CN115290282A - Wind tunnel test method for researching course control efficiency of control surface of flying wing aircraft - Google Patents
Wind tunnel test method for researching course control efficiency of control surface of flying wing aircraft Download PDFInfo
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Abstract
The invention discloses a wind tunnel test method for researching course control efficiency of a control surface of an all-wing aircraft, and belongs to the technical field of low-speed wind tunnel tests. The test model used by the invention is a model with similar shape and yaw axis inertia, adopts a metal frame and a carbon fiber skin structure, can be used for carrying out the test after adopting a conventional dynamic test model for balancing weight, simulates the course motion characteristic of the aircraft through a free yaw motion device, obtains course motion histories under different pitch angles and control plane skewness by utilizing an encoder, can obtain the influence of the control plane deflection on the course motion characteristic through data analysis, and provides a reference basis for analyzing the course control efficiency of the control plane of the flying wing aircraft. The nonlinear and unsteady phenomena of the airflow passing through the model are consistent with those of a real aircraft, and the test result of the course control efficiency of the control surface is reliable. The test method is simple to operate, reliable in result, capable of being used for identifying potential course non-instruction movement and capable of evaluating control surface course control efficiency of a full-size real machine.
Description
Technical Field
The invention belongs to the technical field of low-speed wind tunnel tests, and particularly relates to a wind tunnel test method for researching course control efficiency of an aircraft control surface.
Background
The flying wing layout aircraft has the advantages that the three-axis stability of the flying wing layout aircraft is reduced due to the fact that a horizontal tail and a vertical tail are omitted, the problem of course stability is particularly prominent, and safe flight is guaranteed by the aid of a novel control surface, such as a cracking type resistance rudder, a full-motion wing tip and the like.
At present, whether a course control surface can meet a control requirement is evaluated, and a stability control characteristic analysis is mainly performed by using static aerodynamic force (moment) and dynamic stability derivative data, but a space flow field has a complex unsteady characteristic when a flying wing is distributed at a large attack angle, and a control efficiency evaluation result is influenced because a linear small disturbance assumption is adopted in a dynamic stability derivative obtaining process and the nonlinear and unsteady characteristic description is incomplete. In order to promote the engineering application process of the flying wing layout, a wind tunnel test method for researching the course control efficiency of the control surface of the flying wing aircraft needs to be developed.
Disclosure of Invention
In order to solve the problems, the invention provides a wind tunnel test method for researching the course control efficiency of a control surface of a flying wing aircraft, which adopts a wind tunnel free motion test to directly evaluate and solves the problem that the simulation of real-time nonlinear and unsteady aerodynamic phenomena is not completely described by adopting static aerodynamic force (moment) and dynamic stability derivative construction.
The technical scheme adopted by the invention is as follows: a wind tunnel test method for researching course control efficiency of a control surface of an aerovane comprises the following steps:
(a) Processing a course inertia characteristic similar test model with a variable control surface deflection, wherein the test model is provided with a metal framework and a carbon fiber skin, and balancing weights of the model according to an inertia similarity relation, and the similarity relation meets the following formula:
in the formula I z,model For testing the rotational inertia of yaw axis of model, I z,airplane Rotational inertia of yaw axis of true machine, b model For the test model reference length, b airplane The reference length is the real machine reference length;
(b) A free yaw movement device is arranged in the test model, a support rod penetrates through the back of the test model and is connected with the free yaw movement device, the support rod is fixed on the wind tunnel pitch angle movement mechanism, and an encoder data line in the free yaw movement device is connected with a wind tunnel acquisition system;
(c) Before a low-speed wind tunnel test, placing the test model in a wind tunnel test section, wherein the axis of a test model body is positioned at the central axis of the wind tunnel, and determining the zero yaw angle position of the model;
(d) The angle of the control surface is changed through the control surface angle sheet, so that the deflection angle of the rudder reaches a target value delta contral ；
(e) Adjusting the pitch angle of the test model through the wind tunnel pitch angle mechanism to enable the pitch angle of the test model to reach a target value theta, and placing the yaw angle of the model at a target initial position psi start Locking the yaw freedom degree of the free yaw device through remote control;
(f) Starting the wind tunnel, starting a wind tunnel data acquisition system after the wind speed reaches a target value V, acquiring the yaw motion history of the test model, remotely releasing the yaw freedom degree of the free yaw device after 10s, stopping acquisition after 60s, and stopping wind;
(g) Repeating the step e and the step f to obtain yaw movement history data under the same control plane deflection and different pitch angles;
(h) Repeating the step d, the step e, the step f and the step g, obtaining yaw movement history data under different control plane skewness, and ending the wind tunnel test of the course control efficiency of the control plane;
(i) Comprehensively analyzing the influence of different control surface combination skewness on the course motion characteristic under different attack angles to obtain the control surface course control efficiency, wherein the comprehensive analysis adopts a quality factor Pp-v value to evaluate the control surface course control efficiency, and the formula is as follows,
where Δ ψ is the absolute value of the amplitude change from the peak to the nearest valley of the yaw oscillation motion, and Δ t is the time required for the corresponding motion.
Furthermore, the free yaw movement device has a movement range of +/-360 degrees, can rotate for multiple circles continuously, and has an encoder measurement precision of 0.09 degrees.
Further, in the step a, the deviation of the inertia quantity of the yaw axis after the test model is weighted is not more than 3% of the theoretical value.
Further, as in step e described above, the yaw angle initial position ψ start The three points are respectively-0.5 degrees, 0 degrees and 0.5 degrees so as to obtain different oscillation balance points of the test model.
Further, as in step f above, the sampling frequency of the acquisition system is 200kHz.
The invention has the advantages that: the invention utilizes the test model with similar appearance and inertia to carry out the course control efficiency test of the control surface in the wind tunnel, the nonlinear and unsteady phenomena of the airflow passing through the model are consistent with the real aircraft, and the course control efficiency test result of the control surface is reliable. The test model can be obtained by adopting a conventional dynamic test model for balancing weight, so that the universality of the wind tunnel test model is improved, and the wind tunnel test cost is reduced. The test method is simple to operate and reliable in result, can be used for identifying potential course non-command movement, evaluating nonlinear and unsteady aerodynamic problems, and evaluating the course control efficiency of the control surface of the full-size real machine through a similar relation.
Drawings
FIG. 1 is a schematic diagram of a wind tunnel test system;
FIG. 2 is a control surface course control efficiency diagram obtained by the present invention;
the device comprises a test model 1, a free yaw movement device 2, a free yaw movement device 3, a support rod 4, a pitch angle movement mechanism 5 and an inlet of a wind tunnel opening test section.
Detailed Description
The invention is further described by way of example in the following description with reference to the accompanying drawings:
example 1
A wind tunnel test method for researching course control efficiency of a control surface of an aerovane comprises the following steps:
(a) Processing a course inertia characteristic similar test model with a variable control surface skewness, wherein the test model is provided with a metal framework and a carbon fiber skin, the model is weighted according to an inertia similarity relation, the deviation of the inertia quantity of a yaw axis after the model is weighted does not exceed 3% of a theoretical value, and the similarity relation meets the following formula:
in the formula I z,model For testing the rotational inertia of the yaw axis of the model, I z,airplane Rotational inertia of yaw axis of true machine, b model For the test model reference length, b airplane The reference length is the real machine reference length;
(b) As shown in fig. 1, a free yaw movement device is arranged in the test model, a supporting rod penetrates through the back of the test model and is connected with the free yaw movement device, the supporting rod is fixed on a wind tunnel pitch angle movement mechanism, an encoder data line in the free yaw movement device is connected with a wind tunnel acquisition system, the movement range of the free yaw movement device is +/-360 degrees, the free yaw movement device can continuously rotate for multiple circles, and the measurement precision of an encoder is 0.09 degrees;
(c) Before a low-speed wind tunnel test, placing the test model in a wind tunnel test section, wherein the axis of a test model body is positioned at the central axis of the wind tunnel, and determining the zero yaw angle position of the model;
(d) The angle of the control surface is changed through the control surface angle sheet, so that the deflection angle of the rudder reaches a target value delta contral ；
(e) Adjusting the pitch angle of the test model through the wind tunnel pitch angle mechanism to enable the pitch angle of the test model to reach a target value theta, and placing the yaw angle of the model at a target initial position psi start By passingRemotely controlling the yaw freedom degree of the locked free yaw device, and the initial position psi of the yaw angle start The test model is divided into three parts, namely-0.5 degrees, 0 degrees and 0.5 degrees respectively, so as to obtain different oscillation balance points of the test model;
(f) Starting the wind tunnel, starting a wind tunnel data acquisition system after the wind speed reaches a target value V, acquiring the yaw movement history of the test model with the sampling frequency of 200kHz, remotely releasing the yaw freedom degree of the free yaw device after 10s, stopping acquisition after 60s, and stopping wind;
(g) Repeating the step e and the step f to obtain yaw movement history data under different pitch angles of the same control surface skewness;
(h) Repeating the step d, the step e, the step f and the step g, obtaining yaw movement history data under different control plane skewness, and ending the wind tunnel test of the course control efficiency of the control plane;
(i) Comprehensively analyzing the influence of different control surface combination skewness on the course motion characteristic under different attack angles to obtain the control surface course control efficiency, wherein the comprehensive analysis adopts a quality factor Pp-v value to evaluate the control surface course control efficiency, the formula is as follows,
where Δ ψ is the absolute value of the amplitude change from the peak to the nearest valley of the yaw oscillation motion, and Δ t is the time required for the corresponding motion.
Example 2
In the embodiment, a Pp-v value diagram of the test model shown in FIG. 2 under different pitch angles and control surface combined deflection is obtained, and it can be seen from the diagram that the Pp-v value is still higher except for the pitch angle of 40 degrees, and the Pp-v values of other pitch angles are controlled by the control surface combination to achieve the expected effect, so that the non-command yaw motion can be effectively reduced by fixing the control surface combination, and the heading flight quality is improved.
Claims (5)
1. A wind tunnel test method for researching course control efficiency of a control surface of an aerovane is characterized by comprising the following steps:
(a) Processing a course inertia characteristic similar test model with a variable control surface deflection, wherein the test model is provided with a metal framework and a carbon fiber skin, and balancing weights of the model according to an inertia similarity relation, and the similarity relation meets the following formula:
in the formula I z,model For testing the rotational inertia of yaw axis of model, I z,airplane Rotational inertia of yaw axis of true machine, b model For the test model reference length, b airplane Is a real machine reference length;
(b) A free yaw movement device is arranged in the test model, a support rod penetrates through the back of the test model and is connected with the free yaw movement device, the support rod is fixed on the wind tunnel pitch angle movement mechanism, and an encoder data line in the free yaw movement device is connected with a wind tunnel acquisition system;
(c) Before a low-speed wind tunnel test, placing the test model in a wind tunnel test section, wherein the axis of a test model body is positioned at the central axis of the wind tunnel, and determining the zero yaw angle position of the model;
(d) The angle of the control surface is changed through the control surface angle sheet, so that the deflection angle of the rudder reaches a target value delta contral ；
(e) Adjusting the pitch angle of the test model through the wind tunnel pitch angle mechanism to enable the pitch angle of the test model to reach a target value theta, and placing the yaw angle of the model at a target initial position psi start Locking the yaw freedom degree of the free yaw device through remote control;
(f) Starting the wind tunnel, starting a wind tunnel data acquisition system after the wind speed reaches a target value V, acquiring the yaw motion history of the test model, remotely releasing the yaw freedom degree of the free yaw device after 10s, stopping acquisition after 60s, and stopping wind;
(g) Repeating the step e and the step f to obtain yaw movement history data under different pitch angles of the same control surface skewness;
(h) Repeating the step d, the step e, the step f and the step g, obtaining yaw movement history data under different control plane skewness, and ending the wind tunnel test of the course control efficiency of the control plane;
(i) Comprehensively analyzing the influence of different control plane combination skewness on the course motion characteristic under different attack angles to obtain the control plane course control efficiency, comprehensively analyzing and evaluating the control plane course control efficiency by adopting a quality factor Pp-v value, wherein the formula is as follows,
where Δ ψ is the absolute value of the amplitude change from the peak to the nearest valley of the yaw oscillation motion, and Δ t is the time required for the corresponding motion.
2. The wind tunnel test method for researching course control efficiency of the control surface of the flying wing aircraft as claimed in claim 1, wherein the free yaw movement device has a movement range of +/-360 degrees and can rotate continuously for multiple cycles, and the encoder has a measurement precision of 0.09 degrees.
3. The wind tunnel test method for researching the course control efficiency of the control surface of the flying wing aircraft according to claim 1, wherein in the step a, the deviation of the inertia quantity of a yaw shaft after the test model is balanced by weight does not exceed 3% of a theoretical value.
4. The wind tunnel test method for researching flight vehicle control surface course control efficiency according to claim 1, wherein in the step e, the yaw angle initial position psi start The three vibration balance points are respectively-0.5 degrees, 0 degrees and 0.5 degrees so as to obtain different vibration balance points of the test model.
5. The wind tunnel test method for researching course control efficiency of the control surface of the flying wing aircraft according to any one of claims 1-4, wherein in the step f, the sampling frequency of the acquisition system is 200kHz.
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