Aircraft wing tip

FIELD: transport.

SUBSTANCE: invention relates to aircraft engineering. Aircraft wing tip comprises channel with inlet and outlet holes. Inlet hole represents an air intake arranged on lower front surface and communicated with conical channel with end cross section with diametre of 0.05 to 0.2 of the length of chord of wing tip section and is located at the distance of 0 to 0.2 of chord length from rear edge along flow direction. Channel axis is located on 0 to 0.2 of chord length above the plane of chords. Channel midsection accommodates disk-like rotary flap with its axis perpendicular to channel axis. Flap can be rotated by airflow.

EFFECT: higher lift and reduced drag.

3 cl, 7 dwg

 

The invention relates to the field of aviation, in particular, to devices for improving the aerodynamic quality and reduce the intensity of the vortex trail, and can be used in the construction of load-bearing surfaces of aircraft, such as aircraft wings, helicopter main rotor, etc.

The goal of improving aerodynamic qualities is traditional for aviation, because it directly affects the efficiency of the aircraft as a vehicle.

The problem of reducing the intensity of the vortex trail became especially acute with the growth takeoff weight of passenger and transport aircraft. Vortex trail, continuing long after the passage of heavy aircraft is very dangerous for flying on the back of the device, especially when a big difference in the flight weight. This circumstance forces to increase the range of motion of the aircraft at major airports, which is one of the main factors that reduce their effectiveness.

Famous wing tip of the aircraft (RF patent No. 2116936 C1, 24.05.1996, IPC VS 23/06), which is an end part of the wing is installed on the generator vortex flow in the opposite direction of rotation of the trailing vortex, with the gondola with the input device and the nozzle, while the gondola is equipped on its the first surface of the intake devices for selecting the part of air flow and the direction of its inward with the tangential component of the velocity vector for the attachment of the internal flow in the direction opposite rotation of the vortex that occurs at the end of the wing and the trailing part of the wing with longitudinal elements made in the form of the upper bounding surface, protruding above the wing to a height approximately equal to the thickness of the profile, and the lower guide surface protruding downward from the lower surface of the wing at ~1/3 of the thickness profile, interacting with the vortex flow created by the generator.

The disadvantage of this design is that the stream flowing from the bottom surfaces of the wing on the upper, the effect on the vortex flow from the generator at a right angle, "crushing" him and creating additional resistance, reducing the effectiveness of the ending. In addition, create additional resistance and suction device. This reduces the gain in aerodynamic quality.

The famous ending (RF patent No. 2233769 C1, 19.03.2003, IPC VS 23/06), in which the forward end section of the wing has a vortex flow generator having a gondola with the input device and the nozzle, and the tail end section of the wing is made with a cylindrical casing with an open radial cross-section forming a longitudinal cut from the top and bottom edges on the outer side surface and fixed middle part at the end of the wing. In front of her the Asti cover made paired inner surface with the outer surface of the nozzle, and at the tail end is an open end.

The disadvantage of this design is that the vortex flow generator has an intrinsic resistance, and flow around sharp edges of the casing followed by vortex formation, which also increases the resistance and reduces the effectiveness of the ending.

Known limit vortex generator (patent RF №2148529 C1, 11.12.1997, IPC VS 23/06)containing the gondola with a high ratio of the squares of the bore and midsection, the output device with a diffuser, nozzle and torquing device. Such a generator mounted on the end portion of the wing creates a vortex flow, the opposite end of the vortex bearing surface, which should improve its aerodynamic characteristics.

The disadvantage of this design is the increased own resistance to twisting device.

The known method and device for the elimination of vorticity in the region of the rear edge and a corresponding reduction in induced drag (patent US 7,134,631 B2, from 14.11.2006 IPC VS 21/06), according to which in the vicinity of the rear edge is a boundary layer suction, and the ending - it managed sduw that allows you to reduce the intensity of the vortex trail and to reduce inductance and noise.

To fake the headed the remainder of this method include the difficulty of its implementation and the high cost of power in organization management boundary layer (in the energy sense is equivalent to increasing aerodynamic drag).

The number of known devices and methods intended to mitigate vortex trail by injection of air into the core of the vortex or in the region of its formation (see, for example, the review of such devices in the work R.Earl Dunham, Jr. "Unsuccessful concepts for aircraft wake vortex minimization,", NASA SP-409, 1977). In particular, the known device and method destabilization of trailing vortices (patent WO 2008/051269 A3, 02.05.2008, IPC VS 23/06), in accordance with which of the wingtips is controlled by blowing jets of air to make disturbances in the emerging end of the vortex and thereby accelerating its destruction.

The disadvantages of all methods of this kind can be considered the complexity of their implementation, the cost of power for the organization of blowing and the difficulty of developing the optimal control algorithm blowing.

Known cylindrical wing tip with spiral slots (patent US 6,892,988 B2, 17.05.2005, IPC B64D 27/02), adopted for the prototype, designed to reduce induced drag and increase aerodynamic quality. In this ending, the air entering through the inlet cylindrical ends, is mixed with air passing through the spiral slit on the side surface of the cylinder and receiving while the vorticity, the opposite end of the vortex. In the total circulation of the vortex is reduced, which should lead to with whom iginio induced drag of the wing and increase glide.

It should be noted that the total circulation of wing vortex sheet is uniquely determined by the value created by the wing lifting force. Attempts to reduce the circulation of the vortex at a given value of the lifting force will only lead to a redistribution of the circulation along the span, which can have both positive and negative effects. In addition, the disadvantage of this design is the additional resistance resulting from separated flow past sharp edges cylindrical ends and edges of spiral slits, which reduces the efficiency of the device.

The objective of this invention is to provide such a terminal device bearing surface (wing, rotor and so on), which, affecting the process of formation of the vortex, would improve aerodynamic characteristics and to increase the level of safety of the flight.

The technical result consists in increasing the lift coefficient and the reduction of the drag coefficient of the bearing surface by reducing tear-off region in the zone of formation of the vortex.

The technical result is achieved by the fact that the ending of the bearing surface of the aircraft, containing a channel with an inlet and outlet, the inlet hole is made in the form of vozduhozabora the ka, located on the lower front surface of the end portion of the bearing surface, and is connected with a tapered channel, the output section which has a diameter equal to 0.05÷0.2 chord length of an end section of the bearing surface and placed at a distance of 0÷0.2 chord length from the trailing edge in the flow direction, and the axis of the channel is located at a distance of 0÷0.2 chord length above the plane of the chord.

The technical result is also achieved by the fact that in the middle section of the channel is set disc-shaped rotary valve, the axis of which is perpendicular to the axis of the channel, and the rotation of the valve about its axis is enforced.

The technical result is also achieved by the fact that the disc-shaped flap is S-shaped for rotation about its own axis under the influence of passing through the flow channel.

The invention is illustrated by drawings:

figure 1 - General view of the wing tip (rotor blades) of the aircraft, made in accordance with the invention;

figure 2 - photograph of the model of a rectangular wing with the proposed tip on tests in a wind tunnel;

figure 3 - photograph of the tested sets of replaceable tips;

figure 4 is a photograph of the valve installed in the channel ending;

5 is a graph illustrating the increase of maximum lift with the crystals when using the ending;

6 is a graph illustrating the reduction of the profile drag of the wing when using the ending;

7 is a graph illustrating the effect of rotating the valve characteristics of the wing.

Experimental studies show that in the vicinity of the rear edge of the wing tip in the zone of formation of the core vortex there is an extensive area of low pressure. In the proposed ending this fact is used to feed into the core of the vortex maloosmyslennoe air from the fore part of the wing tips, ie a kind of passive (no energy cost) blowing in the core of the vortex. To this end, the lower surface of the end portion of the wing in the vicinity of the leading edge 1 is the inlet 2 (figure 1), through which malovany air is introduced into the channel 3 is formed by the ending. In the tail ending this channel has a conical shape, and the size of the outlet channel 4 and its location relative to the rear edge of the wing to be specified for the specific layout of the wing and the selected flight mode so that it was located in the area of maximum rarefaction. In the invention, the diameter of the outlet 4 D=0,05÷0,2 chord length of an end section of the bearing surface and placed at a distance Δ=0÷0.2 chord length from the trailing edge in the flow direction, and the axis to the channel 3 is located at a distance h=0÷0.2 chord length above the plane of the chord. As a result, the air from the nasal part of the ending intensively injected into the emerging core of the vortex and the size of the nucleus increases. Since the circulation of the vortex is determined by the lifting force of the wing, the increase in the size of the kernel leads to a decrease in the maximum peripheral speed, reduce local bevels flow, the reduction of the region of separated flow and, consequently, increase lift and decrease the drag of the wing.

Reducing the maximum peripheral speed significantly reduce the risk when getting in the vortex trail of another aircraft. In the case of such ending on the main rotor blades of the helicopter decrease of the maximum peripheral speed of the vortex leads to a reduction of the local bias flow from the vortex blades on the surface of the next blade and, accordingly, reduction of variable loads on it, which improves the aerodynamic characteristics of the rotor.

The presence of the channel 3 in the tip makes it easy to implement the control of the intensity of the injection of air into the kernel. Set channel 3 rotary disc-shaped valve 5 allows virtually no energy to change the intensity of the injection from zero to maximum. This introduction of perturbations in the core of the vortex can accelerate it in the dissipation and reduce the length of the vortex trail of the aircraft. When forced rotation of the valve 5 electric or other motor, adjustable frequency insertion perturbations to improve their effectiveness. If the valve 5 to give an S-shaped curvature, rotation of the valve 5 can be carried out in the absence of the drive, under the action of flowing through the flow channel. The frequency of rotation of the valve 5 will be determined by its form.

Useful effect of the invention was confirmed in tests in a wind tunnel model of a rectangular wing with aspect ratio (the ratio of wing length to the length of the chord)equal to 5 (figure 2). To study the effect of size and location of the outlet channel model contained replaceable ending 6 (figure 1), the set of which is shown in figure 3. Tests were conducted both with and without flaps 5 and 5 valve, shown in figure 4. When this valve 5 is rotated with controlled frequency electric motor mounted in the compartment of the wing. Tests conducted when the flow velocity V in the range from 20 to 30 m/s (figure 5, 6 marked curves: 7 - speed flow of 20 m/s, 8 - 25 m/s, 9 - 30 m/s)showed that the application of the specified ending at optimum values of the diameter of the outlet Dinleads to an increase of 1÷2% of the maximum lift coefficient Cyamax(figure 5) and a significant decrease in the resistivity profile is tyuleniy wing C xpmin(6). 7 shows the dependence of the aerodynamic coefficients from the frequency of rotation of the flap 5. The numbers represent the frequency of rotation n of the coefficients: 10 - Cha, 11 - mz, 12 - Cstorage, 13 - Cua. The experimental points corresponding to the frequency of rotation n=0, obtained in the absence of valves. It is seen that the use of rotary valves not only increases the drag coefficient of the wing Withstorage(curves 10, 12), but leads to some of its lower and higher lift coefficient

Withua(curve 13). Multiple increases the rate of moment mz(curve 11).

1. Aconcave the bearing surface of the aircraft, containing a channel with an inlet and outlet, wherein the inlet opening is made in the form of an air inlet located on the lower front surface of the end portion of the bearing surface and connected to a tapered channel, the output section which has a diameter equal to 0.05÷-0,2 chord length of an end section of the bearing surface, and placed at a distance of 0÷0.2 chord length from the trailing edge in the flow direction, and the axis of the channel is located at a distance of 0÷0.2 chord length above the plane of the chord.

2. Concave bearing surface of an aircraft according to claim 1, characterized in that on average the treatment channel set disc-shaped rotary valve, the axis of rotation which is perpendicular to the axis of the channel with the possibility of forced rotation.

3. Ending the bearing surface of an aircraft according to claim 2, characterized in that the disc-shaped flap is S-shaped for rotation about its own axis under the influence of passing through the flow channel.



 

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