CN115963854A - Normal overload protection control method considering gravity center change - Google Patents
Normal overload protection control method considering gravity center change Download PDFInfo
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Abstract
The application belongs to the technical field of airplane flight control, and particularly relates to a normal overload protection control method considering gravity center change. The method comprises the steps of S1, converting the steering column displacement Xe 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 unit overload; and S3, superposing 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 airplane to fly based on the elevator deflection instruction. The application realizes the normal overload protection function under the gravity center change and extreme operation working condition, and improves the safety of the airplane.
Description
Technical Field
The application belongs to the technical field of airplane flight control, and particularly relates to a normal overload protection control method considering gravity center change.
Background
Modern airplanes widely adopt fly-by-wire flight control systems, control functions and protection functions are realized through complex control laws, pilot burden is relieved, and flight safety is improved. The normal overload protection function is an important flight boundary protection function for preventing the normal overload of the aircraft from exceeding a limit value. If the normal overload exceeds the limit value, the structure of the airplane body can be damaged, and the airplane can be disassembled in serious cases.
The change of the center of gravity of the airplane has larger influence on normal overload response, and under the same deflection of the elevator, the normal overload response of the rear center of gravity is larger than that of the front center of gravity. The conventional control method adopts a control framework of rod displacement instruction normal overload, and realizes the normal overload protection function under normal flight control (control under a level flight state with unchanged height and speed) at the expense of the stability margin of a flight control system or the maneuverability of an airplane. And the overload of the airplane can exceed a limit value under extreme operating conditions such as quick-dive-after-jump-pull, quick-climb-after-jump-dive, large-pitch-angle-dive-after-jump-pull, flat-flight-acceleration-after-jump-pull and the like.
In the current design process of domestic airplanes, the following defects exist in the design of normal overload protection functions under gravity center change and extreme manipulation:
1. the method has the advantages that the situation that the normal overload protection function covers incomplete working conditions and the airplane has potential danger is assumed that the quick-dive rear step is pulled up, the quick-climb rear step dive and the large-pitch-angle rear step pull-up are performed, the flat-fly acceleration rear step pull-up and other extreme operation working conditions are low in probability, the working conditions are not considered during design and verification, and the normal overload protection function covers incomplete working conditions.
2. In order to take account of the gravity center change, extreme operation conditions such as step pull-up after quick dive, step dive after quick climb, step pull-up after large pitch angle dive, step pull-up after flat flight acceleration and the like, the normal overload value of the rod displacement instruction is reduced, namely the excess allowance of the rear gravity center and the normal overload under the extreme operation conditions is reserved, and the normal overload response of the rod displacement instruction is reduced. The method causes the front center of gravity, the middle center of gravity and normal overload response under normal operation to be reduced, and the sacrifice on the maneuvering capability of the airplane is large.
3. In order to take into account the gravity center change and extreme control working conditions such as step pull-up after quick dive, step dive after quick climb, step pull-up after big pitch angle dive, step pull-up after flat flight acceleration, normal overload response damping characteristic is increased by increasing pitch angle rate feedback gain, thereby reducing the overshoot of normal overload response under rear gravity center and extreme control. Therefore, stability margin of the flight control system is reduced, the anti-interference capability of the control system is reduced, and meanwhile, normal overload response damping of the front center of gravity and the middle center of gravity and normal operation is increased, and normal overload response of the airplane is retarded.
Disclosure of Invention
In order to solve at least one of the technical problems, the application designs a normal overload protection control method considering the change of the gravity center so as to ensure that the normal overload of the airplane does not exceed a limit value under extreme control working conditions such as the change of the gravity center, quick step pull-up after diving, quick step pull-up after climbing, large pitch angle pull-up after diving, and step pull-up after flat flying and accelerating, and the safety of the airplane is improved.
The application provides a normal overload protection control method considering gravity center change, which mainly comprises the following steps:
step S1, converting the driving lever displacement Xe 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 unit overload;
and S3, superposing 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 airplane to fly based on the elevator deflection instruction.
Preferably, the step S1 further includes:
s11, acquiring a stroke range of the displacement Xe of the steering column and simultaneously acquiring an overload limiting range of an airplane structure;
s12, constructing a gradient function for expressing the relation between the displacement and the overload of the steering column based on the travel range and the overload limiting range;
and S13, converting the real-time steering column displacement Xe into a real-time overload command DNy based on the gradient function.
Preferably, step S11 further includes modifying the overload limiting range of the aircraft structure, which includes:
step S111, acquiring a positive overload adjustment coefficient interpolation table when the aircraft pitch angle changes negatively in the maneuvering range of the aircraft and a negative overload adjustment coefficient interpolation table when the aircraft pitch angle changes positively in the maneuvering range of the aircraft, wherein the positive overload adjustment coefficient interpolation table records the corresponding relation between a plurality of negative pitch angles and a plurality of positive overload adjustment coefficients, and the negative overload adjustment coefficient interpolation table records the corresponding relation between a plurality of positive pitch angles and a plurality of negative overload adjustment coefficients;
step S112, interpolating a corresponding adjustment coefficient in the positive overload adjustment coefficient interpolation table or the negative overload adjustment coefficient interpolation table according to the current pitch angle of the airplane;
and a step S113 of correcting the boundary value of the overload limiting range based on the adjustment coefficient.
Preferably, the step S2 further includes:
s21, acquiring a trim rudder deflection De _ trim of the airplane and a rudder deflection De _ ny0 required by normal overload of 0g, wherein g is gravity acceleration;
step S22, determining the gain K as: k = De _ trim-De _ ny0.
Preferably, in step S21, the obtaining of the trim rudder offset De _ trim includes:
s211, acquiring a pitch angle and a roll angle of the airplane;
step S212, calculating normal overload of a body axis based on the pitching angle and the rolling angle of the airplane;
step S213, determining whether the aircraft flies in a steady state or not according to the difference between the feedback value of the normal overload sensor and the calculated normal overload of the body axis system, wherein when the difference is smaller than a threshold value, the aircraft flies in the steady state, the threshold value is 0.04 g-0.06g, and g is gravity acceleration;
and S214, when the airplane is in steady-state flight, forming an airplane trim rudder deviation De _ trim by using the integral rudder deviation command as input through an integrator, otherwise, forming the airplane trim rudder deviation De _ trim by using 0 as input through the integrator.
Preferably, the step S2 further includes:
s23, respectively calculating control surface deflection values required by unit overload under different airplanes, different barometric altitudes and different indicated airspeeds, wherein the different airplanes are airplanes with different flap configurations, weights and gravity centers;
step S24, calculating a first ratio of other indicated airspeeds to the minimum indicated airspeed of each airplane and each barometric altitude by taking the minimum indicated airspeed of the airplane and the barometric altitude as a reference point, calculating a second ratio of a control surface deviation value required by unit overload corresponding to the other indicated airspeeds to a control surface deviation value required by unit overload of the minimum indicated airspeed, constructing a corresponding relation of the first ratio and the second ratio, and obtaining a final corresponding relation of all the airplanes and the barometric altitudes in a mean value calculation mode;
s25, acquiring an indicated airspeed VIAS and a TRIM speed VIAS _ TRIM of the airplane, taking the ratio of the indicated airspeed VIAS to the TRIM speed VIAS _ TRIM as the first ratio, interpolating a second ratio in the corresponding relation, and taking the second ratio as a gain correction coefficient k _ v;
and step S26, performing gain correction on the gain K based on the gain correction coefficient K _ v.
Preferably, in step S25, the TRIM speed VIAS _ TRIM is determined by:
step 251, when the aircraft is in a level flight state, the TRIM speed VIAS _ TRIM is an indicated airspeed VIAS; when the plane is switched from the level flight state to the non-level flight state, locking the TRIM speed VIAS _ TRIM at the indicated airspeed VIAS at the last moment of the level flight state; when the airplane is switched from the non-level flight state to the level flight state, the value of the TRIM speed VIAS _ TRIM is linearly changed into an indicated airspeed VIAS from the TRIM speed VIAS _ TRIM of the previous beat in the set fade time, and when the airplane is in the non-level flight state, the value of the TRIM speed VIAS _ TRIM is the TRIM speed VIAS _ TRIM of the previous beat.
Preferably, step S251 further includes determining whether the aircraft is in a flat flight state, specifically including: when the aircraft flies in a steady state and the displacement of the steering column is smaller than a set value, the aircraft flies in a level flight state, otherwise, the aircraft flies in a non-level flight state.
The application can reserve the traditional control law framework and realize a normal overload protection control function. The application designs a normal direction overload protection control law of considering focus change, introduce pitch angle adjustment overload limit value, it calculates forward passageway gain to obtain aircraft trim rudder in real time partially, and obtain aircraft trim speed in real time, and forward passageway gain to the relation adjustment according to real-time speed and trim speed, thereby solve focus change and step pull-up behind the quick dive, step dive behind the quick climb, step pull-up behind the big pitch angle dive, the problem that aircraft overload exceeded the limit value under extreme operating mode such as step pull-up behind the flat flying acceleration. According to the method, any hardware part of the flight control system does not need to be modified, a normal overload protection control law can be designed according to the method, control law software of the flight control system is changed, modification cost is saved, the normal overload protection function under the gravity center change and extreme operation working conditions is realized, and the safety of the airplane is improved.
Drawings
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 gravity center change.
Fig. 2 is a schematic diagram of a normal overload protection control architecture of an additional trim rudder deflection obtaining module.
Fig. 3 is a schematic diagram of a trim rudder deflection acquisition module.
FIG. 4 is a diagram showing the variation curve of the ratio of the indicated airspeed and the required control surface deflection per unit overload to the reference state at each state point.
FIG. 5 is a schematic diagram of trim speed acquisition logic.
FIG. 6 is a schematic diagram of a logic diagram for determining a level flight status of an aircraft.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the accompanying drawings in the embodiments of the present application. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are some, but not all embodiments of the present application. The embodiments described below with reference to the drawings are exemplary and intended to be used for explaining the present application, and should not be construed as limiting the present application. All other embodiments obtained by a person of ordinary skill in the art without any inventive work based on the embodiments in the present application are within the scope of protection of the present application. Embodiments of the present application will be described in detail below with reference to the drawings.
The application provides a normal overload protection control method considering gravity center change, as shown in fig. 1, which mainly comprises the following steps:
step S1, converting the steering column displacement Xe into an overload command 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 unit overload;
and S3, superposing 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 airplane to fly based on the elevator deflection instruction.
Referring to fig. 1, the normal overload protection control architecture is composed of a forward channel, a proportional channel and an integral channel, wherein the forward channel converts rod displacement Xe into an overload instruction DNy, and then converts the overload instruction DNy into a forward rudder deflection instruction after multiplying by a gain K (rudder deflection required by unit overload); introducing pitch angle rate and normal overload feedback to the proportional channel to improve normal overload dynamic response, and converting the pitch angle rate and the normal overload feedback into proportional rudder deflection instructions after multiplying by gains respectively; and the integral channel realizes the accurate control of the normal overload instruction through an integrator, and multiplies the difference between the actual normal overload instruction and the normal overload instruction of the airplane by a gain to convert the difference into an integral rudder deflection instruction. The forward rudder deflection instruction, the proportional rudder deflection instruction and the integral rudder deflection instruction jointly generate an elevator deflection instruction, and the control of the airplane is realized.
The calculation mode of the proportional rudder deflection instruction and the integral rudder deflection instruction is the existing control law framework, so that the method can be improved on the existing control law framework, any hardware part of a flight control system does not need to be modified, the normal overload protection control law can be designed according to the method, the control law software of the flight control system is changed, the modification cost is saved, the normal overload protection function under the gravity center change and extreme control working conditions is realized, and the safety of the airplane is improved.
In some optional embodiments, step S1 further comprises:
s11, acquiring a stroke range of the displacement Xe of the steering column and simultaneously acquiring an overload limiting range of an airplane structure;
s12, constructing a gradient function for expressing the relation between the displacement of the steering column and the overload based on the travel range and the overload limiting range;
and S13, converting the real-time steering column displacement Xe into a real-time overload instruction DNy based on the gradient function.
In this embodiment, referring to fig. 1, an overload limiting module in the forward path implements conversion of the steering column displacement Xe into overload instructions DNy. For convenience of description, assuming that the Xe range is [ -100,100] mm, the rod displacement is positive and then push rod manipulation, and the rod displacement is negative and then pull rod manipulation, the overload limiting range of the aircraft structure is [ -1,2.5] g, by designing the rod displacement-overload gradient K _ Xe _ Ny (i.e. the gradient function in step S12, the design method is a conventional method and is not described again), the following correspondence can be achieved:
DNy = -1 when Xe = 100;
DNy =2.5 when Xe = -100;
when-100 < Xe < 100, DNy = K _Xe _Ny × Xe.
In this way, the real-time steering column displacement Xe can be converted into the real-time overload command DNy.
In some optional embodiments, step S11 further includes modifying the overload limiting range of the aircraft structure, which includes:
step S111, acquiring a positive overload adjustment coefficient interpolation table when the aircraft pitch angle changes negatively in the maneuvering range of the aircraft and a negative overload adjustment coefficient interpolation table when the aircraft pitch angle changes positively in the maneuvering range of the aircraft, wherein the positive overload adjustment coefficient interpolation table records the corresponding relation between a plurality of negative pitch angles and a plurality of positive overload adjustment coefficients, and the negative overload adjustment coefficient interpolation table records the corresponding relation between a plurality of positive pitch angles and a plurality of negative overload adjustment coefficients;
step S112, interpolating a corresponding adjustment coefficient in the positive overload adjustment coefficient interpolation table or the negative overload adjustment coefficient interpolation table according to the current pitch angle of the airplane;
step S113 corrects the boundary value of the overload restriction range based on the adjustment coefficient.
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 the full tie rod 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. The normal overload command is gradually restored to a maximum as the pitch angle decreases. Assuming that the maneuvering range of the pitching angle of the airplane is [ -30, 30] °, the following table 1 is a positive overload adjustment coefficient interpolation table M1, and the following table 2 is a negative overload adjustment coefficient interpolation table M2. The adjustment coefficients in M1 and M2 may be determined from simulations performed on a particular aircraft.
TABLE 1 interpolation of forward overload adjustment coefficients Table M1
| Angle of pitch (°) | -30 | -20 | -10 | -5 |
| Positive overload adjustment factor | 0.85 | 0.9 | 0.95 | 1 |
TABLE 2 interpolation table of negative overload adjustment coefficient M2
| Angle of pitch (°) | 5 | 10 | 20 | 30 |
| Negative |
1 | 0.95 | 0.9 | 0.85 |
Then, correction values M1 and M2 corresponding to the pitch angles are interpolated from M1 and M2, and DNy corresponding to the rod displacements of 100M and-100M of the above embodiment is adjusted using M1 and M2 as follows:
when Xe =100, dny = -1 × m2;
when Xe = -100, dny =2.5 × m1.
In some optional embodiments, step S2 further comprises:
s21, acquiring a plane trim rudder deflection De _ trim and a rudder deflection De _ ny0 required by 0g normal overload, wherein g is gravity acceleration;
step S22, determining the gain K as: k = De _ trim-De _ ny0.
The gain K in step S2, that is, the gain K (rudder deflection required by unit overload) in the forward channel in fig. 1 is greatly related to the center of gravity of the aircraft, and the gain of the forward center of gravity is greater than the gain of the rear center of gravity, and in order to match the gains with the center of gravity, a calculation module for obtaining the aircraft trim rudder deflection in real time is added to the integral channel in fig. 1, so as to construct a calculation method for the gain K of the forward channel, where the calculation method is:
forward channel gain = trim rudder bias-0 g rudder bias required for normal overload.
Because the trim rudder deflection can reflect the gravity center factor of the airplane, and the rudder deflection required by the normal overload of 0g is irrelevant to the gravity center of the airplane, the method is adopted to calculate the gain of the forward channel, and the matching with the gravity center can be ensured.
It should be noted that De _ ny0 is the rudder deflection required for 0g normal overload. The calculation of De _ ny0 can adopt a traditional calculation method, namely, a rudder bias generating-1 g normal overload is solved, and the detailed description is omitted here.
In some optional embodiments, the obtaining the trim rudder bias De _ trim in step S21 includes:
s211, acquiring a pitch angle and a roll angle of the airplane;
step S212, calculating normal overload of a body axis based on the pitch angle and the roll angle of the airplane;
step S213, determining whether the aircraft flies in a steady state or not according to the difference between the feedback value of the normal overload sensor and the calculated normal overload of the body axis system, wherein when the difference is smaller than a threshold value, the aircraft flies in the steady state, the threshold value is 0.04 g-0.06g, and g is gravity acceleration;
and S214, when the airplane is in steady-state flight, using the integral rudder deflection command as input to form an airplane trim rudder deflection De _ trim through an integrator, otherwise, using 0 as input to form the airplane trim rudder deflection De _ trim through the integrator.
The embodiment mainly describes a trim rudder deflection obtaining method, which specifically includes: and calculating the normal overload during steady-state flight according to the rolling angle and the pitching angle of the airplane, and judging whether the airplane really flies in the steady state or not by combining the value of the normal overload sensor. When the value of the normal overload sensor is basically equal to the normal overload in the steady-state flight obtained by calculation, the aircraft is in the steady-state flight state; otherwise, it represents that the aircraft is not in a steady-state flight condition. The core of a calculation module of the airplane trim rudder deflection is an integrator, an integrated quantity is generated by a numerical value of an integration channel when the airplane is in a steady-state flight state, the airplane trim rudder deflection is generated after passing through the integrator, and the integrated quantity is fed back to the integration channel to counteract the integrated quantity of the integrator of the integration channel, so that the numerical value of the integration channel is reduced. Finally, the output of the aircraft trim rudder deflection calculation module replaces the integral channel when the aircraft is in a steady state flight state. If the airplane is not in a steady-state flying state, the integrated quantity of the integrator of the airplane trim rudder deflection calculation module is 0.
In this embodiment, referring to fig. 2, the calculating module of the trim rudder deflection generates an airplane trim rudder deflection De _ trim and De _ OffLoad according to the integral channel output De _ cmd _ int, where the De _ OffLoad is used to offset the integrated quantity of the integral channel integrator. And the forward rudder deflection instruction, the proportional rudder deflection instruction, the integral rudder deflection instruction and the airplane trim rudder deflection De _ trim are used for generating an elevator deflection instruction together.
The aircraft TRIM rudder offset De _ TRIM is not only used to calculate the gain K, but also directed to the latter synthesizer according to the description of fig. 2, which is mainly determined by the TRIM rudder offset acquisition module itself, see fig. 3, where TRIM _ SW switches are logically as follows:
in steady-state flight, the normal overload of the body axis system is
Wherein pitch is the pitch angle and bank is the roll angle. A normal overload sensor value Ny, defining>
Then: when | Error _ Ny-<0.05g, TRIM _ SW =1, i.e. De _ cmd _ int as input; when | Error _ Ny | ≧ 0.05g, TRIM _ SW =0, that is, 0 is input, and when De _ cmd _ int is input, the output De _ OffLoad value coincides with the input De _ cmd _ int, and after subtraction before the integrator of fig. 2, the value is subjected to subtractionThe De _ cmd _ int value will become 0, so the airplane trim rudder bias De _ trim is taken as input to the back combiner instead of the original integration channel.
In the embodiment, according to the pitch angle, the roll angle, the normal overload sensor measurement accuracy and the zero characteristic, when | Error _ Ny | <0.05g is tentatively determined, the aircraft is considered to be in steady-state flight, and the threshold of 0.05g can be adjusted according to the specific sensor characteristic.
In some optional embodiments, step S2 further comprises:
s23, respectively calculating control plane deflection values required by unit overload of different airplanes, different air pressure altitudes and different indicated airspeeds, wherein the different airplanes are airplanes with different flap configurations, weights and gravity centers;
step S24, calculating a first ratio of other indicated airspeeds to the minimum indicated airspeed of each airplane and each barometric altitude by taking the minimum indicated airspeed of the airplane and the barometric altitude as a reference point, calculating a second ratio of a control surface deviation value required by unit overload corresponding to the other indicated airspeeds to a control surface deviation value required by unit overload of the minimum indicated airspeed, constructing a corresponding relation of the first ratio and the second ratio, and obtaining a final corresponding relation of all the airplanes and the barometric altitudes in a mean value calculation mode;
s25, acquiring an indicated airspeed VIAS and a TRIM speed VIAS _ TRIM of the airplane, taking the ratio of the indicated airspeed VIAS to the TRIM speed VIAS _ TRIM as the first ratio, interpolating a second ratio in the corresponding relation, and taking the second ratio as a gain correction coefficient k _ v;
and step S26, performing gain correction on the gain K based on the gain correction coefficient K _ v.
In the embodiment, state points covering the weight, the gravity center, the configuration of each slat and the speed range of the airplane are selected, the deflection of a control surface required by unit overload under each state point is calculated, the change condition along with the increase of the speed is obtained, and the attenuation coefficient of the forward gain is selected by adopting an averaging method.
Taking a cruise configuration, a weight of 110000kg and a middle gravity center as an example, heights of 3000m and 5000m are selected, an indicated airspeed is selected from small, medium and large speeds in a flight envelope, and the deflection of the control surface required by unit overload is calculated as shown in table 3 (the calculation of the deflection of the control surface required by unit overload belongs to a traditional method and is not described herein any more).
TABLE 3 example of control surface skewness calculation required for unit overload
And (4) taking the minimum speed of each altitude as a reference state, and plotting the ratio of the deviation of the control surface required by the indicated airspeed and the unit overload of other state points at the altitude to the reference state. Taking 3000m altitude as an example, the state point 1 is taken as a reference state, and a curve as shown in fig. 4 is drawn, wherein the abscissa is the ratio of the indicated airspeeds of the state points 2 and 3 and the state point 1, and the ordinate is the ratio of the control surface deflection required by unit overload of the state points 2 and 3 and the state point 1. Similarly, a curve with a height of 5000m is plotted.
Taking the average of the ordinate of the two curves in fig. 4, the attenuation factor k _ v can be determined, as shown in table 4.
TABLE 4 attenuation coefficient k _ v
| VIAS/ |
1 | 1.3 | 1.5 | 1.6 |
| |
1 | 0.57 | 0.42 | 0.2 |
In actual design, curves under a plurality of weights, centers of gravity, heights and configurations need to be drawn, and the average value of the curves is taken to determine k _ v. If the curve has larger difference along with the configuration, etc., k _ v can be made into a multi-dimensional interpolation table and adjusted along with the configuration, etc.
Finally, according to the ratio of the current indicated airspeed VIAS to the TRIM speed VIAS _ TRIM, adjusting the gain K in the forward channel, wherein the gain K is changed into: k = (De _ trim-De _ ny 0) × K _ v.
In some alternative embodiments, in step S25, the TRIM speed VIAS _ TRIM is determined by:
step 251, when the aircraft is in a level flight state, the TRIM speed VIAS _ TRIM is an indicated airspeed VIAS; when the plane is switched from the level flight state to the non-level flight state, locking the TRIM speed VIAS _ TRIM at the indicated airspeed VIAS at the last moment of the level flight state; when the airplane is switched from the non-level flight state to the level flight state, the value of the TRIM speed VIAS _ TRIM is changed from the TRIM speed VIAS _ TRIM of the previous beat to the indication airspeed VIAS within the set desalination time, and when the airplane is in the non-level flight state, the value of the TRIM speed VIAS _ TRIM is the TRIM speed VIAS _ TRIM of the previous beat.
In this embodiment, the set fade time is generally 2s, and referring to fig. 5, when PF _ ST changes from 1 to 0, no fade processing is performed, and via _ TRIM is via _ TRIM of the last beat; when PF _ ST is changed from 0 to 1, fade processing is carried out, fade time is 2s, namely the value of VIAS _ TRIM is that VIAS _ TRIM of the last beat is changed to VIAS linearly through 2 s.
In some optional embodiments, step S251 further includes determining whether the aircraft is in a level flight state, specifically including: when the aircraft flies in a steady state and the displacement of the steering column is smaller than a set value, the aircraft flies in a level flight state, otherwise, the aircraft flies in a non-level flight state.
Referring to fig. 6, it is determined whether the aircraft is in a level flight state through the and gate, and the pilot is considered to be not manipulated when the tentative lever displacement is less than 2mm according to the mechanical characteristics of the pilot lever and the characteristics of the lever displacement sensor, and the value can be adjusted according to the specific pilot lever characteristics. When PF _ ST is 1 in 1s, the plane is in a level flight state; otherwise, the airplane is in a non-level flight state, namely, in a maneuvering state.
It can be understood that when the speed of the airplane changes, the numerical value of the integral channel changes, so that the deviation of the trim rudder also changes, namely the deviation of the trim rudder not only reflects the gravity center factor of the airplane, but also reflects the speed change condition of the airplane. Under the same gravity center, the speed of the airplane is increased, and the deflection of the trim rudder is reduced; the speed of the airplane is reduced, and the offset of the trim rudder is increased. The change of the trim rudder deviation lags behind the change of the airplane speed, when the airplane speed is increased too fast, the trim rudder deviation can be mismatched with the airplane speed, namely, the condition that the trim rudder is slightly larger appears, so that the forward gain is slightly larger, and the normal overload is easy to exceed the limit. In addition, the trim rudder does not change when the aircraft is not in steady-state flight. That is, when the aircraft is not in steady-state flight and the speed is increased, the trim rudder is deflected to be unchanged, so that the forward channel gain is basically unchanged (the rudder deflection required by 0g normal overload is small along with the speed), and the elevator angle generated by operation is increased and is not matched with the aircraft speed, so that the normal overload is over-limited. Therefore, the present application solves this problem through the above-described step S23 to step S26.
Specifically, when the airplane is in a flat flight state, the real-time indicated airspeed of the airplane is taken as the trim speed, and real-time updating is carried out; when the airplane is not in the flat flying state, the trim speed is kept to be the trim speed value of the last beat. In order to ensure that the trim speed does not jump, when the non-level flight state is changed into the level flight state, the trim speed gradually fades from the last beat value to the current indicated airspeed. Adjusting forward gain in a forward channel according to the ratio of the current indicated speed to the trim speed, wherein when the ratio is greater than 1, the current indicated airspeed of the airplane deviates from the trim speed more, and normal overload overrun caused by over-limit of rudder effect is avoided, so that the forward gain needs to be attenuated; when the ratio is less than 1, the current indicated speed of the airplane is less than the trim speed, and the normal overload of the airplane cannot exceed the limit under the condition, and the forward gain does not need to be processed.
The application can reserve the traditional control law framework and realize a normal overload protection control function. The application designs a normal direction overload protection control law of considering focus change, introduce pitch angle adjustment overload limit value, it calculates forward passageway gain to obtain aircraft trim rudder in real time partially, and obtain aircraft trim speed in real time, and forward passageway gain to the relation adjustment according to real-time speed and trim speed, thereby solve focus change and step pull-up behind the quick dive, step dive behind the quick climb, step pull-up behind the big pitch angle dive, the problem that aircraft overload exceeded the limit value under extreme operating mode such as step pull-up behind the flat flying acceleration. According to the method, any hardware part of the flight control system does not need to be modified, a normal overload protection control law can be designed according to the method, the control law software of the flight control system is changed, the modification cost is saved, the functions of gravity center change and normal overload protection under extreme control working conditions are realized, and the safety of the airplane is improved.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (8)
1. A normal overload protection control method considering gravity center change is characterized by comprising the following steps:
step S1, converting the steering column displacement Xe into an overload command 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 unit overload;
and S3, superposing 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 airplane to fly based on the elevator deflection instruction.
2. The normal overload protection control method considering the change of the center of gravity according to claim 1, wherein the step S1 further comprises:
s11, acquiring a stroke range of the displacement Xe of the steering column and acquiring an overload limiting range of the airplane structure;
s12, constructing a gradient function for expressing the relation between the displacement of the steering column and the overload based on the travel range and the overload limiting range;
and S13, converting the real-time steering column displacement Xe into a real-time overload command DNy based on the gradient function.
3. The method for normal overload protection control in consideration of center of gravity change according to claim 2, wherein the step S11 further comprises modifying an overload limiting range of the aircraft structure, which includes:
step S111, acquiring a positive overload adjustment coefficient interpolation table when the pitching angle of the airplane is negatively changed in the maneuvering range of the airplane and a negative overload adjustment coefficient interpolation table when the pitching angle of the airplane is positively changed in the maneuvering range of the airplane, wherein the positive overload adjustment coefficient interpolation table records the corresponding relation between a plurality of negative pitching angles and a plurality of positive overload adjustment coefficients, and the negative overload adjustment coefficient interpolation table records the corresponding relation between a plurality of positive pitching angles and a plurality of negative overload adjustment coefficients;
step S112, interpolating a corresponding adjusting coefficient in the positive overload adjusting coefficient interpolation table or the negative overload adjusting coefficient interpolation table according to the current pitch angle of the airplane;
and a step S113 of correcting the boundary value of the overload limiting range based on the adjustment coefficient.
4. The normal overload protection control method considering the change of the center of gravity according to claim 1, wherein the step S2 further comprises:
s21, acquiring a trim rudder deflection De _ trim of the airplane and a rudder deflection De _ ny0 required by normal overload of 0g, wherein g is gravity acceleration;
step S22, determining the gain K as: k = De _ trim-De _ ny0.
5. The normal overload protection control method considering the change of the center of gravity according to claim 4, wherein the obtaining of the trim rudder deflection De _ trim in step S21 includes:
s211, acquiring a pitch angle and a roll angle of the airplane;
step S212, calculating normal overload of a body axis based on the pitching angle and the rolling angle of the airplane;
step S213, determining whether the aircraft flies in a steady state or not according to the difference between the feedback value of the normal overload sensor and the calculated normal overload of the body axis system, wherein when the difference is smaller than a threshold value, the aircraft flies in the steady state, the threshold value is 0.04 g-0.06g, and g is gravity acceleration;
and S214, when the airplane is in steady-state flight, forming an airplane trim rudder deviation De _ trim by using the integral rudder deviation command as input through an integrator, otherwise, forming the airplane trim rudder deviation De _ trim by using 0 as input through the integrator.
6. The normal overload protection control method considering the change of the center of gravity according to claim 4, wherein the step S2 further comprises:
s23, respectively calculating control plane deflection values required by unit overload of different airplanes, different air pressure altitudes and different indicated airspeeds, wherein the different airplanes are airplanes with different flap configurations, weights and gravity centers;
step S24, for each airplane and each barometric altitude, taking the minimum indicated airspeed of the airplane as a reference point, calculating a first ratio of other indicated airspeeds to the minimum indicated airspeed, calculating a second ratio of control surface deviation values required by unit overload corresponding to the other indicated airspeeds to the control surface deviation values required by unit overload of the minimum indicated airspeed, constructing a corresponding relation of the first ratio and the second ratio, and obtaining a final corresponding relation of all airplanes and the barometric altitudes in a mean value calculation mode;
s25, acquiring an indicated airspeed VIAS and a TRIM speed VIAS _ TRIM of the airplane, taking the ratio of the indicated airspeed VIAS to the TRIM speed VIAS _ TRIM as the first ratio, interpolating a second ratio in the corresponding relation, and taking the second ratio as a gain correction coefficient k _ v;
and step S26, performing gain correction on the gain K based on the gain correction coefficient K _ v.
7. The center-of-gravity variation-considered normal overload protection control method according to claim 6, wherein in step S25, the TRIM speed via _ TRIM is determined by:
step S251, when the aircraft is in a level flight state, the TRIM speed VIAS _ TRIM is an indicated airspeed VIAS; when the plane is switched from the level flight state to the non-level flight state, locking the TRIM speed VIAS _ TRIM at the indicated airspeed VIAS at the last moment of the level flight state; when the airplane is switched from the non-level flight state to the level flight state, the value of the TRIM speed VIAS _ TRIM is changed from the TRIM speed VIAS _ TRIM of the previous beat to the indication airspeed VIAS within the set desalination time, and when the airplane is in the non-level flight state, the value of the TRIM speed VIAS _ TRIM is the TRIM speed VIAS _ TRIM of the previous beat.
8. The method for normal overload protection control in view of center of gravity change as claimed in claim 7, wherein the step S251 further includes determining whether the aircraft is in a level flight state, specifically including: when the aircraft flies in a steady state and the displacement of the steering column is smaller than a set value, the aircraft flies in a level flight state, otherwise, the aircraft flies in a non-level flight state.
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| CN117390774A (en) * | 2023-12-13 | 2024-01-12 | 中国航空工业集团公司西安飞机设计研究所 | Force correction method for aircraft pitching maneuvering control lever |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN117390774A (en) * | 2023-12-13 | 2024-01-12 | 中国航空工业集团公司西安飞机设计研究所 | Force correction method for aircraft pitching maneuvering control lever |
| CN117390774B (en) * | 2023-12-13 | 2024-03-19 | 中国航空工业集团公司西安飞机设计研究所 | Force correction method for aircraft pitching maneuvering control lever |
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