CN112528492B - Fault detection method and device under wing damage condition - Google Patents
Fault detection method and device under wing damage condition Download PDFInfo
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Abstract
The application provides a fault detection method under the condition of wing damage, which comprises the following steps: according to a kinetic equation of the airplane, an expression of aerodynamic force and aerodynamic moment coefficients of the airplane about measurable state quantity is reversely solved, and a lift coefficient identification model is constructed by combining an aerodynamic moment coefficient model of the airplane; identifying the lift coefficient on line by using a limited memory least square method; and comparing the lift coefficient obtained by identification with wind tunnel test data, judging the fault type position and calculating the fault severity. According to the fault detection method, due to the fact that the online parameter identification fault diagnosis method is adopted, after the fault occurs, the system for detecting the fault can quickly and accurately diagnose the fault, subsequently, the remaining flight capacity can be evaluated based on the diagnosis result, the flight control law can be reconstructed, and the flight safety is improved.
Description
Technical Field
The application belongs to the technical field of airplane fault detection, and particularly relates to a fault detection method and device under the condition of wing damage.
Background
When an airplane breaks down, emergency treatment work such as pilot warning, auxiliary decision, control reconstruction and the like needs to be carried out, and the development of the work depends on accurate fault diagnosis results.
In the prior art, fault diagnosis is mainly performed on redundancy management of a flight management system, and the flight management system can only know whether a certain sensor or actuator has a fault, but cannot know the severity of the fault, so that when faults such as control surface actuation limitation and airfoil damage occur, the flight management system is difficult to perform fault emergency treatment in a targeted manner.
Disclosure of Invention
The present application is directed to a method and apparatus for fault detection in the event of wing damage, to address or mitigate at least one of the problems in the background art.
In one aspect, the present application provides a fault detection method in the event of a wing damage, the fault detection method comprising:
according to a kinetic equation of the airplane, an expression of aerodynamic force and aerodynamic moment coefficients of the airplane about measurable state quantity is reversely solved, and a lift coefficient identification model is constructed by combining an aerodynamic moment coefficient model of the airplane;
identifying the lift coefficient on line by using a limited memory least square method;
and comparing the lift coefficient obtained by identification with wind tunnel test data, judging the fault type position and calculating the fault severity.
Further, the process of constructing the lift coefficient identification model includes:
calculating a projection component Z of the total external force borne by the airplane on an OZ axis of a coordinate system of the airplane according to an airplane mass center kinetic equation:
in the formula, m is the mass of the airplane, and u, v and w are the triaxial velocity components under the shafting of the airplane body;
the resultant external force on the airplane is decomposed into aerodynamic force, engine thrust and gravity, and the aerodynamic force, the engine thrust and the gravity are projected on an OZ axis to obtain
In the formula, T is the thrust of the engine,
for engine mounting angle, L ba-3 Third row of the transformation matrix for the axis of the air flow to the axis of the body, F air The angle is a pitch angle, and phi is a roll angle;
in the above formula, L ba-3 ＝[sin α cos β-sin α sin β cos α]
Aerodynamic force F air Expressed as: f air ＝0.5ρV 2 S ref [-C D C C -C L ] T
In the formula, C D Is a coefficient of resistance, C C Is the coefficient of lateral force, C L Is the lift coefficient;
according to the formula:
the aerodynamic coefficient equation obtained after finishing is as follows:
C L-cf ＝ψ L-cf ε L-cf +v L-cf
in the formula:
ψ L-cf ＝0.5ρV 2 S ref [sin α cos β-sin α sin β cos α]，ε L-cf ＝[-C D ,C C ,-C L ] T ，v L-cf to measure noise;
through multiple sampling of flight state parameters, an nth group of aerodynamic coefficient equations are constructed as follows:
based on the derivation, the lift coefficient identification model constructed is:
L L-cf ＝Ψ L-cf ε L-cf +V L-cf
wherein L is L-cf Is a vector formed by the components of the aircraft aerodynamic force on the axis OZ of the body axis system:
Ψ L-cf for the aircraft state quantity information matrix:
ε L-cf vector formed by aerodynamic coefficients to be identified: epsilon L-cf ＝[-C D ,C C ,-C L ] T
V l-cf For the vector consisting of measurement noise:
further, the process of identifying the lift coefficient on line by using the limited memory least square method comprises the following steps:
constructing an index function according to the principle of least squares:
in the formula (I), the compound is shown in the specification,
is epsilon l-cf An estimated value of (d);
to make the index function J L-cf Minimum, estimated value
Satisfies the following conditions:
in the formula, Ψ L-cf + Is Ψ L-cf The generalized inverse of (1).
Further, the process of comparing the lift coefficient obtained by identification with the wind tunnel test data, judging the fault type position and calculating the fault severity degree comprises the following steps:
determining a mathematical model of a lift coefficient obtained by a wind tunnel test:
when the airfoil of the airplane has no fault, the lift coefficient obtained by identification is approximately equal to the lift coefficient in the wind tunnel test data;
when the damage fault of the single-side airfoil surface occurs, the lift coefficient C is identified L * Will be less than the lift coefficient C obtained by the wind tunnel test L With wing damage rateIncrease of C L And C L * The numerical difference of (a) is also getting larger and larger;
identification error coefficient delta C between lift coefficient obtained by identification and lift coefficient obtained by wind tunnel test L Positive correlation is formed with the wing surface damage rate, and the error coefficient delta C is identified through calculation L Establishing an identification error coefficient Δ C L And the functional relation with the wing surface damage rate can estimate the damage rate of the airplane wing surface through the lift coefficient identification result after the fault occurs.
In another aspect, the present application provides a fault detection device in the event of a wing damage, the fault detection device comprising:
the model construction module is used for reversely solving the expression of aerodynamic force and aerodynamic moment coefficients of the airplane about the measurable state quantity according to the kinetic equation of the airplane and constructing a lift coefficient identification model by combining an aerodynamic moment coefficient model of the airplane;
the coefficient identification module is used for identifying the lift coefficient on line according to a limited memory least square method;
and the fault judgment module is used for comparing the lift coefficient obtained by identification with the wind tunnel test data, judging the fault type position and calculating the fault severity.
Further, the process of constructing the lift coefficient identification model by the model construction module comprises:
calculating a projection component Z of the total external force borne by the airplane on an OZ axis of a coordinate system of the airplane according to an airplane mass center kinetic equation:
in the formula, m is the mass of the airplane, and u, v and w are the three-axis velocity components under the shafting of the airplane body;
the resultant external force on the airplane is decomposed into aerodynamic force, engine thrust and gravity, and the aerodynamic force, the engine thrust and the gravity are projected on an OZ axis to obtain
In the formula, T is the thrust of the engine,
for engine mounting angle, L ba-3 Third row of the transformation matrix for the axis of the air flow to the axis of the body, F air The angle is a pitch angle, and phi is a roll angle;
in the above formula, L ba-3 ＝[sin α cos β-sin α sin β cos α]
Aerodynamic force F air Expressed as: f air ＝0.5ρV 2 S ref [-C D C C -C L ] T
In the formula, C D Is a coefficient of resistance, C C Is the coefficient of lateral force, C L Is the lift coefficient;
according to the formula:
the aerodynamic coefficient equation obtained after finishing is as follows:
C L-cf ＝ψ L-cf ε L-cf +v L-cf
in the formula:
ψ L-cf ＝0.5ρV 2 S ref [sin α cos β-sin α sin β cos α]，ε L-cf ＝[-C D ,C C ,-C L ] T ，v L-cf to measure noise;
through multiple sampling of flight state parameters, an nth group of aerodynamic coefficient equations are constructed as follows:
based on the derivation, the lift coefficient identification model constructed is:
L L-cf ＝Ψ L-cf ε L-cf +V L-cf
wherein L is L-cf Is a vector consisting of the components of the aircraft aerodynamic force on the axis OZ of the body axis:
Ψ L-cf for the aircraft state quantity information matrix:
ε L-cf vector formed by aerodynamic coefficients to be identified: epsilon L-cf ＝[-C D ,C C ,-C L ] T
V l-cf For the vector consisting of measurement noise:
further, the process of identifying the lift coefficient on line by the coefficient identification module by limiting a memory least square method comprises the following steps:
constructing an index function according to the principle of least squares:
in the formula (I), the compound is shown in the specification,
is epsilon l-cf An estimated value of (d);
to make the index function J L-cf Minimum, estimated value
Satisfies the following conditions:
in the formula, Ψ L-cf + Is Ψ L-cf The generalized inverse of (1).
Furthermore, the process of comparing the lift coefficient obtained by the fault judgment module through identification with the wind tunnel test data, judging the fault type position and calculating the fault severity degree comprises the following steps:
determining a mathematical model of a lift coefficient obtained by a wind tunnel test:
when the airfoil of the airplane has no fault, the lift coefficient obtained by identification is approximately equal to the lift coefficient in the wind tunnel test data;
identifying the lift coefficient C obtained after the damage fault of the single-side airfoil surface occurs L * Will be less than the lift coefficient C obtained by the wind tunnel test L With increasing airfoil failure rate, C L And C L * The numerical difference of (a) is also getting larger and larger;
identification error coefficient delta C between lift coefficient obtained by identification and lift coefficient obtained by wind tunnel test L Positive correlation is formed with wing surface damage rate, and error coefficient delta C is identified through calculation L Establishing an identification error coefficient Δ C L And the functional relation with the wing surface damage rate can estimate the damage rate of the airplane wing surface through the lift coefficient identification result after the fault occurs.
According to the fault detection method, due to the fact that the online parameter identification fault diagnosis method is adopted, after the fault occurs, the system for detecting the fault can quickly and accurately diagnose the fault, subsequently, the remaining flight capacity can be evaluated based on the diagnosis result, the flight control law can be reconstructed, and the flight safety is improved.
Drawings
In order to more clearly illustrate the technical solutions provided by the present application, the following briefly introduces the accompanying drawings. It is to be understood that the drawings described below are merely exemplary of some embodiments of the application.
Fig. 1 is a schematic flow chart of a fault detection method according to the present application.
Fig. 2 is a schematic diagram of the fault detection apparatus according to the present application.
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 drawings in the embodiments of the present application.
The method for detecting the faults under the condition of wing damage provided by the application adopts a parameter identification method, key parameters such as the control derivative of a control surface, the lift coefficient of an airplane and the like which can reflect the health condition of the airplane are respectively identified in real time, and the diagnosis of the faults is completed by comparing with airborne wind tunnel test data, namely the type, the position and the degree of the faults are confirmed.
As shown in fig. 1, the method for detecting a fault in the case of a wing damage according to the present application mainly includes three steps:
s1, constructing an identification model of the lift coefficient
Firstly, according to a kinetic equation of the airplane, an expression of aerodynamic force and aerodynamic moment coefficients of the airplane about measurable state quantity is reversely solved, and an identification model of a lift coefficient is obtained by combining an aerodynamic moment coefficient model of the airplane.
Specifically, calculating the OZ axis projection component Z of the total external force borne by the airplane in an airplane coordinate system according to an airplane mass center kinetic equation:
in the formula, m is the mass of the airplane, and u, v and w are the three-axis velocity components under the shafting of the airplane body.
The external force borne by the airplane can be decomposed into aerodynamic force, engine thrust and gravity, and the aerodynamic force, the engine thrust and the gravity are projected on an OZ axis:
in the formula, T is the thrust of the engine,
for engine mounting angle, L ba-3 The third row of the transformation matrix from the axis system of the air flow to the axis system, as shown in formula (3), Fair is aerodynamic force, θ,Phi is a pitch angle and a roll angle.
L ba-3 ＝[sin α cos β-sin α sin β cos α] (3)
Aerodynamic force Fair can be expressed as: f air ＝0.5ρV 2 S ref [-C D C C -C L ] T (4)
In the formula, C D Is a coefficient of resistance, C C Is the lateral force coefficient, C L Is the lift coefficient.
The substitution of formulae (1), (3) and (4) for formula (2) can be obtained:
the aerodynamic coefficient equation obtained by sorting the above formula is as follows:
C L-cf ＝ψ L-cf ε L-cf +v L-cf (5)
wherein:
ψ L-cf ＝0.5ρV 2 S ref [sin α cos β-sin α sin β cos α]
ε L-cf ＝[-C D ,C C ,-C L ] T
v L-cf to measure noise.
Through multiple sampling of flight state parameters, an nth group of aerodynamic coefficient equations are constructed as follows:
based on the above derivation, an identification model of the lift coefficient can be constructed as follows:
L L-cf ＝Ψ L-cf ε L-cf +V L-cf (7)
wherein L is L-cf Is a vector formed by the components of the aircraft aerodynamic force on the axis OZ of the body axis system:
Ψ L-cf for the aircraft state quantity information matrix:
ε L-cf vector formed by the aerodynamic coefficients to be identified: epsilon L-cf ＝[-C D ,C C ,-C L ] T
V l-cf For the vector consisting of measurement noise:
s2, identifying lift coefficient on line
And then, identifying key parameters on line by using a limited memory least square method.
The specific process comprises the following steps:
when each flight state parameter sensor of the aircraft is fault-free, the measurement noise V can be considered L-cf Is a high frequency signal that follows a normal distribution. In this case, the least square method identifies the derived roll steering derivative vector
Is to the actual lift coefficient epsilon of the aircraft L-cf An unbiased estimation of (1).
The term "limited memory" refers to the state quantity information matrix Ψ L-cf And lift coefficient information matrix L L-cf The maximum number of sets of stored data is N lim . When the number of groups of the data stored in the information matrix reaches N lim Every time a new set of data is stored, the oldest set of data is deleted. In this way, the number of data sets in the information matrix involved in the recognition calculation will remain unchanged and be the most up-to-date data. This aspect may utilize historical dataSuppressing measurement noise; on the other hand, the effect of historical data before the fault is eliminated quickly, and therefore a new lift coefficient is identified quickly and accurately.
According to the principle of the least square method, the following index functions are constructed:
in the formula (I), the compound is shown in the specification,
is epsilon l-cf An estimate of (d).
To make J L-cf At minimum, then from knowledge of matrix theory, it can be known that:
in the above formula, Ψ L-cf + Is Ψ L-cf The generalized inverse of (1).
And S3, judging the damage degree of the wing.
And finally, comparing the identified key parameters with wind tunnel test data, judging the fault type position and calculating the fault severity.
The mathematical model of the lift coefficient obtained from the wind tunnel test is:
when the airfoil of the airplane has no fault, the lift coefficient obtained by identification is approximately equal to the lift coefficient in the wind tunnel test data.
When the damage fault of the single-side wing surface occurs, parameters such as wing area, mass, aerodynamic coefficient and the like of the airplane can be changed, wherein the maximum influence on the lift force is the reduction of the wing area, and because the wing area in the identification model is taken as the wing reference area of the example airplane and is a constant, the wing reference area cannot be changed in the identification process, the identified lift force system is obtainedNumber C L * Will be less than the lift coefficient C obtained by the wind tunnel test L Along with the increase of the damage rate of the airfoil, the lift coefficient C obtained by the wind tunnel test L And the lift coefficient C obtained by identification L * The numerical difference of (a) is also getting larger and larger. Lift coefficient C obtained by wind tunnel test L And the lift coefficient C obtained by identification L * Identification error coefficient Δ C of L Positive correlation is formed with the wing surface damage rate, and the error coefficient delta C is identified through calculation L Establishing Δ C L And the functional relation with the wing surface damage rate can estimate the damage rate of the airplane wing surface through the lift coefficient identification result after the fault occurs.
According to the fault detection method under the condition of wing damage, due to the design of online parameter identification fault diagnosis, after a fault occurs, the system of the flying pipe can quickly and accurately diagnose the fault, and then the remaining flying capacity can be evaluated and the flying control law can be reconstructed based on the diagnosis result, so that the flying safety is improved.
In addition, the present application also provides a fault detection device in the case of damage to a wing, as shown in fig. 2, the fault detection device 10 includes:
the model building module 11 is used for reversely solving the expression of aerodynamic force and aerodynamic moment coefficients of the airplane about the measurable state quantity according to the kinetic equation of the airplane and building a lift coefficient identification model by combining the aerodynamic moment coefficient model of the airplane;
the coefficient identification module 12 is used for identifying the lift coefficient on line according to a limited memory least square method;
and the fault judgment module 13 is used for comparing the lift coefficient obtained by identification with wind tunnel test data, judging the fault type position and calculating the fault severity.
Finally, there is also provided in the present application a computer device comprising: one or more processors; storage means for storing one or more programs; the one or more programs, when executed by the one or more processors, cause the one or more processors to implement the fault detection method as in any one of the above.
An embodiment of the present application further provides a computer-readable storage medium, which stores a computer program, and the computer program, when executed by a processor, implements the fault detection method according to any one of the above.
It should be understood that portions of the present application may be implemented in hardware, software, firmware, or a combination thereof. In the above embodiments, various steps, methods, apparatuses or modules may be implemented by software or firmware stored in a memory and executed by a suitable instruction execution system. For example, if implemented in hardware, as in another embodiment, any one or combination of the following techniques, which are known in the art, may be used: a discrete logic circuit having a logic gate circuit for implementing a logic function on a data signal, an application specific integrated circuit having an appropriate combinational logic gate circuit, a Programmable Gate Array (PGA), a Field Programmable Gate Array (FPGA), or the like.
It will be understood by those skilled in the art that all or part of the steps carried out in the method of implementing the above embodiments may be implemented by hardware related to instructions of a program, which may be stored in a computer readable storage medium, and the program, when executed, includes one or a combination of the steps of the method embodiments.
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 (3)
1. A method of fault detection in the event of damage to an airfoil, the method comprising:
according to the kinetic equation of the airplane, the expressions of aerodynamic force and aerodynamic moment coefficients of the airplane about measurable state quantities are reversely solved, and a lift coefficient identification model is constructed by combining an aerodynamic moment coefficient model of the airplane, wherein the process comprises the following steps:
calculating a projection component Z of the total external force borne by the airplane on an OZ axis of a coordinate system of the airplane according to an airplane mass center kinetic equation:
in the formula, m is the mass of the airplane, and u, v and w are the triaxial velocity components under the shafting of the airplane body;
the resultant external force applied to the airplane is decomposed into aerodynamic force, engine thrust and gravity, and the aerodynamic force, the engine thrust and the gravity are projected on an OZ axis to obtain
In the formula, T is the thrust of the engine,
for engine mounting angle, L ba-3 Third row of the transformation matrix for the axis of the gas flow to the axis of the body, F air The angle is a pitch angle, and phi is a roll angle;
in the above formula, L ba-3 ＝[sinαcosβ -sinαsinβ cosα]；
Aerodynamic force F air Expressed as: f air ＝0.5ρV 2 S ref [-C D C C -C L ] T ；
In the formula, C D Is a coefficient of resistance, C C Is the coefficient of lateral force, C L Is the coefficient of lift;
according to the formula:
the aerodynamic coefficient equation obtained after finishing is as follows:
C L-cf ＝ψ L-cf ε L-cf +v L-cf ；
in the formula:
ψ L-cf ＝0.5ρV 2 S ref [sinαcosβ -sinαsinβ cosα]，ε L-cf ＝[-C D ,C C ,-C L ] T ，v L-cf to measure noise;
through multiple sampling of flight state parameters, an nth group of aerodynamic coefficient equations are constructed as follows:
based on the derivation, the lift coefficient identification model constructed is:
L L-cf ＝Ψ L-cf ε L-cf +V L-cf ；
wherein L is L-cf Is a vector consisting of the components of the aircraft aerodynamic force on the axis OZ of the body axis:
Ψ L-cf for the aircraft state quantity information matrix:
ε L-cf vector formed by aerodynamic coefficients to be identified: epsilon L-cf ＝[-C D ,C C ,-C L ] T ；
V l-cf For the vector consisting of measurement noise:
the lift coefficient is identified on line by using a limited memory least square method, and the process comprises the following steps:
constructing an index function according to the principle of least squares:
in the formula (I), the compound is shown in the specification,
is epsilon l-cf An estimated value of (d);
to make the index function J L-cf Minimum, estimated value
Satisfies the following conditions:
in the formula, Ψ L-cf + Is Ψ L-cf The generalized inverse of (1);
and comparing the lift coefficient obtained by identification with wind tunnel test data, judging the fault type position and calculating the fault severity.
2. The method for detecting the failure under the condition of wing damage according to claim 1, wherein the process of comparing the lift coefficient obtained by identification with the wind tunnel test data, judging the failure type position and calculating the failure severity degree comprises the following steps:
determining a mathematical model of a lift coefficient obtained by a wind tunnel test:
when the airfoil of the airplane has no fault, the lift coefficient obtained by identification is approximately equal to the lift coefficient in the wind tunnel test data;
when the damage fault of the single-side airfoil surface occurs, the lift coefficient C is identified L * Will be less than the lift coefficient C obtained by the wind tunnel test L With increasing airfoil failure rate, C L And C L * The numerical difference of (a) is also getting larger and larger;
the lift coefficient obtained by identification and the wind tunnel test are obtainedIdentification error coefficient Delta C between the obtained lift coefficients L Positive correlation is formed with the wing surface damage rate, and the error coefficient delta C is identified through calculation L Establishing an identification error coefficient Δ C L And the functional relation with the wing surface damage rate can estimate the damage rate of the airplane wing surface through the lift coefficient identification result after the fault occurs.
3. A fault detection device in the event of a wing damage, the fault detection method comprising:
the model building module is used for reversely solving the expression of aerodynamic force and aerodynamic moment coefficients of the airplane about the measurable state quantity according to the kinetic equation of the airplane, and building a lift coefficient identification model by combining the aerodynamic moment coefficient model of the airplane, and the process comprises the following steps:
calculating a projection component Z of the total external force borne by the airplane on an OZ axis of a coordinate system of the airplane according to an airplane mass center kinetic equation:
in the formula, m is the mass of the airplane, and u, v and w are the triaxial velocity components under the shafting of the airplane body;
the resultant external force on the airplane is decomposed into aerodynamic force, engine thrust and gravity, and the aerodynamic force, the engine thrust and the gravity are projected on an OZ axis to obtain
In the formula, T is the thrust of the engine,
for engine mounting angles, L ba-3 Third row of the transformation matrix for the axis of the air flow to the axis of the body, F air The angle is a pitch angle, and phi is a roll angle;
in the above formula, L ba-3 ＝[sinαcosβ -sinαsinβ cosα]；
Aerodynamic force F air Expressed as: f air ＝0.5ρV 2 S ref [-C D C C -C L ] T ；
In the formula, C D Is a coefficient of resistance, C C Is the coefficient of lateral force, C L Is the coefficient of lift;
according to the formula:
the aerodynamic coefficient equation obtained after finishing is as follows:
C L-cf ＝ψ L-cf ε L-cf +v L-cf ；
in the formula:
ψ L-cf ＝0.5ρV 2 S ref [sinαcosβ -sinαsinβ cosα]，ε L-cf ＝[-C D ,C C ,-C L ] T ，v L-cf to measure noise;
through multiple sampling of flight state parameters, an nth group of aerodynamic coefficient equations are constructed as follows:
based on the derivation, the constructed lift coefficient identification model is as follows:
L L-cf ＝Ψ L-cf ε L-cf +V L-cf ；
wherein L is L-cf Is a vector consisting of the components of the aircraft aerodynamic force on the axis OZ of the body axis:
Ψ L-cf for the aircraft state quantity information matrix:
ε L-cf vector formed by the aerodynamic coefficients to be identified: epsilon L-cf ＝[-C D ,C C ,-C L ] T ；
V l-cf For vectors consisting of measurement noise:
the coefficient identification module is used for identifying the lift coefficient on line according to a limited memory least square method, and the process comprises the following steps:
constructing an index function according to the principle of least squares:
in the formula (I), the compound is shown in the specification,
is epsilon l-cf An estimated value of (d);
to make the index function J L-cf Minimum, estimated value
Satisfies the following conditions:
in the formula, psi L-cf + Is Ψ L-cf The generalized inverse of (1).
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