Autonomous control method and system for fixed-point landing of unpowered return high-speed aircraft
By adopting autonomous control methods on high-speed aircraft without power return, including attitude angle controller, sink rate controller, speed controller and lateral side controller, the linear expansion state observer estimates system interference, the problem of fixed-point landing in the unpowered state is solved, and high-precision and safe landing effect is achieved.
Patent Information
- Application Number
- CN202510194055.9
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-02-21
- Publication Date
- 2025-05-30
AI Technical Summary
High-speed aircraft without power returns face factors such as aerodynamic interference, wind speed changes and ground effects during the landing stage, which affects the aircraft's flight trajectory and landing accuracy. It is difficult for the existing technology to achieve fixed-point landing in a state without power.
The autonomous control method including attitude angle controller, sinking rate controller, speed controller and lateral lateral controller is adopted to estimate system interference in real time through a linear expansion state observer (LESO), and corresponding control instructions are generated in combination with gain and dynamic coefficients to ensure that the aircraft achieves fixed-point landing in a powerless state.
It improves the landing accuracy and safety of the aircraft in a powerless state, can cope with complex weather and aircraft performance deviations, and ensures the accuracy and robustness of fixed-point landing.
Smart Images
Figure CN120066072A_ABST
Abstract
Description
Technical Field
[0001] The present invention relates to the technical field of aircraft control, and particularly relates to an autonomous control method and system for the fixed-point landing of a high-speed aircraft with no power return. Background Art
[0002] High-speed aircraft with no power return face many challenges during the landing phase, including factors such as aerodynamic interference, wind speed changes, and ground effect, which will affect the flight trajectory and landing accuracy of the aircraft. Traditional landing control methods often rely on the power system of the aircraft for adjustment. However, in the powerless state, the aircraft can only achieve a safe landing by controlling its attitude and trajectory. Therefore, developing an efficient autonomous control method to ensure that the aircraft can achieve fixed-point landing in the powerless state has important practical significance. Summary of the Invention
[0003] The present invention provides an autonomous control method and system for the fixed-point landing of a high-speed aircraft with no power return, which can improve the landing accuracy and safety of the aircraft in the powerless state.
[0004] To solve the above technical problems, the present invention provides an autonomous control method for the fixed-point landing of a high-speed aircraft with no power return, including an attitude angle controller, a sink rate controller, a speed controller, and a lateral controller;
[0005] The attitude angle controller includes a pitch angle control loop, a yaw angle control loop, and a roll angle control loop.
[0006] The pitch angle loop takes the deviation between the pitch angle command θ c and the actual pitch angle θ as the input, and estimates the interference in real time through a linear extended state observer LESO θ and combines the pitch angle loop gain k θ and the system dynamic coefficient to generate a pitch rate command Q c and outputs it to the pitch rate loop; the pitch rate loop takes the deviation between the pitch rate command Q c and the actual pitch rate Q as the input, further estimates the system interference through the LESO Q and combines the pitch rate loop gain k Q and the dynamic coefficient to generate an elevator command δ e so as to control the attitude angle of the aircraft;
[0007] The yaw angle loop selects the yaw angular velocity R as the input, multiplies it by the yaw angular velocity gain after passing through a high-pass filter to generate a rudder command δ r ;
[0008] The roll angle loop takes the roll angle command γ cThe deviation from the actual roll angle γ is used as the input, and through the linear extended state observer LESO γ the disturbance is estimated in real time, and combined with the roll angle loop gain k γ and the system dynamic coefficient to generate the roll angle rate command P c and output it to the roll angle rate loop; the roll angle rate loop uses the roll angle rate command P c and the deviation from the actual roll angle rate P as the input, and through the LESO P further estimates the system disturbance, and combines the roll angle rate loop gain k P and the dynamic coefficient to generate the aileron command δ a so as to control the attitude angle of the aircraft;
[0009] The sink rate controller includes a height control loop and a sink rate control loop. The height control loop uses the target height command H c and the deviation from the actual height H as the input, and generates the sink rate command through a proportional-integral controller and outputs it to the sink rate control loop; the sink rate control loop uses the nominal trajectory sink rate and the actual sink rate as the input, estimates the system disturbance through the linear extended state observer LESO, combines the proportional gain and the system dynamic coefficient to generate the pitch angle command θ c so as to control the sink rate;
[0010] The speed control loop of the speed controller uses the target speed command V c and the actual speed V as the input, generates the speed deviation amount, and after proportional-integral operation, combines the spoiler neutral position deflection angle to output the spoiler command so as to control the speed;
[0011] The lateral controller adopts the L 1 tracking guidance algorithm. According to the lateral offset d between the current position of the aircraft and the target straight path and the course deviation Δψ between the speed direction and the target straight direction, it outputs the roll angle command γ c so as to correct the course and the offset range.
[0012] In some of the embodiments, the pitch angle loop LESO θ is designed as follows:
[0013]
[0014] Where: is the estimated value of the pitch angle, is the estimated value of the total disturbance of the pitch angle by the observer, ωθ is the observation bandwidth of the pitch angle expansion state observer, b θ is the pitch angle control input gain.
[0015] The pitch angle rate command is designed as follows:
[0016]
[0017] In the formula: θ c is the pitch angle command, is the estimated value of the total disturbance of the pitch angle by the observer, k θ is the pitch angle loop gain, b θ is the pitch angle control input gain.
[0018] In some of the embodiments, the pitch angle rate loop LESO Q is designed as follows:
[0019]
[0020] In the formula: is the estimated value of the pitch angle rate, is the estimated value of the total disturbance of the pitch angle rate by the observer, ω Q is the observation bandwidth of the pitch angle rate expansion state observer, b Q is the pitch angle rate control input gain.
[0021] The elevator command δ e is designed as follows:
[0022]
[0023] In the formula: is the estimated value of the total disturbance of the pitch angle rate by the observer, k Q is the pitch angle rate loop gain, b Q is the pitch angle rate control input gain.
[0024] In some of the embodiments, the yaw angle loop aileron command δ r is designed as follows:
[0025]
[0026] In some of the embodiments, the roll angle loop LESO γ is designed as follows:
[0027]
[0028] In the formula: is the estimated value of the roll angle, is the estimated value of the total disturbance of the roll angle, ω γ is the observation bandwidth of the roll angle extended state observer, b γ is the roll angle control input gain.
[0029] The roll rate command is designed as follows:
[0030]
[0031] where: γ c is the roll angle command, is the estimated value of the total disturbance of the roll angle by the observer, k γ is the roll angle loop gain, b γ is the roll angle control input gain.
[0032] In some embodiments, the roll rate loop LESO P is designed as follows::
[0033]
[0034] where: is the estimated value of the roll rate, is the estimated value of the total disturbance of the roll rate by the observer, ω P is the observation bandwidth of the roll rate extended state observer, b P is the roll rate control input gain.
[0035] The aileron command δ a is designed as follows:
[0036]
[0037] where: is the estimated value of the total disturbance of the roll rate by the observer, k P is the roll rate loop gain, b P is the roll rate control input gain.
[0038] In some embodiments, the sink rate command is designed as shown below:
[0039]
[0040] In the above formula, is the sink rate command, is the nominal trajectory sink rate, H c is the desired altitude, is the altitude control gain.
[0041] The pitch angle command θ cThe design is as shown in the formula:
[0042]
[0043] In the formula, is the control input gain, is the sink rate control loop gain, is the estimated value of the total uncertainty of the lift rate LESO with respect to the lift rate, is the desired sink rate, is the sink rate.
[0044] In some of the embodiments, the spoiler command is as shown in the following formula:
[0045]
[0046] In the formula, is the spoiler deflection command, is the spoiler neutral position deflection, V c is the speed command, is the speed control gain.
[0047] In some of the embodiments, the roll angle command γ c is calculated through the following steps:
[0048] Step 1: Design the lateral overload command as:
[0049]
[0050] η is a value with a very small magnitude, and it can be approximately obtained that:
[0051] sinη≈η=η 1 +Δψ
[0052]
[0053] Step 2: Design the lateral acceleration command as:
[0054]
[0055] Step 3: Since γ c =gtanθ k , therefore when the trajectory angle θ k is small, the aircraft roll angle command can be approximately expressed as:
[0056]
[0057] The L 1is the distance from the aircraft to the virtual reference point. Specifically, the lateral offset d is the perpendicular distance from the aircraft to the runway centerline, and the heading deviation Δψ is the angle between the velocity direction and the runway landing direction.
[0058] The present invention also provides an autonomous control system for the fixed-point landing of a powerless return high-speed aircraft, which stores the software program corresponding to the above control method.
[0059] The beneficial effects of the present invention compared with the prior art are as follows: By adopting an attitude angle controller based on active disturbance rejection control, the stable control of the aircraft attitude angle is realized; a sink rate control loop is designed and combined with altitude control to ensure that the sink rate of the aircraft meets the requirements before the final touchdown; the speed is controlled by a speed brake to ensure that the speed of the aircraft is controllable during the landing section; the lateral direction adopts an L 1 control method to ensure the accuracy of the aircraft landing point; the control method of the present invention has good robustness, can cope with complex weather and aircraft performance deviations, has high control precision, ensures that the aircraft can land at a designated position, and the system design is simple and easy to be realized in actual engineering. BRIEF DESCRIPTION OF THE DRAWINGS
[0060] Figure 1 Schematic diagram of the control loop of the pitch angle controller;
[0061] Figure 2 Schematic diagram of the control loop of the yaw angle controller;
[0062] Figure 3 Schematic diagram of the control loop of the roll angle controller;
[0063] Figure 4 Schematic diagram of the control loop of the sink rate controller;
[0064] Figure 5 Schematic diagram of the control loop of the speed controller;
[0065] Figure 6 Schematic diagram of the linear tracking L1 guidance scheme. DETAILED DESCRIPTION OF THE EMBODIMENTS
[0066] In order to make the purpose, technical solutions and advantages of the present invention clearer, the present invention will be further described in detail below with reference to the drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present invention and are not used to limit the present invention. In addition, the technical features involved in the various embodiments of the present invention described below can be combined with each other as long as they do not conflict with each other.
[0067] This embodiment provides an autonomous control method for the fixed-point landing of a powerless return high-speed aircraft, including an attitude angle controller, a sink rate controller, a speed controller, and a lateral controller;
[0068] The attitude angle controller includes a pitch angle control loop, a yaw angle control loop, and a roll angle control loop.
[0069] The pitch angle loop takes the deviation between the pitch angle command θ c and the actual pitch angle θ as the input, and estimates the disturbance in real time through the linear extended state observer LESO θ and combines the pitch angle loop gain k θ and the system dynamic coefficient to generate the pitch angle rate command Q c and outputs it to the pitch angle rate loop; the pitch angle rate loop takes the deviation between the pitch angle rate command Q c and the actual pitch angle rate Q as the input, and further estimates the system disturbance through the LESO Q and combines the pitch angle rate loop gain k Q and the dynamic coefficient to generate the elevator command δ e so as to control the attitude angle of the aircraft;
[0070] The yaw angle loop selects the yaw angular velocity R as the input, multiplies it by the yaw angular velocity gain after passing through a high-pass filter to generate the rudder command δ r ;
[0071] The roll angle loop takes the deviation between the roll angle command γ c and the actual roll angle γ as the input, and estimates the disturbance in real time through the linear extended state observer LESO γ and combines the roll angle loop gain k γ and the system dynamic coefficient to generate the roll angle rate command P c and outputs it to the roll angle rate loop; the roll angle rate loop takes the deviation between the roll angle rate command P c and the actual roll angle rate P as the input, and further estimates the system disturbance through the LESO P and combines the roll angle rate loop gain k P and the dynamic coefficient to generate the aileron command δ a so as to control the attitude angle of the aircraft;
[0072] As Figure 1 shown, the pitch angle control loop uses two second-order linear extended observers to achieve longitudinal pitch angle control. The outer loop uses the pitch angle loop, and outputs the pitch angle rate command to enter the inner loop pitch angle rate loop, and the inner loop pitch angle rate loop outputs the elevator command.
[0073] Among them, the pitch angle loop LESO θ is designed as follows:
[0074]
[0075] In the formula: is the estimated value of the pitch angle, is the estimated value of the total disturbance of the pitch angle by the observer, ω θ is the observation bandwidth of the pitch angle extended state observer, b θ is the pitch angle control input gain.
[0076] The pitch rate command is designed as follows:
[0077]
[0078] In the formula: θ c is the pitch angle command, is the estimated value of the total disturbance of the pitch angle by the observer, k θ is the pitch angle loop gain, b θ is the pitch angle control input gain.
[0079] The pitch rate loop LESO Q is designed as follows:
[0080]
[0081] In the formula: is the estimated value of the pitch rate, is the estimated value of the total disturbance of the pitch rate by the observer, ω Q is the observation bandwidth of the pitch rate extended state observer, b Q is the pitch rate control input gain.
[0082] The elevator command δ e is designed as follows:
[0083]
[0084] In the formula: is the estimated value of the total disturbance of the pitch rate by the observer, k Q is the pitch rate loop gain, b Q is the pitch rate control input gain.
[0085] As Figure 2 shown, the input R of the yaw angle control loop is the yaw angular velocity, which is multiplied by after passing through a high-pass filter to generate the rudder command δ r .
[0086] As Figure 3 shown, the roll angle control loop uses two second-order linear extended observers to achieve roll angle control. The outer loop uses the roll angle loop, and the output roll rate command enters the inner loop roll rate loop, and the inner loop roll rate loop outputs the aileron command.
[0087] Among them, the roll angle loop LESO γ is designed as follows:
[0088]
[0089] Where: is the estimated value of the roll angle, is the estimated value of the total disturbance of the roll angle by the observer, ω γ is the observation bandwidth of the roll angle extended state observer, b γ is the control input gain of the roll angle.
[0090] The roll angle rate command is designed as follows:
[0091]
[0092] Where: γ c is the roll angle command, is the estimated value of the total disturbance of the roll angle by the observer, k γ is the roll angle loop gain, b γ is the control input gain of the roll angle.
[0093] The roll angle rate loop LESO P is designed as follows:
[0094]
[0095] Where: is the estimated value of the roll angle rate, is the estimated value of the total disturbance of the roll angle rate by the observer, ω P is the observation bandwidth of the roll angle rate extended state observer, b P is the control input gain of the roll angle rate.
[0096] The aileron command δ a is designed as follows:
[0097]
[0098] Where: is the estimated value of the total disturbance of the roll angle rate by the observer, k P is the roll angle rate loop gain, b P is the control input gain of the roll angle rate.
[0099] The sink rate controller includes a height control loop and a sink rate control loop. The height control loop uses the deviation between the target height command H c and the actual height H as the input, and generates the sink rate command and output to the sinking rate control loop; the sinking rate control loop sinks at the nominal trajectory sinking rate and the actual sinking rate as the input of the deviation, estimate the system disturbance through the linear extended state observer LESO, and combine the proportional gain and the system dynamic coefficient to generate the pitch angle command θ c so as to control the sinking rate;
[0100] As Figure 4 shown, the sinking rate controller is designed based on the second-order linear extended observer. The outer loop is the height control based on PI, which outputs the lift speed command and serves as the input of the inner loop at the same time. The inner loop is the active disturbance rejection control of the sinking rate based on ESO, which outputs the pitch angle command;
[0101] The sinking rate command is designed as follows:
[0102]
[0103] In the above formula, is the sinking rate command, is the nominal trajectory sinking rate, H c is the desired height, is the height control gain.
[0104] The pitch angle command θ c is designed as shown in the formula:
[0105]
[0106] In the formula, is the control input gain, is the sinking rate control loop gain, is the estimated value of the total uncertainty of the lift speed by the lift speed LESO, is the desired sinking rate, is the sinking rate.
[0107] The speed control loop of the speed controller takes the target speed command V c and the actual speed V as the input, generates the speed deviation amount, and after proportional-integral operation, combines the spoiler neutral position deflection angle to output the spoiler command so as to control the speed;
[0108] As Figure 5 shown, the speed controller adopts the traditional PI control, and the spoiler command is as follows:
[0109]
[0110] In the formula, is the spoiler deflection command, is the spoiler neutral position deflection, and V c is the speed command, is the speed control gain.
[0111] The lateral controller adopts the L 1 tracking guidance algorithm. According to the lateral offset d between the current position of the aircraft and the target straight path and the course deviation Δψ between the speed direction and the target straight direction, it outputs the roll angle command γ c so as to correct the course and the offset range.
[0112] When the aircraft lands, it aligns with the runway centerline to prevent deviation from the runway. Therefore, the lateral guidance in the landing section adopts the straight line L1 tracking guidance algorithm. As Figure 5 shown, the calculation process of the roll angle command σ cmd is as follows:
[0113] Design the lateral overload command as:
[0114]
[0115] Assume that η is a value with a very small magnitude, and it can be approximately obtained that:
[0116] sinη≈η = η 1 +Δψ
[0117]
[0118] The lateral acceleration command can be obtained as:
[0119]
[0120] Also, since γ c = gtanθ k , therefore, when the trajectory angle θ k is small, the roll angle command of the aircraft can be approximately expressed as:
[0121]
[0122] In the formula, the L 1 is the distance from the aircraft to the virtual reference point. The lateral offset d is specifically the vertical distance from the aircraft to the runway center line, and the course deviation Δψ is specifically the angle between the speed direction and the runway landing direction.
[0123] Those skilled in the art can easily understand that the above are only the preferred embodiments of the present invention and are not intended to limit the present invention. Any modifications, equivalent replacements, and improvements made within the spirit and principles of the present invention shall be included within the protection scope of the present invention.
Claims
1. An autonomous control method for fixed-point landing of a high-speed aircraft without power return, characterized in that: Including attitude angle controller, sink rate controller, speed controller, lateral controller; The attitude angle controller includes a pitch angle control circuit, a yaw angle control circuit, and a roll angle control circuit; The pitch angle loop uses the pitch angle command θ c The deviation from the actual pitch angle θ is used as input and the linear extended state observer LESO is used θ Real-time estimation of interference, combined with the pitch angle loop gain k θ And the system dynamic coefficient generates the pitch angle rate command Q c And output to the pitch rate loop; the pitch rate loop uses the pitch rate instruction Q c The deviation from the actual pitch rate Q is input through LESO Q Further estimate the system interference and combine the pitch rate loop gain k Q and the dynamic coefficient to generate the elevator command δ e Thereby controlling the pitch angle of the aircraft; The yaw angle loop selects the yaw angular velocity R input, which is multiplied by the yaw angular velocity gain after passing through a high-pass filter. Generate rudder command δ r ; The roll angle loop uses the roll angle command γ c The deviation from the actual roll angle γ is used as input and the linear extended state observer LESO is used. γ The disturbance is estimated in real time and combined with the roll angle loop gain k γ The roll angle rate command P is generated by the system dynamic coefficient c And output to the roll angular rate loop; the roll angular rate loop uses the roll angular rate command P c The deviation from the actual roll angular rate P is input and is calculated by LESO P Further estimate the system disturbance and combine it with the roll rate loop gain k P and the dynamic coefficient to generate the aileron command δ a Thereby controlling the roll angle of the aircraft; The sinking rate controller includes a height control loop and a sinking rate control loop. The height control loop uses a target height instruction H c The deviation from the actual height H is used as input to generate the sinking rate command through the proportional-integral controller And output to the sinking rate control loop; the sinking rate control loop uses the nominal trajectory sinking rate and actual sinking rate The deviation is taken as input, and the system disturbance is estimated by the linear extended state observer LESO, combined with the proportional gain and system dynamic coefficient Generate pitch angle command θ c Thus controlling the sinking rate; The speed control loop of the speed controller uses the target speed command V c The actual speed V is used as input to generate the speed deviation, which is combined with the speed brake neutral position angle after proportional-integral operation. Output speed brake command Thus controlling the speed; The lateral controller uses the L1 tracking guidance algorithm to output the roll angle command γ according to the lateral deviation d between the current position of the aircraft and the target straight path and the heading deviation Δψ between the speed direction and the target straight direction. c Thereby correcting the heading and deviation distance.
2. The autonomous control method for fixed-point landing of a high-speed aircraft without power return according to claim 1 is characterized in that: The pitch angle loop LESO θ The design is as follows: Where: is the estimated value of the pitch angle, is the observer's estimate of the total disturbance of the pitch angle, ω θ is the observation bandwidth of the pitch angle expansion state observer, b θ Input gain for pitch angle control. The pitch rate instruction is designed as follows: Where: θ c is the pitch angle command, is the observer’s estimate of the total disturbance of the pitch angle, k θ is the pitch angle loop gain, b θ Input gain for pitch angle control.
3. The autonomous control method for fixed-point landing of a high-speed aircraft without power return according to claim 1, characterized in that: The pitch rate loop LESO Q The design is as follows: Where: is the estimated value of the pitch angular rate, is the observer's estimate of the total disturbance of the pitch rate, ω Q is the observation bandwidth of the pitch rate expansion state observer, b Q Input gain for pitch rate control. The elevator command δ e The design is as follows: Where: is the observer’s estimate of the total disturbance of the pitch rate, k Q is the pitch rate loop gain, b Q Input gain for pitch rate control.
4. The autonomous control method for fixed-point landing of a high-speed aircraft without power return according to claim 1, characterized in that: The aileron command δ r The design is as follows:
5. The autonomous control method for fixed-point landing of a high-speed aircraft without power return according to claim 1, characterized in that: The roll angle loop LESO γ The design is as follows: Where: is the estimated value of the roll angle, is the observer's estimate of the total disturbance of the roll angle, ω γ is the observation bandwidth of the roll angle expansion state observer, b γ Input gain for roll angle control. The roll angle rate command is designed as follows: Where: γ c is the roll angle command, is the observer's estimate of the total disturbance of the roll angle, k γ is the roll angle loop gain, b γ Input gain for roll angle control.
6. The autonomous control method for fixed-point landing of a high-speed aircraft without power return according to claim 5 is characterized in that: The roll rate loop LESO P The design is as follows: Where: is the estimated value of the roll angular rate, is the observer's estimate of the total disturbance of the roll angular rate, ω P is the observation bandwidth of the roll rate extended state observer, b P Input gain for roll rate control. The aileron command δ a The design is as follows: Where: is the observer's estimate of the total disturbance of the rolling angular rate, k P is the roll rate loop gain, b P Input gain for roll rate control.
7. The autonomous control method for fixed-point landing of a high-speed aircraft without power return according to claim 1, characterized in that: The sink rate command The design is as follows: In the above formula, is the sinking rate command, is the nominal trajectory sinking rate, H c is the expected height, For height control gain. The pitch angle command θ c The design is shown in the formula: In the formula, To control the input gain, is the sink rate control loop gain, is the estimate of the total uncertainty of the vertical velocity LESO, is the expected sinking rate, is the sinking rate.
8. The autonomous control method for fixed-point landing of a high-speed aircraft without power return according to claim 1, characterized in that: The speed brake command is as follows: In the formula, is the speed brake deflection command, is the deflection angle of the speed brake in neutral position, V c is the speed command, is the speed control gain.
9. The autonomous control method for fixed-point landing of a high-speed aircraft without power return according to claim 1, characterized in that: The roll angle command γ c The calculation process includes the following steps: Step 1: Design the lateral overload instructions as: η is a very small value, which can be approximated as: sinη≈η=η1+Δψ Step 2: Design the lateral acceleration command as: Step 3: Since γ c =gtanθ k , so at the trajectory angle θ k When it is small, the aircraft roll angle command can be approximately expressed as: The L1 is the distance from the aircraft to the virtual reference point, the side offset d is specifically the vertical distance from the aircraft to the centerline of the runway, and the heading deviation Δψ is specifically the angle between the speed direction and the runway landing direction.
10. An autonomous control system for fixed-point landing of a high-speed aircraft without power return, storing a software program corresponding to the control method described in any one of claims 1 to 9.