Transverse navigation stabilizing structure, aircraft and control method

Abstract

The invention relates to the technical field of aviation aircrafts, and provides a transverse navigation stabilizing structure, an aircraft and a control method. The transverse navigation stabilizing structure comprises an aircraft rear body and a side fin assembly arranged on the side lower portion of the aircraft rear body. The side fin assembly can effectively block transverse airflow and restrain formation of a local low-pressure area on the windward side of a rear aircraft body of the aircraft, so that the negative yawing moment is reduced, the problem of transverse heading instability of the aircraft in the high-thrust and large-sideslip state is solved, and the flight safety is improved. Moreover, the side fin assemblies have a retractable function, can be unfolded to work in take-off, landing and re-flight stages requiring stability enhancement, and can be folded to form a continuous pneumatic surface in a cruise stage requiring low stability requirement and oil consumption reduction, so that the core contradiction between the stability enhancement device and the cruise low-resistance requirement is solved; and the fuel economy of the aircraft is obviously improved.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aviation aircraft. In particular, it relates to a lateral flight stability structure, an aircraft and a control method. BACKGROUND

[0002] Modern civil aircraft play a vital role in the field of aviation. With the continuous development of the aviation industry, the performance and safety of aircraft are increasingly concerned. During the take-off and landing phases, the aircraft needs to obtain sufficient lift by increasing the engine thrust and the flap deflection angle. However, in this process, the aerodynamic characteristics of the fuselage are easily disturbed by the lateral airflow, which significantly aggravates the lateral instability of the aircraft. At present, the mainstream passenger aircraft generally adopts a vertical tail structure to provide directional stability. However, in the high angle of attack state, the fuselage afterbody of the aircraft reduces the directional stability, and the increased flap deflection further destroys the flow field of the tail. For a specific type of aircraft, the traditional stability enhancement method does not achieve satisfactory results in practical application, so it is particularly urgent to develop a new type of aerodynamic compensation device.

[0003] In related technologies, three structural improvement schemes are mainly used to improve the lateral stability of the aircraft. The first is to increase the vertical tail area, which enhances the lateral stability of the aircraft by expanding the size of the vertical stabilizer. The second is to add a dorsal fin, that is, to install a fixed small aspect ratio wing on the back of the fuselage. The third is to install an abdominal fin, that is, to install a short wing-like structure on the belly of the aircraft. These three schemes all attempt to solve the problem of lateral stability of the aircraft to some extent, but they also have some limitations. Increasing the vertical tail area can theoretically enhance the stability, but the actual effect is often not obvious, and it will significantly increase the structural weight of the aircraft. Adding a dorsal fin also has poor effect, and it will also greatly increase the structural weight. Installing an abdominal fin is limited by the space of the landing gear compartment, and it is difficult to achieve the ideal size, and it will also greatly reduce the ground clearance angle of the aircraft. These schemes cannot effectively compensate for the stability of the specific aerodynamic layout of the aircraft, and they also have the common problems of significant weight increase and interference with the original design flow field.

[0004] These conventional means of related technologies have obvious defects. Increasing the vertical tail area will linearly increase the structural weight, which not only increases the manufacturing cost of the aircraft, but also consumes more fuel. Adding a dorsal fin will induce asymmetric flow separation, which will further aggravate the rolling instability of the aircraft. Installing an abdominal fin is restricted by the installation space, and its size cannot reach the theoretical optimal value, resulting in very weak improvement of the lateral stability. These defects make the existing stability enhancement means unable to fundamentally solve the problem of lateral instability of the aircraft in the high thrust state. SUMMARY

[0005] To solve the problem of lateral instability of the aircraft in a large-thrust state in the related art, the application provides a lateral stability structure, an aircraft and a control method.

[0006] In one aspect, the application provides a lateral stability structure, comprising: an aircraft rear body; a side fin assembly arranged below the sides of the aircraft rear body, the side fin assembly being used to block lateral airflow relative to the aircraft rear body to inhibit the formation of a low-pressure area on the aircraft rear body and reduce the negative yawing moment on the aircraft rear body.

[0007] The above technical solution can effectively block lateral airflow and inhibit the formation of a local low-pressure area on the windward side of the rear body, thereby reducing the negative yawing moment and solving the problem of lateral instability of the aircraft in a large-thrust, large-sideslip state, and improving flight safety.

[0008] Optionally, the side fin assembly comprises a first side fin plate arranged on a first side of the aircraft rear body and a second side fin plate arranged on a second side of the aircraft rear body opposite the first side.

[0009] The above technical solution can effectively deal with crosswinds from different directions and symmetrically improve the flow field on both sides of the rear body to ensure stable and balanced lateral stability gain under various sideslip conditions.

[0010] Optionally, the length of the side fin assembly is 1500mm to 3000mm, and the width of the side fin assembly is 150mm to 250mm.

[0011] The above technical solution can optimize the length and width of the side fin assembly to a specific range to provide the optimal aerodynamic action area and most efficiently generate stable vortexes on the premise of ensuring structural strength, rigidity and storage space, thereby maximizing the lateral stability of the aircraft during takeoff and landing.

[0012] Optionally, the deployment angle of the side fin assembly is 100° to 120°.

[0013] The above technical solution sets the deployment angle of the side fin assembly to the optimal range of 100° to 120°, which can most effectively disturb lateral airflow and generate the strongest stable vortexes, thereby most significantly inhibiting the negative yawing moment and improving the lateral damping coefficient, and avoiding the adverse effects of weakened stability improvement due to improper angles (too small or too large).

[0014] Optionally, the side fin assembly is connected to the rear fuselage of the aircraft via a movable connection; wherein, during the takeoff, landing and go-around phases of the aircraft, the side fin assembly deploys to block the lateral airflow relative to the rear fuselage of the aircraft, and during the cruise phase of the aircraft, the side fin assembly retracts to form a continuous aerodynamic surface on the surface of the rear fuselage of the aircraft.

[0015] By adopting the above technical solution, the side fin assembly is retractable, which can be deployed during takeoff, landing and go-around phases where enhanced stability is required, while it can be retracted to form a continuous aerodynamic surface during cruise phase where stability requirements are not high and fuel consumption needs to be reduced. This solves the core contradiction between the stability enhancement device and the requirement for low drag during cruise, and significantly improves the fuel economy of the aircraft.

[0016] Optionally, the side fin assembly is connected to the frame of the rear fuselage of the aircraft via a pivot. The rear fuselage of the aircraft is provided with a storage compartment for accommodating the retracted side fin assembly. The lateral stability structure also includes a drive mechanism, a locking device, and a position sensor. During the takeoff, landing, and go-around phases of the aircraft, the drive mechanism drives the side fin assembly to rotate around the pivot to deploy the side fin assembly. The position sensor monitors the deployment angle of the side fin assembly. When the side fin assembly deploys to the target deployment angle, the locking device locks the side fin assembly in place. During the cruise phase of the aircraft, the locking device is released, and the drive mechanism drives the side fin assembly to rotate around the pivot to retract the side fin assembly into the storage compartment. The shape of the side fin assembly is adapted to the surface shape of the rear fuselage of the aircraft, so that when the side fin assembly is retracted into the storage compartment, the outer surface of the side fin assembly and the surface of the rear fuselage of the aircraft together form a continuous aerodynamic surface.

[0017] By adopting the above technical solution, an automated and precise control of the deployment and retraction of the side fin assembly is achieved through a rotational deployment mechanism that integrates a drive mechanism, a position sensor, and a locking device. The locking device mechanically secures the side fin after deployment, eliminating the need for continuous work from the drive mechanism, thus reducing system energy consumption and ensuring the structural stability of the side fin in the deployed state, thereby improving flight reliability.

[0018] Optionally, the rear fuselage of the aircraft is provided with a sealing structure at the storage compartment, and the sealing structure is located at the junction of the outer surface of the fin side fin assembly and the surface of the rear fuselage of the aircraft when the fin side fin assembly is retracted.

[0019] By adopting the above technical solution and setting a sealing structure at the storage compartment, the gap between the outer surface of the side fin assembly and the surface of the fuselage can be effectively filled when the side fin assembly is retracted, ensuring the continuity and smoothness of the aerodynamic shape of the fuselage in the retracted state, thereby further reducing the cruise drag of the aircraft.

[0020] Optionally, the aircraft has a storage compartment located in the belly of the rear fuselage, and the transverse stabilization structure also includes a guide rail slider mechanism for driving the side fin assembly to slide along the guide rail direction; wherein, during the takeoff, landing and go-around phases of the aircraft, the guide rail slider mechanism drives the side fin assembly to extend from the storage compartment, and during the cruise phase of the aircraft, the guide rail slider mechanism drives the side fin assembly to retract into the storage compartment.

[0021] The above technical solution provides an alternative linear retraction and deployment method based on a guide rail slider mechanism. This solution allows the side fin assembly to adopt a simpler, easier-to-manufacture flat-plate design, providing flexible integration options and layout possibilities for aircraft with different airframe structures.

[0022] On the other hand, this application also provides an aircraft having the aforementioned lateral stabilization structure.

[0023] By adopting the above technical solution, the aircraft integrates this lateral stability structure, thereby having stronger lateral stability and flight safety during takeoff and go-around phases. At the same time, during the cruise phase, the retractable side fins result in lower fuel consumption and higher economy, comprehensively improving the overall performance of the aircraft.

[0024] In another invention, this application also provides an aircraft control method implemented based on the aforementioned lateral stabilization structure, the method comprising: S1. Determine the flight stage of the aircraft based on the airspeed signal and flap angle signal of the aircraft. S2. When it is determined that the aircraft is in the takeoff phase, when it is determined that the aircraft is in the landing phase, and when it is determined that the aircraft is in the go-around phase, control the side fin assembly to deploy to block the lateral airflow relative to the rear of the aircraft. S3. When it is determined that the aircraft is in the cruise phase, control the side fin assembly to retract so that the surface of the rear fuselage of the aircraft forms a continuous aerodynamic surface.

[0025] By adopting the above technical solution, an automated control logic was established that can intelligently determine the flight stage based on the aircraft's key parameters (airspeed, flap angle) and automatically control the deployment and retraction of the side fin components. This ensures that the lateral stability function is precisely activated during the most needed phases and reliably deactivated when not needed, realizing the intelligent and automated operation of the device and reducing the pilot's workload.

[0026] In summary, this application includes at least one of the following beneficial technical effects: 1. By installing side fin components on the lower side of the rear fuselage, the lateral airflow can be effectively blocked, the formation of a local low-pressure area on the windward side of the rear fuselage can be suppressed, thereby reducing the negative yaw moment, solving the problem of lateral instability of the aircraft under high thrust and large sideslip conditions, and improving flight safety.

[0027] 2. Optimizing the length, width, and thickness of the side fin assembly to a specific range can provide the optimal aerodynamic area while ensuring structural strength, stiffness, and storage space, and generate stable vortices most efficiently, thereby maximizing lateral stability during the takeoff and landing phases of the aircraft.

[0028] 3. Setting the deployment angle of the side fin assembly within the optimal range of 100° to 120° can most effectively disturb the lateral airflow and generate the strongest stable vortex, thereby most significantly suppressing the negative yaw moment and improving the directional damping coefficient, avoiding the adverse effects of weakened stability improvement due to improper angles (too small or too large).

[0029] 4. The side fin assembly is retractable, allowing it to operate during takeoff, landing, and go-around phases where enhanced stability is required, while retracting to form a continuous aerodynamic surface during cruise phases where stability requirements are lower and fuel consumption needs to be reduced. This resolves the core contradiction between the stability enhancement device and the low drag requirement during cruise, significantly improving the aircraft's fuel economy. Attached Figure Description

[0030] Figure 1 This is a schematic diagram of the structure of the aircraft according to an embodiment of this application; Figure 2 This is a schematic diagram of the transverse stabilization structure according to an embodiment of this application; Figure 3 This is a schematic diagram of the pressure distribution on the rear fuselage of an aircraft based on related technologies; Figure 4 This is a schematic diagram of the pressure distribution on the rear fuselage of the aircraft according to an embodiment of this application; Figure 5 This is a schematic diagram of the side fin assembly of the transverse stabilization structure in the retracted state according to an embodiment of this application; Figure 6 This is a perspective view of the side fin assembly of the transverse stabilization structure of another embodiment of this application in the retracted state; Figure 7 This is a flowchart illustrating the aircraft control method according to an embodiment of this application.

[0031] Explanation of reference numerals in the attached drawings: 10, rear fuselage of the aircraft; 20, side fin assembly; 21, first side fin; 22, second side fin. Detailed Implementation

[0032] The following is in conjunction with the appendix Figures 1-7 This application will be described in further detail.

[0033] This application discloses an aircraft.

[0034] Figure 1 This is a schematic diagram of the structure of an aircraft according to an embodiment of this application. (Refer to...) Figure 1 The aircraft has a lateral stabilization structure for improving the lateral stability of the aircraft, the lateral stabilization structure including the rear fuselage 10 of the aircraft and the side fin assembly 20.

[0035] Figure 2 This is a schematic diagram of a transverse stabilizing structure according to an embodiment of this application. (Refer to...) Figure 2 The transverse stabilization structure includes a rear fuselage 10 of the aircraft and a side fin assembly 20 disposed on the lower side of the rear fuselage 10. The side fin assembly 20 includes a first side fin 21 disposed on a first side of the rear fuselage 10 and a second side fin 22 disposed on a second side of the rear fuselage 10 opposite to the first side.

[0036] During takeoff, landing, and go-around phases, the rear fuselage 10 of the aircraft is prone to generating a negative yaw moment, which in turn affects the overall directional stability of the aircraft. Experimental verification on aircraft of related technologies and the aircraft of this application embodiment revealed that installing a side fin assembly 20 on the lower side of the rear fuselage 10 can effectively improve this phenomenon. The experimental conditions included flap deployment (simulating takeoff, landing, and go-around phases with flaps deployed), high slipstream intensity (simulating takeoff, landing, and go-around phases with the aircraft's engine using high thrust), and high sideslip (simulating takeoff, landing, and go-around phases with the aircraft encountering a right-front side wind).

[0037] Figure 3 This is a schematic diagram of the pressure distribution on the rear fuselage of an aircraft using related technologies. (Refer to...) Figure 3 During the takeoff, landing, and go-around phases of the aircraft, when there is crosswind, the downwash will increase due to the deployment of flaps and the increase in slipstream, resulting in a localized low-pressure area on the windward side of the rear fuselage 10 of the aircraft. Figure 3 The shaded area (as shown in the image) creates a negative yawing moment on the rear fuselage 10 of the aircraft, thus affecting the directional stability of the aircraft. Here, the term "low-pressure area" is a relative concept, referring to an area where the pressure is lower than other areas.

[0038] Figure 4 This is a schematic diagram of the pressure distribution on the rear fuselage of the aircraft according to an embodiment of this application. (Refer to...)Figure 4 Since a side fin assembly 20 is provided on the lower side of the rear fuselage 10 of the aircraft, the side fin assembly 20 can block part of the lateral airflow, thereby weakening the low-pressure area, reducing the negative yaw moment on the rear fuselage 10 of the aircraft, and increasing the directional stability of the aircraft.

[0039] To maximize the lateral stability of the aircraft under specific operating conditions, this application embodiment experimentally verifies the core dimensions and deployment angle of the side fin assembly 20 to achieve optimized design. The length of the side fin assembly 20 is preferably 1500mm to 3000mm to provide sufficient lateral aerodynamic area while ensuring structural feasibility and storage space, thereby effectively disturbing lateral airflow. The width of the side fin assembly 20 is preferably 150mm to 250mm to maintain a small frontal area to reduce drag in the retracted state while also considering the vortex generation efficiency in the deployed state. The thickness of the side fin assembly 20 is preferably 2mm to 20mm, aiming to balance structural strength, stiffness, and aerodynamic shape, reduce its own drag, and ensure no deformation under high-speed airflow. The deployment angle of the side fin assembly 20 (i.e. the angle between the side fin assembly 20 and the fuselage of the rear body 10 of the aircraft) is preferably 100° to 120°, with the optimal angle being 110°. This angle determines the degree of airflow disturbance and the intensity of vortices when the side fin is deployed.

[0040] The experimental verification process will be described in detail below.

[0041] The test conditions included flap deployment (simulating takeoff and go-around phases with the aircraft flaps deployed), high slipstream intensity (simulating takeoff and go-around phases with the aircraft's engine using high thrust), and high sideslip (simulating takeoff and go-around phases with the aircraft encountering a 10° right front wind).

[0042] The test measurement index is the yaw moment coefficient C. n and the heading damping coefficient C nβ Yaw moment coefficient C n This measure is used to assess the tendency of an aircraft to rotate about its vertical axis. A negative value indicates a negative yawing moment, causing the aircraft to yaw further in the sideslip direction, leading to directional instability. The closer the value is to zero or a positive value (in the sideslip direction), the better the directional stability. Heading damping coefficient C nβ This measure assesses an aircraft's ability to resist yaw motion. A larger absolute value indicates a stronger resistance to yaw oscillations and better directional stability. Eddy intensity or vortex distribution is visualized and quantified through CFD simulation or Particle Image Velocimetry (PIV) wind tunnel testing. Test data from comparative experiments are shown in Table 1.

[0043] Table 1 According to the test data, the control group data (without side fin assembly 20) shows that, under the calculated operating conditions, the aircraft's yaw moment coefficient is -0.012 and the directional damping coefficient is -0.0032. This indicates that without side fins, a local low-pressure area is generated on the windward side of the rear fuselage, resulting in a significant negative yaw moment. Consequently, the aircraft's directional stability is insufficient, making it susceptible to lateral airflow interference and deflection.

[0044] Compared to the control group, all test groups with side fins (test groups 1-4) showed positive yaw moment coefficients and positive directional damping coefficients. This means that the addition of side fins can effectively weaken the local low-pressure area on the windward side of the rear fuselage, thereby reducing (or even reversing) the negative yaw moment and increasing directional damping, significantly improving the aircraft's directional stability.

[0045] According to the test data, when the side fin length is 3000mm, width is 200mm, and thickness is 10mm, the deployment angle between 100° and 120° shows a significant improvement in directional stability. Test group 2 (side fin length 3000mm, width 200mm, thickness 10mm, deployment angle 110°) achieved the best results: the yaw moment coefficient reached 0.043, which is the largest positive value among all test groups, indicating that it generated the largest positive yaw moment, effectively offsetting the unfavorable negative yaw moment. The directional damping coefficient reached 0.0042, which is the largest positive value in absolute terms among all test groups, showing that the aircraft has the strongest ability to resist yaw oscillations and significantly enhanced directional stability.

[0046] The yaw moment coefficient of test group 1 (100° deployment angle) was 0.028, and the directional damping coefficient was 0.0027. The yaw moment coefficient of test group 3 (130° deployment angle) was 0.034, and the directional damping coefficient was 0.0034. This indicates that when the deployment angle deviates from 110°, the improvement in directional stability decreases, confirming that approximately 110° is the optimal deployment angle for this length and width. This angle allows the fins to optimally generate stable vortices when disturbing lateral airflow, thereby effectively weakening the low-pressure area at the rear of the fuselage.

[0047] Test group 4 (side fin length 2500mm, width 200mm, thickness 10mm, deployment angle 110°) had a yaw moment coefficient of 0.037 and a directional damping coefficient of 0.0036. While the results were good, the yaw moment coefficient and directional damping coefficient were slightly lower compared to test group 2 (3000mm in length). This indicates that within the preferred length range (1500mm to 3000mm), a length of 3000mm (under otherwise identical conditions) provides superior directional stability enhancement.

[0048] In conclusion, the experimental data strongly demonstrates the crucial role of the optimal side fin length (1500mm to 3000mm, with 3000mm being particularly preferred), width (200mm), and deployment angle (100° to 120°, with 110° being particularly preferred) in improving the aircraft's lateral stability. This combination of optimal parameters can maximally reduce the local low-pressure area on the windward side of the rear fuselage, decrease the negative yawing moment, and significantly increase directional damping, thereby effectively improving the aircraft's directional stability during critical phases such as takeoff / go-around and ensuring flight safety.

[0049] Furthermore, in a preferred embodiment of this application, the side fin assembly 20 can be retracted during the cruise phase of the aircraft, or other phases where the requirements for the directional stability of the aircraft are not high, to form a continuous aerodynamic surface with the surface of the rear fuselage 10 of the aircraft, thereby reducing drag and improving fuel economy.

[0050] Figure 5 This is a schematic diagram of the side fin assembly of the transverse stabilization structure according to an embodiment of this application in its retracted state. (Refer to...) Figure 2 and Figure 5 The side fin assembly 20 (the first side fin 21 and the second side fin 22) can be movably connected to the rear fuselage 10 of the aircraft. For example, the side fin assembly 20 may be provided with a pivot, and the side fin assembly 20 is connected to the frame of the rear fuselage 10 of the aircraft via the pivot. The lateral stabilization structure also includes a drive mechanism, a locking device, and a position sensor. The rear fuselage 10 of the aircraft is provided with a storage compartment for accommodating the side fin assembly 20 in its retracted state.

[0051] During takeoff, landing, and go-around phases of the aircraft, the drive mechanism drives the side fin assembly 20 to rotate around the pivot axis, causing the side fin assembly 20 to deploy. The position sensor monitors the deployment angle of the side fin assembly 20 in real time. When the side fin assembly 20 deploys to the target deployment angle, the locking device locks and secures the side fin assembly 20. During the cruise phase of the aircraft, the locking device is released, and the drive mechanism drives the side fin assembly 20 to rotate around the pivot axis, causing the side fin assembly 20 to retract into the storage compartment. The shape of the side fin assembly 20 needs to be adapted to the surface shape of the rear fuselage 10 of the aircraft, so that when the side fin assembly 20 is retracted into the storage compartment, the outer surface of the side fin assembly 20 can form a continuous aerodynamic surface with the surface of the rear fuselage 10 of the aircraft.

[0052] The rotating shaft can be made of 30CrMnSiA alloy steel with quenching treatment and a suitable hard chrome plating thickness (e.g., 0.05mm). This treatment improves the shaft's hardness, wear resistance, and corrosion resistance. Alternatively, titanium alloy can be used, which has advantages such as low density, high strength, and good corrosion resistance. The drive mechanism can be connected between the reinforcing ribs of the side fin assembly 20 and the fuselage ribs of the rear fuselage 10 of the aircraft. The drive mechanism can be a linear actuator, preferably an electro-hydraulic servo type, which can precisely control the deployment and retraction of the fins. When the fins need to be deployed, the linear actuator extends, pushing the side fin assembly 20 to rotate around the rotating shaft; when the side fin assembly 20 needs to be retracted, the linear actuator retracts, driving the side fin assembly 20 back into the storage compartment. Alternatively, a ball screw electric actuator can be used instead, which has advantages such as high transmission efficiency, high positioning accuracy, and good reliability.

[0053] The locking device includes an electromagnetically driven locking pin, which inserts into a preset positioning hole to achieve mechanical locking when the side fin assembly 20 reaches the target position. When the side fin assembly 20 unfolds to the target unfolding angle, the locking pin, driven by electromagnetic force, inserts into the preset positioning hole to fix the side fin assembly 20 in place, preventing the side fin assembly 20 from swaying during flight and affecting the stability of the aircraft. The holding force of the locking pin must meet certain requirements, such as a holding force greater than or equal to 5kN, to ensure the reliability of the locking.

[0054] The rear fuselage 10 of the aircraft is equipped with a sealing structure at the storage compartment. This sealing structure is located at the junction of the outer surface of the fin-side fin assembly 20 and the surface of the rear fuselage 10 when the fin-side fin assembly 20 is retracted. The sealing structure can be a rubber sealing strip, which has good elasticity and sealing performance, effectively ensuring the continuity of the aerodynamic surface in the retracted state. Alternatively, the rubber sealing strip can be replaced with silicone, which has better high-temperature resistance and aging resistance.

[0055] Figure 6 This is a perspective view of the side fin assembly of a transverse stabilizing structure according to another embodiment of this application in its retracted state. (Refer to...) Figure 2 and Figure 6In addition to being connected to the rear fuselage 10 of the aircraft via a rotatable connection, the side fin assembly 20 can also be deployed and retracted using other methods. For example, the rear fuselage 10 can have a storage compartment located within the fuselage, and the transverse stabilization structure further includes a guide rail slider mechanism that drives the side fin assembly 20 to slide along the guide rail. During takeoff, landing, and go-around phases, the guide rail slider mechanism drives the side fin assembly 20 to extend from the storage compartment. During cruise phase, the guide rail slider mechanism drives the side fin assembly 20 to retract into the storage compartment. The storage compartment can have an inflatable sealing strip. When the side fin assembly 20 is retracted, the inflatable sealing strip inflates and expands to provide a seal; when the side fin assembly 20 is deployed, the inflatable sealing strip deflates and contracts, without affecting the movement of the side fin assembly 20. In this case, the shape of the side fin assembly 20 does not need to conform to the shape of the surface of the rear fuselage 10 of the aircraft. It can be constructed as a flat plate so that it can extend and retract from the storage compartment. The structure of the flat plate side fin assembly 20 is also simpler and easier to manufacture and install.

[0056] The implementation principle of the lateral stabilization structure in this application embodiment is as follows: By setting the side fin assembly 20 below the rear fuselage 10 of the aircraft, it can effectively block the lateral airflow and suppress the formation of a local low-pressure area on the windward side of the rear fuselage, thereby reducing the negative yaw moment and solving the problem of lateral instability of the aircraft under high thrust and large sideslip conditions, thus improving flight safety. Furthermore, the side fin assembly 20 is retractable, allowing it to operate during takeoff and go-around phases where enhanced stability is required, while retracting to form a continuous aerodynamic surface during cruise phases where stability requirements are lower and fuel consumption needs to be reduced. This resolves the core contradiction between the stability enhancement device and the requirement for low drag during cruise, significantly improving the aircraft's fuel economy.

[0057] This application also discloses an aircraft control method based on the aforementioned lateral stabilization structure.

[0058] Figure 7 This is a flowchart illustrating an aircraft control method according to an embodiment of this application. (Refer to...) Figure 7 The method includes the following steps: S1. Determine the flight stage of the aircraft based on the airspeed signal and flap angle signal.

[0059] S2. When it is determined that the aircraft is in the takeoff phase, the aircraft is in the landing phase, and the aircraft is in the go-around phase, the side fin assembly 20 is controlled to deploy to block the lateral airflow relative to the rear fuselage 10 of the aircraft.

[0060] S3. When it is determined that the aircraft is in the cruise phase, control the side fin assembly 20 to retract so that the surface of the rear fuselage 10 of the aircraft forms a continuous aerodynamic surface.

[0061] The implementation principle of the aircraft control method in this application embodiment is as follows: an automated control logic is established that can intelligently determine the flight stage based on the aircraft's key parameters (airspeed, flap angle) and automatically control the deployment and retraction of the side fin assembly 20. This ensures that the lateral stability function is precisely activated during the most needed phase and reliably deactivated when not needed, realizing the intelligent and automated operation of the device and reducing the pilot's workload.

[0062] The above are all preferred embodiments of this application, and are not intended to limit the scope of protection of this application. Therefore, all equivalent changes made in accordance with the structure, shape and principle of this application should be covered within the scope of protection of this application.

Claims

1. A transverse stabilizing structure, characterized in that, include: Rear fuselage of the aircraft (10); Side fin assembly (20) is disposed on the lower side of the rear fuselage (10) of the aircraft. The side fin assembly (20) is used to block the lateral airflow relative to the rear fuselage (10) of the aircraft to suppress the low-pressure area formed on the rear fuselage (10) of the aircraft and reduce the negative yaw moment on the rear fuselage (10).

2. The transverse stabilizing structure according to claim 1, characterized in that, The side fin assembly (20) includes a first side fin (21) disposed on a first side of the rear fuselage (10) of the aircraft and a second side fin (22) disposed on a second side of the rear fuselage (10) of the aircraft opposite to the first side.

3. The transverse stabilizing structure according to claim 1, characterized in that, The side fin assembly (20) has a length of 1500 mm to 3000 mm and a width of 150 mm to 250 mm.

4. The transverse stabilizing structure according to claim 1, characterized in that, The deployment angle of the side fin assembly (20) is 100° to 120°.

5. The transverse stabilizing structure according to claim 1, characterized in that, The side fin assembly (20) is connected to the rear fuselage (10) of the aircraft via a movable connection; wherein, during the takeoff, landing and go-around phases of the aircraft, the side fin assembly (20) is deployed to block the lateral airflow relative to the rear fuselage (10), and during the cruise phase of the aircraft, the side fin assembly (20) is retracted to form a continuous aerodynamic surface on the surface of the rear fuselage (10).

6. The transverse stabilizing structure according to claim 5, characterized in that, The side fin assembly (20) is connected to the frame of the rear fuselage (10) of the aircraft via a pivot. The rear fuselage (10) of the aircraft is provided with a storage compartment for accommodating the retracted side fin assembly (20). The lateral stability structure also includes a drive mechanism, a locking device, and a position sensor. During the takeoff, landing, and go-around phases of the aircraft, the drive mechanism drives the side fin assembly (20) to rotate around the pivot to deploy the side fin assembly (20). The position sensor monitors the deployment angle of the side fin assembly (20). When the side fin assembly (20) deploys... When the target deployment angle is reached, the locking device locks and fixes the side fin assembly (20); during the cruise phase of the aircraft, the locking device is unlocked, and the drive mechanism drives the side fin assembly (20) to rotate around the pivot, so that the side fin assembly (20) is retracted into the storage compartment. The shape of the side fin assembly (20) is adapted to the surface shape of the rear fuselage (10) of the aircraft, so that when the side fin assembly (20) is retracted into the storage compartment, the outer surface of the side fin assembly (20) and the surface of the rear fuselage (10) of the aircraft together form a continuous aerodynamic surface.

7. The transverse stabilizing structure according to claim 6, characterized in that, The rear fuselage (10) of the aircraft is provided with a sealing structure at the storage compartment. The sealing structure is located at the junction of the outer surface of the fin side fin assembly (20) and the surface of the rear fuselage (10) of the aircraft when the fin side fin assembly (20) is retracted.

8. The transverse stabilizing structure according to claim 5, characterized in that, The rear fuselage (10) of the aircraft is provided with a storage compartment located in the belly of the fuselage. The transverse stabilization structure also includes a guide rail slider mechanism, which is used to drive the side fin assembly (20) to slide along the guide rail direction. During the takeoff, landing and go-around phases of the aircraft, the guide rail slider mechanism drives the side fin assembly (20) to extend from the storage compartment. During the cruise phase of the aircraft, the guide rail slider mechanism drives the side fin assembly (20) to retract into the storage compartment.

9. An aircraft, characterized in that, It has a transverse stabilizing structure as described in any one of claims 1-8.

10. An aircraft control method, characterized in that, Implemented based on the transverse stabilizing structure as described in any one of claims 1-8, the method includes: S1. Determine the flight stage of the aircraft based on the airspeed signal and flap angle signal of the aircraft. S2. When it is determined that the aircraft is in the take-off phase, when it is determined that the aircraft is in the landing phase, and when it is determined that the aircraft is in the go-around phase, the side fin assembly (20) is controlled to deploy to block the lateral airflow relative to the rear fuselage (10) of the aircraft. S3. When it is determined that the aircraft is in the cruise phase, control the side fin assembly (20) to retract so that the surface of the rear body (10) of the aircraft forms a continuous aerodynamic surface.

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