A normal overload protection control method considering change of barycenter

Abstract

The application belongs to the technical field of aircraft flight control, and particularly relates to a normal overload protection control method considering the change of the center of gravity. The method comprises the following steps: S1, converting the displacement Xe of the control stick into an overload instruction DNy; S2, multiplying the overload instruction DNy by a gain K to form a forward rudder deflection instruction, wherein the gain K is the rudder deflection required by a unit overload; S3, superimposing the forward rudder deflection instruction with a proportional rudder deflection instruction and an integral rudder deflection instruction to generate an elevator deflection instruction, and controlling the flight of the aircraft based on the elevator deflection instruction. The application realizes the normal overload protection function under the change of the center of gravity and extreme operating conditions, and improves the safety of the aircraft.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aircraft flight control, and particularly relates to a normal overload protection control method considering center of gravity change. BACKGROUND

[0002] Modern aircraft widely adopt fly-by-wire flight control systems, and realize control functions and protection functions through complex control laws, thereby reducing the burden of pilots and improving flight safety. The normal overload protection function is an important flight boundary protection function, and is used for preventing the normal overload of an aircraft from exceeding a limit value. If the normal overload exceeds the limit value, the aircraft structure will be damaged, and the aircraft will be disintegrated in a serious case.

[0003] The center of gravity change of an aircraft has a great influence on the normal overload response. Under the same elevator deflection, the normal overload response of a rear center of gravity is greater than that of a front center of gravity. A traditional control method adopts a lever displacement command normal overload control architecture, and realizes the normal overload protection function under normal flight control (maneuvering under a constant height and speed) at the cost of sacrificing the stability margin of a flight control system or the maneuvering ability of the aircraft. Moreover, under extreme control working conditions such as rapid dive followed by step pull-up, rapid climb followed by step dive, large pitch angle dive followed by step pull-up, and constant speed flight followed by step pull-up, the overload of the aircraft will exceed the limit value.

[0004] At present, in the design process of domestic aircraft, the following deficiencies exist in the design of the normal overload protection function under the center of gravity change and extreme control:

[0005] 1. It is assumed that the occurrence probability of extreme control working conditions such as rapid dive followed by step pull-up, rapid climb followed by step dive, large pitch angle dive followed by step pull-up, and constant speed flight followed by step pull-up is low, and these working conditions are not considered in the design and verification, thereby leading to incomplete coverage of the normal overload protection function working conditions, and potential danger of the aircraft.

[0006] 2. In order to take into account the center of gravity change and the extreme control working conditions such as rapid dive followed by step pull-up, rapid climb followed by step dive, large pitch angle dive followed by step pull-up, and constant speed flight followed by step pull-up, the normal overload value of the lever displacement command is reduced, that is, the normal overload exceeding margin under the rear center of gravity and the extreme control working conditions is left, and the normal overload response of the lever displacement command is reduced. This method causes the normal overload response of the front center of gravity, the intermediate center of gravity, and normal control to be reduced, and the maneuvering ability of the aircraft is greatly sacrificed.

[0007] 3. In order to take into account the change of gravity center and the rapid dive after step pull-up, rapid climb after step dive, large pitch angle dive after step pull-up, and flat flight acceleration after step pull-up, the pitch angle rate feedback gain is increased, the normal overload response damping characteristic is increased, thereby reducing the rear gravity center and the normal overload response overshoot under extreme manipulation. This will lead to the reduction of the stability margin of the flight control system, the reduction of the anti-interference ability of the control system, and the increase of the normal overload response damping of the front gravity center and the intermediate gravity center under normal manipulation, and the delay of the normal overload response of the aircraft. SUMMARY

[0008] In order to solve at least one of the above technical problems, the application designs a normal overload protection control method considering the change of gravity center, to ensure that the normal overload of the aircraft does not exceed the limit value under the extreme manipulation conditions of the change of gravity center and the rapid dive after step pull-up, rapid climb after step dive, large pitch angle dive after step pull-up, and flat flight acceleration after step pull-up, and improve the safety of the aircraft.

[0009] The normal overload protection control method considering the change of gravity center provided by the application mainly includes:

[0010] Step S1, converting the joystick displacement Xe into an overload instruction DNy;

[0011] Step S2, multiplying the overload instruction DNy by a gain K to form a forward rudder deflection instruction, wherein the gain K is the rudder deflection required per unit overload;

[0012] Step S3, superimposing the forward rudder deflection instruction with a proportional rudder deflection instruction and an integral rudder deflection instruction to generate an elevator deflection instruction, and controlling the flight of the aircraft based on the elevator deflection instruction.

[0013] Preferably, step S1 further includes:

[0014] Step S11, obtaining the travel range of the joystick displacement Xe, and obtaining the overload limit range of the aircraft structure;

[0015] Step S12, based on the travel range and the overload limit range, constructing a gradient function for representing the relationship between the joystick displacement and the overload;

[0016] Step S13, converting the real-time joystick displacement Xe into a real-time overload instruction DNy based on the gradient function.

[0017] Preferably, in step S11, the overload limit range of the aircraft structure is further corrected, which includes:

[0018] Step S111, obtaining a positive overload adjustment coefficient interpolation table when the aircraft pitch angle is negatively changed in its maneuvering range, and a negative overload adjustment coefficient interpolation table when the aircraft pitch angle is positively changed in its maneuvering range, the positive overload adjustment coefficient interpolation table recording a plurality of negative pitch angles and a plurality of corresponding positive overload adjustment coefficients, and the negative overload adjustment coefficient interpolation table recording a plurality of positive pitch angles and a plurality of corresponding negative overload adjustment coefficients;

[0019] Step S112, interpolating a corresponding adjustment coefficient from the positive overload adjustment coefficient interpolation table or the negative overload adjustment coefficient interpolation table according to the current pitch angle of the aircraft;

[0020] Step S113, correcting the boundary value of the overload limit range based on the adjustment coefficient.

[0021] Preferably, step S2 further comprises:

[0022] Step S21, obtaining a trim rudder deflection De_trim and a rudder deflection De_ny0 required for 0g normal overload, wherein g is the gravity acceleration;

[0023] Step S22, determining a gain K as: K=De_trim-De_ny0.

[0024] Preferably, in step S21, obtaining the trim rudder deflection De_trim comprises:

[0025] Step S211, obtaining an aircraft pitch angle and a roll angle;

[0026] Step S212, calculating a body axis normal overload based on the aircraft pitch angle and the roll angle;

[0027] Step S213, determining whether the aircraft is in steady flight according to a difference between a feedback value of a normal overload sensor and the calculated body axis normal overload, the aircraft being in steady flight when the difference is less than a threshold value, and the threshold value being 0.04g-0.06g, g being the gravity acceleration;

[0028] Step S214, when the aircraft is in steady flight, taking the integral rudder deflection command as an input to form the trim rudder deflection De_trim of the aircraft through an integrator, or taking 0 as the input to form the trim rudder deflection De_trim of the aircraft through the integrator.

[0029] Preferably, step S2 further comprises:

[0030] Step S23, calculating a rudder deflection value required for unit overload under different aircrafts, different barometric altitudes and different indicated airspeeds, wherein the different aircrafts refer to aircrafts with different flap slat wing configurations, weights and gravity centers.

[0031] Step S24, for each aircraft and air pressure height, taking the minimum indicated airspeed as the reference point, calculating the first ratio of other indicated airspeeds to the minimum indicated airspeed, and calculating the second ratio of the deflection value of the control surface required by the unit overload corresponding to the other indicated airspeed to the deflection value of the control surface required by the unit overload of the minimum indicated airspeed, constructing the corresponding relationship of the first ratio and the second ratio, and obtaining the final corresponding relationship by calculating the average value of the corresponding relationships of all aircraft and air pressure heights;

[0032] Step S25, obtaining the indicated airspeed VIAS and the trim speed VIAS_TRIM of the aircraft, taking the ratio of the indicated airspeed VIAS to the trim speed VIAS_TRIM as the first ratio, interpolating the second ratio in the corresponding relationship, and taking the second ratio as the gain correction coefficient k_v;

[0033] Step S26, gain correcting the gain K based on the gain correction coefficient k_v.

[0034] Preferably, in step S25, the trim speed VIAS_TRIM is determined by the following steps:

[0035] Step S251, when the aircraft is in the steady state, the trim speed VIAS_TRIM is the indicated airspeed VIAS; when the aircraft switches from the steady state to the non-steady state, the trim speed VIAS_TRIM is locked at the indicated airspeed VIAS at the last moment of the steady state; when the aircraft switches from the non-steady state to the steady state, the value of the trim speed VIAS_TRIM linearly changes from the last trim speed VIAS_TRIM to the indicated airspeed VIAS within a set fade time, and when the aircraft is in the non-steady state, the value of the trim speed VIAS_TRIM is the last trim speed VIAS_TRIM.

[0036] Preferably, step S251 further comprises determining whether the aircraft is in the steady state, specifically comprising: when the aircraft is in the steady state and the displacement of the steering column is less than a set value, the aircraft is in the steady state, otherwise the aircraft is in the non-steady state.

[0037] The application can retain the traditional control law architecture and realize a normal overload protection control function. The normal overload protection control law considering the center of gravity change is designed, the pitch angle adjustment overload limit value is introduced, the forward channel gain before the aircraft trim rudder deflection calculation is obtained in real time, and the aircraft trim speed is obtained in real time. According to the relationship between the real-time speed and the trim speed, the forward channel gain is adjusted, so as to solve the problem that the aircraft overload exceeds the limit value in the extreme operating conditions such as the center of gravity change and the step pull-up after rapid diving, the step dive after rapid climbing, the step pull-up after large pitch angle diving, and the step pull-up after level flight acceleration. The method does not need to modify any hardware part of the flight control system. The normal overload protection control law can be designed according to the method, the flight control system control law software is changed, the modification cost is saved, the normal overload protection function in the center of gravity change and the extreme operating conditions is realized, and the safety of the aircraft is improved. BRIEF DESCRIPTION OF DRAWINGS

[0038] Figure 1 FIG. 1 is a schematic diagram of a normal overload protection control architecture of a preferred embodiment of the normal overload protection control method considering the center of gravity change of the application.

[0039] Figure 2 FIG. 2 is a schematic diagram of a normal overload protection control architecture with a trim rudder deflection obtaining module.

[0040] Figure 3 FIG. 3 is a schematic diagram of a trim rudder deflection obtaining module.

[0041] Figure 4 FIG. 4 is a schematic diagram of the change curve of the ratio of the indicated airspeed and the unit overload required rudder deflection to the reference state at each state point.

[0042] Figure 5 FIG. 5 is a schematic diagram of a trim speed obtaining logic.

[0043] Figure 6 FIG. 6 is a schematic diagram of a level flight state judgment logic of the aircraft. DETAILED DESCRIPTION

[0044] In order to make the purpose, technical scheme and advantages of the application clearer, the technical scheme of the embodiments of the application will be described in more detail below with reference to the drawings of the embodiments of the application. In the drawings, the same or similar notations represent the same or similar elements or elements with the same or similar functions throughout. The described embodiments are part of the embodiments of the application, not all of the embodiments. The embodiments described below with reference to the drawings are exemplary and are intended to explain the application, and cannot be understood as limiting the application. Based on the embodiments in the application, all other embodiments obtained by those skilled in the art without creative labor fall within the scope of protection of the application. The embodiments of the application will be described in detail below with reference to the drawings.

[0045] This application provides a normal overload protection control method that takes into account changes in the center of gravity, such as... Figure 1 As shown, it mainly includes:

[0046] Step S1: Convert the control stick displacement Xe into an overload command DNy;

[0047] Step S2: Multiply the overload command DNy by the gain K to form a forward rudder deflection command, wherein the gain K is the rudder deflection required per unit overload.

[0048] Step S3: The forward rudder deflection command is superimposed with the proportional rudder deflection command and the integral rudder deflection command to generate the elevator deflection command, and the aircraft is controlled to fly based on the elevator deflection command.

[0049] refer to Figure 1 The normal overload protection control architecture consists of a forward channel, a proportional channel, and an integral channel. The forward channel converts the stick displacement Xe into an overload command DNy, which is then multiplied by a gain K (the rudder deflection required per unit overload) to generate a forward rudder deflection command. The proportional channel introduces pitch rate and normal overload feedback to improve the dynamic response to normal overload, and converts these into a proportional rudder deflection command after multiplying by a gain. The integral channel achieves precise control of the normal overload command through an integrator, subtracting the actual normal overload from the normal overload command and multiplying by a gain to generate an integral rudder deflection command. The forward rudder deflection command, the proportional rudder deflection command, and the integral rudder deflection command together generate the elevator deflection command, thereby controlling the aircraft.

[0050] The calculation methods for proportional rudder deflection command and integral rudder deflection command are based on the existing control law architecture. Therefore, this application can improve upon the existing control law architecture without modifying any hardware components of the flight control system. The normal overload protection control law can be designed according to this method, the flight control system control law software can be modified, the modification cost can be saved, and the normal overload protection function under center of gravity changes and extreme handling conditions can be realized, thereby improving aircraft safety.

[0051] In some alternative implementations, step S1 further includes:

[0052] Step S11: Obtain the travel range of the control stick displacement Xe, and at the same time obtain the overload limit range of the aircraft structure;

[0053] Step S12: Based on the travel range and the overload limit range, construct a gradient function to represent the relationship between the control stick displacement and the overload;

[0054] Step S13: Based on the gradient function, convert the real-time control stick displacement Xe into a real-time overload command DNy.

[0055] In this embodiment, reference Figure 1The overload limiting module in the forward channel converts the control stick displacement Xe into an overload command DNy. For ease of explanation, assume that Xe ranges from [-100, 100] mm, a positive stick displacement indicates push-stick operation, and a negative stick displacement indicates pull-stick operation. The overload limiting range of the aircraft structure is [-1, 2.5] g. By designing the stick displacement-overload gradient K_Xe_Ny (i.e., the gradient function in step S12; this design method is a traditional method and will not be elaborated further), the following correspondence can be achieved:

[0056] When Xe = 100, DNy = -1;

[0057] When Xe = -100, DNy = 2.5;

[0058] When -100<Xe<100, DNy=K_Xe_Ny×Xe.

[0059] Therefore, the real-time control stick displacement Xe can be converted into a real-time overload command DNy.

[0060] In some alternative implementations, step S11 further includes modifying the overload limit range of the aircraft structure, which includes:

[0061] Step S111: Obtain the positive overload adjustment coefficient interpolation table when the aircraft pitch angle changes negatively within its maneuvering range, and the negative overload adjustment coefficient interpolation table when the aircraft pitch angle changes positively within its maneuvering range. The positive overload adjustment coefficient interpolation table records the correspondence between multiple negative pitch angles and multiple positive overload adjustment coefficients, and the negative overload adjustment coefficient interpolation table records the correspondence between multiple positive pitch angles and multiple negative overload adjustment coefficients.

[0062] Step S112: Based on the aircraft's current pitch angle, interpolate the corresponding adjustment coefficient from the positive overload adjustment coefficient interpolation table or the negative overload adjustment coefficient interpolation table;

[0063] Step S113: Correct the boundary value of the overload limit range based on the adjustment coefficient.

[0064] In this embodiment, the maximum normal overload command is attenuated according to the negative pitch angle. That is, as the pitch angle increases, the maximum normal overload command corresponding to full stick pull is attenuated, and as the pitch angle decreases, the attenuation coefficient gradually decreases. Similarly, the minimum normal overload command is attenuated according to the positive pitch angle. As the pitch angle decreases, the normal overload command gradually recovers to its maximum. Assuming the aircraft's pitch angle maneuver range is [-30, 30]°, Table 1 below is the positive overload adjustment coefficient interpolation table M1, and Table 2 below is the negative overload adjustment coefficient interpolation table M2. The adjustment coefficients in M1 and M2 can be determined by simulation based on the specific aircraft.

[0065] Table 1 Interpolation Table of Positive Overload Adjustment Coefficient M1

[0066] Pitch angle (°) -30 -20 -10 -5 Positive over-ride trim coefficient 0.85 0.9 0.95 1

[0067] Table 2 Interpolation Table of Negative Overload Adjustment Coefficient M2

[0068] Pitch angle (°) 5 10 20 30 Negative over-ride trim coefficient 1 0.95 0.9 0.85

[0069] Then, based on M1 and M2, the corresponding pitch angle correction values ​​m1 and m2 are interpolated. m1 and m2 are then used to adjust DNy corresponding to the rod displacements of 100m and -100m in the above embodiment, as follows:

[0070] When Xe = 100, DNy = -1 × m2;

[0071] When Xe=-100, DNy=2.5×m1.

[0072] In some alternative implementations, step S2 further includes:

[0073] Step S21: Obtain the aircraft trim rudder deflection De_trim and the rudder deflection De_ny0 required for 0g normal overload, where g is the gravitational acceleration;

[0074] Step S22: Determine the gain K as: K = De_trim - De_ny0.

[0075] The gain K in step S2 is also known as Figure 1 The gain K (rudder deflection required per unit overload) in the forward channel is highly dependent on the aircraft's center of gravity; the forward center of gravity gain is greater than the aft center of gravity gain. To match the gain to the center of gravity, in Figure 1 A calculation module for real-time acquisition of aircraft trim rudder deflection is added to the integral channel, thereby constructing a method for calculating the gain K of the forward channel. The calculation method is as follows:

[0076] Forward channel gain = rudder deflection trim - rudder deflection required for 0g normal overload.

[0077] Since the rudder deflection for trim can reflect the aircraft's center of gravity, the rudder deflection required for 0g normal overload is independent of the aircraft's center of gravity. Therefore, the method described above can be used to calculate the forward channel gain to ensure matching with the center of gravity.

[0078] It should also be noted that De_ny0 is the rudder deflection required for a 0g normal overload. De_ny0 can be calculated using traditional methods, i.e., finding the rudder deflection required to produce a -1g normal overload, which will not be elaborated here.

[0079] In some alternative implementations, step S21, obtaining the trim rudder offset De_trim, includes:

[0080] Step S211: Obtain the aircraft's pitch and roll angles;

[0081] Step S212: Calculate the body axis normal overload based on the aircraft pitch and roll angles;

[0082] Step S213: Determine whether the aircraft is in steady-state flight based on the difference between the feedback value of the normal overload sensor and the calculated body axis normal overload. When the difference is less than a threshold value, the aircraft is in steady-state flight. The threshold value is 0.04g to 0.06g, where g is the gravitational acceleration.

[0083] Step S214: When the aircraft is in steady-state flight, the integral rudder deflection command is used as input to form the aircraft trim rudder deflection De_trim through the integrator; otherwise, 0 is used as input to form the aircraft trim rudder deflection De_trim through the integrator.

[0084] This embodiment primarily describes a method for obtaining trim deflection, specifically including: calculating the normal overload during steady-state flight based on the aircraft's roll and pitch angles, and combining this with the normal overload sensor values ​​to determine whether the aircraft is truly in steady-state flight. When the normal overload sensor value is approximately equal to the calculated normal overload during steady-state flight, it indicates that the aircraft is in a steady-state flight state; otherwise, it indicates that the aircraft is not in a steady-state flight state. The core of the aircraft trim deflection calculation module is an integrator. The integrated quantity is generated from the value of the integration channel when the aircraft is in a steady-state flight state. After passing through the integrator, the aircraft trim deflection is generated, and the integrated quantity is fed back to the integration channel to cancel the integrated quantity of the integrator, thus reducing the value of the integration channel. Finally, when the aircraft is in a steady-state flight state, the output of the aircraft trim deflection calculation module replaces the integration channel. If the aircraft is not in a steady-state flight state, the integrated quantity of the integrator in the aircraft trim deflection calculation module is 0.

[0085] In this embodiment, reference Figure 2 The trim rudder deflection calculation module generates the aircraft trim rudder deflection De_trim and De_OffLoad based on the integral channel output De_cmd_int. De_OffLoad is used to offset the integrated quantity of the integral channel integrator. The forward rudder deflection command, proportional rudder deflection command, integral rudder deflection command, and aircraft trim rudder deflection De_trim together generate the elevator deflection command.

[0086] Aircraft trim rudder deflection De_trim is used not only to calculate the gain K, but also according to Figure 2 The description also points to the subsequent synthesizer, which is mainly determined by the trim rudder deflection acquisition module itself, see reference. Figure 3 The TRIM_SW switch logic is as follows:

[0087] During steady-state flight, the normal overload of the body axis system is Where pitch is the tilt angle and bank is the roll angle. The normal overload sensor value is Ny, defined as follows: Then: when |Error_Ny| < 0.05g, TRIM_SW = 1, meaning De_cmd_int is used as input; when |Error_Ny| ≥ 0.05g, TRIM_SW = 0, meaning 0 is used as input. When De_cmd_int is used as input, the output De_OffLoad value is consistent with the input De_cmd_int. Figure 2 After the subtraction operation before the integrator, the value of De_cmd_int will become 0. Therefore, the aircraft trim rudder deflection De_trim is used as the input of the subsequent synthesizer to replace the input of the original integration channel.

[0088] In this embodiment, based on the measurement accuracy and zero-position characteristics of the pitch angle, roll angle, and normal overload sensors, it is tentatively set that when |Error_Ny| < 0.05g, the aircraft is considered to be in steady-state flight. The threshold of 0.05g can be adjusted according to the specific sensor characteristics.

[0089] In some alternative implementations, step S2 further includes:

[0090] Step S23: Calculate the control surface deflection value required for unit overload under different aircraft, different pressure altitudes, and different indicated airspeeds. The different aircraft refer to aircraft with different flap and slat configurations, weights, and centers of gravity.

[0091] Step S24: For each aircraft and barometric altitude, using its minimum indicated airspeed as a reference point, calculate the first ratio of other indicated airspeeds to the minimum indicated airspeed. At the same time, calculate the second ratio of the control surface deflection value required per unit overload for other indicated airspeeds to the control surface deflection value required per unit overload for the minimum indicated airspeed. Construct the correspondence between the first ratio and the second ratio. Obtain the final correspondence by averaging the correspondences for all aircraft and barometric altitudes.

[0092] Step S25: Obtain the aircraft indicated airspeed VIAS and trim speed VIAS_TRIM, take the ratio of indicated airspeed VIAS to trim speed VIAS_TRIM as the first ratio, interpolate the second ratio in the correspondence, and take the second ratio as the gain correction coefficient k_v.

[0093] Step S26: Based on the gain correction coefficient k_v, perform gain correction on the gain K.

[0094] In this embodiment, state points covering aircraft weight, center of gravity, various flap and slat configurations, and speed range are selected. The control surface deflection required per unit overload at each state point is calculated to obtain the change with increasing speed. The attenuation coefficient of the forward gain is selected by taking the average value.

[0095] Taking a cruise configuration, a weight of 110,000 kg, and a center of gravity as an example, we selected altitudes of 3,000 m and 5,000 m, and selected small, medium, and large speeds within the flight envelope for the indicated airspeed. The calculation of the control surface deflection required per unit overload is shown in Table 3 (the calculation of the control surface deflection required per unit overload is a traditional method and will not be elaborated here).

[0096] Table 3 Example of Calculation of Required Control Surface Deflection per Unit Overload

[0097]

[0098] Using the minimum speed at each altitude as a reference state, plot the ratios of the indicated airspeed and the control surface deflection required per unit overload at other state points at that altitude to the reference state. Taking an altitude of 3000m as an example, with state point 1 as the reference state, plot as follows... Figure 4 The curves shown represent the ratios of the indicated airspeed at state points 2 and 3 to that at state point 1 on the horizontal axis, and the ratios of the control surface deflection required per unit overload at state points 2 and 3 to that at state point 1 on the vertical axis. Similarly, curves are plotted at an altitude of 5000m.

[0099] Pick Figure 4 The attenuation coefficient k_v can be determined by the average of the ordinates of the two curves, as shown in Table 4.

[0100] Table 4 Attenuation coefficient k_v

[0101] VIAS / VIAS_TRIM 1 1.3 1.5 1.6 k_v 1 0.57 0.42 0.2

[0102] In actual design, it is necessary to draw curves for multiple weights, centers of gravity, heights, and configurations, and take the average of multiple curves to determine k_v. If the curves vary greatly with the configuration, k_v can be made into a multidimensional interpolation table and adjusted according to the configuration.

[0103] Finally, based on the ratio of the current indicated airspeed VIAS to the trim speed VIAS_TRIM, the gain K in the forward channel is adjusted, and the gain K becomes: K=(De_trim-De_ny0)×k_v.

[0104] In some alternative implementations, in step S25, the trimming speed VIAS_TRIM is determined by the following steps:

[0105] Step S251: When the aircraft is in level flight, the trim speed VIAS_TRIM is the indicated airspeed VIAS; when the aircraft switches from level flight to non-level flight, the trim speed VIAS_TRIM is locked at the indicated airspeed VIAS at the last moment of level flight; when the aircraft switches from non-level flight to level flight, the value of the trim speed VIAS_TRIM changes linearly from the previous trim speed VIAS_TRIM to the indicated airspeed VIAS within a set fade time; when the aircraft is in non-level flight, the value of the trim speed VIAS_TRIM is the trim speed VIAS_TRIM of the previous frame.

[0106] In this embodiment, the set fade time is generally 2 seconds, for reference. Figure 5 When PF_ST changes from 1 to 0, no fade processing is performed, and VIAS_TRIM is the VIAS_TRIM of the previous frame; when PF_ST changes from 0 to 1, fade processing is performed, and the fade time is 2s, that is, the value of VIAS_TRIM is the VIAS_TRIM of the previous frame, which changes linearly to VIAS after 2s.

[0107] In some alternative implementations, step S251 further includes determining whether the aircraft is in level flight, specifically: when the aircraft is in steady flight and the displacement of the control stick is less than a set value, the aircraft is in level flight; otherwise, the aircraft is in non-level flight.

[0108] refer to Figure 6 The system uses an AND gate to determine whether the aircraft is in level flight. Based on the mechanical characteristics of the control stick and the characteristics of the stick displacement sensor, a stick displacement of less than 2mm is tentatively considered as the control stick not being manipulated. This value can be adjusted according to the specific characteristics of the control stick. When PF_ST is 1 for 1 second, it indicates that the aircraft is in level flight; otherwise, the aircraft is in a non-level flight state, that is, in a maneuver.

[0109] It is understandable that when the aircraft speed changes, the value of the integral channel changes, and thus the trim rudder deflection also changes. That is, the trim rudder deflection reflects not only the aircraft's center of gravity but also changes in aircraft speed. At the same center of gravity, as the aircraft speed increases, the trim rudder deflection should decrease; as the aircraft speed decreases, the trim rudder deflection should increase. The change in trim rudder deflection lags behind the change in aircraft speed. When the aircraft speed increases too rapidly, the trim rudder deflection will mismatch with the aircraft speed, resulting in a slightly larger trim rudder deflection, leading to a slightly larger forward gain and making it easier to exceed the normal overload limit. Furthermore, when the aircraft is not in steady-state flight, the trim rudder deflection does not change. In other words, when the aircraft is not in steady-state flight and the speed increases, the trim rudder deflection remains unchanged, resulting in a basically unchanged forward channel gain (the rudder deflection required for 0g normal overload changes very little with speed). This causes the elevator angle generated by the control to increase, which is mismatched with the aircraft speed, leading to exceeding the normal overload limit. Therefore, this application solves this problem through the above steps S23-S26.

[0110] Specifically, when the aircraft is in level flight, the real-time indicated airspeed is used as the trim speed and updated in real time. When the aircraft is not in level flight, the trim speed remains the value of the previous trim speed. To ensure that the trim speed does not jump, when transitioning from a non-level flight state to a level flight state, the trim speed gradually fades from the previous value to the current indicated airspeed. Based on the ratio of the current indicated airspeed to the trim speed, the forward gain in the forward channel is adjusted. If the ratio is greater than 1, it indicates that the current indicated airspeed deviates significantly from the trim speed, and to avoid excessive control effect causing normal overload, forward gain attenuation is required. If the ratio is less than 1, it indicates that the current indicated airspeed is less than the trim speed, and in this case, the aircraft's normal overload will not exceed the limit, so forward gain adjustment is not necessary.

[0111] This application retains the traditional control law architecture to implement a normal overload protection control function. The normal overload protection control law designed in this application considers changes in the center of gravity, introduces pitch angle adjustment overload limits, obtains real-time aircraft trim rudder deflection to calculate forward channel gain, and obtains real-time aircraft trim speed. Based on the relationship between real-time speed and trim speed, the forward channel gain is adjusted to solve the problem of aircraft overload exceeding limits under extreme handling conditions such as changes in the center of gravity, rapid dive followed by step-up, rapid climb followed by step-dive, large pitch angle dive followed by step-up, and level flight acceleration followed by step-up. This method does not require any modification to the flight control system hardware. The normal overload protection control law can be designed according to this method, and the flight control system control law software can be modified, saving modification costs and achieving normal overload protection under changes in the center of gravity and extreme handling conditions, thus improving aircraft safety.

[0112] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A normal overload protection control method considering changes in the center of gravity, characterized in that, include: Step S1: Convert the control stick displacement Xe into an overload command DNy; Step S2: Multiply the overload command DNy by the gain K to form a forward rudder deflection command, wherein the gain K is the rudder deflection required per unit overload. Step S3: Superimpose the forward rudder deflection command with the proportional rudder deflection command and the integral rudder deflection command to generate the elevator deflection command, and control the aircraft flight based on the elevator deflection command. Step S2 further includes: Step S21: Obtain the aircraft trim rudder deflection De_trim and the rudder deflection De_ny0 required for 0g normal overload, where g is the gravitational acceleration; Step S22: Determine the gain K as: K = De_trim - De_ny0; In step S21, obtaining the trim rudder offset De_trim includes: Step S211: Obtain the aircraft's pitch and roll angles; Step S212: Calculate the body axis normal overload based on the aircraft pitch and roll angles; Step S213: Determine whether the aircraft is in steady-state flight based on the difference between the feedback value of the normal overload sensor and the calculated body axis normal overload. When the difference is less than a threshold value, the aircraft is in steady-state flight. The threshold value is 0.04g~0.06g, where g is the gravitational acceleration. Step S214: When the aircraft is in steady-state flight, the integral rudder deflection command is used as input to form the aircraft trim rudder deflection De_trim through the integrator; otherwise, 0 is used as input to form the aircraft trim rudder deflection De_trim through the integrator.

2. The normal overload protection control method considering center of gravity changes as described in claim 1, characterized in that, Step S1 further includes: Step S11: Obtain the travel range of the control stick displacement Xe, and at the same time obtain the overload limit range of the aircraft structure; Step S12: Based on the travel range and the overload limit range, construct a gradient function to represent the relationship between the control stick displacement and the overload; Step S13: Based on the gradient function, convert the real-time control stick displacement Xe into a real-time overload command DNy.

3. The normal overload protection control method considering center of gravity changes as described in claim 2, characterized in that, Step S11 further includes modifying the overload limit range of the aircraft structure, which includes: Step S111: Obtain the positive overload adjustment coefficient interpolation table when the aircraft pitch angle changes negatively within its maneuvering range, and the negative overload adjustment coefficient interpolation table when the aircraft pitch angle changes positively within its maneuvering range. The positive overload adjustment coefficient interpolation table records the correspondence between multiple negative pitch angles and multiple positive overload adjustment coefficients, and the negative overload adjustment coefficient interpolation table records the correspondence between multiple positive pitch angles and multiple negative overload adjustment coefficients. Step S112: Based on the aircraft's current pitch angle, interpolate the corresponding adjustment coefficient from the positive overload adjustment coefficient interpolation table or the negative overload adjustment coefficient interpolation table; Step S113: Correct the boundary value of the overload limit range based on the adjustment coefficient.

4. The normal overload protection control method considering center of gravity changes as described in claim 1, characterized in that, Step S2 further includes: Step S23: Calculate the control surface deflection value required for unit overload under different aircraft, different pressure altitudes, and different indicated airspeeds. The different aircraft refer to aircraft with different flap and slat configurations, weights, and centers of gravity. Step S24: For each aircraft and barometric altitude, using its minimum indicated airspeed as a reference point, calculate the first ratio of other indicated airspeeds to the minimum indicated airspeed. At the same time, calculate the second ratio of the control surface deflection value required per unit overload for other indicated airspeeds to the control surface deflection value required per unit overload for the minimum indicated airspeed. Construct the correspondence between the first ratio and the second ratio. Obtain the final correspondence by averaging the correspondences for all aircraft and barometric altitudes. Step S25: Obtain the aircraft indicated airspeed VIAS and trim speed VIAS_TRIM, take the ratio of indicated airspeed VIAS to trim speed VIAS_TRIM as the first ratio, interpolate the second ratio in the correspondence, and take the second ratio as the gain correction coefficient k_v. Step S26: Based on the gain correction coefficient k_v, perform gain correction on the gain K.

5. The normal overload protection control method considering center of gravity changes as described in claim 4, characterized in that, In step S25, the trimming speed VIAS_TRIM is determined through the following steps: Step S251: When the aircraft is in level flight, the trim speed VIAS_TRIM is the indicated airspeed VIAS; when the aircraft switches from level flight to non-level flight, the trim speed VIAS_TRIM is locked at the indicated airspeed VIAS at the last moment of level flight; when the aircraft switches from non-level flight to level flight, the value of the trim speed VIAS_TRIM changes linearly from the previous trim speed VIAS_TRIM to the indicated airspeed VIAS within a set fade time; when the aircraft is in non-level flight, the value of the trim speed VIAS_TRIM is the trim speed VIAS_TRIM of the previous frame.

6. The normal overload protection control method considering center of gravity changes as described in claim 5, characterized in that, Step S251 further includes determining whether the aircraft is in level flight, specifically: when the aircraft is in steady flight and the displacement of the control stick is less than a set value, the aircraft is in level flight; otherwise, the aircraft is in non-level flight.

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