Aero-engine ultra-compact combustion chamber with stepped air intake structure

Abstract

The application discloses an aviation engine super-compact combustion chamber with a ladder type air inlet structure and belongs to the technical field of aviation gas turbine devices. The aviation engine super-compact combustion chamber comprises an annular cavity, an outer fuel ring, an inner fuel ring and a dilated type mixing and afterburning structure, the outer side of the annular cavity is provided with a ladder type outer ring air inlet, the inner side of the annular cavity is provided with an inner ring air inlet, the ladder type outer ring air inlet has at least two stepped flow guide surfaces distributed along the axial direction, the outer fuel ring is arranged on the rear wall surface of the annular cavity (1), the inner fuel ring is arranged on the front wall surface of the annular cavity, the dilated type mixing and afterburning structure is connected to the downstream of the annular cavity, the outer side of the dilated type mixing and afterburning structure is provided with an outer afterburning hole and an outer mixing hole, the inner side of the dilated type mixing and afterburning structure is provided with an inner afterburning hole (14) and an inner mixing hole (15), the axis of the ladder type outer ring air inlet and the inner ring air inlet is at an angle of 30-40 degrees with the radial direction of the annular cavity, and the axis is uniformly distributed in the circumferential direction. The aviation engine super-compact combustion chamber accelerates flame propagation by using a supergravity environment and forms a stable supergravity combustion area in the annular cavity by means of tangential injection of air, thereby ensuring combustion performance, greatly shortening the axial length of the combustion chamber and improving the thrust-to-weight ratio and the power-to-weight ratio.

Description

Technical Field

[0001] This invention relates to a miniature aircraft engine combustion chamber, belonging to the technical field of aircraft gas turbine devices. Background Technology

[0002] The combustion chamber is one of the three core components of an aero-engine, responsible for converting the chemical energy of fuel into kinetic energy. Its performance directly affects the overall performance of the engine and even the entire aircraft. By utilizing the principle of accelerating flame propagation in a hypergravity environment, the axial dimension of the combustion chamber can be significantly shortened, reducing the overall weight of the aircraft and thus improving the thrust-to-weight ratio and power-to-weight ratio.

[0003] With the rapid development of UAV technology and the national strategic layout for the low-altitude economy, the demand for high-performance small turbojet engines is becoming increasingly urgent. However, existing micro-compact combustor designs for aero-engines still have certain technical problems. Currently, all micro-compact combustor designs used in small aero-engines employ an outer ring intake scheme, leading to the following issues: (1) High total pressure loss: Due to the small opening area of ​​the annular cavity, air flow is restricted, resulting in a high total pressure loss; (2) High temperature zone is concentrated on the inner wall: The outer ring air intake method causes the high temperature zone in the annular cavity to be mainly distributed on the inner wall, which easily leads to high temperature corrosion of the wall surface and accelerates the aging of the components; Therefore, to address the aforementioned issues, it is necessary to further optimize the air intake method and combustion organization structure of the annular cavity in order to improve the performance of the ultra-compact combustion chamber of a small turbojet engine. Summary of the Invention

[0004] This invention provides an ultra-compact combustion chamber for an aero-engine with a stepped intake structure, comprising an annular cavity (1), an outer fuel ring (5), an inner fuel ring (6), and an expansion-type mixing and afterburning structure (7). The annular cavity (1) has a stepped outer ring intake (2) on its outer side and an inner ring intake (3) on its inner side. The stepped outer ring intake (2) has at least two stepped guide surfaces distributed axially. The outer fuel ring (5) is located on the rear wall of the annular cavity (1), and the inner fuel ring (6) is located on the front wall of the annular cavity (1). The expansion-type mixing and afterburning structure (7) is connected downstream of the annular cavity (1), and has an outer afterburning hole (8) and an outer mixing hole (9) on its outer side, and an inner afterburning hole (14) and an inner mixing hole (15) on its inner side. A spark plug (10) is located on the front wall of the annular cavity (1).

[0005] The axes of the stepped outer ring air inlet (2) and the inner ring air inlet (3) are at an angle of 30°-40° to the radial direction of the annular cavity, and are evenly distributed circumferentially.

[0006] Preferably, the stepped outer ring air inlet (2) is provided with an outer ring air inlet baffle (4) extending 3 mm into the annular cavity.

[0007] Preferably, the annular cavity (1) is provided with a flame stabilizing step (16) on the inner side, and the height of the flame stabilizing step is 2.2 mm.

[0008] Preferably, the front wall of the annular cavity (1) is provided with at least three combustion chamber mounting rods (12) fixedly connected thereto. The combustion chamber mounting rods extend along the axial direction of the annular cavity and are evenly spaced along the circumference of the annular cavity.

[0009] Preferably, the inner fuel ring (6) is connected to the front wall of the annular cavity (1) via an inner fuel ring mounting bracket (11), and the outer fuel ring (5) is connected to the rear wall of the annular cavity (1) via an outer fuel ring mounting bracket (13).

[0010] Preferably, the stepped annular cavity (1), the outer annular air inlet (2) and the inner annular air inlet (3) can be arranged in a periodic 30° angle.

[0011] Therefore, the ultra-compact combustion chamber of an aero-engine using the aforementioned stepped intake structure has the following advantages compared to the prior art: (1) This invention applies the mechanism of hypergravity combustion acceleration flame to the main combustion chamber of a turbojet engine, which can form a hypergravity combustion environment in the annular cavity, thereby changing the traditional combustion chamber along the axial direction to combustion along the circumference of the annular cavity. This can increase the residence time of fuel in the annular cavity, significantly shorten the axial length of the combustion chamber, thereby reducing the weight of the combustion chamber and the whole engine, and improving the thrust-to-weight ratio and power-to-weight ratio.

[0012] (2) The present invention adopts a double-sided air intake method with inner and outer rings, which not only greatly increases the opening area of ​​the combustion chamber and reduces the total pressure loss, but also forms a low temperature area in the inner ring, which better solves the problem that the high temperature area of ​​the existing micro-aero-engine ultra-compact combustion chamber is concentrated on the inner wall.

[0013] (3) The outer ring of the annular cavity of the present invention is provided with a flame stabilizing step, which can improve the stability of combustion in the annular cavity.

[0014] The technical solution of the present invention will be further described below with reference to the accompanying drawings and specific embodiments.

[0015] Figure 1 This is an overall isometric schematic diagram of an ultra-compact combustion chamber for an aero-engine with a stepped intake structure according to the present invention. Figure 2 This is a front view of an ultra-compact combustion chamber for an aero-engine with a stepped intake structure according to the present invention. Figure 3 This is a cross-sectional view along the front view of an ultra-compact combustion chamber for an aero-engine with a stepped air intake structure according to the present invention. Figure 4This is a circumferential cross-sectional view along the middle surface of the annular cavity of an ultra-compact combustion chamber of an aero-engine with a stepped air intake structure according to the present invention.

[0016] Figure 5 This is a comparison diagram of the performance of the ultra-compact combustion chamber of an aero-gas turbine with a stepped air intake structure according to the present invention and the prototype.

[0017] Figure 6 This is a combustion chamber temperature distribution diagram obtained from numerical simulation of the ultra-compact combustion chamber of an aero-gas turbine with a stepped intake structure according to the present invention.

[0018] In the diagram, 1-annular cavity, 2-stepped outer annular air inlet, 3-inner annular air inlet, 4-outer annular air inlet baffle, 5-outer fuel ring, 6-inner fuel ring, 7-expansion type mixing and afterburning structure, 8-outer afterburning hole, 9-outer mixing hole, 10-spark plug, 11-inner fuel ring mounting bracket, 12-combustion chamber mounting bracket, 13-outer annular fuel ring mounting bracket, 14-inner afterburning hole, 15-inner mixing hole, 16-flame stabilizing step

[0019] To further clarify the technical solution of the present invention, a detailed description is now provided in conjunction with the accompanying drawings and specific embodiments.

[0020] Unless otherwise expressly defined, the technical or scientific terms used in this invention shall be interpreted as meaning commonly understood by those skilled in the art.

[0021] Figure 1 This is an overall isometric schematic diagram of an ultra-compact combustion chamber for an aero-engine with a stepped intake structure according to the present invention. Figure 2 Front view of the combustion chamber; Figure 3 This is a cross-sectional view of the combustion chamber along the frontal viewing direction; Figure 4 This is a circumferential cross-sectional view of the combustion chamber along the middle surface of the annular cavity; Figure 5 The results show a comparison of the numerical simulation performance of the combustion chamber and the prototype. Figure 6 This is a temperature distribution diagram of the combustion chamber obtained from numerical simulation.

[0022] like Figure 1-4 As shown, this invention provides an ultra-compact combustion chamber for an aero-engine with a stepped intake structure, comprising an annular cavity 1, an outer fuel ring 5, an inner fuel ring 6, and an expansion-type mixing and afterburning structure 7. The annular cavity 1 serves as the main combustion zone, with a stepped outer annular intake 2 on the outer side and an inner annular intake 3 on the inner side. The stepped outer annular intake 2 has at least two stepped guide surfaces distributed axially. Corresponding numerical calculations show that using four stepped guide surfaces can achieve a total combustion efficiency of 94.2%. Figure 5The outer fuel ring 5 is located on the rear wall of the annular cavity 1, and the inner fuel ring 6 is located on the front wall of the annular cavity 1. The expanded blending and afterburning structure 7 is located downstream of the annular cavity 1 as an afterburning and blending zone, and forms a main combustion-afterburning-blending structure in the annular cavity. It has an outer afterburning hole 8 and an outer blending hole 9 on its outer side, and an inner afterburning hole 14 and an inner blending hole 15 on its inner side.

[0023] In addition, a spark plug 10 is provided on the front wall of the annular cavity 1 to generate an electric spark to ignite the fuel; the outer annular air inlet 2 and the inner annular air inlet 3 are evenly distributed on the outer and inner sides of the annular cavity 1 at a certain angle, respectively, and a high gravity combustion environment is constructed in the annular cavity by enhancing the swirling effect.

[0024] In addition, the axes of the stepped outer ring air inlet 2 and the inner ring air inlet 3 form an angle of 30°-40° with the radial direction of the annular cavity, and are evenly distributed circumferentially. Figure 5 The numerical simulation results are for when the axis of the stepped outer ring air inlet 2 and the inner ring air inlet 3 makes a 35° angle with the radial direction of the annular cavity. This angle has better overall performance.

[0025] In addition, the stepped outer ring air inlet 2 is provided with an outer ring air inlet baffle 4 extending 3 mm into the annular cavity.

[0026] In addition, a flame-stabilizing step 16 is provided on the inner side of the annular cavity 1, and the height of the flame-stabilizing step is 2.2 mm.

[0027] In addition, to enhance structural stability, the front wall of the annular cavity 1 is provided with at least three combustion chamber mounting rods 12 that are fixedly connected to it. The combustion chamber mounting rods extend along the axial direction of the annular cavity and are evenly spaced along the circumference of the annular cavity. The center of the mounting rod is provided with a threaded mounting hole coaxial with it, which can be installed in the casing by bolts.

[0028] In addition, the inner fuel ring 6 is connected to the front wall of the annular cavity 1 via the inner fuel ring mounting bracket 11, and the outer fuel ring 5 is connected to the rear wall of the annular cavity 1 via the outer fuel ring mounting bracket 13; the inner fuel ring mounting bracket 11 and the outer fuel ring mounting bracket 13 are integrally formed or welded with the combustion ring on the inner side, and the outer side is a snap-fit ​​structure for fixing the fuel ring.

[0029] In addition, a flame stabilizing step 16 is provided on the inner side of the annular cavity 1 to improve the combustion stability in the annular cavity. The outer and inner ring fuels can be gaseous or liquid fuels, which are injected into the annular cavity in the co-current direction after the flame stabilizing step 16, thereby optimizing the fuel mixing and atomization characteristics.

[0030] In addition, the stepped annular cavity 1, the outer annular air inlet 2, and the inner annular air inlet 3 can be arranged in a 30° periodic pattern.

[0031] Working Principle: Air flowing axially from the centrifugal compressor enters the combustion chamber of this invention primarily through three paths: a stepped outer annular inlet, an inner annular inlet, and internal and external afterburning and mixing orifices. Air enters the annular cavity circumferentially at a certain angle through the stepped outer and inner annular inlets, forming a strong swirling flow within the annular cavity, thus creating a hypergravity combustion environment. In this environment, the combustion and atomization rates of the fuel are significantly improved, and the denser fuel moves towards the outer diameter of the annular cavity, while the less dense intermediate products concentrate in the inner diameter region. Therefore, the residence time of the fuel in the annular cavity is only related to its mass, ensuring complete combustion without requiring a long axial length. After the fuel and air are thoroughly mixed, they burn rapidly under the ignition of the electric spark generated by the spark plug. Subsequently, the combustion products flow out of the annular cavity and enter the afterburning and mixing zone, where they further burn and mix with the air entering through the internal and external afterburning and mixing orifices before finally exiting the combustion chamber.

[0032] Figure 5 The numerical simulation results show a comparison between the ultra-compact combustion chamber of an aero-gas turbine with a stepped intake structure as described in this invention and the KJ66 prototype. Although the total pressure loss (ξ) of this invention increases slightly by 0.88% and the combustion efficiency (η) decreases by 0.9%, its axial length is shortened by nearly 1 / 3, its total weight is reduced by 7.7%, its thrust-to-weight ratio (T / W) is increased by 8.3%, and its combustion chamber outlet temperature distribution (OTDF) performance is superior. Figure 6 The figure shows the combustion chamber temperature distribution obtained from numerical simulation of the ultra-compact combustion chamber of an aero-engine with a stepped intake structure according to the present invention. As can be seen from the figure, the present invention can form a continuous low-temperature region on the inner wall of the annular cavity, effectively alleviating the problem of high-temperature regions concentrating on the inner wall of the annular cavity.

[0033] Therefore, the ultra-compact combustion chamber of aero-engines with a stepped intake structure proposed in this invention can meet the requirements of small and micro aero-engines for high-performance combustion chambers, and effectively solve the problems of large total pressure loss and high temperature concentration of the annular cavity on the inner wall in current small and micro turbojet engines, providing a new design idea for the development of future high-performance aero-engines.

[0034] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and do not constitute a limitation thereof. Although the present invention has been described in detail with reference to preferred embodiments, those skilled in the art can still make appropriate modifications or equivalent substitutions without departing from the scope of the technical solutions of the present invention, and all such modifications or equivalent substitutions should be considered to fall within the protection scope of the present invention.

Claims

1. An ultra-compact combustion chamber for an aero-engine with a stepped intake structure, comprising an annular cavity (1), an outer fuel ring (5), an inner fuel ring (6), and an expansion-type mixing and afterburning structure (7), characterized in that: The annular cavity (1) is provided with a stepped outer ring air inlet (2) on the outside and an inner ring air inlet (3) on the inside. The stepped outer ring air inlet (2) has at least two stepped guide surfaces distributed along the axial direction. The outer fuel ring (5) is disposed on the rear wall of the annular cavity (1), and the inner fuel ring (6) is disposed on the front wall of the annular cavity (1); The expanded mixing and afterburning structure (7) is connected downstream of the annular cavity (1), with an external afterburning hole (8) and an external mixing hole (9) on its outer side, and an internal afterburning hole (14) and an internal mixing hole (15) on its inner side. The axes of the stepped outer ring air inlet (2) and the inner ring air inlet (3) are at an angle of 30°-40° to the radial direction of the annular cavity, and are evenly distributed circumferentially.

2. The combustion chamber according to claim 1, characterized in that: The stepped outer ring air inlet (2) is provided with an outer ring air inlet baffle (4) extending 3 mm into the inner side of the annular cavity.

3. The combustion chamber according to claim 1, characterized in that: The annular cavity (1) has a flame-stabilizing step (16) on its inner side, and the flame-stabilizing step has a height of 2.2 mm.

4. The combustion chamber according to claim 1, characterized in that: The annular cavity (1) has at least three combustion chamber mounting rods (12) fixedly connected to it on the front wall surface. The combustion chamber mounting rods extend along the axial direction of the annular cavity and are evenly spaced along the circumference of the annular cavity.

5. The combustion chamber according to claim 1, characterized in that: The inner fuel ring (6) is connected to the front wall of the annular cavity (1) via the inner fuel ring mounting bracket (11), and the outer fuel ring (5) is connected to the rear wall of the annular cavity (1) via the outer fuel ring mounting bracket (13).

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