CN112849389B - Dynamic stall control method based on dynamic droop of wing leading edge - Google Patents
Dynamic stall control method based on dynamic droop of wing leading edge Download PDFInfo
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- CN112849389B CN112849389B CN202110113563.1A CN202110113563A CN112849389B CN 112849389 B CN112849389 B CN 112849389B CN 202110113563 A CN202110113563 A CN 202110113563A CN 112849389 B CN112849389 B CN 112849389B
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Abstract
The invention discloses a dynamic stall control method based on dynamic droop of a wing leading edge, and belongs to the technical field of aircraft flow control. In the pitching oscillation process of the wing, when the attack angle of the wing is increased to a set value, controlling the droop of the wing leading edge to reduce the local effective attack angle of the leading edge; and in the process of reducing the attack angle of the wing, the leading edge of the wing is controlled to recover from a sagging state, so that the aerodynamic characteristics of the original wing profile are maintained. The invention effectively controls the dynamic stall in the pitching oscillation process of the wing by reasonably controlling the dynamic droop of the wing leading edge.
Description
Technical Field
The invention belongs to the technical field of aircraft flow control, and particularly relates to a dynamic stall control method based on dynamic droop of a wing leading edge.
Background
The dynamic stall refers to unsteady flow separation and stall phenomena of an aerodynamic component in a dynamic motion process, and can occur on aerodynamic devices such as helicopter rotors, wind driven generator blades and the like. For helicopters in the forward flight regime, dynamic stall is likely to occur due to the cyclic variation of the angle of attack (i.e. pitch oscillation) of the rotor blades during cyclic pitching, and at relatively low incoming flow velocities. Dynamic stall produces pressure fluctuations and sudden load changes on the upper surface of the wing, which may cause a significant increase in normal force and a large low head moment, which has a large impact on the aerodynamic characteristics of the aircraft. There is therefore a need for flow control to suppress dynamic stall and improve aircraft aerodynamics.
There are many current dynamic stall flow control techniques, which can be divided into two broad categories, passive and active, depending on whether energy is consumed, the former including vortex generators, airfoil leading edge deformation, etc.; the latter include synthetic jets, slat/flap dynamic deflection, etc.
In various dynamic stall flow control technologies, the vortex generator is single in optimal working condition and is not suitable for complex and variable flight environments; the synthetic jet control method has low jet momentum and weak control effect.
Disclosure of Invention
In view of this, the invention provides a dynamic stall control method based on dynamic droop of a wing leading edge, which effectively controls dynamic stall in a wing pitching oscillation process by reasonably controlling the dynamic droop of the wing leading edge.
A dynamic stall control method based on dynamic droop of a wing leading edge is characterized in that in the pitching oscillation motion process of a wing, when the attack angle of the wing is increased to a set value, the droop of the wing leading edge is controlled to reduce the local effective attack angle of the leading edge; and in the process of reducing the attack angle of the wing, the leading edge of the wing is controlled to recover from a sagging state, so that the aerodynamic characteristics of the original wing profile are maintained.
Further, the attack angle of the wing is larger than a certain preset value alpha def When the airplane wing needs to be deformed, the front 1/4 chord length part of the wing droops in a body coordinate system, the wing remains static in an incoming flow coordinate system, and the local attack angle at the front part of the wing is fixed to be the attack angle alpha at the deformation starting moment def The local angle of attack at the front of the wing remains unchanged during deformation.
Further, in the coordinate system with the body, the displacement perpendicular to the chord direction of the wing at each point of the front part of the wing:
in the formula, alpha 1 Is the amplitude of the pitching oscillation movement of the wing, omega is the angular frequency of the pitching oscillation of the wing,
the deformation initiation phase angle, t is time.
Further, the change law of the angle of attack is as follows:
wherein: alpha is alpha 0 Is the average angle of attack of the airfoil pitch oscillation.
Further, the amount of deformation of the airfoil leading edge droop does not offset the change in the angle of attack of the airfoil.
Has the beneficial effects that:
1. according to the dynamic stall control method based on the dynamic droop of the wing leading edge, the effective attack angle of the front part of the wing is reduced, so that the pressure distribution of the wing is more reasonable, the adverse pressure gradient near the leading edge is reduced, the flow separation is not easy to occur, and the dynamic stall phenomenon in the pitching oscillation process of the wing can be effectively inhibited.
2. The invention can effectively control the dynamic stall under different flight conditions by adopting different leading edge droop forms and control parameters.
Drawings
FIG. 1 is a schematic view of a leading edge droop mode and coordinate system definition;
FIG. 2 is a schematic view of the variation law of the angle of attack of the leading edge portion of the airfoil;
FIG. 3 is a graph of lift coefficient with and without control;
FIG. 4 is a graph of drag coefficient with and without control;
fig. 5 is a graph of pitch moment coefficient versus angle of attack hysteresis.
Detailed Description
The invention is described in detail below by way of example with reference to the accompanying drawings.
The invention provides a dynamic stall control method based on dynamic droop of a wing leading edge, which is characterized in that in the pitching oscillation motion process of a wing, when the attack angle of the wing is increased to a certain value, the wing leading edge is enabled to droop, and the local effective attack angle of the leading edge is reduced, so that the pressure distribution of the wing is more reasonable, the counter pressure gradient near the leading edge is reduced, and the separation is not easy to occur; and in the process of reducing the attack angle of the wing, the wing leading edge is recovered from the sagging state, so that the aerodynamic characteristics of the original wing profile are well maintained.
The wing leading edge droop pattern is shown in figure 1. Fig. 2 shows the variation of the angle of attack of the whole wing and the leading edge part with time, and the horizontal axis represents the phase angle, which is the product of time and angular frequency, phi = ω t, so that the time can be expressed under the condition of constant angular frequency. The vertical axis represents angle of attack; the sine curve in the figure represents the change rule of the attack angle of the whole wing, and the straight line segment AC represents the change rule of the attack angle of the leading edge part of the wing in the deformation process. By definition, during the pitching oscillatory movement of the wing, when the angle of attack increases above a certain predetermined value α def At the front 1/4 chord length part of the wingDrooping in a coordinate system of the body, but keeping still in a coordinate system of the incoming flow, and the local attack angle of the front part of the wing is fixed as the attack angle (alpha) of the deformation starting moment def ). The local angle of attack at the front of the wing remains constant during deformation, as shown by the straight line segment AC in fig. 2. While the overall angle of attack is reduced from a maximum value to alpha when the wing is in pitch oscillation def When the wing is in use, the front part of the wing is restored from a drooping state and synchronously moves along with the whole.
Formula (1) is the displacement perpendicular to the chord direction of the wing of each point in the front part of the wing in a random coordinate system:
in the formula, alpha 1 Is the amplitude of the pitching oscillation of the wing, omega is the angular frequency of the pitching oscillation of the wing,
the deformation initiation phase angle, t is time.
Fig. 3 to 5 show that the flow control is performed by using the dynamic stall control method based on the dynamic droop of the leading edge of the wing, which is obtained by numerical simulation calculation, and under the condition of subsonic flight (mach number Ma = 0.3), the NACA 0012 airfoil makes pitching oscillation motion according to the following attack angle change law:
and comparing a hysteretic curve of the airfoil lift coefficient, the drag coefficient and the pitching moment coefficient, which are obtained by numerical simulation calculation along with the change of the attack angle, with an uncontrolled condition, wherein the parameters of the controller are as follows: phase angle of onset of deformation
It can be seen that the change of aerodynamic coefficient of the airfoil profile with the attack angle after the flow control is added shows a more gradual hysteresis loop compared with the uncontrolled condition; angle of attack in pitch oscillatory motionThe lift coefficient does not drop suddenly in the process of reduction; in the process of increasing the attack angle, the resistance divergence and the pitching moment negative peak are effectively inhibited, which shows that the dynamic stall control method based on the dynamic droop of the wing leading edge can effectively inhibit the dynamic stall phenomenon in the pitching oscillation process of the wing.
According to the invention, the wing leading edge dynamically sags, and the deformation mode of the leading edge in the sagging process can be defined differently, for example, the sagging deformation of the leading edge can not be offset with the change of the wing attack angle, the sections A-C in the attached figure 2 can not be straight lines, and the deformation expression is not limited to the formula (1).
In summary, the above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention. 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 (4)
1. A dynamic stall control method based on dynamic droop of a wing leading edge is characterized in that in the pitching oscillation process of a wing, when the attack angle of the wing is increased to a set value, the droop of the wing leading edge is controlled to reduce the local effective attack angle of the leading edge; in the process of reducing the attack angle of the wing, the leading edge of the wing is controlled to recover from a drooping state, so that the aerodynamic characteristics of the original wing profile are maintained; in a random coordinate system, the displacement amount of each point at the front part of the wing, which is vertical to the chord direction of the wing, is as follows:
in the formula, alpha 1 Is the amplitude of the pitching oscillation movement of the wing, omega is the angular frequency of the pitching oscillation of the wing,
the deformation initiation phase angle, t is time.
2. The machine-based of claim 1The dynamic stall control method for the dynamic droop of the leading edge of the wing is characterized in that the attack angle of the wing is larger than a preset value alpha def When the wing is in use, the front 1/4 chord length part of the wing droops in a following coordinate system, the wing is kept static in an incoming flow coordinate system, and the local attack angle at the front part of the wing is fixed as the attack angle alpha at the deformation starting moment def The local angle of attack at the front of the wing remains unchanged during deformation.
3. The dynamic stall control method based on dynamic droop of the leading edge of the airfoil as claimed in claim 2, wherein the angle of attack variation law is as follows:
wherein: alpha is alpha 0 The average angle of attack of the airfoil pitch oscillations.
4. The method for dynamic stall control based on dynamic droop of the leading edge of an airfoil as claimed in claim 3, wherein the amount of deformation of the dynamic droop of the leading edge of the airfoil does not offset the change in angle of attack of the airfoil.
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| IL121164A (en) * | 1997-06-26 | 2002-03-10 | Univ Ramot | Airfoil with dynamic stall control by oscillatory forcing |
| GB0700604D0 (en) * | 2007-01-11 | 2007-02-21 | Airbus Uk Ltd | A leading edge structure for an aerofoil |
| RU2361776C1 (en) * | 2008-03-14 | 2009-07-20 | Сергей Федорович Бокарев | Method of controlling airflow about wing to improve basic aerodynamic characteristics, and aircraft to this end |
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| CN111392037B (en) * | 2020-03-30 | 2021-05-18 | 南京航空航天大学 | Helicopter rotor dynamic stall control method and system |
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