A mixer structure for an air turbine rocket engine

Abstract

The present invention provides a mixer structure for an air-turbine rocket engine, which includes a jet annulus, a swirl vane, and a center cone. Along the gas flow direction, the front end of the center cone is a hemisphere, which serves to reduce the total pressure loss. The rear part of the center cone is a circumferential groove, and the circumferential groove is covered by an annular partition to form an internal flow passage. The annular partition smoothly transitions with the outer wall surface of the center cone, and the overall shape is a frustum of a cone. A jet annulus is formed between the annular partition and the outer wall surface of the center cone, and the jet annulus connects the internal flow passage with the external flow passage outside the annular partition and the center cone. The internal flow passage is a gas flow passage, and the external flow passage is an air flow passage. The gas in the internal flow passage can enter the air flow passage from the jet annulus and mix with the air. The present invention is a passive mixing structure. Compared with the active mixing structure, this type of structure is simple, reliable, and has a good mixing effect. In addition, compared with other passive mixing structures, this structure can achieve full fuel mixing within a shorter distance while ensuring a lower total pressure loss.

Description

Technical Field

[0001] The present invention relates to the technical field of air turbo rocket engines, and particularly to a mixer structure suitable for an air turbo rocket engine. Background Art

[0002] An air turbo rocket (ATR) is an air-breathing combined cycle engine. It uses a gas generator to produce high-temperature fuel-rich gas to drive a turbine, which drives a compressor to compress air. The outside air is pressurized by the compressor and then mixed with the fuel-rich gas at the turbine outlet and undergoes secondary combustion in the main combustion chamber. The generated high-temperature gas is fully expanded through a nozzle and discharged to produce thrust. Compared with conventional turbine engines, the air turbo rocket engine has a wider operating range, can achieve a flight speed of Ma = 0 - 4 and a flight altitude of H = 0 - 30 km, and has advantages such as simple structure, light weight, and high thrust-to-weight ratio. This engine uses outside air as an oxidant, so its specific impulse is much higher than that of a rocket engine, and it can also be recycled, achieving better economy. Compared with a ramjet engine, the air turbo rocket engine overcomes the disadvantage of poor self-starting performance and can accelerate autonomously from a stationary state at sea level to a high-altitude high-speed cruise state. As an organic integration of a rocket engine and an aeroengine, the air turbo rocket engine has characteristics such as high comprehensive performance, good technical inheritance, and strong realizability, and is a new power system that can be applied to near-space vehicles in China in a relatively short time.

[0003] However, to achieve engineering applications, there are still many key technologies to be overcome for the air turbo rocket engine. A key problem faced in the research of this engine is the combustion problem of high-temperature gas and air in the combustion chamber. The quality of combustion has a crucial impact on the performance of the entire engine. The more complete the combustion, the more energy is released, the higher the thrust and specific impulse of the engine, and at the same time, the lower the pollutant emissions. Uniform mixing of gas and air before entering the combustion chamber is a prerequisite and an important factor for complete fuel combustion. In addition, the primary flow and secondary flow after the turbine of the air turbo rocket engine are mixed. This process involves the interaction between high-temperature, high-speed fuel-rich gas and relatively low-temperature, low-speed air, as well as the exchange of momentum, energy, and mass. The mixing effect directly affects the combustion efficiency, and thus affects the performance of the engine. Any total pressure loss will reduce the expansion work ability of the gas flow in the nozzle, resulting in a decrease in the thrust and economy of the engine. Therefore, it is necessary to achieve full mixing of gas and air in a shorter distance while reducing the total pressure loss, which poses higher requirements for the design of the mixing structure of the air turbo rocket engine mixer.

[0004] Numerous scholars at home and abroad have conducted extensive research on the mixing problem of gaseous fuel and air. According to its mechanism, the mixing methods are mainly divided into active mixing and passive mixing. Active mixing mainly relies on large-scale self-excitation. Common active mixing methods mainly include pulsed jets and wavy walls, etc. For example, the sinusoidal wavy wall mixing structure designed by Li et al. with a frequency of 1200 Hz. Passive mixing, on the other hand, is based on the axial vortex structure induced by different jet methods or configurations. Common passive mixing methods mainly include jet hole structures, lobe mixers, and swirlers, etc. For example, the injection mixing device invented by Tian Guangzhi et al., the lobe mixer designed by Chen Xin et al. based on the inlet structure parameters of the air turbine rocket engine combustion chamber, and the ternary swirl mixer invented by Sun Lin et al. Although active mixing can also enhance the mixing effect, it requires adding additional active control structures in the engine, increasing the complexity of the engine structure, and its reliability still needs further research and improvement. In contrast, passive mixing promotes the uniform mixing of fluids through special structures. Its structure is simple, reliable, and has a good mixing effect, and it does not require adjusting the structure parameters according to the working conditions changes. It is still the mainstream mixing method at present. Summary of the Invention

[0005] In an air turbine rocket engine, before the fuel enters the combustion chamber for combustion, it needs to be fully mixed with air in the mixer. The mixing of the core gas and the bypass air has always been one of the key problems to be solved in the research of air turbine rocket engines. The existing passive mixing structures all have relatively large axial dimensions, and it is difficult to achieve full fuel mixing within a short distance while ensuring a low total pressure loss. To improve the mixing problem in air turbine rocket engines, the present invention proposes a mixer structure for an air turbine rocket engine, which overcomes the circumferential non-uniformity during the mixing process and can achieve full fuel mixing within a short distance while ensuring a low total pressure loss.

[0006] The technical solution of the present invention is as follows:

[0007] The mixer structure for an air turbine rocket engine described includes a jet annulus, swirl vanes, and a central cone;

[0008] Along the gas flow direction, the front end of the central cone is a hemisphere, which plays a role in reducing the total pressure loss; the rear part of the central cone is a circumferential groove, and the circumferential groove is covered with an annular partition to form an inner flow path. The annular partition is smoothly transitioned with the outer wall surface of the central cone, and the whole is a frustum of a cone. A jet annulus is formed between the annular partition and the outer wall surface of the central cone. The jet annulus connects the inner flow path with the annular partition and the outer flow path outside the central cone. The inner flow path is a gas flow path, and the outer flow path is an air flow path. The gas in the inner flow path can enter the air flow path from the jet annulus and mix with the air;

[0009] The ratio of the outlet area A1 of the jet annulus to the inlet area A2 of the gas flow path is 0.1 to 0.25;

[0010] The swirl vanes are installed between the mixer shroud and the central cone, or between the mixer shroud and the annular baffle;

[0011] The inlet angle α1 of the swirl vanes is not greater than 20°; the outlet angle α2 of the swirl vanes is 30° to 60°; the number of swirl vanes is 8 to 16;

[0012] The angle α3 between the generatrix of the frustum and the central axis is 15° to 25°.

[0013] Furthermore, the annular baffle is fixedly supported by three support plates that are 120° apart from each other.

[0014] Furthermore, the ratio of the projected length L1 of the swirl vanes in the axial direction to the length L2 of the central cone is 0.4 to 0.5.

[0015] Advantageous Effects

[0016] The present invention integrates the jet structure, the swirl structure, and the central cone, making the structure simpler and more compact. It combines the mixing characteristics of the three and overcomes their respective disadvantages, enabling the fuel and air to be fully mixed within a shorter distance. Compared with pure jet mixing and pure swirl mixing, the present invention significantly shortens the mixing distance and has a lower total pressure loss.

[0017] The present invention uses a jet annulus structure to replace the traditional jet hole structure to solve the problem of uneven mixing in the circumferential direction. The fuel is injected into the air flow path through the annulus, which helps to enhance the penetration depth of the fuel and form an entrainment effect with the air, thus greatly enhancing the mixing effect. The swirl vanes with a certain angle and curvature can further enhance the swirl effect and reduce the total pressure loss. The airflow passing through the swirl vanes forms a swirl with a rotational expansion characteristic, which can increase the turbulence intensity and extend the actual mixing distance. The central cone plays a role in decelerating and expanding the pressure, which is beneficial to the stable combustion of the engine combustion chamber. At the same time, the total pressure loss of the central cone structure is small and the generated recirculation effect helps to promote mixing.

[0018] The present invention also presents the key parameter relationships that affect mixing and total pressure loss: the ratio of the outlet area of the jet annulus to the inlet area of the gas flow path, the outlet angle of the swirl vanes, the number of swirl vanes, and the angle between the generatrix of the frustum at the front end of the central cone and the central axis. By designing the key parameters and cooperating with the relevant structures, the fuel can be fully mixed within a shorter distance while ensuring a lower total pressure loss.

[0019] The additional aspects and advantages of the present invention will be partially given in the following description, partially become obvious from the following description, or be understood through the practice of the present invention. Description of the Drawings

[0020] The above and / or additional aspects and advantages of the present invention will become apparent and be readily understood from the description of the embodiments in conjunction with the following drawings, wherein:

[0021] Figure 1 : Overall sectional three-dimensional schematic diagram of the present invention.

[0022] Figure 2 : Overall three-dimensional schematic diagram of the mixer.

[0023] Figure 3 : Overall sectional three-dimensional schematic diagram of the mixer.

[0024] Figure 4 : Front view of the mixer.

[0025] Figure 5 : Sectional schematic diagram of the mixer.

[0026] Figure 6 : Left view of the mixer.

[0027] Figure 7 : Flow field streamline diagram of Embodiment 1.

[0028] Figure 8 : Flow field streamline diagram of Embodiment 2.

[0029] In the figure: 1. Jet annular gap, 2. Swirl vane, 3. Central cone, 4. Annular partition, 5. Mixer outer cover. Specific embodiments

[0030] The mixer structure of the air turbine rocket engine proposed by the present invention mainly includes a jet annular gap, a swirl vane, and a central cone.

[0031] As Figures 1 to 4 shown, along the gas flow direction, the front end of the central cone is a hemispherical body, which can reduce the total pressure loss; the rear part of the central cone is a circumferential groove, and the circumferential groove is covered by an annular partition to form an internal flow path. The annular partition is fixedly supported on the central body by three support plates that are 120° apart from each other. The annular partition and the outer wall surface of the central cone are smoothly transitioned, and the overall shape is a frustum of a cone. The angle α3 between the generatrix of the frustum of the cone and the central axis is 15° to 25°.

[0032] A jet annular gap is formed between the annular partition and the outer wall surface of the central cone. The jet annular gap connects the internal flow path with the external flow path outside the annular partition and the central cone; the internal flow path is a gas flow path, and the external flow path is an air flow path. The gas in the internal flow path can enter the air flow path from the jet annular gap and mix with the air; the ratio of the outlet area A1 of the jet annular gap to the inlet area A2 of the gas flow path is 0.1 to 0.25.

[0033] The swirl vanes are installed between the mixer shroud and the central cone, or between the mixer shroud and the annular baffle, and play a role in fixedly supporting the central cone. The inlet angle α1 of the swirl vanes is not greater than 20°; the outlet angle α2 of the swirl vanes is 30° to 60°; the number of swirl vanes is 8 to 16; the ratio of the projection length L1 of the swirl vanes in the axial direction to the length L2 of the central cone is 0.4 to 0.5.

[0034] The specific parameters here are determined through the following analysis process:

[0035] To intuitively analyze the mixing ability of the mixer, the concept of the discrete ratio, which measures the degree of deviation from the average value, is defined. The formula for the discrete ratio is as follows:

[0036]

[0037] Among them, x represents the fuel, ω represents the mass fraction at a certain point on the cross-section, represents the average mass fraction of the cross-section. Disc x The smaller the value, the more uniform the fuel distribution.

[0038] If the ratio of the outlet area A1 of the jet annulus to the inlet area A2 of the gas flow path is too small, it will cause congestion of the gas at the annulus and excessive total pressure loss. The larger the ratio, the larger the discrete ratio. Therefore, the preferred ratio is 0.1 to 0.25.

[0039] Table 1 shows the discrete ratio and total pressure recovery coefficient calculated through numerical simulation by only changing the parameter of the outlet angle α2 of the swirl vanes and only changing the number of swirl vanes at the aspect ratio of 4 / 3.

[0040] Table 1

[0041]

[0042] It can be obtained from the data analysis in Table 1 that the larger the outlet angle α2 of the swirl vanes, the smaller the discrete ratio, but the greater the total pressure loss; too many or too few swirl vanes will cause the discrete ratio to be too large. Therefore, the preferred outlet angle α2 of the swirl vanes is 30° to 60°; the number of swirl vanes is 8 to 16.

[0043] If the angle α3 between the generatrix of the frustum at the front end of the central cone and the central axis is too small, it is not conducive to the deceleration and pressure expansion of the combustible mixture, and thus affects the combustion organization in the combustion chamber at the rear end of the mixer; the larger α3 is, the larger the discrete ratio and the total pressure loss will be. Therefore, the preferred α3 is 15° to 25°.

[0044] The embodiments of the present invention will be described in detail below. The embodiments are exemplary and are intended to explain the present invention and should not be construed as limiting the present invention.

[0045] In Embodiment 1, the jet annular gap is arranged at the front end of the swirl vane, the ratio of A1 to A2 is 0.23, the number of vanes is 8, and the vane outlet angle is 60°.

[0046] As Figure 7 shown. The mixing process is as follows. The high-temperature gas flowing out of the turbine enters the gas flow passage and is then injected into the air flow passage through the jet annular gap, which helps to enhance the penetration depth of the fuel and form an annular gas enrichment zone. After the air flowing out of the compressor enters the air flow passage, it immediately pre-mixes with the gas ejected from the annular gap and forms an entrainment effect, thus greatly enhancing the mixing effect. The pre-mixed gas then enters the swirl vane, and at the same time, the central cone plays a role in decelerating and expanding the pressure of the mixed gas, which helps to reduce the flow loss of the gas flow and is beneficial to the organization of combustion in the rear-end combustion chamber. After the mixed gas flows through the swirl vane, it forms a swirl flow under the action of the swirl vane, which has the characteristics of rotational expansion, can increase the turbulence intensity, and extend the actual mixing distance. Finally, due to the structure of the central cone, a local recirculation zone is generated at the outlet of the mixer, and under the combined action of entrainment, swirl and recirculation, uniform mixing of gas and air is achieved within a short distance.

[0047] In Embodiment 2, the jet annular gap is arranged at the rear end of the swirl vane, the ratio of A1 to A2 is 0.19, the number of vanes is 16, and the vane outlet angle is 34°.

[0048] As Figure 8 shown. The mixing process is as follows. The bypass air enters the mixer, first passes through the swirl vane, and at the same time, the central cone plays a role in decelerating and expanding the pressure of the air, which helps to reduce the flow loss of the air and is beneficial to the organization of combustion in the rear-end combustion chamber. The air forms a strong swirl under the action of the swirl vane, which has the characteristics of rotational expansion, can increase the turbulence intensity, and extend the actual mixing distance. After the air leaves the swirl vane, it immediately mixes strongly with the gas ejected from the jet annular gap and forms an entrainment effect, thus greatly enhancing the mixing effect. The gas after preliminary mixing generates a recirculation after leaving the central cone, and the gas performs secondary mixing in the recirculation zone, and finally the gas and air are uniformly mixed within a short distance.

[0049] The air turbine rocket engine often adopts the lobe mixer scheme, so the mixer structure of the present invention is compared with the lobe mixer. Table 2 shows the discrete ratio and total pressure recovery coefficient calculated by numerical simulation for Embodiment 1, Embodiment 2 and the lobe mixer at the aspect ratio of 4 / 3.

[0050] Table 2

[0051]

[0052]

[0053] As can be seen from the data in Table 2, the mixer structure of the present invention is superior to the lobe mixer in terms of mixing uniformity and total pressure loss, and the performance of Example 1 is more excellent.

[0054] Although the embodiments of the present invention have been shown and described above, it can be understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those of ordinary skill in the art can make changes, modifications, substitutions, and variations to the above embodiments within the scope of the present invention without departing from the principles and spirit of the present invention.

Claims

1. A mixer structure for an air-turbine rocket engine, characterized in that: It includes a jet annular slit, a swirl vane, and a central cone; In the air flow direction, the front end of the central cone is a hemisphere; the rear part of the central cone is a circumferential groove, and the circumferential groove is covered by an annular partition to form an inner flow passage. The annular partition is smoothly transitioned with the outer wall surface of the central cone, and the whole is a frustum of a cone. A jet annular slit is formed between the annular partition and the outer wall surface of the central cone. The jet annular slit connects the inner flow passage with the annular partition and the outer flow passage outside the central cone. The inner flow passage is a gas flow passage, and the outer flow passage is an air flow passage. The gas in the inner flow passage can enter the air flow passage from the jet annular slit and be mixed with the air; The ratio of the outlet area A1 of the jet annular slit to the inlet area A2 of the gas flow path is 0.1 - 0.25; The swirl vane is installed between the mixer outer cover and the central cone; The inlet angle α1 of the swirl vane is not greater than 20°; the outlet angle α2 of the swirl vane is 30° - 60°; the number of swirl vanes is 8 - 16; The angle α3 between the generatrix of the frustum of the cone and the central axis is 15° - 25°.

2. The structure of the mixer of an air-turbine rocket engine according to claim 1, wherein: The annular partition is fixedly supported by three support plates that are 120° apart from each other.

3. The mixing device structure of an air turbo rocket engine according to claim 1, characterized in that: The ratio of the projected length L1 of the swirl vane in the axial direction to the length L2 of the central cone is 0.4 - 0.

5.

4. A mixer structure of an air-turbine rocket engine, characterized in that: It includes a jet annular slit, a swirl vane, and a central cone; In the air flow direction, the front end of the central cone is a hemisphere; the rear part of the central cone is a circumferential groove, and the circumferential groove is covered by an annular partition to form an inner flow passage. The annular partition is smoothly transitioned with the outer wall surface of the central cone, and the whole is a frustum of a cone. A jet annular slit is formed between the annular partition and the outer wall surface of the central cone. The jet annular slit connects the inner flow passage with the annular partition and the outer flow passage outside the central cone. The inner flow passage is a gas flow passage, and the outer flow passage is an air flow passage. The gas in the inner flow passage can enter the air flow passage from the jet annular slit and be mixed with the air; The ratio of the outlet area A1 of the jet annular slit to the inlet area A2 of the gas flow path is 0.1 - 0.25; The swirl vane is installed between the mixer outer cover and the annular partition; The inlet angle α1 of the swirl vane is not greater than 20°; the outlet angle α2 of the swirl vane is 30° - 60°; the number of swirl vanes is 8 - 16; The angle α3 between the generatrix of the frustum of the cone and the central axis is 15° - 25°.

5. The structure of the mixer of an air-turbine rocket engine according to claim 4, wherein: The annular partition is fixedly supported by three support plates that are 120° apart from each other.

6. The mixing device structure of an air turbine rocket engine according to claim 4, wherein: The ratio of the projected length L1 of the swirl vane in the axial direction to the length L2 of the central cone is 0.4 - 0.5.

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