A flip and end control semi-physical simulation test method

Abstract

This invention relates to a hardware-in-the-loop (HIL) simulation method for rollover and terminal control. The method designs corresponding control logic based on the control requirements of the aircraft at different stages. In the initial launch phase, the heading angle, pitch angle, heading line-of-sight angle, and pitch line-of-sight angle are inertially smoothed. In the terminal guidance phase, the roll angle is exponentially smoothed according to a certain rule based on the missile-target distance. In the terminal attack phase, the heading and pitch line-of-sight angles are reassigned based on the changes in these angles to control the heading and pitch line-of-sight angles of the five-axis turntable. This invention effectively solves the problem in hardware-in-the-loop simulation experiments where the aircraft cannot perform rollover maneuvers and terminal attack simulations due to the limitations of the five-axis turntable's roll frame and the acceleration limitations of each frame angle.

Description

Technical Field

[0001] This invention belongs to the field of aircraft ground test technology. It is a semi-physical simulation test method for aircraft rollover and terminal control. The contribution of this invention to the prior art is to solve the limitation of the five-axis turntable during the rollover of a certain type of aircraft and the simulation interruption caused by the angular acceleration exceeding the limit during the terminal attack. Background Technology

[0002] During the hardware-in-the-loop simulation test, the simulator, which calculates the aircraft's trajectory in real time, controls the five-axis turntable via a real-time fiber optic network. The five-axis turntable simulates the aircraft's three attitude angles and the line-of-sight angle relative to the target. A certain type of aircraft is mounted upside down on the carrier aircraft; therefore, during the hardware-in-the-loop simulation test, there is a 180° roll maneuver after launch. The roll frame of the five-axis turntable, which houses the detector and target simulator, cannot complete this maneuver due to limitations. Simultaneously, the five-axis turntable has an angular acceleration over-limit protection mechanism; during the terminal attack phase of this type of aircraft, the turntable malfunctions due to excessive angular acceleration, preventing the completion of the hardware-in-the-loop simulation test.

[0003] Currently, there are no roll maneuvers in the aircraft under development, so existing simulation technology cannot meet the requirements of model development. Moreover, during terminal attack tests in hardware-in-the-loop simulation, the five-axis turntable often malfunctions due to excessive angular acceleration, causing hardware-in-the-loop simulation tests to be unable to proceed smoothly and seriously affecting the model development cycle. Summary of the Invention

[0004] Purpose of the invention: This invention provides a hardware-in-the-loop simulation test method for flipping and end-of-pipe control. Its purpose is to effectively solve the problem that the aircraft cannot perform flipping maneuvers and end-of-pipe attack simulations due to the limitation of the five-axis turntable roll frame and the acceleration limitation of each frame angle in hardware-in-the-loop simulation test.

[0005] Technical solution: A hardware-in-the-loop simulation test method for flipping and end control, wherein the test method designs corresponding control logic according to the control requirements of different stages of the aircraft;

[0006] During the initial launch phase, the heading angle, pitch angle, heading line-of-sight angle, and pitch line-of-sight angle are subjected to inertial smoothing.

[0007] During the terminal guidance phase, the roll angle is exponentially smoothed according to a certain rule based on the distance between the missile and the target.

[0008] During the terminal attack phase, the heading and pitch line-of-sight angles are reassigned based on the changes in these angles to control the heading and pitch line-of-sight angles of the five-axis turntable.

[0009] Furthermore, the inertial smoothing process for the heading angle, pitch angle, heading line-of-sight angle, and pitch line-of-sight angle during the initial launch phase is as follows:

[0010] Ktheta=(1-e -T / 3 )·Ktheta1

[0011] psi=(1-e -T / 3 )·psi1

[0012] qb=(1-e -T / 3 )·qb1

[0013] qc=(1-e -T / 3 )·qc1

[0014] Wherein, Ktheta is the pitch angle control of the five-axis turntable; Ktheta1 is the actual pitch angle of the aircraft during flight; psi is the yaw angle control of the five-axis turntable; psi1 is the actual yaw angle of the aircraft during flight; qb is the pitch line-of-sight angle control of the five-axis turntable; qb1 is the actual pitch line-of-sight angle of the aircraft during flight; qc is the yaw line-of-sight angle control of the five-axis turntable; qc1 is the actual yaw line-of-sight angle of the aircraft during flight.

[0015] Furthermore, in the terminal guidance phase, the roll angle is subjected to exponential smoothing processing according to a certain rule based on the distance between the missile and the target.

[0016]

[0017] Where gamma is the roll angle control of the five-axis turntable; gamma1 is the actual roll angle of the aircraft during flight; and T is the real-time flight time of the aircraft.

[0018] Furthermore, to address the issue of turntable malfunctions due to excessive angular acceleration during terminal attacks, the yaw angle, pitch angle, yaw line-of-sight angle, and pitch line-of-sight angle of the simulator-controlled five-axis turntable are subjected to exponential smooth transitions to prevent angle overshoot.

[0019] Ktheta=(1-e -T / 3 )·Ktheta1

[0020] psi=(1-e -T / 3 )·psi1

[0021] qb=(1-e -T / 3 )·qb1

[0022] qc=(1-e -T / 3 )·qc1

[0023] Wherein, Ktheta is the pitch angle control of the five-axis turntable; Ktheta1 is the actual pitch angle of the aircraft during flight; psi is the yaw angle control of the five-axis turntable; psi1 is the actual yaw angle of the aircraft during flight; qb is the pitch line-of-sight angle control of the five-axis turntable; qb1 is the actual pitch line-of-sight angle of the aircraft during flight; qc is the yaw line-of-sight angle control of the five-axis turntable; qc1 is the actual yaw line-of-sight angle of the aircraft during flight.

[0024] Even after adopting the above experimental methods, occasional over-limits in the heading and pitch line-of-sight angles still caused malfunctions in the five-axis turntable. Therefore, further simulation test methods were studied. The heading and pitch line-of-sight angles without exponential smoothing were taken, qbq = qb, qcq = qc. After exponential smoothing, the changes in the pitch and heading line-of-sight angles before and after processing were calculated as qbc and qcc, respectively. Using the horizontal distance as a criterion, if it was less than 100m, the output values ​​of the heading and pitch line-of-sight angles controlling the five-axis turntable were processed. The changes in the heading and pitch line-of-sight angles were assigned values ​​based on |20*0.001| as the evaluation standard. The specific formulas are shown below:

[0025] qbc = qb - qbq

[0026] qcc = qc - qcq

[0027]

[0028]

[0029] Where qbq is the pitch line-of-sight angle calculated by the aircraft in the previous cycle, and qcq is the heading line-of-sight angle calculated by the aircraft in the previous cycle.

[0030] Beneficial technical effects: The flipping and end-effector control test method overcomes the limitations of existing test techniques in addressing angle and acceleration exceeding limits, thus preventing turntable damage due to malfunctions. The simulation method design eliminates the need for turntable design modifications. In summary, this invention saves manpower and resources, ensuring the safety and reliability of testing while facilitating model development. Attached Figure Description

[0031] Figure 1 This is a flowchart of the roll angle processing of the present invention;

[0032] Figure 2 This is a flowchart illustrating the processing of pitch angle, heading angle, pitch line-of-sight angle, and heading line-of-sight angle in this invention. Detailed Implementation

[0033] The technical solution of the present invention will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of the present invention. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0034] Currently, there are no roll maneuvers in aircraft under development, so existing simulation technology cannot meet the requirements of model development. Moreover, during terminal attacks in hardware-in-the-loop simulation tests, five-axis turntable malfunctions often occur due to excessive angular acceleration, causing hardware-in-the-loop simulation tests to be unable to proceed smoothly and seriously affecting the model development cycle. In view of this, this invention specifically studies the test method and proposes a hardware-in-the-loop simulation test method for roll and terminal control.

[0035] The specific implementation method of the present invention is as follows:

[0036] 1) To address the issue of limited roll axis position of the five-axis turntable during spacecraft rollover, technical research was conducted in conjunction with the working modes of the spacecraft's detectors;

[0037] 2) Conduct technical research on the issue of turntable malfunctions caused by changes in angular velocity of heading angle, pitch angle, heading line-of-sight angle, and pitch line-of-sight angle exceeding turntable limits during terminal attacks of aircraft;

[0038] 3) Draw the corresponding flowchart;

[0039] 4) Write the source code in the software that drives the rotary table;

[0040] 5) Conduct thorough simulation verification of this test method in a hardware-in-the-loop simulation experiment;

[0041] 6) During the verification process, it was found that after the exponential smoothing process, the heading line of sight angle and pitch line of sight angle caused the turntable to malfunction due to excessive changes in angular velocity on certain specific trajectories, thus making simulation impossible.

[0042] 7) Conduct further technical research to address the problems identified in 6).

[0043] 8) Draw the corresponding flowchart and write the source code in the simulator that drives the turntable;

[0044] 9) This test method is fully validated in a hardware-in-the-loop simulation test;

[0045] 10) All flight paths can be successfully completed in the hardware-in-the-loop simulation test. The aircraft can perform roll maneuvers and end control normally in the hardware-in-the-loop simulation test, and the simulation test method has been fully verified.

[0046] The processing steps at each stage are as follows:

[0047] 1) Research on experimental methods to address the limitation of the five-axis turntable in completing semi-physical simulation tests during aircraft roll maneuvers due to platform constraints. In semi-physical simulation tests, the simulator and the five-axis turntable interact via fiber optics to control the turntable's movement. However, during roll maneuvers, the roll axis is prone to angle over-limit failures; therefore, improvements are made to the roll axis control within the simulator. For example... Figure 1 As shown, the time limit T0 is set at the point when the aircraft is 6km away from the target. Before this point, the roll angle is 0°. For the next 40 seconds, an exponential smoothing transition is performed, followed by outputting the calculated roll angle to the simulator. The specific method is as follows. Figure 1 And the following formula:

[0048]

[0049] Where gamma is the roll angle control of the five-axis turntable; gamma1 is the actual roll angle of the aircraft during flight; and T is the real-time flight time of the aircraft.

[0050] Note: Since the spacecraft's detector only intervenes in control when entering terminal guidance, the five-axis turntable roll angle control output of 0 when the spacecraft is less than 6km from the target does not affect the simulation; the time-taking program is executed only once, and then jumps out to execute other programs.

[0051] 2) To address the issue of excessive angular acceleration during terminal attacks, the yaw angle, pitch angle, yaw line-of-sight angle, and pitch line-of-sight angle of the five-axis turntable controlled by the simulator are smoothly transitioned exponentially to prevent angles from exceeding limits.

[0052] Ktheta=(1-e -T / 3 )·Ktheta1

[0053] psi=(1-e -T / 3 )·psi1

[0054] qb=(1-e -T / 3 )·qb1

[0055] qc=(1-e -T / 3 )·qc1

[0056] Wherein, Ktheta is the pitch angle control of the five-axis turntable; Ktheta1 is the actual pitch angle of the aircraft during flight; psi is the yaw angle control of the five-axis turntable; psi1 is the actual yaw angle of the aircraft during flight; qb is the pitch line-of-sight angle control of the five-axis turntable; qb1 is the actual pitch line-of-sight angle of the aircraft during flight; qc is the yaw line-of-sight angle control of the five-axis turntable; qc1 is the actual yaw line-of-sight angle of the aircraft during flight.

[0057] 3) Even after adopting the above testing methods, occasional exceedances of the heading and pitch line-of-sight angles and angular velocities still caused malfunctions of the five-axis turntable. Therefore, further research on simulation testing methods is needed. Figure 2 As shown, the heading and pitch line-of-sight angles without exponential smoothing are taken, qbq = qb, qcq = qc. After exponential smoothing, the changes in the pitch and heading line-of-sight angles before and after processing are calculated as qbc and qcc, respectively. Using horizontal distance as the criterion, if it is less than 100m, the heading and pitch line-of-sight angle outputs of the five-axis control turntable are processed. The changes in the heading and pitch line-of-sight angles are assigned values ​​based on |20*0.001|, as shown in the specific formulas below:

[0058] qbc = qb - qbq

[0059] qcc = qc – qcq

[0060]

[0061]

[0062] Where qbq is the pitch line-of-sight angle calculated by the aircraft in the previous cycle, and qcq is the heading line-of-sight angle calculated by the aircraft in the previous cycle.

[0063] The above specific embodiments or examples are only used to explain the technical solutions of the present invention and are not intended to limit the present application. Parts not described in detail are considered to be conventional technical means or common knowledge in the field. Those skilled in the art should understand that, based on the design concept of the present application, it is possible to make adaptive modifications to the technical solutions described in the foregoing embodiments or to make equivalent substitutions for some or all of the technical features. These modifications or substitutions do not cause the essence of the corresponding technical solutions to deviate from the scope of the technical solutions of the various embodiments of the present invention.

Claims

1. A hardware-in-the-loop simulation test method for flipping and end control, characterized in that, The experimental method designs corresponding control logic based on the control requirements of the aircraft at different stages; During the initial launch phase, the heading angle, pitch angle, heading line-of-sight angle, and pitch line-of-sight angle are subjected to inertial smoothing. During the terminal guidance phase, the roll angle is exponentially smoothed according to a certain rule based on the distance between the missile and the target. During the terminal attack phase, the heading and pitch line of sight angles are reassigned based on the changes in heading and pitch line of sight angles in order to control the heading and pitch line of sight angles of the five-axis turntable. The inertial smoothing process for the heading angle, pitch angle, heading line-of-sight angle, and pitch line-of-sight angle during the initial launch phase is as follows: Ktheta = (1- ) Ktheta1 dogs = (1- ) dogs1 qb =(1- ) QB1 qc =(1- ) qc1 Wherein, Ktheta is the pitch angle control of the five-axis turntable; Ktheta1 is the actual pitch angle of the aircraft during flight; psi is the yaw angle control of the five-axis turntable; psi1 is the actual yaw angle of the aircraft during flight; qb is the pitch line-of-sight angle control of the five-axis turntable; qb1 is the actual pitch line-of-sight angle of the aircraft during flight; qc is the yaw line-of-sight angle control of the five-axis turntable; qc1 is the actual yaw line-of-sight angle of the aircraft during flight. During the terminal guidance phase, the roll angle is exponentially smoothed according to a certain rule based on the distance between the missile and the target. gamma = Where gamma is the roll angle control of the five-axis turntable; gamma1 is the actual roll angle of the aircraft during flight; and T is the real-time flight time of the aircraft. To address the issue of excessive angular acceleration during terminal attacks, the yaw, pitch, yaw line-of-sight, and pitch line-of-sight angles of the simulator-controlled five-axis turntable are smoothly transitioned exponentially to prevent angle overshoot. Smooth transitions are handled according to the following exponential rules: Ktheta = (1- ) Ktheta1 dogs = (1- ) dogs1 qb =(1- ) QB1 qc =(1- ) qc1 Wherein, Ktheta is the pitch angle control of the five-axis turntable; Ktheta1 is the actual pitch angle of the aircraft during flight; psi is the yaw angle control of the five-axis turntable; psi1 is the actual yaw angle of the aircraft during flight; qb is the pitch line-of-sight angle control of the five-axis turntable; qb1 is the actual pitch line-of-sight angle of the aircraft during flight; qc is the yaw line-of-sight angle control of the five-axis turntable; qc1 is the actual yaw line-of-sight angle of the aircraft during flight. Take the unexponentially smoothed heading and pitch line-of-sight angles, qbq=qb, qcq=qc. After exponential smoothing, calculate the changes in the pitch and heading line-of-sight angles before and after processing, respectively: qbc and qcc. Using horizontal distance as a criterion, if it is less than 100m, process the heading and pitch line-of-sight angle outputs of the five-axis control turntable, assigning values ​​to the changes in heading and pitch line-of-sight angles using |20*0.001| as the evaluation standard. Using horizontal distance as a criterion, if it is greater than or equal to 100m, no secondary processing of the heading and pitch line-of-sight angle outputs of the five-axis control turntable is required. The specific formula for assignment is shown below: qbc = qb - qbq qcc = qc - qcq qb = qc = Where qbq is the pitch line-of-sight angle calculated by the aircraft in the previous cycle, and qcq is the heading line-of-sight angle calculated by the aircraft in the previous cycle.

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