Rocket-based rotating detonation engine test simulation device based on adjustable thermal throat, operation method and application thereof

Abstract

The present invention discloses a rocket-based rotating detonation engine test simulation device based on an adjustable thermal throat, its operation method, and application, and belongs to the field of rocket engine research technology. The device includes a device throat, a first expansion section, an isolation section, a constant direct injection oil section, an expansion-concave section, and a thermal throat section with adjustable injection volume, which are connected in sequence; the isolation section and the constant direct injection oil section are equipped with a central rocket; the constant direct injection oil section is provided with four groups of first ring injection injection points, the central rocket is provided with two groups of second ring injection points, and the expansion-concave section is provided with two groups of third ring injection points; the thermal throat section is provided with two groups of thermal throat ring jets. The built-in central rocket in the present invention can provide thrust for the aircraft at zero ground speed, 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; at the same time, the use of the adjustable thermal throat avoids the "back dead weight" problem caused by the installation of a variable geometry throat.

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 test simulation device based on an adjustable thermal throat, an operation method and an 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), a combination of rotating detonation and rocket-based combined cycle engines can achieve greater performance gains. Generally speaking, the engine with the above combination method needs to be equipped with a variable geometry throat for wide speed range performance matching, but the geometric throat adjustment mechanism is complex and the structural mass is large, which reduces the reliability of the engine and has a "back dead weight" effect on the aircraft.

[0005] The solution to the "dead weight" problem of variable geometry throats is an "adjustable thermal throat." Traditional thermal throats require a distributed heat release mechanism, significantly impacting the engine's specific impulse. The rotating detonation wave front, with its concentrated heat release and high heat release rate, is an ideal combustion mode for thermal congestion. Therefore, adapting adjustable thermal throats to rocket-based rotating detonation combined cycle engines has become a new research hotspot. Summary of the Invention

[0006] In response to the above-mentioned problems, the present invention aims to provide a rocket-based rotating detonation engine test simulation device based on an adjustable thermal throat, as well as its operation method and application. 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 rotating detonation combustion at the boundary of rotating detonation combustion extinction; at the same time, the use of the adjustable thermal throat avoids the "back dead weight" problem caused by the installation of a variable geometry throat.

[0007] In order to achieve the above object, the technical solution adopted by the present invention is as follows:

[0008] A rocket-based rotating detonation engine test simulator based on an adjustable thermal throat comprises an equipment throat, a first expansion section, and an isolation section connected in sequence;

[0009] The rear end of the isolation section is connected in sequence with the direct injection section, the expansion-concave cavity section and the thermal throat section with adjustable injection amount; 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;

[0010] The 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; the thermal throat section is provided with two groups of thermal throat ring jets, and the support plate is provided with a group of wall injections.

[0011] Furthermore, the two groups of thermal throat annular sprays are arranged along the thermal throat section close to the front end of the thermal throat section and are arranged in parallel along the axial direction of the thermal throat section;

[0012] The phase of the thermal throat section is divided based on the airflow direction, where the 0° phase area and the 180° phase area are the fuel injection area, the 30° phase area is the thermal throat initiation area, the 90° phase area and the -90° phase area are the third window area, the -45° phase area is the third dynamic pressure sensor installation area, and the -135° phase area is the third static pressure sensor installation area.

[0013] 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;

[0014] 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.

[0015] Furthermore, a transition inner column matching the cylindrical matching area is provided in the isolation section.

[0016] 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.

[0017] 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.

[0018] 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;

[0019] The concave cavity section is provided with multiple sets of sensor interfaces and optical glass windows.

[0020] Furthermore, the phase of the expansion-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.

[0021] Furthermore, the present invention also provides a method for operating the aforementioned rocket-based rotating detonation engine test simulator based on the adjustable thermal throat, comprising the following steps:

[0022] S1: After the direct injection section and the expansion-concave section are inducted, the third group of first annular injection points of the direct injection section and the two groups of thermal throat annular injection points of the thermal throat section first inject fuel and ignite, forming rotating detonation combustion in the thermal throat section;

[0023] 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 operation reaches the Ma1.5-Ma2.0 stage, the center rocket gradually closes, and pure rotating detonation combustion occurs in the flow channel. In this mode, the amount of fuel injected into the thermal throat section is relatively small.

[0024] S3: When the Mach number is Ma2.0-Ma3.0, the fuel injection in the thermal throat section is increased, 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 fuel injection is turned on;

[0025] S4: When the Mach number is Ma4.0-Ma5.0, the injection amount of the thermal throat section continues to increase, and the injection points are located at the four groups of first ring injection points on the equal direct injection section;

[0026] S5: When the Mach number is Ma6.0, the fuel injection amount in the thermal throat section increases, and the injection points are located at the four groups of first ring injection points and the support plate on the direct injection section;

[0027] 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.

[0028] Furthermore, the present invention also provides the application of the rocket-based rotating detonation engine test simulation device based on the adjustable thermal throat as described above in a rocket-based rotating detonation engine.

[0029] The beneficial effects of the present invention are as follows: compared with the prior art, the improvement of the present invention is that:

[0030] 1. The present invention proposes a rocket-based rotating detonation engine test simulation device based on an adjustable thermal throat, in which a rocket-based rotating detonation combustion chamber flow channel is formed inside. 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.

[0031] 2. The rotating detonation ramjet combustion flow channel of the present invention is provided with a total of 11 groups of injection points, corresponding to the relevant injection requirements under Ma0-Ma6 flight conditions and the relevant injection requirements of the thermal throat. By controlling the injection area, combustion area, and thermal throat, stable operation of rotating detonation combustion in a wide speed range and a wide altitude range is achieved.

[0032] 3. The thermal throat section of the present invention is used to adjust the combustion pressure in the flow channel of the combustion chamber, providing pressure guarantee for the wide range operation of the combustion chamber.

[0033] 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

[0034] Figure 1 This is a schematic diagram of the external structure of the rocket-based rotating detonation engine test simulation device based on the adjustable thermal throat of the present invention.

[0035] Figure 2 This is a cross-sectional view of the internal structure of the rocket-based rotating detonation engine test simulation device based on the adjustable thermal throat of the present invention.

[0036] Figure 3 This is a front view of the internal structure of the rocket-based rotating detonation engine test simulation device based on the adjustable thermal throat of the present invention.

[0037] Figure 4 This is a schematic diagram of the external structure of the central rocket of the present invention.

[0038] Figure 5 It is a cross-sectional view of the internal structure of the central rocket of the present invention.

[0039] Figure 6 It is a schematic diagram of the internal structure of the rocket support plate of the present invention.

[0040] 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.

[0041] Figure 8 It is a left sectional view of the central rocket of the present invention.

[0042] Figure 9 This is a schematic diagram of the external structure of the direct injection oil section of the present invention.

[0043] Figure 10 This is a cross-sectional view of the internal structure of the direct injection oil section of the present invention.

[0044] 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.

[0045] Figure 12 This is a schematic diagram of the external structure of the expansion-concave cavity section of the present invention.

[0046] Figure 13 This is a cross-sectional view of the internal structure of the expansion-concave cavity section of the present invention.

[0047] 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.

[0048] 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.

[0049] 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.

[0050] Figure 17 This is a schematic diagram of the external structure of the thermal throat section of the present invention.

[0051] Figure 18 This is a cross-sectional view of the internal structure of the thermal throat section of the present invention.

[0052] Figure 19 It is a schematic diagram of the phase partitioning of the thermal throat section in the side section 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. Second window area; 50231, 6 0. Oil inlet hole of injection group; 5024. Photomultiplier installation area; 5025. Water-cooled sensor installation area; 5026. Second static pressure sensor installation area; 5027. Heat flow sensor installation area; 6. Thermal throat section; 601. Thermal throat ring injection; 602. Fuel injection area; 603. Thermal throat detonation area; 6031. Detonation interface; 604. Third window area; 605. Third dynamic pressure sensor installation area; 606. Third static pressure sensor installation area; 7. Center rocket; 701. Cylindrical matching area; 702. Support plate installation area; 7021. Rocket support plate; 7022. Support support plate; 7023. Wall injection; 703. Fuel injection area; 7031. Second ring injection point; 704. Baffle installation area; 8. 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] The present invention provides a rocket-based rotating detonation engine test simulation device based on an adjustable thermal throat, as shown in the attached Figure 1-19 As shown. The device comprises, from front to back, a device throat 1, a first expansion section 2, and an isolation section 3, which are connected in sequence. The rear end of the isolation section 3 is connected in sequence to a constant direct injection section 4, an expansion-recessed cavity section 5, and a thermal throat section 6 with adjustable injection volume. The isolation section 3 and the constant direct injection section 4 are built with a central rocket 7, and the central rocket 7 is connected to the isolation section 3 via a support plate.

[0056] The equal direct injection oil section 4 is provided with four groups of first ring jet injection points 4011, the central rocket 7 is provided with two groups of second ring jet injection points 7031, the expansion-concave cavity section 5 is provided with two groups of third ring jet injection points 5011; the thermal throat section 6 is provided with two groups of thermal throat ring jets 601; and a group of wall injections 7023 is opened on the support plate.

[0057] Specifically, the throat 1, first expansion section 2 and isolation section 3 of the equipment can adopt the existing throat, expansion section and isolation section structure. The isolation section 3 is provided with a transfer inner column 301, and the inner surface of the transfer inner column 301 is matched with the cylindrical surface of the central rocket 7 series for installing the central rocket 7.

[0058] Furthermore, the central rocket 7 is divided into four functional areas from front to back, namely the cylinder matching area 701, the support plate installation area 702, the injection area 703 and the baffle installation area 704, as shown in the attached figure. Figure 7 As shown. The cylindrical matching area 701 is cylindrically matched with the inner surface of the transfer inner column 301, and a negative tolerance is adopted; the central rocket 7 is connected to the isolation section 3 through a support plate in the support plate installation area 702. The support plate specifically includes a rocket support plate 7021 and a support support plate 7022. The support plate installation area 702 is provided with a rocket support plate installation groove and a support support plate installation groove, which are respectively used to install the rocket support plate 7021 and the support support plate 7022. As shown in the attached figure Figure 8 As shown. A set of wall injection 7023 is provided on the rocket support plate 7021 and the support support plate 7022. Figure 6 As shown (attached Figure 6 The internal structure of rocket support plate 7021 is shown in the figure. Support plate 7022 has the same internal structure as rocket support plate 7021. Rocket support plate 7021 functions as its own fuel injection and rocket gas flow channel; support plate 7022 functions as a fuel injection supply for central rocket 7, its own fuel injection, and support for central rocket 7. All support plates are made of 304 stainless steel.

[0059] Two sets of second ring injection points 7031 are located in the injection area 703, one of which is a 40×Φ0.3 injection set and the other is an 80×Φ0.3 injection set. The baffle installation area 704 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 7 should adopt positive tolerance.

[0060] The direction from the cylindrical matching area 701 to the baffle installation area 704 is determined to be the phase perspective. Under this perspective, the central rocket 7 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 7021, and the 120° phase zone and the -120° phase zone are the installation areas of the support support plate 7022. 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.

[0061] The 0° phase area is the installation area of ​​rocket support plate 7021. The outer contour of the rocket support plate installation groove adopts positive tolerance. The positive tolerance range should consider the positioning dimension tolerance zone on the basis of being consistent with the support plate installation groove of isolation section 3 to ensure smooth installation.

[0062] The 120° phase area and the -120° phase area are both installation areas for the support plate 7022. 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.

[0063] The outer diameter of the oil inlet hole of the two groups of second ring injection points 7031 in the injection area 703 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 path of the two groups of second ring injection points 7031 is an internal flow path, which can be completed by a segmented processing-welding method. Before welding, special attention should be paid to the coaxiality of the fuel flow path 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. In particular, check whether the nozzle of the second ring injection point 7031 is blocked by welding.

[0064] The graphite lining positioning groove on the flange surface of the baffle mounting area 704 should adopt positive tolerance. The center rocket 7 is machined from stainless steel.

[0065] 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.

[0066] 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.

[0067] 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.

[0068] 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.

[0069] 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.

[0070] 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.

[0071] 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 (positive tolerance). The sealing groove width is 4.1 (positive tolerance), and the groove depth is 3.6 (negative 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 8.

[0072] 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.

[0073] 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 sets of sensor interfaces and an optical glass window 8.

[0074] 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-cooled 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 8, as shown in the attached figure. Figure 14 shown.

[0075] 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.

[0076] 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.

[0077] 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.

[0078] The 90° and -90° phase zones are both second windowed areas 5023. The left side of the second windowed area 5023 in the 90° phase zone is the 60-hole injection group oil inlet hole 50231, which is Φ6 and drilled through to the outer diameter of the liquid collecting chamber. The bottom surface of the 60-hole injection group oil inlet hole 50231 is flattened, and the flattened surface is 143mm from the center axis. The flat surface width is 48mm. The thread size of the flange screw holes in the second windowed area 5023 is 12×M6×1.0, drill 15 and tap 12. The locating pin hole size is Φ5 and drill 5. Positive tolerance should be used, as shown in the attached figure. Figure 16 shown.

[0079] 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."

[0080] 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.

[0081] 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 flux sensor installation area 5027. The relevant requirements are the same as those for the second dynamic pressure sensor installation area 5021. After installation, the heat flux sensor is fixed and encapsulated using the heat flux sensor plug.

[0082] Furthermore, the thermal throat section 6 is located at the rear end of the expansion-concave cavity section 5, and the phase of the thermal throat section 6 is divided based on the airflow direction, wherein the 0° phase area and the 180° phase area are the fuel injection area 602, the 30° phase area is the thermal throat detonation area 603, the 90° phase area and the -90° phase area are the third window area 604, the -45° phase area is the third dynamic pressure sensor installation area 605, and the -135° phase area is the third static pressure sensor installation area 606, as shown in the attached figure. Figure 19 As shown, two sets of thermal throat ring sprays 601 are arranged axially in parallel along the thermal throat section 6.

[0083] The 0° phase area and the 180° phase area are the fuel injection area 602, and two sets of thermal throat ring sprays 601 are arranged in parallel. The spray hole size of the two sets of thermal throat ring sprays 601 is 40×Φ0.3, the liquid collecting cavity is 4×3, and the oil inlet size is Φ6-drilled to the outer diameter of the liquid collecting cavity.

[0084] The 30° phase zone is the thermal throat initiation zone 603. Two sets of initiation ports 6031 are provided, one for establishing initiation in two different thermal throat positions. The initiation ports 6031 are Φ10 drilled through and then stepped to Φ17. After machining, a DN10 male connector is welded on.

[0085] The 90° phase and −90° phase regions constitute the third windowing region 604 , and are also provided with an optical glass window 8 .

[0086] The -45° phase area is the third dynamic pressure sensor installation area 605, which is provided with a dynamic pressure sensor installation hole with a size of 4×M10×1.0-tapped. The installation surface is spotfaced.

[0087] The -135° phase area is the third static pressure sensor installation area 606, which is provided with corresponding pressure-inducing holes. The size of the pressure-inducing holes 6081 is 20×Φ2-drilled, and the extended pressure rod is welded after processing.

[0088] The working principle of the rocket-based rotating detonation engine test simulator based on the adjustable thermal throat in the present invention is as follows:

[0089] After air enters the combustion chamber (the direct injection section 4 and the expansion-recess section 5), fuel is first injected and ignited by the third group of first annular injection points 4011 of the direct injection section 4 and the two groups of thermal throat annular injection points 601 of the thermal throat section 6, forming rotating detonation combustion in the thermal throat section 6. After the rotating detonation wave in the thermal throat section 6 stabilizes, fuel injection in front of the ramjet flow channel is initiated. When the fuel reaches the front of the thermal throat section 6, it is ignited by the rotating detonation wave and propagates back to the combustion zone in front of the flow channel. This combustion zone can then conduct slow-burn ramjet combustion or rotating detonation combustion, depending on the situation. When rotating detonation combustion is selected, a detonation wave is injected into the initiation point in front of the flow channel, converting the combustion into rotating detonation combustion.

[0090] When operating at very low Mach numbers (Ma0-Ma2.0) in the subsonic and supersonic ranges, center rocket 7 opens to provide thrust and create an ejection effect to draw in air. The fuel from center rocket 7, along with the intake air and kerosene, undergoes rotating detonation combustion. When operating in the Ma1.5-Ma2.0 range, center rocket 7 gradually closes, and pure rotating detonation combustion occurs in the flow path. In this mode, the amount of fuel injected into thermal throat section 6 is relatively low.

[0091] When working at low Mach numbers (Ma2.0-Ma3.0), the fuel injection of the thermal throat section 6 is increased, and the injection point is selected as the second ring injection point 7031 on the central rocket 7, and the combustion area is located in the concave cavity section 502; when approaching the flame stabilization equivalence ratio limit, the rocket support plate 7021 and the support support plate 7022 are turned on to achieve flame stabilization over a larger range.

[0092] When working at medium Mach numbers (Ma4.0-Ma5.0), the injection amount of the thermal throat section 6 continues to increase, 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;

[0093] When operating at high Mach number (Ma6.0), the fuel injection amount of thermal throat section 6 increases, and 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. The combustion area is located in the small-sized combustion zone;

[0094] 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 7 can be opened as needed.

[0095] 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 test simulation device based on an adjustable thermal throat, comprising an equipment throat (1), a first expansion section (2) and an isolation section (3) connected in sequence; characterized in that: The rear end of the isolation section (3) is connected in sequence to a constant direct injection section (4), an expansion-concave cavity section (5), and a thermal throat section (6) with adjustable injection amount; a central rocket (7) is built into the isolation section (3) and the constant direct injection section (4), and the central rocket (7) 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 (7) is provided with two groups of second annular injection points (7031), and the expansion-concave cavity section (5) is provided with two groups of third annular injection points (5011); the thermal throat section (6) is provided with two groups of thermal throat annular injections (601), and the support plate is provided with a group of wall injections (7023); The two groups of thermal throat ring sprays (601) are arranged along the thermal throat section (6) close to the front end of the thermal throat section (6) and are arranged in parallel along the axial direction of the thermal throat section (6); The phase of the thermal throat section (6) is divided based on the airflow direction, wherein the 0° phase area and the 180° phase area are the fuel injection area (602), the 30° phase area is the thermal throat detonation area (603), the 90° phase area and the -90° phase area are the third window area (604), the -45° phase area is the third dynamic pressure sensor installation area (605), and the -135° phase area is the third static pressure sensor installation area (606).

2. The rocket-based rotating detonation engine test simulator based on an adjustable thermal throat according to claim 1, characterized in that: The central rocket (7) is divided into a cylindrical matching area (701), a support plate installation area (702), a fuel injection area (703) and a baffle installation area (704) from front to back; The two groups of second annular injection points (7031) are located in the fuel injection area (703), and the fuel flow channels of the two groups of second annular injection points (7031) are inner flow channels.

3. The rocket-based rotating detonation engine test simulator based on an adjustable thermal throat according to claim 2, characterized in that: A transition inner column (301) matching the column matching area (701) is provided in the isolation section (3).

4. The rocket-based rotating detonation engine test simulator based on an adjustable thermal throat 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 (8).

5. The rocket-based rotating detonation engine test simulator based on an adjustable thermal throat 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 test simulator based on an adjustable thermal throat 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 (8).

7. The rocket-based rotating detonation engine test simulator based on an adjustable thermal throat 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. The method for operating the rocket-based rotating detonation engine test simulator based on an adjustable thermal throat according to any one of claims 1 to 7, characterized in that: The following steps are included: S1: After the direct injection section (4) and the expansion-concave cavity section (5) are fed with air, the third group of first annular injection points (4011) of the direct injection section (4) and the two groups of thermal throat annular injections (601) of the thermal throat section (6) first inject oil and ignite, forming rotating detonation combustion in the thermal throat section (6); S2: When the Mach number is Ma0-Ma2.0, the central rocket (7) is opened to provide thrust and form an ejection effect to inhale air. The fuel gas of the central rocket (7) 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 (7) is gradually closed, and pure rotating detonation combustion is performed in the flow channel. In this mode, the amount of fuel injected into the thermal throat section (6) is relatively small. S3: When the Mach number is Ma2.0-Ma3.0, the fuel injection of the thermal throat section (6) is increased, and the injection point is selected at the second ring injection point (7031) on the central rocket (7); when it approaches the flame stabilization equivalence ratio limit, the support plate fuel injection is turned on; S4: When the Mach number is Ma4.0-Ma5.0, the injection amount of the thermal throat section (6) continues to increase, 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 injection amount of the thermal throat section (6) increases, and the injection points are located at the four groups of first ring injection points (4011) and the support plate on the 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 center rocket (7) is opened as appropriate.

9. Application of the rocket-based rotating detonation engine test simulation device based on the adjustable thermal throat as described in any one of claims 1 to 7 in a rocket-based rotating detonation engine.

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