An airplane maneuver load simulation method and device considering the downwash effect

Abstract

The application belongs to the technical field of aircraft simulation, and particularly relates to an aircraft maneuver load simulation method and device considering the downwash effect. 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, superimposing the aerodynamic force coefficient increment caused by the downwash angle of attack on the aerodynamic force coefficient of the aircraft body to determine the aircraft aerodynamic force; S3, performing dynamic response solution on the aircraft aerodynamic force to obtain the aerodynamic angle of attack, the flight speed and the pitch angular velocity; S4, calculating the current elevator deflection according to the elevator command input; and S5, re-determining the aircraft aerodynamic force according to the flight speed, the pitch angular velocity, the aerodynamic angle of attack and the elevator deflection until the simulation is completed. The application reasonably considers the effect of the downwash in the process of determining the aircraft maneuver load, the simulation process is simple and has high precision, and the engineering practice needs are met.

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 aircraft maneuvering loads taking into account downwash effects. Background Art

[0002] In traditional aircraft flight load design, the aircraft is typically in freestream, unaffected by the downwash of other objects. However, when a large aircraft is in front of an aircraft, the aircraft behind it will be affected by the downwash of the preceding aircraft. This can alter the aerodynamic characteristics and aerodynamic load distribution of the aircraft behind it, necessitating consideration of downwash in flight load design.

[0003] The downwash generated by the leading aircraft is a complex aerodynamic disturbance, dependent on the flight speed, flight attitude, and spatial relationship of the preceding and following aircraft. When this downwash is strong, it can significantly impact the flight loads and flight safety of the following aircraft. Therefore, a method for determining aircraft maneuvering loads that accounts for downwash is needed to meet engineering needs. Summary of the Invention

[0004] In order to solve the above problems, the first aspect of the present application provides an aircraft maneuvering load simulation method considering downwash, which mainly includes:

[0005] 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;

[0006] Step S2: adding the aerodynamic coefficient increment caused by the downwash angle of attack to the aerodynamic coefficient of the aircraft body to obtain an aerodynamic coefficient correction value, and determining the aerodynamic force of the aircraft based on the aerodynamic coefficient correction value;

[0007] Step S3, solving the dynamic response of the aerodynamic force of the aircraft to obtain the aerodynamic angle of attack, flight speed and pitch angular velocity;

[0008] Step S4, calculating the elevator deflection at the current moment according to the elevator command input;

[0009] Step S5: Return to step S2 based on the flight speed, pitch angular velocity, aerodynamic angle of attack, and elevator deflection, redetermine the aerodynamic force of the aircraft, and repeat the above steps until the simulation is completed.

[0010] Preferably, step S1 further comprises:

[0011] S11, obtaining input aircraft flight speed V, flight altitude H, aircraft mass m, and aircraft aerodynamic characteristics data set;

[0012] S12, calculating the atmospheric density ρ at the flight altitude H, and further calculating the velocity pressure Q of the aircraft;

[0013] 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.

[0014] Preferably, in step S12, the speed pressure Q of the aircraft is calculated by the following formula:

[0015] Q=0.5ρV.

[0016] Preferably, step S2 further comprises:

[0017] Step S21: interpolate the aerodynamic characteristic data set of the aircraft body 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 ;

[0018] Step S22: interpolating the data set consisting of the aerodynamic characteristic data of the aircraft body and the downwash angle of attack to obtain a downwash aerodynamic coefficient increment;

[0019] Step S23: The drag coefficient C of the aircraft D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq Based on the above, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated by adding the downwash aerodynamic coefficient increment. 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 downwash aerodynamic coefficient increment is obtained by interpolation using the following formula:

[0024]

[0025] Among them, according to the flight Mach number Ma, aerodynamic angle of attack α and downwash angle of attack Interpolate the drag coefficient increment of the aircraft Lift coefficient increment Pitching moment coefficient increment

[0026] In step S23, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated according to the following formulas: y :

[0027]

[0028] 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.

[0029] Preferably, step S3 further comprises:

[0030] 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;

[0031] Step S32: Determine a new aerodynamic angle of attack α and flight Mach number Ma.

[0032] Preferably, in step S31, the dynamic response is solved by the following formula:

[0033]

[0034] 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;

[0035] 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:

[0036]

[0037] In step S32, the new aerodynamic angle of attack α and the flight Mach number Ma are determined by the following formula:

[0038]

[0039] A second aspect of the present application provides an aircraft maneuvering load simulation device taking into account downwash effects, mainly comprising:

[0040] 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;

[0041] an aerodynamic force calculation module, configured to superimpose the aerodynamic force coefficient increment caused by the downwash angle of attack onto the aerodynamic force coefficient of the aircraft body to obtain an aerodynamic force coefficient correction value, and determine the aerodynamic force of the aircraft based on the aerodynamic force coefficient correction value;

[0042] 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, flight speed and pitch angular velocity;

[0043] An elevator deflection calculation module is used to calculate the elevator deflection at the current moment according to the elevator command input;

[0044] The 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 elevator deflection.

[0045] Preferably, the parameter initialization module includes:

[0046] 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;

[0047] 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;

[0048] 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.

[0049] Preferably, in the speed-pressure calculation unit, the speed-pressure Q of the aircraft is calculated by the following formula:

[0050] Q=0.5ρV.

[0051] Preferably, the aerodynamic force calculation module includes:

[0052] The first interpolation unit is used to interpolate the aerodynamic characteristic data set of the aircraft body 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 ;

[0053] A second interpolation unit is configured to interpolate a downwash aerodynamic coefficient increment from a data set consisting of the aerodynamic characteristic data of the aircraft body and the downwash angle of attack;

[0054] Aerodynamic parameter calculation unit, 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 Based on the above, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated by adding the downwash aerodynamic coefficient increment. y .

[0055] Preferably, in the first 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 :

[0056]

[0057] 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 ;

[0058] In the second interpolation unit, the downwash aerodynamic coefficient increment is obtained by interpolation using the following formula:

[0059]

[0060] Among them, according to the flight Mach number Ma, aerodynamic angle of attack α and downwash angle of attack Interpolate the drag coefficient increment of the aircraft Lift coefficient increment Pitching moment coefficient increment

[0061] 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 :

[0062]

[0063] 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.

[0064] Preferably, the dynamic response solving module includes:

[0065] 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;

[0066] The longitudinal maneuvering load calculation unit is used to determine the new aerodynamic angle of attack α and the flight Mach number Ma.

[0067] Preferably, in the dynamic response solving unit, the dynamic response is solved by the following formula:

[0068]

[0069] 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;

[0070] 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:

[0071]

[0072] In the longitudinal maneuvering load calculation unit, the new aerodynamic angle of attack α and the flight Mach number Ma are determined by the following formula:

[0073]

[0074] 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 maneuvering load simulation method considering downwash as described above.

[0075] This application reasonably considers the role of downwash in determining the maneuvering load of the aircraft. The simulation process is simple and accurate, meeting the needs of engineering practice. BRIEF DESCRIPTION OF THE DRAWINGS

[0076] Figure 1 It is a flow chart of a preferred embodiment of the aircraft maneuvering load simulation method considering downwash effect of the present application. DETAILED DESCRIPTION

[0077] 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.

[0078] The first aspect of the present application provides a method for simulating aircraft maneuvering loads taking into account downwash effects, such as Figure 1 As shown, it mainly includes:

[0079] 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.

[0080] 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.

[0081] In some optional embodiments, step S1 further includes:

[0082] S11, obtaining input aircraft flight speed V, flight altitude H, aircraft mass m and aircraft aerodynamic characteristics data set;

[0083] S12, calculate the atmospheric density ρ at the flight altitude H, and further calculate the speed pressure Q of the aircraft;

[0084] 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.

[0085] In some optional implementations, in step S12, the speed pressure Q of the aircraft is calculated using the following formula:

[0086] Q=0.5ρV.

[0087] 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.

[0088] 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.

[0089] Step S2: superimposing the aerodynamic coefficient increment caused by the downwash angle of attack onto the aerodynamic coefficient of the aircraft body to obtain an aerodynamic coefficient correction value, and determining the aerodynamic force of the aircraft based on the aerodynamic coefficient correction value.

[0090] With the aerodynamic angle of attack α and the elevator deflection δe, the aerodynamic force can be calculated.

[0091] In some optional embodiments, step S2 further includes:

[0092] Step S21: interpolate the aerodynamic characteristic data set of the aircraft body 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 ;

[0093] Step S22: interpolating the data set consisting of the aerodynamic characteristic data of the aircraft body and the downwash angle of attack to obtain a downwash aerodynamic coefficient increment;

[0094] Step S23: The drag coefficient C of the aircraft D , lift coefficient CL , pitching moment coefficient C my and pitch damping coefficient C mq Based on the above, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated by adding the downwash aerodynamic coefficient increment. y .

[0095] 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 :

[0096]

[0097] 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 .

[0098] In step S22, the downwash aerodynamic coefficient increment is obtained by interpolation using the following formula:

[0099]

[0100] Among them, according to the flight Mach number Ma, aerodynamic angle of attack α and downwash angle of attack Interpolating the drag coefficient of an aircraft Lift coefficient increment Pitching moment coefficient increment

[0101] In step S23, 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 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 based on the current flight Mach number Ma, aerodynamic angle of attack α, and elevator control deflection δe of the aircraft, the aerodynamic coefficient of the aircraft body can be interpolated. Secondly, by obtaining the downwash angle of attack In step S22, the angle of attack and Mach number of the aircraft body are connected to interpolate the downwash aerodynamic coefficient increment, including the drag coefficient increment Lift coefficient increment Pitching moment coefficient increment Among them, downwash angle of attack It is related to the flight state of the preceding aircraft that causes downwash. Finally, 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 to obtain the aerodynamic angle of attack, flight speed and pitch angular velocity.

[0106] This step is mainly used to update the aerodynamic angle of attack so that step S2 can be used cyclically.

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

[0108] 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;

[0109] Step S32: Determine the new aerodynamic angle of attack α and normal overload n z And flight Mach number Ma.

[0110] In some optional implementations, in step S31, the dynamic response is solved using the following formula:

[0111]

[0112] 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;

[0113] 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:

[0114]

[0115] In step S32, the new aerodynamic angle of attack α and the flight Mach number Ma are determined by the following formula:

[0116]

[0117] 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.

[0118] Step S4: Calculate the elevator deflection at the current moment according to the elevator command input.

[0119] Step S5: Return to step S2 based on the flight speed, pitch angular velocity, aerodynamic angle of attack, and elevator deflection, redetermine the aerodynamic force of the aircraft, and repeat the above steps until the simulation is completed.

[0120] refer to Figure 1 At the new simulation moment, the downwash angle of attack, the pitch angular velocity q in step S31, the aerodynamic angle of attack α in step S32, the flight Mach number Ma, and the elevator deflection δe in step S4 are returned to step S2 to calculate the aircraft aerodynamic force at the current moment. This cycle continues until the simulation ends.

[0121] This application reasonably considers the role of downwash in determining the maneuvering load of the aircraft. The simulation process is simple and accurate, meeting the needs of engineering practice.

[0122] A second aspect of the present application provides an aircraft maneuvering load simulation device taking into account downwash effects corresponding to the above method, mainly comprising:

[0123] 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;

[0124] an aerodynamic force calculation module, configured to superimpose the aerodynamic force coefficient increment caused by the downwash angle of attack onto the aerodynamic force coefficient of the aircraft body to obtain an aerodynamic force coefficient correction value, and determine the aerodynamic force of the aircraft based on the aerodynamic force coefficient correction value;

[0125] 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, flight speed and pitch angular velocity;

[0126] An elevator deflection calculation module is used to calculate the elevator deflection at the current moment according to the elevator command input;

[0127] The 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 elevator deflection.

[0128] In some optional implementations, the parameter initialization module includes:

[0129] 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;

[0130] 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;

[0131] 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.

[0132] In some optional implementations, in the speed-pressure calculation unit, the speed-pressure Q of the aircraft is calculated using the following formula:

[0133] Q=0.5ρV.

[0134] In some optional embodiments, the aerodynamic force calculation module includes:

[0135] The first interpolation unit is used to interpolate the aerodynamic characteristic data set of the aircraft body 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 ;

[0136] A second interpolation unit is configured to interpolate a downwash aerodynamic coefficient increment from a data set consisting of the aerodynamic characteristic data of the aircraft body and the downwash angle of attack;

[0137] Aerodynamic parameter calculation unit, 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 Based on the above, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated by adding the downwash aerodynamic coefficient increment. y .

[0138] In some optional embodiments, in the first 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 :

[0139]

[0140] 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 ;

[0141] In the second interpolation unit, the downwash aerodynamic coefficient increment is obtained by interpolation using the following formula:

[0142]

[0143] Among them, according to the flight Mach number Ma, aerodynamic angle of attack α and downwash angle of attack Interpolate the drag coefficient increment of the aircraft Lift coefficient increment Pitching moment coefficient increment

[0144] 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 :

[0145]

[0146] 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.

[0147] In some optional implementations, the dynamic response solving module includes:

[0148] 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;

[0149] The longitudinal maneuvering load calculation unit is used to determine the new aerodynamic angle of attack α and the flight Mach number Ma.

[0150] In some optional implementations, in the dynamic response solving unit, the dynamic response is solved by the following formula:

[0151]

[0152] 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;

[0153] F xfis 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:

[0154]

[0155] In the longitudinal maneuvering load calculation unit, the new aerodynamic angle of attack α and the flight Mach number Ma are determined by the following formula:

[0156]

[0157] 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 maneuvering load simulation method considering downwash as described above.

[0158] 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 aircraft maneuvering loads taking into account downwash, 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: adding the aerodynamic coefficient increment caused by the downwash angle of attack to the aerodynamic coefficient of the aircraft body to obtain an aerodynamic coefficient correction value, and determining the aerodynamic force of the aircraft based on the aerodynamic coefficient correction value; Step S3, solving the dynamic response of the aerodynamic force of the aircraft to obtain the aerodynamic angle of attack, flight speed and pitch angular velocity; Step S4, calculating the elevator deflection at the current moment according to the elevator command input; Step S5: Return to step S2 to re-determine the aerodynamic force of the aircraft based on the flight speed, pitch angular velocity, aerodynamic angle of attack, and elevator deflection, and repeat the above steps until the simulation ends. Step S2 further comprises: Step S21: interpolate the aerodynamic characteristic data set of the aircraft body 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 ; Step S22: interpolating the data set consisting of the aerodynamic characteristic data of the aircraft body and the downwash angle of attack to obtain a downwash aerodynamic coefficient increment; Step S23: The drag coefficient C of the aircraft D , lift coefficient C L , pitching moment coefficient C my and pitch damping coefficient C mq Based on the above, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated by adding the downwash aerodynamic coefficient increment. y .

2. The aircraft maneuvering load simulation method considering downwash as claimed in 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 aircraft maneuvering load simulation method considering downwash as claimed in claim 2, characterized in that: In step S12, the speed pressure Q of the aircraft is calculated using the following formula: Q=0.5ρV.

4. The aircraft maneuvering load simulation method considering downwash as claimed in claim 1, characterized in that: 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 downwash aerodynamic coefficient increment is obtained by interpolation using the following formula: Among them, according to the flight Mach number Ma, aerodynamic angle of attack α and downwash angle of attack Interpolate the drag coefficient increment of the aircraft Lift coefficient increment Pitching moment coefficient increment In step S23, 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.

5. The aircraft maneuvering load simulation method considering downwash as claimed in claim 4, characterized in that: Step S3 further comprises: 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 a new aerodynamic angle of attack α and flight Mach number Ma.

6. The aircraft maneuvering load simulation method considering downwash as claimed in claim 5, 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 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 the flight Mach number Ma are determined by the following formula: Among them, atan is the inverse tangent function, V S The speed of sound.

7. An aircraft maneuvering load simulation device taking into account downwash, 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 force calculation module, configured to superimpose the aerodynamic force coefficient increment caused by the downwash angle of attack onto the aerodynamic force coefficient of the aircraft body to obtain an aerodynamic force coefficient correction value, and determine the aerodynamic force of the aircraft based on the aerodynamic force coefficient correction value; 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, flight speed and pitch angular velocity; An elevator deflection calculation module is used to calculate the elevator deflection at the current moment according to the elevator command input; A parameter update module is used to re-determine the aerodynamic force of the aircraft based on the flight speed, pitch rate, aerodynamic angle of attack and elevator deflection; Wherein, the aerodynamic force calculation module includes: The first interpolation unit is used to interpolate the aerodynamic characteristic data set of the aircraft body 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 ; A second interpolation unit is configured to interpolate a downwash aerodynamic coefficient increment from a data set consisting of the aerodynamic characteristic data of the aircraft body and the downwash angle of attack; Aerodynamic parameter calculation unit, 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 Based on the above, the aerodynamic drag D, lift L and pitching moment m of the aircraft are calculated by adding the downwash aerodynamic coefficient increment. y .

8. The aircraft according to claim 7 taking into account the downwash effect The maneuverable load simulation device is 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.

9. The aircraft maneuvering load simulation device taking into account downwash as claimed in claim 8, 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.

10. The aircraft maneuvering load simulation device taking into account downwash as claimed in claim 7, characterized in that: In the first 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 second interpolation unit, the downwash aerodynamic coefficient increment is obtained by interpolation using the following formula: Among them, according to the flight Mach number Ma, aerodynamic angle of attack α and downwash angle of attack Interpolate the drag coefficient increment of the aircraft Lift coefficient increment Pitching moment coefficient increment 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. The aircraft maneuvering load simulation device considering downwash as claimed in claim 10, characterized in that: The dynamic response solving 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; The longitudinal maneuvering load calculation unit is used to determine the new aerodynamic angle of attack α and the flight Mach number Ma.

12. The aircraft maneuvering load simulation device taking into account downwash as claimed in claim 11, 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 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 maneuvering load calculation unit, the new aerodynamic angle of attack α and the flight Mach number Ma are determined by the following formula: Among them, atan is the inverse tangent function, V S The speed of sound.

13. 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 maneuvering load simulation method considering downwash as claimed in any one of claims 1 to 6.

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