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

Abstract

The invention discloses a rocket-based rotary detonation engine test simulation device based on an adjustable thermal throat and an operation method and application thereof, and belongs to the technical field of rocket engine research. The device comprises an equipment throat, a first expansion section, an isolation section, an equal-direction oil injection section, an expansion-concave cavity section and a thermal throat section with adjustable oil injection quantity which are connected in sequence, a central rocket is arranged in the isolation section and the equal direct oil injection section; four groups of first ring-jet injection points are arranged on the equal-straight oil injection section, two groups of second ring-jet injection points are arranged on the central rocket, and two groups of third ring-jet injection points are arranged on the expansion-concave cavity section; and two groups of thermal throat annular sprayers are arranged on the thermal throat section. The built-in center rocket can provide thrust for the aircraft when the ground is at zero speed, provide additional thrust when the cruising thrust of the aircraft is insufficient, and provide a flame stabilizing mechanism for rotary detonation combustion at the rotary detonation combustion flameout boundary; and meanwhile, the problem of dead weight of the back caused by additionally arranging a variable geometric throat is avoided by using the adjustable thermal throat.

Description

Technical Field

[0001] The 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, and an operation method and application thereof. Background Art

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

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

[0004] In view of the above 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 engines of the above combination need to be equipped with variable geometry throats for wide speed range performance matching, but the geometry 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 "adjustable thermal throats". Traditional thermal throats need to work in a distributed heat release mode, which has a great impact on the engine specific impulse. The rotating detonation wave front has the characteristics of concentrated heat release and high heat release rate, which is an ideal combustion mode for forming thermal congestion. Therefore, how to use adjustable thermal throats to adapt rocket-based rotating detonation combined cycle engines has become a new research hotspot. Summary of the invention

[0006] In view of 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, and 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, provide additional thrust when the aircraft's cruising thrust is insufficient, and provide a flame stabilization mechanism for the rotating detonation combustion at the rotating detonation combustion extinction boundary. 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: A rocket-based rotating detonation engine test simulation device based on an adjustable thermal throat includes an equipment throat, a first expansion section and an isolation section connected in sequence; The rear end of the isolation section is connected in sequence with the equal direct injection section, the expansion-concave cavity section and the thermal throat section with adjustable injection amount; the isolation section and the equal direct injection section are built with a central rocket, and the central rocket is connected to the isolation section through a support plate; The equal direct injection oil section is provided with four groups of first ring jet injection points, the central rocket is provided with two groups of second ring jet injection points, the expansion-concave cavity section is provided with two groups of third ring jet injection points; 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.

[0008] Furthermore, two groups of thermal throat ring sprays are arranged along the thermal throat section close to the front end of the thermal throat section and in parallel along the axial direction of the thermal throat section; The phase of the thermal throat section is divided based on the airflow direction, where the 0° phase zone and the 180° phase zone are the fuel injection zone, the 30° phase zone is the thermal throat ignition zone, the 90° phase zone and the -90° phase zone are the third windowing zone, the -45° phase zone is the third dynamic pressure sensor installation zone, and the -135° phase zone is the third static pressure sensor installation zone.

[0009] Furthermore, the central rocket is divided into a cylindrical matching area, a support plate installation area, a fuel injection area and a baffle installation area from front to back; The two groups of second ring-jet injection points are located in the injection area, and the fuel flow channels of the two groups of second ring-jet injection points are inner flow channels.

[0010] Furthermore, a transition inner column matching with the cylindrical matching area is provided in the isolation section.

[0011] Furthermore, the direct injection oil section adopts a direct injection oil section-windowing method, is provided with four groups of first ring injection points with different injection point parameters, and is provided with an optical glass window.

[0012] Furthermore, the direct injection section is distributed in six phases according to functional areas, the 0° phase area and the 180° phase area are the fuel inlet of the injection group, the 90° phase area and the -90° phase area are the first window area, the 45° phase area is the first static pressure sensor installation area, and the -45° phase area is the first dynamic pressure sensor installation area.

[0013] Furthermore, the expansion-concave cavity section is divided into a second expansion section and a concave cavity section according to the airflow direction, and the second expansion section is provided with two groups of the third annular injection points; The concave cavity section is provided with multiple sets of sensor interfaces and optical glass windows.

[0014] Furthermore, the phase of the expansion-concave cavity section is divided from the perspective of the airflow direction. The 0° phase zone and the 180° phase zone are the second dynamic pressure sensor installation area, the 45° phase zone is the ignition area, the 90° phase zone and the -90° phase zone are the second window area, the 135° phase zone is the photoelectric multiplier installation area, the -150° phase zone is the water-cooled sensor installation area, the -135° phase zone is the second static pressure sensor installation area, and the -45° phase zone is the heat flow sensor installation area.

[0015] Furthermore, the present invention also provides an operation method of the rocket-based rotating detonation engine test simulation device based on the adjustable thermal throat as described above, comprising the following steps: S1: After the direct injection section and the expansion-concave cavity section are inhaled, the injection point of the third group of first ring injections in the direct injection section and the two groups of thermal throat ring injections in the thermal throat section are first injected and ignited, forming rotating detonation combustion in the thermal throat section; S2: When the Mach number is Ma0-Ma2.0, the center rocket opens to provide thrust and forms an ejection effect to inhale air. The center rocket gas, intake air and kerosene perform rotating detonation combustion together. When the operation reaches the Ma1.5-Ma2.0 stage, the center rocket is gradually closed, and pure rotating detonation combustion is performed in the flow channel. In this mode, the amount of fuel injected in the thermal throat section is small. S3: When the Mach number is Ma2.0-Ma3.0, the fuel injection in the thermal throat section increases, and the injection point is selected at the injection point of the second ring on the central rocket; when it approaches the limit of the flame stabilization equivalence ratio, the support plate injection is turned on; S4: When the Mach number is Ma4.0-Ma5.0, the injection amount of the thermal throat section continues to increase, and the injection point is located at the four groups of first ring injection points on the equal direct injection section; S5: When the Mach number is Ma6.0, the 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 equal direct injection section; S6: When the altitude increases or approaches the boundary of rotating detonation flameout, in order to maintain the stability of the rotating detonation wave in the combustion chamber, the fuel injection is moved forward and the center rocket is opened as appropriate.

[0016] 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.

[0017] The beneficial effects of the present invention are as follows: compared with the prior art, the present invention has the following improvements: 1. The present invention proposes a rocket-based rotating detonation engine 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 rotating detonation combustion extinction boundary.

[0018] 2. A total of 11 groups of injection points are arranged in the rotating detonation ramjet combustion flow channel of the present invention, which respectively correspond to the relevant injection requirements under Ma0-Ma6 flight conditions and the relevant injection requirements of the thermal throat. The stable operation of rotating detonation combustion in a wide speed range and a wide height range is achieved by controlling the injection area, the combustion area and the thermal throat.

[0019] 3. The thermal throat section in 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.

[0020] 4. The engine of the present invention adopts rocket jet / pre-detonator coordinated detonation technology to solve the problem of difficulty in ignition of rotating detonation with low dynamic pressure. BRIEF DESCRIPTION OF THE DRAWINGS

[0021] Figure 1 It 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.

[0022] 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.

[0023] 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.

[0024] Figure 4 It is a schematic diagram of the external structure of the central rocket of the present invention.

[0025] Figure 5 It is a cross-sectional view of the internal structure of the center rocket of the present invention.

[0026] Figure 6 It is a schematic diagram of the internal structure of the rocket support plate of the present invention.

[0027] Figure 7 It is a schematic diagram of phase partitioning of the front cross-sectional view of the central rocket of the present invention.

[0028] Figure 8 It is a left sectional view of the central rocket of the present invention.

[0029] Fig. 9 It is a schematic diagram of the external structure of the direct injection oil section of the present invention.

[0030] Fig.10 It is a cross-sectional view of the internal structure of the direct injection oil section of the present invention.

[0031] Fig.11 It is a schematic diagram of the phase division of the direct injection section in the present invention.

[0032] Fig.12 It is a schematic diagram of the external structure of the expansion-concave cavity section of the present invention.

[0033] Fig.13 It is a cross-sectional view of the internal structure of the expansion-concave cavity section of the present invention.

[0034] Fig.14 It is a schematic diagram of the phase partitioning of the expansion-concave cavity section side section of the present invention.

[0035] Fig.15 It is a schematic diagram of the external structure of the 0° phase zone of the expansion-concave cavity section of the present invention.

[0036] Fig.16 It is a schematic diagram of the external structure of the 45° phase zone and the 90° phase zone of the expansion-concave cavity section of the present invention.

[0037] Fig.17 It is a schematic diagram of the external structure of the thermal throat section of the present invention.

[0038] Fig.18 It is a cross-sectional view of the internal structure of the thermal throat section of the present invention.

[0039] Fig.19 It is a schematic diagram of the phase partitioning of the thermal throat section side section of the present invention.

[0040] Among them: 1. Equipment throat; 2. First expansion section; 3. Isolation section; 301. Transfer inner column; 4. Direct injection section; 401. Inlet of injection group; 4011. First ring injection injection point; 402. First window area; 403. First static pressure sensor installation area; 404. First dynamic pressure sensor installation area; 5. Expansion-concave cavity section; 501. Second expansion section; 5011. Third ring injection injection point; 502. Concave cavity section; 5021. Second dynamic pressure sensor installation area; 50211. Second dynamic pressure sensor installation hole; 50212. 80-hole injection group oil inlet hole; 5022. Detonation area; 50221. Rocket jet detonation interface; 50222. Pre-detonation tube detonation interface; 5023. Second window area; 50231, 6 0 hole injection group oil inlet hole; 5024, photoelectric multiplier installation area; 5025, water cooling sensor installation area; 5026, second static pressure sensor installation area; 5027, thermal flux sensor installation area; 6, thermal throat section; 601, thermal throat ring spray; 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 spray injection point; 704, baffle installation area; 8, optical glass window. DETAILED DESCRIPTION

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

[0042] 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 The device comprises, from front to back, a device throat 1, a first expansion section 2 and an isolation section 3 connected in sequence, the rear end of the isolation section 3 is connected in sequence to an equal direct injection section 4, an expansion-recessed cavity section 5 and a thermal throat section 6 with adjustable injection amount; the isolation section 3 and the equal direct injection section 4 are built with a central rocket 7; and the central rocket 7 is connected to the isolation section 3 through a support plate The equal direct injection oil section 4 is provided with four groups of first ring jet injection points 4011, the central rocket 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 are opened on the support plate.

[0043] Specifically, the throat 1, the first expansion section 2 and the 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 central rocket 7 series for the installation of the central rocket 7.

[0044] Furthermore, the central rocket 7 is divided into four functional areas from front to back, namely, a cylinder matching area 701, a support plate installation area 702, a fuel injection area 703 and a baffle installation area 704, as shown in the attached figure. Figure 7 The cylindrical matching area 701 is cylindrically matched with the inner profile 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, and 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 A set of wall injections 7023 are provided on the rocket support plate 7021 and the support support plate 7022, as shown in the attached Figure 6 As shown (attached Figure 6 The internal structure of the rocket support plate 7021 is shown in the figure. The support plate 7022 has the same internal structure as the rocket support plate 7021). The rocket support plate 7021 has the functions of self-injection and rocket gas flow channel; the support plate 7022 has the functions of supplying fuel to the central rocket 7, self-injection and supporting the central rocket 7. The support plate series are all made of 304 stainless steel.

[0045] Two groups of second ring injection points 7031 are located in the injection area 703, one of which is a 40×Φ0.3 injection group and the other is an 80×Φ0.3 injection group; the baffle installation area 704 is used to install the central rocket graphite liner baffle and the central rocket sealing cover plate, and the installation threaded hole size is 6×M6×1.0, and the thread depth is at least 8mm. After the thread processing is completed, the chip cleaning should be carried out, and the installation test should be carried out after the chip cleaning. In order to ensure the smooth installation with the central rocket graphite liner, the inner surface of the central rocket 7 should adopt a positive tolerance.

[0046] The direction from the cylindrical matching area 701 to the baffle installation area 704 is determined as 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. 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. 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.

[0047] 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 be consistent with the support plate installation groove of the isolation section 3, and the positioning dimension tolerance zone should be considered to ensure smooth installation.

[0048] The 120° phase zone and the -120° phase zone 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.

[0049] The outer diameter of the oil inlet hole of the two groups of second ring-jet injection points 7031 in the injection area 703 is changed from Φ12 to Φ6, of which the Φ12 part should adopt a positive tolerance, chamfer C1.5, and the bottom surface of the support plate mounting groove should be used as the positioning reference. The fuel flow channels of the two groups of second ring-jet injection points 7031 are internal flow channels, and the fuel flow channel processing can be completed by the segmented processing-welding method. Before welding, special attention should be paid to the coaxiality of the fuel flow channel to ensure that the fuel injection pressure is not affected by the sudden change in area caused by welding misalignment. After welding, it is required to clean the weld to ensure that the outer surface of the center rocket 7 is smooth and flat, and pay special attention to check whether the spray hole of the second ring-jet injection point 7031 is blocked due to welding.

[0050] The flange surface graphite liner positioning groove of the baffle mounting area 704 should adopt positive tolerance. The center rocket 7 is machined from stainless steel.

[0051] Furthermore, the direct injection section 4 in the present invention adopts a direct injection section-windowing method and is made of 304 stainless steel.

[0052] 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.

[0053] According to the airflow direction, the direct injection section 4 is distributed in six phases according to the functional areas. The 0° phase area and the 180° phase area are the injection group oil inlet 401, the 90° phase area and the -90° phase area are the first window area 402, the 45° phase area is the first static pressure sensor installation area 403, and the -45° phase area is the first dynamic pressure sensor installation area 404. Fig.11 shown.

[0054] The 0° phase zone of the equal direct injection section 4 is the injection group oil inlet 401, and the oil inlet hole size of the injection group oil inlet 401 is 4×Φ6-. Drill through to the outer diameter of the liquid collecting cavity, weld the DN6 injection rod after the oil inlet hole is processed, and the 180° phase zone is symmetrically distributed with the 0° phase zone, which is also the injection group oil inlet 401, and the relevant size is the same. Along the axial direction of the equal direct injection section 4, four groups of first ring injection injection points 4011 are arranged in parallel.

[0055] The -45° phase zone of the direct injection section 4 is the first dynamic pressure sensor installation area 404. The size of the dynamic pressure sensor installation hole is 5×M10×1.0-tap. The bottom surface of the installation area is required to be flattened. The flattened plane is 120 mm from the center axis and the plane width is 44 mm.

[0056] The 45° phase area of ​​the direct injection section 4 is the first static pressure sensor installation area 403, and the pressure-inducing hole has a size of 11×Φ2-drilled. After processing, the static pressure sensor extension rod is welded to the pressure-inducing hole.

[0057] The 90° phase zone is the first window zone 402. The size of the threaded hole of the window zone cover plate mounting flange is 12×M6×1.0-drill 15 tap 12, the size of the positioning hole is Φ5 drill 5-positive tolerance, the sealing groove width is 4.1-positive tolerance, and the groove depth is 3.6-negative tolerance. The -90° phase zone is similar in structure to the 90° phase zone, with the same parameters of the threaded hole and the sealing groove, and the same fixed size of the positioning hole, but the positioning holes of the two phase zones are asymmetric. Both first window zones 402 are installed with optical glass windows 8.

[0058] The flange hole size at both ends of the direct injection section 4 is 12×Φ11-through hole, the left end flange sealing groove size is groove width 6.3mm-positive tolerance, groove depth 4.1mm-negative tolerance, stopper outer diameter negative tolerance. The right end flange stopper outer diameter positive tolerance.

[0059] Furthermore, the expansion-concave section 5 is divided into a second expansion section 501 and a concave section 502 according to the airflow direction, and two groups of third annular injection points 5011 are located on the second expansion section 501, one group of third annular injection points 5011 is a 60×Φ0.3 injection group, and the other group of third annular injection points 5011 is an 80×Φ0.3 injection group. The concave section 502 is provided with multiple groups of sensor interfaces and optical glass windows 8.

[0060] The phase of the expansion-concave cavity section 5 is divided from the perspective of the flow direction of the airflow. The 0° phase area and the 180° phase area are the second dynamic pressure sensor installation area 5021, the 45° phase area is the detonation area 5022, the 90° phase area and the -90° phase area are the second window area 5023, the 135° phase area is the photomultiplier installation area 5024, the -150° phase area 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 8, as shown in the attached Fig.14 shown.

[0061] More specifically, the 0° phase zone is the second dynamic pressure sensor installation area 5021, which requires the bottom surface to be flattened, with the flattened plane 143mm from the center axis and a plane width of 48mm. The second dynamic pressure sensor installation hole 50211 is a 4×M10×1.0 fine-thread through hole, which requires the use of M10×1.0×20 fine-thread screws for installation testing after processing to ensure that the fine-thread has no stripping or broken threads. The bottom of the second dynamic pressure sensor installation area 5021 is the 80-hole injection group oil inlet hole 50212, with a hole diameter of Φ6, which is drilled to the outer diameter of the liquid collecting cavity, as shown in the attached Fig.15 shown.

[0062] The 45° phase zone is the detonation zone 5022. The detonation zone 5022 is provided with a rocket jet detonation interface 50221 and a pre-detonation tube detonation interface 50222. The pre-detonation tube detonation interface 50222 is welded with a DN10 male head. The flange hole thread size of the rocket jet detonation interface 50221 is 4×M6×1.0. A liner positioning groove is provided in the rocket jet detonation interface 50221. The liner positioning groove is a stepped structure of Φ35 to Φ20. Positive tolerances are adopted, and a graphite liner is installed inside the liner positioning groove.

[0063] The size of the detonation wave flow channel in the pre-detonation tube detonation interface 50222 is Φ10-drilled. A welding step is reserved on the outside, and the welding step is a Φ17 hole concentric with the detonation wave flow channel. After the pre-detonation tube detonation interface 50222 is processed, a DN10 male connector is welded. It is required to pay special attention to the reinforcement of the welding root, and pay attention to cleaning the detonation wave flow channel after welding.

[0064] The 90° phase zone and the -90° phase zone are both the second window zone 5023. The left side of the second window zone 5023 of the 90° phase zone is the 60-hole injection group oil inlet hole 50231, with a size of Φ6-drilled to the outer diameter of the liquid collecting cavity. The bottom surface of the 60-hole injection group oil inlet hole 50231 is flattened, and the flattened plane is 143mm away from the center axis. The plane width is 48mm. The thread size of the flange screw hole of the second window zone 5023 is 12×M6×1.0-drill 15 tap 12, and the size of the positioning pin hole is Φ5 drill 5. Positive tolerance should be used, as shown in the attached Fig.16 shown.

[0065] The 135° phase zone is the photomultiplier installation area 5024. The bottom surface of the installation area is flattened. The flattened plane is 140mm from the center axis and the plane width is 75.5mm. The thread size of the flange hole in the installation area is 24×M6×1.0-drill 6 tap 5, the step in the installation area is Φ20.2-deep 6 turn Φ6-drill through, the Φ20.2 part adopts negative tolerance, and the finished size shall not be less than Φ20.0. The Φ6 part adopts positive tolerance. There are five kinds of workpieces in the photomultiplier optical path series, namely "photomultiplier glass lens_large", "photomultiplier glass lens_small", "photomultiplier glass mirror graphite gasket_large", "photomultiplier glass mirror graphite gasket_small" and "photomultiplier M10 gasket".

[0066] The 180° phase zone is also the second dynamic pressure sensor installation area 5021, and the relevant requirements are consistent with those of the 0° phase zone.

[0067] 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 consistent with the second dynamic pressure sensor installation area 5021. After the heat flux sensor is installed, it is fixed and packaged through the heat flux sensor plug.

[0068] 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 Fig.19 As shown; two sets of thermal throat ring sprays 601 are arranged axially in parallel along the thermal throat section 6.

[0069] The 0° phase zone and the 180° phase zone are the fuel injection zone 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.

[0070] The 30° phase zone is the thermal throat detonation zone 603, and two sets of detonation interfaces 6031 are provided, which are used for thermal throat detonation establishment under two different thermal throat position conditions. The size of the detonation interface 6031 is Φ10-drilled through and turned to Φ17 steps, and a DN10 male connector is welded after processing.

[0071] The 90° phase and −90° phase regions are the third windowing region 604 , which is also provided with an optical glass window 8 .

[0072] The -45° phase area is the third dynamic pressure sensor installation area 605, which is provided with a dynamic pressure sensor installation hole, and the size of the dynamic pressure sensor installation hole is 4×M10×1.0-tap and punch. The installation surface is flattened.

[0073] The -135° phase zone is the third static pressure sensor installation zone 606, which is provided with corresponding pressure-inducing holes. The size of the pressure-inducing hole 6081 is 20×Φ2-drilled, and the extended pressure rod is welded after processing.

[0074] The working principle of the rocket-based rotating detonation engine test simulation device based on the adjustable thermal throat in the present invention is: After the combustion chamber (the direct injection section 4 and the expansion-concave section 5) is inhaled, 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 a rotating detonation combustion in the thermal throat section 6. After the rotating detonation wave of the thermal throat section 6 is stabilized, the injection in front of the ram flow channel is started, and when the fuel flows to the front of the thermal throat section 6, it is ignited by the rotating detonation wave, and the flame is transmitted back to the combustion zone in front of the flow channel. The combustion zone in front of the flow channel can perform slow combustion ramjet combustion or rotating detonation combustion according to the situation. When rotating detonation combustion is selected, a detonation wave is injected into the ignition point in front of the flow channel to convert the combustion into rotating detonation combustion.

[0075] When working at extremely low Mach numbers (Ma0-Ma2.0) in the subsonic and supersonic sections, the center rocket 7 opens to provide thrust and forms an ejection effect to inhale air. The gas from the center rocket 7, the intake air, and kerosene undergo a rotating detonation combustion. When working at the Ma1.5-Ma2.0 stage, the center 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.

[0076] When working at low Mach numbers (Ma2.0-Ma3.0), the fuel injection of the thermal throat section 6 increases, and the injection point selects the second ring injection point 7031 on the central rocket 7, and the combustion area is located at 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.

[0077] When working at medium Mach number (Ma4.0-Ma5.0), the injection amount of thermal throat section 6 is continuously increased, the injection point is located at the four groups of first ring injection points 4011 on the equal direct injection section 4, and the combustion area is located at the concave cavity section 502; When working at high Mach number (Ma6.0), the fuel injection amount of thermal throat section 6 increases, and the fuel injection points are located at the four groups of first ring injection points 4011 and rocket support plate 7021 and support support plate 7022 on the equal direct injection section 4, and the combustion area is located in the small-size combustion area; When the altitude increases or approaches the boundary of rotating detonation flameout, in order to maintain the stability of the rotating detonation wave in the combustion chamber, the fuel injection is moved forward and the center rocket 7 can be opened as appropriate.

[0078] The above shows and describes the basic principles, main features and advantages of the present invention. It should be understood by those skilled in the art that the present invention is not limited to the above embodiments, and the above embodiments and descriptions are only for explaining the principles of the present invention. Without departing from the spirit and scope of the present invention, the present invention may have various changes and improvements, which fall within the scope of the present invention to be protected. The scope of protection of the present invention is defined by the attached claims and their equivalents.

Claims

1. A rocket-based rotating detonation engine 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 direct injection section (4), an expansion-recess 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 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 ring jet injection points (4011), the central rocket (7) is provided with two groups of second ring jet injection points (7031), and 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 the support plate is provided with a group of wall injections (7023).

2. The rocket-based rotating detonation engine test simulator based on an adjustable thermal throat according to claim 1, characterized in that: 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 zone and the 180° phase zone are the fuel injection zone (602), the 30° phase zone is the thermal throat detonation zone (603), the 90° phase zone and the -90° phase zone are the third window zone (604), the -45° phase zone is the third dynamic pressure sensor installation zone (605), and the -135° phase zone is the third static pressure sensor installation zone (606).

3. 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.

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

5. 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).

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

7. 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 third annular injection points (5011); The concave cavity section (502) is provided with multiple sets of sensor interfaces and optical glass windows (8).

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

9. 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 8, characterized in that: The steps include: S1: After the direct injection section (4) and the expansion-concave cavity section (5) are inducted, the third group of first annular injection points (4011) of the direct injection section (4) and the two groups of thermal throat annular injections (601) of the thermal throat section (6) first inject fuel and ignite, forming rotating detonation combustion in the thermal throat section (6); S2: When the Mach number is Ma0-Ma2.0, the center rocket (7) is opened to provide thrust and form an ejection effect to inhale air. The combustion gas of the center 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 center 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 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 (7) is opened as appropriate.

10. Application of the rocket-based rotating detonation engine test simulation device based on adjustable thermal throat as described in any one of claims 1 to 8 in a rocket-based rotating detonation engine.

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