A detonation combustion chamber and a turbine engine based on a combined compressor

Abstract

The application discloses a knock combustion chamber and a turbine engine based on a combined compressor, comprising: a combustion chamber body, which is internally provided with a through fluid passage and is suitable for transmitting a first fluid; a fuel nozzle, which is arranged at a front section of the combustion chamber body; and a fluid guiding structure, which is arranged outside the combustion chamber body and is communicated with the fluid passage, and is suitable for transmitting a second fluid, which is opposite to the first fluid in flow direction, from a rear section of the combustion chamber body to the front section of the combustion chamber body, and spraying the second fluid into the combustion chamber body to form a fluid barrier. The application forms the fluid barrier by introducing the second fluid through the fluid guiding structure, dynamically changes the blockage ratio by adjusting the pressure or flow of the second fluid, and solves the problem that the parameters of the conventional solid barrier are fixed and cannot adapt to the requirements of different working conditions of the engine. Compared with the solid barrier, the fluid barrier has smaller flow resistance, reduces the energy loss of airflow in the combustion chamber, and is beneficial to improving the overall performance of the engine.

Description

Technical Field

[0001] This invention relates to the field of aero-engine technology, specifically to a detonation combustion chamber and a turbine engine based on a combined compressor. Background Technology

[0002] Faced with the increasing complexity of the current flight environment, future high-performance combat aircraft place higher demands on their power plants. However, existing conventional turbine engines, based on isobaric combustion, have already reached their component efficiency limits, making further improvement difficult. Therefore, to improve engine performance and achieve a leapfrog development of my country's traditional aero-engines, it is essential to explore new power plants. Detonation combustion has an extremely short reaction time and rapid propagation speed, and can generally be approximated as an isochoric combustion process. Therefore, under the same engine inlet parameters, detonation turbine engines have higher cycle thermal efficiency and lower fuel consumption. Furthermore, their self-pressurization characteristics are beneficial for improving engine thrust-to-weight ratio and other performance characteristics.

[0003] When detonation combustion is applied to a traditional turbine engine, a detonation turbine engine is formed. It generally consists of components such as an intake, compressor, axial and radial diffusers, multi-tube detonation combustion chamber, turbine, power take-off shaft, and exhaust system. Its working principle is as follows: After the outside air is compressed by the compressor, its temperature and pressure increase. After passing through the radial and axial diffusers, a portion of the outlet air flows axially into the multi-tube detonation combustion chamber, where it is atomized and mixed with the injected fuel and then ignited, generating a shock wave. The shock wave continuously superimposes in the detonation combustion chamber to form a detonation wave. The detonation wave pressurizes the gas, and the resulting high-temperature, high-pressure gas impacts the turbine to do work. The power generated is used to drive the compressor and accessory transmission devices. Finally, the high-temperature, high-pressure gas is discharged into the outside atmosphere through the exhaust system.

[0004] When detonation turbine engines are applied to practical engineering applications, there are some key technical problems. The obstruction in the detonation combustion chamber is usually a Shchelkin spiral (solid obstruction). Once it is manufactured, parameters such as the blockage ratio cannot be changed, which cannot meet the needs of the engine under different operating conditions. At the same time, the flow resistance loss caused by the solid obstruction is large, which is not conducive to improving engine performance. Summary of the Invention

[0005] In view of this, the present invention provides a detonation combustion chamber and a turbine engine based on a combined compressor to solve the problem that the detonation combustion chamber usually uses solid obstacles, which cannot meet the requirements of different engine operating conditions and results in large flow resistance losses.

[0006] In a first aspect, the present invention provides a detonation combustion chamber, comprising:

[0007] The combustion chamber body has a through fluid channel inside, which is suitable for transporting a first fluid; A fuel nozzle, disposed at the front section of the combustion chamber body, is used to supply fuel and atomize and mix it with a first fluid; A fluid guiding structure is disposed outside the combustion chamber body and communicates with the fluid channel. The fluid guiding structure is adapted to transmit a second fluid, which flows in the opposite direction to the first fluid, from the rear section of the combustion chamber body to the front section of the combustion chamber body, and to inject the second fluid into the combustion chamber body to form a fluid obstruction.

[0008] The beneficial effects of the aforementioned knock combustion chamber are as follows: By introducing a second fluid through a fluid guiding structure to form a fluid obstruction, the blockage ratio can be dynamically changed by adjusting the pressure or flow rate of the second fluid, thus solving the problem that traditional solid obstructions have fixed parameters and cannot adapt to different engine operating conditions. The fluid obstruction has lower flow resistance than solid obstructions, reducing energy loss of airflow within the combustion chamber and contributing to improved overall engine performance.

[0009] The second fluid flows in the opposite direction to the first fluid and is injected into the combustion chamber, which can enhance the airflow disturbance in the combustion chamber, promote the interaction between the combustible premixed gas and the flame front, accelerate the transformation of the slow combustion wave into the detonation wave, and improve the stability and reliability of the detonation wave formation.

[0010] Meanwhile, the flow of the second fluid can create a cooling effect on the combustion chamber walls (especially when the temperature of the second fluid is lower than the temperature of the combustion gas), reducing the heat load and extending the service life of the combustion chamber.

[0011] In one alternative embodiment, the fluid guiding structure includes: The outer casing of the combustion chamber is sleeved outside the combustion chamber body and forms an airflow cavity between the combustion chamber body and the combustion chamber body. The air inlet of the airflow cavity is arranged in the rear section of the outer casing of the combustion chamber. At least one connecting channel group, the connecting channel group including at least one connecting channel, the connecting channel being disposed on the outer wall of the combustion chamber body, the fluid channel and the airflow cavity being connected through the connecting channel.

[0012] The beneficial effects of the above technical solution are as follows: the air inlet of the airflow cavity is arranged in the rear section of the outer cylinder of the combustion chamber, thereby allowing the second airflow entering the airflow cavity to be transported from the rear section to the front section of the outer cylinder of the combustion chamber. The second fluid in the airflow cavity can exchange heat with the outer wall of the combustion chamber body, reducing the wall temperature of the combustion chamber body. In addition, multiple connecting channel groups can be evenly arranged in the axial direction of the combustion chamber body. The connecting channels in each group are distributed circumferentially or arranged spirally along the axial direction, so that the injection of the second fluid covers different axial sections of the combustion chamber, improving the spatial uniformity of fluid obstruction.

[0013] In one alternative implementation, the connection channel is a slit or a slot; When the connecting channel group is a slit group, there are multiple slit groups. Adjacent slit groups are arranged along the axial direction of the combustion chamber body. Each slit group includes multiple slits arranged circumferentially along the combustion chamber body or spirally along the axial direction of the combustion chamber body.

[0014] In a second aspect, the present invention provides a turbine engine based on a combined compressor, comprising: A compressor system includes at least two compressor stages, which are divided into a pre-compressor and a post-compressor, wherein the pre-compressor is connected to an intake passage. The aforementioned detonation combustion chamber is used to induce detonation combustion in the fuel-air mixture; An airflow splitting structure is provided, comprising a return channel, a first airflow distribution pipe, and a second airflow distribution pipe. The return channel is connected to the outlet of the front compressor, and the outlet of the return channel is divided into a first airflow distribution pipe and a second airflow distribution pipe. The first airflow distribution pipe is connected to the fluid channel of the detonation combustion chamber, and the second airflow distribution pipe is connected to the inlet of the rear compressor. A turbine assembly, which is connected to the compressor system and disposed within the turbine chamber, is used to drive the compressor system. In at least two-stage compressors, the outlet of the first-stage compressor outputs a first gas flow and transmits the first gas flow to the fluid passage of the combustion chamber body, while the outlet of the second-stage compressor outputs a second gas flow through a first flow channel and transmits the second gas flow to the fluid guiding structure.

[0015] In one optional embodiment, the compressor system includes a first compressor and a second compressor arranged in sequence. The first compressor is connected to the fluid passage of the combustion chamber body through a first airflow distribution pipe. The first compressor is connected to the air inlet of the second compressor through a second airflow distribution pipe. The air outlet of the second compressor is connected to a fluid guiding structure. The second compressor is used to introduce the pressurized second airflow into the combustion chamber body through the fluid guiding structure.

[0016] In one optional embodiment, the first airflow distribution pipe and the second airflow distribution pipe are respectively provided with adjustment structures, and the airflow of the first airflow distribution pipe and the second airflow distribution pipe are adjusted by the adjustment structures to adapt to different engine operating conditions.

[0017] In one optional embodiment, multiple detonation combustion chambers are provided, and the multiple detonation combustion chambers are arranged in a multi-tube circumferential manner. During operation, at least some of the detonation combustion chambers work symmetrically in pairs or all of them work simultaneously. The first airflow distribution pipe is provided in multiple ways, and each first airflow distribution pipe is respectively connected to the air inlet of the detonation combustion chamber; The exhaust ports of each of the aforementioned detonation combustion chambers are connected to the turbine chamber via a confluence structure.

[0018] In one optional embodiment, the turbine assembly is disposed within a turbine chamber, which communicates with the outlet of the detonation combustion chamber; the outlet of the subsequent compressor is further connected to a second flow channel, which communicates with the blade cavity of the turbine assembly, so that the third airflow output from the subsequent compressor enters the blades of the turbine assembly and forms a cooling film on the blade surface; the turbine assembly includes: A gas turbine is mounted on a power output outer shaft, which is a hollow shaft and connected to a pre-stage compressor and a post-stage compressor. One end of the power output outer shaft is rotatably mounted in a bearing cavity via a first bearing, and the other end of the power output outer shaft is rotatably mounted in a casing via a third bearing. A power turbine is mounted on an inner power output shaft, which passes through an outer power output shaft. Both ends of the inner power output shaft extend from the outer power output shaft and are rotatably mounted in a bearing cavity via a second bearing.

[0019] In one optional embodiment, the outlet of the subsequent compressor is also connected to a third flow channel, which is connected to the bearing cavity through an airflow transmission path, so that the fourth airflow output by the subsequent compressor enters the bearing cavity. And / or, the end of the turbine chamber is connected to the exhaust device.

[0020] In one alternative implementation, the end of the power output inner shaft away from the power turbine is connected to an accessory drive and a starter generator.

[0021] In summary, the technical solution of the present invention has the following advantages: The engine of this invention employs a combined compressor, which allows part of the airflow from the compressor outlet to flow into the detonation combustion chamber to participate in combustion, while the other part is further pressurized by the subsequent compressor. The pressurized air then flows into the detonation combustion chamber to act as a fluid obstruction, which helps to reduce the flow resistance of the combustion chamber, adapt to different engine operating conditions, and at the same time reduce the thermal load on the combustion chamber wall, improving the reliability of the detonation combustion chamber under long-term operation. In addition, the pressurized air can solve problems such as bleed air sealing and blade cooling in detonation turbine engines. Attached Figure Description

[0022] To more clearly illustrate the specific embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0023] Figure 1 A schematic diagram of a turbine engine based on a combined compressor provided by the present invention; Figure 2 This is a schematic diagram of the structure of a detonation combustion chamber provided by the present invention.

[0024] Explanation of reference numerals in the attached figures: 1. Generator; 2. Accessory transmission device; 3. Intake passage; 4. First compressor; 5. Radial diffuser; 6. Airflow splitting structure; 61. Return passage; 62. First airflow distribution pipe; 63. Second airflow distribution pipe; 7. Second compressor; 71. First flow channel; 72. Second flow channel; 73. Third flow channel; 8. Ignition device; 10. Detonation combustion chamber; 101. Combustion chamber body; 102. Fluid passage; 103. Fluid guiding structure; 1031. Outer cylinder of combustion chamber; 1032. Airflow cavity; 1033. Connecting passage; 104. Fuel nozzle; 11. Gas turbine; 12. Power turbine; 13. Exhaust device; 15. Power output outer shaft; 16. Power output inner shaft; 18. First bearing; 19. Second bearing; 20. Bearing cavity; 21. Casing; 211. Turbine chamber; 22. Third bearing. Detailed Implementation

[0025] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, 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, 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.

[0026] When detonation combustion is applied to a traditional turbine engine, a detonation turbine engine is formed. It generally consists of components such as an intake, compressor, axial and radial diffusers, multi-tube detonation combustion chamber, turbine, power take-off shaft, and exhaust system. Its working principle is as follows: After the outside air is compressed by the compressor, its temperature and pressure increase. After passing through the radial and axial diffusers, a portion of the outlet air flows axially into the multi-tube detonation combustion chamber, where it is atomized and mixed with the injected fuel and then ignited, generating a shock wave. The shock wave continuously superimposes in the detonation combustion chamber to form a detonation wave. The detonation wave pressurizes the gas, and the resulting high-temperature, high-pressure gas impacts the turbine to do work. The power generated is used to drive the compressor and accessory transmission devices. Finally, the high-temperature, high-pressure gas is discharged into the outside atmosphere through the exhaust system.

[0027] When detonation turbine engines are applied to practical engineering applications, several key technical problems exist: 1) In the existing technical solutions, the knock combustion chamber generally uses solid obstructions (Shchelkin spirals). Once the obstruction ratio and other parameters are completed, they cannot be changed, which cannot meet the needs of different engine operating conditions. At the same time, the flow resistance loss generated by such obstructions is large, the structural stress is too concentrated, the reliability and stability under long-term operation are low, and the structural rigidity and adaptability are poor, making it difficult to adaptively match the needs of different engine operating conditions.

[0028] 2) Due to the self-pressurizing characteristics of detonation combustion, the resulting combustion gas pressure will be higher than the compressor outlet gas pressure. At this point, bleed air from the compressor outlet cannot meet the pressure requirements for blade cooling and bearing cavity sealing. The highest pressure point in a detonation turbine engine is at the detonation combustion chamber outlet. If bleed air from the compressor outlet at this point, it will not meet the pressure requirements for turbine blade cooling and bearing cavity sealing, leading to combustion gas intrusion, which in turn causes blade erosion and bearing failure.

[0029] 3) Under prolonged operation, the wall temperature of the detonation combustion chamber will rise sharply. The resulting heat load and stress concentration will shorten the life of the detonation combustion chamber and reduce its safety and reliability.

[0030] Based on this, the present invention provides a detonation combustion chamber and a turbine engine based on a combined compressor, aiming to specifically solve the above-mentioned key technical problems and improve the engine's adaptability to operating conditions, reliability and service life.

[0031] The specific embodiments of the present invention will now be described in detail with reference to the detonation combustion chamber of the first aspect of the present invention and the turbine engine based on the combined compressor of the second aspect.

[0032] According to an embodiment of the present invention, in a first aspect, a detonation combustion chamber is provided, combined with... Figures 1 to 2 As shown, it includes a combustion chamber body 101, a fuel nozzle 104, and a fluid guiding structure 103.

[0033] The combustion chamber body 101 has a through fluid channel 102 inside, which is suitable for transporting the first fluid.

[0034] The fuel nozzle 104 is located at the front of the combustion chamber body 101 and is used to supply fuel and atomize and mix it with the first fluid.

[0035] The fluid guiding structure 103 is disposed outside the combustion chamber body 101 and communicates with the fluid channel 102. The fluid guiding structure 103 is adapted to transmit a second fluid, which flows in the opposite direction to the first fluid, from the rear section of the combustion chamber body 101 to the front section of the combustion chamber body 101, and to inject the second fluid into the combustion chamber body 101 to form a fluid obstruction.

[0036] The aforementioned detonation combustion chamber introduces a second fluid through the fluid guiding structure 103 to form a fluid obstruction. The obstruction ratio can be dynamically changed by adjusting the pressure or flow rate of the second fluid, solving the problem that traditional solid obstructions (such as the Shchelkin spiral) have fixed parameters and cannot adapt to different engine operating conditions. Compared with solid obstructions, the fluid obstruction has less flow resistance, reducing energy loss of airflow in the combustion chamber and improving the overall engine performance.

[0037] The second fluid flows in the opposite direction to the first fluid and is injected into the combustion chamber, which can enhance the airflow disturbance in the combustion chamber, promote the interaction between the combustible premixed gas and the flame front, accelerate the transformation of the slow combustion wave into the detonation wave, and improve the stability and reliability of the detonation wave formation.

[0038] Meanwhile, the flow of the second fluid can create a cooling effect on the combustion chamber walls (especially when the temperature of the second fluid is lower than the temperature of the combustion gas), reducing the heat load and extending the service life of the combustion chamber.

[0039] The specific structure, position and fuel supply method of the fuel nozzle 104 can be changed. The fuel nozzle 104 can be a pressure atomizing nozzle, an air atomizing nozzle, etc.

[0040] In some embodiments, the fluid guiding structure 103 includes an outer combustion chamber shell 1031 and a connecting channel assembly. The outer combustion chamber shell 1031 is sleeved on the outside of the combustion chamber body 101 and forms an airflow cavity 1032 between it and the combustion chamber body 101. The connecting channel assembly is provided with at least one connecting channel, including at least one connecting channel 1033, which is disposed on the outer wall of the combustion chamber body. The fluid channel 102 and the airflow cavity 1032 are connected through the connecting channel 1033.

[0041] In this embodiment, the air inlet of the airflow cavity 1032 is arranged at the rear section of the outer casing 1031 of the combustion chamber, thereby allowing the second airflow entering the airflow cavity 1032 to be transported from the rear section of the outer casing 1031 to the front section of the outer casing 1031. The second fluid in the airflow cavity 1032 can exchange heat with the outer wall of the combustion chamber body, reducing the wall temperature of the combustion chamber body.

[0042] In addition, multiple connecting channel groups can be evenly arranged in the axial direction of the combustion chamber body 101. The connecting channels 1033 in each group are distributed circumferentially or arranged spirally along the axial direction, so that the injection of the second fluid covers different axial sections of the combustion chamber and improves the spatial uniformity of the fluid obstacle.

[0043] In some embodiments, the connecting channel 1033 is a slit or a slot. When the connecting channel group is a slit group, multiple slit groups are set, and adjacent slit groups are arranged along the axial direction of the combustion chamber body 101. Each slit group includes multiple slits arranged circumferentially along the combustion chamber body 101 or spirally arranged along the axial direction of the combustion chamber body 101. In this embodiment, the detonation combustion chamber has a certain number of slits. The size and spacing of the slits are not specifically required and can be adapted according to engine requirements or detonation combustion chamber design specifications. In addition, an outer combustion chamber outer cylinder 1031 is added outside the detonation combustion chamber. The outer combustion chamber outer cylinder 1031 needs to surround all slits. The distance and form between the outer combustion chamber outer cylinder 1031 and the combustion chamber are not limited, and the structure is as follows. Figure 2 As shown.

[0044] The working principle of the aforementioned detonation combustion chamber is as follows: Air flows into the detonation combustion chamber 10 from the air distribution port of the return channel 61. The fuel sprayed from the nozzle is atomized and mixed with the air to form a combustible premixed gas. After ignition and combustion by the igniter, a slow combustion wave is generated. At this time, a portion of the air pressurized by the rear compressor is transmitted from the tail to the head of the detonation combustion chamber and is ejected through the slit to form a fluid obstruction. Due to a certain pressure loss in the airflow from the tail to the head, the airflow pressure ejected near the head slit in the detonation combustion chamber is lower than that at the tail. As a result, the slow combustion wave will not completely block the channel during propagation, but will form an obstruction with a certain blockage ratio. After the slow combustion wave collides with the wall of the detonation combustion chamber and the obstruction, it forms a... The reflected shock wave collides at the central axis of the detonation combustion chamber, further intensifying the chemical reaction and promoting the chemical reaction with the unburned mixture. This accelerates the coupling speed between the chemical reaction front and the shock wave front, until a detonation wave is formed at a certain location in the detonation combustion chamber. The detonation wave propagates to both ends of the detonation combustion chamber. The upstream detonation wave enters the return channel through the gas distributor, while the downstream detonation wave impacts the turbine to do work. After the detonation wave exits the combustion chamber, an expansion wave propagating upstream is generated at the outlet, gradually reducing the pressure at the head of the detonation combustion chamber until the compressor outlet airflow can smoothly flow into the detonation combustion chamber. At this point, the filling process of the next cycle begins.

[0045] According to an embodiment of the present invention, in a second aspect, a turbine engine based on a combined compressor is provided, comprising a casing 21, a compressor system, a detonation combustion chamber 10, an airflow splitting structure 6, and a turbine assembly. The compressor system includes at least two compressor stages, which are divided into a pre-stage compressor and a post-stage compressor. The pre-stage compressor is connected to an intake passage 3. The detonation combustion chamber 10 is used to induce detonation combustion in a fuel-air mixture. The airflow splitting structure 6 includes a return passage 61, a first airflow distribution pipe 62, and a second airflow distribution pipe 63. The return passage 61 is connected to the outlet of the pre-stage compressor, and the outlet of the return passage 61 is divided into the first airflow distribution pipe 62 and the second airflow distribution pipe 63. The first airflow distribution pipe 62 is connected to a fluid passage 102 of the detonation combustion chamber 10, and the second airflow distribution pipe 63 is connected to the inlet of the post-stage compressor. The turbine assembly is connected to the compressor system and disposed within a turbine chamber 211 for driving the compressor system. In at least two-stage compressors, the outlet of the first-stage compressor outputs a first airflow and transmits the first airflow to the fluid passage 102 of the combustion chamber body 101, and the outlet of the second-stage compressor outputs a second airflow through the first flow channel 71 and transmits the second airflow to the fluid guiding structure 103.

[0046] In this embodiment, the first airflow output from the front compressor directly enters the fluid channel of the combustion chamber body, providing the core combustion medium for detonation combustion; the second airflow output from the rear compressor is assisted by the fluid guiding structure to adjust the flow field distribution in the combustion chamber, effectively enhancing the mixing uniformity and combustion stability of the fuel-air mixture.

[0047] In a preferred embodiment, the compressor system includes a first compressor 4 and a second compressor 7 arranged sequentially. The first compressor 4 is connected to the fluid passage 102 of the combustion chamber body 101 via a first airflow distribution pipe 62. The first compressor 4 is connected to the air inlet of the second compressor 7 via a second airflow distribution pipe 63. The air outlet of the second compressor 7 is connected to the fluid guiding structure 103. The second compressor 7 is used to introduce the pressurized second airflow into the combustion chamber body 101 through the fluid guiding structure 103. This embodiment has a front and rear combined compressor, both driven by a gas turbine. The dimensions and other parameters of the front and rear compressors may be different, and their pressure ratios may also be different, depending on specific requirements. In this embodiment, the first compressor 4 is a centrifugal compressor, and the second compressor 7 is a mixed-flow compressor.

[0048] As an alternative implementation, the combined compressor can be replaced with a multi-stage centrifugal or axial compressor, etc.

[0049] Furthermore, radial diffusers 5 are respectively installed between the outlet of the first compressor 4 and the return channel 61, and between the second airflow distribution pipe 63 and the inlet of the second compressor 7.

[0050] In some embodiments, the first airflow distribution pipe 62 and the second airflow distribution pipe 63 are respectively provided with adjustment structures. The airflow rates of the first airflow distribution pipe 62 and the second airflow distribution pipe 63 are adjusted by the adjustment structures to adapt to different engine operating conditions. That is, in this embodiment, the area ratio of the two airflow channels at the air distribution port can be determined according to the flow rate required by the knock combustion chamber. The air distribution port can be made into an adjustable area structure to adapt to different engine operating conditions. The adjustment structure can be an adjustable blade assembly, an electric throttle valve, or a pneumatic damper valve, etc.

[0051] In this embodiment, by adjusting the structure to precisely control the airflow of the first and second airflow distribution pipes, the combined compressor system can flexibly allocate the airflow ratio entering the combustion chamber and the second compressor according to the actual needs of the engine under different operating conditions. This ensures that the detonation combustion chamber always maintains a highly efficient and stable detonation combustion state, which not only helps to improve the performance and fuel economy of the turbine engine, but also reduces component wear caused by unreasonable airflow distribution, extends the service life of the engine, and enhances its adaptability and reliability under complex operating conditions.

[0052] In this embodiment, the gas pressurized by the subsequent compressor is diverted to different locations, and the airflow rate distributed at each location is determined according to the actual working requirements of the engine.

[0053] In some embodiments, multiple knock combustion chambers 10 are provided, and the multiple knock combustion chambers adopt a multi-tube circumferential arrangement. The specific number of arrangements depends on the internal space of the engine. When the knock combustion chambers are working, they usually work symmetrically in pairs or all at the same time, so that the force generated when the gas impacts the turbine is symmetrical, which can avoid the adverse effects caused by the asymmetry of the force.

[0054] In this embodiment, fuel is evenly distributed to each detonation combustion chamber by compression or other means, wherein the type of fuel is not limited and can be gaseous fuel or liquid fuel.

[0055] Multiple first airflow distribution pipes 62 are provided, each corresponding to an air inlet of a knock combustion chamber 10. That is, the return channel air distribution port and the knock combustion chamber inlet have an annular to multi-pipe airflow splitting structure to match the two and ensure smooth distribution of the compressor outlet airflow to each knock combustion chamber. Simultaneously, the air outlets of each knock combustion chamber 10 are connected to the turbine chamber 211 through a confluence structure. That is, an airflow confluence structure is provided between the knock combustion chamber outlet and the gas turbine inlet to integrate the knock combustion chamber outlet airflow and allow them to jointly impact the turbine to perform work.

[0056] In some embodiments, the outlet of the subsequent compressor is also connected to a second flow channel 72, which communicates with the inner cavity of the turbine assembly blades, so that the third airflow output by the subsequent compressor enters the blades of the turbine assembly, and then exits from the blade surface, forming a cooling film on the blade surface to protect the blades from being burned.

[0057] The turbine assembly includes a gas turbine 11 and a power turbine 12. The gas turbine 11 is mounted on a power output outer shaft 15, which is a hollow shaft connected to both the upstream and downstream compressors. One end of the power output outer shaft 15 is rotatably mounted in a bearing cavity 20 via a first bearing 18, and the other end is rotatably mounted in a casing 21 via a third bearing 22. The power turbine 12 is mounted on a power output inner shaft 16, which passes through the power output outer shaft 15. Both ends of the inner shaft 16 extend from the outer shaft 15 and are rotatably mounted in the bearing cavity 20 via second bearings 19.

[0058] In some embodiments, the outlet of the subsequent compressor is also connected to a third flow channel 73, which is connected to the bearing cavity 20 through an airflow transmission path, so that the fourth airflow output by the subsequent compressor enters the bearing cavity 20, preventing the high-temperature and high-pressure gas discharged from the detonation combustion chamber outlet from entering the bearing cavity and causing bearing failure.

[0059] Furthermore, the airflow transmission path includes a transmission channel and an airflow orifice. One end of the transmission channel is connected to the third flow channel 73, and the other end is connected to the turbine chamber 211. The airflow orifice is formed on the wall of the power output outer shaft 15 located within the turbine chamber 211. There is a transmission gap between the power output outer shaft 15 and the power output inner shaft 16, and the end of the transmission gap is connected to the bearing cavity. The fourth airflow output from the third flow channel 73 then enters the turbine chamber 211 through the transmission channel, then enters the transmission gap through the airflow orifice, and finally is transmitted to the bearing cavity through the transmission gap.

[0060] In some embodiments, the end of the turbine chamber 211 is connected to the exhaust device 13, through which airflow is discharged.

[0061] In some embodiments, the end of the power output inner shaft 16 away from the power turbine 12 is connected to an accessory drive 2 and a starter 1. The starter 1 is connected to the accessory drive 2, and the accessory drive 2 is connected to the power output inner shaft 16. Taking a two-stage compressor system as an example, a dual-rotor structure is adopted, in which the first compressor, the second compressor, and the gas turbine form one rotor, and the power turbine, the power output inner shaft, and the accessory drive form another rotor.

[0062] The aforementioned turbine engine based on a combined compressor can be applied to different types of engines, such as turboshaft, turboprop, turbojet, and turbofan engines. When applied to turboshaft and turboprop engines, the engine has a power output shaft, through which power is transmitted to drive auxiliary transmission devices, rotors, etc.; when applied to turbojet or turbofan engines, it does not have a power output shaft.

[0063] Taking a two-stage compressor system as an example, the working principle of the turbine engine based on the combined compressor is as follows: The generator 1 is started to drive the first compressor 4 and the second compressor 7, and outside air is immediately drawn into the intake passage 3. After being compressed by the first compressor 4, the airflow temperature and pressure rise synchronously, flowing along the radial diffuser 5 and the return passage 61. At this time, a portion of the airflow enters the second compressor 7 through the second airflow distribution pipe 63, where it performs work again, further increasing the pressure. Subsequently, this high-pressure airflow is divided into three paths: First, it flows into the airflow cavity 1032 between the outer cylinder 1031 and the combustion chamber body 101, and is injected into the combustion chamber through the slit, forming a fluid obstruction, which not only increases the turbulence in the combustion chamber, but also promotes the stable formation of the detonation wave; Second, it flows into the inside of the turbine blades and is ejected through the air film holes on the blade surface, forming a protective air film on the blade surface, effectively preventing the blades from being burned by high-temperature combustion gas; Third, it flows into the bearing cavity, which can prevent the high-temperature and high-pressure combustion gas discharged from the detonation combustion chamber outlet from entering the cavity, preventing the bearing from failing due to overheating.

[0064] Another portion of the airflow directly enters the detonation combustion chamber, where it mixes with the atomized fuel sprayed from the fuel nozzle 104 to form a combustible premixed gas. After ignition and combustion by the igniter 8, the resulting slow-burning wave gradually develops into a stable detonation wave in a certain area of ​​the combustion chamber under the influence of fluid obstacles. The resulting high-temperature, high-pressure gas impacts the gas turbine 11 and the power turbine 12. The power generated by the gas turbine 11 drives the first compressor 4 and the second compressor 7; the power generated by the power turbine 12 is transmitted through the power output shaft to drive the accessory transmission device 2 and the rotor (not shown) and other components. Finally, the high-temperature, high-pressure gas is discharged into the outside atmosphere through the exhaust system.

[0065] Although embodiments of the invention have been described in conjunction with the accompanying drawings, those skilled in the art can make various modifications and variations without departing from the spirit and scope of the invention, and such modifications and variations all fall within the scope defined by the appended claims.

Claims

1. A detonation combustion chamber, characterized in that, include: The combustion chamber body (101) has a through fluid channel (102) inside, which is adapted to transport a first fluid; A fuel nozzle (104) is disposed at the front section of the combustion chamber body (101) for supplying fuel and atomizing and mixing it with a first fluid; A fluid guiding structure (103) is disposed outside the combustion chamber body (101) and communicates with the fluid channel (102). The fluid guiding structure (103) is adapted to transmit a second fluid, which flows in the opposite direction to the first fluid, from the rear section of the combustion chamber body (101) to the front section of the combustion chamber body (101), and to inject the second fluid into the combustion chamber body (101) to form a fluid obstruction.

2. The detonation combustion chamber according to claim 1, characterized in that, The fluid guiding structure (103) includes: The outer casing (1031) of the combustion chamber is sleeved outside the combustion chamber body (101) and forms an airflow cavity (1032) between the combustion chamber body (101) and the airflow cavity (1032). The air inlet of the airflow cavity (1032) is arranged in the rear section of the outer casing (1031). At least one connecting channel group, the connecting channel group including at least one connecting channel (1033), the connecting channel (1033) being disposed on the outer wall of the combustion chamber body, the fluid channel (102) being connected to the airflow cavity (1032) through the connecting channel (1033).

3. The detonation combustion chamber according to claim 2, characterized in that, The connecting channel (1033) is a slit or a slot; When the connecting channel group is a slit group, the slit group is configured as multiple, and two adjacent slit groups are arranged along the axial direction of the combustion chamber body (101). Each slit group includes multiple slits arranged circumferentially along the combustion chamber body (101) or spirally arranged along the axial direction of the combustion chamber body (101).

4. A turbine engine based on a combined compressor, characterized in that, include: The compressor system includes at least two compressor stages, which are divided into a pre-compressor and a post-compressor. The pre-compressor is connected to the intake passage (3). The detonation combustion chamber (10) according to any one of claims 1-3 is used to cause detonation combustion of a mixture of fuel and air; The airflow splitting structure (6) includes a return channel (61), a first airflow distribution pipe (62), and a second airflow distribution pipe (63). The return channel (61) is connected to the outlet of the front compressor. The outlet of the return channel (61) is divided into a first airflow distribution pipe (62) and a second airflow distribution pipe (63). The first airflow distribution pipe (62) is connected to the fluid channel (102) of the detonation combustion chamber (10), and the second airflow distribution pipe (63) is connected to the inlet of the rear compressor. A turbine assembly, which is connected to the compressor system and disposed in the turbine chamber (211), is used to drive the compressor system to operate; In at least two-stage compressors, the outlet of the first-stage compressor outputs a first airflow and transmits the first airflow to the fluid passage (102) of the combustion chamber body (101), and the outlet of the second-stage compressor outputs a second airflow through the first flow channel (71) and transmits the second airflow to the fluid guiding structure (103).

5. The turbine engine based on a combined compressor according to claim 4, characterized in that, The compressor system includes a first compressor (4) and a second compressor (7) arranged in sequence. The first compressor (4) is connected to the fluid passage (102) of the combustion chamber body (101) through a first airflow distribution pipe (62). The first compressor (4) is connected to the air inlet of the second compressor (7) through a second airflow distribution pipe (63). The air outlet of the second compressor (7) is connected to the fluid guiding structure (103). The second compressor (7) is used to introduce the pressurized second airflow into the combustion chamber body (101) through the fluid guiding structure (103).

6. The turbine engine based on a combined compressor according to claim 4, characterized in that, The first airflow distribution pipe (62) and the second airflow distribution pipe (63) are respectively provided with adjustment structures. The airflow of the first airflow distribution pipe (62) and the second airflow distribution pipe (63) are adjusted by the adjustment structures to adapt to different engine operating conditions.

7. The turbine engine based on a combined compressor according to claim 4, characterized in that, The detonation combustion chamber (10) is provided in multiple ways. The multiple detonation combustion chambers are arranged in a multi-tube circumferential manner. During operation, at least some of the detonation combustion chambers work symmetrically in pairs or all of them work simultaneously. The first airflow distribution pipe (62) is provided in multiple ways, and each first airflow distribution pipe (62) is connected to the air inlet of one of the detonation combustion chambers (10); The outlet of each of the detonation combustion chambers (10) is connected to the turbine chamber (211) through a confluence structure.

8. The turbine engine based on a combined compressor according to claim 4, characterized in that, The turbine chamber (211) is connected to the outlet of the detonation combustion chamber (10); the outlet of the subsequent compressor is also connected to a second flow channel (72), which is connected to the blade cavity of the turbine assembly, so that the third airflow output by the subsequent compressor enters the blade of the turbine assembly and forms a cooling film on the blade surface; the turbine assembly includes: A gas turbine (11) is mounted on a power output outer shaft (15). The power output outer shaft (15) is a hollow shaft and is connected to the front compressor and the rear compressor. One end of the power output outer shaft (15) is rotatably mounted in the bearing cavity (20) through a first bearing (18), and the other end of the power output outer shaft (15) is rotatably mounted in the casing (21) through a third bearing (22). The power turbine (12) is mounted on the inner power output shaft (16), which passes through the outer power output shaft (15). Both ends of the inner power output shaft (16) extend from the outer power output shaft (15) and are rotatably mounted in the bearing cavity (20) through the second bearing (19).

9. The turbine engine based on a combined compressor according to claim 8, characterized in that, The outlet of the subsequent compressor is also connected to a third flow channel (73), which is connected to the bearing cavity (20) through an airflow transmission path so that the fourth airflow output by the subsequent compressor enters the bearing cavity (20). And / or, the end of the turbine chamber (211) is connected to the exhaust device (13).

10. The turbine engine based on a combined compressor according to claim 8, characterized in that, The end of the power output inner shaft (16) away from the power turbine (12) is connected to the accessory transmission device (2) and the generator (1).

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