A continuous detonation turbine engine based on coaxial shear injection
Through the systematic design of the coaxial shear injection structure and fuel supply path, efficient pre-evaporation and uniform injection of liquid aviation kerosene are achieved, solving the stability problem of continuous detonation combustion under liquid kerosene conditions and ensuring stable propagation of detonation waves and uniform combustion.
Patent Information
- Authority / Receiving Office
- CN · China
- Patent Type
- Patents(China)
- Current Assignee / Owner
- PEKING UNIV
- Filing Date
- 2026-03-09
- Publication Date
- 2026-05-12
AI Technical Summary
The existing injection method of liquid aviation kerosene in aero engines makes it difficult to form a uniform combustible mixture in a very short time, resulting in unstable continuous detonation combustion. Furthermore, the insufficient evaporation of fuel leads to unstable detonation wave propagation, making it difficult to achieve stable continuous detonation combustion under liquid kerosene conditions.
It adopts a coaxial shear injection structure, and performs pre-evaporation treatment through thermal coupling between the fuel supply path and the flame tube wall. It forms a uniform combustible mixture layer in both the axial and circumferential directions at the inlet of the combustion zone, and achieves rapid mixing by utilizing the shearing action of high-speed gaseous fuel and compressed air.
This approach significantly shortens the evaporation and reaction preparation time of fuel in the main combustion zone, improves the uniformity of the mixture at the combustion zone inlet, ensures the stable propagation of detonation waves and the reliability of continuous detonation combustion, and provides an engineering-feasible structural implementation path.
Smart Images
Figure CN121782057B_ABST
Abstract
Description
Technical Field
[0001] This invention belongs to the field of continuous detonation engine technology, specifically relating to a continuous detonation turbine engine based on coaxial shear injection. Background Technology
[0002] Continuous detonation engines are a new type of power plant that utilizes the continuous propagation of detonation waves within the combustion chamber to organize the combustion process. Unlike the isobaric combustion method based on slow combustion in traditional aero engines, detonation combustion completes fuel heat release in a very short time through the strong coupling of shock waves and chemical reactions. Its thermodynamic process is closer to isochoric combustion, and theoretically it has higher cycle efficiency and total pressure gain potential.
[0003] Introducing continuous detonation combustion into the main combustion zone of aero-engines not only promises to overcome the inherent limitations of traditional slow-burning flame tubes in terms of thermodynamic efficiency, but also provides a new technical path for simplifying the overall engine structure and improving performance. However, under actual engineering conditions using liquid aviation kerosene as fuel, the establishment and maintenance of continuous detonation combustion face significant challenges. Aviation kerosene's low volatility, slow evaporation process, and long reaction induction time make it difficult to form a homogeneous combustible mixture with compressed air that meets the requirements for detonation propagation within the limited combustion chamber size and extremely short reaction time of aero-engines. These contradictions have become one of the key technical bottlenecks that urgently need to be overcome in the engineering application of continuous detonation combustion as the main combustion mode of aero-engines.
[0004] In traditional aero-engine combustion chambers, fuel supply and injection typically involve directly injecting liquid aviation kerosene into the combustion zone in liquid phase. The evaporation and mixing processes rely primarily on heat transfer and flow mixing within the main combustion zone. This method is suitable for combustion modes dominated by slow combustion. However, continuous detonation combustion places much stricter requirements on mixing uniformity and reaction induction time; the formation of the combustible mixture must be completed within an extremely short timescale. Due to the slow evaporation process of liquid aviation kerosene, the significant droplet size, and the evaporation lag effect, it is difficult to form a stable, homogeneous combustible mixture layer at the combustion zone inlet that meets the detonation establishment conditions, thus restricting the reliable initiation and sustained propagation of the detonation wave.
[0005] In existing aero-engine combustion chambers, liquid kerosene is typically atomized and injected directly into the combustion zone through radial or axial nozzles, relying on turbulent mixing and heat transfer to complete evaporation and combustion. This type of injection method can achieve stable combustion under slow combustion conditions, but under the extremely short reaction induction time and high mixing uniformity requirements of detonation combustion, it often leads to insufficient fuel evaporation, large fluctuations in local stoichiometry, difficulty in detonation establishment, or even instability.
[0006] Meanwhile, existing injection structures mostly employ radial nozzles, swirling nozzles, or asymmetric arrangements. Their mixing mechanisms primarily rely on large-scale vortex structures and unsteady recirculation regions, making it difficult to achieve a uniform mixing distribution around the combustion chamber. This mixing inhomogeneity is further amplified under continuous detonation combustion conditions, easily introducing equivalence ratio fluctuations and flow field inhomogeneities in the circumferential direction. This leads to problems such as unstable detonation wave propagation, propagation velocity fluctuations, mode jumps, and even localized detonation failure, severely impacting the stability and engineering feasibility of continuous detonation as the primary combustion mode for aero-engines. Traditional injection structures often use simple radial nozzles or swirling nozzles, whose injection scale and mixing mechanisms are primarily designed for slow combustion, making it difficult to form a uniformly mixed layer in the circumferential direction suitable for continuous detonation propagation under high-pressure, high-speed incoming flow conditions. These problems make it difficult for existing technologies to achieve stable continuous detonation combustion under liquid kerosene conditions without significantly increasing the incoming flow temperature or adding hydrogen or oxygen.
[0007] Furthermore, in existing technologies, the fuel supply path and the combustor wall are typically independent of each other. Before entering the combustion zone, the fuel fails to undergo effective thermal pretreatment using the high-temperature environment of the combustor wall, resulting in a high concentration of fuel evaporation and phase transition near the main combustion zone inlet. This strong coupling between the evaporation process and the detonation reaction process not only prolongs the reaction induction time but also introduces significant thermodynamic inhomogeneities at the detonation wave leading edge, weakening the self-sustaining capability of the detonation wave and hindering the stable establishment and propagation of continuous detonation waves.
[0008] Therefore, how to achieve efficient pre-evaporation and controllable injection of liquid kerosene within the limited space of the aero-engine flame tube without significantly increasing system complexity, and how to construct an injection-mixing structure that is conducive to continuous detonation establishment and stable propagation, has become a key technical problem that urgently needs to be solved in this field. Summary of the Invention
[0009] This invention provides a continuous detonation turbine engine based on coaxial shear injection. Through systematic design of the fuel supply path and injection structure, the engine ensures that aviation kerosene undergoes sufficient and controllable pre-evaporation before entering the main combustion zone, forming a combustible mixture layer with good axial and circumferential uniformity at the combustion zone inlet. By introducing a coaxial shear injection structure, the strong shearing action between high-speed gaseous fuel and compressed air achieves a rapid, efficient, and stable mixing process within a limited flame tube length. This satisfies the comprehensive requirements of continuous detonation combustion for mixing uniformity, reaction induction time, and combustion stability, creating favorable conditions for the establishment and stable operation of continuous detonation combustion.
[0010] To achieve the above objectives, the present invention adopts the following specific technical solution:
[0011] A continuous detonation turbine engine based on coaxial shear injection, the continuous detonation turbine engine includes an outer casing, a central shaft, an air intake, a tail nozzle, a compressor, a diffuser, a flame tube, a turbine, and fuel lines.
[0012] The central shaft is coaxially mounted within the outer casing; one end of the outer casing is fixedly connected to the air intake, and the other end is fixedly connected to the tail nozzle; the compressor, diffuser, flame tube, and turbine are arranged sequentially along the axial direction of the central shaft; the compressor is fixedly mounted on the central shaft and located at the connection between the air intake and the outer casing; the turbine includes a turbine stator and a turbine rotor fixedly mounted on the central shaft; the turbine rotor is located at the connection between the tail nozzle and the outer casing; the diffuser and turbine stator are both fixedly mounted on the inner wall of the outer casing; the central shaft is rotatably mounted on the diffuser and turbine stator via bearings and bushings.
[0013] The flame tube is fixedly connected between the diffuser and the turbine stator, and a coaxial shear injection structure is provided at the end facing the diffuser; the coaxial shear injection structure has a fuel manifold and a multi-row array of coaxial shear injection units; the fuel manifold is used for fuel distribution and temporary storage; each coaxial shear injection unit consists of a central air passage located in the center and an annular fuel jet slit located on the outer periphery of the outlet end of the central air passage;
[0014] The fuel pipeline includes a main fuel supply line, a fuel manifold, and multiple fuel distribution pipes attached to the outer wall of the flame tube, used to vaporize aviation kerosene through heat exchange; the fuel manifold is located on the outer periphery of the outlet end of the flame tube and is connected to the main fuel supply line and the fuel distribution pipes, used to distribute aviation kerosene from the main fuel supply line to the multiple fuel distribution pipes; the multiple fuel distribution pipes are distributed along the circumference of the flame tube, used to connect the fuel manifold and the fuel manifold chamber;
[0015] The annular fuel jet slit is connected to the fuel manifold, enabling the airflow injected from the central air channel and the gasified fuel injected from the annular fuel jet slit to achieve coaxial shear mixing.
[0016] Furthermore, the fuel manifold is formed by the outer cylinder of the flame tube, the inner cylinder of the flame tube, metal disc one, and metal disc two surrounding it;
[0017] The outer and inner cylinders of the flame tube are coaxially arranged between the outer casing and the central shaft; an outer bypass duct is formed between the outer cylinder and the outer casing; a continuous detonation main combustion chamber is formed between the outer and inner cylinders; an inner duct is formed on the inner side of the inner cylinder; compressed air from the diffuser outlet enters the outer bypass duct, the continuous detonation main combustion chamber, and the inner duct respectively; the air entering the continuous detonation main combustion chamber through the central air passage forms an axial high-speed central air jet, which is used for coaxial shear injection and participates in continuous detonation combustion;
[0018] Metal disc one and metal disc two are fixedly connected between the outer cylinder and the inner cylinder of the flame tube and are located at the end of the flame tube facing the diffuser; the central air passage runs through metal disc one and metal disc two; the annular fuel jet slit runs through metal disc two.
[0019] Furthermore, the central air channel is a stepped hole, and is formed by a metal disc and a tube integrally formed with the metal disc.
[0020] The tail end of the tube is inserted into the through hole of the second metal disk, forming an annular fuel jet gap between the tube and the through hole, and forming a shear mixing zone of gasified fuel and gas flow in the through hole outside the tail end of the tube.
[0021] Furthermore, the stepped orifice is formed by connecting a large orifice at the air inlet end and a small orifice at the air outlet end; the diameter of the large orifice is larger than the diameter of the small orifice.
[0022] Furthermore, both the outer and inner sections of the flame tube are equipped with multiple dilution and mixing holes that penetrate the wall thickness.
[0023] Furthermore, the fuel manifold is a ring-shaped pipe.
[0024] Furthermore, both the outer and inner cylinders of the flame tube are provided with multiple rows of dilution and mixing holes distributed circumferentially.
[0025] Furthermore, each row of dilution and mixing pores has a different pore size.
[0026] Furthermore, multiple fuel diversion tubes are evenly distributed along the circumference of the flame tube.
[0027] Furthermore, both the main fuel supply line and the fuel distribution pipe extend from the head to the tail along the axial direction of the flame tube.
[0028] Compared with the prior art, the continuous detonation turbine engine of the present invention has the following beneficial effects:
[0029] 1. Significantly improves the reaction preparation conditions of liquid aviation kerosene under continuous detonation conditions. In existing technologies, liquid aviation kerosene is typically preheated by external equipment to meet the extremely short reaction induction time required for continuous detonation in aero-engines, which is insufficient for engineering applications. This invention utilizes an evaporator-type flame tube structure to deeply couple the fuel supply path with the thermal environment of the flame tube's outer wall. During the axial and circumferential flow of the aviation kerosene along the flame tube, it undergoes sufficient heat exchange with the flame tube wall, resulting in a significant temperature rise and phase transition from liquid to gas before entering the injection unit. This creates a predominantly gaseous fuel state and allows for sufficient and controllable pre-evaporation before entering the main combustion zone. Furthermore, a combustible mixture with good axial and circumferential uniformity is formed at the combustion zone inlet, allowing the fuel to be introduced into the main combustion zone at the nozzle in gaseous or quasi-gase form, significantly reducing the difficulty of forming a detonation wave between aviation kerosene and compressed air. The continuous detonation turbine engine of the present invention significantly shortens the evaporation and reaction preparation time of fuel in the main combustion zone, thereby reducing the difficulty of detonation establishment under liquid kerosene conditions from the source.
[0030] 2. Effectively improves the spatial uniformity of the combustible mixture injection at the combustion zone inlet. The continuous detonation turbine engine of this invention, by arranging multiple rows of arrayed coaxial shear injection units at the head of the flame tube, achieves a high-flow-rate reactant supply while refining and uniformly distributing the injection scale circumferentially. The coaxial shearing action formed between the high-speed central air jet and the outer ring fuel vapor jet allows for efficient mixing of fuel and air near the nozzle outlet, thereby forming a combustible mixture layer at the combustion zone inlet, which helps improve the stability of the detonation mode.
[0031] 3. It provides an engineering-feasible structural path for continuous detonation as the main combustion mode in aero-engines. The continuous detonation turbine engine of this invention does not rely on significantly increasing the incoming flow temperature, adding hydrogen, or enriching oxygen to reduce the difficulty of detonation establishment. Instead, it systematically designs the fuel supply path and injection structure through the flame tube, solving the core engineering problem of achieving continuous detonation combustion under liquid aviation kerosene conditions while maintaining the overall architectural compatibility of the aero-engine. This provides an engineering-feasible structural path for the application of continuous detonation as the main combustion mode. Attached Figure Description
[0032] Figure 1 This is a perspective view of the continuous detonation turbine engine of the present invention;
[0033] Figure 2 This is a cross-sectional view of the continuous detonation turbine engine of the present invention;
[0034] Figure 3 for Figure 2 A partially enlarged structural diagram of part A in the middle;
[0035] Figure 4 A schematic diagram of the three-dimensional structure of the upstream side of the flame tube;
[0036] Figure 5 A three-dimensional structural diagram of the downstream side of the flame tube;
[0037] Figure 6 This is a cross-sectional view of the flame tube;
[0038] Figure 7 for Figure 6 A magnified schematic diagram of part B in the middle section.
[0039] Figure label:
[0040] 1-Outer casing, 2-Central shaft, 3-Intake duct, 4-Tail nozzle, 5-Compressor, 6-Diffuser, 7-Flame tube, 8-Turbine stator, 9-Turbine rotor, 10-Outer bypass duct, 11-Continuous detonation main combustion chamber, 12-Inner duct, 13-Shaft sleeve, 14-Coaxial shear injection unit, 101-Fuel manifold, 102-Central air passage, 103-Annular fuel jet slit, 104-Outer cylinder of flame tube, 105-Inner cylinder of flame tube, 106-Metal disc one, 107-Metal disc two, 108-Pipe body, 109-Shear mixing zone, 110-Dilution mixing hole, 201-Main fuel supply line, 202-Fuel manifold, 203-Fuel splitter tube. Detailed Implementation
[0041] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.
[0042] In the description of this invention, it should be understood that the terms "first," "second," etc., are used for descriptive purposes only and should not be construed as indicating or implying relative importance. Those skilled in the art can understand the specific meaning of these terms in this invention based on the specific circumstances. Furthermore, in the description of this invention, unless otherwise stated, "multiple" refers to two or more. "And / or" describes the relationship between related objects, indicating that three relationships can exist. For example, A and / or B can represent: A existing alone, A and B existing simultaneously, and B existing alone. The character " / " generally indicates that the preceding and following related objects have an "or" relationship.
[0043] This embodiment provides a continuous detonation turbine engine based on coaxial shear injection. This continuous detonation turbine engine uses a continuous detonation combustor as the main combustion unit of the aero-engine. Its overall configuration arranges the compressor 5, the continuous detonation combustor, and the turbine components sequentially along the engine's axial direction, forming a complete propulsion energy conversion system. The continuous detonation combustor is located between the compressor 5 and the turbine, used to achieve efficient detonation combustion during engine operation and output high-kinetic-energy gas to drive the downstream turbine.
[0044] like Figure 1 and Figure 2 As shown in the structure, the continuous detonation turbine engine includes an outer casing 1, a central shaft 2, an air intake 3, a tail nozzle 4, a compressor 5, a diffuser 6, a flame tube 7, a turbine, and fuel lines. The compressor 5 and turbine are mechanically connected via the central shaft 2, and the outer casing 1 provides structural support and overall enclosure for the internal functional components. It should be noted that, in this embodiment of the invention, [the following is omitted as it is not directly related to the preceding text]. Figure 2 The middle air intake 3 is upstream on one side and downstream on the tail nozzle 4 on the other side. At the same time, the head or front end of each component points towards the end of the air intake 3, and the tail or end of each component points towards the end of the tail nozzle 4.
[0045] like Figure 2 As shown, the central shaft 2 is coaxially arranged with the outer casing 1 and is located inside the outer casing 1. One end of the outer casing 1 is fixedly connected to the air intake 3, and the other end is fixedly connected to the tail nozzle 4. The compressor 5, diffuser 6, flame tube 7, and turbine are arranged sequentially along the axial direction of the central shaft 2. The compressor 5 is fixedly mounted on the central shaft 2 and rotates synchronously with the central shaft 2. The compressor 5 is located at the connection between the air intake 3 and the outer casing 1. The diffuser 6 is fixedly mounted on the inner wall of the outer casing 1. The turbine includes a turbine stator 8 fixedly mounted on the inner wall of the outer casing 1 and a turbine rotor 9 fixedly mounted on the central shaft 2. The turbine rotor 9 is located at the connection between the tail nozzle 4 and the outer casing 1. The central shaft 2 is rotatably mounted on the diffuser 6 and the turbine stator 8 via bearings and a bushing 13.
[0046] like Figure 3 , Figure 4 and Figure 5As shown, the flame tube 7 is fixedly connected between the diffuser 6 and the turbine stator 8, that is, one end of the flame tube 7 is fixedly connected to the diffuser 6, and the other end is fixedly connected to the turbine stator 8; a coaxial shear injection structure is provided at the end of the flame tube 7 facing the diffuser 6; the compressed air and fuel vapor entering the flame tube 7 to participate in combustion both enter the coaxial shear injection structure to participate in injection and mixing. The coaxial shear injection structure has a fuel manifold 101 and a multi-row array coaxial shear injection unit 14. In this embodiment, a three-row array coaxial shear injection unit 14 is used as an example for description. The fuel manifold 101 is used for fuel distribution and temporary storage; the aviation kerosene after pre-evaporation treatment enters the fuel manifold 101 mainly in gas phase form; each coaxial shear injection unit 14 consists of a central air channel 102 located in the center and an annular fuel jet slit 103 provided on the outer periphery of the outlet end of the central air channel 102. Figure 6 and Figure 7 As shown, the fuel manifold 101 is formed by an outer cylinder 104, an inner cylinder 105, a first metal disc 106, and a second metal disc 107 surrounding the flame tube. The outer cylinder 104 and the inner cylinder 105 are coaxially arranged between the outer casing 1 and the central shaft 2. The first metal disc 106 and the second metal disc 107 are fixedly connected between the outer cylinder 104 and the inner cylinder 105 and are located at the end of the flame tube 7 facing the diffuser 6. The central air passage 102 passes through the first metal disc 106 and the second metal disc 107. The annular fuel jet slit 103 passes through the second metal disc 107. Figure 3 As shown, an outer bypass duct 10 is formed between the outer tube 104 of the flame tube and the outer casing 1; a continuous detonation main combustion chamber 11 is formed between the outer tube 104 of the flame tube and the inner tube 105 of the flame tube; and an inner bypass duct 12 is formed inside the inner tube 105 of the flame tube. Figure 2 , Figure 3 and Figure 6 As shown, the compressed air from the diffuser 6 outlet enters the outer bypass duct 10, the continuous detonation main combustion chamber 11, and the inner bypass duct 12, respectively, with the airflow direction as follows: Figure 6As indicated by the middle arrow, compressed air from compressor 5 enters the outer bypass duct 10, the continuous detonation main combustion chamber 11, and the inner duct 12 of the flame tube 7. Air entering the continuous detonation main combustion chamber 11 participates in continuous detonation combustion. Air entering the continuous detonation main combustion chamber 11 through the central air channel 102 of the coaxial shear injection unit 14 forms an axial high-speed central air jet, providing the necessary airflow conditions for subsequent coaxial shear injection and the construction of the combustible mixture layer, and is used for coaxial shear injection and participation in continuous detonation combustion. The downstream sections of both the outer cylinder 104 and the inner cylinder 105 of the flame tube are provided with multiple dilution and mixing holes 110 penetrating the wall thickness, with each row of dilution and mixing holes 110 having a different aperture. The outer bypass duct 10 and the inner duct 12 together constitute an airflow channel for film cooling and cooling of the flame tube 7 wall surface and for mixing and cooling of the turbine inlet gas. At the outlet of the coaxial shear injection unit 14, a strong coaxial shearing effect is formed between the high-speed central airflow and the outer ring fuel vapor flow, which enables the fuel and air to be efficiently mixed and injected into the main combustion zone of the flame tube 7.
[0047] like Figure 5As shown, the fuel pipeline includes a main fuel supply line 201, a fuel manifold 202, and multiple fuel distribution pipes 203 attached to the outer wall of the flame tube 7, used to vaporize aviation kerosene through heat exchange; the main fuel supply line 201 serves as the inlet for aviation kerosene outside the engine and is connected to the fuel distribution structure inside the flame tube 7. The fuel manifold 202 is located on the outer periphery of the outlet end of the flame tube 7 and is connected to the main fuel supply line 201 and the fuel distribution pipes 203, used to distribute aviation kerosene from the main fuel supply line 201 to the multiple fuel distribution pipes 203; the multiple fuel distribution pipes 203 are distributed circumferentially along the flame tube 7, used to connect the fuel manifold 202 and the fuel manifold cavity 101. In this embodiment, the fuel manifold 202 is an annular pipe, and multiple fuel diversion caps 203 are evenly distributed along the circumference of the flame tube 7. Up to 12 fuel diversion caps 203 can be provided. Both the main fuel supply line 201 and the fuel diversion caps 203 extend from the head to the tail along the axial direction of the flame tube 7. The annular fuel jet slit 103 communicates with the fuel manifold 101, enabling the airflow injected from the central air channel 102 to coaxially shear and mix with the gasified fuel injected from the annular fuel jet slit 103. After entering through the main fuel supply line 201, the aviation kerosene flows through the fuel channels arranged axially and circumferentially along the outer cylinder 104 of the flame tube, continuously exchanging heat with the wall of the flame tube 7. During this process, the fuel absorbs heat from the wall of the flame tube 7, actively cooling the wall while simultaneously achieving a significant temperature increase and undergoing a phase transition, resulting in the fuel being in a predominantly gaseous or quasi-gase state before entering the injection unit. This fuel supply method directly serves the continuous detonation main combustion chamber 11, thereby introducing fuel into the detonation combustion zone in gaseous or quasi-gase form at the nozzle of the injection unit, significantly reducing the difficulty of the mixing and reaction conditions required for aviation kerosene and compressed air to establish a detonation wave. Figure 5 As shown, the main fuel supply line 201 is arranged axially from the head of the flame tube 7 and extends to the tail of the flame tube 7, connecting with the fuel manifold 202 located at the outlet of the flame tube 7. The fuel manifold 202 redistributes the aviation kerosene from the main fuel supply line 201 to multiple fuel distribution capillary pipes 203. The fuel distribution capillary pipes 203 are attached to the outer wall of the flame tube 7 and eventually reconnect to each injection unit at the head of the flame tube 7.
[0048] In the aforementioned continuous detonation turbine engines, such as Figure 7As shown, the central air passage 102 is a stepped orifice, formed by a metal disk 106 and a tube 108 integrally formed with the metal disk 106. The tail end of the tube 108 is inserted into the through hole of the metal disk 107, forming an annular fuel jet gap 103 between the tube 108 and the through hole, and forming a shear mixing region 109 of gasified fuel and airflow in the through hole outside the tail end of the tube 108. The stepped orifice is formed by connecting a large orifice at the air inlet end and a small orifice at the air outlet end; the diameter of the large orifice is larger than the diameter of the small orifice.
[0049] The working principle of the aforementioned continuous detonation turbine engine is as follows: Based on the stringent requirements of continuous detonation combustion for mixing uniformity, reaction induction time, and reactant mixing quality, the fuel supply path, pre-evaporation method, and injection structure are designed in a coordinated manner to ensure that most of the fuel undergoes a phase transition from liquid to gas before entering the main combustion zone. The pre-evaporated fuel vapor is introduced into the fuel manifold 101 of the coaxial shear injection structure. High-pressure air from the compressor 5, in addition to entering the inner duct 12 and outer bypass duct 10 of the flame tube 7, enters the central air passage 102 of the array-type coaxial shear injection unit 14 through the coaxial shear injection structure at the head of the flame tube 7. At the nozzle exit, the high-speed air jet in the center and the outer ring fuel vapor jet form a strong coaxial shearing action in the axial direction, rapidly mixing the fuel and air. This creates an injection and mixing structure with coaxial shear characteristics at the combustion zone inlet, providing the necessary initial conditions for the establishment and stable propagation of the detonation wave.
[0050] The aforementioned continuous detonation turbine engine has the following advantages in continuous detonation applications: First, it adopts an evaporator-type flame tube 7 structure, meaning that aviation kerosene first enters the fuel manifold 202 located at the tail of the combustion chamber through the main fuel supply line 201 of the flame tube 7, and is then distributed to multiple fuel distribution tubes 203 arranged axially along the outer cylinder 104 of the flame tube. This allows the fuel to continuously exchange heat with the flame tube 7 wall, which is in a high-temperature environment, during its transport, thus completing sufficient heating and pre-evaporation treatment. Therefore, the fuel is already in a predominantly gaseous state when it reaches the injection position at the head of the flame tube 7, and this pretreatment process significantly shortens the evaporation and reaction preparation time of the fuel in the main combustion zone. Second, it is equipped with a multi-row array coaxial shear injection unit 14, which, while achieving a large flow rate of reactant supply, significantly improves the spatial uniformity of the injected reactants compared to traditional kerosene nozzles with larger volume and fewer arrangement. Finally, this coaxial shear injection structure enhances the mixing process of gaseous fuel and air through high-speed shearing, enabling the combustible mixture to achieve high mixing uniformity, thereby meeting the stringent requirements of continuous detonation waves for mixing uniformity and reaction induction time.
[0051] Therefore, the aforementioned continuous detonation turbine engine achieves efficient pretreatment and stable injection of liquid aviation kerosene in continuous detonation main combustion applications, providing key structural and mechanistic support for the engineering application of continuous detonation combustion as the main combustion mode of aero engines.
[0052] Obviously, those skilled in the art can make various modifications and variations to the embodiments of the present invention without departing from the spirit and scope of the invention. Therefore, if these modifications and variations fall within the scope of the claims of the present invention and their equivalents, the present invention also intends to include these modifications and variations.
[0053] In summary, the above are merely preferred embodiments of the present invention and are not intended to limit the scope of protection of the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.
Claims
1. A continuous detonation turbine engine based on coaxial shear injection, characterized in that, It includes the outer casing, central shaft, air intake, tail nozzle, compressor, diffuser, flame tube, turbine, and fuel lines; The central shaft is coaxially mounted within the outer casing; one end of the outer casing is fixedly connected to the air intake, and the other end is fixedly connected to the tail nozzle; the compressor, diffuser, flame tube, and turbine are arranged sequentially along the axial direction of the central shaft; the compressor is fixedly mounted on the central shaft and located at the connection between the air intake and the outer casing; the turbine includes a turbine stator and a turbine rotor fixedly mounted on the central shaft; the turbine rotor is located at the connection between the tail nozzle and the outer casing; the diffuser and turbine stator are both fixedly mounted on the inner wall of the outer casing; the central shaft is rotatably mounted on the diffuser and turbine stator via bearings and bushings. The flame tube is fixedly connected between the diffuser and the turbine stator, and a coaxial shear injection structure is provided at the end facing the diffuser; the coaxial shear injection structure has a fuel manifold and a multi-row array of coaxial shear injection units; the fuel manifold is used for fuel distribution and temporary storage; each coaxial shear injection unit consists of a central air passage located in the center and an annular fuel jet slit located on the outer periphery of the outlet end of the central air passage; The fuel pipeline includes a main fuel supply line, a fuel manifold, and multiple fuel distribution pipes attached to the outer wall of the flame tube, used to vaporize aviation kerosene through heat exchange; the fuel manifold is located on the outer periphery of the outlet end of the flame tube and is connected to the main fuel supply line and the fuel distribution pipes, used to distribute aviation kerosene from the main fuel supply line to the multiple fuel distribution pipes. Multiple fuel diversion tubes are distributed circumferentially along the flame tube to connect the fuel manifold and the fuel manifold cavity; The annular fuel jet slit is connected to the fuel manifold, enabling the airflow injected from the central air channel and the gasified fuel injected from the annular fuel jet slit to achieve coaxial shear mixing.
2. The continuous detonation turbine engine as described in claim 1, characterized in that, The fuel manifold is formed by the outer cylinder of the flame tube, the inner cylinder of the flame tube, metal disk one, and metal disk two surrounding it; The outer and inner cylinders of the flame tube are coaxially arranged between the outer casing and the central shaft; an outer bypass duct is formed between the outer cylinder and the outer casing; a continuous detonation main combustion chamber is formed between the outer and inner cylinders; an inner duct is formed on the inner side of the inner cylinder; compressed air from the diffuser outlet enters the outer bypass duct, the continuous detonation main combustion chamber, and the inner duct respectively; the air entering the continuous detonation main combustion chamber through the central air passage forms an axial high-speed central air jet, which is used for coaxial shear injection and participates in continuous detonation combustion; Metal disc one and metal disc two are fixedly connected between the outer cylinder and the inner cylinder of the flame tube and are located at the end of the flame tube facing the diffuser; the central air passage runs through metal disc one and metal disc two; the annular fuel jet slit runs through metal disc two.
3. The continuous detonation turbine engine as described in claim 2, characterized in that, The central air channel is a stepped hole, and is formed by a metal disc and a tube integrally formed with the metal disc. The tail end of the tube is inserted into the through hole of the second metal disk, forming an annular fuel jet gap between the tube and the through hole, and forming a shear mixing zone of gasified fuel and gas flow in the through hole outside the tail end of the tube.
4. The continuous detonation turbine engine as described in claim 3, characterized in that, The stepped orifice is formed by connecting a large orifice at the air inlet and a small orifice at the air outlet; the diameter of the large orifice is larger than the diameter of the small orifice.
5. The continuous detonation turbine engine as described in claim 2, characterized in that, The fuel manifold is a ring-shaped pipe.
6. The continuous detonation turbine engine as described in any one of claims 2-5, characterized in that, Both the outer and inner cylinders of the flame tube are provided with multiple rows of dilution and mixing holes distributed circumferentially.
7. The continuous detonation turbine engine as described in claim 6, characterized in that, Each row of dilution and mixing holes has a different pore size.
8. The continuous detonation turbine engine as described in claim 6, characterized in that, Multiple fuel diversion tubes are evenly distributed along the circumference of the flame tube.
9. The continuous detonation turbine engine as described in claim 8, characterized in that, Both the main fuel supply line and the fuel distribution pipe extend from the head to the tail along the axial direction of the flame tube.