Pulse detonation engine and aircraft based on pressure exchanger bleed air supercharging

Abstract

The application discloses a pulse detonation turbine engine and an aircraft based on pressure exchanger air bleed pressurization, which comprises an air inlet, a compressor, a gas turbine, an accessory device and a power output shaft, the accessory device extracts power through the power output shaft, and the engine further comprises: a pressure exchanger, which is used for pressure exchange of high-pressure gas and low-pressure gas; a detonation combustion chamber, which is used for mixing and igniting air and fuel atomization, and forming a detonation wave in the detonation combustion chamber under the action of shock wave superposition; and a gas shunt device, which is used for introducing the airflow output by the compressor into the pressure exchanger and the detonation combustion chamber respectively, introducing part of high-temperature and high-pressure gas generated by the detonation combustion chamber into the pressure exchanger, and introducing the airflow at the outlet end of the detonation combustion chamber into the gas turbine after integration. Through cooperation of the gas shunt device and the pressure exchanger, technical problems such as bearing sealing of the pulse detonation engine and turbine blade cooling difficulty are effectively solved.

Description

Technical Field

[0001] This invention relates to the field of engine technology, and in particular, to a pulse detonation turbine engine based on bleed air pressurization via a pressure exchanger. Furthermore, this invention also relates to an aircraft comprising the aforementioned pulse detonation turbine engine based on bleed air pressurization via a pressure exchanger. Background Technology

[0002] Conventional turbocharged engine structure such as Figure 1 As shown, it consists of an intake manifold, compressor, isobaric combustion chamber, gas turbine, gas supply, exhaust system, accessory devices, and power take-off shaft.

[0003] The working principle of a conventional turbine engine is as follows: outside air flows into the engine through the intake manifold, is compressed by the compressor, and the gas pressure and temperature increase. The outlet airflow is split into multiple streams. One stream flows into the turbine for bleed air sealing and blade cooling, while another stream flows into the isobaric combustion chamber, where it mixes with the injected fuel and is ignited. The resulting high-temperature, high-pressure gas impacts the gas turbine and the power turbine, converting thermal energy into the turbine's kinetic energy. The power generated by the gas turbine drives the compressor, while the power generated by the power turbine drives accessory devices (including fuel pumps, oil pumps, generators, etc.) and the helicopter rotor. Subsequently, the high-temperature, high-pressure gas is discharged into the outside atmosphere through the exhaust system.

[0004] The compressor and turbine rotor of an engine rely on bearings for positioning and rotation. These bearings require lubrication and have temperature limitations. The combustion gases produced after ignition are extremely hot. To prevent bearing failure, it's necessary to prevent the high-temperature combustion gases from flowing into the bearing cavity. Therefore, the gas pressure inside the bearing cavity needs to be higher than the combustion gas pressure in the engine's flow path. For conventional turbine engines, the airflow pressure at the compressor outlet is the highest pressure point within the entire engine, and its temperature is within the bearing's limiting range. Therefore, introducing some gas from the compressor outlet as a sealing gas can prevent the high-temperature combustion gases from entering the bearing cavity, thus protecting the bearings. Furthermore, to prevent turbine blade erosion, airflow from the compressor outlet needs to be introduced to form a cooling film on the blade surface, isolating the high-temperature combustion gases.

[0005] Under the same engine inlet parameters, compared with conventional turbine engines based on constant pressure combustion, pulse detonation turbine engines based on isochoric combustion have advantages such as high cycle thermal efficiency, low fuel consumption, and high thrust-to-weight ratio. Specifically, due to the self-pressurization characteristics of detonation combustion, the air-fuel mixture pressure increases after passing through the detonation chamber. With the same thermodynamic cycle parameters, pulse detonation turbine engines can reduce the number of compressor stages and improve the engine's thrust-to-weight ratio.

[0006] However, this brings a series of problems that restrict the practical application of pulse detonation turbine engines. These include the need for bleed-through sealing of the flow channels to prevent bearing failure due to high temperatures, and the requirement for turbine blade cooling to prevent blade burnout. In conventional turbine engines, the pressure of the air-fuel mixture gradually decreases after passing through the combustion chamber and turbine, making the compressor outlet the point of highest engine pressure. Figure 2 As shown, the high-pressure air drawn from the compressor outlet at this time can meet the pressure requirements for flow channel bleed-out sealing and turbine blade cooling. However, the pulse detonation turbine engine is different. Due to its self-pressurization characteristics, the pressure ratio before and after the combustion wave can reach 15 to 55 times. At this time, the gas pressure at the detonation chamber outlet is much higher than the air pressure at the compressor outlet, and the bleed-out air from the compressor outlet cannot meet the pressure requirements for sealing and cooling.

[0007] In gas turbine engines, the gas flow rate used for cooling and sealing accounts for approximately 3-20% of the total compressor inlet flow rate. In pulse detonation turbine engines, a relatively low-temperature, high-pressure bleed gas is required for cooling and sealing, but conventional gas pressurization methods, such as bladed compressors, cannot achieve gas pressurization with a small flow rate and high pressure ratio.

[0008] (1) Existing conventional engines based on isobaric combustion have basically reached their limits in terms of compressor, turbine efficiency, and combustion chamber efficiency. At the same time, given the upper limit of the cycle thermal efficiency under isobaric combustion, the engine performance cannot be significantly improved. However, pulse detonation engines based on isochoric combustion theoretically have higher cycle thermal efficiency than isobaric cycles. After solving problems such as bearing sealing and blade cooling, they have great potential for practical application.

[0009] (2) Conventional engines draw air from the compressor outlet for sealing the bearing cavity and cooling the turbine blades. However, given the self-pressurization characteristics of detonation combustion in pulse detonation engines, the gas pressure at the outlet of the detonation combustion chamber is higher than the air pressure at the compressor outlet, so this method is no longer applicable.

[0010] (3) When conventional mechanical devices (axial flow, centrifugal compressor, etc.) are used to pressurize the airflow at the compressor outlet, the efficiency is not high and the pressure ratio is difficult to meet the requirements. If a multi-stage compressor is used, the engine weight will be greatly increased. In addition, the power turbine needs to provide some power to drive the mechanical device in order to achieve the pressurization effect, which will reduce the power-to-weight ratio of the engine. Summary of the Invention

[0011] This invention provides a pulse detonation turbine engine and aircraft based on pressure exchanger bleed air pressurization, in order to solve the technical problems of difficulty in improving the performance of conventional turbine engines and the inability of pulse detonation turbine engines to meet sealing and cooling requirements, thus hindering their practical application and promotion.

[0012] According to one aspect of the present invention, a pulse detonation turbine engine based on pressure exchanger bleed air boosting is provided, comprising an intake manifold, a compressor, a gas turbine, an accessory device, and a power output shaft, wherein the accessory device extracts power via the power output shaft, and the turbine engine further comprises:

[0013] A pressure exchanger is used to exchange pressure between incoming high-pressure gas and low-pressure gas.

[0014] The detonation combustion chamber is used to atomize and mix the incoming air and fuel and ignite them for combustion. Under the superposition of shock waves, a detonation wave is formed in the detonation combustion chamber, and the pressure pulsation generated by the detonation combustion chamber changes synchronously with the pressure pulsation at the outlet of the pressure exchanger.

[0015] The gas splitting device is used to introduce the gas flow output from the compressor into the pressure exchanger and the detonation combustion chamber respectively, and to introduce part of the high-temperature and high-pressure gas generated in the detonation combustion chamber into the pressure exchanger. It is also used to integrate the gas flow at the outlet of the detonation combustion chamber and introduce it into the gas turbine.

[0016] As a further improvement to the above technical solution, the pressure exchanger includes a hub and a stator disposed outside the end of the hub. The hub is provided with airflow channels evenly distributed in the circumferential direction. The stator is provided with an end plate, and the end plate has a port for connecting the airflow channels and the gas diversion device respectively. The pressure exchanger is used to periodically connect the airflow channels with the port in the rotating state.

[0017] As a further improvement to the above technical solution, the pressure exchanger is connected to the drive mechanism, or the pressure exchanger is connected to the accessory device through a transmission mechanism.

[0018] As a further improvement to the above technical solution, the transmission mechanism includes an input shaft connected to the hub and a transmission gear set disposed between the input shaft and the accessory device.

[0019] As a further improvement to the above technical solution, the gas splitting device includes a first gas splitting structure and a second gas splitting structure. The inlet end of the pressure exchanger is provided with an end plate serving as a low-pressure air inlet and an end plate serving as a high-pressure gas inlet. The outlet end of the pressure exchanger is provided with an end plate serving as a high-pressure air outlet and an end plate serving as a low-pressure gas outlet. The first gas splitting structure is connected to the outlet end of the compressor, the inlet end of the detonation combustion chamber, and the low-pressure air inlet, respectively. The second gas splitting structure is connected to a preset position of the detonation combustion chamber and the high-pressure gas inlet, respectively. The gas splitting device is also used to connect the high-pressure air outlet and the bearing cavity of the gas turbine. The gas splitting device is also used to connect the low-pressure gas outlet and the inlet end of the gas turbine.

[0020] As a further improvement to the above technical solution, the preset position on the detonation combustion chamber is matched with the position where the detonation wave is formed.

[0021] As a further improvement to the above technical solution, a one-way valve is provided at the inlet end of the detonation combustion chamber.

[0022] As a further improvement to the above technical solution, the engine is a turboshaft engine, including a power turbine and an exhaust device, the power output shaft is connected to the power turbine, and the exhaust device is used to discharge the combustion gas.

[0023] As a further improvement to the above technical solution, multiple pressure exchangers are evenly arranged along the circumference of the detonation combustion chamber.

[0024] According to another aspect of the invention, an aircraft is also provided, which includes the aforementioned pulse detonation turbine engine based on pressure exchanger bleed air pressurization.

[0025] The present invention has the following beneficial effects:

[0026] Outside air enters through the intake duct, is compressed by the compressor, increasing its temperature and pressure. A gas splitting device directs the low-pressure air from the compressor outlet into the detonation combustion chamber and the pressure exchanger. Part of the low-pressure air flows into the detonation combustion chamber and mixes with the fuel injected from the combustion chamber. Once the combustible mixture is fully filled, ignition occurs. Under the continuous superposition of shock waves, a detonation wave is formed at a predetermined position in the detonation chamber. Part of the resulting high-temperature, high-pressure gas is introduced into the pressure exchanger for pressure and heat exchange, while the remaining high-temperature, high-pressure gas flows to the gas turbine. After pressure and heat exchange in the pressure exchanger, the high-pressure air can be introduced into the bearing cavity for gas sealing and turbine blade cooling. The low-pressure gas mixes with the high-temperature, high-pressure gas flowing from the detonation chamber to the gas turbine, jointly impacting the gas turbine. The turbine converts the internal energy of the gas into kinetic energy. This engine should... The detonation combustion chamber possesses self-pressurizing characteristics, thereby reducing the number of compressor stages for the same specifications, significantly improving thermal cycle efficiency, and effectively enhancing engine performance. It employs a pneumatic method to pressurize air, effectively avoiding the complexity and insufficient pressure ratio issues caused by traditional mechanical structures such as bladed compressors, achieving high-pressure-ratio gas pressurization with low flow rate. Through the cooperation of a gas splitter and a pressure exchanger, a portion of the combustion gas from the pulse detonation combustion chamber outlet and a portion of the air from the compressor outlet are exchanged within the pressure exchanger, increasing the pressure of the compressor outlet gas. Furthermore, the pressure pulsation at the pressure exchanger outlet is synchronized with the pressure pulsation generated by the detonation combustion chamber, allowing the high-pressure gas at its outlet to synchronously enter the bearing cavity for engine bleed air and sealing, effectively solving technical challenges such as bearing sealing and turbine blade cooling difficulties in pulse detonation engines.

[0027] In addition to the objectives, features, and advantages described above, the present invention has other objectives, features, and advantages. The invention will now be described in further detail with reference to the figures. Attached Figure Description

[0028] The accompanying drawings, which form part of this application, are used to provide a further understanding of the invention. The illustrative embodiments of the invention and their descriptions are used to explain the invention and do not constitute an undue limitation of the invention. In the drawings:

[0029] Figure 1 This is a schematic diagram of a conventional turbine engine;

[0030] Figure 2 This is a diagram showing the airflow pressure changes inside a conventional turbine engine;

[0031] Figure 3 This is a schematic diagram of the structure of a turbine engine according to a preferred embodiment of the present invention;

[0032] Figure 4This is a diagram showing the airflow pressure variation of a turbine engine according to a preferred embodiment of the present invention;

[0033] Figure 5 This is a schematic diagram of the wave rotor structure of a preferred embodiment of the present invention;

[0034] Figure 6 yes Figure 5 The three-dimensional structure unfolded of the wave rotor shown. Figure 1 ;

[0035] Figure 7 yes Figure 5 The three-dimensional structure unfolded of the wave rotor shown. Figure 2 ;

[0036] Figure 8 This is a two-dimensional diagram illustrating the working principle of the wave rotor according to a preferred embodiment of the present invention.

[0037] Legend:

[0038] 1. Accessory device 2. Intake duct 3. Compressor 4. Detonation combustion chamber 5. Pressure exchanger 51. Low-pressure air inlet 52. High-pressure gas inlet 53. High-pressure air outlet 54. Low-pressure gas outlet 55. Hub 56. Stator component 57. End plate 58. Airflow channel 61. First air distribution structure 62. Second air distribution structure 7. Gas turbine 8. Power turbine 9. Exhaust device 10. Power output shaft. Detailed Implementation

[0039] The embodiments of the present invention will be described in detail below with reference to the accompanying drawings. However, the present invention can be implemented in many different ways as defined and covered below.

[0040] Figure 1 This is a schematic diagram of a conventional turbine engine; Figure 2 This is a diagram showing the airflow pressure changes inside a conventional turbine engine; Figure 3 This is a schematic diagram of the structure of a turbine engine according to a preferred embodiment of the present invention; Figure 4 This is a diagram showing the airflow pressure variation of a turbine engine according to a preferred embodiment of the present invention; Figure 5 This is a schematic diagram of the wave rotor structure of a preferred embodiment of the present invention; Figure 6 yes Figure 5 The three-dimensional structure unfolded of the wave rotor shown. Figure 1 ; Figure 7 yes Figure 5 The three-dimensional structure unfolded of the wave rotor shown. Figure 2 ; Figure 8 This is a two-dimensional diagram illustrating the working principle of the wave rotor according to a preferred embodiment of the present invention.

[0041] like Figures 2 to 8As shown, the pulse detonation turbine engine based on pressure exchanger bleed air boosting in this embodiment includes an intake duct 2, a compressor 3, a gas turbine 7, an accessory device 1, and a power output shaft 10. The accessory device 1 extracts power via the power output shaft 10. This turbine engine also includes:

[0042] Pressure exchanger 5 is used to exchange pressure between incoming high-pressure gas and low-pressure gas.

[0043] The detonation combustion chamber 4 is used to atomize and mix the incoming air and fuel and ignite them for combustion. Under the superposition of shock waves, a detonation wave is formed in the detonation combustion chamber 4, and the pressure pulsation generated by the detonation combustion chamber 4 changes synchronously with the pressure pulsation at the outlet of the pressure exchanger.

[0044] The gas splitting device is used to introduce the gas flow output from the compressor 3 into the pressure exchanger 5 and the detonation combustion chamber 4 respectively, and to introduce part of the high-temperature and high-pressure gas generated by the detonation combustion chamber 4 into the pressure exchanger 5. It is also used to integrate the gas flow at the outlet of the detonation combustion chamber 4 and introduce it into the gas turbine 7.

[0045] The compressor 3 can be a centrifugal compressor 3, an axial compressor 3, or a combined compressor 3, etc. The pressure exchanger 5 in this embodiment is a wave rotor structure, which can be a through-flow, recirculation, external combustion, or internal combustion wave rotor structure. It is a component that generates a shock wave by instantaneously contacting high-pressure gas and low-pressure gas in a single component to increase the pressure of low-pressure gas. It can operate at low flow rates and has a high boost pressure ratio. The turbine engine in this embodiment takes a turboshaft engine as an example, which has a power output shaft and a power turbine. In other embodiments, it can also be a turbojet, turbofan, or turboprop engine, etc. It should be noted that turbojet and turbofan engines do not require a power output shaft and a power turbine. On the other hand, under the premise of ensuring successful detonation, the configuration of the detonation combustion chamber 4 is not limited. It can be a recirculation type that is conducive to shortening the axial length of the engine, or a straight pipe type that is simple in structure and easy to process, etc. At the same time, the type of obstacle in the detonation chamber is also not restricted. It can be an orifice plate, a groove, or a Shchelkin spiral, etc.

[0046] The working principle of this engine is as follows: Outside air enters through the intake duct 2, and is compressed by the compressor 3, increasing the temperature and pressure of the airflow. A gas splitting device is installed to introduce the low-pressure air from the compressor 3 outlet into the detonation combustion chamber 4 and the pressure exchanger 5. Part of the low-pressure air flows into the detonation combustion chamber 4 and mixes with the fuel injected from the combustion chamber. After the combustible mixture is fully filled, it is ignited and burned. Under the continuous superposition of shock waves, a detonation wave is formed at a preset position in the detonation chamber. Part of the generated high-temperature, high-pressure gas is introduced into the pressure exchanger 5 for pressure and heat exchange, while the remaining high-temperature, high-pressure gas flows to the gas turbine 7. After pressure and heat exchange in the pressure exchanger 5, the high-pressure air can be introduced into the bearing cavity for gas sealing and turbine blade cooling. The low-pressure gas mixes with the high-temperature, high-pressure gas flowing from the detonation chamber to the gas turbine 7, and together they impact the gas turbine 7, converting the internal energy of the gas into kinetic energy. This engine utilizes a detonation combustion chamber 4, whose detonation combustion has self-pressurization characteristics, thereby reducing the number of compressor stages 3 under the same specifications, significantly improving thermal cycle efficiency, and effectively improving engine performance. It employs a pneumatic method to pressurize air, effectively avoiding the complexity and insufficient pressure ratio problems caused by traditional mechanical structures such as bladed compressors 3, achieving gas pressurization with a small flow rate and a large pressure ratio. Through the cooperation of a gas splitting device and a pressure exchanger 5, a portion of the combustion gas from the outlet of the pulse detonation combustion chamber 4 and a portion of the air from the outlet of the compressor 3 are pressure-exchanged within the pressure exchanger 5, increasing the pressure of the gas at the outlet of the compressor 3. Furthermore, the pressure pulsation at the outlet of the pressure exchanger 5 is synchronized with the pressure pulsation generated by the detonation combustion chamber, allowing the high-pressure gas at its outlet to synchronously enter the bearing cavity for engine bleed air and sealing, effectively solving technical problems such as bearing sealing and turbine blade cooling difficulties in pulse detonation engines.

[0047] It should be noted that the pressure pulsation at the outlet end of the wave rotor during operation can be synchronized with the pressure pulsation generated by the detonation combustion chamber by matching the design of the wave rotor structure.

[0048] In this embodiment, the engine is a turboshaft engine, which also includes a power turbine 8 and an exhaust device 9. The power output shaft 10 is connected to the power turbine 8, and the exhaust device 9 is used to discharge the gas. It can be understood that this engine adopts a dual rotor structure, with the gas turbine 7 and the compressor 3 forming a gas generator rotor, and the power turbine 8, the power output shaft, and the accessory device 1 forming another rotor structure. The power generated by the gas turbine 7 is used to drive the compressor 3, and the power generated by the power turbine 8 drives the accessory device 1 through the power output shaft 10. The accessory device 1 may include a reducer, a generator, etc., and then the gas is discharged into the outside atmosphere through the exhaust device 9.

[0049] In this embodiment, the wave rotor includes a hub 55 and a stator 56 disposed outside the end of the hub 55. The hub 55 is provided with airflow channels 58 evenly distributed in the circumferential direction. The stator 56 is provided with an end plate 57. The end plate 57 has a port for connecting the airflow channels 58 and the gas diversion device respectively. The pressure exchanger 5 is used to periodically connect the airflow channels 58 with the port in the rotating state, thereby controlling the flow of fluid. The fluid generates compression waves and expansion waves in the wave rotor channel, thereby realizing fluid energy exchange.

[0050] It is understandable that a portion of the combustion gas from the outlet of the pulse detonation combustion chamber 4 and a portion of the air from the outlet of the compressor 3 undergo pressure exchange within the wave rotor. This increases the pressure of the gas at the outlet of compressor 3, which is then used for bleed air and sealing of the engine. The gas after pressure exchange is then depressurized and mixed with the combustion gas at the outlet of the detonation chamber to drive the low-pressure turbine or is directly discharged. The gas pressure within the wave rotor flow channel is as follows: Figure 4 As shown, the air pressure after being boosted by the wave rotor is higher than the gas pressure at the outlet of the detonation chamber, which effectively solves the problems of bleed air sealing and blade cooling in pulse detonation turbine engines, and promotes the further practical engineering application of pulse detonation turbine engines.

[0051] Specifically, the gas splitting device includes a first gas splitting structure 61 and a second gas splitting structure 62. The inlet end of the pressure exchanger 5 is provided with an end plate 57 serving as a low-pressure air inlet 51 and an end plate 57 serving as a high-pressure gas inlet 52. The outlet end of the pressure exchanger 5 is provided with an end plate 57 serving as a high-pressure air outlet 53 and an end plate 57 serving as a low-pressure gas outlet 54. The first gas splitting structure 61 is connected to the outlet end of the compressor 3, the inlet end of the detonation combustion chamber 4, and the low-pressure air inlet 51, respectively. The second gas splitting structure 62 is connected to the preset position of the detonation combustion chamber 4 and the high-pressure gas inlet 52, respectively. The gas splitting device is also used to connect the high-pressure air outlet 53 and the bearing cavity of the gas turbine 7. The gas splitting device is also used to connect the low-pressure gas outlet 54 and the inlet end of the gas turbine 7.

[0052] Among them, the preset position on the detonation combustion chamber 4 matches the position where the detonation wave is formed; part of the gas at the outlet of the compressor 3 flows into the wave rotor through the low-pressure air inlet 51, and part of the gas flows into the detonation combustion chamber 4. The gas diversion device is provided with an air inlet near the position where the detonation wave is formed, and part of the high-temperature and high-pressure gas is introduced into the wave rotor through the high-pressure gas inlet 52. After pressure and heat exchange in the wave rotor, the pressure of the high-pressure air formed is higher than the outlet pressure of the detonation combustion chamber 4. It is introduced into the bearing cavity for sealing and for cooling the turbine blades. The remaining high-temperature and high-pressure gas flows to the turbine. The low-pressure gas is mixed with the high-temperature and high-pressure gas flowing out of the detonation chamber and together impacts the gas turbine 7 and the power turbine 8.

[0053] The specific working principle of the pressure exchanger 5 in this embodiment is as follows: (Refer to...) Figure 5 The wave rotor is cut along the circumference of the hub cylinder 55. Figure 5 (as shown by the dashed line), the cross-sectional line is as follows Figure 5 As shown, the process is gradually unfolded along the cross-section, and the rotational operation of the three-dimensional wave rotor is simplified into a uniform translation process of a two-dimensional channel on a plane, ultimately unfolding into a two-dimensional diagram. (Refer to...) Figure 8 And obtain the wave system diagram inside the rotor channel at a certain moment. It can be understood that the three-dimensional channel inside the hub 55 is the rectangular structure in the two-dimensional diagram. Figure 8 The uppermost and lowermost channels are connected in the 3D diagram (i.e., Figure 8 (The two two-dimensional channel cross-sections are on the same section). The rotation of the hub 55 corresponds to the up-and-down movement of the channel, while the stationary components 56 and end plates 57 on both sides remain fixed. As the airflow channel 58 moves, its two ends pass through different ports, thus opening and closing at both ends according to a certain pattern. (Reference) Figure 8 The space-time relationship of a single working cycle of the wave rotor is shown. Figure 8 The shock wave and expansion wave trajectories shown are the lines connecting the positions of the shock wave and expansion wave in each channel of the wave rotor at that moment (both the shock wave and expansion wave are positive waves in the channel, i.e., the wave surface is perpendicular to the side wall). The wave rotor's operation starts from the bottom, at which point both ends of the channel are closed, and there is low-pressure, low-temperature flow inside the channel. The wave system generated during the operation of the wave rotor can be divided into low-pressure and high-pressure parts of the working cycle. The low-pressure cycle begins when the port corresponding to the low-pressure gas outlet 54 opens. After the gas is discharged from the pipeline, an expansion wave is generated and propagates upstream. As the wave rotor rotates, the expansion wave just reaches the port corresponding to the low-pressure air inlet 51. The pressure at this port decreases, air flows in, and reflection generates a rightward expansion wave. As the wave rotor rotates, the channel where the expansion wave is located reaches the closed end of the stator disk, and the port corresponding to the low-pressure gas outlet 54 closes. Then, a leftward shock wave is generated (the leftward shock wave in the diagram is the inlet end of the wave rotor). After the shock wave sweeps over, the gas stops, and at this time, the port corresponding to the low-pressure air inlet 51 is just closed, thus ending the low-pressure cycle. The high-pressure cycle begins when the port corresponding to the high-pressure gas inlet 52 opens. The wave rotor is connected to the detonation chamber with higher pressure. After the high-pressure gas enters the channel, a shock wave is triggered from the lower edge of the port corresponding to the high-pressure gas inlet 52, which compresses the air trapped in the channel, causing the air pressure to rise sharply. When the shock wave reaches the right end (the right end in the diagram is the wave rotor outlet end), the port corresponding to the high-pressure air outlet 53 opens and connects to the bearing cavity. At this time, the high-pressure air formed in the wave rotor flows into the bearing cavity, isolating the gas at the outlet of the detonation chamber and forming a leftward expansion wave. After reaching the left end, the port corresponding to the high-pressure gas inlet 52 closes, triggering a rightward shock wave. After the shock wave sweeps through, the gas in the channel becomes still. When it reaches the right end, the port corresponding to the high-pressure gas inlet 52 closes, and the wave rotor cycle ends.

[0054] In some embodiments, a plurality of pressure exchangers 5 are evenly arranged along the circumference of the detonation combustion chamber 4; in this embodiment, two are used as an example.

[0055] It should be noted that one or more pressure exchangers can be provided, and one or more detonation combustion chambers can also be provided. In one embodiment, a pressure exchanger can be provided, and the high-pressure gas from the pulse detonation combustion chamber can be integrated and flowed into the pressure exchanger inlet through a gas diversion device, and then the high-pressure air can be flowed into the bearing cavity and the gas turbine inlet end through the gas diversion device.

[0056] In some embodiments, the pressure exchanger 5 is connected to the drive mechanism, or the pressure exchanger 5 is connected to the accessory device 1 through the transmission mechanism; the pressure exchanger 5 can be directly driven by a drive mechanism such as a motor, or power can be extracted from the accessory device 1 through the transmission mechanism.

[0057] Specifically, the transmission mechanism includes an input shaft connected to the hub 55 and a transmission gear set disposed between the input shaft and the accessory device 1. The pressure exchanger 5 and the input end have a preset transmission ratio, so that the working speed of the wave rotor matches the operating conditions.

[0058] In some embodiments, a one-way valve is provided at the inlet end of the detonation combustion chamber 4. The detonation wave propagates to both ends. By setting the one-way valve, the pressure of the detonation combustion chamber 4 is prevented from being transmitted back. The detonation wave propagating downstream is discharged towards the turbine. After the detonation combustion chamber 4 discharges the gas, the resulting expansion wave propagates upstream of the detonation combustion chamber 4. The pressure at the head of the detonation combustion chamber 4 gradually decreases until the gas flow at the outlet end of the compressor 3 can flow smoothly into the detonation combustion chamber 4, and the detonation combustion chamber 4 begins the next cycle.

[0059] On the other hand, this embodiment also provides an aircraft that uses the above-mentioned pulse detonation turbine engine based on pressure exchanger bleed air pressurization.

[0060] Example 1

[0061] Tests and comparisons were conducted on the turbine engine of this preferred embodiment. After replacing the conventional isobaric combustion chamber with a pulse detonation combustion chamber, the thermal cycle efficiency can be increased from 27% to 49%, and the unit fuel consumption rate of the detonation cycle is 29% of that of the isobaric cycle.

[0062] The above description is merely a preferred embodiment of the present invention and is not intended to limit the invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A pulse detonation turbine engine based on bleed air booster via a pressure exchanger, characterized in that, The engine includes an intake manifold (2), a compressor (3), a gas turbine (7), an accessory device (1), and a power output shaft (10), wherein the accessory device (1) extracts power via the power output shaft (10), and the engine further includes: Pressure exchanger (5) is used to exchange pressure between incoming high-pressure gas and low-pressure gas; The detonation combustion chamber (4) is used to atomize and mix the incoming air with fuel and ignite it for combustion. Under the superposition of shock waves, a detonation wave is formed in the detonation combustion chamber (4), and the pressure pulsation generated by the detonation combustion chamber (4) changes synchronously with the pressure pulsation at the outlet end of the pressure exchanger (5). The gas splitting device is used to introduce the low-pressure air output by the compressor (3) into the pressure exchanger (5) and the detonation combustion chamber (4) respectively, and to introduce part of the high-temperature and high-pressure gas generated by the detonation combustion chamber (4) into the pressure exchanger (5), and to integrate the airflow at the outlet end of the detonation combustion chamber (4) and introduce it into the gas turbine (7). In this process, after pressure and heat exchange in the pressure exchanger (5), the pressure of the high-pressure air formed is higher than the outlet pressure of the detonation combustion chamber (4), and the high-pressure air is introduced into the bearing cavity.

2. The pulse detonation turbine engine based on pressure exchanger bleed air pressurization according to claim 1, characterized in that, The pressure exchanger (5) includes a hub (55) and a stationary component (56) disposed outside the end of the hub (55). The hub (55) is provided with airflow channels (58) evenly distributed in the circumferential direction. The stationary component (56) is provided with an end plate (57). The end plate (57) has a port for connecting the airflow channels (58) and the gas diversion device respectively. The pressure exchanger (5) is used to periodically connect the airflow channels (58) with the port in the rotating state.

3. The pulse detonation turbine engine based on pressure exchanger bleed air pressurization according to claim 2, characterized in that, The pressure exchanger (5) is connected to the drive mechanism, or the pressure exchanger (5) is connected to the accessory device (1) through the transmission mechanism.

4. The pulse detonation turbine engine based on pressure exchanger bleed air pressurization according to claim 3, characterized in that, The transmission mechanism includes an input shaft connected to the hub (55) and a transmission gear set disposed between the input shaft and the accessory device (1).

5. The pulse detonation turbine engine based on pressure exchanger bleed air pressurization according to claim 2, characterized in that, The gas splitting device includes a first gas splitting structure (61) and a second gas splitting structure (62). The inlet end of the pressure exchanger (5) is provided with an end plate (57) serving as a low-pressure air inlet (51) and an end plate (57) serving as a high-pressure gas inlet (52). The outlet end of the pressure exchanger (5) is provided with an end plate (57) serving as a high-pressure air outlet (53) and an end plate (57) serving as a low-pressure gas outlet (54). The first gas splitting structure (61) is connected to the outlet end of the compressor (3), the inlet end of the detonation combustion chamber (4), and the low-pressure air inlet (51), respectively. The second gas splitting structure (62) is connected to the preset position of the detonation combustion chamber (4) and the high-pressure gas inlet (52); the gas splitting device is also used to connect the high-pressure air outlet (53) and the bearing cavity of the gas turbine (7); the gas splitting device is also used to connect the low-pressure gas outlet (54) and the inlet end of the gas turbine (7).

6. The pulse detonation turbine engine based on pressure exchanger bleed air pressurization according to claim 5, characterized in that, The preset position on the detonation combustion chamber (4) matches the position where the detonation wave is formed.

7. The pulse detonation turbine engine based on pressure exchanger bleed air pressurization according to claim 1, characterized in that, A one-way valve is provided at the inlet end of the detonation combustion chamber (4).

8. The pulse detonation turbine engine based on bleed air pressurization via a pressure exchanger according to any one of claims 1-7, characterized in that, The engine is a turboshaft engine, including a power turbine (8) and an exhaust device (9). The power output shaft (10) is connected to the power turbine (8), and the exhaust device (9) is used to discharge the combustion gas.

9. The pulse detonation turbine engine based on bleed air pressurization via a pressure exchanger according to any one of claims 1-7, characterized in that, Multiple pressure exchangers (5) are evenly arranged along the circumference of the detonation combustion chamber (4).

10. An aircraft, characterized in that, Including the pulse detonation turbine engine based on pressure exchanger bleed air pressurization as described in any one of claims 1-9.

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