A method for expanding control of a v-tail rudder surface heading control

Abstract

The present application belongs to the technical field of flight dynamics, and particularly relates to a V-tail rudder surface heading control expansion method, which comprises the following steps: when the left rudder surface deflection angle reaches the upper limit, the right rudder surface deflection angle continues to deflect to the left by a certain angle, the right tail wing generates an upward right lift increment perpendicular to the flow direction; the right tail wing simultaneously generates a vertical downward force and a lifting moment relative to the whole aircraft gravity center; a downward lowering moment is generated through longitudinal periodic pitch of the rotor and rotor tilt angle, so that the pitching moment is balanced; and the body angle of attack is adjusted to balance the vertical force of the whole aircraft.

Description

TECHNICAL FIELD

[0001] The present application belongs to the technical field of flight mechanics, and particularly relates to a V-tail rudder surface heading control expansion method. BACKGROUND

[0002] Compared with a conventional tail, a V-tail is increasingly applied to the design of a tilt-rotor aircraft due to its smaller wet area and lighter structure weight. The V-tail is composed of two wings arranged in a V shape at the tail of the fuselage, and each wing is composed of a fixed stabilizer and a hinged rudder. The V-tail functions to control the longitudinal and heading directions of the aircraft. When the longitudinal direction is controlled, the two rudders are deflected in the same direction, and when the heading direction is controlled, the two rudders are deflected in opposite directions. The final deflection angle of the rudder is the result of the combination of the longitudinal and heading controls.

[0003] When the tilt-rotor aircraft is in a steady and straight flight without sideslip, the deflection angles of the two tail rudders are consistent. At this time, the lateral force components generated by the dihedral angles of the two tails cancel each other out, and only the longitudinal resultant force is generated. If a lateral gust is encountered at this time, heading control is required. Due to the opposite deflection of the two rudders, the deflection angle of one side of the rudder increases, and the deflection angle of the other side of the rudder decreases. The side with the increased deflection angle will reach the upper limit of the deflection angle range first. At this time, the heading control capability reaches the limit. Although the deflection angle of the rudder of the other tail does not reach the limit, it cannot continue to be controlled to generate a lateral force and a yawing moment because the change in the deflection angle of the single tail rudder will simultaneously generate a lateral force and a longitudinal force. The longitudinal force generates a longitudinal moment, which makes the aircraft unable to complete longitudinal trim, thereby losing balance.

[0004] Currently, the heading control limit is generally determined by the deflection angle of one side of the tail rudder reaching the upper limit. In fact, the other side of the tail rudder still has control space. How to balance the longitudinal moment caused by the deflection of the single rudder becomes the key to expanding the heading control capability of the V-tail rudder. SUMMARY

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

[0006] To achieve the above-mentioned purpose, the present application adopts the following technical solutions.

[0007] A V-tail rudder heading control expansion method, when a tilt-rotor aircraft is in a steady and straight flight, the two sides of the V-tail are simultaneously deflected downward by the same angle. The method comprises the following steps:

[0008] S1, when the tilt-rotor aircraft encounters right side wind, the fuselage generates right yawing moment, in order to re-achieve balance, the left V-tail control surface is deflected downward, the right V-tail control surface is deflected upward, and the deflection angle of the left V-tail control surface reaches the upper limit of the deflection angle range first, and the deflection angle of the right V-tail control surface is deflected upward by a first angle, then the right V-tail generates a first lift increment which is perpendicular to the flow direction and downward to the right, and the first lift increment is decomposed into horizontal right lateral force and left yawing moment, and vertical downward vertical force and lifting moment relative to the center of gravity of the whole aircraft; the left yawing moment is used to offset the right yawing moment generated by the fuselage due to the right side wind;

[0009] S2, in order to balance the horizontal right lateral force, the vertical downward vertical force and the lifting moment relative to the center of gravity of the whole aircraft generated by the right V-tail, the longitudinal periodic pitch of the rotor and the tilt angle of the rotor are changed to make the fuselage generate downward pitching moment, so as to achieve the balance of pitching moment; the roll angle of the fuselage is adjusted to achieve the balance of the lateral force of the whole aircraft, and the angle of attack of the fuselage is adjusted to achieve the balance of the vertical force of the whole aircraft;

[0010] S3, when the tilt-rotor aircraft encounters left side wind, the fuselage generates left yawing moment, in order to re-achieve balance, the left V-tail control surface is deflected upward, the right V-tail control surface is deflected downward, and the deflection angle of the right V-tail control surface reaches the upper limit of the deflection angle range first, and the deflection angle of the left V-tail control surface is deflected upward by a second angle, then the left V-tail generates a second lift increment which is perpendicular to the flow direction and downward to the left, and the second lift increment is decomposed into horizontal left lateral force and right yawing moment, and vertical downward vertical force and lifting moment relative to the center of gravity of the whole aircraft, the right yawing moment is used to offset the left yawing moment generated by the fuselage due to the left side wind;

[0011] S4, in order to balance the horizontal left lateral force, the vertical downward vertical force and the lifting moment relative to the center of gravity of the whole aircraft generated by the left V-tail, the longitudinal periodic pitch of the rotor and the tilt angle of the rotor are changed to make the fuselage generate downward pitching moment, so as to achieve the balance of pitching moment; the roll angle of the fuselage is adjusted to achieve the balance of the lateral force of the whole aircraft, and the angle of attack of the fuselage is adjusted to achieve the balance of the vertical force of the whole aircraft.

[0012] Further, in S1: when the tilt-rotor aircraft encounters right side wind, the right V-tail generates a first lift increment which is perpendicular to the flow direction and downward to the right , specifically:

[0013]

[0014] In the formula, is the air density, is the relative velocity of the right V-tail to the flow, is the reference area of the right V-tail, The lift coefficient increment when the right V-tail rudder surface deflection angle is continuously deflected upward by a first angle The lift coefficient increment when the right V-tail rudder surface deflection angle is continuously deflected upward by a first angle

[0015] Further, in S2: the vertical force downward Is:

[0016]

[0017] The lifting moment of the relative center of gravity of the whole machine Is:

[0018]

[0019] In the formula, The upper angle of the V-tail, The distance from the aerodynamic center of the V-tail to the center of gravity of the whole machine.

[0020] Further, in S2: by changing the rotor longitudinal cyclic pitch, rotor tilt angle, the fuselage generates a downward pitching moment, so as to achieve the balance of the pitching moment, specifically:

[0021] By changing the rotor longitudinal cyclic pitch , the rotor tilt angle , the rotor generates a downward force , and further makes the fuselage generate a downward pitching moment , as follows:

[0022]

[0023]

[0024] In the formula, The vertical control derivative of the rotor longitudinal cyclic pitch, The rotor tension, The number of rotors, The pitch control derivative of the rotor longitudinal cyclic pitch in the pitch direction, The distance from the rotor tension point to the center of gravity of the whole machine.

[0025] Further, in S2: by adjusting the angle of attack of the fuselage to achieve the vertical force balance of the whole machine, specifically:

[0026] By adjusting the angle of attack of the fuselage The lift And the pitching moment , as follows

[0027]

[0028]

[0029] wherein, is the airspeed of the fuselage, is the reference area of the fuselage, is the increment of the lift coefficient of the fuselage after the change of the angle of attack of the fuselage, is the derivative of the pitching moment of the fuselage with respect to the angle of attack of the fuselage.

[0030] Further, the method further comprises:

[0031] the deflection angle of the right V-tail surface is deflected upward by a first angle then, the rotor longitudinal cyclic pitch is changed to balance the force and moment of the fuselage , the rotor tilt angle and the change of the angle of attack of the fuselage satisfy the following balance equation:

[0032]

[0033] .

[0034] Further,

[0035] the control coupling relationship between the rotor longitudinal cyclic pitch and the rotor tilt angle is set as follows:

[0036]

[0037] wherein, is the control coefficient of the rotor longitudinal cyclic pitch with respect to the rotor tilt angle;

[0038] the control coupling relationship between the rotor longitudinal cyclic pitch and the rotor tilt angle is brought into the balance equation, and the rotor longitudinal cyclic pitch , the rotor tilt angle and the angle of attack of the fuselage are solved.

[0039] Further, the control coefficient of the rotor longitudinal cyclic pitch with respect to the rotor tilt angle is valued as follows: when the control effect of the rotor longitudinal cyclic pitch is greater than the control effect of the rotor tilt angle, the control coefficient takes a large value, and when the control effect of the rotor tilt angle is greater than the control effect of the rotor longitudinal cyclic pitch, the control coefficient takes a small value;

[0040] the control coefficient is valued according to the following formula: ;

[0041] wherein, is the control derivative of the rotor tilt angle in the pitch direction.

[0042] The V-tail rudder surface heading control expansion method of the application can expand the heading control capability of the aircraft, enhance the sideslip capability of the aircraft in flight, improve the anti-crosswind capability in the air, expand the sideslip envelope of the aircraft, and especially in the case of small flight speed and low tail rudder surface efficiency, when the aircraft needs to be stably sideslipped in the presence of crosswind, the tail rudder surface deflection angle requirement is large, and the deflection angle of one side of the tail rudder surface is easy to reach the limit. The method of the application can effectively improve the sideslip capability of the V-tail aerodynamic layout aircraft. BRIEF DESCRIPTION OF DRAWINGS

[0043] Fig. 1 The V-tail rudder surface initial deflection angle schematic diagram provided by the embodiment of the application;

[0044] Fig. 2 The V-tail rudder surface deflection angle change schematic diagram when sideslipping provided by the embodiment of the application;

[0045] Fig. 3 The rudder surface deflection angle change schematic diagram when the V-tail left rudder surface deflection angle reaches the limit and continues to be heading controlled provided by the embodiment of the application. DETAILED DESCRIPTION

[0046] The technical solutions of the application will be described in detail below with reference to the accompanying drawings.

[0047] The principle of the technical solutions of the application will be described in detail below. Since the resistance is always perpendicular to the incoming flow and backward, the change of the lift size and direction is mainly analyzed below, and the forces and moments on all the tail surfaces should consider the change of the resistance. Taking the right sideslip of the tilt rotor aircraft as an example for analysis, when there is no sideslip stable level flight, the deflection angles of the two sides of the tail surface are the same, and the initial deflection angle of the rudder surface is for example, the deflection angles of the left and right rudder surfaces are , for

[0048]

[0049] At this time, the left tail produces a lift perpendicular to the incoming flow direction and upward, and the right tail produces a lift perpendicular to the incoming flow direction and leftward. The lift components in the transverse direction of the left and right tail surfaces cancel each other out, and the lift components in the longitudinal direction add up to produce a vertical upward force.

[0050] On this basis, when encountering a right crosswind, for the heading static stable aircraft, the fuselage and tail surface produce a right yawing moment, and the left foot is stepped to realize the heading balance. The deflection angles of the two sides of the rudder surface are simultaneously deflected to the left by an angle of , the left side rudder is deflected downward, the right side rudder is deflected upward, according to the convention, the general rudder is deflected downward as positive, deflected upward as negative, that is, the left side rudder deflection angle is

[0051]

[0052] the right side rudder deflection angle is

[0053]

[0054] At this time, the left side tail wing generates an upward lift increment perpendicular to the flow direction, the right side tail wing generates a downward lift increment perpendicular to the flow direction, the lift increments of the left and right side tail wings in the lateral direction are superimposed on each other, generating a rightward lateral force and a leftward yawing moment relative to the center of gravity, to balance the leftward force and the rightward yawing moment of the fuselage and tail wing stabilizer due to the right side slip; the increments of the left and right side tail wing forces in the longitudinal direction are counteracted, the vertical resultant force remains unchanged, maintaining the vertical force and the pitch moment balance.

[0055] With the increase of the right side slip angle, the left side rudder deflection angle first reaches the upper limit, at this time the right side rudder still has steering space. With the continuous increase of the right side slip angle, the rightward yawing moment of the fuselage and tail wing stabilizer increases, and the right side rudder deflection angle needs to be further deflected to the left , the left side rudder deflection angle remains unchanged at the upper limit, that is, the right side rudder deflection angle is

[0056]

[0057] At this time, the right side tail wing generates a downward lift increment perpendicular to the flow direction, as follows:

[0058]

[0059] In the formula, is the air density, is the relative flow velocity of the right side tail wing, is the reference area of the right side tail wing, is the lift coefficient increment of the right side tail wing rudder deflection . The lateral component of this lift increment generates a rightward force and a leftward yawing moment, to balance the leftward force and the rightward yawing moment of the fuselage and tail wing stabilizer due to the right side slip, and the longitudinal component generates a vertical downward force and a lifting moment relative to the center of gravity, as follows

[0060]

[0061]

[0062] In the formula, is the tail up angle, is the distance from the aerodynamic center of the tail to the center of gravity of the whole aircraft.

[0063] The rotor longitudinal cyclic pitch is changed The rotor tilt angle is changed The rotor generates a downward force and a nose down moment, as follows

[0064]

[0065]

[0066] where, is the vertical control derivative of the rotor longitudinal cyclic pitch, is the rotor tension, is the number of rotors, is the vertical control derivative of the rotor longitudinal cyclic pitch, is the distance from the rotor tension point to the center of gravity of the whole aircraft.

[0067] The fuselage angle of attack is changed The fuselage lift and the pitch moment are adjusted, as follows

[0068]

[0069]

[0070] where, is the relative speed of the fuselage to the flow, is the reference area of the whole aircraft, is the increment of the lift coefficient of the fuselage after the angle of attack is changed, is the derivative of the pitch moment of the fuselage to the angle of attack.

[0071] The control coupling relationship between the rotor longitudinal cyclic pitch and the rotor tilt angle is set, as follows

[0072]

[0073] where, is the control coefficient of the rotor longitudinal cyclic pitch to the rotor tilt angle. This control coefficient can be adjusted according to the control efficiency of the rotor longitudinal cyclic pitch and the rotor tilt angle to the pitch moment. If the control efficiency of the rotor longitudinal cyclic pitch is greater than that of the rotor tilt angle, the control coefficient takes a large value. If the control efficiency of the rotor tilt angle is greater than that of the rotor longitudinal cyclic pitch, the control coefficient takes a small value;

[0074] The control coefficient is taken as follows: ;

[0075] wherein, is the control derivative of the rotor tilt angle in the pitch direction.

[0076] By balancing the total force and moment, the right side rudder deflection angle continues to deflect leftward by the following formula After that, the change of the rotor longitudinal cyclic pitch, the rotor tilt angle and the change of the body angle of attack are realized.

[0077]

[0078]

[0079] The rudder deflection angle change of the V-tail rudder heading control expansion method is shown in Figs. 1 to 3 . Taking the right side flight of the tilt rotor aircraft as an example, when the steady flight without sideslip, the deflection angles of the rudders on both sides are the same, and are , as shown in Fig. 1 ; on this basis, the right sideslip is performed, and the left foot is stepped on. For the aircraft with static stability in the heading direction, the deflection angles of the rudders on both sides deflect leftward by an angle at the same time, the left rudder deflects downward, and the right rudder deflects upward. According to the convention, the deflection of the rudder downward is positive, and the deflection upward is negative, that is, the deflection angle of the left rudder is , and the deflection angle of the right rudder is , as shown in Fig. 2 ; with the increase of the right sideslip angle, taking the initial deflection angle as positive as an example, the deflection angle of the left rudder reaches the upper limit first, at this time, the deflection angle of the right rudder still has a control space, the right sideslip angle continues to increase, the deflection angle of the control right rudder continues to deflect leftward by an angle , and the deflection angle of the left rudder remains unchanged, so that the heading balance is realized, as shown in Fig. 3 ; at this time, the right tail will generate a head-lifting moment at the same time, by changing the rotor longitudinal cyclic pitch and tilting by an angle , a downward head-lowering moment is generated, so that the pitch moment is balanced.

[0080] In the technical scheme of the present application, when the tilt rotor aircraft with the V-tail aerodynamic layout encounters crosswind or active sideslip, the heading control is performed to make the aircraft fly stably in the sideslip, after the deflection angle of one side rudder reaches the limit, the heading control can be continued to change the deflection angle of the rudder on the other side, to realize the stable sideslip by a larger angle; the pitch moment balance is realized by changing the rotor longitudinal cyclic pitch and the rotor tilt angle.

Claims

1. A method of V-tail fin surface heading control augmentation, comprising: The method comprises the following steps of: S1, when the tilt-rotor aircraft encounters right wind, the fuselage generates right yawing moment, in order to re-achieve balance, the left V-tail surface is downward deflected, the right V-tail surface is upward deflected, and the deflection angle of the left V-tail surface reaches the upper limit of the deflection angle range first, and the deflection angle of the right V-tail surface is upward deflected by a first angle, then the right V-tail surface generates a first lift increment which is perpendicular to the flow direction and downward to the right, and the first lift increment is decomposed into horizontal lateral force to the right, left yawing moment, vertical downward vertical force and lifting moment relative to the center of gravity of the whole aircraft, and the left yawing moment is used to offset the right yawing moment generated by the fuselage due to the right wind; S2, in order to balance the horizontal lateral force to the right, the vertical downward vertical force and the lifting moment relative to the center of gravity of the whole aircraft generated by the left V-tail surface, the longitudinal periodic pitch of the rotor and the tilt angle of the rotor are changed to make the fuselage generate downward pitching moment, so as to achieve balance of the pitching moment, the roll angle of the fuselage is adjusted to achieve balance of the lateral force of the whole aircraft, and the attack angle of the fuselage is adjusted to achieve balance of the vertical force of the whole aircraft; S3, when the tilt-rotor aircraft encounters left wind, the fuselage generates left yawing moment, in order to re-achieve balance, the left V-tail surface is upward deflected, the right V-tail surface is downward deflected, and the deflection angle of the right V-tail surface reaches the upper limit of the deflection angle range first, and the deflection angle of the left V-tail surface is upward deflected by a second angle, then the left V-tail surface generates a second lift increment which is perpendicular to the flow direction and downward to the left, and the second lift increment is decomposed into horizontal lateral force to the left, right yawing moment, vertical downward vertical force and lifting moment relative to the center of gravity of the whole aircraft, and the right yawing moment is used to offset the left yawing moment generated by the fuselage due to the left wind; S4, in order to balance the horizontal lateral force to the left, the vertical downward vertical force and the lifting moment relative to the center of gravity of the whole aircraft generated by the left V-tail surface, the longitudinal periodic pitch of the rotor and the tilt angle of the rotor are changed to make the fuselage generate downward pitching moment, so as to achieve balance of the pitching moment, the roll angle of the fuselage is adjusted to achieve balance of the lateral force of the whole aircraft, and the attack angle of the fuselage is adjusted to achieve balance of the vertical force of the whole aircraft.

2. The method of claim 1, wherein, In S1: When the tiltrotor aircraft encounters a right side wind, the right V-tail generates a first lift increment perpendicular to the incoming flow direction, right down Specifically: wherein is the air density, is the right V-tail relative velocity, is the right V-tail reference area, is the right V-tail elevator deflection angle, and is the lift coefficient increment when the right V-tail elevator deflection angle continues to deflect by a first angle.

3. The method of claim 2, wherein, In S2: the vertically downward vertical force is: The relative full machine center of gravity lifting moment Is: wherein is the dihedral angle of the V-tail, is the distance from the aerodynamic center of the V-tail to the center of gravity of the aircraft.

4. The method of claim 3, wherein, In S2, the longitudinal periodic pitch of the rotor and the tilt angle of the rotor are changed to make the fuselage generate downward pitching moment, so as to achieve balance of the pitching moment, and the specific process is as follows: By changing the longitudinal cyclic pitch of the rotor the angle of tilt of the rotor to cause the rotor to produce a downward force and in turn a downward pitching moment on the airframe as follows: wherein is the rotor longitudinal cyclic pitch in the vertical direction, is the rotor tension, is the number of rotors, is the rotor longitudinal cyclic pitch in the vertical direction, is the distance from the rotor tension point to the center of gravity of the helicopter.

5. The method of claim 4, wherein, In S2, the attack angle of the fuselage is adjusted to achieve balance of the vertical force of the whole aircraft, and the specific process is as follows: By adjusting the angle of attack of the fuselage The lift generated And the pitch moment As follows wherein is the relative velocity of the fuselage to the flow, is the reference area of the whole aircraft, is the increment of the lift coefficient of the fuselage after the change of the angle of attack of the fuselage, is the derivative of the pitching moment of the fuselage to the angle of attack of the fuselage.

6. The method of claim 5, wherein, The method further comprises: The right V-tail rudder surface deflection angle is upwardly deflected by a first angle After, the rotor longitudinal cyclic pitch is changed to achieve the fuselage force and moment trim , the rotor tilt angle , and the change amount of the fuselage angle of attack satisfy the following balance equation: 。 7. The V-tail surface heading control expansion control method according to claim 6, characterized in that, Setting a rotor longitudinal cyclic pitch and a rotor tilt angle in a coupled relationship as follows: wherein is the rotor longitudinal cyclic pitch-to-tilt control effectiveness Coupling the rotor longitudinal cyclic pitch and rotor tilt angle manipulation into the balance equation yields the rotor longitudinal cyclic pitch , rotor tilt angle , and fuselage angle of attack .

8. The method of claim 7, wherein, 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; manipulation coefficient is valued according to the following formula: ; wherein is the control derivative of the rotor tilt angle in the pitch direction.

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