An air turbine rocket engine

Abstract

The application discloses an air turbine rocket engine, which comprises a pulse detonation system and a air-breathing engine system with a compressor and a mixing chamber, the pulse detonation system is provided with a pulse detonation combustion chamber, a gas turbine and a rich fuel gas delivery pipeline, the gas turbine is connected with the compressor, and the rich fuel gas delivery pipeline is communicated with the pulse detonation combustion chamber and the mixing chamber. The air turbine rocket engine utilizes the characteristics of high temperature and high pressure of the outlet gas of the pulse detonation combustion chamber to realize high-efficiency work of the gas turbine.

Description

Technical Field

[0001] This invention relates to the field of engine technology, and in particular to an air turbine rocket engine. Background Technology

[0002] To achieve wide-speed-range flight from the ground to near space, combined-power engines have emerged. The most representative combined-power engines are the turbine-based combined cycle engine (TBCC) and the rocket-based combined cycle engine (RBCC).

[0003] Air Turbo Rockets (ATRs), as a special case combining TBCCs and RBCCs, mainly include two types: Air Turbo Rocket Expander-cycles (ATREXs) and Gas Generator Air Turbo Rockets (ATR-GGs). ATR-GGs are air-breathing engine systems combining a gas generator system and a turbojet engine system. ATR-GGs require carrying their own oxidizer, increasing the engine's weight; simultaneously, carrying their own oxidizer results in a lower specific impulse and lower efficiency. Because the gas turbine is driven by the high-temperature gas within the gas generator, its power output is lower than that of a conventional turbojet engine's gas turbine.

[0004] Therefore, how to provide an air turbine rocket engine that solves the above-mentioned technical problems is a technical problem that urgently needs to be solved by those skilled in the art. Summary of the Invention

[0005] The purpose of this invention is to provide an air turbine rocket engine that utilizes the high temperature and high pressure characteristics of the gas exiting the pulse detonation combustion chamber to achieve efficient operation of the gas turbine.

[0006] To achieve the above objectives, the present invention provides an air turbine rocket engine, including a pulse detonation system and an air-breathing engine system having a compressor and a mixing chamber. The pulse detonation system is provided with a pulse detonation combustion chamber, a gas turbine, and a fuel-rich gas delivery pipeline. The gas turbine is connected to the compressor, and the fuel-rich gas delivery pipeline connects the pulse detonation combustion chamber and the mixing chamber.

[0007] Preferably, the gas turbine and the fuel-rich gas delivery pipeline are located at the tail end of the pulse detonation combustion chamber, and the pulse detonation system is also provided with a vent valve and a spark plug located at the head end of the pulse detonation combustion chamber.

[0008] Preferably, the pulse detonation system is further provided with an oxidizer tank and a fuel tank, which are connected to the pulse detonation combustion chamber through the venting valve.

[0009] Preferably, a main valve is provided in the passage between the vent valve and the oxidant tank, and in the passage between the vent valve and the fuel tank.

[0010] Preferably, the pulse detonation combustion chamber has an inner cavity and an outer duct surrounding the inner cavity, and the venting valve is disposed in the inner cavity.

[0011] Preferably, the fuel tank has a first branch communicating with the inner cavity and a second branch communicating with the outer duct of the detonation chamber.

[0012] Preferably, both the first branch and the second branch are equipped with throttle valves.

[0013] Preferably, the air-breathing engine system further includes an air intake duct located at the front end of the compressor and a combustion chamber and a tail nozzle located at the end of the mixing chamber.

[0014] Compared with the above-mentioned background technology, the air turbine rocket engine provided by the present invention includes a pulse detonation system and an air-breathing engine system. The air-breathing engine system has a compressor and a mixing chamber. The pulse detonation system is provided with a pulse detonation combustion chamber, a gas turbine and a fuel-rich gas delivery pipeline. The gas turbine is connected to the compressor, and the fuel-rich gas delivery pipeline connects the pulse detonation combustion chamber and the mixing chamber.

[0015] The purpose of the aforementioned air turbine rocket engine is to solve the problems of low gas turbine power and high oxidizer mass carried by the ATR-GG engine. It replaces the gas generator of the traditional ATR-GG engine with a pulse detonation combustor, and utilizes the high temperature and high pressure characteristics of the gas at the outlet of the pulse detonation combustor to increase the power of the gas turbine. At the same time, it reduces the required oxidizer mass without changing the engine's unit thrust and specific impulse, so as to achieve the coordinated and efficient operation of the two engines. Attached Figure Description

[0016] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on the provided drawings without creative effort.

[0017] Figure 1 This is a schematic diagram of the structure of an air turbine rocket engine provided in an embodiment of the present invention;

[0018] Figure 2 This is a schematic diagram of the operation of an air turbine rocket engine provided in an embodiment of the present invention;

[0019] Figure 3 This is a two-dimensional schematic diagram of a pulse detonation engine;

[0020] Figure 4 The ideal thermodynamic cycle pv diagram for the existing ATR-GG engine;

[0021] Figure 5 The ideal thermodynamic cycle pv diagram of an air turbine rocket engine provided in an embodiment of the present invention;

[0022] Figure 6 This is the first comparison image;

[0023] Figure 7 This is the second comparison image;

[0024] Figure 8 This is the third comparison image.

[0025] in:

[0026] 1-Intake duct, 2-Compressor, 3-Mixing chamber, 4-Combustion chamber, 5-Main flow duct outside the main flow path, 6-Tail nozzle, 7-Oxidizer tank, 8-Fuel tank, 9-Main valve, 10-Throttle valve, 11-Blast valve, 12-Spark plug, 13-Knock duct outside the main flow path, 14-Pulse knock combustion chamber, 15-Gas turbine, 16-Rich gas delivery pipeline. Detailed Implementation

[0027] The technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, and not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0028] To enable those skilled in the art to better understand the present invention, the present invention will be further described in detail below with reference to the accompanying drawings and specific embodiments.

[0029] Please refer to Figures 1 to 8 ,in, Figure 1 This is a schematic diagram of the structure of an air turbine rocket engine provided in an embodiment of the present invention. Figure 2 This is a schematic diagram of the operation of an air turbine rocket engine provided in an embodiment of the present invention. Figure 3 This is a two-dimensional schematic diagram of a pulse detonation engine. Figure 4 The ideal thermodynamic cycle pv diagram for the existing ATR-GG engine is shown. Figure 5 This is a pv diagram of the ideal thermodynamic cycle of an air turbine rocket engine provided in an embodiment of the present invention. Figure 6 The first comparison image is shown. Figure 7 This is the second comparison image. Figure 8 This is the third comparison image.

[0030] In a first specific embodiment, the air turbine rocket engine provided by the present invention includes a pulse detonation system and an air-breathing engine system. The air-breathing engine system has a compressor 2 and a mixing chamber 3. The pulse detonation system is provided with a pulse detonation combustion chamber 14, a gas turbine 15 and a fuel-rich gas delivery pipe 16. The gas turbine 15 is connected to the compressor 2, and the fuel-rich gas delivery pipe 16 connects the pulse detonation combustion chamber 14 and the mixing chamber 3.

[0031] It should be noted that the conventional ATR-GG engine has a relatively small flow of fuel-rich gas driving the turbine, resulting in low turbine power, which adversely affects the compressor's boosting capability. This limits the application of the ATR-GG engine in wide-speed-range aircraft. This embodiment proposes a hybrid pulse detonation engine based on the existing ATR-GG, which is an improvement on the existing ATR-GG engine.

[0032] Specifically, in response to the problems of traditional gas generator air turbine rocket engines having a large amount of oxidizer and low gas turbine power, this embodiment replaces the gas generator of the traditional gas turbine rocket engine with a pulse detonation combustion chamber 14.

[0033] Unlike the existing ATR-GG, which combines a gas generator system and a turbojet engine system, the hybrid pulse detonation engine based on ATR-GG in this embodiment combines a pulse detonation system and an air-breathing engine system. This is equivalent to replacing the gas generator in the ATR-GG with a pulse detonation combustion chamber 14. By utilizing the high temperature and high pressure characteristics of the gas outlet of the pulse detonation combustion chamber 14, the output power of the gas turbine is increased, as well as the temperature and pressure of the fuel-rich gas before the gas turbine inlet are increased, thereby increasing the turbine pressure ratio and power of the gas turbine and improving the engine performance. Furthermore, the mass of oxidizer carried by the engine itself is reduced without affecting the overall engine performance, achieving synergistic and efficient operation of the two engines.

[0034] In the working principle of the pulse detonation system, the pulse detonation engine (PDE) can be referenced. As a new type of engine that uses periodic pulse detonation waves to generate thrust, the combustion in the combustion chamber (Pulse Detonation Combustion, PDC) can be approximated as an isochoric process, which can produce gas with higher pressure and temperature, thereby making up for the low power of the ATR-GG gas turbine. At the same time, it can reduce the weight of the oxidizer carried by the engine itself and improve the engine performance.

[0035] In addition, the air-breathing engine system also has an intake duct 1 located at the front end of the compressor 2 and a combustion chamber 4 and a tail nozzle 6 located at the end of the mixing chamber 3.

[0036] In this embodiment, the air-breathing engine system is a turbojet engine system, consisting of an intake duct 1, a compressor 2, a mixing chamber 3, a combustion chamber 4, a main flow bypass duct 5, and a tail nozzle 6; the pulse detonation system is a pulse detonation engine system, and the engine requires liquid fuel and oxidizer to ensure normal operation.

[0037] Specifically, the pulse detonation system is also equipped with an oxidizer tank 7 and a fuel tank 8. The oxidizer tank 7 supplies oxygen to the pulse detonation combustion chamber 14, and the fuel tank 8 supplies hydrogen to the pulse detonation combustion chamber 14.

[0038] In addition, the pulse detonation system is also equipped with a vent valve 11 and a spark plug 12. The vent valve 11 and the spark plug 12 are located at the beginning of the pulse detonation combustion chamber 14, and the gas turbine 15 and the fuel-rich gas delivery pipe 16 are located at the end of the pulse detonation combustion chamber 14.

[0039] In this embodiment, the oxidizer tank 7 and the fuel tank 8 are connected to the pulse detonation combustion chamber 14 via the venting valve 11.

[0040] Furthermore, main valves 9 are installed in the channels between the vent valve 11 and the oxidant storage tank 7, as well as between the vent valve 11 and the fuel storage tank 8.

[0041] In this embodiment, the number of main valves 9 depends on the oxidant tank 7 and the fuel tank 8. Taking the case where there is only one oxidant tank 7 and one fuel tank 8, a main valve 9 is installed on the pipeline of the oxidant tank 7 to control whether oxygen is supplied in the oxidant tank 7. A main valve 9 is also installed on the pipeline of the fuel tank 8 to control whether hydrogen is supplied in the fuel tank 8.

[0042] Furthermore, the pulse detonation combustion chamber 14 has an inner cavity and a detonation outer duct 13 located around the inner cavity, and the vent valve 11 is disposed in the inner cavity.

[0043] In this embodiment, the main valve 9 is opened, and the fuel and oxidant in the oxidant tank 7 and the fuel tank 8 are transported through the fuel delivery pipeline. After the fuel and oxidant in the inner cavity of the pulse detonation combustion chamber 14 are mixed evenly, the vent valve 11 is closed. At this time, the spark plug 12 is ignited, and the mixed fuel burns near the closed head end and undergoes a series of complex changes to form high temperature and high pressure gas.

[0044] Furthermore, the fuel tank 8 has a first branch communicating with the inner cavity and a second branch communicating with the detonation chamber duct 13.

[0045] In this embodiment, the fuel output from the fuel tank 8 has two paths: one is to enter the inner cavity of the pulse detonation combustion chamber 14 through the first branch and mix with the oxidant for combustion; the other is to enter the outer duct 13 of the pulse detonation combustion chamber 14 through the second branch, cool the wall of the pulse detonation combustion chamber 14, and mix with the high-temperature and high-pressure gas to form fuel-rich gas.

[0046] Furthermore, both the first and second branches are equipped with throttle valves 10.

[0047] In this embodiment, taking the example that there is one first branch and one second branch, the first branch is equipped with a throttle valve 10 and the second branch is also equipped with a throttle valve 10. The throttle valve 10 can regulate the fuel flow rate that is output from the fuel storage tank 8 and sent into the inner cavity and the knockout outer duct 13.

[0048] Please refer to Figure 1 and Figure 2 In a specific work process description:

[0049] When the main valve 9 opens, the fuel and oxidizer in the oxidizer tank 7 and fuel tank 8 enter the pulse detonation system through the fuel delivery pipeline. All the oxidizer enters the pulse detonation combustion chamber 14 through the vent valve 11, while the fuel, regulated by the throttle valve 10, partially enters the inner cavity of the pulse detonation combustion chamber 14 through the vent valve 11 to participate in combustion, and the other part enters the outer duct 13 of the detonation chamber to cool the walls of the pulse detonation combustion chamber 14. After the fuel and oxidizer in the inner cavity are evenly mixed, the vent valve 11 closes, and the spark plug 12 ignites the fuel mixture, which burns near the closed head end. The combustion of the fuel mixture forms two detonation waves that propagate to the left and right ends respectively. The detonation wave propagating towards the head end... After being reflected at the head end, the combustion wave propagates towards the tail end of the opening. The propagation speed of this combustion wave is relatively fast, and it eventually catches up with the combustion wave propagating towards the opening end. After multiple combustion waves are superimposed, a detonation wave is formed. After passing through the detonation wave, the pressure and temperature of the mixed gas increase rapidly and it spontaneously combusts, causing the reaction to continue and eventually forming high-temperature and high-pressure gas. The gas at the outlet of the pulse detonation combustion chamber 14 mixes with the fuel in the outer duct 13 of the detonation chamber to form a fuel-rich gas. The fuel-rich gas drives the gas turbine 15 to rotate, and the turbine drives the compressor 2 to rotate. The compressor 2 decelerates and pressurizes the air entering from the intake duct 1. The pressurized air mixes with the fuel-rich gas in the mixing chamber 3 and is ignited in the combustion chamber 4 to form gas. The high-temperature and high-pressure gas expands through the tail nozzle 6 and is accelerated out, forming thrust.

[0050] The aforementioned air turbine rocket engine combines the advantages of ATR-GG's high specific impulse with PDE's high thrust and high combustion efficiency, leveraging the strengths of both engines. It replaces the traditional ATR-GG's gas generator with a pulse detonation combustion chamber 14 to compensate for the traditional ATR-GG engine's lower gas turbine power and larger oxidizer mass, thereby improving the engine's overall performance.

[0051] In contrast, in a conventional ATR-GG, combustion within the gas generator is approximately isobaric, with the combustion chamber outlet pressure remaining nearly constant. However, the pulse detonation combustion chamber 14 approximates an isochoric process, where pressure increases during combustion, resulting in a higher outlet gas pressure compared to conventional gas generators. When the turbine outlet gas pressure is constant, the same mass flow rate of fuel-rich gas produces a higher turbine pressure ratio, increasing the gas turbine's power. Conversely, when the turbine pressure ratio is constant, a smaller mass flow rate of fuel-rich gas is required, thus reducing the amount of oxidizer carried by the engine itself.

[0052] Please refer to Figure 6 to Figure 8 In this embodiment, the air turbine rocket engine corresponds to the graph of the new ATR-GG, while the existing ATR-GG corresponds to the graph of the conventional ATR-GG. Figure 6This shows the change in the ratio of oxidant mass flow rate to air mass flow rate as a function of the air-fuel ratio for the conventional ATR-GG and the new ATR-GG engines, both with the same turbine power. Figure 7 Figure 8 shows a comparison of the specific thrust of the conventional ATR-GG engine and the hybrid pulse detonation engine based on ATR-GG.

[0053] according to Figure 6 It can be seen that, when the turbine pressure ratio and compressor pressure ratio are constant, compared with the traditional ATR-GG, the new ATR-GG engine, which replaces the gas generator with a pulse detonation combustion chamber 14, consumes about half the oxidizer flow rate, proving that the new ATR-GG engine can reduce the mass of oxidizer it carries. Meanwhile, from... Figure 7 and Figure 8 It can be seen that when the air-fuel ratio (defined here as the ratio of fuel mass flow rate to air mass flow rate) is relatively small, the new ATR-GG produces better thrust and specific impulse than the traditional ATR-GG, thus improving engine performance. However, when the air-fuel ratio increases further, the new ATR-GG produces less thrust and specific impulse, but the overall difference is not significant.

[0054] contrast Figure 4 and Figure 5 In the diagram: 7 / 8-11 represents the isochoric pressurization process of fuel; 11-14 represents the isochoric heating process of fuel in the gas generator; 14-15 represents the isentropic expansion process of fuel-rich gas in the gas turbine; 0-1 represents the isentropic compression process of air in the intake manifold; 1-2 represents the isentropic compression process of air in the compressor; 2-3 and 15-3 represent the isobaric mixing processes of air and fuel-rich gas in the mixing chamber; 3-4 represents the isobaric combustion process of the mixture in the main combustion chamber; 4-6 represents the isentropic expansion process of gas in the exhaust nozzle. 7 / 8-11-14-15-3-4-6-0' represents the circulation process of fuel-rich gas; 0-1-2-3-4-6-0 represents the circulation process of air.

[0055] The air turbine rocket engine in this embodiment has the following advantages:

[0056] 1. Compared to the traditional ATR-GG engine, this engine, under a certain mass flow rate of fuel-rich gas, increases the temperature and pressure of the outlet gas through the self-pressurization effect of the knock chamber, thereby increasing the power of the gas turbine;

[0057] 2. This engine can reduce the mass of oxidizer consumed while maintaining a constant gas turbine power, and has a relatively small impact on the engine's specific impulse and unit thrust;

[0058] 3. This engine can improve fuel utilization efficiency, increase compressor pressure ratio and the pressure and temperature of air at the main combustion chamber inlet, thereby increasing the pressure and temperature of the combustion gas in the main combustion chamber;

[0059] 4. This engine can reduce the noise of a "pure" pulse detonation engine and improve the comfort of near-space vehicles.

[0060] In this embodiment, replacing the gas generator of the ATR-GG engine with a pulse detonation engine can increase the turbine pressure ratio and rotational power of the turbine under the condition of a certain fuel-rich gas mass flow rate, thereby increasing the compressor pressure ratio and the gas pressure in the main combustion chamber; it can also reduce the oxidizer consumption by about 50% under the condition of a certain gas turbine power, thereby reducing the mass of oxidizer carried by the engine itself, while the engine's thrust and specific impulse remain almost unchanged.

[0061] It should be noted that in this specification, relational terms such as first and second are used only to distinguish one entity from several other entities, and do not necessarily require or imply any such actual relationship or order between these entities.

[0062] The air turbine rocket engine provided by this invention has been described in detail above. Specific examples have been used to illustrate the principles and implementation methods of this invention. The descriptions of the embodiments above are merely for the purpose of helping to understand the method and core ideas of this invention. It should be noted that those skilled in the art can make various improvements and modifications to this invention without departing from its principles, and these improvements and modifications also fall within the protection scope of the claims of this invention.

Claims

1. An air-turbine rocket engine, characterized in that, The system includes a pulse detonation system and an air-breathing engine system with a compressor (2) and a mixing chamber (3). The pulse detonation system is provided with a pulse detonation combustion chamber (14), a gas turbine (15), and a fuel-rich gas delivery pipe (16). The gas turbine (15) is connected to the compressor (2), and the fuel-rich gas delivery pipe (16) connects the pulse detonation combustion chamber (14) and the mixing chamber (3). The pulse detonation system is also provided with a vent valve (11) and a spark plug (12) located at the beginning of the pulse detonation combustion chamber (14). The pulse detonation system is also provided with an oxidizer tank (7) and a fuel tank (8), which are connected to the pulse detonation combustion chamber (14) through the venting valve (11); The pulse detonation combustion chamber (14) has an inner cavity and a detonation outer duct (13) surrounding the inner cavity, and the vent valve (11) is located in the inner cavity; The fuel tank (8) has a first branch that communicates with the inner cavity and a second branch that communicates with the detonation chamber outside the culvert (13).

2. The air turbine rocket engine according to claim 1, characterized in that, The gas turbine (15) and the fuel-rich gas delivery pipe (16) are located at the tail end of the pulse detonation combustion chamber (14).

3. The air turbine rocket engine according to claim 1, characterized in that, A main valve (9) is provided in the passage between the vent valve (11) and the oxidant tank (7) and between the vent valve (11) and the fuel tank (8).

4. The air turbine rocket engine according to claim 1, characterized in that, Both the first branch and the second branch are equipped with throttle valves (10).

5. The air turbine rocket engine according to any one of claims 1 to 4, characterized in that, The air-breathing engine system also has an intake duct (1) located at the front end of the compressor (2) and a combustion chamber (4) and a tail nozzle (6) located at the end of the mixing chamber (3).

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