V-shaped empennage control surface course control expansion control method

Abstract

The invention belongs to the technical field of flight mechanics, and particularly relates to a V-shaped empennage control surface course control extension control method which comprises the steps that a left control surface deflection angle reaches an upper limit, a right control surface deflection angle continues to deflect leftwards, and a right empennage generates right lower lift increment perpendicular to the incoming flow direction; the right empennage can simultaneously generate a vertically downward force and a head-up torque relative to the gravity center of the whole aircraft; a downward head lowering moment is generated through the longitudinal periodic pitch change of the rotor wings and the tilt angle of the rotor wings, so that the pitching moment is balanced; the vertical force balance of the whole aircraft is achieved by adjusting the attack angle of the aircraft body.

Description

Technical Field

[0001] The present invention belongs to the technical field of flight mechanics, and in particular relates to a V-tail rudder surface heading manipulation expansion control method. Background Art

[0002] Compared with conventional tail fins, V-tail fins are increasingly used in tilt-rotor aircraft designs due to their smaller wetted area and lighter structural weight. V-tail fins are two wing surfaces arranged in a V shape at the tail of the fuselage. The wing surfaces consist of fixed stabilizers and articulated rudders. Their function is to control the longitudinal direction and heading of the aircraft. When controlling the longitudinal direction, the two rudder surfaces deflect in the same direction, and when controlling the heading, the two rudder surfaces deflect in opposite directions. The final deflection angle of the rudder surface is the result of the combination of longitudinal and heading control.

[0003] When a tilt-rotor aircraft is in stable level flight without sideslip, the deflection angles of the two tail rudder surfaces are consistent. At this time, the lateral force components generated by the upper dihedral angles of the two tail rudders cancel each other out, and only a longitudinal resultant force is generated. If a side gust of wind encounters this time, heading control is required on this basis. Because the two rudder surfaces deflect in opposite directions, the deflection angle of one side increases and the deflection angle of the other side decreases. The side with the increased deflection angle will first reach the upper limit of the rudder surface deflection range. At this time, the heading control capability reaches the limit. Although the deflection angle of the other side's tail rudder surface has not reached the limit, it cannot continue to be manipulated to generate lateral force and yaw moment, because the change in the deflection angle of the unilateral tail rudder surface will simultaneously generate lateral force and longitudinal force, and the longitudinal force generates a longitudinal moment, making it impossible for the aircraft to complete longitudinal balance, thereby losing balance.

[0004] At present, the heading control limit is generally determined by the tail rudder on one side reaching the upper limit of the deflection angle. In fact, there is still room for control of the tail rudder on the other side. How to balance the longitudinal torque caused by the deflection of the unilateral rudder surface has become the key to expanding the heading control capability of the V-type tail rudder surface. Summary of the Invention

[0005] The V-tail rudder heading control expansion control method of the present invention is that after the deflection angle of the V-tail rudder on one side reaches an upper limit, the heading control can continue to change the deflection angle of the rudder on the other side to generate a yaw moment, and the longitudinal moment thus generated is balanced by changing the longitudinal cyclic pitch and the rotor tilt angle.

[0006] In order to achieve the above-mentioned purpose, the present invention adopts the following technical solutions to achieve it.

[0007] A method for controlling the heading manipulation of a V-tail rudder surface is provided. When a tiltrotor aircraft is in stable level flight, the V-tails on both sides are simultaneously deflected downward by the same angle. The method comprises: S1, when a tiltrotor aircraft encounters a right-hand wind, the fuselage generates a right yawing moment. To regain balance, the left V-tail rudder is manipulated to deflect downward, and the right V-tail rudder upward. The left V-tail rudder deflection angle first reaches the upper limit of the rudder deflection angle range. The right V-tail rudder deflection angle is then manipulated upward by a first angle. The right V-tail then generates a first lift increment perpendicular to the incoming flow direction and to the lower right. This first lift increment is decomposed into a horizontal rightward lateral force and a left yawing moment, as well as a vertical downward vertical force and a pitch-up moment relative to the center of gravity of the entire aircraft. The left yawing moment is used to offset the right yawing moment of the fuselage generated by the right-hand wind. S2, to balance the horizontal rightward lateral force, vertical downward vertical force, and pitch moment relative to the aircraft's center of gravity generated by the right V-tail, by changing the longitudinal cyclic rotor pitch and rotor tilt angle to generate a downward nose-down moment on the fuselage, thereby achieving pitch moment balance; lateral force balance is achieved by adjusting the fuselage roll angle, and vertical force balance is achieved by adjusting the fuselage angle of attack; S3: When the tilt-rotor aircraft encounters a left-hand wind, a left yawing moment is generated on the fuselage. To regain balance, the left V-tail rudder is manipulated to deflect upward and the right V-tail rudder downward. The right V-tail rudder deflection angle first reaches the upper limit of the rudder deflection angle range. The left V-tail rudder deflection angle is then manipulated upward by a second angle. The left V-tail then generates a second lift increment perpendicular to the incoming flow and directed downward to the left. This second lift increment is decomposed into a horizontal leftward lateral force and a right yawing moment, as well as a vertical downward vertical force and a pitch-up moment relative to the center of gravity of the aircraft. The right yawing moment is used to offset the left yawing moment of the fuselage generated by the left-hand wind. S4, in order to balance the horizontal left lateral force, vertical downward vertical force and pitch-up moment relative to the center of gravity of the entire aircraft generated by the left V-shaped tail, the longitudinal cyclic pitch and rotor tilt angle are changed to make the fuselage generate a downward nose-down moment, thereby achieving pitch moment balance; the lateral force balance of the entire aircraft is achieved by adjusting the fuselage roll angle, and the vertical force balance of the entire aircraft is achieved by adjusting the fuselage angle of attack.

[0008] Furthermore, in S1: When the tilt-rotor aircraft encounters right-side wind, the first lift increment generated by the right V-tail is perpendicular to the incoming flow direction and to the right and below. , specifically:

[0009] Where, is the air density, is the relative incoming flow velocity of the right V-tail, is the reference area of the right V-tail, The right V-tail rudder deflection angle continues to deflect upward to the first angle The increase in lift coefficient at .

[0010] Furthermore, in S2: the vertical force for:

[0011] The lifting moment relative to the center of gravity of the whole aircraft for:

[0012] Where, is the dihedral angle of the V-tail. It is the distance from the aerodynamic center of the V-tail to the center of gravity of the entire aircraft.

[0013] Furthermore, in S2: by changing the longitudinal cyclic pitch of the rotor and the rotor tilt angle, the fuselage generates a downward nose-down moment, thereby achieving pitch moment balance, specifically: By changing the longitudinal cyclic pitch of the rotor , rotor tilt angle , causing the rotor to generate a downward force , which in turn causes the fuselage to generate a downward nose moment , as follows:

[0014]

[0015] Where, is the vertical control derivative of the rotor longitudinal cyclic pitch change, is the rotor thrust, is the number of rotors, is the control derivative of the rotor longitudinal cyclic pitch change in the pitch direction, It is the distance from the point where the rotor thrust acts to the center of gravity of the entire aircraft.

[0016] Furthermore, in S2: vertical force balance of the entire aircraft is achieved by adjusting the fuselage angle of attack, specifically: By adjusting the fuselage angle of attack The lift generated and pitching moment , as follows

[0017]

[0018] Where, is the relative velocity of the fuselage, is the reference area of the entire machine, is the lift coefficient increment after the fuselage angle of attack changes, is the derivative of the fuselage pitching moment with respect to the fuselage angle of attack.

[0019] Furthermore, the method further comprises: The right V-tail rudder deflection angle is deflected upward to the first angle Finally, the rotor longitudinal cyclic pitch variation with fuselage force and moment balance is achieved , rotor tilt angle and fuselage angle of attack The change in satisfies the following equilibrium equation:

[0020] .

[0021] Further, Setting the rotor longitudinal cyclic pitch and rotor tilt angle The manipulation coupling relationship is as follows

[0022] Where, is the control coefficient of the rotor longitudinal cyclic pitch change to the rotor tilt angle; The longitudinal cyclic pitch of the rotor and rotor tilt angle The control coupling relationship is brought into the balance equation to solve the rotor longitudinal cyclic pitch variation. , rotor tilt angle and fuselage angle of attack .

[0023] Furthermore, the control coefficient of the rotor longitudinal cyclic pitch change on the rotor tilt angle is The value rule of: When the longitudinal cyclic pitch control effect of the rotor is greater than the control effect of the rotor tilt angle, the control coefficient Take the larger value. When the control effect of the rotor tilt angle is greater than the control effect of the rotor longitudinal cyclic pitch change, the control coefficient Take the smaller value; Control coefficient The value is obtained according to the following formula: ; in, is the control derivative of the rotor tilt angle in the pitch direction.

[0024] The V-tail rudder heading control expansion control method of the present invention can expand the heading control capability of the aircraft, enhance the sideslip capability of the aircraft during flight, improve the ability to resist crosswinds in the air, and expand the sideslip envelope of the aircraft. Especially when the flight speed is low and the efficiency of the tail rudder is low, when encountering crosswinds and stable sideslipping is required, the tail rudder deflection angle requirement is large, and the deflection angle of the tail rudder on one side can easily reach the limit. The method of the present invention can effectively improve the sideslip capability of an aircraft with a V-tail aerodynamic layout. BRIEF DESCRIPTION OF THE DRAWINGS

[0025] Figure 1 A schematic diagram of the initial deflection angle of a V-tail rudder provided in an embodiment of the present invention;

[0026] Figure 2 A schematic diagram of the change in deflection angle of the V-tail rudder surface during sideslip provided by an embodiment of the present invention;

[0027] Figure 3 A schematic diagram of the change in the rudder surface deflection angle of the V-tail left rudder surface during sideslip provided by an embodiment of the present invention when continuing heading control after the deflection angle reaches a limit. DETAILED DESCRIPTION

[0028] The technical solution of the present invention is described in detail below with reference to the accompanying drawings.

[0029] The principle of the technical solution of the present invention is explained in detail below. Since the drag is always perpendicular to the incoming flow and backward, the following analysis mainly focuses on the change in the magnitude and direction of the lift. All forces and moments on the tail should take the change in drag into account. Take the right side sliding of a tilt-rotor aircraft as an example for analysis. In stable level flight without sideslip, the deflection angles of the tail rudders on both sides are the same. For example, the left and right rudder angles are positive. 、 for

[0030]

[0031] At this time, the left tail wing generates a lift perpendicular to the incoming flow direction and to the upper right, and the right tail wing generates a lift perpendicular to the incoming flow direction and to the upper left. The lateral lift components of the left and right tail wings cancel each other out, and the longitudinal lift components are superimposed on each other to generate a vertical upward resultant force.

[0032] On this basis, if the wind is blowing from the right, the fuselage and tail stabilizer will generate a right yaw moment for the statically stable aircraft. The left pedal is used to achieve heading balance, and the rudder angles on both sides are deflected to the left at the same time. , the left rudder deflects downward, and the right rudder deflects upward. According to convention, the downward deflection of the rudder is positive, and the upward deflection is negative, that is, the left rudder deflection angle for

[0033]

[0034] Right rudder angle for

[0035]

[0036] At this time, the left tail generates a lift increment perpendicular to the incoming flow direction and to the upper right, and the right tail generates a lift increment perpendicular to the incoming flow direction and to the lower right. The lateral components of the lift increments of the left and right tail wings are superimposed on each other, generating a lateral force to the right and a yaw moment to the left relative to the center of gravity, in order to balance the left force and right yaw moment generated by the right side sliding of the fuselage and tail stabilizer; the longitudinal components of the increments of the left and right tail wing forces cancel each other out, and the vertical resultant force remains unchanged, keeping the vertical force and pitching moment balanced.

[0037] As the right slip angle increases, the left rudder deflection angle reaches the upper limit first, and at this time the right rudder still has room for manipulation. As the right slip angle continues to increase, the right yaw moment of the fuselage and tail stabilizer increases, and the right rudder deflection angle needs to be manipulated to continue to deflect to the left. , the left rudder angle remains unchanged at the upper limit, that is, the right rudder angle for

[0038]

[0039] At this time, the right tail generates a lift increment perpendicular to the incoming flow direction and to the lower right, as shown in the following formula:

[0040]

[0041] Where, is the air density, is the relative flow velocity of the right tail wing, is the reference area of the right tail wing, Deflection of the right tail rudder The lift coefficient increment. The lateral component of this lift increment generates a rightward force and a left yawing moment to balance the leftward force and right yawing moment generated by the right side slip of the fuselage and tail stabilizer. The longitudinal component generates a vertical downward force and a nose-up moment relative to the center of gravity, as shown in the following formula:

[0042]

[0043]

[0044] Where, is the tail anhedral angle, It is the distance from the aerodynamic center of the tail to the center of gravity of the entire aircraft.

[0045] By changing the longitudinal cyclic pitch of the rotor , rotor tilt angle , so that the rotor generates a downward force and a nose-down moment, as shown below

[0046]

[0047]

[0048] Where, is the vertical control derivative of the rotor longitudinal cyclic pitch change, is the rotor thrust, is the number of rotors, is the control derivative of the rotor longitudinal cyclic pitch change in the pitch direction, It is the distance from the point where the rotor thrust acts to the center of gravity of the entire aircraft.

[0049] By fuselage angle of attack Adjust the lift and pitch moment of the fuselage as follows

[0050]

[0051]

[0052] Where, is the relative velocity of the fuselage, is the reference area of the entire machine, is the lift coefficient increment after the fuselage angle of attack changes, is the derivative of the fuselage pitching moment with respect to the fuselage angle of attack.

[0053] Setting the rotor longitudinal cyclic pitch and rotor tilt angle The manipulation coupling relationship is as follows

[0054]

[0055] Where, It is the control coefficient of the rotor longitudinal cyclic pitch change to the rotor tilt angle. This control coefficient can be adjusted according to the control effect of the rotor longitudinal cyclic pitch change and the rotor tilt angle on the pitch moment. The control effect of the rotor longitudinal cyclic pitch change is greater than the control effect of the rotor tilt angle. Taking the larger value, the control effect of the rotor tilt angle is greater than the control effect of the rotor longitudinal cyclic pitch change, and the control coefficient Take the smaller value;

[0056] Control coefficient The value is obtained according to the following formula: ;

[0057] in, is the control derivative of the rotor tilt angle in the pitch direction.

[0058] By balancing the total force and moment of the aircraft, the right rudder deflection angle can be calculated as follows: Finally, the changes in the longitudinal cyclic pitch of the rotor, the rotor tilt angle and the fuselage angle of attack are achieved.

[0059]

[0060]

[0061] The rudder angle change of the V-tail rudder heading control expansion control method is as follows Figures 1 to 3 As shown. Taking the tilt-rotor aircraft flying on the right side as an example, when the aircraft is in stable level flight without sideslip, the deflection angles of the tail rudder surfaces on both sides are the same, both ,like Figure 1 As shown; on this basis, slide to the right and push the left pedal. For a statically stable aircraft, the rudder angles on both sides will deflect to the left at the same time. , the left rudder deflects downward, and the right rudder deflects upward. According to convention, the downward deflection of the rudder is positive, and the upward deflection is negative, that is, the deflection angle of the left rudder is , the right rudder angle is ,like Figure 2 As shown; as the right side sliding angle increases, the initial deflection angle For example, if the left rudder angle reaches the upper limit first, the right rudder still has room for manipulation, and the right slip angle continues to increase. The right rudder angle is manipulated to continue to deflect to the left. , the left tail rudder deflection angle remains at the upper limit position unchanged to balance the heading, such as Figure 3 As shown; at this time, the right tail will also generate a nose-up moment, through the longitudinal periodic pitch change of the rotor and tilting to a certain angle , generating a downward nose-down moment to balance the pitching moment.

[0062] In the technical solution of the present invention, the tilt-rotor aircraft with a V-shaped tail aerodynamic layout performs heading control to ensure stable sideslip flight when encountering crosswind or active sideslip. After the deflection angle of the tail rudder surface on one side reaches the limit, the heading control can continue to change only the deflection angle of the tail rudder surface on the other side to achieve a larger angle of stable sideslip; the pitch moment balancing is achieved by changing the longitudinal cyclic pitch of the rotor and the rotor tilt angle.

Claims

1. A V-tail rudder heading control method, characterized in that: When the tilt-rotor aircraft is in stable level flight, the V-shaped tail fins on both sides are simultaneously deflected downward by the same angle, and the method includes: S1, when a tiltrotor aircraft encounters a right-hand wind, the fuselage generates a right yawing moment. To regain balance, the left V-tail rudder is manipulated to deflect downward, and the right V-tail rudder upward. The left V-tail rudder deflection angle first reaches the upper limit of the rudder deflection angle range. The right V-tail rudder deflection angle is then manipulated upward by a first angle. The right V-tail then generates a first lift increment perpendicular to the incoming flow direction and to the lower right. This first lift increment is decomposed into a horizontal rightward lateral force and a left yawing moment, as well as a vertical downward vertical force and a pitch-up moment relative to the center of gravity of the entire aircraft. The left yawing moment is used to offset the right yawing moment of the fuselage generated by the right-hand wind. S2, to balance the horizontal rightward lateral force, vertical downward vertical force, and pitch moment relative to the aircraft's center of gravity generated by the right V-tail, by changing the longitudinal cyclic rotor pitch and rotor tilt angle to generate a downward nose-down moment on the fuselage, thereby achieving pitch moment balance; lateral force balance is achieved by adjusting the fuselage roll angle, and vertical force balance is achieved by adjusting the fuselage angle of attack; S3: When the tilt-rotor aircraft encounters a left-hand wind, a left yawing moment is generated on the fuselage. To regain balance, the left V-tail rudder is manipulated to deflect upward and the right V-tail rudder downward. The right V-tail rudder deflection angle first reaches the upper limit of the rudder deflection angle range. The left V-tail rudder deflection angle is then manipulated upward by a second angle. The left V-tail then generates a second lift increment perpendicular to the incoming flow and directed downward to the left. This second lift increment is decomposed into a horizontal leftward lateral force and a right yawing moment, as well as a vertical downward vertical force and a pitch-up moment relative to the center of gravity of the aircraft. The right yawing moment is used to offset the left yawing moment of the fuselage generated by the left-hand wind. S4, in order to balance the horizontal left lateral force, vertical downward vertical force and pitch-up moment relative to the center of gravity of the entire aircraft generated by the left V-shaped tail, the longitudinal cyclic pitch and rotor tilt angle are changed to make the fuselage generate a downward nose-down moment, thereby achieving pitch moment balance; the lateral force balance of the entire aircraft is achieved by adjusting the fuselage roll angle, and the vertical force balance of the entire aircraft is achieved by adjusting the fuselage angle of attack.

2. A V-tail rudder control method according to claim 1, characterized in that: In S1: When the tilt-rotor aircraft encounters right-side wind, the first lift increment generated by the right V-tail is perpendicular to the incoming flow direction and to the lower right. , specifically: Where, is the air density, is the relative incoming flow velocity of the right V-tail, is the reference area of the right V-tail, The right V-tail rudder deflection angle continues to deflect upward to the first angle The increase in lift coefficient at .

3. A V-tail rudder heading control method according to claim 2, characterized in that: In S2: the vertical force pointing downward for: The lifting moment relative to the center of gravity of the whole aircraft for: Where, is the dihedral angle of the V-tail. It is the distance from the aerodynamic center of the V-tail to the center of gravity of the entire aircraft.

4. A V-tail rudder heading control method according to claim 3, characterized in that: In S2, by changing the longitudinal cyclic pitch and the rotor tilt angle, the fuselage generates a downward nose-down moment, thereby achieving pitch moment balance. Specifically: By changing the longitudinal cyclic pitch of the rotor , rotor tilt angle , causing the rotor to generate a downward force , which in turn causes the fuselage to generate a downward nose moment , as follows: Where, is the vertical control derivative of the rotor longitudinal cyclic pitch change, is the rotor thrust, is the number of rotors, is the control derivative of the rotor longitudinal cyclic pitch change in the pitch direction, It is the distance from the point where the rotor thrust acts to the center of gravity of the entire aircraft.

5. A V-tail rudder control method according to claim 4, characterized in that: In S2, vertical force balance of the entire aircraft is achieved by adjusting the fuselage angle of attack, specifically: By adjusting the fuselage angle of attack The lift generated and pitching moment , as follows Where, is the relative velocity of the fuselage, is the reference area of the entire machine, is the lift coefficient increment after the fuselage angle of attack changes, is the derivative of the fuselage pitching moment with respect to the fuselage angle of attack.

6. A V-tail rudder control method according to claim 5, characterized in that: The method further comprises: The right V-tail rudder deflection angle is deflected upward to the first angle Finally, the rotor longitudinal cyclic pitch variation with fuselage force and moment balance is achieved , rotor tilt angle and fuselage angle of attack The change in satisfies the following equilibrium equation: 。 7. A V-tail rudder control method according to claim 6, characterized in that: Setting the rotor longitudinal cyclic pitch and rotor tilt angle The manipulation coupling relationship is as follows Where, is the control coefficient of the rotor longitudinal cyclic pitch change to the rotor tilt angle; The longitudinal cyclic pitch of the rotor and rotor tilt angle The control coupling relationship is brought into the balance equation to solve the rotor longitudinal cyclic pitch variation. , rotor tilt angle and fuselage angle of attack .

8. A V-tail rudder heading control method according to claim 7, characterized in that: Control coefficient of rotor longitudinal cyclic pitch change on rotor tilt angle The value rule of: When the longitudinal cyclic pitch control effect of the rotor is greater than the control effect of the rotor tilt angle, the control coefficient Take the larger value. When the control effect of the rotor tilt angle is greater than the control effect of the rotor longitudinal cyclic pitch change, the control coefficient Take the smaller value; Control coefficient The value is obtained according to the following formula: ; in, is the control derivative of the rotor tilt angle in the pitch direction.

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