A method and device for flight simulation under gust load

Abstract

The application 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, performing aircraft 1g level flight trimming to determine the aircraft trimming angle of attack and the aircraft trimming elevator deflection; S2, determining the aerodynamic force of the aircraft; S3, performing dynamic response solution on the aerodynamic force of the aircraft to obtain the aerodynamic angle of attack, the flight speed and the pitch angular velocity, and determining the normal overload of the aircraft after superimposing the flight speed on the wind speed of the vertical gust field; S4, obtaining the elevator deflection calculated by the stability augmentation system, superimposing the elevator deflection at a simulation moment to obtain the actual required elevator deflection; and S5, re-determining the aerodynamic force of the aircraft according to the flight speed, the pitch angular velocity, the aerodynamic angle of attack and the actual required elevator deflection until the simulation is completed. The application considers the role of the stability augmentation system in the determination of the aircraft gust load, and the simulation process is simple and has high precision.

Description

Technical Field

[0001] This application belongs to the field of aircraft simulation technology, and specifically relates to a flight simulation method and device under gust load. Background Technology

[0002] Traditional aircraft typically have sufficient longitudinal static stability margin, and usually do not need to manipulate the elevator during gust load calculations, and generally do not need to consider the impact of stabilization systems on gust loads.

[0003] Modern aircraft employ relaxed static stability design, making safe flight increasingly reliant on stability augmentation systems to achieve superior flight quality and safety. The basic principle of a longitudinal stability augmentation system is as follows: sensors measure signals such as the aircraft's angular rate and overload; the flight control computer calculates the elevator's motion commands according to a predetermined control law and drives the elevator to deflect, generating a reasonable and sufficient aerodynamic torque to provide additional motion damping and stability to the aircraft.

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

[0005] To address the aforementioned problems, the first aspect of this application provides a flight simulation method under gust loads, mainly comprising:

[0006] Step S1: Perform 1g level flight trim of the aircraft to determine the trim angle of attack and the trim elevator deflection of the aircraft, which will serve as the initial aerodynamic angle of attack and elevator deflection for the simulation.

[0007] Step S2: Determine the aerodynamic forces of the aircraft;

[0008] Step S3: Solve the aerodynamic response of the aircraft to obtain the aerodynamic angle of attack, flight speed and pitch rate. Then, add the flight speed to the wind speed of the vertical gust field to determine the normal overload of the aircraft.

[0009] Step S4: Obtain the elevator deflection calculated by the stabilization system based on the aerodynamic angle of attack and normal overload, and superimpose it with the elevator deflection at the previous simulation moment to obtain the actual required elevator deflection.

[0010] Step S5: Based on the flight speed, pitch rate, aerodynamic angle of attack, and the required elevator deflection, return to step S2, redetermine the aerodynamic forces of the aircraft, and repeat the above steps until the simulation ends.

[0011] Preferably, step S1 further includes:

[0012] S11. Obtain the input dataset of aircraft flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristics;

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

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

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

[0016] Q = 0.5ρV.

[0017] Preferably, step S2 further includes:

[0018] Step S21: Interpolate the aerodynamic characteristic dataset to obtain the drag coefficient C of the aircraft 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 aircraft's drag coefficient C D Lift coefficient C L Pitch 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 using the following formula. D Lift coefficient C L Pitch moment coefficient C my and pitch damping coefficient C mq :

[0021]

[0022] The drag coefficient C of the aircraft is interpolated based on the flight Mach number Ma, aerodynamic angle of attack α, and elevator deflection δe. D Lift coefficient C L Pitch moment coefficient C my Based on the flight Mach number Ma and the aircraft trim angle of attack α, the pitch damping coefficient C is interpolated. 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 rate of the aircraft, V is the flight speed of the aircraft, and c is the mean aerodynamic chord length.

[0026] Preferably, step S3 further includes:

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

[0028] Step S32: Add the aircraft speed to the wind speed of the vertical gust field and update the aircraft speed component.

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

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

[0031]

[0032] Where m is the mass of the aircraft; g is the acceleration due to gravity; u and w are the projections of the aircraft velocity onto the X and Z axes of the body axis, respectively; θ is the pitch angle of the aircraft; I yy Let Y be the moment of inertia of the aircraft relative to the Y-axis of the body axis system;

[0033] F xf X represents the projection of the aerodynamic force onto the X-axis of the body shaft system and the projection of the engine thrust onto the X-axis of the body shaft system. T The sum of; F zf Z represents the projection of the aerodynamic force onto the Z-axis of the body shaft system and the projection of the engine thrust onto the Z-axis of the body shaft system. T The sum of M; yf The aerodynamic torque m on the Y-axis of the body axis system y The torque m of the engine thrust on the Y-axis of the body shaft system T The sum is calculated using the following formula:

[0034]

[0035] In step S32, the aircraft's velocity components are updated using the following formula:

[0036]

[0037] Among them, u gThis represents the current vertical gust wind speed.

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

[0039]

[0040] The second aspect of this application provides a flight simulation device under gust loads, mainly comprising:

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

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

[0043] The dynamic response solution module is used to solve the aerodynamic forces of the aircraft, obtain the aerodynamic angle of attack, flight speed and pitch rate, and determine the normal overload of the aircraft by superimposing the flight speed with the wind speed of the vertical gust field.

[0044] The stability augmentation calculation module is used to obtain the elevator deflection calculated by the stability augmentation system based on the aerodynamic angle of attack and normal overload, and to superimpose the elevator deflection at the previous simulation moment to obtain the actual required elevator deflection.

[0045] The parameter update module is used to redetermine the aerodynamic forces of the aircraft based on flight speed, pitch rate, aerodynamic angle of attack, and the actual required elevator deflection.

[0046] Preferably, the parameter initialization module includes:

[0047] The flight parameter acquisition unit is used to acquire the input data of the aircraft's flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristics.

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

[0049] The trim unit is used to perform 1g trim on the aircraft based on the above parameters to obtain the aircraft trim angle of attack and the aircraft trim elevator deflection.

[0050] Preferably, in the velocity-pressure calculation unit, the velocity-pressure Q of the aircraft is calculated using the following formula:

[0051] Q = 0.5ρV.

[0052] Preferably, the aerodynamic calculation module includes:

[0053] Interpolation units are used to interpolate the drag coefficient C of the aircraft from the aircraft aerodynamic characteristic dataset 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 pitching moment m of the aircraft. y .

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

[0056]

[0057] The drag coefficient C of the aircraft is interpolated based on the flight Mach number Ma, aerodynamic angle of attack α, and elevator deflection δe. D Lift coefficient C L Pitch moment coefficient C my Based on the flight Mach number Ma and the aircraft trim angle of attack α, the pitch damping coefficient C is interpolated. mq .

[0058] 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 :

[0059]

[0060] Where Q is the velocity pressure of the aircraft, S is the reference area of ​​the aircraft, q is the pitch rate of the aircraft, V is the flight speed of the aircraft, and c is the mean aerodynamic chord length.

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

[0062] The dynamic response solution unit is used to solve the dynamic response of the aircraft using the fourth-order Runge-Kutta method, and to obtain the projection values ​​of the aircraft velocity on the X and Z axes of the body axis system, pitch angle and pitch angular velocity.

[0063] The speed update unit is used to superimpose the aircraft speed with the wind speed of the vertical gust field to update the aircraft's speed components.

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

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

[0066]

[0067] Where m is the mass of the aircraft; g is the acceleration due to gravity; u and w are the projections of the aircraft velocity onto the X and Z axes of the body axis, respectively; θ is the pitch angle of the aircraft; I yy Let Y be the moment of inertia of the aircraft relative to the Y-axis of the body axis system;

[0068] F xf X represents the projection of the aerodynamic force onto the X-axis of the body shaft system and the projection of the engine thrust onto the X-axis of the body shaft system. T The sum of; F zf Z represents the projection of the aerodynamic force onto the Z-axis of the body shaft system and the projection of the engine thrust onto the Z-axis of the body shaft system. T The sum of M; yf The aerodynamic torque m on the Y-axis of the body axis system y The torque m of the engine thrust on the Y-axis of the body shaft system T The sum is calculated using the following formula:

[0069]

[0070] In the speed update unit, the aircraft's speed components are updated using the following formula:

[0071]

[0072] Among them, u g This represents the current vertical gust wind speed.

[0073] In the longitudinal motion load calculation unit, the new aerodynamic angle of attack α and normal overload n are determined by the following formulas. z and flight Mach number Ma:

[0074]

[0075] A third aspect of this application provides a computer device including 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 stabilization system in the determination of gust loads on aircraft. The simulation process is simple and highly accurate, meeting the needs of engineering practice. Attached Figure Description

[0077] Figure 1 This is a flowchart of a preferred embodiment of the flight simulation method under gust loads of this application.

[0078] Figure 2 This application Figure 1 The vertical gust wind speed time history diagram of the embodiment shown.

[0079] Figure 3 This application Figure 1 Elevator deflection time history of the embodiment shown.

[0080] Figure 4 This application Figure 1 The normal overload time history diagram at the center of mass of the aircraft in the illustrated embodiment. Detailed Implementation

[0081] To make the objectives, technical solutions, and advantages of this application clearer, the technical solutions in the embodiments of this application will be described in more detail below with reference to the accompanying drawings. In the drawings, the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The described embodiments are only some, not all, of the embodiments of this application. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain this application, and should not be construed as limiting this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application. The embodiments of this application will be described in detail below with reference to the accompanying drawings.

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

[0083] Step S1: Perform 1g level flight trim of the aircraft to determine the trim angle of attack and the trim elevator deflection of the aircraft, which will serve as the initial aerodynamic angle of attack and elevator deflection for the simulation.

[0084] This step is primarily used to obtain the initial aerodynamic angle of attack and elevator deflection through 1g leveling for simulation loops. In addition, reference... Figure 1 Before the simulation begins, it is necessary to obtain parameters such as the aircraft's speed, altitude, mass, and aerodynamic characteristics dataset for use in the simulation calculations.

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

[0086] S11, obtain the input dataset of aircraft flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristics;

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

[0088] S13. Based on the above parameters, the aircraft is trimmed by 1g to obtain the aircraft trim angle of attack and the aircraft trim elevator deflection.

[0089] In some alternative implementations, in step S12, the velocity pressure Q of the aircraft is calculated using the following formula:

[0090] Q = 0.5ρV.

[0091] In this embodiment, step S13 is the 1g balancing step, which requires the parameters given in step S11, such as the aircraft's flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristic dataset. It also requires the parameter velocity-pressure Q from step S12, which needs to be calculated based on the atmospheric density ρ at flight altitude H. In addition, step S12 also needs to determine the speed of sound V at flight altitude H. S This is to calculate the Mach number Ma for a given aircraft's flight speed V.

[0092] After the calculation in step S13, the aircraft trim angle of attack α can be obtained. trim and the elevator offset δe trim This is used as the aerodynamic angle of attack α and elevator deflection δe in subsequent simulation cycles.

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

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

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

[0096] Step S21: Interpolate the aerodynamic characteristic dataset to obtain the drag coefficient C of the aircraft 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 aircraft's drag coefficient C D Lift coefficient C LPitch 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 .

[0098] In some alternative implementations, in step S21, the drag coefficient C of the aircraft is interpolated using the following formula. D Lift coefficient C L Pitch moment coefficient C my and pitch damping coefficient C mq :

[0099]

[0100] The drag coefficient C of the aircraft is interpolated based on the flight Mach number Ma, aerodynamic angle of attack α, and elevator deflection δe. D Lift coefficient C L Pitch moment coefficient C my Based on the flight Mach number Ma and the aircraft trim angle of attack α, the pitch damping coefficient C is interpolated. mq .

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

[0102]

[0103] Where Q is the velocity pressure of the aircraft, S is the reference area of ​​the aircraft, q is the pitch rate of the aircraft, V is the flight speed of the aircraft, and c is the mean aerodynamic chord length.

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

[0105] Step S3: Solve the aerodynamic response of the aircraft to obtain the aerodynamic angle of attack, flight speed and pitch rate. Then, determine the normal overload of the aircraft by superimposing the flight speed with the wind speed of the vertical gust field.

[0106] This step is mainly used to update the aerodynamic angle of attack, and the calculated normal overload will also be used to update the elevator deflection so that step S2 can be used repeatedly.

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

[0108] Step S31: Use the fourth-order Runge-Kutta method to solve the dynamic response of the aircraft and obtain the projection values ​​of the aircraft velocity on the X-axis and Z-axis of the body axis, pitch angle and pitch velocity.

[0109] Step S32: Add the aircraft speed to the wind speed of the vertical gust field and update the aircraft speed component.

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

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

[0112]

[0113] Where m is the mass of the aircraft; g is the acceleration due to gravity; u and w are the projections of the aircraft velocity onto the X and Z axes of the body axis, respectively; θ is the pitch angle of the aircraft; I yy Let Y be the moment of inertia of the aircraft relative to the Y-axis of the body axis system;

[0114] F xf X represents the projection of the aerodynamic force onto the X-axis of the body shaft system and the projection of the engine thrust onto the X-axis of the body shaft system. T The sum of; F zf Z represents the projection of the aerodynamic force onto the Z-axis of the body shaft system and the projection of the engine thrust onto the Z-axis of the body shaft system. T The sum of M; yf The aerodynamic torque m on the Y-axis of the body axis system y The torque m of the engine thrust on the Y-axis of the body shaft system T The sum is calculated using the following formula:

[0115]

[0116] In step S32, the aircraft's velocity components are updated using the following formula:

[0117]

[0118] Among them, u g This represents the current vertical gust wind speed.

[0119] In step S33, the new aerodynamic angle of attack α and normal overload n are determined using the following formulas. z and 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 wind speed on the aircraft is considered, and the aircraft's velocity components are updated. Finally, in step S33, the aerodynamic angle of attack α is updated. In the formula of step S33, atan is the arctangent function.

[0122] Step S4: Obtain the elevator deflection calculated by the stability augmentation system based on the aerodynamic angle of attack and normal overload, and superimpose it with the elevator deflection at the previous simulation moment to obtain the actual required elevator deflection.

[0123] refer to Figure 1 At the new simulation moment, the new aerodynamic angle of attack α and normal overload n calculated in step S32 will be used. z By transmitting the aircraft's stability augmentation system data, the required elevator deflection δe2 can be obtained. Then, by superimposing the elevator deflection δe1 from the previous moment, the actual required elevator deflection δe can be obtained.

[0124] Step S5: Based on the flight speed, pitch rate, aerodynamic angle of attack, and the required elevator deflection, return to step S2, redetermine the aerodynamic forces 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 loop also include the Mach number Ma updated in step S33 and the pitch rate q updated in step S31.

[0126] Figure 2 This is a time history graph of vertical gust wind speed, with time on the horizontal axis and wind speed on the vertical axis. Figure 3 This is the elevator deflection time history, with time on the horizontal axis and elevator deflection on the vertical axis. Without considering the stability augmentation system, the elevator maintains a trimmed deflection. With the stability augmentation system in mind, the elevator needs to be driven to deflect according to the control law. Figure 4 This is a time history graph of normal overload at the center of mass of the aircraft. The horizontal axis represents time, and the vertical axis represents the normal overload at the center of mass of the aircraft. After considering the effect of the stabilization system, the normal overload is reduced, especially at the second peak of the normal overload, where there is a significant reduction.

[0127] This application reasonably considers the role of the stabilization system in the determination of gust loads on aircraft. The simulation process is simple and highly accurate, meeting the needs of engineering practice.

[0128] The second aspect of this application provides a flight simulation device under gust loads 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 aircraft trim angle of attack and the aircraft trim elevator deflection, and use them as the initial aerodynamic angle of attack and elevator deflection for simulation.

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

[0131] The dynamic response solution module is used to solve the aerodynamic forces of the aircraft, obtain the aerodynamic angle of attack, flight speed and pitch rate, and determine the normal overload of the aircraft by superimposing the flight speed with the wind speed of the vertical gust field.

[0132] The stability augmentation calculation module is used to obtain the elevator deflection calculated by the stability augmentation system based on the aerodynamic angle of attack and normal overload, and to superimpose the elevator deflection at the previous simulation moment to obtain the actual required elevator deflection.

[0133] The parameter update module is used to redetermine the aerodynamic forces of the aircraft based on flight speed, pitch rate, aerodynamic angle of attack, and the actual required elevator deflection.

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

[0135] The flight parameter acquisition unit is used to acquire the input data of the aircraft's flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristics.

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

[0137] The trim unit is used to perform 1g trim on the aircraft based on the above parameters to obtain the aircraft trim angle of attack and the aircraft trim elevator deflection.

[0138] In some alternative embodiments, the velocity pressure Q of the aircraft is calculated in the velocity pressure calculation unit using the following formula:

[0139] Q = 0.5ρV.

[0140] In some alternative implementations, the aerodynamic calculation module includes:

[0141] Interpolation units are used to interpolate the drag coefficient C of the aircraft from the aircraft aerodynamic characteristic dataset 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 pitching moment m of the aircraft. y .

[0143] In some alternative embodiments, the drag coefficient C of the aircraft is interpolated in the interpolation unit using the following formula. D Lift coefficient C L Pitch moment coefficient C my and pitch damping coefficient C mq :

[0144]

[0145] The drag coefficient C of the aircraft is interpolated based on the flight Mach number Ma, aerodynamic angle of attack α, and elevator deflection δe. D Lift coefficient C L Pitch moment coefficient C my Based on the flight Mach number Ma and the aircraft trim angle of attack α, the pitch damping coefficient C is interpolated. mq .

[0146] 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 :

[0147]

[0148] Where Q is the velocity pressure of the aircraft, S is the reference area of ​​the aircraft, q is the pitch rate of the aircraft, V is the flight speed of the aircraft, and c is the mean aerodynamic chord length.

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

[0150] The dynamic response solution unit is used to solve the dynamic response of the aircraft using the fourth-order Runge-Kutta method, and to obtain the projection values ​​of the aircraft velocity on the X and Z axes of the body axis system, pitch angle and pitch angular velocity.

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

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

[0153]

[0154] Where m is the mass of the aircraft; g is the acceleration due to gravity; u and w are the projections of the aircraft velocity onto the X and Z axes of the body axis, respectively; θ is the pitch angle of the aircraft; I yy Let Y be the moment of inertia of the aircraft relative to the Y-axis of the body axis system;

[0155] F xf X represents the projection of the aerodynamic force onto the X-axis of the body shaft system and the projection of the engine thrust onto the X-axis of the body shaft system. T The sum of; F zf Z represents the projection of the aerodynamic force onto the Z-axis of the body shaft system and the projection of the engine thrust onto the Z-axis of the body shaft system. T The sum of M; yf The aerodynamic torque m on the Y-axis of the body axis system y The torque m of the engine thrust on the Y-axis of the body shaft system T The sum is calculated using the following formula:

[0156]

[0157] In step S32, the aircraft's velocity components are updated using the following formula:

[0158]

[0159] Among them, u g This represents the current vertical gust wind speed.

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

[0161]

[0162] A third aspect of this application provides a computer device including 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 description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the technical scope disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A flight simulation method under gust load, characterized in that, include: Step S1: Perform 1g level flight trim of the aircraft to determine the trim angle of attack and the trim elevator deflection of the aircraft, which will serve as the initial aerodynamic angle of attack and elevator deflection for the simulation. Step S2: Determine the aerodynamic forces of the aircraft; Step S3: Solve the aerodynamic response of the aircraft to obtain the aerodynamic angle of attack, flight speed and pitch rate. Then, add the flight speed to the wind speed of the vertical gust field to determine the normal overload of the aircraft. Step S4: Obtain the elevator deflection calculated by the stabilization system based on the aerodynamic angle of attack and normal overload, and superimpose it with the elevator deflection at the previous simulation moment to obtain the actual required elevator deflection. Step S5: Based on the flight speed, pitch rate, aerodynamic angle of attack, and the actual required elevator deflection, return to step S2, redetermine the aerodynamic forces of the aircraft, and repeat the above steps until the simulation ends. Step S2 further includes: Step S21: Interpolate the aerodynamic characteristic dataset to obtain the drag coefficient C of the aircraft 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 aircraft's drag coefficient C D Lift coefficient C L Pitch 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 ; Step S3 further includes: Step S31: Use the fourth-order Runge-Kutta method to solve the dynamic response of the aircraft and obtain the projection values ​​of the aircraft velocity on the X-axis and Z-axis of the body axis, pitch angle and pitch velocity. Step S32: Add the aircraft speed to the wind speed of the vertical gust field and update the aircraft speed component. Step S33: Determine the new aerodynamic angle of attack α and normal overload n based on the updated aircraft velocity components. z And the flight Mach number Ma.

2. The flight simulation method under gust load as described in claim 1, characterized in that, Step S1 further includes: S11. Obtain the input dataset of aircraft flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristics; S12. Calculate the atmospheric density ρ at flight altitude H, and further calculate the velocity pressure Q of the aircraft; S13. Based on the above parameters, perform 1g trim on the aircraft to obtain the aircraft trim angle of attack and the aircraft trim elevator deflection.

3. The flight simulation method under gust load as described in claim 2, characterized in that, In step S12, the velocity pressure Q of the aircraft is calculated using the following formula: Q = 0.5ρV.

4. The flight simulation method under gust load as described in claim 1, characterized in that, In step S21, the drag coefficient C of the aircraft is interpolated using the following formula. D Lift coefficient C L Pitch moment coefficient C my and pitch damping coefficient C mq : The drag coefficient C of the aircraft is interpolated based on the flight Mach number Ma, aerodynamic angle of attack α, and elevator deflection δe. D Lift coefficient C L Pitch moment coefficient C my Based on the flight Mach number Ma and the aircraft trim angle of attack α, the pitch damping coefficient C is interpolated. 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 rate of the aircraft, V is the flight speed of the aircraft, and c is the mean aerodynamic chord length.

5. The flight simulation method under gust load as described in claim 1, characterized in that, In step S31, the dynamic response is solved using the following formula: Where m is the mass of the aircraft; g is the acceleration due to gravity; u and w are the projections of the aircraft velocity onto the X and Z axes of the body axis, respectively; θ is the pitch angle of the aircraft; I yy Let Y be the moment of inertia of the aircraft relative to the Y-axis of the body axis system; F xf X represents the projection of the aerodynamic force onto the X-axis of the body shaft system and the projection of the engine thrust onto the X-axis of the body shaft system. T The sum of; F zf Z represents the projection of the aerodynamic force onto the Z-axis of the body shaft system and the projection of the engine thrust onto the Z-axis of the body shaft system. T The sum of M; yf The aerodynamic torque m on the Y-axis of the body axis system y The torque m of the engine thrust on the Y-axis of the body shaft system T The sum is calculated using the following formula: In step S32, the aircraft's velocity components are updated using the following formula: Among them, u g This represents the current vertical gust wind speed. In step S33, the new aerodynamic angle of attack α and normal overload n are determined using the following formulas. z and flight Mach number Ma:

6. 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 aircraft trim angle of attack and the aircraft trim elevator deflection, and use them as the initial aerodynamic angle of attack and elevator deflection for simulation. Aerodynamic calculation module, used to determine the aerodynamic forces of the aircraft; The dynamic response solution module is used to solve the aerodynamic forces of the aircraft, obtain the aerodynamic angle of attack, flight speed and pitch rate, and determine the normal overload of the aircraft by superimposing the flight speed with the wind speed of the vertical gust field. The stability augmentation calculation module is used to obtain the elevator deflection calculated by the stability augmentation system based on the aerodynamic angle of attack and normal overload, and to superimpose the elevator deflection at the previous simulation moment to obtain the actual required elevator deflection. The parameter update module is used to redetermine the aerodynamic forces of the aircraft based on the flight speed, pitch rate, aerodynamic angle of attack, and the actual required elevator deflection. The aerodynamic calculation module includes: Interpolation units are used to interpolate the drag coefficient C of the aircraft from the aircraft aerodynamic characteristic dataset 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 pitching moment m of the aircraft. y ; The dynamic response solving module includes: The dynamic response solution unit is used to solve the dynamic response of the aircraft using the fourth-order Runge-Kutta method, and to obtain the projection values ​​of the aircraft velocity on the X and Z axes of the body axis system, pitch angle and pitch angular velocity. The speed update unit is used to superimpose the aircraft speed with the wind speed of the vertical gust field to update the aircraft's speed components. The longitudinal maneuver load calculation unit is used to determine the new aerodynamic angle of attack α and normal overload n based on the updated aircraft velocity components. z And the flight Mach number Ma.

7. The flight simulation device under gust load as described in claim 6, characterized in that, The parameter initialization module includes: The flight parameter acquisition unit is used to acquire the input data of the aircraft's flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristics. The velocity-pressure calculation unit is used to calculate the atmospheric density ρ at flight altitude H, and further calculate the velocity-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 aircraft trim angle of attack and the aircraft trim elevator deflection.

8. The flight simulation device under gust load as described in claim 7, characterized in that, In the velocity-pressure calculation unit, the velocity-pressure Q of the aircraft is calculated using the following formula: Q = 0.5ρV.

9. The flight simulation device under gust load as described in claim 6, characterized in that, In the interpolation unit, the drag coefficient C of the aircraft is interpolated using the following formula. D Lift coefficient C L Pitch moment coefficient C my and pitch damping coefficient C mq : The drag coefficient C of the aircraft is interpolated based on the flight Mach number Ma, aerodynamic angle of attack α, and elevator deflection δe. D Lift coefficient C L Pitch moment coefficient C my Based on the flight Mach number Ma and the aircraft trim angle of attack α, the pitch damping coefficient C is interpolated. 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 rate of the aircraft, V is the flight speed of the aircraft, and c is the mean aerodynamic chord length.

10. The flight simulation device under gust load as described in claim 6, characterized in that, In the dynamic response solving unit, the dynamic response is solved using the following formula: Where m is the mass of the aircraft; g is the acceleration due to gravity; u and w are the projections of the aircraft velocity onto the X and Z axes of the body axis, respectively; θ is the pitch angle of the aircraft; I yy Let Y be the moment of inertia of the aircraft relative to the Y-axis of the body axis system; F xf X represents the projection of the aerodynamic force onto the X-axis of the body shaft system and the projection of the engine thrust onto the X-axis of the body shaft system. T The sum of; F zf Z represents the projection of the aerodynamic force onto the Z-axis of the body shaft system and the projection of the engine thrust onto the Z-axis of the body shaft system. T The sum of M; yf The aerodynamic torque m on the Y-axis of the body axis system y The torque m of the engine thrust on the Y-axis of the body shaft system T The sum is calculated using the following formula: In the speed update unit, the aircraft's speed components are updated using the following formula: Among them, u g This represents the current vertical gust wind speed. In the longitudinal motion load calculation unit, the new aerodynamic angle of attack α and normal overload n are determined by the following formulas. z and flight Mach number Ma:

11. A computer device, characterized in that, It 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 load as described in any one of claims 1-5.

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