CN107271136A

CN107271136A - A kind of wind tunnel test system

Info

Publication number: CN107271136A
Application number: CN201710508256.7A
Authority: CN
Inventors: 赵霞; 宗宁; 黎军
Original assignee: Shenyang Aircraft Design and Research Institute Aviation Industry of China AVIC; AVIC Shenyang Aerodynamics Research Institute
Current assignee: Shenyang Aircraft Design and Research Institute Aviation Industry of China AVIC; AVIC Shenyang Aerodynamics Research Institute
Priority date: 2015-10-28
Filing date: 2015-10-28
Publication date: 2017-10-20
Anticipated expiration: 2035-10-28
Also published as: CN107389292A; CN107271137B; CN107340117A; CN107152996B; CN107271135B; CN107389296A; CN107389296B; CN107340117B; CN107152996A; CN105203291A; CN107271135A; CN107356403B; CN105203291B; CN107271137A; CN107389292B; CN107356403A; CN107271136B

Abstract

The invention discloses a kind of wind tunnel test system, for promoting model aircraft that experiment is blowed vector in wind-tunnel, the vector promotes the fuselage interior of model aircraft to be provided with a first vector propelling nozzle and a second vector propelling nozzle, current velocity controller is provided with the first vector propelling nozzle and the second vector propelling nozzle, the wind tunnel test system includes one and is fixedly connected with the floor of the wind-tunnel and top plate and vertically disposed support column and a pole for being used to support the vector propulsion model aircraft.The wind tunnel test system of the present invention discharges to form jet effect to obtain jet power using the pressure-air of compressed air source by pipeline to vector propelling nozzle, the jet state of vector propelling motor is simulated, the defect of aerodynamics situation of model aircraft can not be promoted in tunnel simulation vector by overcoming prior art.

Description

Wind tunnel test system

Technical Field

The invention relates to an aeronautical aerodynamic test device, in particular to a wind tunnel test system for a vector propulsion aircraft model.

Background

The wind tunnel test is based on the principle of aerodynamics, an airplane model or parts thereof, such as a fuselage, wings and the like, are fixed in the wind tunnel, and artificial airflow is applied to flow through the airplane model or the parts thereof, so that various complex flight states in the air are simulated, and test data are obtained. The wind tunnel is the most basic test equipment for aerodynamic research and aircraft development, and a large number of tests are required to be carried out in the wind tunnel for developing each novel aircraft. The main purpose of the wind tunnel test is to obtain the change rule of various aerodynamic parameters of the airplane model. Each aircraft was evaluated for flight performance, one of the most important criteria being the aerodynamic performance of the aircraft, in addition to factors such as speed, altitude, aircraft weight and engine thrust. The whole airplane wind tunnel test needs to support the whole airplane model in the wind tunnel, and pressure distribution data of all parts of the whole airplane model under specific flight conditions are measured through pressure testing equipment in an artificial airflow environment, so that the power characteristics of the airplane are obtained.

The vector propulsion technology is a technology for controlling the flight of an airplane in real time by replacing a control surface of an original airplane or enhancing the control function of the airplane through a thrust component generated by the deflection of the thrust of an airplane engine through a spray pipe or a tail jet flow. The vector propulsion technology can change a part of the thrust of the engine into the control force to replace or partially replace a control surface, thereby greatly reducing the radar reflection area; the aircraft can be steered by using the part of the steering force no matter how large the angle of attack is and how low the flying speed is, thereby increasing the maneuverability of the aircraft. Because the steering force is generated directly and the magnitude and the direction are variable, the agility of the airplane is increased, and therefore, the vertical fin can be reduced or eliminated appropriately, and other control surfaces can be replaced. This is advantageous in reducing the detectability of the aircraft and also enables the drag of the aircraft to be reduced and the structural weight to be reduced. Therefore, the use of vector boosting techniques is the best option to resolve design conflicts.

However, in the process of performing a full-aircraft wind tunnel test of an aircraft model, due to the limitations of the size of the wind tunnel and the size of the aircraft model, it is impossible to install a real engine inside the aircraft model, and thus, for the aircraft model using a vector propulsion technology, it is still impossible to simulate the aerodynamic conditions of the vector propulsion aircraft model in the wind tunnel test. That is, in the existing wind tunnel test technology, the airplane model is statically supported (sometimes, the attitude of the airplane can be adjusted, but the airplane cannot simulate the situation with power) in the wind tunnel blowing test process, the airplane model has no power, and the flying state is simulated by using the speed of the flowing air flow relative to the airplane model in the wind tunnel test process. However, for an airplane model adopting a vector propulsion technology, when the magnitude and direction of the propulsion of an airplane engine are changed, a wind tunnel test under a static condition can only simulate the aerodynamic condition under one state. The existing wind tunnel test system can only obtain discrete state data by utilizing a large number of static tests, and then obtain approximate dynamic data for continuously adjusting the magnitude and direction of the propelling force in an interpolation mode, so that the test quantity is huge, time and labor are wasted, the cost is high, the test result is still approximate, and the accuracy is poor.

Disclosure of Invention

The technical problem to be solved by the present invention is to provide a wind tunnel test system to reduce or avoid the aforementioned problems.

The wind tunnel test system is used for carrying out a blowing test on a vector propulsion aircraft model in a wind tunnel, a first vector propulsion spray pipe and a second vector propulsion spray pipe are arranged inside a fuselage of the vector propulsion aircraft model, a first tail spray pipe and a second tail spray pipe which extend out of the tail part of the fuselage of the vector propulsion aircraft model and can adjust the air injection direction are arranged at the tail ends of the first vector propulsion spray pipe and the second vector propulsion spray pipe, and flow speed control devices are arranged in the first vector propulsion spray pipe and the second vector propulsion spray pipe; the flow rate control device includes: a plurality of reduced-bore control panels symmetrically disposed about the inner sidewalls of the first and second vectoring nozzle; a plurality of caliber enlargement control plates symmetrically disposed about the inner sidewalls of the first and second vectoring thrust jets; and an elastic skin covering the caliber reducing control plate and the caliber enlarging control plate; the wind tunnel test system comprises a support column which is fixedly connected with the floor and the top plate of the wind tunnel and is vertically arranged, and a support rod which is used for supporting the vector propulsion aircraft model; one end of the supporting rod is connected with the supporting column, and the other end of the supporting rod is fixed on the airplane body between the first vector propulsion spray pipe and the second vector propulsion spray pipe of the vector propulsion airplane model.

The wind tunnel test system of the invention utilizes the high-pressure air of a compressed air source to release to the vector propulsion spray pipe through the pipeline to form a jet effect so as to obtain jet power, simulates the jet state of a vector propulsion engine, overcomes the defect that the aerodynamic condition of a vector propulsion aircraft model cannot be simulated in a wind tunnel in the prior art, greatly reduces the number of wind tunnel tests by simulating the dynamic state with the propulsion, and has the advantages of closer real condition and higher result accuracy. In addition, the invention also adopts various measures such as entering of the pipeline from the wing, adjustment of the pipeline, electric heating wires, flow rate control devices and the like, thereby further reducing the test cost and improving the test precision.

Drawings

The drawings are only for purposes of illustrating and explaining the present invention and are not to be construed as limiting the scope of the present invention. Wherein,

FIG. 1 is a schematic diagram of a vector-propelled aircraft model according to an embodiment of the present invention;

FIG. 2 shows a side view of a wind tunnel testing system according to an embodiment of the present invention;

FIG. 3 shows a top view of a wind tunnel testing system according to another embodiment of the present invention;

FIG. 4 is an enlarged schematic view of a flow rate control device in a wind tunnel testing system according to yet another embodiment of the present invention;

fig. 5 is a sectional view a-a of the flow rate control device shown in fig. 4.

Detailed Description

In order to more clearly understand the technical features, objects, and effects of the present invention, embodiments of the present invention will now be described with reference to the accompanying drawings. Wherein like parts are given like reference numerals.

Fig. 1 shows a schematic structural view of a vectored-propulsion aircraft model according to an embodiment of the present invention, a first vectored-propulsion nozzle 11 and a second vectored-propulsion nozzle 15 are arranged inside a fuselage of the vectored-propulsion aircraft model 10, and the ends of the first vectored-propulsion nozzle 11 and the second vectored-propulsion nozzle 15 are provided with a first jet nozzle 12 and a second jet nozzle 16 which can adjust the jet direction and extend out of the tail of the fuselage of the vectored-propulsion aircraft model 10. In other words, in order to overcome the defect that the prior art can not simulate the aerodynamic condition of a vector propulsion aircraft model in a wind tunnel, the invention provides the aircraft model with a special structure, and the model can simulate the jet state of a vector propulsion engine and is used for generating jet power in the wind tunnel test process. That is, in the above-described vectoring propulsion aircraft model 10, two vectoring propulsion nozzles 11, 15 are provided, the two vectoring propulsion nozzles 11, 15 can generate jet streams like jet engines, and the jet directions of the tail nozzles 12, 16 thereof are adjustable. Of course, it will be understood by those skilled in the art that the vectoring nozzle 11, 15 of the present invention can only inject a jet similar to a jet engine, without its own rotating parts, and not a real jet engine, and therefore the two exhaust nozzles 12, 16 are not true structural direction-adjustable exhaust nozzles, and the two exhaust nozzles 12, 16 are simply tapered ducts of fixed shape, which are mounted at the rear end of the vectoring nozzle 11, 15, and whose deflection angle, and thus the direction of the jet, can be controlled by conventional hydraulic or electromagnetic operating means (not shown in the figures). The control of the jet direction of the jet nozzles 12, 16 may be performed by conventional techniques, and is not the focus of the present invention, and will not be described in detail herein.

FIG. 2 shows a side view of a wind tunnel testing system according to an embodiment of the present invention; as shown in the drawings, the wind tunnel test system of the present invention may be used for performing a blowing test on the vector propulsion aircraft model 10 of the present invention shown in fig. 1 in a wind tunnel, and the wind tunnel test system includes a support column 400 fixedly connected and vertically arranged with a floor 200 and a ceiling 300 of the wind tunnel, and a support rod 500 for supporting the vector propulsion aircraft model 10; strut 500 is attached at one end to support column 400 and at the other end to the fuselage of vectored propulsive aircraft model 10 between first vectored propulsive nozzle 11 and second vectored propulsive nozzle 15 (fig. 3).

Further, as shown in FIG. 3, a top view of a wind tunnel testing system according to another embodiment of the present invention is shown; the wind tunnel test system further comprises a compressed air source 4 arranged outside the wind tunnel, and a first pipeline 51 and a second pipeline 52 for connecting the compressed air source 4 with the first vectoring propulsion nozzle 11 and the second vectoring propulsion nozzle 15 respectively. That is, in order to simulate the jet engine to generate jet air by the two vectoring nozzle 11, 15, the present invention provides the compressed air source 4, and the high-pressure air of the compressed air source 4 is released to the vectoring nozzle 11, 15 through the pipes 51, 52 to form high-speed air flow, so as to form a jet effect to obtain jet power. In fig. 3, for the sake of clarity, two compressed air sources 4 are shown, which two compressed air sources 4 can be shared in practice, i.e. only one compressed air source 4 is required. Of course, those skilled in the art will appreciate that in the actual wind tunnel test, the pressure level of the compressed air source 4, the lengths, diameters, etc. of the conduits 51, 52 and vectoring propulsion nozzles 11, 15 will need to be precisely calculated and controlled to form the jet air stream to achieve the desired flow rate and flow rate. Those skilled in the art can perform further calculation and control according to the actual situation on the basis of the concept proposed by the present invention, and such calculation and control can adopt the existing conventional technical means, which is not the focus of the present invention and is not described in detail.

In order to avoid excessive disturbance of the wind tunnel flow field by the arrangement of the ducts 51, 52, in a preferred embodiment, the first duct 51 and the second duct 52 enter the fuselage interior of the vector propulsion aircraft model 10 from both ends of the two wings 20 of the vector propulsion aircraft model 10 and are connected to the first vector propulsion nozzle 11 and the second vector propulsion nozzle 15, respectively. As can be seen from fig. 2, with this arrangement of the present embodiment, no additional components are added in the vertical direction of the wind tunnel, and only the horizontally farthest ends of the vector propulsion aircraft model 10 are connected with the pipes 51 and 52 (fig. 3), so that the disturbance generated to the aerodynamic shape of the vector propulsion aircraft model 10 is minimized, which is beneficial for obtaining more accurate test data.

In another embodiment, as shown in FIG. 3, a trim line 30 is connected between first vectoring nozzle 11 and second vectoring nozzle 15, and a flow rate regulating solenoid valve 40 is disposed in trim line 30. The purpose of this arrangement is that the air flow and pressure through the first and second conduits 51 and 52 into the first and second vectoring nozzle 11 and 15 may vary somewhat due to conduit size, connection tightness, etc. if it is desired to simulate conditions of equal thrust from two engines, it is necessary to control the air flow and flow rate through the first and second conduits 51 and 52 very precisely, which is a very cumbersome task and requires a very high level of equipment and personnel. With the arrangement of the present embodiment, the air pressure in the first vectoring nozzle 11 and the air pressure in the second vectoring nozzle 15 can be made to be consistent only by opening the adjusting pipe 30 through the solenoid valve 40, and the same thrust control can be easily achieved through the small design, thereby reducing the control requirements and greatly saving the control time and cost.

In another embodiment, an electrical heating wire 50 is provided around the outside of the first and second vectoring propulsion nozzles 11, 15. The high temperature conditions of the jet engine can be simulated simply by means of the electrical heating wire 50, and of course, more importantly, the first vectoring propulsion nozzle 11 and the second vectoring propulsion nozzle 15 can be heated by means of the electrical heating wire 50, so that the high pressure air flow therein expands when heated to increase the jet velocity. That is, it is very difficult to provide a continuous supersonic jet speed only by the compressed air source 4, and the requirement for the equipment for generating compressed air is very high, and the present embodiment can partially reduce the requirement for the equipment of the compressed air source 4 by the arrangement of the electric heating wire 50, which saves the cost.

Likewise, to further increase the jet velocity within the first and second vectoring nozzle 11, 15, in yet another embodiment, the present invention further provides a flow rate control device 60 within the first and second vectoring nozzle 11, 15, as shown in FIGS. 4 and 5.

That is, fig. 4 is an enlarged schematic view showing a flow rate control device in a wind tunnel test system according to still another embodiment of the present invention; fig. 5 shows a cross-sectional view a-a of the flow control device of fig. 4. it is clear from fig. 4 and 5 that the flow control device 60 of the present invention simulates the configuration of a laval nozzle, resulting in a convergent-mid-divergent-end accelerated flow, the principle of which is well known in the art and will not be described.

Specifically, the flow rate control device 60 of the present invention includes: a plurality of reduced diameter control plates 62 symmetrically disposed about an inner sidewall 61 of the first and second vectoring nozzle 11, 15; a plurality of enlarged bore control plates 63 symmetrically disposed about the inner sidewalls 61 of the first and second vectoring nozzle 11, 15; and an elastic skin 64 covering the aperture reduction control plate 62 and the aperture enlargement control plate 63. The cross-sectional view in fig. 5 schematically shows four enlarged-caliber control plates 63 symmetrically disposed about the inner sidewall 61, from which one skilled in the art can surmise that four reduced-caliber control plates 62 may also be symmetrically disposed about the inner sidewall 61, and for clarity of illustration, the cross-sectional view in fig. 5 does not show the elastomeric skin. The elastic skin 64 can be made of metal aluminum skin with larger elasticity or rubber skin with better resilience, so as to cover the gaps between the control plates 62 and 63 and form a smoother inner wall of the nozzle, which is more favorable for avoiding flow velocity loss and improving the air velocity.

Further, although the structure of the flow rate control device 60 of the present invention can be readily seen in the drawings, for purposes of descriptive and protective claims, the present invention will be described in further detail with reference to the drawings wherein a reduced diameter control plate 62 is disposed in front of an enlarged diameter control plate 63 in the direction of airflow of the first and second vectoring nozzle 11, 15, as indicated by arrows F in fig. 2-4.

Still further, the forward end of the port reduction control plate 62 facing the direction of airflow of the first and second vectoring nozzle 11, 15 is hinged to the inboard wall 61 and the rearward end of the port reduction control plate 62 facing away from the direction of airflow of the first and second vectoring nozzle 11, 15 is hinged to the first hydraulic ram 65.

Further, the rear end of the enlarged diameter control plate 63 facing away from the direction of airflow of the first and second vectoring propulsion nozzles 11, 15 is hinged on the inner sidewall 61, and the front end of the enlarged diameter control plate 63 facing the direction of airflow of the first and second vectoring propulsion nozzles 11, 15 is hinged on the second hydraulic ram 66.

In the wind tunnel test process, in order to control the air flow speed of the first vectoring nozzle 11 and the second vectoring nozzle 15, the first hydraulic rod 65 and the second hydraulic rod 66 can be controlled to adjust the angles of the caliber reducing control plate 62 and the caliber enlarging control plate 63, so that the shape of the laval structure of the nozzles is controlled, and the purpose of flexibly controlling the air flow speed is achieved.

In a word, the wind tunnel test system of the invention utilizes the high-pressure air of the compressed air source to release to the vector propulsion spray pipe through the pipeline to form the jet effect so as to obtain jet power, simulates the jet state of the vector propulsion engine, overcomes the defect that the aerodynamic condition of a vector propulsion aircraft model cannot be simulated in the wind tunnel in the prior art, greatly reduces the number of wind tunnel tests by simulating the dynamic state with propulsion, and has the advantages that the wind tunnel tests are closer to the real condition and the result accuracy is higher. In addition, the invention also adopts various measures such as entering of the pipeline from the wing, adjustment of the pipeline, electric heating wires, flow rate control devices and the like, thereby further reducing the test cost and improving the test precision.

It should be appreciated by those of skill in the art that while the present invention has been described in terms of several embodiments, not every embodiment includes only a single embodiment. The description is given for clearness of understanding only, and it is to be understood that all matters in the embodiments are to be interpreted as including technical equivalents which are related to the embodiments and which are combined with each other to illustrate the scope of the present invention.

The above description is only an exemplary embodiment of the present invention, and is not intended to limit the scope of the present invention. Any equivalent alterations, modifications and combinations can be made by those skilled in the art without departing from the spirit and principles of the invention.

Claims

1. A wind tunnel test system is used for carrying out a blowing test on a vector propulsion aircraft model (10) in a wind tunnel, and is characterized in that a first vector propulsion nozzle (11) and a second vector propulsion nozzle (15) are arranged inside a fuselage of the vector propulsion aircraft model (10), a first tail nozzle (12) and a second tail nozzle (16) which can adjust the air injection direction and extend out of the tail part of the fuselage of the vector propulsion aircraft model (10) are arranged at the tail ends of the first vector propulsion nozzle (11) and the second vector propulsion nozzle (15), and flow speed control devices (60) are arranged in the first vector propulsion nozzle (11) and the second vector propulsion nozzle (15); the flow rate control device (60) includes: a plurality of port reduction control plates (62) symmetrically disposed about an inner sidewall (61) of said first (11) and second (15) vectoring nozzle; a plurality of caliber enlargement control plates (63) symmetrically disposed about an inner sidewall (61) of said first (11) and second (15) vectoring nozzle; and an elastic skin (64) covering the caliber reducing control plate (62) and the caliber enlarging control plate (63); the wind tunnel test system comprises a support column (400) which is fixedly connected with the floor (200) and the top plate (300) of the wind tunnel and is vertically arranged, and a support rod (500) which is used for supporting the vector propulsion aircraft model (10); one end of the supporting rod (500) is connected with the supporting column (400), and the other end of the supporting rod is fixed on a fuselage between the first vector propulsion nozzle (11) and the second vector propulsion nozzle (15) of the vector propulsion aircraft model (10).