Visual expression method, system and equipment of flight ability feasible region and medium

Abstract

The invention relates to a visual expression method, system and device for a flight ability feasible region and a medium, and the method comprises the steps: carrying out the trimming of an airplane under a specified crosswind condition, and enabling the airplane to be kept in a balanced state; selecting all high-dimensional points when the aircraft is kept in a balanced state according to given limiting conditions, and reserving generalized state points meeting the requirements of the limiting conditions; and taking a region formed by all generalized state points as a flight ability feasible region, and performing visual expression. The method is suitable for evaluating the anti-crosswind capability in the landing process of the fixed-wing aircraft and estimating the flight feasible region in the crosswind environment, and has the advantages that various limiting factors are visually expressed, the limiting factors are extracted, designers can specifically optimize the anti-crosswind performance of the aircraft, and risk evaluation in the flight process is facilitated.

Description

Technical Field

[0001] The present invention belongs to the technical field of automatic aircraft landing, and in particular relates to a method, system, device and medium for visually expressing a feasible domain of flight capability. Background Art

[0002] Crosswinds can have a great impact on the landing process of an aircraft. Low-altitude wind shear and turbulent wind fields may both produce crosswinds. Crosswinds can bring about a sideslip angle. Due to the aircraft's own heading static stability (weather vane stability), the aircraft has a tendency to reduce the sideslip angle, which means that the aircraft's symmetry plane (nose pointing) has a tendency to be consistent with the aircraft's airspeed, and the aircraft's symmetry plane is no longer aligned with the airport runway. In this process, the crosswind will generate a large side force, so the rudder surface needs to be controlled, otherwise the aircraft will deviate from the original trajectory. Therefore, in a crosswind environment, the pilot needs to correct the impact of the crosswind. There are currently three commonly used correction methods: the sideslip correction method, the crab correction method, and the combination of the two. From the pilot's perspective, the sideslip method is more convenient in operation and can more easily maintain the aircraft's landing track, allowing the pilot to respond quickly under complex crosswind conditions. However, from the perspective of passengers, the yaw law provides higher comfort and can effectively reduce the discomfort of passengers caused by the lateral force generated during landing. When the crosswind speed increases to a certain critical value, a single control method is obviously not enough to meet the requirements of the aircraft during crosswind landing. In this case, the pilot must not only consider the maintenance of the track, but also the comfort of the passengers. Therefore, it is necessary to propose a method for comprehensively evaluating the three options, so as to help pilots make more reasonable decisions under different crosswind conditions, reduce their decision-making pressure in a tense environment, and thus improve flight safety and passengers' flight experience. Through this method, pilots can flexibly adjust the control strategy according to the actual situation, ensuring a safe and smooth landing of the aircraft while minimizing the discomfort of passengers. Summary of the invention

[0003] In order to overcome the problems existing in the prior art, the present invention provides a method, system, device and medium for visually expressing the feasible domain of flight capability, which are used to overcome the current defects.

[0004] A method for visually expressing a feasible domain of flight capability, the method comprising the steps of:

[0005] S1. Trim the aircraft to keep it in a balanced state under specified crosswind conditions;

[0006] S2. Select all high-dimensional points when the aircraft is in a balanced state according to given constraints, and retain the generalized state points that meet the constraints;

[0007] S3. The area formed by all generalized state points is regarded as the feasible area of ​​flight capability and is visualized.

[0008] According to the aspects described above and any possible implementation method, an implementation method is further provided, wherein the balancing in S1 includes: given a set of state vectors and input parameters, the aircraft's total external torque is made zero, the state vectors include: the aircraft's ground speed, attitude angle, angular velocity and position, and the input parameters include: aileron rudder amount, elevator rudder amount, flap rudder amount and throttle opening.

[0009] According to the aspects described above and any possible implementation method, an implementation method is further provided, wherein the given restriction conditions include: aircraft attitude angle restriction, control surface control margin restriction, track sideslip angle restriction, landing speed restriction and sideslip angle restriction.

[0010] According to the above aspects and any possible implementation, an implementation is further provided, wherein the sideslip angle is limited within a range of plus or minus 8 degrees.

[0011] According to the aspects described above and any possible implementation manner, there is further provided an implementation manner, wherein the control margin limit of the rudder surface includes taking 75% of the design value of the maximum deflection angle of the rudder surface as the margin limit.

[0012] According to the aspects described above and any possible implementation, an implementation is further provided, wherein the track sideslip angle limit is limited to a yaw angle of plus or minus 10 degrees.

[0013] According to the above aspects and any possible implementation, an implementation is further provided, wherein the landing speed limit is within a landing speed range of plus or minus 9 m / s of the approach speed.

[0014] The present invention also provides a visual expression system of the feasible domain of flight capability, wherein the system implements the method described and comprises the following modules:

[0015] The trim module is used to trim the aircraft under specified crosswind conditions to keep the aircraft in a balanced state;

[0016] A selection module is used to select all high-dimensional points when the aircraft is in a balanced state according to given constraints, and retain generalized state points that meet the constraints;

[0017] The visualization expression module is used to visualize the area formed by all generalized state points as a feasible area of ​​flight capability.

[0018] The present invention further provides an electronic device, comprising:

[0019] A memory storing executable instructions;

[0020] A processor is used to execute the executable instructions in the memory to implement the method.

[0021] The present invention also provides a computer storage medium, on which a computer program is stored, and the computer program is executed by a processor to implement the method described.

[0022] Compared with the prior art, the present invention has the following beneficial effects:

[0023] The method for visually expressing the feasible domain of flight capability of the present invention comprises the following steps: trimming the aircraft under a specified crosswind condition so that the aircraft is kept in a balanced state; selecting all high-dimensional points when the aircraft is kept in a balanced state according to given constraints, and retaining generalized state points that meet the constraints; and visually expressing the region formed by all generalized state points as the feasible region of flight capability. The present invention is suitable for evaluating the crosswind resistance capability of fixed-wing aircraft during landing, and estimating the feasible domain of flight under a crosswind environment. The advantages of the present invention are: visually expressing various constraint factors, extracting the constrained factors, facilitating designers to optimize the crosswind resistance performance of the aircraft in a targeted manner, and facilitating risk assessment during flight. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] Figure 1 It is a schematic diagram of the aircraft attitude angle limitation range of the present invention;

[0025] Figure 2 It is a structural diagram of a simulation model under crosswind interference of the present invention;

[0026] Figure 3 A schematic diagram of a mapping of the aircraft attitude angle limit range of the present invention;

[0027] Figure 4 It is a visual schematic diagram of the feasible domain of the flight capability of the present invention. DETAILED DESCRIPTION

[0028] In order to better understand the technical solution of the present invention, the content of the present invention includes but is not limited to the specific implementation methods described below, and similar technologies and methods should be considered to be within the scope of protection of the present invention. In order to make the technical problems, technical solutions and advantages to be solved by the present invention clearer, the following will be described in detail with reference to the accompanying drawings and specific embodiments.

[0029] It should be clear that the embodiments described in the present invention are only part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0030] A method for visually expressing a feasible domain of flight capability of the present invention comprises the following steps:

[0031] S1. Trim the aircraft to keep it in a balanced state under specified crosswind conditions;

[0032] S2. Select all high-dimensional points when the aircraft is in a balanced state according to given constraints, and retain the generalized state points that meet the constraints;

[0033] S3. The area formed by all generalized state points is regarded as the feasible area of ​​flight capability and is visualized.

[0034] Preferably, the balancing in S1 of the present invention includes: giving a set of state vectors and input parameters to make the aircraft's combined external torque zero, the state vectors include: the aircraft's ground speed, attitude angle, angular velocity and position, and the input parameters include: aileron rudder amount, elevator rudder amount, flap rudder amount and throttle opening.

[0035] Preferably, the given restriction conditions include: aircraft attitude angle restriction, control surface control margin restriction, track sideslip angle restriction, landing speed restriction and sideslip angle restriction.

[0036] Preferably, the sideslip angle is limited to a range of plus or minus 8 degrees.

[0037] Preferably, the control margin limit of the rudder surface includes taking 75% of the design value of the maximum deflection angle of the rudder surface as the margin limit.

[0038] Preferably, the track sideslip angle limit is limited to a yaw angle of plus or minus 10 degrees.

[0039] Preferably, the landing speed limit is within the landing speed range of plus or minus 9 m / s of the approach speed.

[0040] Specifically, the specific implementation process of the present invention is as follows:

[0041] The automatic approach and landing process of an aircraft generally includes five stages: automatic approach, initial approach, intermediate approach, final approach (slide, level, drift and run) and missed approach. In the automatic approach and landing process, in view of the crosswind situation encountered by the aircraft, the present invention proposes a visual expression method of the feasible domain of flight capability to assist landing decision-making. The three correction methods are described as follows:

[0042] Sideslip correction method: The goal of the sideslip method is to keep the aircraft's body axis X, ground speed, and runway centerline coincident in the final stages of approach and landing. At this time, the aircraft lands along the runway centerline and the aircraft maintains force and torque balance. When using the sideslip method, the aircraft has a large sideslip angle. Only when the aileron and rudder are properly adjusted can the torque caused by the crosswind be balanced, which puts a burden on the pilot. Through force analysis, it can be seen that the side force has the greatest impact on the balance of the lateral channel. In order to balance the side force, the aircraft needs to maintain a large roll angle and use the lift component to balance the side force.

[0043] Crab correction method: a way to keep the aircraft level during the approach phase, with the nose of the aircraft pointing in the direction of the incoming flow (airspeed). At this time, the aircraft's track sideslip angle (the angle between the ground speed and the symmetry plane) is equal to the drift angle (the angle between the airspeed and the ground speed). At this time, the aircraft has no sideslip and the pilot does not need to operate the ailerons and rudder, but the ground speed direction is inconsistent with the nose, so the pilot cannot judge the direction of the descent based on the nose. At the moment of landing, the pilot needs to operate the rudder to straighten the aircraft, eliminate the track sideslip angle, and reduce the lateral load on the landing gear. In addition, the different local flow velocities on both sides of the wings caused by the yaw rotation will produce a rolling torque, so the ailerons need to be operated to eliminate the roll.

[0044] Crab and sideslip combination method: There are generally two combination methods: (1) Use a crab approach in the initial stage of landing. When the aircraft descends to a lower altitude, use the sideslip approach method until the landing is completed. (2) When using the sideslip method to touch down, it is restricted by constraints and cannot keep the track sideslip angle at zero. When only the crab correction method is used, the track sideslip angle is too large. At this time, a certain track sideslip angle can be maintained while maintaining a certain sideslip angle, so that the sum of the sideslip angle and the track sideslip angle is equal to the drift angle. At this time, the aircraft rolls appropriately to generate side force. That is, in the case of strong crosswinds, the sideslip method needs to be combined with a certain track sideslip angle. When the upwind main wheel touches down, reduce the rudder amount to keep the fuselage along the centerline of the runway, and at the same time properly control the aileron to keep the wings level after touching down.

[0045] Considering different crosswind sizes (considering only uniform wind fields), the strategies that an aircraft can adopt under different wind speed conditions are different. The visual expression method of the feasible domain of flight capability of the present invention evaluates the applicability of the three schemes, which is used to evaluate the aircraft's landing anti-crosswind performance and derive the trajectory envelope range.

[0046] Definition

[0047] From the analysis, it can be seen that the sideslip correction method is mainly constrained by the following conditions: rudder and aileron rudder constraints. If the control amount of a certain stick is insufficient, the attitude stability cannot be maintained. Roll angle limit. When the side force is too large, the aircraft roll angle φ will increase accordingly. The specific crosswind roll slope limit depends on multiple factors, including the aircraft model, the aircraft manufacturer's recommendations and the operator's operating procedures. The crab correction method is mainly constrained by the following conditions: the limitation of the lateral load capacity of the landing gear. Different aircraft models have different limitations. Airport runway conditions limit. Since the rudder needs to be used to straighten the aircraft at the moment of touchdown, the side force will cause the aircraft to deviate from the original touchdown point. Therefore, the crab correction method is generally not used when conditions permit, such as landing on a dry runway with a large crosswind. When the runway is relatively slippery, considering the requirements of ground taxiing braking, the crab correction method is faster than the skid measurement method and has a smaller load on the pilot.

[0048] The constraints used in the steps are described as follows:

[0049] (1) Aircraft attitude angle limit: During the landing process, the flight attitude of the aircraft is required to be within a certain limit range. The limit depends on the geometric size of the aircraft model to be evaluated. For example, large civil aircraft have a large wingspan, and when the roll angle is large, there will be a dangerous situation where the wing tip touches the ground or the engine nacelle touches the ground. Therefore, for different aircraft models, it is necessary to obtain the geometric collision diagram that the aircraft model can allow based on the geometric shape of the aircraft. The present invention takes the 737-300 aircraft as an example and obtains the following: Figure 1 The geometric collision diagram shown can be drawn similarly for different aircraft models, although the spans, engine layouts and landing gear layouts are different. Figure 1The horizontal coordinate is the roll angle of the aircraft, and the vertical coordinate is the pitch angle of the aircraft, among which the coordinates of point A are (0, 10~12), the coordinates of point B are (6~10, 12~14), the coordinates of point C are (12~15, 0~2), the coordinates of point D are (7~10, -2~-4), and the coordinates of point E are (0, 0). The straight line segment connecting point A and point B indicates that the excessive pitch angle causes the rear fuselage to collide with the ground before the main landing gear. Therefore, the straight line segment connecting point A and point B and the area above are the areas where the rear fuselage of the aircraft collides, and the area below is the area where the rear fuselage does not collide; the straight line segment connecting point B and point C indicates that when the pitch angle is large, the large roll angle will cause the wing tip to collide with the ground. Therefore, the straight line segment connecting point B and point C and the area to the right are the areas where the flap guide fairing collides, and the area on the left is the area where the flap guide fairing does not collide; the straight line connecting point C and point D The segment indicates that when the pitch angle is small or negative, the excessive roll angle causes the engine nacelle to collide with the ground. Therefore, the straight line segment connecting point C and point D and the area below it are the areas where the engine nacelle collides, and the area above is the area where the engine nacelle does not collide; the straight line segment connecting point D and point E indicates that when the pitch angle is negative, the front landing gear contacts the ground earlier than the main landing gear. Therefore, the straight line segment connecting point D and point E and the area below it are the areas where the front landing gear collides, and the area above is the area where the front landing gear does not collide; the straight line segment connecting point E and point A is used as a symmetric boundary segment. According to the principle of symmetry, when the aircraft rolls negatively or positively, the restricted range formed is symmetrical about the AE boundary line. The present invention chooses to analyze the positive roll. Therefore, the positive roll corresponds to the right side of the boundary line AE, and the area surrounded by points ABCDE is the restricted range of the flight attitude corresponding to the positive roll of the aircraft of this model. The attitude within this range can ensure that the main landing gear of the aircraft contacts the ground first. The attitude of the aircraft includes three, namely, roll angle, pitch angle and yaw angle. Here, only the collision relationship between the aircraft and the ground is considered, and the yaw angle does not affect the collision relationship. Therefore, the "contact state of the aircraft with the ground" has a binary relationship with the roll angle and the pitch angle. In the ground coordinate system, a geometric point is defined. When the roll angle and the pitch angle change, the geometric point can be made not to appear on the main landing gear. The line where all these geometric points are located is the boundary line. The area formed by the points ABCDE uses the above five straight line segments AB, BC, CD, DE, and EA as boundaries to restrict the aircraft attitude. The pitch angle and roll angle of the aircraft are within this area.

[0050] (2) Limitation of the control margin of elevator, aileron and other control surfaces: Flight safety requires that a certain safety margin be left for the control surfaces. The present invention uses 75% of the maximum deflection angle design value of the control surface as the margin limit. Taking the Boeing 737-300 model as an example, the aileron and rudder are limited to plus or minus 20 degrees, and the elevator is limited to [-17, 17] degrees.

[0051] (3) Track sideslip angle limit: When the aircraft ground speed vector is aligned with the earth axis OX g When they overlap, the track sideslip angle is the same as the yaw angle. When the aircraft touches down, the yaw angle needs to be limited to a certain range. The allowable angle varies depending on the airport runway and the aircraft model. A larger yaw angle means that the landing gear bears more lateral loads during landing, and there is a risk of running off the runway. Different aircraft models have different restrictions. According to actual conditions, the present invention takes the limit range of the track sideslip angle during touchdown as [-10,10] degrees.

[0052] (4) Landing speed limit: The speed limit of the aircraft during approach. Calculate the speed parameters of the aircraft during approach. According to the stall speed calculation formula:

[0053]

[0054] Where ρ is the air density, s is the wing area, (C L ) max is the maximum lift coefficient, which determines the approach speed:

[0055] V app =1.3V stall

[0056] Where V app is the approach speed in m / s, V stall The stall speed in the landing phase. The aircraft needs to control the landing speed within a certain range of the standard approach speed. The present invention takes the range of [-9,9] m / s of the approach speed as the landing speed range.

[0057] (5) Side slip angle limitation: During the leveling phase before the aircraft touches down, the nose of the aircraft is aligned with the runway and the wing slope is reduced. At this time, the lateral force on the aircraft is not balanced. A larger sideslip angle will bring a larger side force to the aircraft, causing lateral deviation. The sideslip angle is subject to different limitations for different aircraft models and runway conditions. The present invention takes 8 degrees for illustration.

[0058] When the wind field is constant, the dynamic equation of the rigid body mass center line motion in the airflow system is:

[0059]

[0060] Where: V g is the aircraft airspeed; W is the angular velocity of the airflow system relative to the inertial system, m is the mass of the aircraft, and F is the projection of the resultant force acting on the aircraft in the airflow system.

[0061] Kinematic equations for the downline motion of the aircraft ground system:

[0062]

[0063] La2g It represents the coordinate transformation matrix from the airflow system to the ground system. u, v, w are the projected components of the aircraft ground speed in the aircraft system. x, y, z are the coordinates of the aircraft center of mass in the ground system.

[0064] Kinematic equations for angular motion of an aircraft system:

[0065]

[0066] p, q, and r are the projected components of the angular velocity of the aircraft body relative to the ground system in the body coordinate system.

[0067] The dynamic equations of angular motion of the aircraft system:

[0068]

[0069] H x ,H y ,H z is the projection component of the aircraft angular momentum in the aircraft system, and the superscript ss represents the vector cross product. L, M, Nz are the projection components of the moment of the resultant external force on the center of mass in the aircraft system.

[0070] The above four groups of equations define a total of twelve state variables, namely: u, v, w, φ, θ, ψ, p, q, r, x, y, z. Each group of three represents the ground speed, attitude angle, angular velocity and position of the aircraft respectively. These twelve state variables constitute a vector, which is called the state of the aircraft. In addition, the aircraft's control surface and throttle lever opening are input variables for controlling the aircraft. The present invention sets the input parameter U to have four components: aileron rudder amount, elevator rudder amount, rudder amount, throttle opening, such as Figure 2 In the aircraft simulation structure diagram shown in the figure, based on the current state quantity engine thrust and other instructions, input quantity δ a / δ e / δ r The aircraft dynamics and kinematics module calculates the derivatives of the current state quantity based on the control surface instructions, etc. Next, these derivatives are used to update the current state within a predetermined step size. After completing this process, the next round of solution begins. In this way, it is possible to obtain the situation where the state quantity changes with the input quantity when the initial value is determined. The above process is an existing algorithm, and the specific process will not be repeated.

[0071] The specific visualization process of the present invention is as follows:

[0072] 1. In order to express the aircraft's ability to resist crosswind, the aircraft is first trimmed under specified crosswind conditions, that is, a set of state vectors X (u, v, w, φ, θ, ψ, p, q, r, x, y, z) and inputs U (aileron rudder delta_a, elevator rudder delta_e, rudder delta_r, throttle opening thrust) are found to make the aircraft's combined external force zero, and then the combined external torque zero, which means that when the aircraft's state quantity is X, the control input U can be given to keep the aircraft's state at the state quantity X, that is, to maintain "balance". The simulation trimming process generally uses a numerical solution method to iteratively adjust the control input U and the state vector X until the trimming condition is met. The present invention uses a sequential quadratic programming (SQP) algorithm to implement this process, converting the trimming problem into an optimization problem, with the goal of minimizing a cost function, that is, the sum of the squares of the deviations between the state vector and the input. When the state vectors φ, p, q, r, y, and z are all given, the values ​​of the four input components and the remaining six components are found through numerical solution methods, so that the total external force of the aircraft under the action of these twelve state quantities and four input quantities is zero. Then these twelve state quantities and the corresponding four input quantities are called a balancing state point. The SQP algorithm is an existing mature technology, and the specific calculation process will not be repeated. According to the above description, each balanced state corresponds to one or more groups of control input quantities. The state vector X and the control input U can be combined and regarded as a series of sixteen-dimensional vectors. Each sixteen-dimensional vector is a balancing state point. Under different conditions, on the plane with speed as the horizontal coordinate and roll angle as the vertical coordinate, each point can find the corresponding sixteen-dimensional balancing state point;

[0073] 2. After balancing in step 1, a grid is divided on the plane with speed as the horizontal coordinate and roll angle as the vertical coordinate (the density of the sampling grid determines the detail of the final image). Each grid point corresponds to a 16-dimensional balancing state point, thus obtaining a series of high-dimensional points. Each high-dimensional point indicates that the aircraft has a set of inputs that keep the aircraft state unchanged under the crosswind condition. According to the five restrictions (1)-(5) described above, this series of high-dimensional points is screened and selected. The five restrictions are actually restrictions on the state quantity and input quantity. The 16-dimensional high-dimensional points are checked, and the components of the high-dimensional points corresponding to the five restrictions are checked. If the restrictions are met, the high-dimensional points are retained, and if they are not met, they are discarded. This process is described below.

[0074] 3. According to the aircraft attitude angle limit, the aircraft's roll angle and pitch angle need to be Figure 1 In the ABCDE area shown, the restricted area is used to screen the generalized state points that meet the conditions. Figure 3 The horizontal axis is the aircraft's airspeed, and the vertical axis is the track sideslip angle. Therefore, Figure 3The ABCDE area in is Figure 1 The corresponding ABCDE areas in the image are mapped. The two are just different in their horizontal and vertical coordinates. Therefore, only when the horizontal and vertical coordinates are Figure 3 The high-dimensional state points in the ABCDE region shown satisfy the restriction conditions.

[0075] 4. According to the control margin limit of elevator and aileron, the input value U must meet the control margin limit of elevator and aileron. Here, the elevator limit is taken as an example. Figure 3 As shown, a grayscale bar graph is obtained, that is, the darker the grayscale, the smaller the elevator steering amount, such as the elevator steering amount value near the area where point B in the figure is located is smaller.

[0076] 5. Based on the side slip angle limit of the track, Figure 3 The dashed lines H and I in the middle ordinate are the dashed line H of the maximum value of the track sideslip angle and the dashed line I of the minimum value of the track sideslip angle respectively. The generalized state points inside between these two dashed lines meet the restrictions.

[0077] 6. Landing speed limit: According to the landing speed limit, select the ground speed range as

[0078] [61.2,79.4]m / s, according to the airspeed vector triangle relationship: airspeed + wind speed = ground speed, when the crosswind speed is 20kn, according to the above triangle relationship, the airspeed value range after conversion calculation is [61.9,80.1]m / s. Figure 3 It is given in the form of horizontal axis.

[0079] 7. Sideslip angle β limitation, consider the situation where the aircraft ground speed vector points to the runway direction, at this time the sideslip angle β and the track sideslip angle β k The sum is the equivalent sideslip angle β w , according to the formula:

[0080] β w =arctan(crosswind speed / ground speed)

[0081] Then we can get:

[0082] β k =β w -β

[0083] That is, the limitation of the sideslip angle β is converted into the side slip angle β of the track k restrictions, Figure 3 The high-dimensional points in embody this limitation. Figure 3 The dotted line FG in the figure represents a sideslip angle limit of positive 8 degrees. The dotted line FG serves as the sideslip angle limit when touching down. From the trigonometric relationship of the airspeed vector, it can be seen that when the aircraft speed is low, the equivalent sideslip angle is larger. At this time, if the track sideslip angle is too small, the sideslip angle will be too large.

[0084] right Figure 3 The area shown in FIG. 1 is further processed, that is, the area below the dotted line FG and the area below the track side slip angle of zero are discarded, and the retained part is the visualization part. The visualization area is finally as shown in FIG. Figure 4 As shown in the figure, it is the feasible or flyable area of ​​the aircraft. Since each point in the figure is a high-dimensional state point, each point contains roll angle information, and the desired roll angle information can be plotted in the figure in the form of contour lines. Therefore, Figure 3 or Figure 4 The roll angle value is displayed. In the feasible area, the roll angle of the AB segment is 0. The corresponding crab correction method is used. The aircraft maintains a certain track sideslip angle (about 8-10°) without sideslip. The smaller the speed, the larger the track sideslip angle, indicating that the greater the impact of crosswind disturbance. On the horizontal straight line with the vertical coordinate track sideslip angle of 0 in the figure, the corresponding sideslip correction method is used. At this time, the aircraft maintains a certain roll attitude, and the roll angle is about 9°. The lift component generated by the roll is used to offset the impact of crosswind disturbance on the aircraft movement. In other feasible areas except the above two straight lines or curves, the crab and sideslip combination method is used.

[0085] As an embodiment disclosed by the present invention, the present invention further discloses a visual expression system of a feasible domain of flight capability, wherein the system implements the method described, and comprises the following modules:

[0086] The trim module is used to trim the aircraft under specified crosswind conditions to keep the aircraft in a balanced state;

[0087] A selection module is used to select all high-dimensional points when the aircraft is in a balanced state according to given constraints, and retain generalized state points that meet the constraints;

[0088] The visualization expression module is used to visualize the area formed by all generalized state points as a feasible area of ​​flight capability.

[0089] As an embodiment disclosed in the present invention, the present invention further discloses an electronic device, the electronic device comprising:

[0090] A memory storing executable instructions;

[0091] A processor is used to execute the executable instructions in the memory to implement the method described in the present invention.

[0092] As an embodiment disclosed in the present invention, the present invention further discloses a computer storage medium, on which a computer program is stored, and the computer program is executed by a processor to implement the method described in the present invention.

[0093] The terms used in the embodiments of the present invention are only for the purpose of describing specific embodiments, and are not intended to limit the present invention. The singular forms "a", "said" and "the" used in the embodiments of the present invention and the appended claims are also intended to include plural forms, unless the context clearly indicates other meanings.

[0094] The above description shows and describes several preferred embodiments of the present invention, but as mentioned above, it should be understood that the present invention is not limited to the form disclosed herein, and should not be regarded as excluding other embodiments, but can be used in various other combinations, modifications and environments, and can be modified within the scope of the application concept described herein through the above teachings or the technology or knowledge of the relevant field. The changes and modifications made by those skilled in the art shall not depart from the spirit and scope of the present invention, and shall be within the scope of protection of the claims attached to the present invention.

Claims

1. A method for visually expressing the feasible domain of flight capability, characterized in that: The method comprises the steps of: S1. Trim the aircraft to keep it in a balanced state under specified crosswind conditions; S2. Select all high-dimensional points when the aircraft is in a balanced state according to given constraints, and retain the generalized state points that meet the constraints; S3. The area formed by all generalized state points is regarded as the feasible area of ​​flight capability and is visualized.

2. The visual expression method according to claim 1, characterized in that: The balancing in S1 includes: giving a set of state vectors and input parameters to make the aircraft's combined external torque zero, the state vectors include: the aircraft's ground speed, attitude angle, angular velocity and position, and the input parameters include: aileron rudder amount, elevator rudder amount, flap rudder amount and throttle opening.

3. The visual expression method according to claim 1, characterized in that: The given restriction conditions include: aircraft attitude angle restriction, control surface control margin restriction, track sideslip angle restriction, landing speed restriction and sideslip angle restriction.

4. The visual expression method according to claim 3, characterized in that: The sideslip angle is limited to within the range of plus or minus 8 degrees.

5. The visual expression method according to claim 3, characterized in that: The control margin limit of the rudder surface includes taking 75% of the design value of the maximum deflection angle of the rudder surface as the margin limit.

6. The visual expression method according to claim 3, characterized in that: The track sideslip angle limit is limited to plus or minus 10 degrees of the yaw angle.

7. The visual expression method according to claim 3, characterized in that: The landing speed limit is within the range of ±9 m / s of the approach speed.

8. A visual expression system for the feasible domain of flight capability, characterized in that: The system implements the method described in any one of claims 1 to 7, and includes the following modules: The trim module is used to trim the aircraft under specified crosswind conditions to keep the aircraft in a balanced state; A selection module is used to select all high-dimensional points when the aircraft is in a balanced state according to given constraints, and retain generalized state points that meet the constraints; The visualization expression module is used to visualize the area formed by all generalized state points as a feasible area of ​​flight capability.

9. An electronic device, characterized in that: The electronic device comprises: A memory storing executable instructions; A processor, wherein the processor runs the executable instructions in the memory to implement the method according to any one of claims 1 to 7.

10. A computer storage medium, characterized in that: The medium stores a computer program, and the computer program is executed by a processor to implement the method according to any one of claims 1 to 7.

Images