A high-temperature compact aero-engine combustor

Abstract

This invention discloses a high-temperature, compact aero-engine combustor, relating to the application field of aero-engine combustor model design. To meet the design requirements of next-generation engines, this invention optimizes the swirler and combustor structure, achieving stable and effective combustion within a compact space and high-efficiency combustion under high fuel-to-air ratio conditions. This improves the combustor's thermal strength, resulting in a combustor that operates efficiently within a compact space under high fuel-to-air ratio conditions. The aero-engine includes a combustor casing, a diffuser, a swirler, and a flame tube; the diffuser is installed at the front of the combustor casing; the flame tube is installed inside the combustor casing, with its rear connected to the combustor casing. A swirler is installed at the center of the front part of the flame tube, and several cooling holes and several mixing holes are installed on the side wall of the flame tube from front to back.

Description

Technical Field

[0001] This invention relates to the field of aero-engine combustion chamber model design and application. Technical Background

[0002] As a core thermodynamic component of power plants such as aero-engine gas turbine engines and ground-based gas turbines, the combustion chamber's combustion efficiency, temperature rise level, structural compactness, hot-end component reliability, and pollutant emission characteristics directly determine the overall engine's thrust-to-weight ratio, power density, fuel economy, and service life. It is one of the key indicators for evaluating the advancement of a power system. With the iteration of advanced aero-engines towards higher thrust-to-weight ratios, higher specific thrust, and lighter weight, and the stringent requirements for high efficiency and low emissions in ground-based gas turbines, the combustion chamber faces multi-objective constraints of "high temperature rise, compactness, and high stability." Currently, the combustion chamber temperature rise requirement for advanced military aero-engines has reached 1200K–1400K, requiring an axial length compressed to within 300mm and a thermal intensity increased to 10... 6 kW / m³ and above. Traditional combustor designs are no longer adequate to meet the aforementioned performance upgrade requirements. To achieve these goals, high-temperature compact aero-engine combustors must meet a series of key technical requirements: first, wide-condition combustion stability, adaptable to the entire operating range from idle to maximum power, ensuring ignition and flame stabilization capabilities under extreme conditions; second, uniformity of the outlet temperature field, avoiding thermal shock to turbine blades caused by localized high temperatures.

[0003] While existing high-temperature rise combustors have made some progress in structural optimization and combustion organization, they have not fundamentally resolved the contradiction between "compactness and efficient combustion, and high temperature rise and thermal protection." They also struggle to fully meet key technical requirements such as stability under wide operating conditions and uniform outlet temperature, and are unable to adapt to the multi-objective design needs of next-generation power plants. Therefore, developing a high-temperature rise compact combustor capable of achieving rapid flame stabilization, uniform mixing, and efficient combustion within an extremely short axial dimension has become a key bottleneck in aero-engine technology breakthroughs and has significant engineering application value. Summary of the Invention

[0004] To address the aforementioned problems, this invention discloses a compact aero-engine combustor with a centrally staged high-temperature rise based on a three-stage cyclone separator. This invention can satisfy flame stability and exit temperature field uniformity across a wide operating range.

[0005] The technical solution of the present invention is as follows: the aero-engine includes a combustion chamber casing 1, a diffuser 2, a swirler 3, and a flame tube 4; the diffuser 2 is installed at the front of the combustion chamber casing 1; the flame tube 4 is installed inside the combustion chamber casing 1, and its rear part is connected to the combustion chamber casing 1; a swirler 3 is installed at the center of the front part of the flame tube 4; and a plurality of cooling holes 5 and a plurality of mixing holes 6 are installed on the side wall of the flame tube 4 from front to back.

[0006] Furthermore, the cyclone separator is a three-stage cyclone separator, comprising, from front to back, a pre-combustion stage nozzle 31, a first-stage blade 32, a second-stage blade 33, a main combustion stage nozzle 34, and a third-stage blade 35. The first-stage blades 32, 33, and 35 are evenly arranged along the axial direction of the cyclone separator 3. The angle between the first-stage blade 32 and the tangent at its outer edge is 30°, and there are 8 of them; the angle between the second-stage blade 33 and the tangent at its outer edge is 45°, and there are 12 of them; the angle between the third-stage blade 35 and the tangent at its outer edge is 45°, and there are 16 of them.

[0007] Furthermore, the air entering the flame tube 4 through the cyclone separator 3 accounts for 55% of the total air volume; the air entering the flame tube through the cooling hole 5 accounts for 36% of the total air volume; and the air entering the flame tube through the mixing hole 6 accounts for 9% of the total air volume.

[0008] Furthermore, some of the cooling holes 5 are evenly arranged around the head of the flame tube 4 around the vortex generator 3; others are evenly distributed on the flame tube 4 from front to back, with adjacent rows of cooling holes 5 staggered and the hole diameter is 1mm.

[0009] Furthermore, the mixing holes 6 are distributed downstream of the flame tube 4, and the hole diameter is 5 mm.

[0010] This invention, through technological innovation and optimized design, achieves stable and efficient operation of the combustion chamber under compact conditions, providing strong support for the research of next-generation aero-engines. Under high-temperature and high-pressure airflow conditions, the optimized design of the swirler creates a stable recirculation zone within a compact space, which is beneficial for combustion organization. Compared with existing technologies, this invention has the following beneficial and significant technical effects: 1. This invention adopts an optimized combustion organization method, achieving higher heat release intensity and total temperature rise within the same space, reaching a temperature rise of 1400K with an air-fuel ratio of 0.048. This meets the core requirement of high thrust-to-weight ratio and high power density power plants for high temperature rise. 2. While ensuring combustion performance and stable operation, this invention effectively shortens the axial and radial dimensions, with a flame tube height of 108mm and a length of 180mm, reducing the overall size and weight of the engine, improving engine space utilization, and adapting to lightweight and highly compact installation requirements. 3. This invention results in more uniform combustion heat release and lower outlet temperature field distortion. With an air-fuel ratio of 0.048, the outlet temperature distribution coefficient is only 0.17, and the radial outlet temperature distribution coefficient curve meets the requirements of mainstream combustion chambers. This effectively reduces the thermal shock and local thermal load of turbine components, improving the reliability and service life of hot-end components. Fourth, the flame stabilization mechanism of this invention is more reasonable and has a wider flame stabilization range. By adjusting the fuel intake of the main combustion stage nozzle and the pre-combustion stage nozzle, the combustion chamber can operate stably when the air-fuel ratio is 0.02-0.05, making it less prone to flameout and oscillating combustion, and improving the reliability of ignition and restart performance. Attached Figure Description

[0011] Figure 1 This is a schematic diagram of the combustion chamber structure in this case. Figure 2 This is a three-dimensional schematic diagram of the combustion chamber structure in this case. Figure 3 This is a schematic diagram of the airflow distribution in the combustion chamber of this case. Figure 4 This is a partially enlarged schematic diagram of the hydrocyclone in this case. Figure 5 This is a three-dimensional schematic diagram of the hydrocyclone in this case. Figure 6 This is a schematic diagram of the fuel and air flow direction of the cyclone in this case; In the diagram, 1 is the combustion chamber casing, 2 is the diffuser, 3 is the cyclone separator, 4 is the flame tube, 5 is the cooling hole, 6 is the mixing hole, 31 is the pre-combustion stage nozzle, 32 is the first-stage blade, 33 is the second-stage blade, 34 is the main combustion stage nozzle, and 35 is the third-stage blade. Detailed Implementation

[0012] To clearly illustrate the technical features of this patent, the following detailed description is provided through specific embodiments and in conjunction with the accompanying drawings.

[0013] Addressing numerous limitations and shortcomings of existing aero-engines, such as the contradiction between "compactness and efficient combustion, high temperature rise and thermal protection," and the difficulty in fully meeting key technical requirements like stability under wide operating conditions and uniform outlet temperature, this invention aims to provide a high-temperature-rise compact aero-engine combustor to solve at least one of the aforementioned and other technical problems in the prior art. By optimizing the swirler and combustor structure, this invention achieves stable and effective combustion organization within a compact space, realizing efficient combustion under high fuel-air ratio conditions, thereby improving the combustor's thermal capacity and obtaining a combustor that operates efficiently within a compact space under high fuel-air ratio conditions. This invention is a compact combustor with a flame tube height of 108 mm and a length of 180 mm.

[0014] The aero-engine includes a combustion chamber casing 1, a diffuser 2, a swirler 3, and a flame tube 4; the diffuser 2 is installed at the front of the combustion chamber casing 1; the flame tube 4 is installed inside the combustion chamber casing 1, and its rear is connected to the combustion chamber casing 1; a swirler 3 is installed at the center of the front part of the flame tube 4; and several cooling holes 5 and several mixing holes 6 are installed on the side wall of the flame tube 4 from front to back.

[0015] Furthermore, the cyclone separator 3 is a three-stage cyclone separator, comprising a pre-combustion stage nozzle 31, a first-stage blade 32, a second-stage blade 33, a main combustion stage nozzle 34, and a third-stage blade 35, arranged from front to back. The first-stage blades 32, 33, and 35 are evenly distributed along the axial direction of the cyclone separator 3. The angle between the first-stage blade 32 and the tangent at its outer edge is 30°, and there are 8 of them; the angle between the second-stage blade 33 and the tangent at its outer edge is 45°, and there are 12 of them; the angle between the third-stage blade 35 and the tangent at its outer edge is 45°, and there are 16 of them.

[0016] The airflow entering the combustion chamber through the swirler 3 accounts for 55% of the total airflow. Under the action of the swirler 3, this air forms a strong swirling region in the main combustion zone of the flame tube 4, providing sufficient conditions for combustion and flame stability. Through optimization during the design process, the structure of the swirler 3 ensures localized fuel richness in the pre-combustion zone, guaranteeing stable combustion in the combustion chamber under low operating conditions. Simultaneously, the main combustion stage nozzle 34 can be activated under high operating conditions, achieving stable combustion under high operating conditions by adjusting the stage ratio.

[0017] The downstream of the flame tube 4 has uniformly distributed mixing holes 6 with a diameter of 5 mm along the circumference. The air flow into the flame tube 4 through the mixing holes 6 accounts for 9% of the total airflow. The high-temperature combustion products formed in the flame tube 4 mix with the relatively low-temperature air entering the flame tube through the mixing holes 6, which reduces the temperature of the downstream section of the flame tube 4 and optimizes the temperature distribution of the combustion gas.

[0018] Cooling holes 5, each with a diameter of 1 mm, are evenly arranged circumferentially on the wall of the flame tube 4. Gas entering the flame tube 4 through the cooling holes 5 forms a cooling gas film along its wall, thereby meeting the cooling requirements of the flame tube wall and minimizing the impact on the combustion process inside the flame tube 4.

[0019] Figure 3 The diagram illustrates the internal aerodynamics of the combustion chamber: Air from the compressor enters at point a, flows through the diffuser, and reaches point b. At the head of the flame tube, it splits into three streams. One stream enters the flame tube through a vortex generator, forming a recirculation zone at point c to stabilize combustion. The remaining two streams flow downstream from the flame tube and the inner and outer rings of the casing, respectively. When these two streams reach point d, a portion of the air enters the flame tube through a mixing orifice, forming a mixing zone at point e downstream of the flame tube. The last stream enters the flame tube through film cooling holes on the flame tube wall, forming a film on the inner wall to cool the wall. Finally, all the air flows out of the combustion chamber at point e.

[0020] Figure 6 The diagram illustrates the distribution of air and fuel channels. The air channel passes through the first-stage, second-stage, and third-stage blades. The pre-combustion stage fuel channel is injected through the nozzle in the middle of the swirler. The main combustion stage fuel channel is injected through the oil gap between the second and third stages. Three main combustion stage fuel channels are arranged circumferentially.

[0021] There are many specific ways to implement this invention. The above description is only a preferred embodiment of this invention. It should be noted that for those skilled in the art, several improvements can be made without departing from the principle of this invention, and these improvements should also be considered within the scope of protection of this invention.

Claims

1. A high-temperature compact aero-engine combustion chamber, characterized in that, The aero-engine includes a combustion chamber casing (1), a diffuser (2), a vortex generator (3), and a flame tube (4); the diffuser (2) is installed at the front of the combustion chamber casing (1); the flame tube (4) is installed inside the combustion chamber casing (1), and its rear is connected to the combustion chamber casing (1). A vortex generator (3) is installed at the center of the front of the flame tube (4), and several cooling holes (5) and several mixing holes (6) are installed on the side wall of the flame tube (4) from front to back.

2. The high-temperature compact aero-engine combustion chamber according to claim 1, characterized in that, The cyclone is a three-stage cyclone, which is provided from front to back with a pre-combustion stage nozzle (31), a first-stage blade (32), a second-stage blade (33), a main combustion stage nozzle (34), and a third-stage blade (35). The first-stage blade (32), second-stage blade (33), and third-stage blade (35) are evenly arranged along the axial direction of the hydrocyclone (3); the angle between the first-stage blade (32) and the tangent at the outer edge is 30°; the angle between the second-stage blade (33) and the tangent at the outer edge is 45°; the angle between the third-stage blade (35) and the tangent at the outer edge is 45°.

3. The high-temperature compact aero-engine combustion chamber according to claim 1, characterized in that, The air entering the flame tube (4) through the cyclone separator (3) accounts for 55% of the total air volume; the air entering the flame tube through the cooling hole (5) accounts for 36% of the total air volume; and the air entering the flame tube through the mixing hole (6) accounts for 9% of the total air volume.

4. The high-temperature compact aero-engine combustion chamber according to claim 1, characterized in that, A portion of the cooling holes (5) are evenly arranged around the cyclone separator (3) at the head of the flame tube (4); a portion of the cooling holes (5) are evenly distributed from front to back on the flame tube (4), with adjacent rows of cooling holes (5) staggered and the hole diameter is 1mm.

5. The high-temperature compact aero-engine combustion chamber according to claim 1, characterized in that, The mixing holes (6) are distributed downstream of the flame tube (4) and have a diameter of 5 mm.

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