CN116358821A

CN116358821A - Three-degree-of-freedom virtual flight wind tunnel test system

Info

Publication number: CN116358821A
Application number: CN202310267904.XA
Authority: CN
Inventors: 唐鹏; 张艳来; 冯杨; 郭介东; 张正
Original assignee: Beihang University
Current assignee: Beihang University
Priority date: 2023-03-15
Filing date: 2023-03-15
Publication date: 2023-06-30

Abstract

The invention discloses a three-degree-of-freedom virtual flight wind tunnel test system, which relates to the technical field of wind tunnel test, and utilizes a microminiature flight control module to control deflection of a scaled aircraft model to realize three rotational degrees of freedom motions, a support mechanism module is fixed in a low-speed closed direct-flow wind tunnel module, the scaled aircraft model is supported by the support mechanism module to perform three rotational degrees of freedom motions in the low-speed closed direct-flow wind tunnel module, and the low-speed closed direct-flow wind tunnel module is used for providing a wind tunnel virtual flight environment, so that the scaled aircraft model realizes wind tunnel virtual flight in the low-speed closed direct-flow wind tunnel module, simulates rotational motions of a real aircraft on a boundary state, and realizes research on aerodynamics under an uncontrolled law and research control under a controlled law. The three-degree-of-freedom virtual flight wind tunnel test system disclosed by the invention has the advantages of low cost, simple flow and low risk, and can be used for researching the pneumatics under the control law and researching the control under the control law.

Description

Three-degree-of-freedom virtual flight wind tunnel test system

Technical Field

The invention relates to the technical field of wind tunnel test, in particular to a three-degree-of-freedom virtual flight wind tunnel test system.

Background

In aircraft design, wind tunnels (wind tunnel laboratories) are commonly used to perform blowing tests to obtain aerodynamic properties of an aircraft, such as various aerodynamic derivatives. The aerodynamic derivatives and other data are then used to model dynamics, and then the controller design is performed. In the unmanned aerial vehicle industry, wind tunnel test is often omitted, estimated data or Computational Fluid Dynamics (CFD) computational data are directly used, and design verification of a controller is usually performed by using a scaled aircraft for free flight test, so that the test flow is complex, the aircraft customization cost is high, and the test risk is high. Based on the test difficulty of the free flight test, if the wind tunnel test (wind tunnel test) with relatively low cost can be used for carrying out design verification of the unmanned aerial vehicle controller, the test flow is simplified, and the test risk is reduced.

Wind tunnel testing is omitted in the unmanned aerial vehicle industry because: at present, the unmanned aerial vehicle industry is concentrated on a small unmanned aerial vehicle, such as a large-scale unmanned aerial vehicle, the requirements of the unmanned aerial vehicle are not high, the pneumatic data under a small attack angle can obtain a relatively accurate result through estimation, various methods exist for pneumatic estimation data, CFD (computational fluid dynamics) is a method for obtaining a relatively accurate pneumatic data field, the data are relatively accurate in a small attack angle state, the estimation result is inaccurate when the attack angle is large, the model of the pneumatic data is insufficient to describe aerodynamic force and aerodynamic moment when the attack angle is large, the estimation data are inaccurate and even difficult to converge, and an aircraft controller is designed based on the pneumatic data, and incorrect pneumatic can cause incorrect control; on the other hand, wind tunnel static/dynamic experiments (tests) which can obtain more accurate results are too expensive; therefore, the current unmanned aerial vehicle industry generally keeps the aircraft in a state with relatively accurate pneumatic data as far as possible through a control limiting means, and is far away from a part with inaccurate estimation, and on the other hand, the aircraft is well lost due to the experience of a flight crew.

However, with the gradual development of the unmanned aerial vehicle industry towards large-scale, the flight boundary of the aircraft needs to meet the corresponding airworthiness standard, and the flight boundary has the corresponding requirement, so that a plurality of methods are needed to study the aerodynamic characteristics of the aircraft at the flight boundary, and further, based on the aerodynamic characteristics, how the controller is needed is studied, and the controller at the flight boundary needs to be subjected to higher-requirement experiments, such as free flight of a wind tunnel, real test flight and the like. However, in the aspect of aerodynamics, the existing wind tunnel test system mainly adopts static tests, and only static and aerodynamic coefficient measurement can be performed. However, the static aerodynamic coefficient is inaccurate in the boundary state, and some corresponding special wind tunnels exist for the aerodynamic condition of the boundary state, but the static aerodynamic coefficient is not so much, and is mainly concentrated on obtaining various derivatives of the aerodynamic model of the aircraft, and the aerodynamic model is built by splitting the aerodynamic coefficient, so that the influence of the large influence terms (neglected small influence terms) is known approximately, and therefore, the specific influence is obtained through experiments. In the aspect of flight control, a common wind tunnel has no related method, and at present, a special wind tunnel mainly comprises a free flight wind tunnel, wherein the free flight wind tunnel has six degrees of freedom (three translational motions and three rotational motions), and the free flight wind tunnel has high price, complex flow and high risk.

In summary, there is a need in the art for a wind tunnel test system that can study both pneumatic and controlled under no control law, with low cost, simple flow, and low risk.

Disclosure of Invention

The invention aims to provide a three-degree-of-freedom virtual flight wind tunnel test system which can study and control under the control law and has low cost, simple flow and low risk.

In order to achieve the above object, the present invention provides the following solutions:

the system comprises a low-speed closed direct current wind tunnel module, a supporting mechanism module, a scaled aircraft model and a microminiature flight control module;

the scaled aircraft model is an equal-ratio model of the appearance of a real aircraft; the microminiature flight control module is arranged in the scaled aircraft model; the scaled aircraft model is connected with the microminiature flight control module; the miniature flight control module is used for controlling the scaled aircraft model to deflect so as to realize three-degree-of-freedom rotation;

the supporting mechanism module is fixed in the low-speed closed direct-current wind tunnel module; the supporting mechanism module is used for supporting the scaled aircraft model to perform three-degree-of-freedom rotation in the low-speed closed direct current wind tunnel module;

the low-speed closed direct current wind tunnel module is used for providing a virtual wind tunnel flight environment, so that the scaled aircraft model realizes the virtual wind tunnel flight in the low-speed closed direct current wind tunnel module; the wind tunnel virtual flight is used for simulating a flight boundary on three degrees of freedom of rotation; the low-speed closed direct-current wind tunnel module simulates an airflow field around the scaled aircraft model by generating and controlling airflow, so that the scaled aircraft model receives aerodynamic moment, generates corresponding angular velocity, simulates the rotational motion of a real aircraft on a boundary state, and realizes the study of aerodynamics under the control law and the study of control under the control law.

Optionally, the low-speed closed direct current wind tunnel module comprises an inlet stabilizing section, a contraction section, an experiment section and a fan section which are connected in sequence;

the inlet stabilizing section is an airflow inlet of the wind tunnel; the inlet stabilizing section comprises a section of pipeline with a uniform section and provided with a rectifying facility; the rectifying facility comprises a honeycomb device and a damping net;

the contraction section is a section of pipeline which contracts in an equal ratio; the maximum cross section of the contraction section is equal to the cross section of a pipeline provided with a rectifying facility; the contraction section is used for uniformly accelerating the airflow;

the experiment section is a section of pipeline with a uniform cross section and fixed with the supporting mechanism module; the supporting mechanism module is used for supporting the scaled airplane model to perform three-degree-of-freedom rotation in the experimental section; the section of the experimental section is equal to the minimum section of the contraction section in size; the experimental section is used for simulating airflow fields around the scaled aircraft model;

the fan section is an airflow outlet of the wind tunnel; the fan section comprises a section of pipeline with a uniform section and provided with a fan; the fan section is used for providing power for the flow of the airflow.

Optionally, the support mechanism module comprises a connecting piece, an upper bracket, a lower bracket and a base;

the base is fixed in the experimental section; the lower bracket is fixed on the base; the lower bracket adopts a streamline interface; the lower bracket is used for supporting the upper bracket; the upper bracket adopts a cylindrical design; the upper bracket is used for supporting the scaled airplane model; the lower bracket is also used for enabling the scaled airplane model to be located at the center of the experimental section in height;

the uppermost end of the upper bracket is connected with the connecting piece through a section of cylinder with the cross section smaller than that of the upper bracket; the connecting piece is used for connecting the scaled airplane model with the upper bracket; the connecting piece is used for supporting the scaled airplane model to perform angular displacement free motion around the mass center so as to realize three-degree-of-freedom rotation; the upper bracket is also used for limiting the pitching movement of the scaled airplane model.

Optionally, the connecting piece comprises an X-direction bearing, a Y-direction bearing, a Z-direction bearing, an inner frame, a middle bracket and a Z plug;

the Z-direction bearing is connected with the upper bracket through a cylinder with the cross section smaller than that of the upper bracket; the Z plug is arranged at one end of the Z-direction bearing, which is far away from the upper bracket; the middle support and the inner support are used for supporting the X-direction bearing and the Y-direction bearing; the X-direction bearing, the Y-direction bearing and the Z-direction bearing are used for supporting the scaled airplane model to perform angular displacement free movement around the mass center, so that three-degree-of-freedom rotation is realized.

Optionally, the scaled aircraft model includes steering engines, rudders, elevators and ailerons;

the steering engine is connected with the microminiature flight control module; the steering engine is controlled by a control law in the microminiature flight control module, and the steering engine further respectively controls the deflection of the rudder, the elevator, the aileron and the control surface.

Optionally, the microminiature flight control module comprises a flight control board unit and a code rapid prototype implementation unit based on Matlab/Simulink;

the flight control board unit is connected with the steering engine through a data line; the flight control board unit is used for controlling the steering engine, and the steering engine is used for respectively controlling the deflection of the rudder, the elevator, the aileron and the control surface to realize the deflection of each control surface of the scaled aircraft model;

the Matlab/Simulink-based code rapid prototype implementation unit is used for converting a control law model built in the Matlab/Simulink into a C language to be burnt into the flight control board unit, so that control of a control law on a control surface is realized.

Optionally, the flight control board unit adopts a micro onboard embedded control system.

Optionally, the flight control board unit comprises a miniature angle of attack sideslip angle sensor, an inertia measurement unit and a computer.

According to the specific embodiment provided by the invention, the invention discloses the following technical effects:

the invention discloses a three-degree-of-freedom virtual flight wind tunnel test system, which utilizes a microminiature flight control module to control deflection of a scaled aircraft model to realize three rotational degrees of freedom motions, and utilizes the supporting mechanism module to support the scaled aircraft model to perform three rotational degrees of freedom motions in the low-speed closed direct-flow wind tunnel module by fixing the supporting mechanism module in the low-speed closed direct-flow wind tunnel module, and provides a wind tunnel virtual flight environment through the low-speed closed direct-flow wind tunnel module, so that the scaled aircraft model realizes wind tunnel virtual flight in the low-speed closed direct-flow wind tunnel module, and one characteristic of wind tunnel virtual flight is that the wind tunnel virtual flight simulation system can simulate a flight boundary in three rotational degrees of freedom at lower cost, and can simulate the rotational motion of an aircraft (a real aircraft) in a boundary state in a translational motion relatively truly, so that the study and the study control in a control law are realized.

Drawings

In order to more clearly illustrate the embodiments of the present invention or the technical solutions of the prior art, the drawings that are needed in the embodiments will be briefly described below, it being obvious that the drawings in the following description are only some embodiments of the present invention, and that other drawings may be obtained according to these drawings without inventive effort for a person skilled in the art.

FIG. 1 is a block diagram of an embodiment of a three degree-of-freedom virtual flying wind tunnel test system of the present invention;

FIG. 2 is a schematic diagram of the three-degree-of-freedom virtual flying wind tunnel test system of the present invention;

FIG. 3 is a schematic view of the structure of the upper, lower and base of the support mechanism of the present invention;

FIG. 4 is a right side view of the upper and lower brackets and base of the support mechanism of the present invention;

FIG. 5 is a top view of a cross-section of a base of the support mechanism of the present invention;

FIG. 6 is a top view of a cross-section of a lower bracket of the support mechanism of the present invention;

FIG. 7 is a schematic view of a support mechanism connector according to the present invention;

FIG. 8 is a schematic diagram of a low-speed DC closed wind tunnel module according to the present invention;

FIG. 9 is a schematic view of a wind tunnel geometry model according to the present invention;

FIG. 10 is a rear view of a wind tunnel geometry model of the present invention;

FIG. 11 is a left side view of the wind tunnel geometry model of the present invention;

FIG. 12 is a top view of a wind tunnel geometry model of the present invention;

FIG. 13 is a schematic view of the upper and lower equiangular axes of the present invention.

Detailed Description

The following description of the embodiments of the present invention will be made clearly and completely with reference to the accompanying drawings, in which it is apparent that the embodiments described are only some embodiments of the present invention, but not all embodiments. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.

In order that the above-recited objects, features and advantages of the present invention will become more readily apparent, a more particular description of the invention will be rendered by reference to the appended drawings and appended detailed description.

FIG. 1 is a block diagram of an embodiment of a three degree-of-freedom virtual flying wind tunnel test system of the present invention. Referring to fig. 1, the three-degree-of-freedom virtual flight wind tunnel test system comprises a low-speed closed direct current wind tunnel module 101, a supporting mechanism module 102, a scaled aircraft model 103 and a microminiature flight control module (not shown in the figure).

The scaled aircraft model 103 is an equal-scale model of the exterior of a real aircraft; the microminiature flight control module is arranged inside the scaled aircraft model 103; the scaled aircraft model 103 is connected with a miniature flight control module; the microminiature flight control module is used for controlling the deflection of the scaled airplane model 103 to realize three-degree-of-freedom rotation.

The supporting mechanism module 102 is fixed in the low-speed closed direct current wind tunnel module 101; the support mechanism module 102 is used for supporting the scaled aircraft model 103 to perform three degrees of rotational freedom motion in the low-speed closed direct current wind tunnel module 101.

The low-speed closed direct-current wind tunnel module 101 is used for providing a virtual wind tunnel flight environment, so that the scaled aircraft model 103 realizes the virtual wind tunnel flight in the low-speed closed direct-current wind tunnel module 101; the wind tunnel virtual flight is used for simulating a flight boundary on three degrees of freedom of rotation; the low-speed closed direct-current wind tunnel module 101 simulates an airflow field around the scaled aircraft model 103 by generating and controlling airflow, so that the scaled aircraft model 103 receives aerodynamic moment to generate corresponding angular velocity, simulate the rotational motion of a real aircraft on a boundary state, and realize research on aerodynamics under a non-control law and research and control under a control law.

The low-speed closed direct current wind tunnel module 101 comprises an inlet stabilizing section, a contraction section, an experiment section and a fan section which are connected in sequence.

The inlet stabilizing section is an airflow inlet of the wind tunnel; the inlet stabilizing section comprises a section of pipeline with a uniform section and provided with a rectifying facility; the rectifying means comprises a honeycomb and a damping net.

The contraction section is a section of pipeline which contracts in an equal ratio; the maximum cross section of the contraction section is equal to the cross section of the pipeline provided with the rectifying facility; the constriction serves to accelerate the air flow uniformly.

The experimental section is a section of pipeline with a constant cross section and fixed with a supporting mechanism module 102; the supporting mechanism module 102 is used for supporting the scaled airplane model 103 to perform three-degree-of-freedom rotation in the experimental section; the section of the experimental section is equal to the minimum section of the contraction section; the experimental section was used to simulate the airflow field around the scaled aircraft model 103.

The fan section is an airflow outlet of the wind tunnel; the fan section comprises a section of pipeline with a uniform section and provided with a fan; the fan section is used for providing power for the flow of the air flow.

The support mechanism module 102 includes a connector, an upper bracket, a lower bracket, and a base. The base is fixed in the experimental section; the lower bracket is fixed on the base; the lower bracket adopts a streamline interface; the lower bracket is used for supporting the upper bracket; the upper bracket adopts a cylindrical design; the upper bracket is used for supporting the scaled aircraft model 103; the lower support also serves to center the scaled airplane model 103 in height in the experimental section. The uppermost end of the upper bracket is connected with the connecting piece through a section of cylinder with the cross section smaller than that of the upper bracket; the connecting piece connects the scaled aircraft model 103 with the upper bracket; the connecting piece is used for supporting the scaled airplane model 103 to perform angular displacement free motion around the mass center so as to realize three-degree-of-freedom rotation; the upper bracket also serves to unrestricted pitching movement of the scaled aircraft model 103.

The connecting piece comprises an X-direction bearing, a Y-direction bearing, a Z-direction bearing, an inner frame, a middle bracket and a Z plug. The Z-direction bearing is connected with the upper bracket through a cylinder with the cross section smaller than that of the upper bracket; the Z plug is arranged at one end of the Z-direction bearing, which is far away from the upper bracket; the middle bracket and the inner bracket are used for supporting an X-direction bearing and a Y-direction bearing; the X-direction bearing, the Y-direction bearing and the Z-direction bearing are used for supporting the scaled airplane model 103 to perform angular displacement free motion around the mass center, so that three-degree-of-freedom rotation is realized.

The scaled aircraft model 103 includes steering engines, rudders, elevators and ailerons. The steering engine is connected with the microminiature flight control module; the control law in the microminiature flight control module controls the steering engine, and the steering engine controls the deflection of the rudder, the elevator and the aileron and the deflection of the control surface respectively.

The microminiature flight control module comprises a flight control board unit and a code rapid prototype implementation unit based on Matlab/Simulink. The flight control board unit adopts a miniature onboard embedded control system. The flight control board unit (miniature onboard embedded control system) comprises a miniature attack angle sideslip angle sensor, an inertia measurement unit and a computer. The flight control board unit is connected with the steering engine through a data line; the flight control board unit is used for controlling a steering engine, and the steering engine is used for respectively controlling deflection of a rudder, an elevator and an aileron and deflection of a control surface, so that deflection of each control surface of the scaled aircraft model 103 is realized. The code rapid prototype implementation unit based on Matlab/Simulink is used for converting a control law model established in Matlab/Simulink into a C language to be burnt into the flight control board unit, so that control of the control law on the control surface is realized.

The technical scheme of the invention is described in the following by a specific embodiment:

the invention discloses a three-degree-of-freedom virtual flight wind tunnel test system, which is a three-degree-of-freedom virtual flight small teaching wind tunnel experimental device. The invention develops a three-degree-of-freedom test system (three-degree-of-freedom virtual flight wind tunnel test system) based on a small wind tunnel based on the test difficulty of free flight test, and can be used as a control law design test bed of a nominal model. Fig. 2 is a schematic diagram of the three-degree-of-freedom virtual flight wind tunnel test system (wind tunnel test system) according to the present invention, and as shown in fig. 2, the three-degree-of-freedom virtual flight wind tunnel test system according to the present invention is composed of a low-speed closed direct current wind tunnel module (low-speed direct current wind tunnel), a supporting mechanism module (supporting device), a scaled aircraft model (aircraft scaled model), and a microminiature flight control module.

The microminiature flight control module is used for controlling steering engines of the scaled aircraft model to realize deflection of each control surface of the scaled aircraft model. The miniature flight control module is positioned in the scaled aircraft model and comprises a hardware part and a software part, wherein the hardware part is a flight control board unit, the flight control board unit adopts a miniature onboard embedded control system in the market, and the flight control board unit comprises a miniature attack angle sideslip angle sensor, an inertia measurement unit, a computer and the like and is connected with a steering engine on the scaled aircraft model through a data line; the software part is a code rapid prototype implementation unit based on Matlab/Simulink, and has the function of converting a control law model built in Matlab/Simulink into a C language to be burnt into a flight control board (flight control board unit) to realize control law control on an aircraft control surface, namely, control law design is carried out based on Simulink in MATLAB, and codes supported by a flight control system can be rapidly generated through modularized design of a Simulink end controller model through a flight control development support package carried by a Simulink platform, so that the C language codes are rapidly generated and burnt into the flight control board.

The scaled aircraft model is an equal-ratio model of the appearance of a real aircraft, and is provided with a steering engine, wherein the steering engine is connected with a flight control board unit of a microminiature flight control module, a control law in the flight control board unit controls the steering engine, and the steering engine respectively controls deflection of a rudder, an elevator and an aileron in the scaled aircraft model and deflection of a control surface.

The support mechanism (support mechanism module) is used for supporting the scaled aircraft model and comprises four parts, namely a connecting piece, an upper bracket, a lower bracket and a base, wherein the support mechanism bracket and the base model are shown in fig. 3, 4, 5 and 6, and fig. 3 shows a structural schematic diagram of the upper bracket, the lower bracket and the base; fig. 4 shows a right side view of the upper bracket, the lower bracket and the base, and fig. 6 and 5 illustrate a cross-sectional top view of the support mechanism lower bracket and the base. The connecting piece is connected with the scaled airplane model and the supporting mechanism and supports the scaled airplane model to move in three degrees of freedom, and the connecting piece mainly comprises three direction bearings (an X direction bearing, a Y direction bearing and a Z direction bearing), an inner frame, a middle support and a Z plug, and the supporting mechanism three degrees of freedom connecting piece model is shown in FIG. 7. The upper bracket supports the scaled airplane model and ensures that the pitching motion of the scaled airplane model is not limited, so that the upper bracket adopts a cylindrical design, and the uppermost end of the upper bracket is connected with the connecting piece by a section of smaller cylinder. The lower bracket supports the upper bracket, so that the aircraft is positioned in the center of the wind tunnel in height, the test is facilitated, and a streamline interface is adopted to reduce air flow separation; the whole supporting mechanism is fixed in the wind tunnel by the base, and the bracket is stabilized, so that the bracket cannot generate integral movement interference test. The supporting mechanism needs to enable the scaled aircraft model to perform angular displacement free motion around the mass center, meet the bearing capacity of the system, reduce the influence of supporting components (all component parts of the supporting mechanism) on air flow interference as much as possible, and reduce friction damping and structural vibration. The support mechanism and the scaled airplane model are directly placed in an experimental section of the low-speed closed direct-current wind tunnel module.

The low-speed direct current closed wind tunnel module (low-speed closed direct current wind tunnel module) provides an environment of a test system, and simulates the air flow condition around an aircraft by generating and controlling air flow, and fig. 8 shows a 2D model diagram of the wind tunnel test system of the invention, and the low-speed direct current closed wind tunnel module is composed as shown in fig. 8 and comprises an inlet stabilizing section (stabilizing section), a contraction section, an experimental section and a fan section. The inlet stabilizing section is a wind tunnel airflow inlet and is a section of pipeline with uniform cross section provided with rectifying facilities (a honeycomb device and a damping net); the pipeline of the contraction section contracts in an equal ratio, so that the airflow is uniformly accelerated, and the flow field quality of the experimental section is improved; the experimental section simulates a surrounding flow field of the aircraft, is a core unit of the low-speed direct-current closed wind tunnel module, and is also the position of the supporting mechanism and the aircraft scaling model; the fan section is a fan installation position and provides power for the flow of air flow, and meanwhile, the fan section is an air flow outlet of the wind tunnel. The wind tunnel geometric model of the invention is shown in fig. 9, and the rear view, the left view, the upper view and the upper and lower equiangular axial schematic views of the wind tunnel geometric model of the invention are shown in fig. 10, 11, 12 and 13.

Specifically, the design and main parameters of the low-speed direct current wind tunnel (virtual flying wind tunnel) are as follows:

the direct-flow wind tunnel (direct-flow wind tunnel) is also called an open-circuit wind tunnel, and is characterized in that air flow is discharged out of the wind tunnel after passing through an experimental section, and no special pipeline system is used for guiding back. The small direct current wind tunnel is generally built in a laboratory, and two sections of the large direct current wind tunnel are directly connected with the atmosphere. The invention discloses a three-degree-of-freedom virtual flying wind tunnel test system, which adopts a closed direct current wind tunnel (low-speed direct current wind tunnel) which is divided into an inlet stabilizing section, a contraction section, an experimental section and a fan section, wherein the closed refers to the sealing of the experimental section.

Experiment section:

in order to simulate a prototype flow field, the quality of the experimental section flow field must meet certain requirements. The parameters of the wind tunnel experimental section are as follows: the section is 600mm rectangle, pipeline size D ₁ The length of the direct current closed wind tunnel experimental section is generally 1.5-2.5D in the range of value of 600mm ₁ The wind tunnel of the invention takes L ₁ ＝2D ₁ =1200 mm. The pressure regulating seam is designed at the outlet of the experimental section and is used for supplementing air into the wind tunnel so as to keep the pressure of the experimental section and the pressure of the atmosphere outside the wind tunnel basically equal, thereby simplifying the sealing device of the experimental section.

Stabilizing section:

the stabilizing section is typically a constant cross-section pipe that houses rectifying means (honeycomb and damping net). The front inlet section of the stabilizing section is 1200mm rectangular, and the pipeline dimension D ₂ =1200 mm; the section of the stabilizing section is 1000 mm-1000 mm rectangle, and the dimension D of the pipeline is ₃ =1000 mm; length of the stable segment is L ₃ /D ₃ Between=1.0 and 1.5. The wind tunnel of the invention takes L ₃ ＝0.6*D ₃ ＝600mm。

Generally at the corners, the corner guide vane turning radius r= (0.1-0.2) wind tunnel duct size D; the chord length C of the guide vane is about 0.25D; the distance between the guide vanes is (0.3-0.6) C. The wind tunnel of the present invention does not use corner deflectors but can be used to calculate a honeycomb for guiding and reducing the vortex dimensions, where d=d ₃ 1000mm, c=250 mm, baffle spacing=0.5×250=125 mm. The distance between two adjacent corner guide sheets with the honeycomb lattice size M of 10 percent and the length L ₂ ＝(12～20 M). When L ₂ Near 12M, this is most effective in reducing turbulence. Thus, m=0.1×125mm=12.5 mm can be found as the honeycomb length.

The damping net has the main functions of stabilizing and regulating the airflow in the stable section and reducing the turbulence in the experimental section, and the mesh size is 1/5-1/15 of the grille size of honeycomb. The net has a solidity of 35% and is effective in suppressing turbulence. The literature suggests that the first layer of damping mesh is located 20 to 40 times the cell grid size downstream of the cell to ensure adequate damping distance for turbulence after the cell. Thus, given a first layer of damping mesh at 30 x 12.5 mm=375 mm after the honeycomb, the mesh size of the damping mesh was 15mm x 0.2=3 mm, the mesh solidity was 35%, the number of damping meshes per direction=1200/3-1=399, the total mesh number was 798, h=0.53 mm was found from 798 x 1200mm x H/(1200 mm x 1200 mm) =0.35, and H was the damping mesh thickness.

Shrink section:

the contraction section is used for uniformly accelerating air flow and is beneficial to improving the flow field quality of the experimental section;

general L ₄ /D ₄ =0.75 to 1.25, where D ₄ To reduce the inlet cross-section size, D ₄ =1000 mm. The wind tunnel contraction section of the invention is taken as L ₄ ＝0.6*D ₄ =600 mm, shrinkage curve is:

the center of the section of the inlet end is taken as the origin of coordinates, the vertical axis is taken as the x axis, the horizontal axis is taken as the r axis, the length of the contracted section is L, eta is the contraction ratio, eta is equal to 2.78, and D is the size of the section of the inlet section of the contracted section.

A fan section:

the fan section is the fan installation position, and the cross section diameter of the fan section designed by the wind tunnel is D ₅ Length L = 619.6mm ₅ ＝300mm。

The summary parameters are shown in table 1, with detailed dimensions shown in fig. 8.

Table 1 dc closed wind tunnel geometric parameter table

Front inlet section of stabilizing section	1200*1200mm
		Section of stabilizing section	1000*1000mm
Length of stabilizing section	600mm
		Inlet cross section of constriction section	1000*1000mm
Outlet cross section of constriction section	600*600mm
		Length of contracted section	600mm
Section of experimental section	600*600mm
		Experimental section length	1200mm
Diameter of section of fan section	619.6mm
		Test section-fan section length	150mm
Full length	2991.42mm

The wind tunnel test parameters of the invention are set as follows:

according to the actual test conditions, fluid simulation test conditions are determined, pressure parameter information can be obtained by setting condition simulation in CFD calculation software, and corresponding fan power can be calculated. Taking the test condition that the pressure boost H=90 Pa at the fan boundary and the wind speed of the experimental section is 9.8m/s as an example, the parameters at each section of the fluid simulation model are obtained through CFD calculation as shown in table 2, and the theoretical fan power can be calculated correspondingly.

TABLE 2 wind tunnel section State parameters under selected experimental conditions

Cross section of	Inlet cross section	Section of experimental section	Front section of fan	Outlet cross section
					Total pressure (Pa)	0	-2.03453	-3.10199	84.9223
Static pressure (Pa)	-6.64172	-61.319	-86.27	0
					Speed (m/s)	3.25868	9.80412	11.6227	11.7189
Sectional area (m 2)	1.08977	0.36	0.30579	0.301402

According to the energy equation, there is an airflow energy equation at the outlet cross section:

where subscript 1 represents the inlet cross section, subscript 2 represents the fan front cross section, and subscript 3 represents the outlet cross section. H is the full pressure of the fan (the air flow is dependent on the energy obtained by the power of the fan); ρ is the density of the gas; v (V) ₃ The flow speed of the air flow at the outlet section of the fan;

K _2i corresponding to the local drag coefficient and the loss coefficient along the way, K, on each of the inlet section to the fan front section ₂ ＝∑K _2i Is the energy loss coefficient of the suction end; ΔH ₂₃ Energy lost for the compressed air section, +.>

K ₃ The energy loss coefficient of the compressed air section is; ΔH ₁₃ ＝ΔH ₁₂ +ΔH ₂₃ Energy is lost for the airflow in the wind tunnel.

The total pressure relation is as follows:

P ₁₀ -ΔH ₁₃ +H＝P ₃₀

wherein P is ₁₀ Is the air pressure at the inlet cross section; p (P) ₃₀ Is the air pressure at the outlet cross section.

The method can obtain:

ΔH ₁₃ ＝P ₁₀ -P ₃₀ +H＝5.0777Pa

relative error:

the diameter of the fan is selected: 24 inches.

The effective power of the fan is as follows:

N＝H×Q＝90×(0.6×0.6×10)＝324W

wherein Q is the flow rate.

The fan power after considering the fan efficiency is:

efficiency parameter selection principle:

η ₀ the internal efficiency of the fan is generally 0.75-0.85, the small fan takes a low value, and the large fan takes a high value.

η ₁ -mechanical efficiency:

1. the fan and the motor are directly connected 1;

2. coupling is 0.95-0.98;

3. 0.9 to 0.95 percent of the V-belt is connected and taken;

4. take 0.85 with a flat belt drive.

Here, η is taken according to the characteristics of the wind tunnel dynamic system according to the invention ₀ ＝0.8，η ₁ ＝0.98。

And a power system meeting the requirements can be correspondingly configured according to the calculated fan power so as to meet the test conditions.

The main characteristics of the system of the invention are as follows:

1. the three-degree-of-freedom free rotation of the airplane model can be supported;

2. virtual flight supporting mechanism triaxial rotation scope: pitch + -40 deg., roll + -40 deg., yaw + -20 deg.;

3. the system comprises a microminiature flight control system with a built-in model;

4. supporting a code rapid prototype implementation technology, and automatically generating an airborne code;

5. the system is provided with a high-speed remote control link and a data receiving and transmitting link;

6. the ground station detects key parameters such as the position, the gesture, the power state and the like of the servo in real time;

7. the display mode of the data can be customized: curves, pointers, numbers, meters, etc.;

8. the data of the flying process can be recorded as a file for subsequent analysis;

9. the platform can support dynamic base flight simulation and flight control related experimental tests and experimental teaching;

10. the wind tunnel can adjust airspeed to adapt to different experimental conditions.

The working flow of the three-degree-of-freedom virtual flight wind tunnel test system is as follows:

after the fan is started, the air flow enters the wind tunnel from the inlet section and is uniform and stable after passing through the rectifying facilities (the honeycomb device and the damping net). The uniform airflow increases the speed in the contraction section until the experimental section reaches the target speed, and a flow field is simulated around the aircraft scaling model, so that the aircraft scaling model receives aerodynamic moment to generate corresponding angular speed, and the correctness of the aircraft model data and the correctness of the control law obtained by the CFD method can be verified. After the air flow passes through the airplane, the air flow is discharged out of the direct current wind tunnel through the fan section.

Wind tunnel test is a feedback loop, for a new aircraft, first a test is performed that analyzes (boundary) aerodynamic characteristics, but if the aircraft aerodynamic model is already mature, the control law part can naturally be directly performed: firstly, feeding back pneumatic data, namely CFD pneumatic data, namely an aircraft model, an uncontrolled law experiment and an analysis result, and correcting the pneumatic data; if the pneumatic data reach the level considered to be accurate enough, feedback circulation of the control law is carried out, and the control law is corrected by the pneumatic data which are accurate enough, namely an aircraft model, an aircraft control law, a control law experiment with the control law and an analysis result; and if the control law reaches the condition of meeting the standard, the control law is considered to be designed.

The three-degree-of-freedom virtual flight wind tunnel test system is directly directed to the dynamic characteristics of the aircraft on the boundary, can simulate the rotational movement of the aircraft on the boundary state more truly, namely, the result of the movement of the aircraft is directly obtained, then the model (the pneumatic model of the aircraft) is corrected to meet the result, namely, the result of the movement of the aircraft is obtained firstly, the model is not matched when the model is not matched, and the ignored items are too much and too large. The virtual flying tunnel only reserves three rotational degrees of freedom, and because the translational motion has little meaning on the boundary motion of the aircraft, the pneumatic research and the control research are mainly carried out on the rotation, and after the translational motion is canceled, the aircraft is fixed, so that the test is very safe, the risk is lower, and the flow is simple.

The method for obtaining the aircraft model data through the CFD method comprises the following steps: solidworks, catia, and the like, establishes a model of the aircraft, establishes a fluid space, then points with the like grid software to draw grids, and then obtains pneumatic data through the fluent and the like CFD software. The control law is not obtained by CFD, and the control law is designed based on pneumatic data obtained by CFD.

The three-degree-of-freedom virtual flight wind tunnel test system is used for verifying the correctness of the aircraft model data and the correctness of the control law, and is specifically as follows:

the method comprises the steps of overall design of the aircraft, appearance design of the aircraft, obtaining aerodynamic data by CFD, real object of the aircraft or real object of the scaling model, designing control law of the aircraft, and realizing real operation of the aircraft by the control law.

The invention relates to a three-degree-of-freedom virtual flight wind tunnel test system: 1. verifying CFD to obtain pneumatic data and making a certain correction; 2. completing a control law manipulation aircraft scaling model; 3. and verifying the designed aircraft control law.

The control law is based on pneumatic data, steering engine models, etc. Without control laws, control laws can be designed based on aerodynamic characteristics, then through the present invention the design is revised, and existing control laws can be verified through the present invention. The invention relates to a three-degree-of-freedom virtual flight wind tunnel test system which is mainly used for verifying the correctness of designed aircraft model data and the correctness of control laws.

The three-degree-of-freedom virtual flight wind tunnel test system provided by the invention is provided with a platform for rapidly developing a control law, can perform automatic code generation, is downloaded into a controller in an aircraft, and performs the most critical longitudinal, transverse and heading attitude ring control law design verification in a low-speed direct-current closed wind tunnel (low-speed direct-current wind tunnel). After the gesture control design is verified, the free flight test is performed, so that the test flight risk can be reduced by 80%.

In the present specification, each embodiment is described in a progressive manner, and each embodiment is mainly described in a different point from other embodiments, and identical and similar parts between the embodiments are all enough to refer to each other.

The principles and embodiments of the present invention have been described herein with reference to specific examples, the description of which is intended only to assist in understanding the methods of the present invention and the core ideas thereof; also, it is within the scope of the present invention to be modified by those of ordinary skill in the art in light of the present teachings. In view of the foregoing, this description should not be construed as limiting the invention.

Claims

1. The three-degree-of-freedom virtual flight wind tunnel test system is characterized by comprising a low-speed closed direct-current wind tunnel module, a supporting mechanism module, a scaled aircraft model and a microminiature flight control module;

2. The three-degree-of-freedom virtual flying wind tunnel test system of claim 1, wherein the low-speed closed direct current wind tunnel module comprises an inlet stabilizing section, a contraction section, an experiment section and a fan section which are connected in sequence;

3. The three degree of freedom virtual flying wind tunnel test system of claim 2 wherein the support mechanism module comprises a connector, an upper bracket, a lower bracket and a base;

4. The three degree of freedom virtual flying wind tunnel test system of claim 3 wherein the connector comprises an X-direction bearing, a Y-direction bearing, a Z-direction bearing, an inner frame, a middle frame and a Z-plug;

5. The three degree of freedom virtual flying wind tunnel test system of claim 1 wherein the scaled aircraft model includes steering engines, rudders, elevators and ailerons;

6. The three-degree-of-freedom virtual flying wind tunnel test system according to claim 5, wherein the microminiature flying control module comprises a flying control board unit and a code rapid prototyping unit based on Matlab/Simulink;

7. The three degree of freedom virtual flying wind tunnel test system of claim 6 wherein the flight control board unit employs a miniature on-board embedded control system.

8. The three degree of freedom virtual flying wind tunnel test system of claim 7 wherein the flight control board unit comprises a miniature angle of attack sideslip angle sensor, an inertial measurement unit and a computer.