A method for controlling the climbing trajectory of an aircraft based on energy matching

Abstract

The present invention discloses a method for controlling the climbing trajectory of an aircraft based on energy matching, which comprises the following steps: S1: Establish the transfer function of the longitudinal channel of the target aircraft; S2: Design the stability augmentation loop of the control system and establish the transfer function of the stability augmentation loop; S3: Input the gains K ω , K α and K ny into the control system of the target aircraft, and the target aircraft climbs according to the gains K ω , K α and K ny ; S4: Calculate the difference ΔMa between the Mach number Ma of the actual flight state of the target aircraft and the Mach number of the reference ballistic trajectory, use the difference ΔMa and the Mach number Ma as the interpolation table of the dynamic limit value of the target aircraft, adjust the deviation limit value at different altitude positions according to the interpolation table, calculate the reasonable climbing rate in the climbing flight section of the target aircraft, and control the aircraft to climb according to the climbing rate. The present invention designs a reasonable climbing trajectory during the climbing flight of the target aircraft, calculates the reasonable climbing rate in the climbing flight section of the target aircraft, and prevents the occurrence of stall.

Description

Technical Field

[0001] The present invention relates to the field of aircraft fuselages, and in particular to a method for controlling the climbing trajectory of an aircraft based on energy matching. Background Art

[0002] During flight, the forces acting on the target aircraft mainly include gravity, thrust, and aerodynamic force. If the ballistic inclination angle of the target aircraft during climbing is too large and the climbing rate is relatively high, then the longitudinal component of the thrust is relatively large, and most of the thrust will be used to overcome the work done by gravity, reducing the utilization efficiency of the thrust; if the ballistic inclination angle of the climbing section is too low, then the target aircraft will climb very slowly. It takes a long time to climb to the predetermined height, resulting in excessive energy consumption of the target aircraft. Therefore, the design of the climbing trajectory during the takeoff and climbing stage of the target aircraft is crucial for the entire flight process.

[0003] For this reason, the present invention first proposes a method for controlling the climbing trajectory of an aircraft based on energy matching, and designs a reasonable climbing trajectory during the climbing flight of the target aircraft. If the speed of the target aircraft is relatively low during the climbing stage, then during the speed increase section, the ballistic inclination angle of the climbing flight of the target aircraft is relatively small, the thrust of the target aircraft is mainly used for speed increase, the longitudinal component of the thrust is relatively small, and the aerodynamic lift is also relatively small. At this time, the target aircraft climbs relatively "gently". After experiencing the speed increase section, the speed of the target aircraft is relatively large, and it enters the altitude increase section. The flight speed of the target aircraft in the altitude increase section is relatively large, and the generated lift is relatively large. The aerodynamic lift of the target aircraft is used to overcome gravity, and the thrust is used to overcome resistance. In the altitude increase section, the ballistic inclination angle of the target aircraft is relatively large, and it climbs to the predetermined height more "steeply". Summary of the Invention

[0004] In view of the above deficiencies of the prior art, the present invention provides a method for controlling the climbing trajectory of an aircraft based on energy matching.

[0005] To achieve the above invention object, the technical solution adopted by the present invention is as follows:

[0006] Provide a method for controlling the climbing trajectory of an aircraft based on energy matching, which includes the following steps:

[0007] S1: Establish a dynamic model of the target aircraft at different altitudes and Mach numbers during the climbing section, and perform small perturbation linearization and Laplace transform on the dynamic model to obtain the transfer function of the longitudinal channel of the target aircraft;

[0008] S2: Design the stability augmentation loop of the control system, establish the transfer function of the stability augmentation loop, and calculate the gain K of the angular rate feedback loop ω , the control gain K of the pseudo angle of attack feedback loop α and the control gain K at different altitudes and different Mach numbers ny ;

[0009] S3: Input the gains K ω , K α and K ny into the target drone control system. The target drone climbs according to the gains K ω , K α and K ny ;

[0010] S4: Calculate the difference ΔMa between the Mach number Ma of the actual flight state of the target drone and the Mach number of the reference trajectory. Use the difference ΔMa and the Mach number Ma as an interpolation table for the dynamic limit value of the target drone. Adjust the deviation limit value at different altitude positions according to the interpolation table, calculate the reasonable climb rate in the climbing flight section of the target drone, and control the climb of the aircraft according to the climb rate.

[0011] Further, step S1 includes:

[0012] S11: Establish the dynamic equation of the longitudinal channel of the target drone:

[0013]

[0014] where: θ is the ballistic inclination angle, m is the mass of the target drone, V is the speed of the target drone, P is the thrust of the target drone, α is the angle of attack, X is the axial force, Y is the normal force, J z is the pitch moment of inertia, ω z is the pitch rate, M z is the pitch moment, is the pitch angle, t is the flight time, g is the acceleration due to gravity, y is the normal flight distance, and x is the axial flight distance;

[0015] In the dynamic equation:

[0016]

[0017] where, C y is the lift coefficient of the target drone, q is the current dynamic pressure of the target drone, q = 1 / 2ρV 2 , s is the reference area of the target drone, Ma is the Mach number, δ and z is the pitch rudder deflection angle;

[0018] S12: Perform small perturbation linearization on the dynamic equation to obtain a differential equation set:

[0019]

[0020] In the differential equation set a4 = (P + Y α ) / (mV),

[0021] where, a1, a2, a3, a4, and a5 are all dynamic coefficients, and Y αis the normal force in the angle of attack direction, is the normal force in the pitch rudder deflection angle direction;

[0022] S13: By performing Laplace transform on the differential equation system and neglecting the small quantities of the dynamic coefficients reflecting the downwash of the projectile body and the control surface, the transfer function of the longitudinal channel of the target drone is obtained:

[0023]

[0024] where n y is the normal overload;

[0025] Furthermore, step S2 includes:

[0026] S21: Design the stability augmentation loop of the autopilot of the target drone. The stability augmentation loop includes an angular rate loop and a pseudo angle of attack loop;

[0027] S22: According to the control parameters and transfer function of the stability augmentation loop, establish the open-loop transfer function ω z _open(s) and the closed-loop transfer function ω z _close(s):

[0028]

[0029]

[0030] where, is the gain of the angular rate feedback loop;

[0031] S23: Establish the open-loop transfer function α_open(s) and the closed-loop transfer function α_close(s) of the pseudo angle of attack loop:

[0032]

[0033]

[0034] where K α is the control gain of the pseudo angle of attack feedback loop

[0035] S24: Establish the characteristic polynomial equation det(s) corresponding to the ideal poles:

[0036]

[0037] where ξ and ω n are the damping ratio and frequency corresponding to the ideal poles respectively;

[0038] S25: Utilize the characteristic polynomial equation det(s), the open-loop transfer function ω z _open(s), the closed-loop transfer function ωz _close(s), the open-loop transfer function α_open(s) of the pseudo angle of attack loop, and the closed-loop transfer function α_close(s) of the pseudo angle of attack loop are obtained as follows:

[0039]

[0040] S26: The gain of the angular rate feedback loop of the control system of the autopilot and the control gain K of the pseudo angle of attack feedback loop α are:

[0041]

[0042] S27: Set the control gain K of the target drone at different altitudes and different Mach numbers according to the amplitude margin, phase margin, and rise time T of the control system ny :

[0043]

[0044] Furthermore, step S4 includes:

[0045] S41: Calculate the difference ΔMa between the Mach number Ma of the actual flight state of the target drone and the Mach number of the reference trajectory. Use the difference ΔMa and the Mach number Ma as the interpolation table of the dynamic limit value of the target drone. The Mach number Ma is the horizontal axis variable of the interpolation table, and the difference ΔMa is the vertical axis variable of the interpolation table;

[0046] S42: Collect the Mach number Ma of the target drone during the climbing phase of flight. When the Mach number Ma is within the interval [Ma 标准 - Ma 波动 , Ma 标准 + Ma 波动 , then judge the magnitude of the difference ΔMa:

[0047] If ΔMa ≥ ΔMa 阈值 , then output the interpolation table through the interpolation module and select a smaller deviation limit value in the interpolation table to ensure the stable flight of the aircraft;

[0048] If ΔMa < ΔMa 阈值 , then output the interpolation table through the interpolation module and select a larger deviation limit value in the interpolation table to ensure that the aircraft has sufficient energy to track the reference trajectory;

[0049] S43: Calculate the reasonable climb rate of the target drone during the climbing flight section according to the obtained deviation limit value, and control the aircraft to climb according to the climb rate.

[0050] Furthermore, it also includes:

[0051] S5: Select several feature points on the flight trajectory of the target aircraft, collect different altitudes and different Mach numbers at different feature points, conduct margin checks on the performance of the autopilot according to different altitudes and different Mach numbers, and compare the amplitude margin, phase margin, cut-off frequency, and rise time at all feature points with the corresponding expected indicators respectively, so that the amplitude margin, phase margin, cut-off frequency, and rise time at all feature points meet the expected indicators.

[0052] The beneficial effects of the present invention are as follows: The present invention proposes a method for controlling the climbing trajectory of an aircraft based on energy matching, designs a reasonable climbing trajectory during the climbing flight of the target aircraft. If the speed of the target aircraft is relatively low during the climbing stage, then during the speed increase section, the ballistic inclination angle of the climbing flight of the target aircraft is relatively small, and the thrust of the target aircraft is mainly used for speed increase, the longitudinal component of the thrust is relatively small, and the aerodynamic lift is also relatively small. At this time, the target aircraft climbs relatively "gently". After experiencing the speed increase section, the speed of the target aircraft is relatively large, and it enters the altitude increase section. The flight speed of the target aircraft in the altitude increase section is relatively large, and the generated lift is relatively large. The aerodynamic lift of the target aircraft is used to overcome gravity, and the thrust is used to overcome resistance. In the altitude increase section, the ballistic inclination angle of the target aircraft is relatively large, and it climbs to the predetermined altitude more "steeply".

[0053] According to the actual flight Mach number Ma during the climbing flight, the Mach number of the reference trajectory, and the error ΔMa between the actual flight Mach number, design a two-dimensional interpolation table with dynamic limit values to obtain the real-time changing limit values, select different altitude position deviation limit values according to different flight states of the target aircraft to match the energy possessed by the target aircraft in real time, and calculate the reasonable climbing rate of the target aircraft during the climbing flight section to prevent stall. Description of the Drawings

[0054] Figure 1 It is the structure diagram of the altitude control loop.

[0055] Figure 2 It is the structure block diagram of the dynamic limit value.

[0056] Figure 3 It is the Mach number curve.

[0057] Figure 4 It is the altitude curve.

[0058] Figure 5 It is the ballistic inclination angle curve. Detailed Embodiment

[0059] The specific embodiments of the present invention will be described below to facilitate the understanding of those skilled in the art of the present invention. However, it should be clear that the present invention is not limited to the scope of the specific embodiments. For those of ordinary skill in the art, as long as various changes are within the spirit and scope of the present invention defined and determined by the appended claims, these changes are obvious, and all inventions made using the concept of the present invention are within the scope of protection.

[0060] The method for controlling the climbing trajectory of an aircraft based on energy matching in this solution includes the following steps:

[0061] S1: Establish the dynamic model of the target aircraft at different altitudes and Mach numbers during the climbing section, and perform small perturbation linearization and Laplace transform on the dynamic model to obtain the transfer function of the longitudinal channel of the target aircraft; Step S1 includes:

[0062] S11: Establish the dynamic equation of the longitudinal channel of the target aircraft:

[0063]

[0064] Where: θ is the ballistic inclination angle, m is the mass of the target aircraft, V is the speed of the target aircraft, P is the thrust of the target aircraft, α is the angle of attack, X is the axial force, Y is the normal force, J z is the pitch moment of inertia, ω z is the pitch rate, M z is the pitch moment, is the pitch angle, t is the flight time, g is the acceleration due to gravity, y is the normal flight distance, and x is the axial flight distance;

[0065] In the dynamic equation:

[0066]

[0067] Where, C y is the lift coefficient of the target aircraft, q is the current dynamic pressure of the target aircraft, q = 1 / 2ρV 2 , s is the reference area of the target, γ V is the speed roll angle, is the ballistic deflection angle, Ma is the Mach number, δ z is the pitch rudder deflection angle;

[0068] S12: Perform small perturbation linearization on the dynamic equation to obtain a system of differential equations:

[0069]

[0070] In the system of differential equations a4 = (P + Y α ) / (mV),

[0071] Among them, a1, a2, a3, a4, and a5 are all kinetic coefficients, and Y α is the normal force in the angle of attack direction, and

[0072] S13: By performing Laplace transform on the differential equation system and neglecting the small quantities of the kinetic coefficients of the reaction projectile and the downwash of the rudder surface, the transfer function of the longitudinal channel of the target aircraft is obtained:

[0073]

[0074] where n y is the normal overload;

[0075] S2: Design the stability augmentation loop of the control system. As Figure 1 shown, establish the transfer function of the stability augmentation loop through the structural diagram of the altitude control loop, and calculate the gain K ω of the angular rate feedback loop, the control gain K α of the pseudo angle of attack feedback loop, and the control gain K ny at different altitudes and different Mach numbers;

[0076] Step S2 includes:

[0077] S21: Design the stability augmentation loop of the autopilot of the target aircraft. The stability augmentation loop includes an angular rate loop and a pseudo angle of attack loop;

[0078] S22: According to the control parameters and transfer function of the stability augmentation loop, establish the open-loop transfer function ω z _open(s) and the closed-loop transfer function ω z _close(s) of the angular rate loop:

[0079]

[0080]

[0081] where is the gain of the angular rate feedback loop,

[0082] S23: Establish the open-loop transfer function α_open(s) and the closed-loop transfer function α_close(s) of the pseudo angle of attack loop:

[0083]

[0084]

[0085] where K α is the control gain of the pseudo angle of attack feedback loop;

[0086] S24: Establish the characteristic polynomial equation det(s) corresponding to the ideal poles:

[0087]

[0088] where ξ and ω n are the damping ratio and frequency corresponding to the ideal poles respectively;

[0089] S25: Use the characteristic polynomial equation det(s), the open-loop transfer function ω z _open(s), the closed-loop transfer function ω z _close(s), the open-loop transfer function α_opoen(s) of the pseudo angle of attack loop and the closed-loop transfer function α_close(s) of the pseudo angle of attack loop to obtain:

[0090]

[0091] S26: The gain of the angular rate feedback loop and the control gain K of the pseudo angle of attack feedback loop of the control system of the autopilot α are:

[0092]

[0093] S27: Set the control gain K of the target drone at different altitudes and different Mach numbers according to the amplitude margin, phase margin and rise time T of the control system ny :

[0094]

[0095] where V is the speed.

[0096] S3: Input the gains K ω , K α and K ny into the control system of the target drone, and the target drone climbs according to the gains K ω , K α and K ny ;

[0097] S4: Calculate the difference ΔMa between the Mach number Ma of the actual flight state of the target drone and the Mach number of the reference trajectory, use the difference ΔMa and the Mach number Ma as the interpolation table of the dynamic limit value of the target drone, and adjust the deviation limit value at different altitude positions according to the interpolation table. As Figure 2 shown, calculate the reasonable climb rate in the climbing flight section of the target drone, and control the aircraft to climb according to the climb rate. Step S4 includes:

[0098] S41: Calculate the difference ΔMa between the Mach number Ma of the actual flight state of the target aircraft and the Mach number of the reference trajectory. Use the difference ΔMa and the Mach number Ma as the interpolation table of the dynamic limit value of the target aircraft. The Mach number Ma is the horizontal axis variable of the interpolation table, and the difference ΔMa is the vertical axis variable of the interpolation table;

[0099] S42: Collect the Mach number Ma of the target aircraft during the climbing phase of flight. When the Mach number Ma is within the interval [Ma 标准 - Ma 波动 , Ma 标准 + Ma 波动 , then judge the magnitude of the difference ΔMa:

[0100] If ΔMa ≥ ΔMa 阈值 , then output the interpolation table through the interpolation module, and select a smaller deviation limit value in the interpolation table to ensure the stable flight of the aircraft;

[0101] If ΔMa < ΔMa 阈值 , then output the interpolation table through the interpolation module, and select a larger deviation limit value in the interpolation table to ensure that the aircraft has sufficient energy to track the reference trajectory;

[0102] S43: Calculate the reasonable climb rate of the target aircraft during the climbing flight section according to the obtained deviation limit value, and control the aircraft to climb according to the climb rate.

[0103] Through the interpolation table of the dynamic limit value, the vertical axis variable of the interpolation table is the difference ΔMa between the Mach number Ma of the reference trajectory and the Mach number of the actual flight state, and the horizontal axis variable of the interpolation table is the Mach number Ma of the actual flight of the target aircraft. When the flight Mach number of the target aircraft during the climbing section reaches a certain value, if the value of ΔMa is large, a smaller limit value will be obtained through the interpolation table. If the value of ΔMa is small, at this time the actual flight Mach number of the target aircraft is relatively close to the Mach number of the reference trajectory, then a larger limit value will be obtained through the interpolation table.

[0104] It also includes:

[0105] S5: Select several characteristic points on the flight trajectory of the target aircraft, collect different altitudes and different Mach numbers at different characteristic points, check the performance margin of the autopilot according to different altitudes and different Mach numbers, and compare the amplitude margin, phase margin, cut-off frequency and rise time at all characteristic points with the corresponding expected indicators respectively, so that the amplitude margin, phase margin, cut-off frequency and rise time at all characteristic points meet the expected indicators.

[0106] This embodiment is verified by simulation to verify the effectiveness of the invention effect as follows:

[0107] The module for selecting the dynamic limit value is added to the six-degree-of-freedom model to verify the effectiveness and feasibility of the method. According to a certain type of target drone, during the pulling-offset simulation climb process, due to the increased resistance and decreased thrust, the Mach number rises slowly during the climb. In order to prevent the target drone from stalling, a dynamic limit is added after the altitude position deviation. The dynamic limit value is output by comparing the actual flight Mach number with the Mach number of the reference trajectory. The simulation results are as Figures 3 - 5 shown.

Claims

1. A method for controlling the climbing trajectory of an aircraft based on energy matching, characterized in that, It includes the following steps: S1: Establish the dynamic model of the target aircraft at different altitudes and Mach numbers during the climb phase, and perform small perturbation linearization and Laplace transform on the dynamic model to obtain the transfer function of the longitudinal channel of the target aircraft; S2: Design the stability augmentation loop of the control system, establish the transfer function of the stability augmentation loop, and calculate the gain of the angular rate feedback loop , the control gain of the pseudo angle of attack feedback loop and the control gains at different altitudes and different Mach numbers ; S3: Apply the gain , and to the target drone control system, and the target drone climbs according to the gain , and ; S4: Calculate the Mach number of the actual flight state of the target aircraft Ma The difference Δ between the Mach number of the reference ballistic trajectory Ma , and use the difference Δ Ma and the Mach number Ma as the interpolation table of the dynamic limit value of the target aircraft. Adjust the deviation limit value at different altitude positions according to the interpolation table, calculate the reasonable climb rate of the target aircraft during the climbing flight section, and control the aircraft climb according to the climb rate; The said step S2 includes: S21: Design the stability augmentation loop of the autopilot of the target aircraft. The stability augmentation loop includes an angular rate loop and a pseudo angle of attack loop; S22: Establish the open-loop transfer function and closed-loop transfer function of the angular rate loop according to the control parameters and transfer function of the stability augmentation loop. and the closed-loop transfer function : Among them, is the gain of the angular rate feedback loop, a 1, a 2, a 3, a 4 are all kinetic coefficients; S23: Establish the open-loop transfer function of the pseudo angle of attack loop and the closed-loop transfer function of the pseudo angle of attack loop : Among them, is the control gain of the pseudo angle of attack feedback loop; S24: Establish the characteristic polynomial equation corresponding to the ideal pole point det( s ): Among them, and are the damping ratio and frequency corresponding to the ideal poles, respectively; S25: Using the characteristic polynomial equation det( s ), the open-loop transfer function , the closed-loop transfer function , the open-loop transfer function of the pseudo angle of attack loop and the closed-loop transfer function of the pseudo angle of attack loop to obtain: S26: The gain of the angular rate feedback loop and the control gain of the pseudo angle of attack feedback loop of the autopilot control system are as follows: S27: Set the control gains of the target drone at different altitudes and different Mach numbers according to the amplitude margin, phase margin, and rise time of the control system T :​ ； Among them, g is the acceleration due to gravity, V is the speed of the target drone.

2. The method for controlling the climbing trajectory of an aircraft based on energy matching according to claim 1, wherein The said step S1 includes: S11: Establish the dynamic equation of the longitudinal channel of the target aircraft: Wherein: is the ballistic inclination angle, m is the mass of the target drone, P is the thrust of the target drone, is the angle of attack, X is the axial force, Y is the normal force, is the pitch moment of inertia, is the pitch angular rate, is the pitch moment, is the pitch angle, t is the flight time, y is the normal flight distance, x is the axial flight distance; In the dynamic equation: Among them, is the lift coefficient of the target aircraft, is the current dynamic pressure of the target aircraft, , s is the reference area of the target aircraft, Ma is the Mach number, is the pitch rudder deflection angle; S12: Perform small perturbation linearization on the dynamic equation to obtain a differential equation set: In the system of differential equations , , , , ; Among them, a 5 is the kinetic coefficient, Y α is the normal force in the angle of attack direction, is the normal force in the pitch rudder deflection angle direction; S13: Through Laplace transform on the differential equation set and ignoring the small quantities of the dynamic coefficients reflecting the body and flap downwash, obtain the transfer function of the longitudinal channel of the target aircraft: Among them, n y is the normal overload.

3. The method for controlling the climbing trajectory of an aircraft based on energy matching according to claim 1, wherein The said step S4 includes: S41: Calculate the Mach number of the actual flight state of the target aircraft Ma The difference Δ between the Mach number of the actual flight state of the target aircraft and the Mach number of the reference trajectory Ma , and use the difference Δ Ma and the Mach number Ma as the interpolation table of the dynamic limit value of the target aircraft. The Mach number Ma is the horizontal axis variable of the interpolation table, and the difference Δ Ma is the vertical axis variable of the interpolation table; S42: Collect the Mach number of the target aircraft during the climbing phase of flight Ma , when the Mach number Ma is within the interval Ma 标准 -Ma 波动 , Ma 标准 +Ma 波动 , then judge the magnitude of the difference Δ Ma : If ΔMa≥ΔMa 阈值 , the interpolation table is output through the interpolation module, and a smaller deviation limit amplitude value in the interpolation table is selected to ensure the stable flight of the aircraft; If ΔMa < ΔMa 阈值 , the interpolation table is output through the interpolation module, and a relatively large deviation limit amplitude value in the interpolation table is selected, so as to ensure that the aircraft has sufficient energy to track the reference trajectory; S43: Calculate the reasonable climb rate of the target aircraft during the climb flight phase according to the obtained deviation limit amplitude, and control the aircraft climb according to the climb rate.

4. The method for controlling the climbing trajectory of an aircraft based on energy matching according to claim 1, wherein It also includes: S5: Select several characteristic points on the flight trajectory of the target aircraft, collect different altitudes and different Mach numbers at different characteristic points, perform margin verification on the performance of the autopilot according to different altitudes and different Mach numbers, and compare the amplitude margin, phase margin, cut-off frequency and rise time at all characteristic points with the corresponding expected indicators respectively, so that the amplitude margin, phase margin, cut-off frequency and rise time at all characteristic points meet the expected indicators.

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