A rocket-based rotating detonation engine combustion chamber structure and its application

Abstract

The present invention discloses a rocket-based rotating detonation engine combustion chamber structure and its application, belonging to the field of rocket engine research technology. The combustion chamber structure includes an isolation section, an equal direct injection section, and an expansion-concave cavity section connected in sequence; the isolation section and the equal direct injection section are equipped with a central rocket; the equal direct injection section is provided with four groups of first annular injection injection points, the central rocket is provided with two groups of second annular injection injection points, and the expansion-concave cavity section is provided with two groups of third annular injection points. The present invention first injects oil into the expansion-concave cavity section and uses a pre-detonation tube to ignite the oil-gas mixture 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 intake, provide additional thrust when the aircraft's cruising thrust is insufficient, and provide a flame stabilization mechanism for the rotating detonation combustion at the extinction boundary of the rotating detonation combustion, thereby achieving stable operation of the rotating detonation combustion in a wide speed range and a wide altitude range.

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 using rotating detonation combustion have the characteristics of short axial length, light structural weight 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 more difficult to ignite under low dynamic pressure.

[0003] The rocket-based combined cycle (RBCC) engine boasts zero-speed start capability, a wide operating range, and the ability to reach orbit. However, it is large in size and mass, and has low combustion efficiency and a long combustion distance in ramjet mode.

[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 achieve greater performance gains. Therefore, combining traditional rotating detonation and 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, a limited flight altitude range, ignition under low dynamic pressure, and low combustion efficiency and long combustion distance under the RBCC ramjet mode requires further in-depth research. Summary of the Invention

[0005] In response to the above-mentioned problems, the present invention aims to provide a rocket-based rotating detonation engine combustion chamber structure and its application, which first sprays oil inside the expansion-concave cavity section and ignites 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 ejection 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, thereby realizing 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:

[0007] A rocket-based rotating detonation engine combustion chamber structure comprises an isolation section, an equal direct injection section and an expansion-recessed cavity section connected in sequence;

[0008] The isolation section and the direct injection section are equipped with a central rocket; and the central rocket is connected to the isolation section through a support plate;

[0009] 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.

[0010] 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;

[0011] The two groups of second ring-jet injection points are located in the fuel injection area, and the fuel flow channels of the two groups of second ring-jet injection points are inner flow channels.

[0012] Furthermore, a transition inner column matching the cylindrical matching area is provided in the isolation section.

[0013] Furthermore, the direct oil injection section adopts a direct oil injection section-windowing mode, is provided with four groups of first ring injection points with different injection point parameters, and is provided with an optical glass window.

[0014] Furthermore, the direct injection oil section is distributed in six phases according to functional areas, the 0° phase area and the 180° phase area are the oil 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.

[0015] 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;

[0016] The concave cavity section is provided with multiple sets of sensor interfaces and optical glass windows.

[0017] Furthermore, the phase of the expansion-concave cavity section is divided from the perspective of the airflow direction. The 0° phase area and the 180° phase area are the second dynamic pressure sensor installation area, the 45° phase area is the ignition area, the 90° phase area and the -90° phase area are the second window area, the 135° phase area is the photoelectric multiplier installation area, the -150° phase area is the water-cooled sensor installation area, the -135° phase area is the second static pressure sensor installation area, and the -45° phase area is the heat flow sensor installation area.

[0018] 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;

[0019] 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.

[0020] Furthermore, the present invention also provides a method for operating the aforementioned rocket-based rotating detonation engine test simulator based on the adjustable geometry throat, comprising the following steps:

[0021] S1: After the direct injection section and the expansion-concave cavity section are inducted, the third group of first annular injection points of the direct injection section and the two groups of third annular injection points of the expansion-concave cavity section begin to inject fuel, and the pre-detonator ignites the oil-air mixture in the combustion chamber, forming a rotating detonation wave;

[0022] S2: At Mach numbers between Ma0 and Ma2.0, the center rocket opens to provide thrust and creates an ejection effect to inhale air. The center rocket gas, intake air, and kerosene undergo rotating detonation combustion. When the engine reaches Ma1.5 to Ma2.0, the center rocket gradually closes, and pure rotating detonation combustion occurs in the flow channel. In this mode, the variable geometry throat area is maximized.

[0023] 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 approaching the flame stabilization equivalence ratio limit, the strut injection is turned on;

[0024] 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;

[0025] S5: When the Mach number is Ma6.0, the variable geometry throat area continues to decrease, and the injection points are located at the four groups of first ring injection points and support plates on the equal direct injection section;

[0026] 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.

[0027] 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.

[0028] The beneficial effects of the present invention are as follows: compared with the prior art, the improvement of the present invention is that:

[0029] 1. The present invention proposes a rocket-based rotating detonation engine combustion chamber structure, which combines a central rocket and a rotating detonation. The built-in rocket can provide thrust for the aircraft at zero speed on the ground and form an ejection 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 when the rotating detonation combustion is at the extinction boundary.

[0030] 2. The rocket-based rotating detonation engine combustion chamber structure of the present invention is provided with 9 groups of injection points, which respectively correspond to the relevant injection requirements under Ma0-Ma7.0 flight conditions. By controlling the injection area, combustion area and throat area, the stable operation of rotating detonation combustion in a wide speed range and a wide height range is achieved.

[0031] 3. The geometrically variable throat of 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.

[0032] 4. The engine of the present invention adopts rocket jet / pre-blaster coordinated detonation technology to solve the problem of difficulty in ignition of rotating detonation with low dynamic pressure. BRIEF DESCRIPTION OF THE DRAWINGS

[0033] 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.

[0034] 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.

[0035] Figure 3 This is a front view of the internal structure of the combustion chamber of the rocket-based rotating detonation engine in Example 1 of the present invention.

[0036] Figure 4 This is a schematic diagram of the external structure of the central rocket of the present invention.

[0037] Figure 5 It is a cross-sectional view of the internal structure of the central rocket of the present invention.

[0038] Figure 6 It is a schematic diagram of the internal structure of the rocket support plate of the present invention.

[0039] Figure 7 This is a schematic diagram of the phase partitioning of the front cross-sectional view of the central rocket of the present invention.

[0040] Figure 8 It is a left sectional view of the central rocket of the present invention.

[0041] Figure 9 This is a schematic diagram of the external structure of the direct injection oil section of the present invention.

[0042] Figure 10 This is a cross-sectional view of the internal structure of the direct injection oil section of the present invention.

[0043] Figure 11 It is a schematic diagram of the phase division of the direct injection oil section in the side section of the present invention.

[0044] Figure 12 This is a schematic diagram of the external structure of the expansion-concave cavity section of the present invention.

[0045] Figure 13 This is a cross-sectional view of the internal structure of the expansion-concave cavity section of the present invention.

[0046] Figure 14 It is a schematic diagram of the phase partitioning of the expansion-concave cavity section in the side section of the present invention.

[0047] Figure 15 This is a schematic diagram of the external structure of the 0° phase zone of the expansion-concave cavity section of the present invention.

[0048] Figure 16 This 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.

[0049] Figure 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.

[0050] Figure 18 This is a cross-sectional view of the internal structure of the rocket-based rotating detonation engine test simulation device based on the adjustable geometric throat in Example 2 of the present invention.

[0051] Figure 19 This is a cross-sectional view of the internal structure of the variable circular rotating square section in Example 2 of the present invention.

[0052] Figure 20 This is a cross-sectional view of the internal structure of the variable geometry throat in Example 2 of the present invention.

[0053] Among them: 1. Equipment throat; 2. First expansion section; 3. Isolation section; 301. Transfer inner column; 4. Direct injection section; 401. Injection group oil inlet; 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-cavity section; 501. Second expansion section; 5011. Third ring injection injection point; 502. Cavity section; 5021. Second dynamic pressure sensor installation area; 50211. Second dynamic pressure sensor installation hole; 50212. 80-hole injection group oil inlet; 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, central 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

[0054] 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.

[0055] Example 1:

[0056] 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 oil section 4 and an expansion-concave cavity section 5 connected in sequence, and the isolation section 3 and the equal direct injection oil section 4 are equipped with a central rocket 8; and the central rocket 8 is connected to the isolation section 3 through a support plate.

[0057] 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 injection 8023.

[0058] Specifically, a transition inner column 301 is provided in the isolation section 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.

[0059] 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 the cylindrical matching area 801, the support plate installation area 802, the injection area 803 and the baffle installation area 804. The cylindrical matching area 801 is cylindrically matched with the inner surface 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. 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 used to respectively install the rocket support plate 8021 and the support support plate 8022, as shown in the attached figure. Figure 8 As shown. A set of wall injections 8023 are provided on the rocket support plate 8021 and the support support plate 8022. Figure 6 As shown (attached Figure 6 The internal structure of rocket support plate 8021 is shown in the figure. Support support plate 8022 has the same internal structure as rocket support plate 8021. Rocket support plate 8021 functions as its own fuel injection and rocket gas flow channel; support support plate 8022 functions as a fuel injection supply for central rocket 8, its own fuel injection, and support for central rocket 8. All support plates are made of 304 stainless steel.

[0060] Two sets of second ring injection points 8031 ​​are located in the injection area 803, one of which is a 40×Φ0.3 injection set and the other is an 80×Φ0.3 injection set. The baffle installation area 804 is used to install the central rocket graphite liner baffle and the central rocket sealing cover. The installation threaded holes are 6×M6×1.0, and the thread depth is at least 8mm. After thread processing is completed, the chips should be cleaned and then the installation test should be carried out. To ensure smooth installation with the central rocket graphite liner, the inner surface of the central rocket 8 should adopt positive tolerance.

[0061] The direction from the cylindrical matching area 801 to the baffle installation area 804 is determined to be 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. Among them, 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. Among them, 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.

[0062] 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.

[0063] The 120° phase area and the -120° phase area of ​​the central rocket 8 are both installation areas of 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 of the rocket support plate installation grooves.

[0064] The outer diameter of the oil inlet holes of the two sets of second ring injection points 8031 ​​in the injection area 803 is changed from Φ12 to Φ6. The Φ12 portion should adopt a positive tolerance and a chamfer of 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 sets of second ring injection points 8031 ​​are internal flow channels, which can be completed by a segmented processing-welding method. Before welding, special attention should be paid to the coaxiality of the fuel flow channels to ensure that the fuel injection pressure is not affected by the sudden change in area caused by welding misalignment. After welding, the welds must be cleaned to ensure that the outer surface of the center rocket 8 is smooth and flat. In particular, check whether the nozzle holes of the second ring injection points 8031 ​​are blocked by welding.

[0065] The graphite lining positioning groove on the flange surface of the baffle mounting area 804 should adopt positive tolerance. The center rocket 8 is machined from stainless steel.

[0066] Furthermore, the direct oil injection section 4 in the present invention adopts a direct oil injection section-windowing method and is made of 304 stainless steel.

[0067] 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.

[0068] 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. Figure 11 shown.

[0069] The 0° phase zone of the direct injection section 4 serves as the injection group inlet 401. The inlet hole dimensions are 4 x Φ6 mm. Drilled to the outer diameter of the manifold, the inlet hole is then welded to a DN6 injection rod. The 180° phase zone is symmetrically located with the 0° phase zone and also serves as the injection group inlet 401, sharing the same dimensions. Four sets of first ring injection points 4011 are arranged in parallel along the axial direction of the direct injection section 4.

[0070] The -45° phase area of ​​the direct injection section 4 is the first dynamic pressure sensor installation area 404. The dynamic pressure sensor installation hole size is 5×M10×1.0-tap-through. 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.

[0071] The 45° phase area of ​​the direct injection section 4 is the first static pressure sensor installation area 403. The pressure inlet hole has a size of 11×Φ2-drilled. After processing, the static pressure sensor extension rod is welded to the pressure inlet hole.

[0072] The 90° phase zone is the first windowing area 402. The threaded hole dimensions of the windowing area cover mounting flange are 12×M6×1.0 (drill 15, tap 12). The locating hole dimensions are Φ5, drill 5, plus tolerance. The sealing groove width is 4.1 (plus tolerance), and the groove depth is 3.6 (minus tolerance). The -90° phase zone is structurally similar to the 90° phase zone, with identical threaded hole and sealing groove parameters and fixed locating hole dimensions. However, the locating holes in the two phase zones are asymmetrical. Both first windowing areas 402 are fitted with optical glass windows 9.

[0073] The flange holes on both ends of the direct injection section 4 are 12 x Φ11 through holes. The left flange sealing groove has a positive tolerance of 6.3mm in width and a negative tolerance of 4.1mm in depth. The outer diameter of the stopper on the right flange has a positive tolerance.

[0074] Furthermore, the expansion-concave section 5 is divided into a second expansion section 501 and a concave section 502 according to the airflow direction. Two sets of third annular jet injection points 5011 are located on the second expansion section 501. One set of third annular jet injection points 5011 is a 60×Φ0.3 injection set, and the other set of third annular jet injection points 5011 is an 80×Φ0.3 injection set. The concave section 502 is provided with multiple sensor interfaces and an optical glass window 9.

[0075] The phases of the expansion-concave cavity section 5 are divided from the perspective of the airflow direction. 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 figure. Figure 14 shown.

[0076] More specifically, the 0° phase area 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. After processing, it is required to use M10×1.0×20 fine-thread screws for installation testing to ensure that the fine thread has no slippage 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 through to the outer diameter of the liquid collection cavity, as shown in the attached figure. Figure 15 shown.

[0077] 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. The rocket jet detonation interface 50221 is provided with a liner positioning groove. The liner positioning groove is a stepped structure of Φ35 to Φ20. Both adopt positive tolerances, and a graphite lining is installed inside the liner positioning groove.

[0078] The detonation channel in the pre-detonation tube initiation interface 50222 is drilled through with a Φ10 diameter. A welding step is provided on the outside, consisting of a Φ17 diameter hole concentric with the detonation channel. After machining the pre-detonation tube initiation interface 50222, weld a DN10 male connector. Pay particular attention to strengthening the weld root. After welding, clean the detonation channel.

[0079] Both the 90° and -90° phase zones have a second windowed area 5023. To the left of the second windowed area 5023 in the 90° phase zone is a 60-hole injection assembly oil inlet hole 50231, measuring Φ6 and drilled to the outer diameter of the liquid collection chamber. The bottom surface of the 60-hole injection assembly oil inlet hole 50231 is spotfaced, with the spotfaced surface 143 mm from the center axis and a flat width of 48 mm. The flange screw holes in the second windowed area 5023 have thread dimensions of 12 x M6 x 1.0, drilled 15 and tapped 12. The locating pin hole is Φ5 and drilled 5, and positive tolerances should be used.

[0080] The 135° phase zone is the photomultiplier mounting area 5024. The bottom surface of the mounting area is spotfaced, with the spotface plane 140mm from the center axis and a width of 75.5mm. The mounting area flange hole thread size is 24×M6×1.0, drilled 6 mm and tapped 5 mm. The mounting area step is Φ20.2 mm, deep 6 mm and drilled through Φ6 mm. Negative tolerance is applied to the Φ20.2 mm portion, and the finished dimension must not be less than Φ20.0 mm. Positive tolerance is applied to the Φ6 mm portion. The photomultiplier optical path series includes five workpieces: "Photomultiplier Glass Lens - Large," "Photomultiplier Glass Lens - Small," "Photomultiplier Glass Mirror Graphite Gasket - Large," "Photomultiplier Glass Mirror Graphite Gasket - Small," and "Photomultiplier M10 Washer."

[0081] The 180° phase zone is also the second dynamic pressure sensor installation area 5021, and the relevant requirements are the same as those of the 0° phase zone.

[0082] The -150° phase zone is the water-cooled sensor mounting area 5025. This is the installation area for the water-cooled sensor. The sensor mounting hole dimensions are 1 / 2-20 UNF-2B American threaded, with a countersunk head outer diameter of 14.28mm and a countersunk depth of 1.27mm, with a positive tolerance. To match the thread size of the water-cooled sensor adapter, the mounting hole surface must be raised and spotfaced. The spotface is 146.8mm from the center axis and has a width of 30mm.

[0083] The -135° phase region serves as the second static pressure sensor mounting area 5026. This area features nine Φ2 optical apertures, serving as pressure-inducing holes: two in the second expansion section 501 and seven in the cavity section 502. After the expansion-cavity section 5 is machined, a static pressure sensor extension rod is welded to each pressure-inducing hole.

[0084] The -45° phase zone is the heat flux sensor mounting area 5027. The bottom surface of the heat flux sensor mounting area is spotfaced, with the spotface plane 143mm from the center axis and a width of 48mm. The thread size of the heat flux sensor mounting flange hole is 4×M3×0.5 (drill size 10, tap size 5). The mounting hole is Φ12 (drilled through), and a positive tolerance should be used. The positioning groove of the plug should also have a positive tolerance.

[0085] The working principle of the rocket-based rotating detonation engine combustion chamber structure in this embodiment is as follows:

[0086] After the combustion chamber is filled with air, 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 spray oil. After a certain period of time, a detonation wave is injected from 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.

[0087] When working at extremely low Mach numbers (Ma0-Ma2.0) in the subsonic and supersonic sections, the central rocket 8 opens to provide thrust and forms an ejection effect to inhale air. The combustion gas of the central rocket 8, the intake air, and kerosene undergo rotating detonation combustion together. When working to the Ma1.5-Ma2.0 stage, the central rocket 8 is gradually closed, and pure rotating detonation combustion is carried out in the flow channel.

[0088] When working at low Mach numbers (Ma2.0-Ma3.0), 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 to achieve flame stabilization over a larger range.

[0089] When operating at medium Mach numbers (Ma4.0-Ma5.0), the injection points are located at the four groups of first ring injection points 4011 on the equal direct injection section 4, and the combustion area is located at the concave cavity section 502;

[0090] When operating 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, the rocket support plate 8021, and the support support plate 8022, and the combustion area is located in the small-sized combustion zone;

[0091] 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 needed.

[0092] In summary, under Ma0 and Ma1.0 working 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.

[0093] In addition, it should be noted that the combustion chamber structure in this embodiment can be further applied to rocket-based rotating detonation engines.

[0094] Example 2:

[0095] Example 2 applies the combustion chamber structure in Example 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.

[0096] 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 is connected to the first expansion section 2 and the device throat 1 in sequence in Example 1, 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 oil 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 oil section 4, as shown in the attached figure. Figure 17-20 shown.

[0097] Specifically, the equipment throat 1 and the first expansion section 2 may adopt the existing equipment throat and expansion section structures.

[0098] Furthermore, the variable circular-to-square section 6 has a circular structure at one end and a square structure at the other end, and is used to connect the expansion-concave cavity section 5 at the circular end and the variable geometry throat 7 at the square end.

[0099] The variable geometry throat 7 structure is as shown in the attached Figure 19 and 20 As 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 curved 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 side wall of the throat body 701 .

[0100] In order to provide power to the throat block 702, a drive 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 curved 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 curved vertical plate 7022 and the throat body 701. It should be noted here that the drive 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.

[0101] The working principle of the rocket-based rotating detonation engine test simulator based on the adjustable geometric throat in this embodiment is as follows:

[0102] After the combustion chamber is filled with air, 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 spray oil. After a certain period of time, a detonation wave is injected from 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.

[0103] At very low Mach numbers (Ma0-Ma2.0) in the subsonic and supersonic ranges, center rocket 8 opens to provide thrust and create an ejection effect to draw in air. The combustion gases from center rocket 8, along with the intake air and kerosene, undergo rotating detonation combustion. At Ma1.5-Ma2.0, center rocket 8 gradually closes, allowing pure rotating detonation combustion to occur in the flow path. In this mode, the variable geometry throat 7 has its maximum area.

[0104] 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 in 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 to achieve flame stabilization over a larger range.

[0105] When operating at medium Mach numbers (Ma4.0-Ma5.0), 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 combustion area is located at the concave cavity section 502;

[0106] When operating 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 8021 and the support support plate 8022, and the combustion area is located in the small-sized combustion zone;

[0107] 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 needed.

[0108] The basic principles, main features, and advantages of the present invention are shown and described above. Those skilled in the art should understand that the present invention is not limited to the foregoing embodiments. The foregoing embodiments and descriptions are merely illustrative of the principles of the present invention. Various changes and modifications may be made to the present invention without departing from the spirit and scope of the present invention. Such changes and modifications are intended to fall within the scope of the present invention. The scope of protection claimed in the present invention is defined by the appended claims and their equivalents.

Claims

1. A rocket-based rotating detonation engine combustion chamber structure, comprising an isolation section (3), a 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 annular injection points (4011), the central rocket (8) is provided with two groups of second annular injection points (8031), the expansion-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 injection points (8023).

2. The 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 annular injection points (8031) are located in the fuel injection area (803), and the fuel flow channels of the two groups of second annular injection points (8031) are inner flow channels.

3. The rocket-based rotating detonation engine combustion chamber structure according to claim 2, characterized in that: A transition inner column (301) matching the column 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. The rocket-based rotating detonation engine combustion chamber structure according to claim 4, characterized in that: The 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 the third annular injection points (5011); The concave cavity section (502) is provided with multiple sets of sensor interfaces and optical glass windows (9).

7. The 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 airflow direction, 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 photoelectric multiplier 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).

8. A rocket-based rotating detonation engine test simulator with an adjustable throat geometry, characterized by: Comprising a rocket-based rotating detonation engine combustion chamber structure according to 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 geometry 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 an adjustable geometry throat according to claim 8, characterized in that: The following steps are included: 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) begin 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 central rocket (8) opens to provide thrust and forms an ejection effect to inhale air. The fuel gas of the central 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 central 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 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 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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