CN113607378A - Rope system supporting aircraft model forced free angle motion simulation and suppression method - Google Patents

Rope system supporting aircraft model forced free angle motion simulation and suppression method Download PDF

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CN113607378A
CN113607378A CN202110880458.0A CN202110880458A CN113607378A CN 113607378 A CN113607378 A CN 113607378A CN 202110880458 A CN202110880458 A CN 202110880458A CN 113607378 A CN113607378 A CN 113607378A
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CN113607378B (en
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王晓光
王家骏
吴军
林麒
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Xiamen University
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Abstract

A simulation and suppression method for forced free angular motion of a rope system supporting aircraft model belongs to the technical field of electromechanical control. The method comprises the following steps: 1) a mode of supporting the aircraft model by a 5-rope under-constrained tether support system is adopted; 2) establishing a support system dynamic equation containing a control surface; 3) designing a large attack angle forced pitching motion control law; 4) designing a control surface control law 1, compensating a rope tension constraint term and a Coriolis force term in the rolling and yawing directions, and simulating forced pitching + free rolling/yaw angular motion; 5) the control surface control law 2 is designed to suppress angular motion in the roll/yaw directions. The method can release the degrees of freedom in the roll direction and the yaw direction and inhibit the angular motion in the roll direction and the yaw direction while simulating the large-attack-angle forced angular motion of the aircraft in the pitch direction. The invention can be used for carrying out a wind tunnel forced free angle movement test on the aircraft model and researching the pneumatic/movement/control coupling relation under the large angle of attack.

Description

Rope system supporting aircraft model forced free angle motion simulation and suppression method
Technical Field
The invention belongs to the technical field of electromechanical control, and particularly relates to a simulation and suppression method for forced free angle motion of a rope system supporting aircraft model.
Background
The rope system supporting system uses the rope to replace a traditional rigid rod to control the posture of the end effector, has the advantages of large working space, high load capacity, good dynamic performance and the like, has mature application in the aspects of large radio telescopes, rehabilitation robots and the like, and provides a novel supporting mode for wind tunnel tests. The traditional wind tunnel test adopts hard supports, such as an abdominal support, a back support, a tail support and the like, the inertia force is large, and the support can generate large interference on a flow field near a model; the rope system support has the characteristics of small interference to a flow field and multi-degree-of-freedom movement, and can provide support for the wind tunnel dynamic test support technology.
According to the relationship between the number of ropes m and the degree of freedom n, the rope system supporting system can be divided into three types, namely: under-constraint (m < n +1), full constraint (m ═ n +1), and redundant constraint (m > n + 1). In recent years, two types of rope system supporting systems with redundant constraint and complete constraint are applied to low-speed wind tunnel tests, and corresponding theoretical research and test verification are carried out. However, the mechanism can only drive the aircraft model to do forced motion, and the aerodynamic characteristics under a given motion track are researched. In order to research the large attack angle pneumatic/motion/control coupling relation or the full-mode flutter test, the test closer to the real flight condition needs to be carried out, and the forced free angle motion of the aircraft in a specific direction needs to be simulated.
At present, the forced free angle motion of an aircraft in a specific direction is simulated by combining a bearing with a hard support at home and abroad. For example, in a patent CN200910073391.9 forced pitch-free yaw wind tunnel test device, an aircraft model web strut is supported by a bearing, which can realize forced pitch + free yaw movement at a large attack angle, but the web strut generates large interference to a flow field near the aircraft model; in the hypersonic wind tunnel virtual flight test system and method disclosed in patent CN110207943A, a tail-boom type serial three-degree-of-freedom bearing is arranged in an aircraft model, so that free angular motions of pitch, roll and yaw can be realized, but forced motion cannot be directly simulated, and the existence of a support rod can limit the angular motion range of the aircraft model to a certain extent. In view of the above, no motion simulation related to forced pitch and simultaneous free roll and yaw is known, and therefore, a new large-attack-angle forced free motion simulation and suppression method needs to be designed.
Disclosure of Invention
The invention aims to provide a rope system supporting aircraft model forced free angle motion simulation and inhibition method which can perform a wind tunnel large attack angle forced free angle motion test on an aircraft model and study the pneumatic/motion/control coupling relation under a large attack angle.
The invention comprises the following steps:
1) a mode of supporting the aircraft model by a 5-rope under-constrained tether support system is adopted;
2) establishing a support system dynamic equation containing a control surface;
3) designing a large attack angle forced pitching motion control law;
4) designing a control surface control law 1, compensating a rope tension constraint term and a Coriolis force term in the rolling and yawing directions, and simulating forced pitching + free rolling/yaw angular motion;
5) the control surface control law 2 is designed to suppress angular motion in the roll/yaw directions.
In step 1), the 5-rope under-constrained roping support system comprises 4 upper ropes and 1 lower rope; 4 upper ropes are symmetrically arranged at the front and the back of two sides of the aircraft model, and 1 lower rope is arranged at a position passing through the mass center of the aircraft model; based on the rigidity and natural frequency analysis, the first-order natural frequency of the support system in the rolling and yawing directions is as low as possible, and the basic rolling and yawing frequency is as low as possible through the ropes of the mass center; the ropes are in a tensioning state and are not subjected to virtual traction, namely the rope tension T is greater than 0; due to the incomplete restraint of the under-constrained tether support, the aircraft model can still move even under the condition that the length of the tether is given and unchanged, and therefore the 5-tether under-constrained tether support system is selected as a support mode for simulating the forced free angular motion of the aircraft.
In step 2), the dynamic equation of the support system with the control surface adopts the following expression:
Figure BDA0003192054290000021
wherein M (X) is an aircraft modelAn inertia matrix, X is an aircraft model pose vector,
Figure BDA0003192054290000022
in order to be the term of the speed,
Figure BDA0003192054290000023
is an acceleration term;
Figure BDA0003192054290000024
is a nonlinear Coriolis centrifugal force matrix, wgIn the form of a gravity vector, the vector,
Figure BDA0003192054290000025
is an aerodynamic/moment vector, w, related to the pose and rate of change of the aircraft modele2(δr,δa) For aerodynamic/moment vectors, δ, related to rudder and aileron deflections of the aircraft modelrRudder deflection angle, delta, for aircraft modelsaIs the aileron deflection angle of the aircraft model,
Figure BDA0003192054290000026
is the rudder yaw moment coefficient,
Figure BDA0003192054290000027
the aileron rolling moment coefficient is shown, T is a rope tension vector, and J is a Jacobian matrix of the system;
Figure BDA0003192054290000028
the first derivative is represented as a function of,
Figure BDA0003192054290000029
represents the second derivative, ()TRepresenting the transpose of the matrix.
In the step 3), the large attack angle forced pitching motion control law can be designed under the condition that the deflection angle of the control surface is set to be zero, and the large attack angle forced pitching motion control law is designed by considering the uncertainty of the internal modeling of the compensation system and the external interference existing in the actual test; the method specifically comprises the following steps: the system comprises a wavelet neural network, an extended state observer and a nonsingular terminal sliding mode controller; the wavelet neural network is used for compensating uncertainty of internal modeling of the system, the extended state observer is used for compensating external interference existing in an actual test, and the nonsingular terminal sliding mode controller is used for enabling tracking errors of a controlled object to be converged; the master control law of the large attack angle forced pitching motion is as follows:
Figure BDA0003192054290000031
wherein T is rope tension, C is equivalent Jacobian matrix, A0Is an equivalent inertia matrix, B0As an equivalent damping matrix, D0Is a matrix of equivalent coriolis force and gravity,
Figure BDA0003192054290000032
in order to be an estimate of the interference term,
Figure BDA0003192054290000033
estimating a weight for the maximum neural network;
Figure BDA0003192054290000034
is a nonsingular terminal sliding mode surface function, beta is more than 0, gamma is more than 1 and less than 2, and e is X-XdAs error, X is the actual pose, XdFor theoretical pose, | | represents an absolute value, sign () represents a sign function; w is a wavelet basis function matrix; d1、D2Setting a gain value only in the pitching direction for a semi-positive fixed gain matrix; u. of0Is a robust term in the control law; ()+Representing a pseudo-inverse of the matrix;
Figure BDA0003192054290000038
representing a matrix multiplication, the operation is as follows:
Figure BDA0003192054290000035
wherein, wiIs the ith column vector of the matrix W.
In step 4), the expression of the control surface control law 1 is as follows:
Figure BDA0003192054290000036
wherein (C)iRepresenting the ith component of the column vector. The control law can compensate a rope tension constraint term and a Coriolis force term in the rolling and yawing directions, and the forced pitching + free rolling/yawing angular motion of the aircraft model under the condition that the rolling and yawing directions are not constrained is simulated.
In step 5), the expression of the control surface control law 2 is as follows:
Figure BDA0003192054290000037
wherein (C)dThe desired value is k, the control gain is k, psi is the yaw angle, r is the yaw angular velocity, phi is the roll angle, and p is the roll angular velocity; at the moment, the large-attack-angle forced pitching motion control law in the step 3) is still adopted in the pitching direction, the influence and the inhibition of the rudder deflection angle on the aerodynamic force and aerodynamic moment of the aircraft model are analyzed, and a control method for inhibiting the angular motion in the rolling/yawing direction based on control surface control is designed; further, the aircraft flies under the attitude with a large attack angle, and asymmetric airflow is generated, so that a horizontal direction moment is generated, and the horizontal direction moment can be compensated by the control surface control law to restrain the angular motion in the rolling/yawing direction.
Compared with the prior art, the invention has the beneficial effects that:
the simulation and inhibition method for the forced free angular motion of the rope system supported aircraft model provided by the invention can release the degrees of freedom in the rolling and yawing directions and inhibit the angular motion in the rolling and yawing directions while simulating the large-attack-angle forced angular motion of the aircraft in the pitching direction. The invention can be used for carrying out a wind tunnel forced free angle movement test on the aircraft model and researching the pneumatic/movement/control coupling relation under the large angle of attack.
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FIG. 1 is a general flow diagram of an embodiment of the present invention;
FIG. 2 is a schematic view of a tether support system according to an embodiment of the present invention;
FIG. 3 is a flow chart of the control of the forced pitching motion at a large angle of attack according to the embodiment of the present invention;
FIG. 4 is a graph of simulated pitch angle tracking for large angle of attack forced pitch motion with interference according to an embodiment of the present invention;
FIG. 5 is a diagram of simulated roll and yaw free motions for a large angle of attack forced pitch motion with disturbance according to an embodiment of the present invention;
FIG. 6 is a diagram of a simulated roll and yaw direction free motion of a large angle of attack forced pitch motion without interference according to an embodiment of the present invention.
Detailed Description
The present invention will be further described with reference to the following examples, which are intended to illustrate only some, but not all, embodiments of the present invention. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
FIG. 1 is a general flow chart of an embodiment of the present invention. As shown in fig. 1, a method for simulating and suppressing a large-attack-angle forced free-angle motion of a rope-supported aircraft model includes:
1) a mode of supporting the aircraft model by 5-rope under-constrained ropes is adopted;
2) establishing a support system dynamic equation containing a control surface;
3) designing a large attack angle forced pitching motion control law;
4) designing a control surface control law 1, compensating a rope tension constraint term and a Coriolis force term in the rolling and yawing directions, and simulating forced pitching + free rolling/yaw angular motion;
5) the control surface control law 2 is designed to suppress angular motion in the roll/yaw directions.
The method comprises the following specific steps:
1) the aircraft model is supported by a 5-rope under-constrained tether support system shown in figure 2. OXYZ is a ground coordinate system, PX ' Y ' Z ' is a body coordinate system, and the position above a 5-rope under-constrained rope system supporting system4 ropes are arranged in front of and behind the two sides of the aircraft model from BiPoint leading-out, connecting to P of aircraft modeliPoint-on; the lower 1 rope is led out from the point O and is arranged at a position passing through the centroid P of the aircraft model; the ropes are all in tension state and no virtual traction occurs, i.e. the rope tension T>0。
2) Establishing a dynamic equation of the support system with the control surface as shown in FIG. 1:
Figure BDA0003192054290000051
wherein M (X) is an inertia matrix of the aircraft model, X is an attitude vector of the aircraft model,
Figure BDA0003192054290000052
in order to be the term of the speed,
Figure BDA0003192054290000053
is an acceleration term;
Figure BDA00031920542900000515
is a nonlinear Coriolis centrifugal force matrix, wgIn the form of a gravity vector, the vector,
Figure BDA0003192054290000054
is an aerodynamic/moment vector, w, related to the pose and rate of change of the aircraft modele2(δr,δa) For aerodynamic/moment vectors, δ, related to rudder and aileron deflections of the aircraft modelrRudder deflection angle, delta, for aircraft modelsaIs the aileron deflection angle of the aircraft model,
Figure BDA0003192054290000055
is the rudder yaw moment coefficient,
Figure BDA0003192054290000056
the aileron rolling moment coefficient is shown, T is a rope tension vector, and J is a Jacobian matrix of the system;
Figure BDA0003192054290000057
the first derivative is represented as a function of,
Figure BDA0003192054290000058
represents the second derivative, ()TRepresenting the transpose of the matrix.
3) The large-attack-angle forced pitching motion control law shown in fig. 3 is designed under the condition that the rudder deflection angle is set to be zero. The wavelet neural network in fig. 3 is used for compensating uncertainty of internal modeling of the system, the extended state observer is used for compensating external interference existing in an actual test, and the nonsingular terminal sliding mode controller is used for converging a tracking error of a controlled object.
Wherein, the total control law of the pitching forced movement is as follows:
Figure BDA0003192054290000059
wherein T is rope tension, C is equivalent Jacobian matrix, A0Is an equivalent inertia matrix, B0As an equivalent damping matrix, D0Is a matrix of equivalent coriolis force and gravity,
Figure BDA00031920542900000510
in order to be an estimate of the interference term,
Figure BDA00031920542900000511
estimating a weight for the maximum neural network;
Figure BDA00031920542900000512
is a nonsingular terminal sliding mode surface function, beta is more than 0, gamma is more than 1 and less than 2, and e is X-XdAs error, X is the actual pose, XdFor theoretical pose, | | represents an absolute value, sign () represents a sign function; w is a wavelet basis function matrix; d1、D2Setting a gain value only in the pitching direction for a semi-positive fixed gain matrix; u. of0Is a robust term in the control law; ()+Representing a pseudo-inverse of the matrix;
Figure BDA00031920542900000516
representing a matrix multiplication, the operation is as follows:
Figure BDA00031920542900000513
wherein, wiIs the ith column vector of the matrix W.
The case that a small disturbance is applied in the rolling direction while the aircraft model is controlled to do large-attack-angle forced pitch angle motion, but the control is not performed in the rolling direction is taken as an example. Wherein the simulation track is set as:
Figure BDA00031920542900000514
the disturbance magnitude is set as: [ 0; 0; 0; 0.1 × cos (t); 0; 0]. Fig. 4 is a graph of simulated pitch angle tracking of large angle of attack forced pitch motion with interference according to the embodiment of the present invention, and the result shows that: the aircraft model keeps higher tracking precision in the forced pitching direction; fig. 5 is a diagram of simulated roll and yaw free motions of large-attack-angle forced pitching motion when there is interference in the embodiment of the present invention, and fig. 6 is a diagram of simulated roll and yaw free motions of large-attack-angle forced pitching motion when there is no interference in the embodiment of the present invention, and comparing the results, it can be known that: the rope system support has small restriction in the rolling and yawing directions and has the condition of simulating free angle motion.
4) Designing a control surface control law 1, compensating a rope tension constraint term and a Coriolis force term in the rolling and yawing directions:
Figure BDA0003192054290000061
wherein (C)iRepresenting the ith component of the column vector. The control law can compensate the rope tension and the Coriolis force term and simulate forced pitching + free rolling/yaw angle motion of an aircraft model under the condition that the rolling and yaw directions are not constrained.
5) Designing a control surface control law 2, and inhibiting the angular motion in the rolling/yawing direction:
Figure BDA0003192054290000062
where k is the PD control gain, psi is the yaw angle, r is the yaw angular velocity, phi is the roll angle, p is the roll angular velocity, ()dIs a desired value. And at the moment, the large attack angle forced pitching motion control law in the step 3) is still adopted in the pitching direction.
The principle and the implementation of the present invention are explained by applying a specific example in the present invention, and the above description of the embodiment is only used to help understanding the method and the core idea of the present invention; meanwhile, for a person skilled in the art, according to the idea of the present invention, the specific embodiments and the application range may be changed. In view of the above, the present disclosure should not be construed as limiting the invention.

Claims (6)

1. A simulation and inhibition method for forced free angular motion of a rope system supporting aircraft model is characterized by comprising the following steps:
1) a mode of supporting the aircraft model by a 5-rope under-constrained tether support system is adopted;
2) establishing a support system dynamic equation containing a control surface;
3) designing a large attack angle forced pitching motion control law;
4) designing a control surface control law 1, compensating a rope tension constraint term and a Coriolis force term in the rolling and yawing directions, and simulating forced pitching + free rolling/yaw angular motion;
5) the control surface control law 2 is designed to suppress angular motion in the roll/yaw directions.
2. The method for simulating and restraining forced free angular motion of a tethered support aircraft model as defined in claim 1 wherein in step 1) said 5-tether under-constrained tether support system comprises 4 upper tethers and 1 lower tether; 4 upper ropes are symmetrically arranged at the front and the back of two sides of the aircraft model, and 1 lower rope is arranged at a position passing through the mass center of the aircraft model; based on the rigidity and natural frequency analysis, the first-order natural frequency of the support system in the rolling and yawing directions is as low as possible, and the basic rolling and yawing frequency is as low as possible through the ropes of the mass center; the ropes are in a tension state, and virtual traction does not occur, namely the rope tension T is greater than 0.
3. The method for simulating and restraining forced free angular motion of a rope-tied support aircraft model according to claim 1, wherein in the step 2), the dynamic equation of the support system with the control surface adopts the following expression:
Figure FDA0003192054280000011
wherein M (X) is an inertia matrix of the aircraft model, X is an attitude vector of the aircraft model,
Figure FDA0003192054280000012
in order to be the term of the speed,
Figure FDA0003192054280000013
is an acceleration term;
Figure FDA0003192054280000017
is a nonlinear Coriolis centrifugal force matrix, wgIn the form of a gravity vector, the vector,
Figure FDA0003192054280000014
is an aerodynamic/moment vector, w, related to the pose and rate of change of the aircraft modele2(δr,δa) For aerodynamic/moment vectors, δ, related to rudder and aileron deflections of the aircraft modelrRudder deflection angle, delta, for aircraft modelsaIs the aileron deflection angle of the aircraft model,
Figure FDA0003192054280000018
is the rudder yaw moment coefficient,
Figure FDA0003192054280000019
the aileron rolling moment coefficient is shown, T is a rope tension vector, and J is a Jacobian matrix of the system;
Figure FDA0003192054280000015
the first derivative is represented as a function of,
Figure FDA0003192054280000016
represents the second derivative, ()TRepresenting the transpose of the matrix.
4. The method for simulating and restraining forced free angular motion of a rope-tied supported aircraft model according to claim 1, wherein in step 3), the large-attack-angle forced pitching motion control law is designed under the condition that the drift angle of a control plane is set to be zero, and the large-attack-angle forced pitching motion control law is designed by considering uncertainty of modeling inside a compensation system and external interference existing in an actual test; the method specifically comprises the following steps: the system comprises a wavelet neural network, an extended state observer and a nonsingular terminal sliding mode controller; the wavelet neural network is used for compensating uncertainty of internal modeling of the system, the extended state observer is used for compensating external interference existing in an actual test, and the nonsingular terminal sliding mode controller is used for enabling tracking errors of a controlled object to be converged; the master control law of the large attack angle forced pitching motion is as follows:
Figure FDA0003192054280000021
wherein T is rope tension, C is equivalent Jacobian matrix, A0Is an equivalent inertia matrix, B0As an equivalent damping matrix, D0Is a matrix of equivalent coriolis force and gravity,
Figure FDA0003192054280000022
in order to be an estimate of the interference term,
Figure FDA0003192054280000023
estimating for maximum neural networkCalculating a weight value;
Figure FDA0003192054280000024
is a nonsingular terminal sliding mode surface function, beta is more than 0, gamma is more than 1 and less than 2, and e is X-XdAs error, X is the actual pose, XdFor theoretical pose, | | represents an absolute value, sign () represents a sign function; w is a wavelet basis function matrix; d1、D2Setting a gain value only in the pitching direction for a semi-positive fixed gain matrix; u. of0Is a robust term in the control law; ()+Representing a pseudo-inverse of the matrix;
Figure FDA0003192054280000028
representing a matrix multiplication, the operation is as follows:
Figure FDA0003192054280000025
wherein, wiIs the ith column vector of the matrix W.
5. The method for simulating and restraining the forced free angular motion of the rope-tied support aircraft model according to claim 1, wherein in the step 4), the expression of the control surface control law 1 is as follows:
Figure FDA0003192054280000026
wherein (C)iRepresents the ith component of the column vector; the control surface control law 1 is used for compensating a rope tension constraint term and a Coriolis force term in the rolling and yawing directions and simulating forced pitching + free rolling/yawing angular motion of an aircraft model under the unconfined conditions of the rolling and yawing directions.
6. The method for simulating and restraining the forced free angular motion of the rope-tied support aircraft model according to claim 1, wherein in the step 5), the expression of the control surface control law 2 is as follows:
Figure FDA0003192054280000027
wherein (C)dThe desired value is k, the control gain is k, psi is the yaw angle, r is the yaw angular velocity, phi is the roll angle, and p is the roll angular velocity; at the moment, a large attack angle forced pitching motion control law is still adopted in the pitching direction, the influence and the inhibition of the rudder deflection angle on the aerodynamic force and aerodynamic moment of the aircraft model are analyzed, and a control method for inhibiting the angular motion in the rolling/yawing direction based on control surface control is designed; the aircraft flies under the attitude with a large attack angle to generate asymmetric airflow, so that when a lateral direction moment is generated, the control surface control law is used for compensation, and the angular motion in the rolling/yawing direction is restrained.
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CN114838905B (en) * 2022-03-23 2023-05-12 厦门大学 Novel dynamic aerodynamic force measurement method for model of tethered parallel support aircraft
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CN114643584B (en) * 2022-05-17 2022-09-30 中国科学技术大学 Rapid terminal sliding mode synchronous control method for rope traction parallel robot

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