Flight simulation method and device under gust load

Abstract

The invention belongs to the technical field of aircraft simulation, and particularly relates to a flight simulation method and device under gust load. The method comprises the following steps: S1, carrying out 1g level flight balancing on the aircraft, and determining a balancing attack angle of the aircraft and a balancing elevator skewness of the aircraft; s2, determining the aerodynamic force of the aircraft; s3, dynamic response solving is conducted on aerodynamic force of the aircraft, the aerodynamic attack angle, the flight speed and the pitch angle speed are obtained, and normal overload of the aircraft is determined after the flight speed is superposed with the wind speed of the vertical gust field; s4, acquiring the elevator skewness calculated by the stability augmentation system, and superposing the elevator skewness at the previous simulation moment to obtain the actually required elevator skewness; and S5, re-determining the aerodynamic force of the aircraft according to the flight speed, the pitch angle speed, the aerodynamic angle of attack and the actually required elevator skewness until the simulation is finished. The effect of the stability augmentation system in the determination process of the gust load of the aircraft is considered, the simulation process is simple, and the precision is high.

Description

Technical Field

[0001] The present application belongs to the technical field of aircraft simulation, and in particular relates to a flight simulation method and device under gust loads. Background Art

[0002] Traditional aircraft often have sufficient longitudinal static stability margin. The elevator is usually not manipulated during the gust load calculation process, and it is generally not necessary to consider the impact of the stability augmentation system on the gust load.

[0003] Modern aircraft will adopt a relaxed static stability design, and the safe flight of aircraft will rely more on the stability augmentation system to obtain better flight quality and safety. The basic principle of the longitudinal stability augmentation system is: using sensors to measure aircraft angular rate, overload and other signals, the flight control computer solves the elevator motion command according to the predetermined control law, and drives the elevator to deflect, so as to generate reasonable and sufficient aerodynamic torque to provide additional motion damping and stability for the aircraft.

[0004] For aircraft with relaxed static stability, the effect of the stabilization system on gust loads needs to be considered. Therefore, it is necessary to establish a flight simulation method under gust loads that takes into account the effect of the stabilization system to meet the needs of engineering practice. Summary of the invention

[0005] In order to solve the above problems, the first aspect of the present application provides a flight simulation method under gust load, which mainly includes:

[0006] Step S1, performing level flight trim of the aircraft 1g, determining the trim angle of attack of the aircraft 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, obtaining the aerodynamic angle of attack, flight speed and pitch angular velocity, and determining the normal overload of the aircraft by superimposing the flight speed on the wind speed of the vertical gust field;

[0009] Step S4, obtaining the elevator deflection calculated by the stability augmentation system according to the aerodynamic angle of attack and the normal overload, and superimposing the elevator deflection at the previous simulation moment to obtain the actually required elevator deflection;

[0010] Step S5, according to the flight speed, pitch angular velocity, aerodynamic angle of attack and the actually 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 , pitch 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 , pitch moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L and pitch 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 , pitch moment coefficient C my and pitch damping coefficient C mq :

[0021]

[0022] Among them, 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 , pitch 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 pitch moment m of the aircraft are calculated according to the following formula: y :

[0024]

[0025] Among them, 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, using the fourth-order Runge-Kutta method to solve the dynamic response of the aircraft, and 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, superimposing the aircraft velocity with the wind speed of the vertical gust field to update the aircraft velocity component;

[0029] Step S33: Determine the new aerodynamic angle of attack α and normal overload n based on the updated aircraft velocity component. z And flight Mach number Ma.

[0030] Preferably, in step S31, the dynamic response is solved by the following formula:

[0031]

[0032] 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 inertia moment of the aircraft relative to the Y axis of the body axis system;

[0033] F xf is the projection value of the aerodynamic force on the X-axis of the body axis system and the projection value of the engine thrust on the X-axis of the body axis system T The sum of zf is the projection value of the aerodynamic force on the Z axis of the body axis system and the projection value of the engine thrust on the Z axis of the body axis system T The sum of yf is the moment of aerodynamic force on the Y axis of the body axis system m y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows:

[0034]

[0035] In step S32, the velocity component of the aircraft is updated by the following formula:

[0036]

[0037] Among them, u gis the vertical gust field wind speed at the current moment;

[0038] In step S33, the new aerodynamic angle of attack α and normal overload n are determined by the following formulas: z And the flight Mach number Ma:

[0039]

[0040] The second aspect of the present application provides a flight simulation device under gust load, mainly comprising:

[0041] The parameter initialization module is used to perform 1g level flight trim of the aircraft, determine the trim angle of attack of the aircraft and the trim elevator deflection of the aircraft, as the initial aerodynamic angle of attack and elevator deflection of the simulation;

[0042] an aerodynamic calculation module, used to determine the aerodynamic forces of the aircraft;

[0043] A dynamic response solution module is used to perform dynamic response solution on the aerodynamic force of the aircraft, obtain the aerodynamic angle of attack, flight speed and pitch angular velocity, and determine the normal overload of the aircraft by superimposing the flight speed on the wind speed of the vertical gust field;

[0044] The stabilization calculation module is used to obtain the elevator deflection calculated by the stabilization system according to the aerodynamic angle of attack and the normal overload, and to superimpose the elevator deflection at the previous simulation moment to obtain the actually required elevator deflection;

[0045] The parameter updating module is used to redetermine the aerodynamic force of the aircraft according to the flight speed, pitch angular velocity, aerodynamic angle of attack and the actually required elevator deflection.

[0046] Preferably, the parameter initialization module includes:

[0047] A flight parameter acquisition unit, used to acquire input aircraft flight speed V, flight altitude H, aircraft mass m and aircraft aerodynamic characteristics data set;

[0048] The speed pressure calculation unit is used to calculate the atmospheric density ρ at the flight altitude H, and further calculate the speed pressure Q of the aircraft;

[0049] The trimming 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.

[0050] Preferably, in the speed pressure calculation unit, the speed pressure Q of the aircraft is calculated by the following formula:

[0051] Q=0.5ρV.

[0052] Preferably, the aerodynamic force calculation module comprises:

[0053] The interpolation unit is used to interpolate the aircraft 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 , pitch moment coefficient C my and pitch damping coefficient C mq ;

[0054] The aerodynamic parameter calculation unit is used to calculate the drag coefficient C of the aircraft. D , lift coefficient C L , pitch moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L and pitch moment m of the aircraft y .

[0055] Preferably, in the interpolation unit, the drag coefficient C of the aircraft is interpolated by the following formula: D , lift coefficient C L , pitch moment coefficient C my and pitch damping coefficient C mq :

[0056]

[0057] Among them, 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 , pitch moment coefficient C my , according to the flight Mach number Ma, the aircraft trim angle of attack α interpolation pitch damping coefficient C mq .

[0058] In the aerodynamic parameter calculation unit, the aerodynamic drag D, lift L and pitch moment m of the aircraft are calculated according to the following formula: y :

[0059]

[0060] Among them, 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.

[0061] Preferably, the dynamic response solving module includes:

[0062] The dynamic response solving unit is used to solve the dynamic response of the aircraft by 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;

[0063] A velocity updating unit, used to superimpose the aircraft velocity on the wind speed of the vertical gust field to update the aircraft velocity component;

[0064] The longitudinal maneuvering load calculation unit is used to determine the new aerodynamic angle of attack α, normal load n based on the updated aircraft velocity components. z And flight Mach number Ma.

[0065] Preferably, in the dynamic response solving unit, the dynamic response is solved by the following formula:

[0066]

[0067] 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 inertia moment of the aircraft relative to the Y axis of the body axis system;

[0068] F xf is the projection value of the aerodynamic force on the X-axis of the body axis system and the projection value of the engine thrust on the X-axis of the body axis system T The sum of zf is the projection value of the aerodynamic force on the Z axis of the body axis system and the projection value of the engine thrust on the Z axis of the body axis system T The sum of yf is the moment of aerodynamic force on the Y axis of the body axis system m y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows:

[0069]

[0070] In the velocity update unit, the velocity component of the aircraft is updated by the following formula:

[0071]

[0072] Among them, u g is the vertical gust field wind speed at the current moment;

[0073] In the longitudinal maneuvering 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:

[0074]

[0075] 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 flight simulation method under gust loads as described above.

[0076] This application reasonably considers the role of the stability augmentation system in the process of determining the gust load of the aircraft. The simulation process is simple and accurate, meeting the needs of engineering practice. BRIEF DESCRIPTION OF THE DRAWINGS

[0077] Figure 1 It is a flow chart of a preferred embodiment of the flight simulation method under gust load of the present application.

[0078] Figure 2 This application Figure 1 A time history graph of vertical gust wind speed for the illustrated embodiment.

[0079] Figure 3 This application Figure 1 Elevator deflection time history for the illustrated embodiment.

[0080] Figure 4 This application Figure 1 Time history of normal G-load at the center of mass of the vehicle for the illustrated embodiment. DETAILED DESCRIPTION

[0081] In order to make the purpose, technical scheme and advantages of the implementation of this application clearer, the technical scheme in the implementation of this application will be described in more detail in combination with the drawings in the implementation of this application. In the drawings, the same or similar reference numerals 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 cannot 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 combination with the drawings.

[0082] The first aspect of the present application provides a flight simulation method under gust load, such as Figure 1 As shown, it mainly includes:

[0083] Step S1, perform level flight trim of the aircraft 1g, determine the trim angle of attack of the aircraft and the trim elevator deflection of the aircraft, as the initial aerodynamic angle of attack and elevator deflection of the simulation.

[0084] This step is mainly used to obtain the initial aerodynamic angle of attack and elevator deflection through 1g distribution, so as to carry out simulation cycle. In addition, refer to Figure 1 ,Before the simulation starts, it is also necessary to obtain ,parameters such as the aircraft speed, altitude, mass, and aerodynamic ,characteristics data set for use in simulation calculations.

[0085] In some optional implementations, step S1 further includes:

[0086] S11, obtaining input aircraft flight speed V, flight altitude H, aircraft mass m and aircraft aerodynamic characteristics data set;

[0087] S12, calculating the atmospheric density ρ at the flight altitude H, and further calculating the speed pressure Q of the aircraft;

[0088] 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.

[0089] In some optional implementations, in step S12, the speed pressure Q of the aircraft is calculated by the following formula:

[0090] Q=0.5ρV.

[0091] 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, and also requires the use of the parameter speed pressure Q of step S12, which in turn 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.

[0092] 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.

[0093] Step S2: Determine the aerodynamic force of the aircraft.

[0094] With the aerodynamic angle of attack α and the elevator deflection δe, the aerodynamic force can be calculated.

[0095] In some optional implementations, step S2 further includes:

[0096] 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 , pitch moment coefficient C my and pitch damping coefficient C mq ;

[0097] Step S22: Based on the drag coefficient C of the aircraft D , lift coefficient C L, pitch moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L and pitch moment m of the aircraft y .

[0098] 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 , pitch moment coefficient C my and pitch damping coefficient C mq :

[0099]

[0100] Among them, 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 , pitch moment coefficient C my , according to the flight Mach number Ma, the aircraft trim angle of attack α interpolation pitch damping coefficient C mq .

[0101] In step S22, the aerodynamic drag D, lift L and pitch moment m of the aircraft are calculated according to the following formula: y :

[0102]

[0103] Among them, 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.

[0104] In the above embodiment, it can be seen from the formula that the aerodynamic coefficient can be interpolated based on the flight Mach number Ma, aerodynamic angle of attack α, and elevator control deflection δe of the aircraft at the current moment, and then the aerodynamic force can be further obtained based on the aircraft parameters.

[0105] Step S3, solving the dynamic response of the aerodynamic force of the aircraft, obtaining the aerodynamic angle of attack, flight speed and pitch angular velocity, and determining the normal overload of the aircraft by superimposing the flight speed on the wind speed of the vertical gust field.

[0106] 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.

[0107] In some optional implementations, step S3 further includes:

[0108] Step S31, using the fourth-order Runge-Kutta method to solve the dynamic response of the aircraft, and 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;

[0109] Step S32, superimposing the aircraft velocity with the wind speed of the vertical gust field to update the aircraft velocity component;

[0110] Step S33: Determine the new aerodynamic angle of attack α and normal overload n based on the updated aircraft velocity component. z And flight Mach number Ma.

[0111] In some optional implementations, in step S31, the dynamic response is solved by the following formula:

[0112]

[0113] 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 inertia moment of the aircraft relative to the Y axis of the body axis system;

[0114] F xf is the projection value of the aerodynamic force on the X-axis of the body axis system and the projection value of the engine thrust on the X-axis of the body axis system T The sum of zf is the projection value of the aerodynamic force on the Z axis of the body axis system and the projection value of the engine thrust on the Z axis of the body axis system T The sum of yf is the moment of aerodynamic force on the Y axis of the body axis system m y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows:

[0115]

[0116] In step S32, the velocity component of the aircraft is updated by the following formula:

[0117]

[0118] Among them, u g is the vertical gust field wind speed at the current moment;

[0119] In step S33, the new aerodynamic angle of attack α and normal overload n are determined by the following formulas: z And the flight Mach number Ma:

[0120]

[0121] 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 effect of the vertical gust field wind speed on the aircraft is considered and the velocity component of the aircraft is updated. Finally, in step S33, the aerodynamic angle of attack α is updated. The atan in the formula of step S33 is the inverse tangent function.

[0122] Step S4, obtaining the elevator deflection calculated by the stability augmentation system according to the aerodynamic angle of attack and the normal overload, and superimposing the elevator deflection at the previous simulation moment to obtain the actually required elevator deflection.

[0123] refer to Figure 1 At the new simulation time, the new aerodynamic angle of attack α and normal overload n calculated in step S32 are z By transmitting the aircraft's stabilization system, the elevator deflection δe2 required by the stabilization system can be obtained. Then, the elevator deflection δe1 at the previous moment is superimposed to obtain the actual required elevator deflection δe.

[0124] Step S5, according to the flight speed, pitch angular velocity, aerodynamic angle of attack and the actually required elevator deflection, return to step S2, re-determine the aerodynamic force of the aircraft and repeat the above steps until the simulation ends.

[0125] refer to Figure 1 In addition to the aerodynamic angle of attack α updated in step S33 and the elevator deflection δe updated in step S4, the parameters involved in the cycle also include the Mach number Ma updated in step S33 and the pitch angular velocity q updated in step S31.

[0126] Figure 2 It is a time history diagram of vertical gust wind speed, with the horizontal axis being time and the vertical axis being wind speed. Figure 3 It is the time history of elevator deflection, with the horizontal axis representing time and the vertical axis representing elevator deflection. Without considering the stabilization system, the elevator maintains the deflection in a trimmed state; after considering the effect of the stabilization system, the elevator needs to be driven to deflect according to the solution of the control law. Figure 4 It 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, the normal overload is reduced, especially at the second peak of the normal overload, there is a significant reduction.

[0127] This application reasonably considers the role of the stability augmentation system in the process of determining the gust load of the aircraft. The simulation process is simple and accurate, meeting the needs of engineering practice.

[0128] The second aspect of the present application provides a flight simulation device under gust load corresponding to the above method, mainly comprising:

[0129] The parameter initialization module is used to perform 1g level flight trim of the aircraft, determine the trim angle of attack of the aircraft and the trim elevator deflection of the aircraft, as the initial aerodynamic angle of attack and elevator deflection of the simulation;

[0130] an aerodynamic calculation module, used to determine the aerodynamic forces of the aircraft;

[0131] A dynamic response solution module is used to perform dynamic response solution on the aerodynamic force of the aircraft, obtain the aerodynamic angle of attack, flight speed and pitch angular velocity, and determine the normal overload of the aircraft by superimposing the flight speed on the wind speed of the vertical gust field;

[0132] The stabilization calculation module is used to obtain the elevator deflection calculated by the stabilization system according to the aerodynamic angle of attack and the normal overload, and to superimpose the elevator deflection at the previous simulation moment to obtain the actually required elevator deflection;

[0133] The parameter updating module is used to redetermine the aerodynamic force of the aircraft according to the flight speed, pitch angular velocity, aerodynamic angle of attack and the actually required elevator deflection.

[0134] In some optional implementations, the parameter initialization module includes:

[0135] A flight parameter acquisition unit, used to acquire input aircraft flight speed V, flight altitude H, aircraft mass m and aircraft aerodynamic characteristics data set;

[0136] The speed pressure calculation unit is used to calculate the atmospheric density ρ at the flight altitude H, and further calculate the speed pressure Q of the aircraft;

[0137] The trimming 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.

[0138] In some optional implementations, in the speed-pressure calculation unit, the speed-pressure Q of the aircraft is calculated by the following formula:

[0139] Q=0.5ρV.

[0140] In some optional embodiments, the aerodynamic force calculation module includes:

[0141] The interpolation unit is used to interpolate the aircraft 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 , pitch moment coefficient C my and pitch damping coefficient C mq ;

[0142] The aerodynamic parameter calculation unit is used to calculate the drag coefficient C of the aircraft. D , lift coefficient CL , pitch moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L and pitch moment m of the aircraft y .

[0143] 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 , pitch moment coefficient C my and pitch damping coefficient C mq :

[0144]

[0145] Among them, 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 , pitch moment coefficient C my , according to the flight Mach number Ma, the aircraft trim angle of attack α interpolation pitch damping coefficient C mq .

[0146] In the aerodynamic parameter calculation unit, the aerodynamic drag D, lift L and pitch moment m of the aircraft are calculated according to the following formula: y :

[0147]

[0148] Among them, 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.

[0149] In some optional implementations, the dynamic response solving module includes:

[0150] The dynamic response solving unit is used to solve the dynamic response of the aircraft by 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;

[0151] Longitudinal maneuver load calculation module, used to determine the new aerodynamic angle of attack α, normal overload n z And flight Mach number Ma.

[0152] In some optional implementations, in the dynamic response solving unit, the dynamic response is solved by the following formula:

[0153]

[0154] 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 inertia moment of the aircraft relative to the Y axis of the body axis system;

[0155] F xf is the projection value of the aerodynamic force on the X-axis of the body axis system and the projection value of the engine thrust on the X-axis of the body axis system T The sum of zf is the projection value of the aerodynamic force on the Z axis of the body axis system and the projection value of the engine thrust on the Z axis of the body axis system T The sum of yf is the moment of aerodynamic force on the Y axis of the body axis system m y The moment m of the engine thrust on the Y axis of the body axis system T The calculation formula is as follows:

[0156]

[0157] In step S32, the velocity component of the aircraft is updated by the following formula:

[0158]

[0159] Among them, u g is the vertical gust field wind speed at the current moment;

[0160] In step S33, the new aerodynamic angle of attack α and normal overload n are determined by the following formulas: z And the flight Mach number Ma:

[0161]

[0162] A third aspect of the present application provides a computer device, which includes 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 flight simulation method under gust loads as described above.

[0163] The above is only a specific implementation of the present application, but the protection scope of the present application 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 application should be included in the protection scope of the present application. Therefore, the protection scope of the present application shall be based on the protection scope of the claims.

Claims

1. A flight simulation method under gust load, characterized in that: include: Step S1, performing level flight trim of the aircraft 1g, determining the trim angle of attack of the aircraft 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, obtaining the aerodynamic angle of attack, flight speed and pitch angular velocity, and determining the normal overload of the aircraft by superimposing the flight speed on the wind speed of the vertical gust field; Step S4, obtaining the elevator deflection calculated by the stability augmentation system according to the aerodynamic angle of attack and the normal overload, and superimposing the elevator deflection at the previous simulation moment to obtain the actually required elevator deflection; Step S5, according to the flight speed, pitch angular velocity, aerodynamic angle of attack and the actually required elevator deflection, return to step S2, re-determine the aerodynamic force of the aircraft and repeat the above steps until the simulation ends.

2. The flight simulation method under gust load according to claim 1, characterized in that: 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 flight simulation method under gust load as claimed in claim 2, characterized in that: In step S12, the speed pressure Q of the aircraft is calculated by the following formula: Q=0.5ρV.

4. The flight simulation method under gust load according to claim 1, characterized in that: 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 , pitch 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 , pitch moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L and pitch moment m of the aircraft y .

5. The flight simulation method under gust load as claimed in claim 4, characterized in that: In step S21, the drag coefficient C of the aircraft is interpolated by the following formula: D , lift coefficient C L , pitch moment coefficient C my and pitch damping coefficient C mq : Among them, 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 , pitch 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 pitch moment m of the aircraft are calculated according to the following formula: y : Among them, 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. The flight simulation method under gust load according to claim 1, characterized in that: Step S3 further comprises: Step S31, using the fourth-order Runge-Kutta method to solve the dynamic response of the aircraft, and 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, superimposing the aircraft velocity with the wind speed of the vertical gust field to update the aircraft velocity component; Step S33: Determine the new aerodynamic angle of attack α and normal overload n based on the updated aircraft velocity component. z And flight Mach number Ma.

7. The flight simulation method under gust load as claimed in claim 6, characterized in that: In step S31, 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 inertia moment of the aircraft relative to the Y axis of the body axis system; F xf is the projection value of the aerodynamic force on the X-axis of the body axis system and the projection value of the engine thrust on the X-axis of the body axis system T The sum of zf is the projection value of the aerodynamic force on the Z axis of the body axis system and the projection value of the engine thrust on the Z axis of the body axis system T The sum of yf is the moment of aerodynamic force on the Y axis of the body axis system m 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 velocity component of the aircraft is updated by the following formula: Among them, u g is the vertical gust field wind speed at the current moment; In step S33, the new aerodynamic angle of attack α and normal overload n are determined by the following formulas: z And the flight Mach number Ma:

8. A flight simulation device under gust load, 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 of the aircraft and the trim elevator deflection of the aircraft, 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; A dynamic response solution module is used to perform dynamic response solution on the aerodynamic force of the aircraft, obtain the aerodynamic angle of attack, flight speed and pitch angular velocity, and determine the normal overload of the aircraft by superimposing the flight speed on the wind speed of the vertical gust field; The stabilization calculation module is used to obtain the elevator deflection calculated by the stabilization system according to the aerodynamic angle of attack and the normal overload, and to superimpose the elevator deflection at the previous simulation moment to obtain the actually required elevator deflection; The parameter updating module is used to redetermine the aerodynamic force of the aircraft according to the flight speed, pitch angular velocity, aerodynamic angle of attack and the actually required elevator deflection.

9. The flight simulation device under gust loads as claimed in claim 8, characterized in that: The parameter initialization module includes: A flight parameter acquisition unit, used to acquire input aircraft flight speed V, flight altitude H, aircraft mass m and aircraft aerodynamic characteristics data set; The speed pressure calculation unit is used to calculate the atmospheric density ρ at the flight altitude H, and further calculate the speed pressure Q of the aircraft; The trimming 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.

10. The flight simulation device under gust loads according to claim 9, 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.

11. The flight simulation device under gust loads according to claim 8, characterized in that: The aerodynamic force calculation module comprises: The interpolation unit is used to interpolate the aircraft 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 , pitch 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 , pitch moment coefficient C my and pitch damping coefficient C mq Determine the aerodynamic drag D, lift L and pitch moment m of the aircraft y .

12. The flight simulation device under gust loads according to claim 11, 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 , pitch moment coefficient C my and pitch damping coefficient C mq : Among them, 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 , pitch 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 pitch moment m of the aircraft are calculated according to the following formula: y : Among them, 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.

13. The flight simulation device under gust loads according to claim 8, characterized in that: The dynamic response solving module comprises: The dynamic response solving unit is used to solve the dynamic response of the aircraft by 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; A velocity updating unit, used to superimpose the aircraft velocity on the wind speed of the vertical gust field to update the aircraft velocity component; The longitudinal maneuvering load calculation unit is used to determine the new aerodynamic angle of attack α, normal load n based on the updated aircraft velocity components. z And flight Mach number Ma.

14. The flight simulation device under gust loads according to claim 13, characterized in that: 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 inertia moment of the aircraft relative to the Y axis of the body axis system; F xf is the projection value of the aerodynamic force on the X-axis of the body axis system and the projection value of the engine thrust on the X-axis of the body axis system T The sum of zf is the projection value of the aerodynamic force on the Z axis of the body axis system and the projection value of the engine thrust on the Z axis of the body axis system T The sum of yf is the moment of aerodynamic force on the Y axis of the body axis system m 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 velocity update unit, the velocity component of the aircraft is updated by the following formula: Among them, u g is the vertical gust field wind speed at the current moment; In the longitudinal maneuvering 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:

15. 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 flight simulation method under gust loads as claimed in any one of claims 1 to 7.

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