An anti-icing device for aircraft engine air intake and a weather detection aircraft

Abstract

This utility model belongs to the field of engine de-icing, and discloses an anti-icing device for an aircraft engine inlet and a weather detection aircraft. It includes a partition member disposed within an annular cavity inside the lip of the engine inlet. The partition member divides the annular cavity circumferentially into a heating chamber one covering the stagnation area of ​​the lip and a heating chamber two covering the rear sections on both sides of the stagnation area. It also includes an air bleed pipe connected to the engine to draw out compressed hot air, and an oblique nozzle connecting the air bleed pipe and the heating chamber one. This utility model can achieve heating and de-icing of other areas while primarily heating and de-icing the stagnation area, better distributing and utilizing compressed hot air. Furthermore, through the cooperation of the oblique nozzle and the acceleration channel, the compressed hot air can achieve a higher flow velocity within the narrower acceleration channel, thus filling the entire heating chamber one more quickly and making the heating of the lip more uniform in the circumferential direction.

Description

Technical Field

[0001] This utility model belongs to the field of engine de-icing technology, specifically relating to an anti-icing device for an aircraft engine air intake and a meteorological detection aircraft. Background Technology

[0002] During flight, the leading edge of the engine's air intake lip is directly opposite the direction of the incoming airflow. After impact, the airflow velocity drops instantly to 0 m / s. Because the impact rate of supercooled water droplets is the highest (reaching over 60% of the total impact), the ice layer in this area is the thickest (typically 10-15 mm), and it easily forms dense, transparent "bi-shaped clear ice." The thickness of the ice layer decreases from this area towards the sides and rear of the lip. Ice forming on the lip can alter the aircraft's aerodynamic characteristics and increase its weight. It can also easily detach and enter the air intake, potentially causing serious safety incidents.

[0003] In one existing technical solution, hot compressed air is drawn from the turbine engine through a pipeline and input into the annular cavity of the intake lip. The hot compressed air heats the stagnation area and adjacent areas within the annular cavity, thus achieving de-icing. However, because the ice layer on the lip is gradually distributed, with a thicker ice layer in the stagnation area, and the annular cavity is relatively large, the hot compressed air is evenly distributed upon entry, requiring more air to clear the ice layer behind the stagnation area. This makes it impossible to target the stagnation area for effective de-icing. Furthermore, due to the large length and internal space of the annular cavity in the circumferential direction, it takes time for the compressed hot air to reach its radial position after entering through the inlet. Additionally, because the gas input is perpendicular to the circumferential plane, the gas flow velocity is affected. This results in uneven heating of the lip in the circumferential direction, leading to poor de-icing performance. Utility Model Content

[0004] In view of this, the purpose of this utility model is to provide an anti-icing device for an aircraft engine intake and a weather detection aircraft, so as to solve the problem that the existing anti-icing devices have poor de-icing effect due to uneven heating of the engine intake lip.

[0005] To achieve the above objectives, the present invention adopts the following technical solution:

[0006] On the one hand, an anti-icing device for an aircraft engine air intake is provided, comprising:

[0007] A partition member is provided in an annular cavity inside the lip of the engine intake manifold. The partition member divides the annular cavity into a heating cavity one that covers the stagnation area of ​​the lip along the circumferential direction and a heating cavity two that covers the rear sections of both sides of the stagnation area along the circumferential direction.

[0008] Bleed air lines connected to the engine to draw compressed hot air; and

[0009] An oblique nozzle connecting the air intake pipe and the heating chamber one;

[0010] The separating component includes an acceleration channel facing the stagnation area of ​​the lip and flared partitions on both sides of the acceleration channel. The acceleration channel is distributed circumferentially along the separating component to allow compressed hot air to flow circumferentially. The flared partitions are provided with holes that connect heating chamber one and heating chamber two. The oblique nozzle passes through the acceleration channel and has an inclination angle α of 5-30° with the plane where the acceleration channel is located in the circumferential direction.

[0011] In possible implementations, the flared partition is a straight plate structure or an arc-shaped plate structure, and the flared partition is fitted or fixedly connected to the inner wall of the annular cavity.

[0012] In one possible implementation, the inner surface of the acceleration channel is a smoothly transitioning arc-shaped structure.

[0013] In one possible implementation, multiple oblique nozzles are provided circumferentially within the acceleration channel, and each oblique nozzle is inclined circumferentially.

[0014] In one possible implementation, the air intake line is equipped with a control valve and a flow meter.

[0015] In one possible implementation, the separating member is made of a continuous fiber-reinforced resin-based composite material.

[0016] In one possible implementation, a reinforcing connection is provided between the outer wall of the acceleration channel and the flared partition.

[0017] On the other hand, a weather detection aircraft is also provided, including a weather detection system, an airborne control system, and an anti-icing device for the air intake of an aircraft engine as described in any of the above technical solutions. The weather detection system includes an atmospheric temperature sensor, and the atmospheric temperature sensor and the bleed air duct are simultaneously connected to the airborne control system so as to control the airflow or de-icing of the bleed air duct through atmospheric temperature.

[0018] Compared with the prior art, the present invention has the following beneficial effects:

[0019] This invention relates to an anti-icing device for an aircraft engine inlet. Through a separating component, it provides two heating chambers: one covering the stagnation area and the other covering the rear sections on both sides of the stagnation area. An angled nozzle inserted into the first heating chamber primarily heats and de-ices the stagnation area with compressed hot air. A smaller portion of the compressed hot air from the first heating chamber enters the second heating chamber through a perforation to heat the rear sections on both sides of the stagnation area. This allows for simultaneous heating and de-icing of the stagnation area and other areas, resulting in better distribution and utilization of the compressed hot air. Furthermore, the combination of the angled nozzle and the acceleration channel allows for a higher flow velocity of the compressed hot air within the narrower acceleration channel, enabling faster filling of the first heating chamber and a more uniform circumferential distribution of the compressed hot air. This, in turn, achieves uniform heating of the lip. The angled nozzle's angled arrangement also allows the compressed hot air to enter the first heating chamber from the acceleration channel, preventing stagnation of the compressed hot air within the acceleration channel.

[0020] The meteorological observation aircraft of this invention, through its anti-icing device, can cope with de-icing and anti-icing in complex and severe weather environments such as high-altitude areas, thus better solving the anti-icing problem. Attached Figure Description

[0021] Figure 1 A schematic cross-sectional view of an anti-icing device for an aircraft engine air intake.

[0022] Figure 2 A cross-sectional view of one embodiment of a partition component for an anti-icing device for an aircraft engine air intake;

[0023] Figure 3 A cross-sectional view of another embodiment of a partition component of an anti-icing device for an aircraft engine inlet;

[0024] Figure 4 for Figure 3 A front view of the partition member of the structure shown in the illustration;

[0025] Figure 5 for Figure 3 A perspective view of the implemented structure shown;

[0026] Figure 6 This is a schematic diagram showing the angle between the inclined nozzle and the inner surface of the acceleration channel;

[0027] Figure 7 This is a schematic diagram of the control principle of a weather detection aircraft.

[0028] In the diagram: 1-partition; 2-heating chamber one; 3-heating chamber two; 4-separation component; 41-flared separation part; 411-hole part; 42-acceleration flow channel; 43-reinforced connection part; 5-air intake pipe; 51-flow meter; 52-control valve; 6-engine; 7-slanted nozzle; 8-meteorological detection system; 81-atmospheric temperature sensor; 9-airborne control system. Detailed Implementation

[0029] To enable those skilled in the art to better understand the technical solution of this utility model, the present utility model will be further described in detail below with reference to specific embodiments.

[0030] Please refer to Figure 1-6 As shown, an embodiment of this application provides an anti-icing device for an aircraft engine inlet, comprising: a partition member 4, which is disposed in an annular cavity inside the lip of the engine 6 inlet, the partition member 4 dividing the annular cavity into a heating cavity 2 covering the stagnation area of ​​the lip along the circumferential direction and a heating cavity 3 covering the rear sections of both sides of the stagnation area along the circumferential direction; an air bleed pipe 5 connected to the engine 6 to draw out compressed hot air; an oblique nozzle 7 connecting the air bleed pipe 5 and the heating cavity 2; and wherein the partition member 4 includes an acceleration flow channel 42 facing the stagnation area of ​​the lip and flared partition portions 41 disposed on both sides of the acceleration flow channel 42, the acceleration flow channel 42 being distributed circumferentially along the partition member 4 to allow compressed hot air to flow circumferentially, the flared partition portions 41 having a hole 411 connecting the heating cavity 2 and the heating cavity 3, the oblique nozzle 7 passing through the acceleration flow channel 42 and having an inclined angle α of 5-30° with the plane where the acceleration flow channel 42 is located in the circumferential direction.

[0031] The bleed air duct 5 of the aircraft engine inlet anti-icing device can be connected to the core of the engine 6 and is used to draw out compressed hot air from it. This compressed hot air is used to heat and de-ice the inlet lip. An annular cavity is provided inside the inlet lip, and a partition member 4 is disposed within the annular cavity. The partition member 4 divides the space within the annular cavity into heating cavity 2 and heating cavity 3. Heating cavity 2 covers the stagnation area of ​​the lip in a circumferential direction. This stagnation area is a region including the apex position and the area adjacent to the apex, and its cross-sectional structure is roughly V-shaped. Since this region is the main area for icing, effective coverage of the stagnation area can be achieved by covering the apex position and the areas on both sides of the stagnation area with heating cavity 2. Heating cavity 3 is used to heat the rearward area of ​​the stagnation area, that is, the rear section on both sides of the stagnation area. Heating cavity 3 can heat this area, thereby heating and de-icing or preventing icing of the ice layer at the edge. The oblique nozzle 7 is connected to the output end of the induction pipeline and passes through the acceleration channel 42. The acceleration channel 42 is a narrow channel. After the compressed hot air enters the acceleration channel 42 through the oblique nozzle 7, it will have a greater flow velocity than the flow in the annular cavity. Under the action of the greater flow velocity, the compressed hot air can flow more evenly and quickly in the annular direction, avoiding the problem of reduced flow velocity due to obstruction when input perpendicular to the annular plane. This allows the compressed hot air filled in the heating chamber 2 to heat, de-ice, and prevent icing of the stagnation area more effectively and evenly. Furthermore, by configuring the oblique nozzle 7 with an inclination angle α of 5-30° with the plane of the acceleration channel 42 in the annular direction, the compressed hot air can enter the heating chamber 2 with a shorter residence time after being accelerated in the acceleration channel 42, achieving effective filling and acceleration. In this way, the compressed hot air can be distributed evenly in the annular direction more quickly, which is conducive to achieving uniform heating. The flared partition 41 connects heating chamber 2 and heating chamber 3 via its perforation 411. Excess compressed hot air entering heating chamber 2 enters heating chamber 3 to heat and de-ice the rear sections on both sides of the stagnation area. Furthermore, the flared structure of the partition 41 guides the compressed hot air to effectively cover the stagnation area. In specific implementations, an exhaust port 1 communicating with heating chamber 2 is also provided at the lip, and an exhaust port 2 for easy venting can also be provided in heating chamber 3. This facilitates continuous heating and de-icing.

[0032] Through the above technical solution, the partition member 4 can divide the space into two heating chambers that can cover the stagnation area and the rear sections on both sides of the stagnation area. The oblique nozzle 7, which passes through the heating chamber 1 2, can mainly heat and de-ice the stagnation area with compressed hot air. A smaller portion of the compressed hot air in the heating chamber 1 2 enters the heating chamber 2 3 through the hole 411 to heat the rear sections on both sides of the stagnation area. In this way, while mainly heating and de-icing the stagnation area, heating and de-icing of other areas can be achieved, thus better distributing and utilizing the compressed hot air. Furthermore, through the cooperation of the oblique nozzle 7 and the acceleration channel 42, the compressed hot air can have a higher flow velocity in the narrower acceleration channel 42, thereby filling the entire heating chamber 1 2 more quickly. The oblique arrangement of the oblique nozzle 7 also allows the compressed hot air to enter the heating chamber 1 2 from the acceleration channel 42, avoiding the stagnation of the compressed hot air in the acceleration channel 42.

[0033] In one embodiment, the flared partition 41 is a straight plate structure or an arc-shaped plate structure, and the flared partition 41 is fitted or fixedly connected to the inner wall of the annular cavity.

[0034] In this way, the flared partition 41 can adopt a straight plate structure or an arc-shaped plate structure, which can better cover the stagnation area and guide it to both sides of the stagnation area positioning position. In the specific implementation process, the partition member 4 can be connected to the inside of the annular cavity by fitting together or by fixing. When fixing, it can be fixed by fasteners and adhesives. The acceleration channel 42 can be supported by the partition 1 inside the intake channel, and it can also have a connection with the partition 1 for reinforcement and fixation, without limitation.

[0035] To facilitate the better entry of compressed hot air from the acceleration channel 42 into the heating chamber 2, the inner surface of the acceleration channel 42 is further characterized by a smoothly transitioning arc-shaped structure.

[0036] In this way, when the compressed hot air flows along the acceleration channel 42, the airflow can be guided outward by the smoothly transitioning arc structure, so that it can enter the heating chamber 2 better.

[0037] In other embodiments, multiple oblique nozzles 7 are arranged circumferentially within the acceleration channel 42, each oblique nozzle 7 being inclined circumferentially. Two to four oblique nozzles 7 can be arranged at equal intervals circumferentially, thereby further improving the uniformity and flow of compressed hot air in the heating chamber 2.

[0038] In specific implementation, the air intake pipe 5 may be equipped with a control valve 52 and a flow meter 51. The control valve 52 is used to control the opening and closing of the air intake pipe 5, and the flow meter 51 is used to detect the flow rate of compressed hot air entering the annular cavity through the air intake pipe 5, so as to understand the flow status.

[0039] In one embodiment, the separating member 4 is made of a continuous fiber-reinforced resin-based composite material.

[0040] Continuous fiber reinforced resin matrix composites exhibit excellent performance in terms of thermal insulation, strength, and high-temperature resistance. The thermal insulation allows the heating chamber 2 to more effectively heat and de-ice the stagnation zone; the good strength improves its structural stability within the annular cavity; and the good high-temperature resistance allows it to withstand the high temperatures of compressed hot air. Furthermore, it is lighter than metal materials. In practical implementation, the fibers in the continuous fiber reinforced resin matrix composites can be carbon fiber or quartz fiber.

[0041] To further improve the structural stability between the flared partition 41 and the acceleration channel 42, a reinforcing connection 43 is provided between the outer wall of the acceleration channel 42 and the flared partition 41.

[0042] The plateau's main environment is harsh, drastically changing, and difficult to survive in. Furthermore, the surrounding terrain is steep, and the climate is complex and diverse, which makes plateau meteorological land observations face many problems and difficulties.

[0043] Currently, weather observation aircraft that routinely carry meteorological monitoring payloads experience increased icing rates at their air intake lips under adverse weather conditions. For example, in high-altitude environments, heating, de-icing, and anti-icing measures are particularly important in their lodging areas.

[0044] Please combine Figure 7 As shown, embodiments of this application also provide a weather detection aircraft, including a weather detection system 8, an airborne control system 9, and an anti-icing device for the air intake of an aircraft engine as described in any of the above technical solutions. The weather detection system 8 includes an atmospheric temperature sensor 81, and the atmospheric temperature sensor 81 and the bleed air duct 5 are simultaneously connected to the airborne control system 9 to control the airflow or de-icing of the bleed air duct 5 by means of atmospheric temperature.

[0045] The meteorological detection system 8 and the airborne control system 9 are systems inherent to the meteorological aircraft. The meteorological detection system 8 is used to detect meteorological information, including but not limited to meteorological data detected by the atmospheric temperature sensor 81. The airborne control system 9 is used to control the opening and closing of the bleed air pipeline 5 based on the atmospheric temperature data collected by the atmospheric temperature sensor 81, thereby controlling the opening or closing of the anti-icing device in the aircraft engine intake. In specific implementation, the atmospheric temperature sensor 81 is located on the outside of the aircraft fuselage to detect the ambient atmospheric temperature. The control valve 52 is preferably a flow regulating valve, whose opening degree can be adjusted. The correspondence between atmospheric temperature and compressed hot air can be determined by simulation calculations or actual experiments, thereby controlling the opening degree of the flow regulating valve. Of course, the control valve 52 can also be a valve with only opening and closing control functions. The airborne control system 9 sets a preset value for atmospheric temperature. When the atmospheric temperature is higher than the preset value, the bleed air pipeline 5 is controlled to circulate air for heating and de-icing or heating and anti-icing; when it is lower than the preset value, the airflow is stopped. The above two implementation methods can be selected and configured according to actual needs without limitation.

[0046] The above are merely preferred embodiments of this utility model. It should be noted that the above preferred embodiments should not be considered as limitations on this utility model, and the scope of protection of this utility model should be determined by the scope defined in the claims. For those skilled in the art, several improvements and modifications can be made without departing from the spirit and scope of this utility model, and these improvements and modifications should also be considered within the scope of protection of this utility model.

Claims

1. An anti-icing device for an aircraft engine air intake, characterized in that, include: The partition member (4) is located in the annular cavity inside the intake lip of the engine (6). The partition member (4) divides the annular cavity into a heating cavity one (2) that covers the stagnation area of ​​the lip along the circumferential direction and a heating cavity two (3) that covers the rear sections of both sides of the stagnation area along the circumferential direction. A bleed air line (5) connected to the engine (6) to draw compressed hot air; and An oblique nozzle (7) connects the air intake pipe (5) and the heating chamber (2); The partition member (4) includes an acceleration channel (42) facing the stagnation area of ​​the lip and a flared partition (41) on both sides of the acceleration channel (42). The acceleration channel (42) is distributed circumferentially along the partition member (4) to allow compressed hot air to flow circumferentially. The flared partition (41) is provided with a hole (411) that connects heating chamber one (2) and heating chamber two (3). The oblique nozzle (7) passes through the acceleration channel (42) and has an inclination angle α of 5-30° with the plane where the acceleration channel (42) is located in the circumferential direction.

2. The anti-icing device for an aircraft engine air intake as described in claim 1, characterized in that, The flared partition (41) is a straight plate structure or an arc plate structure, and the flared partition (41) is fitted or fixedly connected to the inner wall of the annular cavity.

3. The anti-icing device for an aircraft engine air intake as described in claim 1, characterized in that, The inner surface of the acceleration channel (42) is a smoothly transitioned arc-shaped structure.

4. The anti-icing device for an aircraft engine air intake as described in claim 1, characterized in that, Multiple oblique nozzles (7) are provided circumferentially within the acceleration channel (42), and each oblique nozzle (7) is inclined circumferentially.

5. The anti-icing device for an aircraft engine air intake as described in claim 1, characterized in that, The air intake pipeline (5) is equipped with a control valve (52) and a flow meter (51).

6. The anti-icing device for an aircraft engine air intake as described in claim 1, characterized in that, The separating member (4) is made of continuous fiber reinforced resin matrix composite material.

7. The anti-icing device for an aircraft engine air intake as described in claim 1, characterized in that, A reinforcing connection (43) is provided between the outer wall of the acceleration channel (42) and the flared partition (41).

8. A weather observation aircraft, characterized in that, The system includes a weather detection system (8), an airborne control system (9), and an anti-icing device for an aircraft engine air intake as described in any one of claims 1-7. The weather detection system (8) includes an atmospheric temperature sensor (81), and the atmospheric temperature sensor (81) and the bleed air line (5) are simultaneously connected to the airborne control system (9) to control the bleed air line (5) to circulate or cut off air by means of atmospheric temperature.

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