Combustion chamber structure of rocket-based rotary detonation engine and application of combustion chamber structure

Abstract

The invention discloses a rocket-based rotary detonation engine combustion chamber structure and application thereof, and belongs to the technical field of rocket engine research. The combustion chamber structure comprises an isolation section, an equal-direction oil injection section and an expansion-concave cavity section which are connected in sequence. A central rocket is arranged in the isolation section and the equal direct oil injection section; and four groups of first ring-jet injection points are arranged on the equal straight oil injection section, two groups of second ring-jet injection points are arranged on the central rocket, and two groups of third ring-jet injection points are arranged on the expansion-concave cavity section. Oil is injected into the expansion-concave cavity section firstly, a pre-detonation pipe is adopted for igniting an oil-gas mixture, rotating detonation waves are formed, the built-in center rocket can provide thrust for the aircraft at the ground zero speed and form an injection effect for air suction, extra thrust is provided when the cruising thrust of the aircraft is insufficient, and when the rotating detonation combustion flameout boundary exists, the detonation effect is achieved. A flame stabilizing mechanism is provided for rotary detonation combustion, and stable work of wide-speed-range and wide-height-range rotary detonation combustion is achieved.

Description

Technical Field

[0001] The present invention relates to the technical field of rocket engine research, and in particular to a rocket-based rotating detonation engine combustion chamber structure and application thereof. Background Art

[0002] Rotating detonation combustion has the characteristics of short combustion distance and high combustion efficiency. Compared with traditional slow-burning ramjet engines, ramjet engines that use rotating detonation combustion have the characteristics of short axial length, light structural mass and high combustion efficiency. However, traditional rotating detonation ramjet engines cannot start at zero speed, have a narrow operating Mach number, a limited flight altitude range, and are difficult to ignite under low dynamic pressure.

[0003] The rocket-based combined cycle engine (RBCC) has the characteristics of zero-speed start, wide operating range, and orbital capability. However, its overall size and mass are large, and its combustion efficiency is low in ramjet mode and the combustion distance is long.

[0004] In view of the above-mentioned characteristics of traditional rotating detonation and rocket-based combined cycle engines (RBCC), the combination of rotating detonation and rocket-based combined cycle engines can obtain greater performance gains. Therefore, combining traditional rotating detonation with rocket-based combined cycle engines has become a research hotspot. However, how to simultaneously solve the problems of rotating detonation ramjet engines being unable to start at zero speed, having a narrow operating Mach number, limited flight altitude range, ignition under low dynamic pressure, and low combustion efficiency and long combustion distance in the RBCC ramjet mode requires further in-depth research. Summary of the invention

[0005] In view of the above-mentioned problems, the present invention aims to provide a rocket-based rotating detonation engine combustion chamber structure and its application, by first spraying oil inside the expansion-concave cavity section, and igniting the oil-gas mixture in the combustion chamber by the pre-detonation tube to form a rotating detonation wave, the built-in central rocket can provide thrust for the aircraft at zero speed on the ground and form an induced effect to inhale air, provide additional thrust when the aircraft's cruising thrust is insufficient, and provide a flame stabilization mechanism for the rotating detonation combustion at the flameout boundary of the rotating detonation combustion, thereby achieving stable operation of rotating detonation combustion in a wide speed range and a wide altitude range.

[0006] In order to achieve the above object, the technical solution adopted by the present invention is as follows: A rocket-based rotating detonation engine combustion chamber structure, comprising an isolation section, an equal direct injection section and an expansion-recess section connected in sequence; The isolation section and the direct injection section are provided with a central rocket; and the central rocket is connected to the isolation section through a support plate; The equal direct injection oil section is provided with four groups of first ring-jet injection points, the central rocket is provided with two groups of second ring-jet injection points, the expansion-concave cavity section is provided with two groups of third ring-jet injection points, and the support plate is provided with a group of wall injection.

[0007] Furthermore, the central rocket is divided into a cylindrical matching area, a support plate installation area, a fuel injection area and a baffle installation area from front to back; The two groups of second ring-jet injection points are located in the injection area, and the fuel flow channels of the two groups of second ring-jet injection points are inner flow channels.

[0008] Furthermore, a transition inner column matching with the cylindrical matching area is provided in the isolation section.

[0009] Furthermore, the direct injection oil section adopts a direct injection oil section-windowing method, is provided with four groups of first ring injection points with different injection point parameters, and is provided with an optical glass window.

[0010] Furthermore, the direct injection section is distributed in six phases according to functional areas, the 0° phase area and the 180° phase area are the fuel inlet of the injection group, the 90° phase area and the -90° phase area are the first window area, the 45° phase area is the first static pressure sensor installation area, and the -45° phase area is the first dynamic pressure sensor installation area.

[0011] Furthermore, the expansion-concave cavity section is divided into a second expansion section and a concave cavity section according to the airflow direction, and the second expansion section is provided with two groups of the third annular injection points; The concave cavity section is provided with multiple sets of sensor interfaces and optical glass windows.

[0012] Furthermore, the phase of the expansion-concave cavity section is divided from the perspective of the airflow direction. The 0° phase zone and the 180° phase zone are the second dynamic pressure sensor installation area, the 45° phase zone is the ignition area, the 90° phase zone and the -90° phase zone are the second window area, the 135° phase zone is the photoelectric multiplier installation area, the -150° phase zone is the water-cooled sensor installation area, the -135° phase zone is the second static pressure sensor installation area, and the -45° phase zone is the heat flow sensor installation area.

[0013] Furthermore, the present invention also provides a rocket-based rotating detonation engine test simulation device based on an adjustable geometric throat, comprising the rocket-based rotating detonation engine combustion chamber structure as described above; It also includes a first expansion section and a device throat connected in sequence at the front end of the combustion chamber structure, and a variable circular-to-square section and a variable geometry throat connected in sequence at the rear end of the combustion chamber structure.

[0014] Furthermore, the present invention also provides an operation method of the rocket-based rotating detonation engine test simulation device based on the adjustable geometric throat as described above, comprising the following steps: S1: After the direct injection section and the expansion-concave cavity section are inhaled, the third group of first ring injection points of the direct injection section and the two groups of third ring injection points of the expansion-concave cavity section start to inject fuel, and the pre-detonation tube ignites the oil-gas mixture in the combustion chamber to form a rotating detonation wave; S2: When the Mach number is Ma0-Ma2.0, the center rocket opens to provide thrust and forms an ejection effect to inhale air. The center rocket gas, intake air and kerosene perform rotating detonation combustion together. When the operation reaches the Ma1.5-Ma2.0 stage, the center rocket is gradually closed, and pure rotating detonation combustion is performed in the flow channel. In this mode, the variable geometry throat area is the largest. S3: When the Mach number is Ma2.0-Ma3.0, the variable geometry throat area is reduced, and the injection point is selected at the second ring injection point on the central rocket; when it approaches the flame stabilization equivalence ratio limit, the support plate injection is turned on; S4: When the Mach number is Ma4.0-Ma5.0, the variable geometry throat area continues to decrease, and the injection point is located at the four groups of first ring injection points on the equal direct injection section; S5: When the Mach number is Ma6.0, the variable geometry throat area continues to decrease, and the injection point is located at the four groups of first ring injection points and the support plate on the equal direct injection section; S6: When the altitude increases or approaches the boundary of rotating detonation flameout, in order to maintain the stability of the rotating detonation wave in the combustion chamber, the fuel injection is moved forward and the center rocket is opened as appropriate.

[0015] Furthermore, the present invention also includes the application of the rocket-based rotating detonation engine combustion chamber structure as described above in a rocket-based rotating detonation engine.

[0016] The beneficial effects of the present invention are as follows: compared with the prior art, the present invention has the following improvements: 1. The present invention proposes a rocket-based rotating detonation engine combustion chamber structure, which combines the center of gravity rocket and the rotating detonation. The built-in rocket can provide thrust for the aircraft at zero speed on the ground and form an induced effect to inhale air, provide additional thrust when the aircraft's cruising thrust is insufficient, and provide a flame stabilization mechanism for the rotating detonation combustion at the boundary of the rotating detonation combustion extinction.

[0017] 2. A total of 9 groups of injection points are arranged in the combustion chamber structure of the rocket-based rotating detonation engine in the present invention, which respectively correspond to the relevant injection requirements under Ma0-Ma7.0 flight conditions. The stable operation of rotating detonation combustion in a wide speed range and a wide height range is achieved by controlling the injection area, the combustion area and the throat area.

[0018] 3. The variable geometry throat in the present invention is used to adjust the combustion pressure in the flow channel of the combustion chamber, and can provide pressure guarantee for the wide range operation of the combustion chamber.

[0019] 4. The engine of the present invention adopts rocket jet / pre-detonator coordinated detonation technology to solve the problem of difficulty in ignition of rotating detonation with low dynamic pressure. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] Figure 1 This is a schematic diagram of the external structure of the combustion chamber structure of the rocket-based rotating detonation engine in Example 1 of the present invention.

[0021] Figure 2 This is a cross-sectional view of the internal structure of the combustion chamber structure of the rocket-based rotating detonation engine in Example 1 of the present invention.

[0022] Figure 3 This is a front view of the internal structure of the combustion chamber structure of the rocket-based rotating detonation engine in Example 1 of the present invention.

[0023] Figure 4 It is a schematic diagram of the external structure of the central rocket of the present invention.

[0024] Figure 5 It is a cross-sectional view of the internal structure of the center rocket of the present invention.

[0025] Figure 6 It is a schematic diagram of the internal structure of the rocket support plate of the present invention.

[0026] Figure 7 It is a schematic diagram of phase partitioning of the front cross-sectional view of the central rocket of the present invention.

[0027] Figure 8 It is a left sectional view of the central rocket of the present invention.

[0028] Fig. 9 It is a schematic diagram of the external structure of the direct injection oil section of the present invention.

[0029] Fig.10 It is a cross-sectional view of the internal structure of the direct injection oil section of the present invention.

[0030] Fig.11 It is a schematic diagram of the phase division of the direct injection section in the present invention.

[0031] Fig.12 It is a schematic diagram of the external structure of the expansion-concave cavity section of the present invention.

[0032] Fig.13 It is a cross-sectional view of the internal structure of the expansion-concave cavity section of the present invention.

[0033] Fig.14 It is a schematic diagram of the phase partitioning of the expansion-concave cavity section side section of the present invention.

[0034] Fig.15 It is a schematic diagram of the external structure of the 0° phase zone of the expansion-concave cavity section of the present invention.

[0035] Fig.16 It is a schematic diagram of the external structure of the 45° phase zone and the 90° phase zone of the expansion-concave cavity section of the present invention.

[0036] Fig.17 This is a schematic diagram of the external structure of a rocket-based rotating detonation engine test simulation device based on an adjustable geometric throat in Example 2 of the present invention.

[0037] Fig.18 This is a cross-sectional view of the internal structure of a rocket-based rotating detonation engine test simulation device based on an adjustable geometric throat in Example 2 of the present invention.

[0038] Fig.19 It is a cross-sectional view of the internal structure of the variable circular rotating square section in the second embodiment of the present invention.

[0039] Fig. 20 It is a cross-sectional view of the internal structure of the variable geometry throat in the second embodiment of the present invention.

[0040] Among them: 1. Equipment throat; 2. First expansion section; 3. Isolation section; 301. Transfer inner column; 4. Direct injection section; 401. Inlet of injection group; 4011. First ring injection injection point; 402. First window area; 403. First static pressure sensor installation area; 404. First dynamic pressure sensor installation area; 5. Expansion-concave cavity section; 501. Second expansion section; 5011. Third ring injection injection point; 502. Concave cavity section; 5021. Second dynamic pressure sensor installation area; 50211. Second dynamic pressure sensor installation hole; 50212. 80-hole injection group oil inlet hole; 5022. Detonation area; 50221. Rocket jet detonation interface; 50222. Pre-detonation tube detonation interface; 5023 , the second window area; 50231, 60-hole injection group oil inlet hole; 5024, photomultiplier installation area; 5025, water-cooled sensor installation area; 5026, second static pressure sensor installation area; 5027, heat flux sensor installation area; 6, variable circular square section; 7, variable geometry throat; 701, throat body; 702, throat block; 7021, horizontal plate; 7022, arc-shaped vertical plate; 7023, fan-shaped plate; 8, center rocket; 801, cylindrical matching area; 802, support plate installation area; 8021, rocket support plate; 8022, support support plate; 8023, wall injection; 803, injection area; 8031, second ring injection point; 804, baffle installation area; 9, optical glass window. DETAILED DESCRIPTION

[0041] In order to enable those skilled in the art to better understand the technical solution of the present invention, the technical solution of the present invention is further described below in conjunction with the accompanying drawings and embodiments. Embodiment 1

[0042] Embodiment 1 provides a rocket-based rotating detonation engine combustion chamber structure, as shown in the attached Figure 1-16 As shown, the combustion chamber structure includes an isolation section 3, an equal direct injection section 4 and an expansion-concave cavity section 5 which are connected in sequence, and the isolation section 3 and the equal direct injection section 4 have a central rocket 8 built in; and the central rocket 8 is connected to the isolation section 3 through a support plate.

[0043] The equal direct injection oil section 4 is provided with four groups of first annular injection points 4011, the central rocket 8 is provided with two groups of second annular injection points 8031, the expansion-concave cavity section 5 is provided with two groups of third annular injection points 5011, and the support plate is provided with a group of wall injections 8023.

[0044] Specifically, a transition inner column 301 is provided in the isolation segment 3, and the inner surface of the transition inner column 301 is cylindrically matched with the central rocket 8 for installing the central rocket 8.

[0045] Furthermore, the central rocket 8 adopts a central rocket with a combustion chamber structure, which is divided into four functional areas from front to back, namely, a cylindrical matching area 801, a support plate installation area 802, an injection area 803 and a baffle installation area 804. The cylindrical matching area 801 is cylindrically matched with the inner profile of the transfer inner column 301, and a negative tolerance is adopted; the central rocket 8 is connected to the isolation section 3 through a support plate in the support plate installation area 802, and the support plate specifically includes a rocket support plate 8021 and a support support plate 8022. The support plate installation area 802 is provided with a rocket support plate installation groove and a support support plate installation groove, which are respectively used to correspondingly install the rocket support plate 8021 and the support support plate 8022, as shown in the attached Figure 8 A set of wall injections 8023 are provided on the rocket support plate 8021 and the support support plate 8022, as shown in the attached Figure 6 As shown (attached Figure 6 The internal structure of the rocket support plate 8021 is shown in the figure. The support plate 8022 has the same internal structure as the rocket support plate 8021). The rocket support plate 8021 has the functions of self-injection and rocket gas flow channel; the support plate 8022 has the functions of supplying fuel to the central rocket 8, self-injection and supporting the central rocket 8. The support plate series are all made of 304 stainless steel.

[0046] Two groups of second ring injection points 8031 ​​are located in the injection area 803, one of which is a 40×Φ0.3 injection group and the other is an 80×Φ0.3 injection group; the baffle installation area 804 is used to install the central rocket graphite liner baffle and the central rocket sealing cover plate. The installation threaded hole size is 6×M6×1.0, and the thread depth is at least 8mm. After the thread processing is completed, the chip cleaning should be carried out, and the installation test should be carried out after the chip cleaning. In order to ensure the smooth installation with the central rocket graphite liner, the inner surface of the central rocket 8 should adopt a positive tolerance.

[0047] The direction from the cylindrical matching area 801 to the baffle installation area 804 is determined as the phase perspective. Under this perspective, the central rocket 8 is divided into three phase zones, namely, the 0° phase zone, the 120° phase zone, and the -120° phase zone. The 0° phase zone is the installation area of ​​the rocket support plate 8021, and the 120° phase zone and the -120° phase zone are the installation areas of the support support plate 8022. Correspondingly, the isolation section 3 is divided into phases according to the perspective of the flow direction of the airflow. The 0° phase zone is the installation area of ​​the rocket support plate 8021, the 45° phase zone is the installation area of ​​the static pressure sensor, the 120° phase zone is the installation area of ​​the support support plate 8022, the -45° phase zone is the installation area of ​​the dynamic pressure sensor, and the -120° phase zone is also the installation area of ​​the support support plate 8022. The 0° phase / 120° phase / -120° phase geometric structures are exactly the same, and the processing technology requirements are the same. The specific setting parameters of the isolation section 3 are not repeated here.

[0048] The 0° phase area of ​​the central rocket 8 is the installation area of ​​the rocket support plate 8021. The outer contour of the rocket support plate installation groove adopts positive tolerance. The positive tolerance range should consider the positioning dimension tolerance band on the basis of being consistent with the support plate installation groove of the isolation section 3 to ensure smooth installation.

[0049] The 120° phase zone and the -120° phase zone of the central rocket 8 are both installation areas for the support plate 8022. The geometric dimensions of the support plate installation grooves at the two phases are consistent, and the tolerance requirements are consistent with the relevant requirements for the rocket support plate installation grooves.

[0050] The outer diameter of the oil inlet hole of the two groups of second ring-jet injection points 8031 ​​in the injection area 803 is changed from Φ12 to Φ6, of which the Φ12 part should adopt a positive tolerance, chamfer C1.5, and the bottom surface of the support plate mounting groove should be used as the positioning reference. The fuel flow channels of the two groups of second ring-jet injection points 8031 ​​are internal flow channels, and the fuel flow channel processing can be completed by the segmented processing-welding method. Before welding, special attention should be paid to the coaxiality of the fuel flow channel to ensure that the fuel injection pressure is not affected by the sudden change in area caused by welding misalignment. After welding, it is required to clean the weld to ensure that the outer surface of the center rocket 7 is smooth and flat, and pay special attention to check whether the spray hole of the second ring-jet injection point 8031 ​​is blocked due to welding.

[0051] The flange surface graphite liner positioning groove of the baffle mounting area 804 should adopt positive tolerance. The center rocket 8 is machined from stainless steel.

[0052] Furthermore, the direct injection section 4 in the present invention adopts a direct injection section-windowing method and is made of 304 stainless steel.

[0053] The medium direct injection oil section 4 of the present invention is provided with four groups of first ring injection points 4011, and the injection point parameters along the airflow direction are 80×Φ0.3, 60×Φ0.3, 40×Φ0.3, and 80×Φ0.3 respectively.

[0054] According to the airflow direction, the direct injection section 4 is distributed in six phases according to the functional areas. The 0° phase area and the 180° phase area are the injection group oil inlet 401, the 90° phase area and the -90° phase area are the first window area 402, the 45° phase area is the first static pressure sensor installation area 403, and the -45° phase area is the first dynamic pressure sensor installation area 404. Fig.10 shown.

[0055] The 0° phase zone of the equal direct injection section 4 is the injection group oil inlet 401, and the oil inlet hole size of the injection group oil inlet 401 is 4×Φ6-. Drill through to the outer diameter of the liquid collecting cavity, weld the DN6 injection rod after the oil inlet hole is processed, and the 180° phase zone is symmetrically distributed with the 0° phase zone, which is also the injection group oil inlet 401, and the relevant size is the same. Along the axial direction of the equal direct injection section 4, four groups of first ring injection injection points 4011 are arranged in parallel.

[0056] The -45° phase zone of the direct injection section 4 is the first dynamic pressure sensor installation area 404. The size of the dynamic pressure sensor installation hole is 5×M10×1.0-tap. The bottom surface of the installation area is required to be flattened. The flattened plane is 120 mm from the center axis and the plane width is 44 mm.

[0057] The 45° phase area of ​​the direct injection section 4 is the first static pressure sensor installation area 403, and the pressure-inducing hole has a size of 11×Φ2-drilled. After processing, the static pressure sensor extension rod is welded to the pressure-inducing hole.

[0058] The 90° phase zone is the first window zone 402. The threaded hole size of the flange of the window zone cover plate is 12×M6×1.0-drill 15 tap 12, the positioning hole size is Φ5 drill 5-positive tolerance, the sealing groove width is 4.1-positive tolerance, and the groove depth is 3.6-negative tolerance. The -90° phase zone has a similar structure to the 90° phase zone, with the same threaded hole and sealing groove parameters, the same positioning hole molding size, but the positioning holes of the two phase zones are asymmetric. Both first window zones 402 are installed with optical glass windows 9.

[0059] The flange hole size at both ends of the direct injection section 4 is 12×Φ11-through hole, the left end flange sealing groove size is groove width 6.3mm-positive tolerance, groove depth 4.1mm-negative tolerance, stopper outer diameter negative tolerance. The right end flange stopper outer diameter positive tolerance.

[0060] Furthermore, the expansion-concave section 5 is divided into a second expansion section 501 and a concave section 502 according to the airflow direction, and two groups of third annular injection points 5011 are located on the second expansion section 501, one group of third annular injection points 5011 is a 60×Φ0.3 injection group, and the other group of third annular injection points 5011 is an 80×Φ0.3 injection group. The concave section 502 is provided with multiple groups of sensor interfaces and optical glass windows 9.

[0061] The phase of the expansion-concave cavity section 5 is divided from the perspective of the flow direction of the airflow. The 0° phase area and the 180° phase area are the second dynamic pressure sensor installation area 5021, the 45° phase area is the detonation area 5022, the 90° phase area and the -90° phase area are the second window area 5023, the 135° phase area is the photomultiplier installation area 5024, the -150° phase area is the water cooling sensor installation area 5025, the -135° phase area is the second static pressure sensor installation area 5026, and the -45° phase area is the heat flow sensor installation area 5027. The two second window areas 5023 are also installed with optical glass windows 9, as shown in the attached Fig.14 shown.

[0062] More specifically, the 0° phase zone is the second dynamic pressure sensor installation area 5021, which requires the bottom surface to be flattened, with the flattened plane 143mm from the center axis and a plane width of 48mm. The second dynamic pressure sensor installation hole 50211 is a 4×M10×1.0 fine-thread through hole, which requires the use of M10×1.0×20 fine-thread screws for installation testing after processing to ensure that the fine-thread has no stripping or broken threads. The bottom of the second dynamic pressure sensor installation area 5021 is the 80-hole injection group oil inlet hole 50212, with a hole diameter of Φ6, which is drilled to the outer diameter of the liquid collecting cavity, as shown in the attached Fig.15 shown.

[0063] The 45° phase zone is the detonation zone 5022. The detonation zone 5022 is provided with a rocket jet detonation interface 50221 and a pre-detonation tube detonation interface 50222. The pre-detonation tube detonation interface 50222 is welded with a DN10 male head. The flange hole thread size of the rocket jet detonation interface 50221 is 4×M6×1.0. A liner positioning groove is provided in the rocket jet detonation interface 50221. The liner positioning groove is a stepped structure of Φ35 to Φ20. Positive tolerances are adopted, and a graphite liner is installed inside the liner positioning groove.

[0064] The size of the detonation wave flow channel in the pre-detonation tube detonation interface 50222 is Φ10-drilled. A welding step is reserved on the outside, and the welding step is a Φ17 hole concentric with the detonation wave flow channel. After the pre-detonation tube detonation interface 50222 is processed, a DN10 male connector is welded. It is required to pay special attention to the reinforcement of the welding root, and pay attention to cleaning the detonation wave flow channel after welding.

[0065] The 90° phase zone and the -90° phase zone are both the second window zone 5023. The left side of the second window zone 5023 of the 90° phase zone is the 60-hole injection group oil inlet hole 50231, the size of which is Φ6-drilled to the outer diameter of the liquid collecting cavity. The bottom surface of the 60-hole injection group oil inlet hole 50231 is flattened, and the flattened plane is 143mm away from the center axis. The plane width is 48mm. The thread size of the flange screw hole of the second window zone 5023 is 12×M6×1.0-drill 15 tap 12, and the size of the positioning pin hole is Φ5 drill 5, and positive tolerance should be used.

[0066] The 135° phase zone is the photomultiplier installation area 5024. The bottom surface of the installation area is flattened. The flattened plane is 140mm from the center axis and the plane width is 75.5mm. The thread size of the flange hole in the installation area is 24×M6×1.0-drill 6 tap 5, the step in the installation area is Φ20.2-deep 6 turn Φ6-drill through, the Φ20.2 part adopts negative tolerance, and the finished size shall not be less than Φ20.0. The Φ6 part adopts positive tolerance. There are five kinds of workpieces in the photomultiplier optical path series, namely "photomultiplier glass lens_large", "photomultiplier glass lens_small", "photomultiplier glass mirror graphite gasket_large", "photomultiplier glass mirror graphite gasket_small" and "photomultiplier M10 gasket".

[0067] The 180° phase zone is also the second dynamic pressure sensor installation area 5021, and the relevant requirements are consistent with those of the 0° phase zone.

[0068] The -150° phase zone is the water-cooled sensor installation area 5025, which is the installation area of ​​the water-cooled sensor. The sensor installation hole size is 1 / 2-20 UNF-2B American thread tapping, the countersunk outer diameter is 14.28mm, the countersunk is 1.27mm, and the positive tolerance is 1.27mm. In order to match the thread size of the water-cooled sensor adapter, the installation hole plane is required to be raised and flattened. The flattened plane is 146.8mm from the center axis and the plane width is 30mm.

[0069] The -135° phase zone is the second static pressure sensor installation zone 5026, which has 9 Φ2 optical through holes as pressure-inducing holes, including 2 holes in the expansion section 501 and 7 holes in the cavity section 502. After the expansion-cavity section 5 body is processed, each pressure-inducing hole is welded with a static pressure sensor extension rod.

[0070] The -45° phase zone is the heat flux sensor installation area 5027. The bottom surface of the heat flux sensor installation area is flattened. The flattened plane is 143mm from the center axis and the plane width is 48mm. The thread size of the heat flux sensor installation flange hole is 4×M3×0.5-drill 10 tap 5, and the installation hole is Φ12-drilled through. Positive tolerance should be used. The plug positioning groove adopts positive tolerance.

[0071] The working principle of the rocket-based rotating detonation engine combustion chamber structure in this embodiment is: After the combustion chamber is intaked with air, the third group of first ring injection points 4011 of the direct injection section 4 and the two groups of third ring injection points 5011 of the expansion-concave cavity section 5 start to spray oil. After a certain period of injection, a detonation wave is injected into the rocket jet detonation interface 50221 and the pre-detonation tube detonation interface 50222 to ignite the oil-gas mixture in the combustion chamber and form a rotating detonation wave.

[0072] When working at extremely low Mach numbers (Ma0-Ma2.0) in the subsonic and supersonic sections, the center rocket 8 opens to provide thrust and forms an induced air intake effect. The combustion gas of the center rocket 8, the intake air and kerosene undergo rotational detonation combustion together. When working to the Ma1.5-Ma2.0 stage, the center rocket 8 is gradually closed, and pure rotational detonation combustion is carried out in the flow channel.

[0073] When working at low Mach numbers (Ma2.0-Ma3.0), the injection point selects the second ring injection point 8031 ​​on the central rocket 8, and the combustion area is located at the concave cavity section 502; when approaching the flame stabilization equivalent ratio limit, the rocket support plate 8021 and the support support plate 8022 are turned on to achieve flame stabilization over a larger range.

[0074] When working at medium Mach number (Ma4.0-Ma5.0), the injection point is located at the four groups of first ring injection points 4011 on the equal direct injection section 4, and the combustion area is located at the concave cavity section 502; When working at high Mach numbers (Ma6.0-Ma7.0), the injection points are located at the four groups of first ring injection points 4011 on the equal direct injection section 4 and the rocket support plate 7021 and the support support plate 7022, and the combustion area is located in the small-sized combustion area; When the altitude increases or approaches the boundary of rotating detonation flameout, in order to maintain the stability of the rotating detonation wave in the combustion chamber, the fuel injection is moved forward and the central rocket 8 can be opened as appropriate.

[0075] In summary, under Ma0 and Ma1.0 conditions, the flame stability equivalent ratio is 0.5-0.8 under the support of the central rocket 8 gas, the flame stability range under Ma2.0 conditions is 0.4-1.2, the flame stability range of Ma3.0 is 0.3-1.2, the flame stability range of Ma4.0 is 0.3-1.2, the flame stability range of Ma5.0 is 0.3-1.2, the flame stability range of Ma6.0 is 0.3-1.2, and the flame stability range of Ma7.0 is 0.3-1.2.

[0076] In addition, it should be noted that the combustion chamber structure in this embodiment can be further applied to rocket-based rotating detonation engines. Embodiment 2

[0077] Embodiment 2 applies the combustion chamber structure in embodiment 1 to a rocket-based rotating detonation engine test simulation device based on an adjustable geometric throat, and adjusts the combustion pressure in the combustion chamber flow channel through the adjustable geometric throat to provide pressure guarantee for the wide range operation of the combustion chamber.

[0078] Specifically, in the rocket-based rotating detonation engine test simulation device based on the adjustable geometry throat, the front end of the combustion chamber structure in the first embodiment is connected to the first expansion section 2 and the device throat 1 in sequence, and the rear end of the combustion chamber structure is connected to the variable geometry throat 7 through the variable circular rotation square section 6, thereby forming an overall structure in which the device throat 1, the first expansion section 2, the isolation section 3, the equal direct injection section 4, the expansion-concave cavity section 5, the variable circular rotation square section 6, and the variable geometry throat 7 are connected in sequence from front to back, and the central rocket 8 is built into the isolation section 3 and the equal direct injection section 4, as shown in the attached figure. Figure 17-20 shown.

[0079] Specifically, the equipment throat 1 and the first expansion section 2 may adopt the existing equipment throat and expansion section structures.

[0080] Furthermore, one end of the variable circular-to-square section 6 is a circular structure, and the other end is a square structure, which is used to connect the expansion-recessed cavity section 5 at the circular end and the variable geometry throat 7 at the square end.

[0081] The variable geometry throat 7 structure is as shown in the attached Fig.19 and 20As shown, it includes a throat body 701, which is a cavity structure surrounded by a shell, and its upper end, left inlet end and right outlet end are all open; a throat block 702 is movably provided in the throat body 701, and the throat block 702 includes a horizontal plate body 7021, an arc-shaped vertical plate 7022 and two fan-shaped plate bodies 7023; the two fan-shaped plate bodies 7023 are vertically arranged at the front and rear ends of the horizontal plate body 7021, and the direction is consistent with the direction of the horizontal plate body 7021, and the length is the same; the tips of the two fan-shaped plate bodies 7023 are located at the right end of the horizontal plate body 7021, and the arc-shaped vertical plate 7022 is vertically arranged at the left end of the horizontal plate body 7021, and the curvature of its left outer side wall is consistent with the curvature of the arc surface of the two fan-shaped plate bodies 7023. The throat block 702 is located at the upper opening of the throat body 701 , and the horizontal plate 7021 covers the upper opening of the throat body 701 , and the right end of the horizontal plate 7021 is hingedly connected to the inner wall of the throat body 701 .

[0082] In order to provide power to the throat block 702, a driving system is also provided on the horizontal plate 7021, which is used to drive the horizontal plate 7021 to rotate in a cavity away from or close to the throat body 701 with the hinged end as the fixed end; during the rotation process, the horizontal plate 7021 serves as the top plate of the throat body 701, and the arc-shaped vertical plate 7022 is attached to the left end of the throat body 701, and a cavity of variable size is formed between the horizontal plate 7021, the arc-shaped vertical plate 7022 and the throat body 701. It should be noted here that the driving system in the present invention can adopt any structure that can drive the horizontal plate 7021 to rotate and can realize remote control, which will not be elaborated in the present invention.

[0083] The working principle of the rocket-based rotating detonation engine test simulation device based on the adjustable geometric throat in this embodiment is: After the combustion chamber is intaked with air, the third group of first ring injection points 4011 of the direct injection section 4 and the two groups of third ring injection points 5011 of the expansion-concave cavity section 5 start to spray oil. After a certain period of injection, a detonation wave is injected into the rocket jet detonation interface 50221 and the pre-detonation tube detonation interface 50222 to ignite the oil-gas mixture in the combustion chamber and form a rotating detonation wave.

[0084] When working at extremely low Mach numbers (Ma0-Ma2.0) in the subsonic and supersonic sections, the center rocket 8 opens to provide thrust and forms an ejection effect to inhale air. The gas from the center rocket 8, the intake air, and kerosene undergo a rotating detonation combustion. When working at the Ma1.5-Ma2.0 stage, the center rocket 8 is gradually closed, and pure rotating detonation combustion is performed in the flow channel. In this mode, the variable geometry throat 7 has the largest area.

[0085] When working at low Mach numbers (Ma2.0-Ma3.0), the area of ​​the variable geometry throat 7 is reduced, the injection point is selected as the second ring injection point 8031 ​​on the central rocket 8, and the combustion area is located at the concave cavity section 502; when approaching the flame stabilization equivalence ratio limit, the rocket support plate 8021 and the support support plate 8022 are turned on for injection to achieve flame stabilization over a larger range.

[0086] When working at medium Mach number (Ma4.0-Ma5.0), the area of ​​variable geometry throat 7 is continuously reduced, the injection point is located at four groups of first ring injection points 4011 on the equal direct injection section 4, and the combustion area is located at the concave cavity section 502; When working at high Mach numbers (Ma6.0-Ma7.0), the area of ​​the variable geometry throat 7 continues to decrease, the injection points are located at the four groups of first ring injection points 4011 on the equal direct injection section 4 and the rocket support plate 7021 and the support support plate 7022, and the combustion area is located in the small-sized combustion area; When the altitude increases or approaches the boundary of rotating detonation flameout, in order to maintain the stability of the rotating detonation wave in the combustion chamber, the fuel injection is moved forward and the central rocket 8 can be opened as appropriate.

[0087] The above shows and describes the basic principles, main features and advantages of the present invention. It should be understood by those skilled in the art that the present invention is not limited to the above embodiments, and the above embodiments and descriptions are only for explaining the principles of the present invention. Without departing from the spirit and scope of the present invention, the present invention may have various changes and improvements, which fall within the scope of the present invention to be protected. The scope of protection of the present invention is defined by the attached claims and their equivalents.

Claims

1. A rocket-based rotating detonation engine combustion chamber structure, comprising an isolation section (3), an equal direct injection section (4) and an expansion-recess section (5) connected in sequence; characterized in that: The isolation section (3) and the direct injection section (4) are provided with a central rocket (8); and the central rocket (8) is connected to the isolation section (3) via a support plate; The equal direct injection oil section (4) is provided with four groups of first ring-jet injection points (4011), the central rocket (8) is provided with two groups of second ring-jet injection points (8031), the expansion-concave cavity section (5) is provided with two groups of third ring-jet injection points (5011), and the support plate is provided with a group of wall injections (8023).

2. A rocket-based rotating detonation engine combustion chamber structure according to claim 1, characterized in that: The central rocket (8) is divided into a cylindrical matching area (801), a support plate installation area (802), a fuel injection area (803) and a baffle installation area (804) from front to back; The two groups of second ring-jet injection points (8031) are located in the fuel injection area (803), and the fuel flow channels of the two groups of second ring-jet injection points (8031) are inner flow channels.

3. A rocket-based rotating detonation engine combustion chamber structure according to claim 2, characterized in that: A transition inner column (301) matching the column surface matching area (801) is provided in the isolation section (3).

4. The rocket-based rotating detonation engine combustion chamber structure according to claim 1, characterized in that: The iso-direct oil injection section (4) adopts an iso-direct oil injection section-windowing method, is provided with four groups of first ring injection points (4011) with different injection point parameters, and is provided with an optical glass window (9).

5. A rocket-based rotating detonation engine combustion chamber structure according to claim 4, characterized in that: The iso-direct injection section (4) is distributed in six phases according to functional areas, the 0° phase area and the 180° phase area are the injection group oil inlet (401), the 90° phase area and the -90° phase area are the first window area (402), the 45° phase area is the first static pressure sensor installation area (403), and the -45° phase area is the first dynamic pressure sensor installation area (404).

6. The rocket-based rotating detonation engine combustion chamber structure according to claim 1, characterized in that: The expansion-concave cavity section (5) is divided into a second expansion section (501) and a concave cavity section (502) according to the airflow direction, and the second expansion section (501) is provided with two groups of third annular injection points (5011); The concave cavity section (502) is provided with multiple sets of sensor interfaces and optical glass windows (9).

7. A rocket-based rotating detonation engine combustion chamber structure according to claim 6, characterized in that: The phases of the expansion-concave cavity section (5) are divided from the perspective of the flow direction of the airflow: the 0° phase zone and the 180° phase zone are the second dynamic pressure sensor installation area (5021), the 45° phase zone is the detonation area (5022), the 90° phase zone and the -90° phase zone are the second window area (5023), the 135° phase zone is the photomultiplier installation area (5024), the -150° phase zone is the water-cooled sensor installation area (5025), the -135° phase zone is the second static pressure sensor installation area (5026), and the -45° phase zone is the heat flow sensor installation area (5027).

8. Rocket-based rotating detonation engine test simulator based on adjustable geometry throat, characterized by: A rocket-based rotating detonation engine combustion chamber structure comprising any one of claims 1 to 7; It also includes a first expansion section (2) and a device throat (1) connected in sequence at the front end of the combustion chamber structure, and a variable circular-to-square section (6) and a variable geometric throat (7) connected in sequence at the rear end of the combustion chamber structure.

9. The method for operating the rocket-based rotating detonation engine test simulator based on adjustable geometry throat according to claim 8, characterized in that: The steps include: S1: After the direct injection section (4) and the expansion-concave cavity section (5) are inducted, the third group of first annular injection points (4011) of the direct injection section (4) and the two groups of third annular injection points (5011) of the expansion-concave cavity section (5) start to inject fuel, and the pre-detonation tube ignites the oil-gas mixture in the combustion chamber to form a rotating detonation wave; S2: When the Mach number is Ma0-Ma2.0, the center rocket (8) is opened to provide thrust and form an ejection effect to inhale air. The combustion gas of the center rocket (8) is combined with the intake air and kerosene to perform rotating detonation combustion. When the operation reaches the Ma1.5-Ma2.0 stage, the center rocket (8) is gradually closed, and pure rotating detonation combustion is performed in the flow channel. In this mode, the area of ​​the variable geometry throat (7) is the largest. S3: When the Mach number is Ma2.0-Ma3.0, the area of ​​the variable geometry throat (7) is reduced, and the injection point is selected at the second ring injection point (8031) on the central rocket (8); when approaching the flame stabilization equivalence ratio limit, the support plate injection is turned on; S4: When the Mach number is Ma4.0-Ma5.0, the area of ​​the variable geometry throat (7) continues to decrease, and the injection point is located at the four groups of first ring injection points (4011) on the equal direct injection section (4); S5: When the Mach number is Ma6.0, the area of ​​the variable geometry throat (7) continues to decrease, and the injection point is located at the four groups of first ring injection points (4011) and the support plate on the equal direct injection section (4); S6: When the altitude increases or approaches the boundary of rotating detonation flameout, in order to maintain the stability of the rotating detonation wave in the combustion chamber, the fuel injection is moved forward and the central rocket (8) is opened as appropriate.

10. Application of the rocket-based rotating detonation engine combustion chamber structure as described in any one of claims 1 to 7 in a rocket-based rotating detonation engine.

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