Fault isolation method for active damping device of aircraft landing gear

Abstract

The embodiment of the invention provides a fault isolation method for an active damping device of an aircraft landing gear, and belongs to the technical field of aviation, and the method specifically comprises the steps: carrying out the modeling of a single landing gear, obtaining a vertical-direction kinetic equation, and constructing a state-space equation of a single landing gear model according to the vertical-direction kinetic equation; expanding the state-space equation of the single undercarriage model to a complete machine to obtain a state-space equation of a complete machine undercarriage; fault models of different fault types are established for the actuator and the sensor; according to the fault model of the actuator and the fault model of the sensor, the state-space equation of the complete machine undercarriage is rewritten into a fault-containing state-space equation, and an unknown input observer is designed according to the state-space equation; according to an output residual error of an unknown input observer, a vertical displacement component residual error root-mean-square value is adopted to carry out actuator fault isolation judgment; and judging the fault isolation of the sensor according to the output residual values of the n sub observers. Through the scheme disclosed by the invention, the fault isolation efficiency, accuracy and adaptability are improved.

Description

Technical Field

[0001] The disclosed embodiments relate to the field of aviation technology, and more particularly to a fault isolation method for an active shock absorbing device of an aircraft landing gear. Background Art

[0002] As the most important long-distance transportation tools in the contemporary world, airplanes and high-speed trains have very high moving speeds and their internal integration is also very precise. The safety and stability of these precision parts are also very important. According to domestic and foreign statistics, many aircraft accidents occur during the taxiing phase of the aircraft and are related to the landing gear. As a key component of the aircraft, the landing gear plays a vital role in the aircraft's takeoff, landing, ground taxiing, and emergency response. Not only must it be able to bear the entire weight of the aircraft and ensure the stability of the aircraft when parked; it must also reduce vibration during takeoff, landing, and ground taxiing to ensure the smooth operation of the aircraft body.

[0003] With the continuous development of technology, in order to improve ride comfort, the software and hardware technologies of landing gear are constantly developing, and the control system has also shifted from traditional passive control to active control with better shock absorption performance. However, these have made the landing gear system structure more complex and increased the risk of failure. Therefore, improving the safety of landing gear has become an important topic that needs to be studied urgently.

[0004] When the actuator of the landing gear fails, the actuator cannot output the required active control force according to the requirements of the controller, resulting in a deterioration in the motion state of the aircraft during taxiing. Since the working environment of the landing gear system is relatively harsh and the actuator needs to perform a very large actuation force load, the probability of actuator failure is high. The sensor is an important device for measuring the state of the system. When the sensor fails, it cannot measure the correct data signal, so the data transmitted to the controller is different from the actual value, and the control result given by the controller is also inaccurate. Therefore, actuator failure and sensor failure are a research focus.

[0005] When a landing gear fails, detailed fault information needs to be obtained quickly and processed in time. However, the current research on landing gear fault diagnosis is relatively simple, with little isolation and reconstruction of faults, making it difficult to obtain detailed fault information. Therefore, when a landing gear fails, it is necessary to isolate the actuator or sensor fault based on the fault residual information and determine the specific location of the fault.

[0006] It can be seen that there is an urgent need for a fault isolation method for an aircraft landing gear active shock absorbing device with high fault isolation efficiency, accuracy and adaptability. Summary of the invention

[0007] In view of this, an embodiment of the present disclosure provides a fault isolation method for an active shock absorbing device of an aircraft landing gear, which at least partially solves the problems of poor fault isolation efficiency, accuracy and adaptability in the prior art.

[0008] The present disclosure provides a fault isolation method for an active shock absorbing device of an aircraft landing gear, comprising:

[0009] Step 1, modeling the single landing gear, analyzing the vertical dynamic equation of the single landing gear model according to the law of dynamics, and constructing the state space equation of the single landing gear model accordingly;

[0010] Step 2, combining the coupling of the roll motion, pitch motion and vertical motion of the aircraft fuselage, expanding the state space equation of the single landing gear model to the whole aircraft, and combining the motion equation of each wheel and the vertical motion displacement to obtain the state space equation of the whole aircraft landing gear;

[0011] Step 3, establishing fault models of different fault types for the actuators and sensors of the landing gear;

[0012] Step 4: According to the fault model of the actuator and the fault model of the sensor, the state space equation of the whole landing gear is rewritten into a state space equation including the fault and the unknown input observer is designed accordingly;

[0013] Step 5, when the actuator fails, the actuator fault isolation is determined by using the residual root mean square value of the vertical displacement component according to the output residual of the unknown input observer;

[0014] Step 6: Set n sub-observers. When a sensor fails, the sensor fault isolation is determined based on the output residual values ​​of the n sub-observers.

[0015] According to a specific implementation of the embodiment of the present disclosure, the vertical dynamic equation of the single landing gear model is expressed as:

[0016]

[0017] Where M represents the fuselage mass, m t Mainly includes the quality of wheels and tires, is the vertical acceleration of the fuselage, is the vertical acceleration of the wheel, k s is the landing gear stiffness coefficient, k t is the equivalent stiffness coefficient of the tire, c s is the landing gear damping coefficient, the vertical displacement of the fuselage z s , vertical speed of fuselage Wheel vertical displacement z t , wheel vertical speed is the state variable, that is Control input u d (t) = [u], interference input q d (t) = [q];

[0018] The state space equation of the single landing gear model is expressed as

[0019]

[0020] Among them, the expressions of matrices A0, B0 and D0 are

[0021]

[0022] According to a specific implementation of the embodiment of the present disclosure, the state space equation of the whole aircraft landing gear is expressed as follows:

[0023]

[0024] in, C=I 12×12 , A5=[0 1],A6=[P2l2P2], O m×n represents a zero matrix of m rows and n columns, I m×n Represents the identity matrix with m rows and n columns.

[0025] According to a specific implementation of the embodiment of the present disclosure, the expression of the fault model of the actuator is:

[0026]

[0027] Among them, u zf (t) is the actuator fault output, and u zf (t) is bounded, η i , i=1,2,3 is the actuator gain fault coefficient, and 0≤η i ≤1,α i ,i=1,2,3 is the actuator constant deviation fault coefficient;

[0028] The expression of the fault model of the sensor is:

[0029]

[0030] Among them, y zf (t) is the sensor fault output, and y zf (t) is bounded, δ j ,j=1,2,…,12 is the sensor gain fault coefficient, and 0≤δ j≤1, β j ,j=1,2,…,12 is the constant deviation fault coefficient of the sensor.

[0031] According to a specific implementation of the embodiment of the present disclosure, the expression of the state space equation containing the fault is:

[0032]

[0033] Among them, f u (t), f y (t) represent the actuator fault signal and sensor fault signal respectively;

[0034] The expression of the unknown input observer is:

[0035]

[0036] Among them, w z (t) represents the observer state variable, Is for x z (t) state estimation variable, F∈R n×n , G∈R n×n , H∈R n×l , K∈R n×l is the observer matrix that needs to be solved for the design.

[0037] According to a specific implementation of the embodiment of the present disclosure, the expression of the output residual is:

[0038]

[0039] Among them, when there is no actuator or sensor failure in the landing gear, the output residual r z (t)→0.

[0040] According to a specific implementation of the embodiment of the present disclosure, step 5 specifically includes:

[0041] The fuselage vertical displacement components at each landing gear are selected for analysis. By comparing the residual root mean square values ​​of the fuselage vertical displacement components of each landing gear, the actuator with a residual value greater than that of the other two actuators is determined as the faulty actuator, thereby achieving actuator fault isolation.

[0042] According to a specific implementation of the embodiment of the present disclosure, step 6 specifically includes:

[0043] Set up a set of parallel sub-observers UIO1,…,UIO n , and the jth observer UIO j Do not introduce the jth state variable y j , j = 1, 2, ..., n, by observing the observer UIOj Whether the residual of the output tends to 0, while the residual of other parallel observers output does not tend to 0, the fault isolation of the sensor is realized.

[0044] The fault isolation scheme of the active shock absorbing device of the aircraft landing gear in the embodiment of the present disclosure includes: step 1, modeling a single landing gear, analyzing the vertical direction dynamic equation of the single landing gear model according to the law of dynamics, and constructing the state space equation of the single landing gear model accordingly; step 2, combining the coupling of the roll motion, pitch motion and vertical motion of the aircraft fuselage, expanding the state space equation of the single landing gear model to the whole aircraft, and combining the motion equation of each wheel and the vertical motion displacement to obtain the state space equation of the whole aircraft landing gear; step 3, establishing fault models of different fault types for the actuator and sensor of the landing gear; step 4, rewriting the state space equation of the whole aircraft landing gear into a state space equation containing faults according to the fault model of the actuator and the fault model of the sensor, and designing an unknown input observer accordingly; step 5, when the actuator fails, according to the output residual of the unknown input observer, the vertical displacement component residual root mean square value is used to determine the actuator fault isolation; step 6, setting n sub-observers, when the sensor fails, the sensor fault isolation is determined according to the output residual values ​​of the n sub-observers.

[0045] The beneficial effects of the embodiments of the present disclosure are as follows: through the scheme of the present disclosure, by extending the single landing gear model to the whole aircraft, the coupling relationship of multiple motions is taken into account, making the model more comprehensive and accurate; a fault model for actuators and sensors is established, which can effectively identify different types of faults and enhance the reliability of the system; by rewriting the state space equations and designing unknown input observers, the system can monitor and respond to faults in real time, thereby improving the sensitivity of fault detection; through output residual analysis, the source of the fault can be determined quickly and accurately, reducing the complexity and response time of fault handling; the mechanism of setting multiple sub-observers makes the system have good scalability and can adapt to landing gear systems of different sizes and complexities; by timely identifying and isolating faults, it helps to ensure the safety of the aircraft during take-off and landing, and improves the reliability of the overall operation. BRIEF DESCRIPTION OF THE DRAWINGS

[0046] In order to more clearly illustrate the technical solutions of the embodiments of the present disclosure, the drawings required for use in the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present disclosure. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0047] Figure 1 A schematic flow chart of a fault isolation method for an active shock absorbing device of an aircraft landing gear provided by an embodiment of the present disclosure;

[0048] Figure 2 A landing gear model diagram provided by an embodiment of the present disclosure;

[0049] Figure 3 An actuator fault isolation strategy diagram provided by an embodiment of the present disclosure;

[0050] Figure 4 A sensor fault isolation strategy diagram provided by an embodiment of the present disclosure;

[0051] Figure 5 The residual output of the sub-observer of Experiment 1 provided in the embodiment of the present disclosure;

[0052] Figure 6 The residual output of the sub-observer of Experiment 2 provided in the embodiment of the present disclosure;

[0053] Figure 7 The residual output of the sub-observer of Experiment 3 provided in the embodiment of the present disclosure;

[0054] Figure 8 This is the residual output of the sub-observer of Experiment 4 provided in the embodiment of the present disclosure. DETAILED DESCRIPTION

[0055] The embodiments of the present disclosure are described in detail below with reference to the accompanying drawings.

[0056] The following describes the embodiments of the present disclosure through specific examples, and those skilled in the art can easily understand other advantages and effects of the present disclosure from the contents disclosed in this specification. Obviously, the described embodiments are only a part of the embodiments of the present disclosure, rather than all of the embodiments. The present disclosure can also be implemented or applied through other different specific embodiments, and the details in this specification can also be modified or changed in various ways based on different viewpoints and applications without departing from the spirit of the present disclosure. It should be noted that the following embodiments and features in the embodiments can be combined with each other without conflict. Based on the embodiments in the present disclosure, all other embodiments obtained by ordinary technicians in the field without making creative work are within the scope of protection of the present disclosure.

[0057] It should be noted that various aspects of the embodiments within the scope of the appended claims are described below. It should be apparent that the aspects described herein may be embodied in a wide variety of forms, and any specific structure and / or function described herein is merely illustrative. Based on the present disclosure, it should be understood by those skilled in the art that an aspect described herein may be implemented independently of any other aspect, and two or more of these aspects may be combined in various ways. For example, any number of aspects described herein may be used to implement the device and / or practice the method. In addition, other structures and / or functionalities other than one or more of the aspects described herein may be used to implement this device and / or practice this method.

[0058] It should also be noted that the illustrations provided in the following embodiments are only schematic illustrations of the basic concept of the present disclosure. The illustrations only show components related to the present disclosure rather than being drawn according to the number, shape and size of components in actual implementation. In actual implementation, the form, quantity and proportion of each component may be changed arbitrarily, and the component layout may also be more complicated.

[0059] Additionally, in the following description, specific details are provided to facilitate a thorough understanding of the examples. However, it will be understood by those skilled in the art that the aspects described may be practiced without these specific details.

[0060] An embodiment of the present disclosure provides a fault isolation method for an active shock absorbing device of an aircraft landing gear, and the method can be applied to a fault isolation process of an aircraft system in a transportation scenario.

[0061] See also Figure 1 , is a flow chart of a fault isolation method for an active shock absorbing device of an aircraft landing gear provided by an embodiment of the present disclosure. Figure 1 As shown, the method mainly comprises the following steps:

[0062] Step 1, modeling the single landing gear, analyzing the vertical dynamic equation of the single landing gear model according to the law of dynamics, and constructing the state space equation of the single landing gear model accordingly;

[0063] In the specific implementation, considering the actual situation of the landing gear's vertical movement and the effectiveness of the research, the following reasonable assumptions are made:

[0064] (1) Only the vertical motion of the fuselage and landing gear is considered, and the influence of horizontal forces is not considered;

[0065] (2) Simplify the spring force and damping force models and express them using linearized equations;

[0066] (3) Ignore the internal friction of the buffer, and its movement range is within the limit and is not subject to structural constraints;

[0067] (4) Only consider the elastic characteristics of the tire and ignore its damping characteristics;

[0068] (5) During the movement, the fuselage of the aircraft does not undergo elastic deformation and is regarded as an ideal rigid body, and the position of the center of mass does not change.

[0069] This establishes Figure 2 The whole aircraft landing gear model is shown. First, for one of the single landing gears, M in the figure represents the fuselage mass, m t It mainly includes the mass of the wheels and tires. According to the law of dynamics, the vertical operation equation of the landing gear can be obtained as follows:

[0070]

[0071] Among them, k s is the landing gear stiffness coefficient, k t is the equivalent stiffness coefficient of the tire, c s is the landing gear damping coefficient, the vertical displacement of the fuselage z s , vertical speed of fuselage Wheel vertical displacement z t , wheel vertical speed is the state variable, that is Control input u d (t) = [u]; interference input q d (t) = [q]. According to equation (1.1), the state space equation of the single landing gear model can be constructed as:

[0072]

[0073] In the formula, the matrices are as follows:

[0074]

[0075] Step 2, combining the coupling of the roll motion, pitch motion and vertical motion of the aircraft fuselage, expanding the state space equation of the single landing gear model to the whole aircraft, and combining the motion equation of each wheel and the vertical motion displacement to obtain the state space equation of the whole aircraft landing gear;

[0076] In the specific implementation, the single landing gear model is extended to the whole aircraft by combining the coupling of the roll motion, pitch motion and vertical motion of the aircraft body. According to the laws of dynamics, and I y and I x denote the pitch and roll moments of inertia of the fuselage, respectively, l f , l h are the distances from the center of mass of the fuselage to the front and rear landing gears along the X-axis, L l , L rare the distances from the center of mass of the fuselage to the left and right landing gears along the Y axis. Let i = 1, 2, 3, representing the front landing gear, rear left landing gear, and rear right landing gear, respectively, l1 = -l f ,l2=l3=l h ,L1=0,L2=-L l ,L3=L r , the equations of the vertical motion, pitch motion, and roll motion of the aircraft fuselage can be simplified as follows:

[0077]

[0078] The forces acting on each landing gear are similar to the vertical motion equation of the fuselage in equation (1.1):

[0079]

[0080] The vertical motion equations of the wheels of each landing gear are the same as equation (1.1), which are:

[0081]

[0082] according to Figure 2 The geometric relationship of the fuselage movement can be used to calculate the vertical movement displacement of each landing gear. and the pitch angle θ is small, so for the convenience of research, we take θ≈sinθ, It can be written as:

[0083]

[0084] where z si Respectively represent the vertical displacement components of the fuselage at each landing gear, k si , k ti 、z ti 、c si Same reason.

[0085] Select control input u z (t) = [u1 u2 u3] T , interference input q z (t) = [q1 q2 q3] T , the state variable is Considering the need for active control design and sensor fault detection, the measurement output The state space equation for constructing the landing gear of the whole aircraft is as follows:

[0086]

[0087] in: C=I 12×12 , Where: A5=

[01] ,A6=[P2l2P2], (i=1,2,3), and the matrix O m×n represents a zero matrix of m rows and n columns, I m×n Represents the identity matrix with m rows and n columns.

[0088] Selecting a Class B pavement and a longitudinal taxiing speed of 50 m / s, the parameter values ​​of the above landing gear model are shown in Table 1.

[0089] Table 1

[0090]

[0091] Step 3, establishing fault models of different fault types for the actuators and sensors of the landing gear;

[0092] In the specific implementation, different types of landing gear actuator failures and sensor failures are modeled separately; the actuator output during failure is taken as u f , the output under normal conditions without faults is u. When one or more actuators of the active control landing gear fail, the landing gear actuator fault model can be established as follows:

[0093]

[0094] Among them, u zf (t) is the actuator fault output, and u zf (t) bounded; η i (i=1,2,3) is the actuator gain fault coefficient, and 0≤η i ≤1; α i (i=1,2,3) is the actuator constant deviation fault coefficient. i , α i When different values ​​are taken, the corresponding fault types of the actuator are also different, as shown in Table 2.

[0095] Table 2

[0096]

[0097] When one or more sensors fail, the landing gear sensor failure model can be established as follows:

[0098]

[0099] y zf (t) bounded; δ j (j=1,2,…,12) is the sensor gain fault coefficient, and 0≤δ j≤1; β j (j=1,2,…,12) is the constant deviation fault coefficient of the sensor. The sensor fault types are shown in Table 3.

[0100] Table 3

[0101]

[0102]

[0103] Let matrix I be the identity matrix of appropriate dimension, f u (t), f y (t) represent the actuator fault signal and sensor fault signal respectively, which can be expressed as:

[0104]

[0105]

[0106] Step 4: According to the fault model of the actuator and the fault model of the sensor, the state space equation of the whole landing gear is rewritten into a state space equation including the fault and the unknown input observer is designed accordingly;

[0107] In specific implementation, based on the above expression, the state space equation of the landing gear containing the fault can be written as:

[0108]

[0109] When there is no actuator failure or sensor failure in the system, f u (t) = 0, f y (t) = 0. Substituting the landing gear model parameter values ​​in Table 1 into equation (1.12), we can obtain rank (CD) = rank (D) = 3.

[0110] The following unknown input observer is designed for the above active control landing gear including fault conditions.

[0111]

[0112] Apply the unknown input observer of equation (1.13) to equation (1.12), and the state estimate obtained by the observer is The state estimation error value is obtained as:

[0113]

[0114] Derivative the above formula gives:

[0115]

[0116] in:

[0117]

[0118] make

[0119] (KC-I)D=0(1.17)

[0120] F=A-KCA-H a C(1.18)

[0121] G=I-KC(1.19)

[0122] Then in formula (1.15) and e z (t) All irrelevant terms can be eliminated, and the state estimation error equation can be written as e z (t) = Fe z (t), when the design matrix F makes its eigenvalues ​​all negative, the error will converge to zero, that is, e z (t)→0,

[0123] Since the matrix D is a column-full rank matrix, according to the knowledge of generalized inverse matrix, there exists a left inverse matrix (CD) + , so that (CD) + CD=I, according to the generalized inverse matrix:

[0124] (CD) + =[(CD) T CD] -1 (CD) T (1.20)

[0125] To satisfy the equation (1.18) that is, KCD = D, (CD) + It can be seen that CD=I, D(CD) + CD=D, so we can solve its special solution K * =D(CD) + =D[(CD) T CD] -1 (CD) T Substituting into equation (1.18), we have:

[0126] F=AD[(CD) T CD] -1 (CD) T CA-H a C(1.21)

[0127] Pick but

[0128]

[0129] According to the observability PBH rank criterion, the necessary and sufficient condition for a linear system to be completely observable is that for all eigenvalues ​​λ in the system matrix A i (i=1,2,…,n), the following relations hold:

[0130]

[0131] It has been proven that the array is observable, then there must exist a matrix H a , all eigenvalues ​​of the matrix F can be configured to be negative, meeting the requirements of the unknown input observer.

[0132] By K * =D(CD) + =D[(CD) T CD] -1 (CD) T The sum of equations (1.20) solves the unknown matrices K and G of the observer. Select a suitable matrix H a According to the relationship (1.18) and by configuring the poles, the eigenvalues ​​of the matrix F1 are all negative, thereby solving the matrix F; finally, according to formula (1.22), the matrix H is solved, thus completing the design of the observer.

[0133] H=H a +H b =H a +FK(1.23)

[0134] According to the state estimation of the observer, define the observer output residual r z (t) is:

[0135]

[0136] Substitute the landing gear parameter values ​​into matrices G and K, and select the appropriate matrix H a , the real part of the eigenvalue of matrix F is calculated as:

[0137] Re[λ F ]=(-44.18-44.18-67-67-16.9-16.9-2-2-0.88-0.88-1.33-1.33).

[0138] Step 5, when the actuator fails, the actuator fault isolation is determined by using the residual root mean square value of the vertical displacement component according to the output residual of the unknown input observer;

[0139] In specific implementation, only the failure of a single actuator is considered each time. When the actuator of a landing gear fails, the impact on the motion state of the landing gear will be greater than that of the other two landing gears. The faulty actuator has the largest residual value r[zsi ](i=1,2,3), so we can compare their RMS values ​​R[z si ](i=1,2,3), determine the actuator that has failed. The specific actuator fault isolation strategy is as follows: Figure 3 shown.

[0140] For each actuator of the active control landing gear shown in formula (1.12), the fault setting in Table 4 is selected for simulation, and the residual root mean square value of the fuselage vertical displacement component at each landing gear and the fault isolation results are shown in Table 5. It can be seen that the residual root mean square value of the fuselage vertical displacement component at the landing gear with fault is significantly greater than that at the landing gear without fault, so the faulty device can be accurately identified and the fault isolation of the actuator can be achieved.

[0141] Table 4

[0142]

[0143]

[0144] Table 5

[0145]

[0146] Step 6: Set n sub-observers. When a sensor fails, the sensor fault isolation is determined based on the output residual values ​​of the n sub-observers.

[0147] In the specific implementation, when designing the unknown input observer, all the measured outputs are introduced into the observer, and any sensor failure will cause the change of the observer output residual. Therefore, a set of parallel sub-observers UIO1,…,UIO n , and the jth observer UIO j Do not introduce the jth state variable y j (j=1,2,…,n), if the measured state y j If the sensor fails, the observer UIO j The residual of the output tends to 0, while the residual of other parallel observers does not tend to 0. The specific strategy is as follows Figure 4 shown.

[0148] For the system shown in equation (1.12), taking the fuselage vertical displacement sensor, fuselage vertical velocity sensor, pitch angular velocity sensor, and roll angle sensor failure as examples, four sub-observers are designed without introducing the fuselage vertical displacement, fuselage vertical velocity, pitch angular velocity, and roll angle, as shown in Table 6. The residual output of the sub-observer is obtained through simulation as follows: Figures 5 to 8 shown.

[0149] Table 6

[0150]

[0151] It can be seen that under different fault settings, the observers whose residual signals tend to 0 can be clearly separated into UIO1, UIO2, UIO3, and UIO4, respectively. From this, it can be judged that the faulty sensors are the fuselage vertical displacement sensor, the fuselage vertical velocity sensor, the pitch angular velocity sensor, and the roll angle sensor, respectively.

[0152] The fault isolation method for the active shock absorbing device of the aircraft landing gear provided in this embodiment expands the single landing gear model to the whole aircraft and considers the coupling relationship of multiple motions, so that the model is more comprehensive and accurate; a fault model for the actuator and the sensor is established, which can effectively identify different types of faults and enhance the reliability of the system; by rewriting the state space equation and designing the unknown input observer, the system can monitor and respond to faults in real time, thereby improving the sensitivity of fault detection; through output residual analysis, the source of the fault can be determined quickly and accurately, reducing the complexity and response time of fault handling; the mechanism of setting multiple sub-observers makes the system have good scalability and can adapt to landing gear systems of different sizes and complexities; by timely identifying and isolating faults, it helps to ensure the safety of the aircraft during take-off and landing and improve the reliability of the overall operation.

[0153] It should be understood that various parts of the present disclosure may be implemented in hardware, software, firmware or a combination thereof.

[0154] The above is only a specific implementation of the present disclosure, but the protection scope of the present disclosure is not limited thereto. Any changes or substitutions that can be easily thought of by a person skilled in the art within the technical scope disclosed in the present disclosure should be included in the protection scope of the present disclosure. Therefore, the protection scope of the present disclosure should be based on the protection scope of the claims.

Claims

1. A fault isolation method for an active shock absorbing device of an aircraft landing gear, characterized in that: include: Step 1, modeling the single landing gear, analyzing the vertical dynamic equation of the single landing gear model according to the law of dynamics, and constructing the state space equation of the single landing gear model accordingly; Step 2, combining the coupling of the roll motion, pitch motion and vertical motion of the aircraft fuselage, expanding the state space equation of the single landing gear model to the whole aircraft, and combining the motion equation of each wheel and the vertical motion displacement to obtain the state space equation of the whole aircraft landing gear; Step 3, establishing fault models of different fault types for the actuators and sensors of the landing gear; Step 4: According to the fault model of the actuator and the fault model of the sensor, the state space equation of the whole landing gear is rewritten into a state space equation including the fault and the unknown input observer is designed accordingly; Step 5, when the actuator fails, the actuator fault isolation is determined by using the residual root mean square value of the vertical displacement component according to the output residual of the unknown input observer; Step 6: Set n sub-observers. When a sensor fails, the sensor fault isolation is determined based on the output residual values ​​of the n sub-observers.

2. The method according to claim 1, characterized in that , the vertical dynamic equation of the single landing gear model is expressed as Where M represents the fuselage mass, m t Mainly includes the quality of wheels and tires, is the vertical acceleration of the fuselage, is the vertical acceleration of the wheel, k s is the landing gear stiffness coefficient, k t is the equivalent stiffness coefficient of the tire, c s is the landing gear damping coefficient, the vertical displacement of the fuselage z s , vertical speed of fuselage Wheel vertical displacement z t , wheel vertical speed is the state variable, that is Control input u d (t) = [u], interference input q d (t) = [q]; The state space equation of the single landing gear model is expressed as Among them, the expressions of matrices A0, B0 and D0 are 3. The method according to claim 2, characterized in that , the state space equation of the whole landing gear is expressed as in, C=I 12×12 , O m×n represents a zero matrix of m rows and n columns, I m×n Represents the identity matrix with m rows and n columns.

4. The method according to claim 3, characterized in that ,The expression of the actuator fault model is Among them, u zf (t) is the actuator fault output, and u zf (t) is bounded, η i , i=1,2,3 is the actuator gain fault coefficient, and 0≤η i ≤1,α i ,i=1,2,3 is the actuator constant deviation fault coefficient; The expression of the fault model of the sensor is Among them, y zf (t) is the sensor fault output, and y zf (t) is bounded, δ j ,j=1,2,…,12 is the sensor gain fault coefficient, and 0≤δ j ≤1, β j ,j=1,2,…,12 is the constant deviation fault coefficient of the sensor.

5. The method according to claim 4, characterized in that , the expression of the state space equation containing the fault is Among them, f u (t), f y (t) represent the actuator fault signal and sensor fault signal respectively; The expression of the unknown input observer is: Among them, w z (t) represents the observer state variable, Is for x z (t) state estimation variable, F∈R n×n , G∈R n×n , H∈R n×l , K∈R n×l is the observer matrix that needs to be solved for the design.

6. The method according to claim 5, characterized in that , the expression of the output residual is Among them, when there is no actuator or sensor failure in the landing gear, the output residual r z (t)→0.

7. The method according to claim 6, characterized in that , the step 5 specifically includes: The fuselage vertical displacement components at each landing gear are selected for analysis. By comparing the residual root mean square values ​​of the fuselage vertical displacement components of each landing gear, the actuator with a residual value greater than that of the other two actuators is determined as the faulty actuator, thereby achieving actuator fault isolation.

8. The method according to claim 7, characterized in that , the step 6 specifically includes: Set up a set of parallel sub-observers UIO1,…,UIO n , and the jth observer UIO j Do not introduce the jth state variable y j , j = 1, 2, ..., n, by observing the observer UIO j Whether the residual of the output tends to 0, while the residual of other parallel observers output does not tend to 0, the fault isolation of the sensor is realized.

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