Supersonic turbine engine based on detonation combustion and control method thereof

Abstract

The invention discloses a supersonic turbine engine based on detonation combustion and a control method thereof, a detonation combustion chamber is installed behind a main combustion chamber, the detonation combustion chamber, an engine turbine stator casing and a main spray pipe form a parallel structure, the detonation combustion chamber is nested outside the turbine casing and the main spray pipe, and the tail end of the detonation combustion chamber is provided with a detonation exhaust nozzle. During detonation combustion, the main jet pipe and the detonation exhaust jet pipe exhaust gas separately. High-pressure air flow used for organizing detonation combustion is led from two strands of air flow in a main combustion chamber of the engine. A vent hole is formed in the rear end of an outer casing of the main combustion chamber, and the flow of high-pressure air flowing into the detonation combustion chamber for organizing combustion is controlled by adjusting a vent valve. On the basis of a conventional turbine engine with mature performance, the engine is combined with the detonation combustion chamber in parallel, so that a novel detonation turbine combined engine structure is formed. Through the combined action of detonation combustion exhaust and main combustion exhaust of the engine, the high-altitude supersonic propulsion performance of the engine is greatly improved.

Description

Technical Field

[0001] The present invention belongs to the technical field of supersonic propulsion of gas turbine engines, and relates to a supersonic turbine engine based on continuous detonation combustion, and in particular to a supersonic power device and a control method thereof that combines continuous detonation combustion exhaust gas with turbine-based engine exhaust gas for propulsion. Background Art

[0002] With the development of hypersonic aircraft, the existing turbine-based power plant can no longer meet the needs of hypersonic flight technology. Since turbine engines with rotating parts are difficult to use at Mach numbers above 2.5, and the minimum operating Mach number of the existing ramjet engine is above 3.0, it faces difficulties such as difficulty in starting at low speed and cannot be used as a single propulsion system. Therefore, the turbine-based combined cycle engine may encounter a "thrust trap" in the range of Mach numbers 2.5 to 3.0, that is, the propulsion system cannot provide sufficient thrust in this speed range, thereby limiting the acceleration performance and flight efficiency of the aircraft.

[0003] As a self-pressurized high-efficiency combustion device, the Continuously Rotating Detonation Engine (CRDE) works on the principle of forming periodic detonation waves in an annular or tubular combustion chamber, and realizing rapid combustion and energy release of fuel through the compression of detonation waves. It has the characteristics of high combustion thermodynamic cycle efficiency, fast heat release, low fuel consumption rate, and high propulsion performance. It can start intervention work on the ground or at a lower Mach number, and its unit area thrust is higher than that of turbine engines and ramjet engines in the range of Mach numbers 2 to 3, which helps to eliminate the "thrust trap" that occurs in turbine-based combined cycle engines during mode conversion, and ultimately can provide continuous and stable full-speed acceleration capabilities for near-space high-speed aircraft in the range of Mach numbers 0 to 4. Compared with traditional turbine engines, the new propulsion technology scheme of continuous detonation + turbine combination is theoretically expected to increase the unit area thrust of the engine by 15% to 20% and reduce the fuel consumption rate by 10% to 15%, thereby greatly improving the high-speed propulsion performance of the engine.

[0004] Although continuous detonation engines have many advantages in theory, the existing continuous detonation turbine engine design schemes still have many technical problems in practical applications. For example, the existing continuous detonation turbine engine schemes mainly use detonation combustion chambers to directly replace the main combustion chambers, without considering the interaction between rotating parts and the detonation combustion chambers; in this scheme, the detonation combustion chamber is located between the compressor and the turbine, and the detonation wave propagates periodically in the combustion chamber. The flow field is not uniform in time and space, which will have an adverse effect on the operation of the compressor and turbine, such as blade vibration, flow instability and other problems, thereby reducing the overall performance and reliability of the engine. In short, the existing schemes for continuous detonation turbine engines have low feasibility and are difficult to implement in engineering.

[0005] In summary, although continuous detonation engine technology has significant advantages in theory, how to effectively combine it with traditional turbine engines while avoiding the adverse effects of detonation combustion on the working stability of the engine core is a technical problem that needs to be solved urgently. Solving this problem is of great significance for improving the high-speed propulsion performance of the engine, expanding its working envelope, and achieving hypersonic flight. Summary of the invention

[0006] (I) Purpose of the invention

[0007] The present invention aims at the above defects and deficiencies of the existing continuous detonation turbine engine, such as the detonation combustion chamber directly replacing the main combustion chamber, resulting in unstable operation of the core engine, narrow detonation combustion working envelope, performance improvement and application scope, etc., and in order to solve at least one of the above and other technical problems in the prior art, a new detonation + turbine combination engine and its control method are proposed. The present invention is based on a conventional turbine engine with mature performance, and is combined in parallel with a detonation combustion chamber to form a new detonation turbine combination engine configuration. Through the separate exhaust mode of continuous detonation combustion exhaust and engine main combustion exhaust and the joint action of the two, the adverse effects of detonation combustion on the engine core engine are effectively avoided, the detonation combustion working envelope of the engine is widened, and the "thrust trap" problem that may occur in the turbine-based combined cycle engine in the Mach number range of 2.5 to 3.0 is effectively solved, and finally the high-altitude supersonic propulsion performance of the engine is greatly improved.

[0008] (II) Technical solution

[0009] In order to achieve the above purpose and solve the technical problems, the present invention provides the following technical solutions:

[0010] The first invention object of the present invention is to provide a supersonic turbine engine based on detonation combustion, which is used to improve the high-altitude supersonic propulsion performance of the engine, comprising an engine inlet duct and an inner duct and an outer duct which are axially arranged downstream of the engine inlet duct and are concentrically arranged, wherein at least a compressor rotor and a compressor stator are arranged in the engine inlet duct, wherein:

[0011] At least one main combustion chamber is provided in the upstream section of the inner duct, at least a turbine rotor and a turbine stator are provided downstream of the main combustion chamber, a main tail nozzle is provided at the end of the downstream section of the inner duct and forms an exhaust passage containing the main combustion gas flow, and the compressor rotor, compressor stator, main combustion chamber, turbine rotor and turbine stator constitute the main components of the turbine engine core machine;

[0012] The upstream section of the outer duct is formed into a two-stream airflow outlet channel, the two-stream airflow outlet channel is located outside the outer ring of the main combustion chamber, and is used to lead out the high-temperature and high-pressure air participating in the detonation combustion from the engine intake duct, and the downstream of the two-stream airflow outlet channel extends at least to the rear end of the main combustion chamber in the axial direction, and a bleed valve for adjusting the flow rate of the lead-out gas is provided at the end thereof;

[0013] The midstream section of the outer duct is formed into a two-stream airflow introduction channel, which is used to introduce two airflows participating in detonation combustion from the two-stream airflow outlet channel, and a check valve structure is provided at the downstream end of the two-stream airflow introduction channel, which is used to prevent the detonation pressure wave at the moment of detonation initiation from being transmitted back to the upstream;

[0014] The downstream section of the outer duct is formed into a detonation combustion chamber, and the upstream part of the detonation combustion chamber is provided with at least one oil collecting ring and a high-energy electric ignition device in sequence along the axial direction, and the downstream part of the detonation combustion chamber is formed into a detonation combustion space, wherein the oil collecting ring is used to atomize the fuel and spray it into the detonation combustion chamber, and the high-energy electric ignition device is used to ignite the atomized fuel and realize detonation initiation, and the detonation combustion space is used to accommodate and accelerate the combustion gas and has a geometric configuration suitable for the propagation of detonation waves, and a detonation tail nozzle is installed in series at the end of the detonation combustion chamber, and the throat area of ​​the detonation tail nozzle is adjustable to meet the expansion and discharge requirements of the detonation gas flow under different flight conditions.

[0015] The second object of the present invention is to provide a control method for the above-mentioned supersonic turbine engine based on detonation combustion, wherein the control method comprises at least the following steps when implemented:

[0016] SS1. Start the engine and adjust the air flow distribution

[0017] Initialize the working status of the compressor rotor and stator through the engine control unit, start the turbine engine core, ensure that the engine enters the normal working mode, gradually increase the engine speed to the preset value, and monitor the working parameters of various engine components, including but not limited to the compressor outlet temperature, pressure and turbine inlet temperature;

[0018] SS2. Preparation for switching to detonation combustion working mode

[0019] Monitor the flight parameters in real time. When the flight Mach number exceeds a preset threshold and the various working parameters of the turbine engine core are stable, the engine control unit issues a detonation combustion working mode switching instruction, gradually adjusts the working state of the turbine engine core, and simultaneously opens the bleed valve and adjusts its opening, gradually introduces two air flows from the outer ring of the main combustion chamber into the detonation combustion chamber, and ensures that the air flow in the detonation combustion chamber can meet the detonation combustion initiation conditions;

[0020] SS3. Adjust the detonation tail nozzle throat area

[0021] Before entering the detonation combustion working mode, the engine control unit adjusts the throat area of ​​the detonation tail nozzle in advance according to at least the current flight altitude, flight speed, external air pressure and cavity pressure conditions of the detonation combustion chamber of the aircraft to ensure that the cavity pressure of the detonation combustion chamber can meet the detonation combustion initiation conditions;

[0022] SS4. Start the detonation combustion process

[0023] When the airflow enters the detonation combustion chamber and the airflow rate and chamber pressure reach the detonation conditions, the engine control unit controls the detonation combustion fuel supply device to supply fuel into the oil collecting ring. When the fuel supply flow rate and fuel supply pressure reach the preset level, the high-energy electric ignition device is started to ignite the fuel atomized by the oil collecting ring and sprayed into the detonation combustion chamber to achieve detonation initiation ignition, and monitor the pressure wave changes in the detonation combustion process in real time to ensure that the detonation combustion can continue, and dynamically adjust the fuel injection amount to maintain the stability of the detonation combustion, and monitor the working state of the check valve structure in real time to ensure that it can effectively prevent the detonation pressure wave from propagating upstream, and reduce the adverse effects of the detonation pressure wave at the moment of detonation initiation on the upstream components of the engine;

[0024] SS5. Feedback control and dynamic regulation of knock combustion

[0025] During the detonation combustion process, the engine control unit monitors the indoor cavity pressure of the detonation combustion chamber, the total inlet pressure of the detonation tail nozzle and the local static pressure at the flight altitude of the aircraft in real time, adopts a closed-loop control strategy and combines the flight condition data to adjust the opening of the bleed valve and the outlet area and throat area ratio of the detonation tail nozzle in real time to ensure that the expansion and discharge of the detonation gas flow are always in the optimal state.

[0026] (III) Technical Effect

[0027] Compared with the prior art, the supersonic turbine engine based on detonation combustion and the control method thereof of the present invention have the following technical advantages and beneficial effects:

[0028] (1) The present invention abandons the previous solution of directly replacing the main combustion chamber with a detonation combustion chamber in a detonation turbine engine, thereby ensuring that the high-speed and stable propulsion of the detonation combustion and turbine base combination is achieved without affecting the working stability of the engine core. The detonation combustion chamber of the present invention is installed at the rear end of the main combustion chamber, and the detonation tail nozzle is installed in series behind the detonation combustion chamber. The detonation combustion chamber and the detonation tail nozzle form a parallel structure with the turbine stator casing and the main tail nozzle. When the engine enters the detonation combustion working mode, the outer detonation airflow and the inner main combustion gas flow are each exhausted separately through independent nozzles, which can greatly avoid the adverse effects of the detonation airflow on the main combustion gas flow and the turbine, thereby ensuring the working stability of the engine under the detonation combustion state.

[0029] (2) In the present invention, the high-temperature and high-pressure air participating in the detonation combustion in the detonation combustion chamber is introduced into two air flows outside the outer ring of the main combustion chamber, and the flow rate of the introduced gas is adjusted by the bleed valve to ensure that the detonation combustion of the engine can be successfully initiated under different flight conditions, thereby widening the detonation combustion working envelope of the engine.

[0030] (3) The present invention sets a check valve structure upstream of the detonation combustion chamber, which can effectively prevent the detonation pressure wave at the moment of detonation initiation from being transmitted back to the upstream of the compressor, thereby preventing the compressor components from entering an unstable working state when the engine starts the detonation combustion working mode.

[0031] (4) In view of the "thrust trap" that may occur in current turbine-based combined cycle engines within the Mach number range of 2.5 to 3.0, the present invention provides a technical solution for a supersonic turbine engine based on continuous detonation combustion, which can provide continuous and stable full-speed acceleration capability for near-space high-speed aircraft within the Mach number range of 0 to 4. Compared with the current pre-research or under-research continuous detonation + turbine-based combined cycle engines, a different new configuration solution is adopted. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] Figure 1 It is a schematic diagram of the overall configuration of the supersonic turbine engine based on detonation combustion of the present invention;

[0033] Figure 2 It is a partial structural schematic diagram of the hot end component of the engine of the present invention;

[0034] Figure 3 It is a schematic diagram of the control method of the supersonic turbine engine based on detonation combustion of the present invention.

[0035] Description of reference numerals:

[0036] 1- compressor rotor, 2- compressor stator (including stator blades and casing), 3- main combustion chamber, 4- main combustion chamber casing, 5- bleed valve, 6- turbine stator (including guide vane and turbine casing), 7- turbine rotor, 8- check valve structure, 9- oil collecting ring, 10- high energy electric ignition device, 11- detonation combustion chamber, 12- detonation tail nozzle throat area adjustment mechanism, 13- detonation tail nozzle, 14- main tail nozzle, 15- heat insulation device, 16- detonation combustion space. DETAILED DESCRIPTION

[0037] In order to make the purpose, technical scheme and advantages of the implementation of the present invention clearer, the technical scheme in the embodiment of the present invention will be described in more detail below in conjunction with the drawings in the embodiment of the present invention. In the drawings, the same or similar reference numerals throughout represent the same or similar elements or elements with the same or similar functions. The described embodiments are part of the embodiments of the present invention, not all of the embodiments, and are intended to be used to explain the present invention, but should not be understood as limitations on the present invention. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0038] The present invention aims to provide a supersonic turbine engine based on detonation combustion and a control method thereof, which forms a new detonation turbine combined engine configuration by combining a conventional turbine engine with a detonation combustion chamber in parallel on the basis of a mature performance. Through the separate exhaust mode of continuous detonation combustion exhaust and engine main combustion exhaust and the joint action of the two, the adverse effects of detonation combustion on the engine core are effectively avoided, the detonation combustion working envelope of the engine is widened, and the "thrust trap" problem that may occur in a turbine-based combined cycle engine within the Mach number range of 2.5 to 3.0 is effectively solved.

[0039] Example 1

[0040] As a specific example, Figure 1 , 2As shown, the supersonic turbine engine based on continuous detonation combustion of the present invention comprises a compressor rotor 1, a compressor stator 2, a main combustion chamber 3, a main combustion chamber casing 4, a turbine rotor 7, a turbine stator 6, a main tail nozzle 14, a detonation combustion chamber 11, a detonation combustion space 16, a detonation tail nozzle 13, a venting valve 5, an oil collecting ring 9, a check valve structure 8, a high-energy electric ignition device 10, a detonation tail nozzle throat area adjustment mechanism 12, a heat insulation device 15, etc. The detonation combustion chamber 11 is installed at the rear end of the main combustion chamber 3, the turbine stator 6 and the main tail nozzle 14 are nested inside the detonation combustion chamber 11, forming a parallel structure with the detonation combustion chamber 11, and the detonation tail nozzle 13 is installed in series after the detonation combustion chamber 11.

[0041] Specifically, in the supersonic turbine engine based on continuous detonation combustion of the present invention, a compressor rotor 1 and a compressor stator 2 are provided in the engine inlet duct, and an inner duct and an outer duct are concentrically arranged downstream of the engine inlet duct, wherein: a main combustion chamber 3 is provided in the upstream section of the inner duct, and at least a turbine rotor 7 and a turbine stator 6 are provided downstream of the main combustion chamber 3, a main tail nozzle 14 is provided at the end of the downstream section of the inner duct and forms an exhaust channel containing the main combustion gas flow, and the compressor rotor 1, the compressor stator 2, the main combustion chamber 3, the turbine rotor 7 and the turbine stator 6 constitute the main components of the turbine engine core.

[0042] The upstream section of the duct is formed into two airflow outlet channels, which are located outside the outer ring of the main combustion chamber 3 and are used to lead out high-temperature and high-pressure air participating in detonation combustion from the engine intake duct, and the downstream of the two airflow outlet channels extends at least to the rear end of the main combustion chamber 3 in the axial direction, and a bleed valve 5 for adjusting the flow rate of the lead-out gas is provided at its end. The midstream section of the duct is formed into two airflow introduction channels, which are used to introduce two airflows participating in detonation combustion from the two airflow outlet channels, and a check valve structure 8 is provided at the downstream end of the two airflow introduction channels to prevent the detonation pressure wave at the moment of detonation initiation from being transmitted back upstream. The downstream section of the outer duct is formed as a detonation combustion chamber 11, and the upstream part of the detonation combustion chamber 11 is provided with at least an oil collecting ring 9 and a high-energy electric ignition device 10 in sequence along the axial direction, and the downstream part of the detonation combustion chamber 11 is formed as a detonation combustion space 16, wherein the oil collecting ring 9 is used to atomize the fuel and spray it into the detonation combustion chamber 11, and the high-energy electric ignition device 10 is used to ignite the atomized fuel and realize detonation initiation, and the detonation combustion space 16 is used to accommodate and accelerate the combustion gas and has a geometric configuration suitable for the propagation of detonation waves, and a detonation tail nozzle 13 is installed in series at the end of the detonation combustion chamber 11, and the throat area of ​​the detonation tail nozzle 13 is adjustable to meet the expansion and discharge requirements of the detonation gas flow under different flight conditions.

[0043] The working principle of the supersonic turbine engine based on continuous detonation combustion of the present invention is as follows:

[0044] When the aircraft flies at a low Mach number (Ma=0-2.0), the bleed valve 5 at the inlet of the engine detonation combustion chamber 11 remains completely closed. At this time, the engine is used as a conventional turbojet engine. The air entering the engine is compressed by the compressor rotor 1 and flows only to the main combustion chamber 3. After fuel injection and combustion, it drives the turbine rotor 7 and is finally discharged from the main tail nozzle 14. The two air flows outside the combustion chamber serve as the cooling air flow and the mixed air flow of the main combustion chamber 3, and are finally discharged from the main tail nozzle 14. At this time, the thrust of the engine is only generated by the high-speed exhaust of the main tail nozzle.

[0045] When the aircraft needs to accelerate to a high Mach number (Ma=2.0-4.0) for flight, the engine detonation combustion chamber 11 needs to intervene in the work. At this time, the engine control program will first control the bleed valve 5 to open. The opening of the bleed valve 5 is based on the intake flow required for the engine to successfully detonate the detonation combustion under the flight conditions at that time. The control program can steplessly control the opening of the bleed valve 5. Then, the engine control program will adjust the throat area of ​​the detonation tail nozzle 13 in advance according to the flight conditions by controlling the throat area adjustment mechanism 12 of the detonation tail nozzle to obtain the appropriate detonation combustion detonation chamber pressure in the detonation combustion chamber 11. When the engine control program detects that the chamber pressure in the detonation combustion chamber 11 meets the detonation conditions, the control program will control the detonation combustion fuel supply device to supply fuel to the oil collecting ring 9. When the detonation combustion chamber fuel supply flow and fuel supply pressure reach a certain level, the control program will control the high-energy electric ignition device 10 to start detonation detonation ignition. After the detonation is successfully initiated, the engine control program will use a closed-loop control strategy to adjust the detonation tail nozzle 13 outlet area and throat area ratio in real time based on the monitored total pressure at the inlet of the detonation tail nozzle 13 (approximately equal to the detonation combustion chamber pressure) and the local static pressure at the flight altitude of the engine, so as to ensure the optimal working state of the detonation tail nozzle 13, so that the detonation gas flow reaches a fully expanded state in the detonation tail nozzle 13, thereby increasing the exhaust speed and improving the engine's external detonation thrust. At this time, the total thrust of the engine is jointly generated by the detonation gas exhaust flowing through the detonation tail nozzle 13 and the main combustion exhaust flowing through the main nozzle 14.

[0046] In some preferred embodiments, the high-temperature and high-pressure air entering the detonation combustion chamber 11 and participating in the detonation combustion is introduced into two air flows outside the outer ring of the main combustion chamber 3, and the flow rate of the introduced gas is adjusted by the bleed valve 5. The bleed air flow rate range is preferably 8% to 15% of the total engine intake flow rate, and the bleed valve 5 is communicated with the engine control unit, and the opening of the bleed valve 5 is adjusted according to the real-time flight status and the pressure in the detonation combustion chamber to ensure a stable supply of high-temperature and high-pressure air under different flight conditions.

[0047] In some preferred embodiments, the detonation tail nozzle 13 is a convergent-divergent tail nozzle with adjustable throat area, and its outlet area and throat area ratio are adjusted in real time at least according to the flight conditions and the working state parameters of the detonation combustion chamber to adapt to the acceleration and expansion characteristics of the supersonic exhaust gas flow. In addition, the throat area of ​​the detonation tail nozzle 13 is adjusted by the detonation tail nozzle throat area adjustment mechanism 12 installed on the outer wall of the detonation combustion chamber. When the aircraft is flying at high speed, the outlet area and throat area ratio of the detonation tail nozzle 13 are adjusted to ensure the optimal working state of the detonation tail nozzle, thereby increasing the exhaust speed and improving the detonation thrust of the engine.

[0048] In some preferred embodiments, when the engine enters the detonation combustion working mode, the exhaust mode is separate exhaust, that is, the main combustion gas flow flowing through the engine turbine rotor 7 is discharged through the main tail nozzle 14; the secondary combustion gas flow after detonation combustion in the detonation combustion chamber 11 is discharged through the detonation tail nozzle 13. Since the exhaust is separate, the two exhaust flows do not affect each other, but work together to generate the thrust of the engine. The total thrust gain of the engine generated by the detonation combustion exhaust and the main combustion exhaust is about 10% to 15%.

[0049] In some preferred embodiments, in order to avoid the high gas temperature in the detonation combustion chamber 11 causing thermal damage to the main tail nozzle 14 of the engine during continuous detonation combustion, a heat insulation device 15 is filled between the detonation combustion chamber 11 / detonation tail nozzle 13 and the main tail nozzle 14 to reduce the heat conduction of the two gas flows, thereby ensuring that the detonation combustion chamber does not cause thermal damage to the downstream section of the inner channel and the main tail nozzle when working. At the same time, in order to reduce the wall temperature of the outer wall of the detonation combustion chamber 11, a zirconium oxide thermal insulation coating is sprayed on the inner tube of the detonation combustion chamber 11 to improve the thermal shock resistance of the detonation combustion chamber in a high temperature environment, reduce thermal fatigue damage and extend the service life of the detonation combustion chamber, and at the same time reduce the temperature of the outer wall of the detonation combustion chamber, thereby effectively reducing infrared radiation and improving the stealth performance of the aircraft.

[0050] In some preferred embodiments, in order to prevent the detonation pressure wave from being transmitted back to the upstream of the engine through the vent valve 5 at the moment of detonation, a check valve structure 8 is arranged upstream of the detonation combustion chamber 11. The check valve structure 8 is preferably a double-layer plate design, which has the function of adaptively adjusting the valve opening, and can automatically adjust the valve opening degree according to the intensity of the detonation pressure wave. Under the condition of not affecting the normal air intake of the detonation combustion chamber, the detonation pressure wave at the moment of detonation detonation can effectively reduce the adverse effect of the detonation pressure wave on the operation of the compressor rotor 1 component upstream of the engine, and avoid the compressor from entering an unstable working state such as surging when the engine starts detonation combustion. The check valve structure 8 can adaptively adjust the valve opening size according to the pressure difference between the front and back of the valve.

[0051] In some preferred embodiments, the fuel required for detonation combustion is sprayed from the oil collecting ring 9, which is installed upstream of the detonation combustion chamber 11 and downstream of the check valve structure 8. The oil collecting ring 9 has 20 injection holes evenly distributed around the entire circumference, and the diameter of the injection holes is 0.5mm to 1.0mm. The injection angle is inclined toward the downstream side of the airflow and inward at an acute angle, for example, it can be at an angle of 25° with the engine axis. During detonation initiation and combustion, the fuel is sprayed from the injection holes of the oil collecting ring 9 to form atomized fuel. The atomized fuel is ignited by the high-energy electric ignition device 10 installed in the injection area of ​​the oil collecting ring 9, thereby achieving detonation initiation. In addition, the high-energy electric ignition device 10 preferably adopts a redundant design and is equipped with at least two sets of ignition electrodes. When one set of electrodes fails to work due to a fault or the service life is exhausted, the other set of electrodes can immediately take over the work to ensure the continuous reliability of the detonation combustion process and avoid the failure of detonation combustion or the degradation of engine performance due to a single point failure.

[0052] Example 2

[0053] Based on the above embodiment 1, as a further supplement and refinement, this embodiment provides a control method for the above-mentioned supersonic turbine engine based on detonation combustion, which at least includes the following steps:

[0054] SS1. Start the engine and adjust the air flow distribution

[0055] Initialize the working status of the compressor rotor and stator through the engine control unit, start the turbine engine core, ensure that the engine enters the normal working mode, gradually increase the engine speed to the preset value, and monitor the working parameters of various engine components, including but not limited to the compressor outlet temperature, pressure and turbine inlet temperature;

[0056] SS2. Preparation for switching to detonation combustion working mode

[0057] Monitor flight parameters in real time. When the flight Mach number exceeds the preset threshold (range of 2.0 to 4.0) and the various operating parameters of the turbine engine core are stable, the engine control unit issues a detonation combustion working mode switching command, gradually adjusts the working state of the turbine engine core, and opens the bleed valve and adjusts its opening, gradually introducing two air flows from the outer ring of the main combustion chamber into the detonation combustion chamber to ensure that the air flow in the detonation combustion chamber can meet the detonation combustion initiation conditions.

[0058] SS3. Adjust the detonation tail nozzle throat area

[0059] Before entering the detonation combustion working mode, the engine control unit adjusts the throat area of ​​the detonation tail nozzle in advance according to at least the current flight altitude, flight speed, external air pressure and cavity pressure conditions of the detonation combustion chamber of the aircraft to ensure that the cavity pressure of the detonation combustion chamber can meet the detonation combustion initiation conditions;

[0060] SS4. Start the detonation combustion process

[0061] When the airflow enters the detonation combustion chamber and the airflow rate and chamber pressure reach the detonation conditions, the engine control unit controls the detonation combustion fuel supply device to supply fuel into the oil collecting ring. When the fuel supply flow rate and fuel supply pressure reach the preset level, the high-energy electric ignition device is started to ignite the fuel atomized by the oil collecting ring and sprayed into the detonation combustion chamber to achieve detonation initiation ignition, and monitor the pressure wave changes in the detonation combustion process in real time to ensure that the detonation combustion can continue, and dynamically adjust the fuel injection amount to maintain the stability of the detonation combustion, and monitor the working state of the check valve structure in real time to ensure that it can effectively prevent the detonation pressure wave from propagating upstream, and reduce the adverse effects of the detonation pressure wave at the moment of detonation initiation on the upstream components of the engine;

[0062] SS5. Feedback control and dynamic regulation of knock combustion

[0063] During the detonation combustion process, the engine control unit monitors the indoor cavity pressure of the detonation combustion chamber, the total inlet pressure of the detonation tail nozzle and the local static pressure at the flight altitude of the aircraft in real time, adopts a closed-loop control strategy and combines the flight condition data to adjust the opening of the bleed valve and the outlet area and throat area ratio of the detonation tail nozzle in real time to ensure that the expansion and discharge of the detonation gas flow are always in the optimal state.

[0064] Preferably, the control method also includes step SS6 of switching between the detonation combustion mode and the conventional turbine mode. Specifically: during the flight, the engine control unit switches the engine operating mode according to the changes in the flight conditions. When the flight speed is lower than the preset Mach number, the detonation combustion mode is gradually shut down, and first the fuel injection amount is reduced, then the high-energy electric ignition device is turned off, and finally the air bleed valve is gradually closed to stop the introduction of the two airflows. The airflow only works through the main combustion chamber, so that the engine smoothly transitions to the conventional turbine mode; when the flight speed reaches the requirement of detonation combustion again, it switches back to the detonation combustion mode to maintain the best propulsion performance.

[0065] Preferably, the control method also includes step SS7 of performing fault detection and safety mechanism, specifically: during the entire control process, the operating status parameters of key engine components are continuously monitored and safety thresholds are set. When an abnormality is detected, the emergency handling procedure is immediately initiated to stop the detonation combustion, close the bleed valve, stop the fuel supply, and switch to the conventional turbine mode or execute the emergency shutdown procedure when necessary to ensure the safety of the engine and the aircraft.

[0066] Through the above embodiments, the purpose of the present invention is fully and effectively achieved. Those skilled in the art can understand that the present invention includes but is not limited to the contents described in the drawings and the above specific embodiments. Although the present invention has been described with respect to the most practical and preferred embodiments currently considered, it should be understood that the present invention is not limited to the disclosed embodiments, and any modification that does not deviate from the functional and structural principles of the present invention will be included in the scope of the claims.

Claims

1. A supersonic turbine engine based on detonation combustion, comprising an engine inlet duct and an inner duct and an outer duct which are axially arranged downstream of the engine inlet duct and arranged concentrically, wherein at least a compressor rotor and a compressor stator are arranged in the engine inlet duct, characterized in that: At least one main combustion chamber is provided in the upstream section of the inner duct, at least a turbine rotor and a turbine stator are provided downstream of the main combustion chamber, a main tail nozzle is provided at the end of the downstream section of the inner duct and forms an exhaust passage containing the main combustion gas flow, and the compressor rotor, compressor stator, main combustion chamber, turbine rotor and turbine stator constitute the main components of the turbine engine core machine; The upstream section of the outer duct is formed into a two-stream airflow outlet channel, the two-stream airflow outlet channel is located outside the outer ring of the main combustion chamber, and is used to lead out the high-temperature and high-pressure air participating in the detonation combustion from the engine intake duct, and the downstream of the two-stream airflow outlet channel extends at least to the rear end of the main combustion chamber in the axial direction, and a bleed valve for adjusting the flow rate of the lead-out gas is provided at the end thereof; The midstream section of the outer duct is formed into a two-stream airflow introduction channel, which is used to introduce two airflows participating in detonation combustion from the two-stream airflow outlet channel, and a check valve structure is provided at the downstream end of the two-stream airflow introduction channel, which is used to prevent the detonation pressure wave at the moment of detonation initiation from being transmitted back to the upstream; The downstream section of the outer duct is formed into a detonation combustion chamber, and the upstream part of the detonation combustion chamber is provided with at least one oil collecting ring and a high-energy electric ignition device in sequence along the axial direction, and the downstream part of the detonation combustion chamber is formed into a detonation combustion space, wherein the oil collecting ring is used to atomize the fuel and spray it into the detonation combustion chamber, and the high-energy electric ignition device is used to ignite the atomized fuel and realize detonation initiation, and the detonation combustion space is used to accommodate and accelerate the combustion gas and has a geometric configuration suitable for the propagation of detonation waves, and a detonation tail nozzle is installed in series at the end of the detonation combustion chamber, and the throat area of ​​the detonation tail nozzle is adjustable to meet the expansion and discharge requirements of the detonation gas flow under different flight conditions.

2. A supersonic turbine engine based on detonation combustion as claimed in claim 1, characterized in that: The high-temperature and high-pressure air entering the detonation combustion chamber to participate in the detonation combustion is drawn from two air flows outside the outer ring of the main combustion chamber. The flow rate of the drawn-out gas is regulated by the bleed valve, and the bleed valve is communicated with the engine control unit, and the opening of the bleed valve is adjusted according to the real-time flight status and the pressure in the detonation combustion chamber to ensure the stable supply of high-temperature and high-pressure air and the working stability of the detonation combustion chamber under different flight conditions.

3. A supersonic turbine engine based on detonation combustion as claimed in claim 1, characterized in that: The detonation tail nozzle is a convergent-divergent tail nozzle with adjustable throat area, and its outlet area and throat area ratio are adjusted in real time at least according to flight conditions and working state parameters of the detonation combustion chamber to adapt to the acceleration and expansion characteristics of the supersonic exhaust airflow.

4. A supersonic turbine engine based on detonation combustion as claimed in claim 3, characterized in that: The throat area of ​​the detonation tail nozzle is adjusted by a detonation tail nozzle throat area adjustment mechanism installed on the outer wall of the detonation combustion chamber. The detonation tail nozzle throat area adjustment mechanism is communicated with the engine control unit, and the outlet area and throat area ratio of the detonation tail nozzle are adjusted to ensure the optimal working state of the detonation tail nozzle.

5. A supersonic turbine engine based on detonation combustion as claimed in claim 1, characterized in that: A zirconium oxide high-temperature protective coating is sprayed on the inner wall of the detonation combustion chamber to improve the thermal shock resistance of the detonation combustion chamber in a high-temperature environment, and to reduce infrared radiation by lowering the temperature of the outer wall of the detonation combustion chamber, thereby improving the stealth performance of the aircraft.

6. A supersonic turbine engine based on detonation combustion as claimed in claim 1, characterized in that: A heat insulation device is provided between the detonation combustion chamber and the downstream section of the inner duct, which is used to effectively block the heat transfer between the two in a high-temperature working environment, so as to avoid thermal damage to the downstream section of the inner duct and the main tail nozzle when the detonation combustion chamber is working.

7. A supersonic turbine engine based on detonation combustion as claimed in claim 1, characterized in that: The check valve structure is a double-layer plate design, has the function of adaptively adjusting the valve opening, and can automatically adjust the valve opening degree according to the intensity of the detonation pressure wave to prevent the detonation pressure wave from transmitting back to the upstream of the engine at the moment of detonation.

8. A supersonic turbine engine based on detonation combustion as claimed in claim 1, characterized in that: The fuel required for detonation combustion is sprayed out from the oil collecting ring. The oil collecting ring is installed upstream of the detonation combustion chamber and downstream of the check valve structure. The oil collecting ring is evenly distributed with multiple injection holes around the entire circumference. The injection angle is toward the downstream side of the airflow and inclined inwardly at an acute angle. When detonation initiation and combustion occur, the fuel is sprayed out from the injection holes of the oil collecting ring to form atomized fuel.

9. A supersonic turbine engine based on detonation combustion as claimed in claim 8, characterized in that: The high-energy electric ignition device is arranged downstream of the oil collecting ring to ignite the atomized fuel, thereby achieving detonation initiation. The high-energy electric ignition device adopts a redundant design and is equipped with at least two groups of ignition electrodes. When one group of electrodes fails to work due to failure or exhaustion of service life, the other group of electrodes can immediately take over the work to ensure the continuous reliability of the detonation combustion process and avoid detonation combustion failure or engine performance degradation due to single point failure.

10. A control method for a supersonic turbine engine based on detonation combustion according to any one of claims 1 to 9, characterized in that: The control method comprises at least the following steps when implemented: SS1. Start the engine and adjust the air flow distribution Initialize the working status of the compressor rotor and stator through the engine control unit, start the turbine engine core, ensure that the engine enters the normal working mode, gradually increase the engine speed to the preset value, and monitor the working parameters of various engine components, including but not limited to the compressor outlet temperature, pressure and turbine inlet temperature; SS2. Preparation for switching to detonation combustion working mode Monitor the flight parameters in real time. When the flight Mach number exceeds a preset threshold and the various working parameters of the turbine engine core are stable, the engine control unit issues a detonation combustion working mode switching instruction, gradually adjusts the working state of the turbine engine core, and simultaneously opens the bleed valve and adjusts its opening, gradually introduces two air flows from the outer ring of the main combustion chamber into the detonation combustion chamber, and ensures that the air flow in the detonation combustion chamber can meet the detonation combustion initiation conditions; SS3. Adjust the detonation tail nozzle throat area Before entering the detonation combustion working mode, the engine control unit adjusts the throat area of ​​the detonation tail nozzle in advance according to at least the current flight altitude, flight speed, external air pressure and cavity pressure conditions of the detonation combustion chamber of the aircraft to ensure that the cavity pressure of the detonation combustion chamber can meet the detonation combustion initiation conditions; SS4. Start the detonation combustion process When the airflow enters the detonation combustion chamber and the airflow rate and chamber pressure reach the detonation conditions, the engine control unit controls the detonation combustion fuel supply device to supply fuel into the oil collecting ring. When the fuel supply flow rate and fuel supply pressure reach the preset level, the high-energy electric ignition device is started to ignite the fuel atomized by the oil collecting ring and sprayed into the detonation combustion chamber to achieve detonation initiation ignition, and monitor the pressure wave changes in the detonation combustion process in real time to ensure that the detonation combustion can continue, and dynamically adjust the fuel injection amount to maintain the stability of the detonation combustion, and monitor the working state of the check valve structure in real time to ensure that it can effectively prevent the detonation pressure wave from propagating upstream, and reduce the adverse effects of the detonation pressure wave at the moment of detonation initiation on the upstream components of the engine; SS5. Feedback control and dynamic regulation of knock combustion During the detonation combustion process, the engine control unit monitors the indoor cavity pressure of the detonation combustion chamber, the total inlet pressure of the detonation tail nozzle and the local static pressure at the flight altitude of the aircraft in real time, adopts a closed-loop control strategy and combines the flight condition data to adjust the opening of the bleed valve and the outlet area and throat area ratio of the detonation tail nozzle in real time to ensure that the expansion and discharge of the detonation gas flow are always in the optimal state.

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