aircraft shape, has a simple structure, and is easy to maintain; and can reduce the resistance of the aircraft in high-speed airflow.

[0044] (6) In the present invention, front and rear injection devices are provided; the front injection device is used under high-altitude cruising conditions, so that sufficient combustion can be achieved in a longer combustion chamber length under low-pressure conditions under high-altitude cruising, and the third cavity group is used as secondary auxiliary combustion to improve the combustion efficiency of fuel in high-altitude cruising; the rear injection device injects fuel at low altitudes, so that the thermal throat upstream shock wave surface generated under high-pressure conditions at low altitudes is far away from the entrance of the isolation section, thereby improving the intake performance of the upstream air intake duct and preventing the upstream air intake duct from failing to start. The coordinated use of the front injection device and the rear injection device shortens the length of the combustion chamber. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] Figure 1 A three-dimensional diagram of a combustion chamber of a multi-cavity scramjet engine using aviation kerosene as fuel according to Example 1;

[0046] Figure 2 This is a structural diagram of a combustion chamber of a multi-cavity scramjet engine using aviation kerosene as fuel according to Example 1;

[0047] Figure 3 The combustion flow field diagram of the combustion chamber in the cruise stage when the equivalence ratio is 0.1 and the total pressure of the incoming flow is 1.0Mpa;

[0048] Figure 4 This is a comparison chart of combustion performance under different injection positions;

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

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

[0051] Figure 7 This is a comparison chart of combustion efficiency when the total incoming flow pressure is 0.8MPa to 1.2MPa;

[0052] Figure 8 A three-dimensional diagram of the combustion chamber after the auxiliary injection device is added in Example 3;

[0053] Fig. 9 The combustion chamber structure diagram after adding the auxiliary injection device in Example 3;

[0054] Fig.10 This is a comparison chart of the combustion oscillation control effect of auxiliary injection during the combustion chamber climbing stage of Example 3.

[0055] Among them: 1-isolation section; 2-first concave cavity group; 3-second concave cavity group; 4-third concave 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 DESCRIPTION

[0056] The present invention will be further described in detail below in conjunction with the accompanying drawings and embodiments.

[0057] Embodiment 1:

[0058] The present embodiment provides a multi-cavity scramjet engine combustion chamber with a simple structure, which can adapt to different equivalence ratio ranges and different flight altitudes, and has a short length, low resistance, high combustion performance, low combustion chamber wall heat load, easy maintenance, and convenient aircraft-engine integrated design.

[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, which are the first cavity group 2, the second cavity group 3 and the third cavity group 4 respectively; each cavity group has two oppositely arranged cavities, that is, the two cavities in each cavity group are connected in parallel up and down along the airflow direction (that is, including an upper cavity and a lower cavity); thus, the combustion chamber has a total of six cavities, which are divided into three groups.

[0060] The isolation section 1 is a first square tube with an expansion angle, and the isolation section 1 is connected to the first cavity group 2 (such as connected together by a fixing member); the first cavity group 2 is connected to the second cavity group 3 through a second square tube 8 with an expansion angle (such as connected together by a fixing member); the second cavity group 3 is connected to the third cavity group 4 through a third party tube 9 with an expansion angle (such as connected together by a fixing member), and the third cavity group 4 is connected to the combustion chamber expansion section 5 at the rear (such as connected together by a fixing member). The fixing member can be a screw or a rivet, or a sealing strip. The area between the first cavity group 2 and the third cavity group 4 is the main combustion area.

[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, and the upper walls are upwardly inclined, thereby forming an expansion angle; the upward inclination angle of the upper wall (the angle with the horizontal plane) does not exceed 1°.

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

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

[0064] The upper and lower walls of the two cavities in the second group of cavities 3 are horizontal, and the rear wall has a slope, that is, the surfaces of the two cavities in the second group of cavities 3 connected to the third-party tube 9 are inclined surfaces inclined backward; as an example, the rear walls of the two cavities in the second group of cavities 3 are a 45° slope (that is, the angle between them and the axis of the combustion chamber is 45°).

[0065] The upper and lower walls of the two cavities in the third cavity group 4 are horizontal, and the rear wall has a slope, that is, the surfaces of the two cavities in the third cavity group 4 connected to the combustion chamber expansion section 5 are backward-inclined slopes; as an example, the rear walls of the two cavities in the third cavity group 4 are a 45° slope (that is, the angle between them and the axis of the combustion chamber is 45°).

[0066] The lower wall surface of the combustion chamber expansion section 5 is horizontal (thus being parallel to the lower wall surface of the lower concave cavity in the three concave cavity groups), and the upper wall surface has an expansion angle; as an example, the upper wall surface of the combustion chamber expansion section 5 is an upwardly inclined inclined surface, thereby forming an expansion angle; the upper end surface is inclined upward at an angle (the angle with the horizontal plane) not exceeding 6°.

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

[0068] The airflow enters through the inlet of the isolation section 1, flows through the isolation section 1, and enters the first cavity group 2; the rear wall surfaces of the two cavities in the first cavity group 2 have slopes. The high-speed airflow is mixed with the trapped vortex reflow gas generated in the first cavity group 2, and a vortex is formed in the center of the high-speed airflow. The high-speed airflow gas with the vortex flows through the second square tube 8 with an expansion angle again and enters the second cavity group 3. Its functions are: first, to provide an attachment surface for the upstream shock wave surface of the thermal throat in the scramjet engine, and to separate it from the upstream air inlet, so as to prevent the thermal throat shock wave surface from entering the upstream air inlet and affecting the air intake performance of the air inlet. Secondly, the existence of the first cavity group 2 enables the formation of entrained vortices in the high-speed airflow, improves the turbulent characteristics of the high-speed airflow, and provides a better airflow environment for the fuel mixing in the second cavity group; thirdly, the existence of the second square tube 8 behind the first cavity group 2 enables the entrained vortices in the high-speed airflow to develop fully; fourthly, the existence of the 45° slope wall of the cavity in the first cavity group 2 enables the gas in the vortex recirculation zone in the cavity to form a 45° angle with the incoming gas, thereby achieving a balance between the total pressure loss of the gas in the recirculation zone to the incoming flow and the enhancement of the mixing of the incoming flow. Therefore, the first cavity group 2 is used as a mixing cavity in the high-speed airflow, and cooperates with the second square tube 8 between the first cavity group 2 and the second cavity group 3, so that the entrained vortex is formed in the high-speed airflow, and fully developed, and the mixing is enhanced, while the resistance in the high-speed airflow is not significantly increased.

[0069] In the high-altitude cruising stage, when the equivalent is relatively large, a subcombustion mode is formed in the combustion chamber, and the combustion is mainly concentrated in the central airflow near the second cavity group 3; when the equivalent is relatively small, a supercombustion mode is formed in the combustion, and the combustion is mainly concentrated in the second cavity group 3. At this time, the second cavity group 3 shows a local rich combustion situation, so that the combustion in the second cavity group 3 is not sufficient, and there is still a large amount of fuel and incomplete combustion. Therefore, the third cavity group 4 passes through the trapped vortex recirculation zone in the cavity again, and the trapped vortex in the recirculation zone draws the high-speed incoming gas into the cavity, mixes with the exhaust gas in the second cavity group 3, and then burns again. The high-temperature gas after combustion in the third cavity group 4 is ejected into the center of the high-speed gas through the 45° slope wall of the cavity in the third 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, making the temperature near the pipe wall behind the cavity lower, and also reduces the total pressure loss caused by the excessively large inclination angle between the reflux gas in the cavity and the high-speed incoming gas.

[0070] The upper wall of the combustion chamber expansion section 5 has an expansion angle, through which the high-temperature supersonic gas flowing through the third cavity group 4 is accelerated, thereby obtaining better momentum performance. At the same time, the expansion angle can keep the wall away from the high-temperature and high-speed gas, reduce the heat load of the expansion section wall, and thus reduce the negative mass caused by high heat load cooling. The lower wall of the combustion chamber expansion section is parallel to the lower wall of the lower cavity in the first three cavity groups, and can be inserted into the aircraft to achieve parallelism with the lower surface of the aircraft, reducing the flight resistance caused by the combustion chamber during flight, thereby realizing the aircraft-engine integrated design and improving the performance of the aircraft.

[0071] The injection unit includes: a front injection device 7 and a rear injection device 6, both of which adopt wall surface small hole injection. The front injection device 7 is arranged on the wall surface of the second square tube 8, specifically, the upper and lower walls of the second square tube 8 are relatively provided with the front injection device 7; the rear injection device 6 is arranged on the wall surface of the third tube 9, specifically, the upper and lower walls of the third tube 9 are relatively provided with the rear injection device 6.

[0072] As an example, the front injection device 7 and the rear injection device 6 are both provided with 9 injection holes equally spaced in the transverse direction (i.e., the width direction of the second square tube 8 and the third party tube 9), that is, the upper and lower walls of the second square tube 8 are each provided with nine injection holes equally spaced; the upper and lower walls of the third party tube 9 are each provided with nine injection holes equally spaced. The aperture of the injection hole is 0.3mm-0.6mm, preferably 0.5mm (0.5mm droplet particles have a larger penetration momentum). The distance between the central axis of the front injection device 7 and the front wall of the second cavity group 3 is 55mm-75mm, preferably 65mm, and the distance between the central axis of the rear injection device 6 and the front wall of the third cavity group 4 is 55mm-75mm, preferably 65mm.

[0073] The front injection device 7 is used in the circulation stage in the high-altitude low-pressure environment. In conjunction with the cavity group, the fuel is burned in the second cavity group 3 and the third cavity group 4 to improve the low-pressure combustion performance in the high altitude. The rear injection device 6 is used in the climbing stage in the low-altitude high-pressure environment. The injection position of the rear injection device 6 is far away from the entrance of the isolation section 1, so that the upstream shock wave surface of the thermal throat generated at a higher equivalence ratio is far away from the entrance of the isolation section, thereby improving the performance of the upstream air inlet and avoiding the air inlet from not starting due to the upstream shock wave surface of the thermal throat being too close to the air inlet. At the same time, in the climbing stage, the first cavity group 2 and the second cavity group 3 add entrainment vortices to the high-speed incoming gas for many times, so that the fuel in the climbing stage is more fully mixed with the high-speed incoming gas with more entrainment vortices, thereby improving the combustion heat release and combustion efficiency, reducing the fuel consumption in the climbing stage, and thus improving the range of the scramjet engine.

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

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

[0076] When in the high-altitude cruising stage, the front injection device 7 is opened and the rear injection device 6 is closed. At this time, the first cavity group 2 is used for mixing to improve the mixing efficiency; the second cavity group 3 is used as the main combustion cavity. The third cavity group 4 is used for secondary auxiliary combustion. At this time, after the second cavity group 3 burns, the exhaust gas generated enters the third cavity group 4 and continues to burn, which improves the combustion efficiency. An obvious combustion reaction also occurs in the third cavity group 4, thereby improving the combustion efficiency. Figure 3 As shown (equivalence ratio is 0.1, total pressure of incoming flow is 1.0 MPa). It can be clearly found that heat release is significantly increased at the trailing 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 scramjet combustion chamber can burn efficiently over a wide equivalence ratio range

[0078] like Figure 4 As shown, it can be clearly found that when only the front injection device 7 or only the rear injection device 6 is opened, 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 mixing and thus improve combustion performance

[0080] like Figure 5 As shown, it can be clearly found that when only the front injection device 7 or only the rear injection device 6 is opened, when there is the first cavity group 2, the combustion efficiency is much higher than the combustion efficiency when there is no first cavity group.

[0081] (4) Able to adapt to the climbing and cruising phases

[0082] During the climbing phase, the total pressure of the incoming flow is high and the equivalent is relatively large. If the front injection device 7 is opened, the thermal throat shock wave surface is too close to the isolation section entrance. At this time, the pressure at the isolation section 1 entrance increases, affecting the starting performance of the air intake upstream of the isolation section 1, thereby causing the air intake to fail to start. When the rear injection device 6 is opened, the pressure near the isolation section 1 does not change significantly, and the thermal throat shock wave surface upstream is far away from the isolation section 1 entrance, thereby improving the starting performance of the air intake. Figure 6 As shown, when the total pressure of the incoming flow is 1.0 MPa and the equivalence ratio ER=0.6, the front injection device 7 is opened, and the thermal throat shock wave surface affects the upstream air inlet.

[0083] In addition, the combustion performance is also higher under different total pressures of incoming flow during the acceleration and deceleration stages and the climbing stage. Figure 7As shown, when the total pressure of the incoming flow is 0.8MPa, 0.9MPa, 1.0MPa, 1.1MPa, and 1.2MPa, respectively, the combustion efficiency in the equivalence ratio range ER = 0 to 1 is 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° slope angle of the rear wall of the three cavity groups can keep the high-temperature gas away from the wall while balancing the total pressure loss and fuel mixing. The expansion angle of the square tube with expansion angle between adjacent 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, taking into account cooling.

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

[0087] In summary, the scramjet engine combustion chamber can not only burn efficiently in a wide equivalence ratio range during the cruise phase under a high-altitude and low-pressure environment, but also can adapt to the climb phase under a low-altitude and high-pressure environment, and burn efficiently without affecting the intake conditions under low-altitude and high-pressure conditions. At the same time, the horizontal design of the lower wall of the combustion chamber facilitates the integrated design of the hypersonic aircraft, thereby reducing the aircraft resistance caused by the combustion chamber. The combined design of multiple cavities further improves the combustion performance while shortening the combustion chamber length. In addition, the small expansion angle of the upper wall of the rear pipe of each cavity group in the combustion chamber reduces the wall thickness and the difficulty of thermal protection, thereby reducing the weight of the engine. At present, this device can effectively solve the existing problems that the existing scramjet engine cannot adapt to the high-altitude cruise phase and the low-altitude climb phase at the same time, and burns efficiently in a wide equivalence ratio range, and the combustion chamber is heavy and the thermal protection is difficult.

[0088] Embodiment 2:

[0089] On the basis of the above-mentioned embodiment 1, this embodiment provides an aircraft having the cavity scramjet engine combustion chamber of the above-mentioned embodiment 1 and a control method thereof.

[0090] The cavity scramjet engine combustion chamber of the above-mentioned embodiment 1 is arranged inside the engine of the aircraft.

[0091] When the aircraft using the scramjet engine combustion chamber is working in the low-altitude climbing stage, the valve connected to the front injection device 7 is closed, the valve connected to the rear injection device 6 is opened, and the opening of the valve connected to the rear injection device 6 is adjusted according to the aircraft control requirements to increase the amount of oil. When the aircraft is approaching high-altitude cruising, 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 front injection device 7 is gradually opened (i.e., the opening of the valve connected to the front injection device 7 is gradually increased), so that the aircraft begins to enter the high-altitude cruising stage. Until the valve connected to the rear injection device 6 is completely closed, the valve connected to the front injection device 7 is opened according to the required amount of oil, and the aircraft completely enters the high-altitude cruising stage. Then continue to operate until the aircraft completes the voyage.

[0092] Embodiment 3:

[0093] On the basis of the above-mentioned embodiment 1, Figure 8 and Fig. 9 As shown, the injection unit further includes an auxiliary injection device 10; the auxiliary injection device 10 is arranged on the two concave cavities of the second concave cavity group 3; that is, the auxiliary injection device 10 is respectively arranged on the two concave cavity walls of the second concave cavity group 3. The auxiliary injection device 10 also adopts wall surface small hole injection, including a plurality of injection holes distributed at equal intervals in the transverse direction.

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

[0095] During the low-altitude and high-pressure climbing stage, 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, and the oscillation amplitude may even exceed 100%, causing structural damage. By providing the auxiliary injection device 10, the large-amplitude oscillations can be suppressed, thereby achieving optimization of the oscillation amplitude.

[0096] Specifically, when a large-amplitude pressure oscillation occurs (instantaneous maximum pressure - instantaneous minimum pressure) / average pressure * 100%, it is considered that a large-amplitude oscillation occurs when it is greater than 10%), the valve connected to the auxiliary injection device 10 is opened, and a small portion of the flow of the rear injection device 6 is diverted through the auxiliary injection device 10, so that the auxiliary injection device 10 can also suppress the oscillation in the first cavity group while ensuring that the total equivalence ratio remains unchanged and the overall thrust change is not too obvious, thereby suppressing the large-amplitude combustion oscillation of the engine; Fig.10 As shown ( Fig.10In the figure, the red line is the oscillation amplitude at different positions of the combustion chamber when the equivalence ratio of the rear injection device 6 is 1, and the green line is the oscillation amplitude at different positions of 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 (for example, the equivalence ratio is 0.7), the oscillation amplitude can be greatly reduced by 50%.

[0097] Although the present invention has been described in detail above by general description and specific embodiments, it is obvious to those skilled in the art that some modifications or improvements can be made to the present invention. Therefore, these modifications or improvements made without departing from the spirit of the present invention all belong to the scope of protection claimed by the present invention.

Claims

1. A multi-cavity scramjet combustion chamber, characterized in that: include: An isolation section (1), a first concave cavity group (2), a second concave cavity group (3), a third concave cavity group (4), an injection module and a combustion chamber expansion section (5); Each cavity group has two cavities arranged opposite to each other up and down; The isolation section (1) is a first square tube connected to the first concave cavity group (2); the first concave cavity group (2) is connected to the second concave cavity group (3) via a second square tube (8); the second concave cavity group (3) is connected to the third concave cavity group (4) via a third square tube (9); the third concave cavity group (4) is connected to the combustion chamber expansion section (5) at its rear; The injection unit comprises: a front injection device (7) and a rear injection device (6); the front injection device (7) is disposed on the upper and lower walls of the second square tube in a relative manner; and the rear injection device (6) is disposed on the upper and lower walls of the third square tube in a relative manner.

2. The multi-cavity scramjet engine combustion chamber according to claim 1, characterized in that: The injection unit further comprises an auxiliary injection device (10); Auxiliary injection devices (10) are respectively arranged opposite to each other on the two concave cavities of the second concave cavity group (3).

3. The multi-cavity scramjet engine combustion chamber according to claim 1, characterized in that: The isolation section (1) is a first square tube with an expansion angle, and its upper wall surface is an upwardly inclined surface, and the upward inclination angle does not exceed 1°.

4. The multi-cavity scramjet engine combustion chamber according to claim 1, characterized in that: The upper wall surfaces of the second square tube and the third square tube are upwardly inclined surfaces, and the upwardly inclined angle does not exceed 1°.

5. The multi-cavity scramjet engine combustion chamber according to claim 1, characterized in that: The upper wall surface of the combustion chamber expansion section (5) is inclined upward, and the upward inclination angle does not exceed 6°.

6. The multi-cavity scramjet engine combustion chamber according to any one of claims 1 to 5, characterized in that: The rear wall surfaces of two 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 inclined backwards at 45 degrees.

7. The multi-cavity scramjet engine combustion chamber according to any one of claims 1 to 5, characterized in that: The lower wall surfaces of the isolation section (1), the first concave cavity group (2), the second concave cavity group (3), the third concave cavity group (4) and the combustion chamber expansion section (5) are horizontal.

8. The multi-cavity scramjet engine combustion chamber according to any one of claims 1 to 5, characterized in that: The front injection device (7) and the rear injection device (6) are both provided with a plurality of injection holes distributed at equal intervals in the transverse direction; The diameter of the injection hole is 0.3 mm to 0.6 mm; The distance between the central axis of the front injection device (7) and the front wall surface of the second cavity group (3) is 55 mm to 75 mm; The distance between the central axis of the rear injection device (6) and the front wall surface of the third cavity group (4) is 55 mm to 75 mm.

9. A control method for a multi-cavity scramjet combustion chamber, characterized in that: The engine combustion chamber is a multi-cavity scramjet engine combustion chamber as claimed in any one of claims 1 to 8, and is arranged inside the engine of the aircraft; When the aircraft using the scramjet engine combustion chamber is operating in a low-altitude climbing phase, the valve connected to the front injection device (7) is closed, the valve connected to the rear injection device (6) is opened, and the opening of the valve connected to the rear injection device (6) is adjusted according to the aircraft control requirements to increase the oil volume; When 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, and 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 oil volume, and the aircraft completely enters the high-altitude cruising stage; then it continues to operate until the aircraft completes the voyage.

10. The control method of a multi-cavity scramjet combustion chamber according to claim 9, characterized in that: When auxiliary injection devices (10) are respectively arranged on the two concave cavities of the second concave cavity group (3) in opposite directions: During the low-altitude climbing stage, when a large-amplitude pressure oscillation occurs, the valve connected to the auxiliary injection device (10) is opened, and kerosene with a set equivalence ratio is injected through the auxiliary injection device (10) to suppress the large-amplitude pressure oscillation.

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