CN112098035A

CN112098035A - Aircraft test system

Info

Publication number: CN112098035A
Application number: CN202011302608.1A
Authority: CN
Inventors: 郭洪涛; 路波; 余立; 张昌荣; 吕彬彬; 寇西平; 闫昱; 曾开春; 查俊; 郭鹏
Original assignee: Ultra High Speed Aerodynamics Institute China Aerodynamics Research and Development Center
Current assignee: Ultra High Speed Aerodynamics Institute China Aerodynamics Research and Development Center; High Speed Aerodynamics Research Institute of China Aerodynamics Research and Development Center
Priority date: 2020-11-19
Filing date: 2020-11-19
Publication date: 2020-12-18

Abstract

The invention relates to the field of aircraft testing, in particular to an aircraft testing system based on a half model. The aerodynamic performance testing device for the aircraft comprises a model, wherein the model is arranged in a testing space, the model comprises a fuselage and wings, the testing space comprises a plurality of mutually connected wallboards, the fuselage is arranged close to the wallboards, a plurality of connecting pieces are arranged between the fuselage and the wings, and the fuselage and the wings are arranged in a non-contact manner. The wing is connected with a testing device, the testing device and the wing are kept in the same state and used for testing the aerodynamic performance of the wing, and the testing device and the testing space are in a non-communicated state. The aircraft test system provided by the invention can not only solve the problem that the traditional cushion block method can only isolate the influence of the boundary layer but can not eliminate the influence of the gap flow of the fuselage, but also greatly reduce the influence of the gap flow of the wing, and improve the precision of test data.

Description

Aircraft test system

Technical Field

The invention relates to the field of aircraft testing, in particular to an aircraft testing system based on a half model.

Background

The half model is a common method in aircraft testing, specifically, the half model is that a longitudinal symmetry plane is installed close to a wind tunnel wall plate, in the test process, in order to change the attack angle of the model, the model can freely rotate along with the rotation of a side wall window, a gap with a certain width is reserved between the longitudinal symmetry plane of the model and the side wall of the wind tunnel, airflow flows through the gap inevitably during the blowing test, and the cross flow can interfere the flow-around characteristic of the half model, so that the test result is distorted. In addition, the low-speed and low-energy airflow in the boundary layer of the hole wall can also influence the test result of the semi-model to a certain extent, so that the test result is different from the real aerodynamic characteristic.

In order to solve the above problems, the existing solutions mainly include a boundary layer suction method and a boundary layer blowing method, wherein the boundary layer suction method is a half-model test method that the boundary layer of the hole wall at the position of the half-model and the area near the position is completely or partially sucked by a proper method so as to eliminate or reduce the influence of the boundary layer on the test result. The boundary layer blowing method is characterized in that high-pressure airflow is blown into a boundary layer along a wall surface at the upstream of a half model, so that the energy of the airflow in the boundary layer is increased, the flow speed is increased, the boundary layer is thinned, and the influence of low-energy flow of the boundary layer on the test result of the half model is reduced.

When a boundary layer suction method is adopted, the high-pressure air flow is realized by mainly adopting a vacuum pump, the air extraction rate of the vacuum pump is large enough, or the volume of a vacuum tank is large enough, the vacuum degree is high enough, so that the boundary layer of the cavity wall can be fully sucked, and the boundary layer characteristic of a model zone in the test process is consistent. For longer models, it must be considered that the suction devices are respectively arranged at the front, middle and rear sections to uniformly suck the boundary layer without influencing the flow of the main flow in the test section.

And when the boundary layer blowing method is adopted, the boundary layer thinning effect is weakened along with the increase of the distance from the blowing seam, and the problem of how to uniformly suck the boundary layer for a longer model is solved without influencing the main flow flowing in the test section. Therefore, selecting the height and the shape of the proper blowing guide vane and determining the proper blowing pressure and blowing quantity to ensure that the high-pressure air flow can be uniformly blown in, wherein the flow speed in the boundary layer is close to or equal to but not more than the main flow speed, and the method is one of the key problems to be solved by the boundary layer blowing method.

Therefore, no matter a boundary layer suction method or a boundary layer blowing method is adopted, a complex air suction or blowing device is required, a good blowing/suction disc structure design is selected, and a proper blowing and suction amount is determined, and a test model cannot be too long, otherwise, the technical problems that the boundary layer is uniformly sucked and removed, and the main flow in a test section is not influenced cannot be considered.

In order to eliminate the interference of gap channeling and the boundary layer of the side wall of the wind tunnel on the aerodynamic characteristics of the model, the influence of the boundary layer of the side wall of the wind tunnel on the model streaming can be eliminated by adopting the integrated design of a test model body/boundary layer cushion block. When the model is designed, the thickness of the machine body on one side of the symmetrical surface of the test model is properly increased, so that the actual appearance of the test model can be exposed outside the influence area of the boundary layer of the side wall of the wind tunnel and the clearance cross flow. This is equivalent to adding a cushion block between the test model and the wind tunnel side wall, but the cushion block is integrated with the test model, and the design can avoid forming a new gap between the cushion block and the real model. Although the boundary layer is isolated by adding the cushion block in the traditional cushion block method, a gap still exists between the cushion block and the model, and the influence of gap cross flow cannot be avoided.

And the thickness of the body of the model is indirectly increased by adding the cushion block, so that the appearance of the test model is distorted, and if the aerodynamic force of the body is measured together according to the conventional half-mold force measurement mode, the distortion of the measurement result is caused. Considering that the aim of the half-module static aeroelasticity test is to obtain the influence quantity of the static aeroelasticity of the wing, the fuselage can hardly deform and only plays a role of providing real streaming around the wing. Thus, the semi-static aeroelasticity test may not necessarily measure the aerodynamic forces to which the fuselage is subjected. Therefore, when the model is designed, the fuselage is separated from the wings, the fuselage is directly and fixedly connected with the wind tunnel side wall rotating window, and the wings are connected with the half-mode balance through the balance connector. Therefore, the test aim of only measuring the aerodynamic force borne by the wing and not measuring the aerodynamic force borne by the fuselage is achieved.

Although the adverse effects of the boundary layer of the side wall and the gap cross flow on the static aeroelastic half-die test can be avoided through the cushion block, the half-die static aeroelastic half-die test only requires to measure the aerodynamic force of the wing alone, and in order to prevent the wing and the fuselage from interfering, a certain margin gap still needs to be left between the wing and the fuselage. However, in the high-speed wind tunnel test, the incoming flow velocity of the test can reach more than 300m/s, the gap cross flow velocity can reach dozens of meters per second, and the flow characteristic of the wing is still influenced to a certain extent, so that the influence on the accuracy of the force measurement test of the wing cannot be avoided.

In conclusion, the cushion block method cannot eliminate the influence of the gap channeling of the machine body; the boundary layer suction/blowing method not only needs to be provided with a complex blowing/sucking device, but also has certain limit on the geometric dimension of a test model; the idea of the integrated design of the model fuselage/boundary layer cushion block can avoid the gap cross flow of the fuselage, but cannot avoid the gap cross flow between the wings and the fuselage.

Therefore, a new semi-module support test system for static aeroelasticity test, which not only can effectively isolate boundary layer interference of the side wall of the wind tunnel, but also can effectively avoid gap cross flow influence, needs to be designed.

Disclosure of Invention

The invention aims to provide an aircraft test system which can effectively isolate interference of a boundary layer on the side wall of a wind tunnel and can also effectively avoid gap cross flow influence.

The invention is realized by the following steps:

the utility model provides an aircraft test system for the aerodynamic performance of test aircraft, includes the model, and the model sets up in test space, and the model includes fuselage and wing, and test space includes a plurality of wallboards of interconnect, and the fuselage is close to the wallboard setting, is provided with between fuselage and the wing a plurality of connecting pieces and is the non-contact setting between fuselage and the wing. The wing is connected with a testing device, the testing device and the wing are kept in the same state and used for testing the aerodynamic performance of the wing, and the testing device and the testing space are in a non-communicated state.

Furthermore, the connecting pieces are arranged on gears in the fuselage and also comprise screw rods arranged in the wings, and the screw rods are meshed with the gears.

Furthermore, the screw rod penetrates through an inner hole of the gear and is connected with the gear.

Furthermore, the surface of the inner hole of the gear and the surface of the screw rod are provided with teeth which are matched and connected with each other.

Further, the screw rod is in threaded connection with the outer side of the gear.

Further, the surface of the screw is provided with teeth which are in fit connection with the threads.

Furthermore, the corresponding setting surface of the machine body and the wall plate is provided with a rotating device, and the rotating device realizes the rotation of the machine body.

Furthermore, a filling layer is attached between the machine body and the wall plate.

Furthermore, a rotating device is arranged on the contact surface of the wall plate and the filling layer and connected with the machine body, and the rotating device realizes the rotation of the machine body.

Further, the rotating device comprises a rotating window which is arranged on the same plane with the wall plate, the edge of the rotating window is attached to the edge of the wall plate, and the testing device is arranged on the rotating window in a penetrating mode and connected with the wing.

Further, the rotating window is made of a transparent material.

Furthermore, the testing device comprises a balance and a balance connector, one end of the balance connector is connected with the wing, and the other end of the balance connector is connected with the balance.

The beneficial effect of above-mentioned scheme:

according to the aircraft test system provided by the invention, the connection between the fuselage and the wings can be realized through the connecting piece, the boundary layer of the side wall of the wind tunnel can be effectively isolated, the aerodynamic interference of gap channeling on the static aeroelasticity test half-module can be eliminated, the problem that the influence of the boundary layer can be isolated but the influence of the fuselage gap channeling can not be eliminated in the traditional cushion block method can be solved, the influence of the wing gap channeling can be greatly weakened, and the test data accuracy is improved.

Drawings

In order to more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings needed to be used in the embodiments will be briefly described below, it should be understood that the following drawings only illustrate some embodiments of the present invention and therefore should not be considered as limiting the scope, and for those skilled in the art, other related drawings can be obtained according to the drawings without inventive efforts.

FIG. 1 is a schematic overall structure diagram of a first embodiment of an aircraft testing system provided by the invention;

FIG. 2 is a schematic diagram illustrating the overall structure of a second embodiment of the aircraft testing system provided by the present invention;

FIG. 3 shows a schematic structural view of a connector provided by the present invention;

icon:

100-model;

200-wall plate;

300-a connector;

400-a test device;

500-a rotation device;

110-a fuselage;

120-an airfoil;

210-a filler layer;

310-a gear;

320-screw rod;

410-a balance;

420-balance connector;

510-window rotation.

Detailed Description

In order to make the objects, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are some, but not all, embodiments of the present invention. The components of embodiments of the present invention generally described and illustrated in the figures herein may be arranged and designed in a wide variety of different configurations. Thus, the following detailed description of the embodiments of the present invention, presented in the figures, is not intended to limit the scope of the invention, as claimed, but is merely representative of selected embodiments of the 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.

It should be noted that: like reference numbers and letters refer to like items in the following figures, and thus, once an item is defined in one figure, it need not be further defined and explained in subsequent figures.

In the description of the present invention, it should be noted that the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the drawings or the orientations or positional relationships that the products of the present invention are conventionally placed in use, and are only used for convenience in describing the present invention and simplifying the description, but do not indicate or imply that the devices or elements referred to must have a specific orientation, be constructed and operated in a specific orientation, and thus, should not be construed as limiting the present invention. Furthermore, the terms "first," "second," "third," and the like are used solely to distinguish one from another and are not to be construed as indicating or implying relative importance.

In the description of the present invention, it should also be noted that, unless otherwise explicitly specified or limited, the terms "disposed," "mounted," "connected," and "connected" are to be construed broadly and may, for example, be fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; they may be connected directly or indirectly through intervening media, or they may be interconnected between two elements. The specific meanings of the above terms in the present invention can be understood in specific cases to those skilled in the art.

The following is a detailed description of the wind tunnel experiment system according to the embodiment of the present invention:

referring to FIG. 1, an aircraft testing system for testing the aerodynamic performance of an aircraft includes a model under test 100 and a test space for placing and testing the model 100.

In the present embodiment, the tested model 100 is a test aircraft with a structure of a half model 100, and in the present embodiment, the aircraft is an aircraft with a wing 120 having aerodynamic effect, specifically, an airplane with a half model 100.

The airplane with the semi-model 100 structure comprises wings 120 and a fuselage 110, wherein the fuselage 110 is divided into two parts along a vertical axis in the embodiment, wherein the separated part of the fuselage 110 is connected and fixed with one surface of the wings 120 and a fixed position of a test space.

In this embodiment, the test space is a wind tunnel, and is composed of a plurality of interconnected panels 200, and the body 110 is disposed adjacent to the panels 200. In the present embodiment, the main body 110 is disposed at a side wall plate position, and in other embodiments, the main body 110 may also be disposed at an upper wall plate position and a lower wall plate position.

A plurality of connectors 300 are arranged between the fuselage 110 and the wings 120, and the fuselage 110 and the wings 120 are arranged in a non-contact manner.

Referring to fig. 3, the number of the connecting members 300 in the present embodiment is plural, that is, at least two connecting members are included, and the number may be adjusted in other embodiments. The aircraft comprises a gear 310 arranged in the fuselage 110 and a screw 320 arranged in the wing 120, wherein the screw 320 is meshed with the gear 310.

In this embodiment, the screw 320 penetrates through an inner hole of the gear 310 to be connected with the gear 310, in this embodiment, a tooth structure is disposed on a surface of the inner hole of the gear 310, teeth that are in fit connection with the tooth structure of the inner hole of the gear 310 are disposed on a surface of the screw 320, and connection and relative position fixation between the wing 120 and the fuselage 110 can be achieved through the tooth structures of the screw 320 and the gear 310. The gear 310 in this embodiment is a helical gear 310, and can be connected with the screw 320 in a matching manner.

In other embodiments, the screw 320 may also be a straight rod sleeved with the gear 310, and the sleeved gear 310 is engaged with the gear 310 arranged in the fuselage 110, so as to achieve the connection and the relative position fixation between the wing 120 and the fuselage 110.

In the connecting member 300 provided in this embodiment, the position between the screw 320 or the straight rod and the gear 310 is defined by providing the limiting block at a specific position in the circumferential direction of the screw 320 or the straight rod, that is, the non-contact arrangement between the wing 120 and the fuselage 110 is achieved.

Referring to fig. 3, the present embodiment further provides another structure of the connecting element 300, and the connecting element 300 provided in the present embodiment is also a plurality of connecting elements 300 for connecting the fixed wing 120 and the fuselage 110. In the present embodiment, the connection member 300 includes a screw 320 disposed in the wing 120 and a gear 310 disposed in the fuselage 110, wherein a tooth structure is disposed outside the gear 310, and the screw 320 is circumferentially provided with a tooth pattern matching with the tooth structure. In this embodiment, the gear 310 is a bevel gear 310, and can be coupled with the screw 320.

In other embodiments, the screw 320 may be replaced by a rack, and when the rack is disposed in the wing 120, the gear 310 is a straight gear, and the rack is connected to the gear 310 in a matching manner.

In this embodiment, the body 110 and the wall plate 200 are provided on the installation surface thereof with a rotation device 500, and the rotation device 500 rotates the body 110. The rotating device 500 comprises a rotating window 510 which is arranged on the same plane with the wall plate 200, and the edge of the rotating window 510 is attached to the edge of the wall plate 200. In the present embodiment, the rotating window 510 is made of a transparent material, which enables observation of the rotating state of the fuselage 110 and the state of the wing 120. The rotating window 510 is externally connected with a driving device, the rotating window 510 can rotate at a certain angle, the rotating angle of the rotating window 510 is determined by the space distance of the testing space, and when the transverse and vertical lengths of the testing space are greater than the length of the machine body 110, the rotating window 510 can rotate by 360 degrees.

The wing 120 is connected with the testing device 400, the testing device 400 and the wing 120 are kept in the same state for testing the aerodynamic performance of the wing 120, the testing device 400 and the testing space are in a non-communication state, namely the testing device 400 is not exposed in the testing space, and the airflow generated by testing cannot affect the testing device 400.

The testing device 400 is disposed through the rotating window 510 and connected to the fuselage 110 and the wing 120, and in this embodiment, the testing device 400 is used to individually test the aerodynamic data change of the wing 120 under the airflow condition, so as to individually test the aerodynamic properties of the wing 120.

In this embodiment, the testing device 400 includes a balance 410 and a balance connector 420, wherein one end of the balance connector 420 is connected to the wing 120, and the other end is connected to the balance 410. Wherein balance 410 is the primary testing tool for testing and recording the pneumatic parameters at specific conditions at a specific time. The balance connector 420 is a connector 300.

Example 2

The present embodiment differs from embodiment 1 in that a filler layer 210 is provided between the body 110 and the wall plate 200 in a bonded manner, that is, the body 110 is provided in a bonded manner on one side of the filler layer 210, and the wall plate 200 is provided in a bonded manner on the other side of the filler layer 210.

In this embodiment, the filling layer 210 and the fuselage 110 are of a split structure, that is, when a gap occurs between the fuselage 110 and the wall plate 200, a gap blocking effect can be achieved, so that the overall influence of an airflow generated by the gap on the fuselage 110 is solved, and the influence of flutter generated by the fuselage 110 due to the influence of the airflow on the wings 120 is indirectly solved.

In the present embodiment, the contact surface of the rotating means 500 and the filling layer 210 is disposed in contact, and the horizontal surface of the rotating means 500 is disposed in parallel with the outermost surface of the filling layer 210. The rotating device 500 is connected to the filling layer 210 and the body 110, and the rotating device 500 can rotate the body 110.

In this embodiment, the rotating device 500 includes a rotating window 510, and an edge of the rotating window 510 is disposed to be attached to an edge of the filling layer 210.

The aircraft test system provided by the embodiment is based on embodiment 1, and the gap between the fuselage 110 and the wall plate 200 can be minimized by arranging the filling layer 210, so that the influence caused by the outside airflow is minimized.

In the present embodiment, the above description is only a preferred embodiment of the present invention, and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention. In the present embodiment, the above description is only a preferred embodiment of the present invention, and is not intended to limit the present invention, and various modifications and changes may be made by those skilled in the art. Any modification, equivalent replacement, or improvement made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims

1. An aircraft test system is used for testing the aerodynamic performance of an aircraft and comprises a model, wherein the model is arranged in a test space, the model comprises a fuselage and wings, and the test space comprises a plurality of mutually connected wall plates; the wing is connected with a testing device, the testing device and the wing are kept in the same state and used for testing the aerodynamic performance of the wing, and the testing device and the testing space are in a non-communicated state.

2. The aircraft testing system of claim 1, wherein a plurality of said connectors comprise gears disposed within said fuselage and further comprising screws disposed within said wings, said screws being in meshing engagement with said gears.

3. The aircraft testing system of claim 2, wherein the screw is threaded through the gear bore and is coupled to the gear.

4. The aircraft testing system of claim 3 wherein said gear bore surface and said screw surface are provided with teeth that cooperatively engage one another.

5. The aircraft testing system of claim 2, wherein said screw is threaded outboard of said gear.

6. The aircraft testing system of claim 5 wherein said screw surface is provided with teeth in mating engagement with said threads.

7. The aircraft testing system according to any one of claims 1 to 6, wherein a rotating device is arranged on the corresponding arrangement surface of the fuselage and the wall plate, and the rotating device realizes rotation of the fuselage.

8. The aircraft testing system according to any one of claims 1 to 6, wherein a filling layer is attached between the fuselage and the wall plate.

9. The aircraft testing system of claim 8, wherein the face of the wall plate contacting the infill layer is provided with a rotating device, the rotating device being connected to the fuselage, the rotating device effecting rotation of the fuselage.

10. The aircraft testing system of claim 9, wherein the rotating device comprises a rotating window disposed coplanar with the wall plate, an edge of the rotating window is attached to an edge of the wall plate, and the testing device is disposed through the rotating window and the fuselage and connected to the wing.

11. The aircraft testing system of claim 10, wherein the louver is comprised of a see-through material.

12. The aircraft testing system of claim 10, wherein the testing device comprises a balance and a balance connector, one end of the balance connector is connected to the wing, and the other end of the balance connector is connected to the balance.