Method for implementing variable-rate rotary motion of a tethered support aircraft model

Abstract

The application relates to a method for realizing variable-rate rotary motion of a tethered support airplane model, belonging to the technical field of wind tunnel testing. First, an 8-tether redundant constraint tethered support airplane model is adopted; second, in a three-dimensional rotary space, a relationship between a roll angle velocity and a pitch angle velocity is established, so that a angle of attack change rate is zero; third, a relationship between the roll angle velocity and a yaw angle velocity is further established, so that a side slip angle change rate is zero; subsequently, given an initial angle of attack and an initial side slip angle, a three-rotary-degree-of-freedom fixed coupling ratio motion is completed; then, three rotary angle velocity components around the body axis are synthesized to obtain a total rotary angle velocity, that is, an angle velocity of variable-rate rotation around a velocity axis; finally, according to a kinematic relationship, inverse solution of the airplane model position is obtained to obtain a tether length, which is used for motion control, so as to realize variable-rate rotary motion of the airplane model around the velocity axis. The method can realize rotary motion of the tethered support airplane model.

Description

TECHNICAL FIELD

[0001] The present application belongs to the technical field of wind tunnel test, and particularly relates to a method for realizing variable-rate rotation of a rope-supported aircraft model, which is used in a wind tunnel spin aerodynamic characteristic test. BACKGROUND

[0002] The spin of a high-performance fighter aircraft in a large angle of attack flight is very complex and has always seriously endangered flight safety. Therefore, a large amount of research work on fighter aircraft spin has been carried out at home and abroad; among them, the aerodynamic characteristics of the variable-rate rotation of an aircraft around the speed axis is an important research target.

[0003] The rotating balance test device is an important basic equipment developed for the research of aircraft spin characteristics, and is mainly used for measuring the aerodynamic characteristics of an aircraft model in an equal-speed rotation state at different rates around the speed axis, so as to provide necessary aerodynamic data for the analysis and prediction of aircraft spin characteristics. Abroad, represented by NASA, a plurality of rotating balance test devices have been developed; in China, the China Aerodynamics Research and Development Center and the China Aerodynamics Research Institute have also developed rotating balance test devices respectively. For example, in a method for improving the accuracy of a rotating balance test (CN202010982702.X), a typical rotating balance test device is given, in which a model is fixed on a sliding block through a support rod, the support rod can rotate relative to the sliding block to change the initial roll angle of the model, the sliding block slides along a curved rail to change the initial pitch angle of the model, the curved rail is fixed on a main shaft, and the variable-rate rotation of the model around the speed axis is realized by rotating the main shaft. However, in the operation of this device, the support rod and the sliding block are locked respectively, that is, only constant angle of attack and constant side slip angle experiments can be carried out; and the test device occupies a huge space and has high energy consumption.

[0004] The two-degree-of-freedom dynamic test bed is a kind of dynamic test platform designed by Nanjing University of Aeronautics and Astronautics, which can not only simulate the dynamic flight of an aircraft more accurately, but also make the simulated motion relatively simple and can independently analyze the dynamic parameters of each flight state. The test bed can perform dynamic tests of pitch-rolling coupling motion and yaw-rolling coupling motion at different angles of attack. The aircraft model is installed in a tail support manner, and its motion in two degrees of freedom is controlled by two independent hydraulic systems. In the lateral and longitudinal coupling motion, the condition of the variable-rate rotation of the aircraft model around the speed axis can be obtained by decomposing the rotation vector, that is, the side slip angle rate is zero, so as to obtain the fixed relationship between the roll angle speed and the yaw angle speed. However, this device can at most realize a certain range of motion in two degrees of freedom, and cannot completely simulate the changes of three attitude angles in the spin. SUMMARY

[0005] The application aims to provide a method for realizing variable-rate rotary motion of a tethered support aircraft model.

[0006] To achieve the above-mentioned purpose, the application comprises the following steps:

[0007] 1) using an 8-rope redundant constraint tethered support system to support the aircraft model;

[0008] 2) in a three-dimensional rotary space, the relationship between the roll angular velocity p and the pitch angular velocity q is constructed so that the component Ω of the total rotary angular velocity Ω in the y direction of the velocity coordinate system is zero, i.e., the change rate of the angle of attack yw is zero;

[0009] 3) the relationship between the roll angular velocity p and the yaw angular velocity r is constructed so that the component Ω of the total rotary angular velocity Ω in the z direction of the velocity coordinate system is zero, i.e., the change rate of the side slip angle zw is zero;

[0010] 4) given the initial angle of attack α and the initial side slip angle β, the constant coupling ratio motion of the three rotary degrees of freedom is completed;

[0011] 5) the three angular velocity components p, q, and r of the rotation around the body axis are synthesized to obtain the total rotary angular velocity Ω, i.e., the angular velocity ω of the variable-rate rotation around the velocity axis;

[0012] 6) according to the kinematic relationship, the inverse solution of the pose of the aircraft model is obtained to obtain the rope length, which is used for motion control to realize the variable-rate rotary motion of the aircraft model around the velocity axis.

[0013] In step 1), the 8-rope redundant constraint tethered support system specifically comprises: four upper ropes symmetrically arranged in front and back on both sides of the aircraft model, and four lower ropes symmetrically arranged in front and back on both sides of the aircraft model; the ropes are in a tension state and do not have virtual traction, i.e., the rope tension T>0; due to the small disturbance of the tethered support to the flow field and the much smaller volume and energy consumption of the support system than the rotary balance, the 6-degree-of-freedom coupled motion is realized, and the support system is selected as the support mode for simulating the variable-rate rotary motion of the aircraft around the velocity axis.

[0014] In step 2), the relationship between the roll angular velocity p and the pitch angular velocity q adopts the following expression:

[0015]

[0016] ​​wherein a is the angle of attack of the aircraft model, β is the sideslip angle of the aircraft model, is the rate of change of the angle of attack of the aircraft model, Ω is the total angular velocity of the aircraft model, Ω yw is the component of Ω orthogonally projected onto the y direction of the velocity coordinate system, p is the roll angular velocity of the aircraft model, q is the pitch angular velocity of the aircraft model; simplifying, we have:

[0017] q = p tan β.

[0018] In step 3), the relationship between the roll angular velocity p and the yaw angular velocity r is expressed as follows:

[0019]

[0020] wherein, is the rate of change of the angle of attack of the aircraft model, Ω is the total angular velocity of the aircraft model, Ω zw is the component of Ω orthogonally projected onto the z direction of the velocity coordinate system, r is the yaw angular velocity of the aircraft model; simplifying, we have:

[0021] r = p tan a (1 - tan β).

[0022] In step 4), the defined motion of the constant coupling ratio is expressed as follows:

[0023]

[0024] wherein, η rp is the yaw-roll coupling ratio, η qp is the pitch-roll coupling ratio, | | represents the absolute value.

[0025] In step 5), the synthesis formula of the angular velocity ω is expressed as follows:

[0026] ω = Ω = Ω xw = p cos a cos β + q cos a sin β + r cos β sin a

[0027] wherein, Ω xw is the component of Ω orthogonally projected onto the x direction of the velocity coordinate system.

[0028] In step 6), the kinematic relationship and inverse solution are expressed as follows:

[0029]

[0030] wherein, L i is the vector of the i-th rope, L i is the length of the i-th rope, B i is the vector of the i-th rope and the pulley connection point in the ground coordinate system, X Pis the position vector of the center of mass P of the aircraft model in the ground coordinate system, r i is the position vector of the i-th rope and the traction point P of the aircraft model i is the position vector in the body coordinate system, ( ) T denotes the transpose of the matrix, and R is the rotation transformation matrix from the body coordinate system to the ground coordinate system;

[0031]

[0032] wherein φ, θ, and ψ are the attitude angles of the aircraft model, representing the roll angle, the pitch angle, and the yaw angle of the aircraft model, respectively, and the values are determined by the following formula:

[0033]

[0034] Compared with the prior art, the beneficial effects of the present application are as follows:

[0035] The method for realizing variable-rate rotary motion of the tethered support aircraft model provided by the present application can simulate the variable-rate rotation of the aircraft around the speed axis while reducing the interference of the support system on the flow field around the aircraft model, and the support system used in the method occupies much smaller volume and consumes much less energy than the rotary balance test device. By using the method, wind tunnel variable-rate rotary motion test of the aircraft model around the speed axis can be performed to study the aerodynamic characteristics of the aircraft model in the spin state. BRIEF DESCRIPTION OF DRAWINGS

[0036] Figure 1 is the overall flowchart of the embodiment of the present application;

[0037] Figure 2 is the schematic diagram of the tethered support system of the embodiment of the present application;

[0038] Figure 3 is the schematic diagram of the three-dimensional rotation space of the embodiment of the present application;

[0039] Figure 4 is the kinematic schematic diagram of the tethered support system of the embodiment of the present application;

[0040] Figure 5 is the angular velocity curve of the variable-rate rotation around the speed axis of the embodiment of the present application;

[0041] Figure 6 is the attitude angle tracking curve and the position stability curve of the embodiment of the present application;

[0042] Figure 7 is the rope length change curve of the head-up oscillation motion of the embodiment of the present application;

[0043] Figure 8 is the head-up oscillation motion experiment diagram of the embodiment of the present application. DETAILED DESCRIPTION

[0044] The technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments of the present application. Based on the embodiments in the present application, all the other embodiments obtained by a person of ordinary skill in the art without creative effort belong to the scope of the present application.

[0045] The purpose of the present application is to provide a method for realizing variable-rate rotation motion of a tethered support aircraft model.

[0046] In order to make the above-mentioned purposes, features and advantages of the present application more obvious and easy to understand, the present application will be further described in detail below with reference to the drawings and specific embodiments.

[0047] Figure 1 The present application is a general flowchart of the embodiments. As shown in the figure, a method for realizing variable-rate rotation motion of a tethered support aircraft model includes: Figure 1

[0048] 1) An 8-tether redundant constraint tethered support aircraft model is adopted;

[0049] 2) In a three-dimensional rotation space, the relationship between roll angular velocity p and pitch angular velocity q is constructed, so that the component Ω yw of total rotation angular velocity Ω in the y direction of the velocity coordinate system is zero, that is, the change rate of attack angle is zero;

[0050] 3) The relationship between roll angular velocity p and yaw angular velocity r is constructed, so that the component Ω zw of total rotation angular velocity Ω in the z direction of the velocity coordinate system is zero, that is, the change rate of side slip angle is zero;

[0051] 4) Given the initial attack angle α and the initial side slip angle β, the constant coupling ratio motion of the three rotational degrees of freedom is completed;

[0052] 5) The three angular velocity components p, q, r of rotation around the body axis are synthesized to obtain the total rotation angular velocity Ω, that is, the angular velocity ω of variable-rate rotation around the velocity axis;

[0053] 6) According to the kinematic relationship, the inverse solution is obtained from the pose of the aircraft model, and the tether length is obtained for motion control, so as to realize the variable-rate rotation motion of the aircraft model around the velocity axis.

[0054] The specific steps are as follows:

[0055] 1) Figure 2 The present application is a schematic diagram of a tethered support system. An 8-tether redundant constraint tether is adopted as shown in the figure to support the aircraft model. As shown in the figure, Figure 2 Figure 2 ​​As shown, OXYZ is the ground coordinate system, Pxyz is the body coordinate system, and the 8-rope redundant constraint rope support system has four ropes symmetrically arranged on the front and back sides of the aircraft model at the top, and four ropes symmetrically arranged on the front and back sides of the aircraft model at the bottom, starting from B. i Point P is drawn out and connected to the aircraft model. i At the point; the ropes are all in a taut state, with no loose tension, that is, the rope tension T > 0.

[0056] 2) Figure 3 This is a schematic diagram of the three-dimensional rotation space according to an embodiment of the present invention. Figure 3 Within the three-dimensional rotational space shown, establish the body coordinate system Ox. b y b z b Stable coordinate system Ox s y s z s and velocity coordinate system Ox w y w z w V represents the incoming flow velocity in the wind tunnel, α is the angle of attack of the aircraft model, β is the sideslip angle of the aircraft model, p represents the roll rate of the aircraft model, q represents the pitch rate of the aircraft model, r represents the yaw rate of the aircraft model, and ω represents the angular velocity of the aircraft model's variable-rate rotation about its velocity axis. (Direction and Oy) w (Same axis) indicates the rate of change of the angle of attack of the aircraft model. (Direction and Oy) w (Same axis) indicates the rate of change of sideslip angle of the aircraft model.

[0057] The total rotational angular velocity Ω of the aircraft model is orthogonally projected onto Oy w On the axis, and set it to zero:

[0058]

[0059] Among them, Ω yw Orthogonally project Ω onto Oy w The component on the axis, this formula ensures that the angle of attack α of the aircraft model remains constant. Simplifying the above formula, we can obtain the relationship between the roll angular velocity p and the pitch angular velocity q:

[0060] q = ptanβ

[0061] 3) Figure 3 This is a schematic diagram of the three-dimensional rotation space according to an embodiment of the present invention. Figure 3 Within the three-dimensional rotational space shown, the total rotational angular velocity Ω of the aircraft model is orthogonally projected onto Oz. w On the axis, and set it to zero:

[0062]

[0063] Among them, Ω zw Orthogonally project Ω onto Oz w The component on the axis, this formula ensures that the sideslip angle β of the aircraft model remains constant. Simplifying the above formula, we can obtain the relationship between the roll rate p and the yaw rate r:

[0064] r = ptanα(1-tanβ)

[0065] 4) Figure 3 This is a schematic diagram of the three-dimensional rotation space according to an embodiment of the present invention. Figure 3 Within the three-dimensional rotational space shown, steps 2) and 3) constrain the aircraft model's angle of attack α and sideslip angle β to remain constant. The initial angle of attack α = ±15° (corresponding to the aircraft model's nose-up and nose-down states, respectively) and initial sideslip angle β = 0 are given. Let the aircraft model's roll velocity be:

[0066] p = 2πfAcos(2πft)

[0067] Where A is the amplitude, f is the oscillation frequency, and t represents the motion time, the roll angle approximates a sinusoidal oscillation. Let A = 45°, f = 0.5Hz, then:

[0068]

[0069] From step 2), we can obtain:

[0070] q = 0

[0071] From step 3), we can obtain:

[0072]

[0073] Therefore, the coupling ratio is:

[0074]

[0075] Where, η rp η is the yaw-roll coupling ratio. qp Let || represent the pitch-roll coupling ratio, and || represent the absolute value. The coupling ratio remains constant. Therefore, the initial attitude angle and angular velocities along the three body axes can be determined, enabling constant coupling ratio motion for the three rotational degrees of freedom.

[0076] 5) Figure 3 This is a schematic diagram of the three-dimensional rotation space according to an embodiment of the present invention. Figure 3 Within the three-dimensional rotation space shown, p, q, r are placed in Ox w Adding the components along the axis yields the total rotational angular velocity Ω of the aircraft model, orthogonally projected onto Oz. w Components on the axis Ω xw :

[0077]

[0078] From steps 2) and 3), we obtain:

[0079] Ω yw =Ω zw =0

[0080] Therefore, the total rotational angular velocity Ω of the aircraft model is orthogonally projected onto Oz. w Components on the axis Ω xw This refers to the total rotational angular velocity Ω itself, which is also the angular velocity ω of the aircraft model's variable-rate rotational motion around the velocity axis.

[0081]

[0082] The results are as follows Figure 4 As shown.

[0083] 6) Figure 5 This is a kinematic diagram of the rope support system according to an embodiment of the present invention. Figure 5 In the rope-supported system shown, the ground coordinate system OXYZ and the body coordinate system Pxyz are defined as follows: Figure 2 The same. The kinematic relationship is as follows:

[0084]

[0085] Among them, L i The vector of the i-th rope L i Let B be the length of the i-th rope. i Let be the vector of the connection point between the i-th rope and the pulley in the ground coordinate system. X P Let P be the position vector of the centroid P of the aircraft model in the ground coordinate system. r i Let P be the traction point between the i-th rope and the airplane model. i Position vector in the body coordinate system () T Represents the transpose of the matrix, where R is the rotation transformation matrix from the body coordinate system to the ground coordinate system;

[0086]

[0087] Where φ, θ, and ψ represent the roll, pitch, and yaw angles of the aircraft model, respectively, and their values ​​can be determined by the angle of attack α and sideslip angle β. The conversion relationship between attitude angles φ, θ, and ψ and airflow angles α and β is as follows:

[0088]

[0089] The initial values of the attitude angles φ, θ, ψ are respectively:

[0090]

[0091] According to the kinematic relationship, the reversible solution obtains the rope length L i , which is used for motion control to realize the variable rate rotation motion of the aircraft model around the speed axis.

[0092] Figure 6 The attitude angle tracking curve and the position stability curve diagram of the embodiment of the present application are shown in the following figure. Figure 6 It can be seen from the attitude angle tracking curve and the position stability curve diagram that in the roll direction, the absolute error of trajectory tracking is not more than 0.9°, and the relative error is not more than 7%; in the pitch direction, the absolute error of trajectory tracking is not more than 0.1°, and the relative error is not more than 0.8%; in the yaw direction, the absolute error of trajectory tracking is not more than 0.3°, and the relative error is not more than 4%; in the three translation directions, the changes of the center of mass of the aircraft model tend to 0. In summary, the experiment using the method provided by the present application is continuous and smooth in the whole motion process.

[0093] Figure 7 The rope length change curve diagram of the head-up oscillation motion of the embodiment of the present application is shown in the following figure. Figure 8 It can be seen from the rope length change curve diagram of the head-up oscillation motion that the change of the rope length is continuous and has no mutation when the experiment using the method provided by the present application is carried out, which proves that the control performance of the controller is good.

[0094] Figure 8 The head-up oscillation motion experiment diagram of the embodiment of the present application is shown in the following figure. Figure 8 It can be seen from the head-up oscillation motion experiment diagram that the method provided by the present application can well realize the motion instruction of the aircraft model.

[0095] The principles and implementation manners of the present application are described by using specific examples, and the above embodiments are only used to help understand the method of the present application and its core idea; for those skilled in the art, according to the idea of the present application, the specific implementation manners and application ranges will be changed. In summary, the content of the specification should not be understood as a limitation of the present application.

Claims

1. A method for implementing variable rate rotational motion of a tethered support aircraft model, comprising: The method comprises the following steps: 1) using an 8-cable redundant constraint tethered-cable supported aircraft model; 2) In the three-dimensional rotation space, by constructing the relationship between the roll angular velocity p and the pitch angular velocity q, q = ptanβ, so that the component of the total rotation angular velocity Ω in the y direction of the velocity coordinate system is zero, i.e. the rate of change of the angle of attack yw is zero, specifically including: the relationship between the roll angular velocity p and the pitch angular velocity q adopts the following expression:​ Wherein, a is the angle of attack of the aircraft model, β is the sideslip angle of the aircraft model, is the angle of attack of the aircraft model, Ω is the total angular velocity of the aircraft model, Ω yw is the component of Ω orthogonally projected onto the y direction of the velocity coordinate system, p is the roll angular velocity of the aircraft model, q is the pitch angular velocity of the aircraft model; 3) the relationship between roll angular velocity p and yaw angular velocity r is constructed such that the component of Ω in the z direction of the velocity coordinate system Ω zw is zero, i.e. the rate of change of sideslip angle is zero; 4) given initial angle of attack α and initial sideslip angle β, completing the constant-coupling ratio motion of three rotational degrees of freedom; 5) synthesizing three angular velocity components p, q, r rotating around the body axis to obtain total rotational angular velocity Ω, i.e. angular velocity ω rotating around the velocity axis at a variable rate; 6) according to the kinematic relationship, inversely solving the pose of the aircraft model from the tether length to obtain the tether length, for motion control, and realizing the variable-rate rotation motion of the aircraft model around the velocity axis.

2. The method of claim 1, wherein the variable rate rotational motion of the tethered support model aircraft is achieved by, The relationship between roll angular velocity p and yaw angular velocity r is established as r = p tan a (1 - tan β) so that the component of Ω in the z direction of the velocity coordinate system Ωz is zero, i.e., the side slip angle change rate zw is zero, specifically including: the relationship between roll angular velocity p and yaw angular velocity r adopts the following expression: is zero, specifically including: the relationship between roll angular velocity p and yaw angular velocity r adopts the following expression: wherein is the rate of change of the angle of attack of the aircraft model, Ω zw is the component of Ω orthogonal projected onto the z direction of the velocity coordinate system, r is the yaw angular velocity of the aircraft model.

3. The method of claim 1, wherein the variable rate rotational motion of the tethered support model aircraft is achieved by, Given initial angle of attack α and initial sideslip angle β, completing the constant-coupling ratio motion of three rotational degrees of freedom, specifically comprising: where η rp is the yaw-roll coupling ratio, η qp is the pitch-roll coupling ratio, and | | denotes the absolute value.

4. The method of claim 1, wherein the variable rate rotational motion of the tethered support model aircraft is achieved by, Synthesizing three angular velocity components p, q, r rotating around the body axis to obtain total rotational angular velocity Ω, i.e. angular velocity ω rotating around the velocity axis at a variable rate, specifically comprising: the angular velocity synthesis formula adopts the following expression: ω = Ω = Ω xw = p cos α cos β + q cos α sin β + r cos β sin α where Ω xw is the component of Ω orthogonally projected onto the x-direction of the velocity coordinate system.

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