Aircraft longitudinal maneuver load simulation method and device

Abstract

The application belongs to the technical field of aircraft simulation, and particularly relates to a kind of aircraft longitudinal maneuver load simulation method and device.The method comprises the following steps: step S1, aircraft 1g level flight trimming is carried out, and the aircraft trimming angle of attack and the aircraft trimming elevator deflection are determined;Step S2, the aerodynamic force of the aircraft is determined;Step S3, the dynamic response of the aerodynamic force of the aircraft is solved, and the aerodynamic angle of attack, the normal overload, the flight speed and the pitch angular velocity are obtained;Step S4, the first elevator deflection calculated by the stability augmentation system according to the aerodynamic angle of attack and the normal overload is obtained, the second elevator deflection given by the elevator command is superimposed, and the actual required elevator deflection is obtained;Step S5, according to the flight speed, the pitch angular velocity, the aerodynamic angle of attack and the actual required elevator deflection, the aerodynamic force of the aircraft is determined again until the simulation is finished.The application considers the effect of the stability augmentation system in the process of determining the longitudinal maneuver load of the aircraft, and the simulation process is simple and accurate.

Description

Technical Field

[0001] The present application belongs to the field of aircraft simulation technology, and in particular relates to a method and device for simulating longitudinal maneuvering loads of an aircraft. Background Art

[0002] Traditional aircraft often have sufficient longitudinal static stability margin, and the longitudinal maneuvering load can achieve the target normal overload through elevator trapezoidal / delta control (or command control), and generally there is no need to consider the impact of the stability augmentation system on the flight load.

[0003] Modern aircraft employ designs with relaxed static stability, and their safe flight relies even more heavily on stability augmentation systems (STASs) to achieve superior flight quality and safety. The basic principle of a longitudinal STA is to use sensors to measure signals such as the aircraft's angular rate and g-force. The flight control computer then calculates elevator motion commands according to a predetermined control law and drives the elevator deflection to generate a reasonable and sufficient aerodynamic torque, providing additional motion damping and stability for the aircraft.

[0004] The stability augmentation system will change the original elevator trapezoidal / delta control, which in turn affects the longitudinal maneuvering load design results. Therefore, it is necessary to establish a method for determining the aircraft longitudinal maneuvering load that takes into account the effect of the stability augmentation system to meet the needs of longitudinal maneuvering load simulation. Summary of the Invention

[0005] In order to solve the above problems, the first aspect of the present application provides a method for simulating longitudinal maneuvering loads of an aircraft, which mainly includes:

[0006] Step S1, performing level flight trim of the aircraft 1g, determining the trim angle of attack and the trim elevator deflection of the aircraft, which are used as the initial aerodynamic angle of attack and elevator deflection of the simulation;

[0007] Step S2: determining the aerodynamic force of the aircraft;

[0008] Step S3, solving the dynamic response of the aerodynamic force of the aircraft to obtain the aerodynamic angle of attack, normal overload, flight speed and pitch angular velocity;

[0009] Step S4: obtaining a first elevator deflection calculated by the stability augmentation system based on the aerodynamic angle of attack and the normal overload, and superimposing the second elevator deflection given by the elevator command to obtain the actually required elevator deflection;

[0010] Step S5: Based on the flight speed, pitch angular velocity, aerodynamic angle of attack, and the actual required elevator deflection, return to step S2, re-determine the aerodynamic force of the aircraft, and repeat the above steps until the simulation ends.

[0011] Preferably, step S1 further comprises:

[0012] S11, obtaining input aircraft flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristics data set;

[0013] S12, calculating the atmospheric density ρ at the flight altitude H, and further calculating the velocity pressure Q of the aircraft;

[0014] S13. Perform 1g trim on the aircraft based on the above parameters to obtain the trim angle of attack and the trim elevator deflection of the aircraft.

[0015] Preferably, in step S12, the speed pressure Q of the aircraft is calculated by the following formula:

[0016] Q=0.5ρV.

[0017] Preferably, step S2 further comprises:

[0018] Step S21: interpolate the aircraft aerodynamic characteristic data set to obtain the aircraft drag coefficient C under the current aerodynamic angle of attack and elevator deflection conditions. D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq ;

[0019] Step S22: Based on the drag coefficient C of the aircraft D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L, and pitching moment m of the aircraft y .

[0020] Preferably, in step S21, the drag coefficient C of the aircraft is interpolated by the following formula: D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq :

[0021]

[0022] The drag coefficient C of the aircraft is interpolated according to the flight Mach number Ma, the aerodynamic angle of attack α and the elevator deflection δe D , lift coefficient C L , pitching moment coefficient C my , according to the flight Mach number Ma, the aircraft trim angle of attack α interpolation pitch damping coefficient C mq .

[0023] In step S22, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated according to the following formulas: y :

[0024]

[0025] Where Q is the velocity pressure of the aircraft, S is the reference area of ​​the aircraft, q is the pitch angular velocity of the aircraft, V is the flight speed of the aircraft, and c is the average aerodynamic chord length.

[0026] Preferably, step S3 further comprises:

[0027] Step S31: Use the fourth-order Runge-Kutta method to solve the dynamic response of the aircraft to obtain the projection value of the aircraft velocity on the X-axis and Z-axis of the body axis system, the pitch angle and the pitch angular velocity;

[0028] Step S32: Determine the new aerodynamic angle of attack α and normal overload n z And flight Mach number Ma.

[0029] Preferably, in step S31, the dynamic response is solved by the following formula:

[0030]

[0031] Where m is the mass of the aircraft; g is the acceleration of gravity; u and w are the projection values ​​of the aircraft velocity on the X-axis and Z-axis of the body axis system respectively; θ is the pitch angle of the aircraft; I yy is the moment of inertia of the aircraft relative to the Y axis of the body axis system;

[0032] F xf is the projection value of aerodynamic force on the X axis of the body axis system and the projection value of engine thrust on the X axis of the body axis system T The sum of zf is the projection value of aerodynamic force on the Z axis of the body axis system and the projection value of engine thrust on the Z axis of the body axis system T The sum of M yf is the moment m of the aerodynamic force on the Y axis of the body axis system y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows:

[0033]

[0034] In step S32, the new aerodynamic angle of attack α and normal overload n are determined by the following formulas: z And the flight Mach number Ma:

[0035]

[0036] The second aspect of the present application provides an aircraft longitudinal maneuvering load simulation device, which mainly includes:

[0037] The parameter initialization module is used to perform 1g level flight trim of the aircraft, determine the trim angle of attack and the trim elevator deflection of the aircraft, and use them as the initial aerodynamic angle of attack and elevator deflection of the simulation;

[0038] an aerodynamic calculation module, used to determine the aerodynamic forces of the aircraft;

[0039] The dynamic response solution module is used to solve the dynamic response of the aerodynamic force of the aircraft and obtain the aerodynamic angle of attack, normal overload, flight speed and pitch angular velocity;

[0040] a stabilization calculation module for obtaining a first elevator deflection calculated by the stabilization system based on the aerodynamic angle of attack and normal overload, and superimposing the second elevator deflection given by the elevator command to obtain the actual required elevator deflection;

[0041] The parameter updating module is used to re-determine the aerodynamic force of the aircraft according to the flight speed, pitch angular velocity, aerodynamic angle of attack and the actual required elevator deflection.

[0042] Preferably, the parameter initialization module includes:

[0043] A flight parameter acquisition unit is used to obtain the input aircraft flight speed V, flight altitude H, aircraft mass m and aircraft aerodynamic characteristics data set;

[0044] The speed and pressure calculation unit is used to calculate the atmospheric density ρ at the flight altitude H, and further calculate the speed and pressure Q of the aircraft;

[0045] The trim unit is used to perform 1g trim on the aircraft based on the above parameters to obtain the trim angle of attack and the trim elevator deflection of the aircraft.

[0046] Preferably, in the speed-pressure calculation unit, the speed-pressure Q of the aircraft is calculated by the following formula:

[0047] Q=0.5ρV.

[0048] Preferably, the aerodynamic force calculation module includes:

[0049] The interpolation unit is used to interpolate the aircraft's aerodynamic characteristics data set to obtain the aircraft's drag coefficient C under the current aerodynamic angle of attack and elevator deflection conditions. D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq ;

[0050] The aerodynamic parameter calculation unit is used to calculate the drag coefficient C of the aircraft. D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient Cmq Determine the aerodynamic drag D, lift L, and pitching moment m of the aircraft y .

[0051] Preferably, in the interpolation unit, the drag coefficient C of the aircraft is interpolated by the following formula: D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq :

[0052]

[0053] The drag coefficient C of the aircraft is interpolated according to the flight Mach number Ma, the aerodynamic angle of attack α and the elevator deflection δe D , lift coefficient C L , pitching moment coefficient C my , according to the flight Mach number Ma, the aircraft trim angle of attack α interpolation pitch damping coefficient C mq .

[0054] In the aerodynamic parameter calculation unit, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated according to the following formulas: y :

[0055]

[0056] Where Q is the velocity pressure of the aircraft, S is the reference area of ​​the aircraft, q is the pitch angular velocity of the aircraft, V is the flight speed of the aircraft, and c is the average aerodynamic chord length.

[0057] Preferably, the dynamic response solving module includes:

[0058] The dynamic response solving unit is used to solve the dynamic response of the aircraft using the fourth-order Runge-Kutta method to obtain the projection value of the aircraft velocity on the X-axis and Z-axis of the body axis system, the pitch angle and the pitch angle velocity;

[0059] Longitudinal maneuvering load calculation unit, used to determine the new aerodynamic angle of attack α, normal overload n z And flight Mach number Ma.

[0060] Preferably, in the dynamic response solving unit, the dynamic response is solved by the following formula:

[0061]

[0062] Where m is the mass of the aircraft; g is the acceleration of gravity; u and w are the projection values ​​of the aircraft velocity on the X-axis and Z-axis of the body axis system respectively; θ is the pitch angle of the aircraft; I yy is the moment of inertia of the aircraft relative to the Y axis of the body axis system;

[0063] F xf is the projection value of aerodynamic force on the X axis of the body axis system and the projection value of engine thrust on the X axis of the body axis system T The sum of zf is the projection value of aerodynamic force on the Z axis of the body axis system and the projection value of engine thrust on the Z axis of the body axis system T The sum of M yf is the moment m of the aerodynamic force on the Y axis of the body axis system y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows:

[0064]

[0065] In the longitudinal maneuver load calculation unit, the new aerodynamic angle of attack α, normal overload n are determined by the following formulas: z And the flight Mach number Ma:

[0066]

[0067] A third aspect of the present application provides a computer device comprising a processor, a memory, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the aircraft longitudinal maneuvering load simulation method as described above.

[0068] This application reasonably considers the role of the stability augmentation system in determining the longitudinal maneuvering load of the aircraft. The simulation process is simple and accurate, meeting the needs of engineering practice. BRIEF DESCRIPTION OF THE DRAWINGS

[0069] Figure 1 This is a flow chart of a preferred embodiment of the aircraft longitudinal maneuvering load simulation method of the present application.

[0070] Figure 2 This application Figure 1 A time history of aircraft elevator deflection for the illustrated embodiment.

[0071] Figure 3 This application Figure 1 Time history of normal G-load at the center of mass of the vehicle for the illustrated embodiment.

[0072] Figure 4 This application Figure 1 A time history diagram of the aircraft pitch acceleration of the illustrated embodiment. DETAILED DESCRIPTION

[0073] In order to make the purpose, technical solutions and advantages of the implementation of this application clearer, the technical solutions in the implementation of this application will be described in more detail below in conjunction with the drawings in the implementation of this application. In the drawings, the same or similar numbers throughout represent the same or similar elements or elements with the same or similar functions. The described implementation is a part of the implementation of this application, not all of the implementations. The implementation described below with reference to the drawings is exemplary and is intended to be used to explain this application, and should not be understood as a limitation on this application. Based on the implementation in this application, all other implementations obtained by ordinary technicians in this field without making creative work are within the scope of protection of this application. The implementation of this application is described in detail below in conjunction with the drawings.

[0074] The first aspect of the present application provides a method for simulating longitudinal maneuvering loads of an aircraft, such as Figure 1 As shown, it mainly includes:

[0075] Step S1: Perform level flight trim on the aircraft 1g to determine the trim angle of attack and the trim elevator deflection of the aircraft, which are used as the initial aerodynamic angle of attack and elevator deflection of the simulation.

[0076] This step is mainly used to obtain the initial aerodynamic angle of attack and elevator deflection through 1g equalization in order to carry out the simulation cycle. In addition, refer to Figure 1 ,Before the simulation begins, it is also necessary to obtain ,parameters such as the aircraft speed, altitude, mass, and aerodynamic ,characteristics data set for use in simulation ,calculations.

[0077] In some optional embodiments, step S1 further includes:

[0078] S11, obtaining input aircraft flight speed V, flight altitude H, aircraft mass m and aircraft aerodynamic characteristics data set;

[0079] S12, calculate the atmospheric density ρ at the flight altitude H, and further calculate the speed pressure Q of the aircraft;

[0080] S13, performing 1g trim on the aircraft based on the above parameters to obtain the trim angle of attack and the trim elevator deflection of the aircraft.

[0081] In some optional implementations, in step S12, the speed pressure Q of the aircraft is calculated using the following formula:

[0082] Q=0.5ρV.

[0083] In this embodiment, step S13 is a 1g trim step, which requires the use of the aircraft flight speed V, flight altitude H, aircraft mass m, aircraft aerodynamic characteristics data set and other parameters given in step S11. It also requires the use of the speed pressure Q parameter of step S12. The speed pressure Q parameter of step S12 needs to be calculated based on the atmospheric density ρ at the flight altitude H. In addition, in step S12, the speed of sound V at the flight altitude H also needs to be determined. S , in order to calculate the Mach number Ma for a given aircraft flight speed V.

[0084] After the calculation in step S13, the aircraft trim angle of attack α can be obtained. trim and trim elevator deflection δe trim , and use it as the aerodynamic angle of attack α and elevator deflection δe to participate in the subsequent simulation cycle.

[0085] Step S2: Determine the aerodynamic force of the aircraft.

[0086] With the aerodynamic angle of attack α and the elevator deflection δe, the aerodynamic force can be calculated.

[0087] In some optional embodiments, step S2 further includes:

[0088] Step S21: interpolate the aircraft aerodynamic characteristic data set to obtain the aircraft drag coefficient C under the current aerodynamic angle of attack and elevator deflection conditions. D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq ;

[0089] Step S22: Based on the drag coefficient C of the aircraft D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L, and pitching moment m of the aircraft y .

[0090] In some optional embodiments, in step S21, the drag coefficient C of the aircraft is interpolated by the following formula: D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq :

[0091]

[0092] The drag coefficient C of the aircraft is interpolated according to the flight Mach number Ma, the aerodynamic angle of attack α and the elevator deflection δe D , lift coefficient C L, pitching moment coefficient C my , according to the flight Mach number Ma, the aircraft trim angle of attack α interpolation pitch damping coefficient C mq .

[0093] In step S22, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated according to the following formulas: y :

[0094]

[0095] Where Q is the velocity pressure of the aircraft, S is the reference area of ​​the aircraft, q is the pitch angular velocity of the aircraft, V is the flight speed of the aircraft, and c is the average aerodynamic chord length.

[0096] In the above embodiment, it can be seen from the formula that the aerodynamic coefficient can be interpolated based on the current flight Mach number Ma, aerodynamic angle of attack α, and elevator control deflection δe of the aircraft, and then the aerodynamic force can be further obtained based on the aircraft parameters.

[0097] Step S3: solving the dynamic response of the aerodynamic force of the aircraft to obtain the aerodynamic angle of attack, normal overload, flight speed and pitch angular velocity.

[0098] This step is mainly used to update the aerodynamic angle of attack. At the same time, the calculated normal overload will be used to update the elevator deflection so that step S2 can be used cyclically.

[0099] In some optional embodiments, step S3 further includes:

[0100] Step S31: Use the fourth-order Runge-Kutta method to solve the dynamic response of the aircraft to obtain the projection value of the aircraft velocity on the X-axis and Z-axis of the body axis system, the pitch angle and the pitch angular velocity;

[0101] Step S32: Determine the new aerodynamic angle of attack α and normal overload n z And flight Mach number Ma.

[0102] In some optional implementations, in step S31, the dynamic response is solved using the following formula:

[0103]

[0104] Where m is the mass of the aircraft; g is the acceleration of gravity; u and w are the projection values ​​of the aircraft velocity on the X-axis and Z-axis of the body axis system respectively; θ is the pitch angle of the aircraft; I yy is the moment of inertia of the aircraft relative to the Y axis of the body axis system;

[0105] F xf is the projection value of aerodynamic force on the X axis of the body axis system and the projection value of engine thrust on the X axis of the body axis system TThe sum of zf is the projection value of aerodynamic force on the Z axis of the body axis system and the projection value of engine thrust on the Z axis of the body axis system T The sum of M yf is the moment m of the aerodynamic force on the Y axis of the body axis system y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows:

[0106]

[0107] In step S32, the new aerodynamic angle of attack α and normal overload n are determined by the following formulas: z And the flight Mach number Ma:

[0108]

[0109] In step S31, the aerodynamic angle of attack α at the current simulation moment is used to solve the dynamic response. Then, in step S32, the aerodynamic angle of attack α is updated according to the dynamic response solution result. The atan in the formula of step S32 is the inverse tangent function.

[0110] Step S4: obtaining the first elevator deflection calculated by the stability augmentation system based on the aerodynamic angle of attack and the normal overload, and superimposing the second elevator deflection given by the elevator command to obtain the actually required elevator deflection.

[0111] refer to Figure 1 At the new simulation moment, the new aerodynamic angle of attack α and normal overload n calculated in step S32 are z By passing the aircraft's stability augmentation system (STAR), we can obtain the first elevator deflection δe2 required by the STAR. We then obtain the current second elevator deflection δe1 from the elevator command input. Finally, by adding these two elevator deflections, we obtain the actual required elevator deflection δe.

[0112] Step S5: Based on the flight speed, pitch angular velocity, aerodynamic angle of attack, and the actual required elevator deflection, return to step S2, re-determine the aerodynamic force of the aircraft, and repeat the above steps until the simulation ends.

[0113] refer to Figure 1 In addition to the aerodynamic angle of attack α updated in step S32 and the elevator deflection δe updated in step S4, the parameters involved in the loop also include the Mach number Ma updated in step S32 and the pitch angular velocity q updated in step S31.

[0114] Figure 2 This is a time history diagram of the aircraft's elevator deflection. The horizontal axis is time and the vertical axis is elevator deflection. After considering the role of the stability augmentation system, the elevator is no longer a traditional triangular / trapezoidal control, but a more complex control deflection curve. Figure 3 This is a time history diagram of the normal overload at the center of mass of the aircraft. The horizontal axis is time and the vertical axis is the normal overload at the center of mass of the aircraft. After considering the effect of the stability augmentation system, it takes longer for the normal overload to reach the target value. Figure 4 This is a time history diagram of the aircraft's pitch angle acceleration. The horizontal axis is time, and the vertical axis is the aircraft's pitch angle acceleration. After considering the effect of the stabilization system, the pitch angle acceleration amplitude decreases.

[0115] This application reasonably considers the role of the stability augmentation system in determining the longitudinal maneuvering load of the aircraft. The simulation process is simple and accurate, meeting the needs of engineering practice.

[0116] The second aspect of the present application provides an aircraft longitudinal maneuvering load simulation device corresponding to the above method, mainly comprising:

[0117] The parameter initialization module is used to perform 1g level flight trim of the aircraft, determine the trim angle of attack and the trim elevator deflection of the aircraft, and use them as the initial aerodynamic angle of attack and elevator deflection of the simulation;

[0118] an aerodynamic calculation module, used to determine the aerodynamic forces of the aircraft;

[0119] The dynamic response solution module is used to solve the dynamic response of the aerodynamic force of the aircraft and obtain the aerodynamic angle of attack, normal overload, flight speed and pitch angular velocity;

[0120] a stabilization calculation module for obtaining a first elevator deflection calculated by the stabilization system based on the aerodynamic angle of attack and normal overload, and superimposing the second elevator deflection given by the elevator command to obtain the actual required elevator deflection;

[0121] The parameter updating module is used to re-determine the aerodynamic force of the aircraft according to the flight speed, pitch angular velocity, aerodynamic angle of attack and the actual required elevator deflection.

[0122] In some optional implementations, the parameter initialization module includes:

[0123] A flight parameter acquisition unit is used to obtain the input aircraft flight speed V, flight altitude H, aircraft mass m and aircraft aerodynamic characteristics data set;

[0124] The speed and pressure calculation unit is used to calculate the atmospheric density ρ at the flight altitude H, and further calculate the speed and pressure Q of the aircraft;

[0125] The trim unit is used to perform 1g trim on the aircraft based on the above parameters to obtain the trim angle of attack and the trim elevator deflection of the aircraft.

[0126] In some optional implementations, in the speed-pressure calculation unit, the speed-pressure Q of the aircraft is calculated using the following formula:

[0127] Q=0.5ρV.

[0128] In some optional embodiments, the aerodynamic force calculation module includes:

[0129] The interpolation unit is used to interpolate the aircraft's aerodynamic characteristics data set to obtain the aircraft's drag coefficient C under the current aerodynamic angle of attack and elevator deflection conditions. D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq ;

[0130] The aerodynamic parameter calculation unit is used to calculate the drag coefficient C of the aircraft. D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L, and pitching moment m of the aircraft y .

[0131] In some optional embodiments, in the interpolation unit, the drag coefficient C of the aircraft is interpolated by the following formula: D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq :

[0132]

[0133] The drag coefficient C of the aircraft is interpolated according to the flight Mach number Ma, the aerodynamic angle of attack α and the elevator deflection δe D , lift coefficient C L , pitching moment coefficient C my , according to the flight Mach number Ma, the aircraft trim angle of attack α interpolation pitch damping coefficient C mq .

[0134] In the aerodynamic parameter calculation unit, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated according to the following formulas: y :

[0135]

[0136] Where Q is the velocity pressure of the aircraft, S is the reference area of ​​the aircraft, q is the pitch angular velocity of the aircraft, V is the flight speed of the aircraft, and c is the average aerodynamic chord length.

[0137] In some optional implementations, the dynamic response solving module includes:

[0138] The dynamic response solving unit is used to solve the dynamic response of the aircraft using the fourth-order Runge-Kutta method to obtain the projection value of the aircraft velocity on the X-axis and Z-axis of the body axis system, the pitch angle and the pitch angle velocity;

[0139] Longitudinal maneuvering load calculation unit, used to determine the new aerodynamic angle of attack α, normal overload n z And flight Mach number Ma.

[0140] In some optional implementations, in the dynamic response solving unit, the dynamic response is solved by the following formula:

[0141]

[0142] Where m is the mass of the aircraft; g is the acceleration of gravity; u and w are the projection values ​​of the aircraft velocity on the X-axis and Z-axis of the body axis system respectively; θ is the pitch angle of the aircraft; I yy is the moment of inertia of the aircraft relative to the Y axis of the body axis system;

[0143] F xf is the projection value of aerodynamic force on the X axis of the body axis system and the projection value of engine thrust on the X axis of the body axis system T The sum of zf is the projection value of aerodynamic force on the Z axis of the body axis system and the projection value of engine thrust on the Z axis of the body axis system T The sum of M yf is the moment m of the aerodynamic force on the Y axis of the body axis system y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows:

[0144]

[0145] In the longitudinal maneuver load calculation unit, the new aerodynamic angle of attack α, normal overload n are determined by the following formulas: z And the flight Mach number Ma:

[0146]

[0147] A third aspect of the present application provides a computer device comprising a processor, a memory, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the aircraft longitudinal maneuvering load simulation method as described above.

[0148] The above description is merely a specific embodiment of the present application, but the scope of protection of the present application is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in this application should be included in the scope of protection of the present application. Therefore, the scope of protection of the present application should be based on the scope of protection of the claims.

Claims

1. A method for simulating longitudinal maneuvering loads of an aircraft, characterized in that: include: Step S1, performing level flight trim of the aircraft 1g, determining the trim angle of attack and the trim elevator deflection of the aircraft, which are used as the initial aerodynamic angle of attack and elevator deflection of the simulation; Step S2: determining the aerodynamic force of the aircraft; Step S3, solving the dynamic response of the aerodynamic force of the aircraft to obtain the aerodynamic angle of attack, normal overload, flight speed and pitch angular velocity; Step S4: obtaining a first elevator deflection calculated by the stability augmentation system based on the aerodynamic angle of attack and the normal overload, and superimposing the second elevator deflection given by the elevator command to obtain the actually required elevator deflection; Step S5: Based on the flight speed, pitch rate, aerodynamic angle of attack, and actual required elevator deflection, return to step S2 to re-determine the aerodynamic force of the aircraft and repeat the above steps until the simulation ends. Among them, S3 further includes: Step S31: Use the fourth-order Runge-Kutta method to solve the dynamic response of the aircraft to obtain the projection value of the aircraft velocity on the X-axis and Z-axis of the body axis system, the pitch angle and the pitch angular velocity; Step S32: Determine the new aerodynamic angle of attack α and normal overload n z And the flight Mach number Ma; In step S31, the dynamic response is solved using the following formula: Where m is the mass of the aircraft; g is the acceleration of gravity; u and w are the projection values ​​of the aircraft velocity on the X-axis and Z-axis of the body axis system respectively; θ is the pitch angle of the aircraft; I yy is the moment of inertia of the aircraft relative to the Y axis of the body axis system; F xf is the projection value of aerodynamic force on the X axis of the body axis system and the projection value of engine thrust on the X axis of the body axis system T The sum of zf is the projection value of aerodynamic force on the Z axis of the body axis system and the projection value of engine thrust on the Z axis of the body axis system T The sum of M yf is the moment m of the aerodynamic force on the Y axis of the body axis system y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows: In step S32, the new aerodynamic angle of attack α and normal overload n are determined by the following formulas: z And the flight Mach number Ma:

2. The method for simulating aircraft longitudinal maneuvering loads according to claim 1, wherein: Step S1 further comprises: S11, obtaining input aircraft flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristics data set; S12, calculating the atmospheric density ρ at the flight altitude H, and further calculating the velocity pressure Q of the aircraft; S13. Perform 1g trim on the aircraft based on the above parameters to obtain the trim angle of attack and the trim elevator deflection of the aircraft.

3. The method for simulating aircraft longitudinal maneuvering loads according to claim 2, wherein: In step S12, the speed pressure Q of the aircraft is calculated using the following formula: Q=0.5ρV.

4. The method for simulating aircraft longitudinal maneuvering loads according to claim 1, wherein: Step S2 further comprises: Step S21: interpolate the aircraft aerodynamic characteristic data set to obtain the aircraft drag coefficient C under the current aerodynamic angle of attack and elevator deflection conditions. D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq ; Step S22: Based on the drag coefficient C of the aircraft D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L, and pitching moment m of the aircraft y .

5. The method for simulating aircraft longitudinal maneuvering loads according to claim 4, wherein: In step S21, the drag coefficient C of the aircraft is interpolated by the following formula: D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq : The drag coefficient C of the aircraft is interpolated according to the flight Mach number Ma, the aerodynamic angle of attack α and the elevator deflection δe D , lift coefficient C L , pitching moment coefficient C my , according to the flight Mach number Ma, the aircraft trim angle of attack α interpolation pitch damping coefficient C mq ; In step S22, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated according to the following formulas: y : Where Q is the velocity pressure of the aircraft, S is the reference area of ​​the aircraft, q is the pitch angular velocity of the aircraft, V is the flight speed of the aircraft, and c is the average aerodynamic chord length.

6. An aircraft longitudinal maneuvering load simulation device, characterized in that: include: The parameter initialization module is used to perform 1g level flight trim of the aircraft, determine the trim angle of attack and the trim elevator deflection of the aircraft, and use them as the initial aerodynamic angle of attack and elevator deflection of the simulation; an aerodynamic calculation module, used to determine the aerodynamic forces of the aircraft; The dynamic response solution module is used to solve the dynamic response of the aerodynamic force of the aircraft and obtain the aerodynamic angle of attack, normal overload, flight speed and pitch angular velocity; a stabilization calculation module for obtaining a first elevator deflection calculated by the stabilization system based on the aerodynamic angle of attack and normal overload, and superimposing the second elevator deflection given by the elevator command to obtain the actual required elevator deflection; A parameter update module is used to re-determine the aerodynamic force of the aircraft based on the flight speed, pitch angular velocity, aerodynamic angle of attack and the actual required elevator deflection; Wherein, the dynamic response solution module includes: The dynamic response solving unit is used to solve the dynamic response of the aircraft using the fourth-order Runge-Kutta method to obtain the projection value of the aircraft velocity on the X-axis and Z-axis of the body axis system, the pitch angle and the pitch angle velocity; Longitudinal maneuvering load calculation unit, used to determine the new aerodynamic angle of attack α, normal overload n z And the flight Mach number Ma; In the dynamic response solving unit, the dynamic response is solved by the following formula: Where m is the mass of the aircraft; g is the acceleration of gravity; u and w are the projection values ​​of the aircraft velocity on the X-axis and Z-axis of the body axis system respectively; θ is the pitch angle of the aircraft; I yy is the moment of inertia of the aircraft relative to the Y axis of the body axis system; F xf is the projection value of aerodynamic force on the X axis of the body axis system and the projection value of engine thrust on the X axis of the body axis system T The sum of zf is the projection value of aerodynamic force on the Z axis of the body axis system and the projection value of engine thrust on the Z axis of the body axis system T The sum of M yf is the moment m of the aerodynamic force on the Y axis of the body axis system y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows: In the longitudinal maneuver load calculation unit, the new aerodynamic angle of attack α, normal overload n are determined by the following formulas: z And the flight Mach number Ma:

7. The aircraft longitudinal maneuver load simulation device according to claim 6, characterized in that: The parameter initialization module includes: A flight parameter acquisition unit is used to obtain the input aircraft flight speed V, flight altitude H, aircraft mass m and aircraft aerodynamic characteristics data set; The speed and pressure calculation unit is used to calculate the atmospheric density ρ at the flight altitude H, and further calculate the speed and pressure Q of the aircraft; The trim unit is used to perform 1g trim on the aircraft based on the above parameters to obtain the trim angle of attack and the trim elevator deflection of the aircraft.

8. The aircraft longitudinal maneuver load simulation device according to claim 7, characterized in that: In the speed-pressure calculation unit, the speed-pressure Q of the aircraft is calculated by the following formula: Q=0.5ρV.

9. The aircraft longitudinal maneuvering load simulation device according to claim 6, characterized in that: The aerodynamic force calculation module includes: The interpolation unit is used to interpolate the aircraft's aerodynamic characteristics data set to obtain the aircraft's drag coefficient C under the current aerodynamic angle of attack and elevator deflection conditions. D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq ; The aerodynamic parameter calculation unit is used to calculate the drag coefficient C of the aircraft. D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L, and pitching moment m of the aircraft y .

10. The aircraft longitudinal maneuvering load simulation device according to claim 9, characterized in that: In the interpolation unit, the drag coefficient C of the aircraft is interpolated by the following formula: D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq : The drag coefficient C of the aircraft is interpolated according to the flight Mach number Ma, the aerodynamic angle of attack α and the elevator deflection δe D , lift coefficient C L , pitching moment coefficient C my , according to the flight Mach number Ma, the aircraft trim angle of attack α interpolation pitch damping coefficient C mq ; In the aerodynamic parameter calculation unit, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated according to the following formulas: y : Where Q is the velocity pressure of the aircraft, S is the reference area of ​​the aircraft, q is the pitch angular velocity of the aircraft, V is the flight speed of the aircraft, and c is the average aerodynamic chord length.

11. A computer device, characterized in that: The method comprises a processor, a memory, and a computer program stored in the memory and executable on the processor, wherein the processor executes the computer program to implement the aircraft longitudinal maneuvering load simulation method according to any one of claims 1 to 5.

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