JPH0539092A - Transient flight posture control method for vertical posture landing aircraft - Google Patents

Abstract

PURPOSE:To control a posture in the optimum conditions by a simple method wherein an aircraft having aerodynamic control faces and a thrust deflecting device to control the posture of an airframe as for three axes by obtaining the angular deviations of posture as for the three axes, and obtaining angular acceleration command values on the basis of the deviations. CONSTITUTION:A vertical posture landing aircraft is provided with right and left elevons 43b, 43a to control a pitch axis and a roll axis and a rudder 42 to control a yaw axis as aerodynamic control faces, and thrust deflecting vanes 44a-44b cross-likely are arranged in rear of the exhaust port of an engine as a thrust deflecting device. In the case of controlling these respective control elements 42-44d with a control device, deviations between respective posture angle commands and actual posture angles as for the three axes, consisting of the pitch axis, the roll axis, and the yaw axis, are obtained, and angular acceleration commands are obtained according to the control rule based on these deviations. Next, a control gain scheduled as a function against a pitch posture angle or a pitch posture angle command is used together with the angular acceleration command so as to obtain the control face angle command.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a method for controlling the flight attitude of a vertical attitude landing aircraft which makes a landing by transitioning from a cruise status in a horizontal attitude to a hover status in a vertical attitude.

[0002]

2. Description of the Related Art As a conventional technique for controlling the flight of a flying object, there is one disclosed in Japanese Patent Laid-Open No. 63-75500. This technique is intended to perform good control while controlling a flying object flying toward a target by adapting to flight conditions that change from moment to moment. The flight control gain calculator is mounted on the flying object, and the feedback gain is calculated by solving a state equation that describes how the velocity and flight attitude change depending on the operation amount of the aerodynamic control surface and the thrust deflector. Then, the flight speed and the attitude are fed back to the operation amounts of the aerodynamic control surface and the thrust deflector.

[0003]

However, such a conventional technique has the following problems. Flight conditions change from moment to moment, and the amount of calculation for solving the equation of state is large, leading to an increase in cost.

In the control surface control, a dynamic pressure schedule gain is generally used. This is based on the aerodynamic characteristics that have been clarified in advance, and the control gain is set as a function of the dynamic pressure that determines the magnitude of the aerodynamic force in particular with the state quantity representing the flight condition. Then, the dynamic pressure is measured during flight, and the control gain is changed based on the measured value. However, in the process of the vertical attitude landing gear transiting from the horizontal attitude to the vertical attitude, the angle of attack and the dynamic pressure change rapidly in addition to the large angle of attack. Therefore, it is extremely difficult to measure the dynamic pressure during transition flight, and it is difficult to control the flight attitude.

The present invention has been made in view of the above circumstances, and a transitional flight attitude control of a vertical attitude lander capable of optimally controlling an aerodynamic control surface and a thrust deflecting device by adapting to changes in flight conditions during transitional flight. The purpose is to provide a method.

[0006]

SUMMARY OF THE INVENTION The present invention provides attitude control during transition flight of a vertical attitude lander having an aerodynamic control surface and thrust deflectors for attitude control of the aircraft with respect to pitch, roll and yaw axes. A method of obtaining a deviation between each attitude angle command and an actual attitude angle with respect to these three axes, a step of obtaining an angular acceleration command according to a control law based on the obtained deviation, and a pitch attitude angle or pitch A step of obtaining a steering angle command using a control gain and an angular acceleration command scheduled as a function of the attitude angle command, and operating the aerodynamic control surface and the thrust deflector based on the obtained steering angle command to perform attitude control. And the step of performing.

[0007]

With respect to the pitch axis, the roll axis, and the yaw axis, the deviation between the attitude angle command and the actual attitude angle is calculated, the angular acceleration command is calculated according to the control law, and the pitch attitude angle or a function to the pitch attitude angle command is obtained. The steering angle command is obtained from the scheduled control gain and the angular acceleration command. The attitude control is performed by operating the aerodynamic control surface and the thrust deflector based on the steering angle command. in this way,
During transitional flight, a simple method is used because it uses a pitch attitude angle that is easily measured or a control gain that is scheduled as a function for a pitch attitude angle command that is obtained by easy calculation, without using dynamic pressure that is difficult to measure. Attitude control can be optimized.

[0008]

DESCRIPTION OF THE PREFERRED EMBODIMENTS An embodiment of the present invention will be described below with reference to the drawings. First, the appearance of a vertical attitude landing gear to which the present embodiment is applied is shown in FIG. The left and right elevons 43 that control the pitch axis and roll axis as aerodynamic control surfaces
A and 43b and a ladder 42 for controlling the yaw axis are provided. Further, as the thrust deflecting device, thrust deflecting vanes 44a to 44d are provided behind the exhaust port of the engine. The thrust deflection vanes 44a to 44d are four vanes arranged in a cross shape, and each vane independently moves like a moving blade. The left and right elevons 43a to 43b as aerodynamic control surfaces and the rudder 42 are used for flight in a horizontal attitude, and thrust deflection vanes 44a to 44d as a thrust deflector are used in flight in a vertical hover attitude. And during the transition flight, the left and right elevons 43a-4
3b and the rudder 42 are used in combination with the thrust deflection vanes 44a to 44d. Here, the elevons 43a and 43b are oriented in the direction of arrow A, the vanes 44a and 44b are oriented in the direction of arrow A, and the vanes 44c and 44d are oriented in the direction of arrow B. The ladder 42 takes the direction of the arrow B to be positive.

The axes to be controlled include a pitch axis, a roll axis and a yaw axis, and the attitude response characteristic when controlling one of these axes is shown in the block diagram of FIG.
The attitude command angle θ _C is input to the comparator 11, and the integrator 16
The deviation θ _C −θ from the actual posture angle θ output from is input to the calculator 12. In the calculator 12, the transfer function G _C1 of the control law is used for the calculation, and the steering angle δ _{A of the} aerodynamic control surface is calculated.
And the operating angle δ _T of the thrust deflector are output. here,
It is assumed that the steering angle command and the steering angle are equal, ignoring the dynamic characteristics of the steering surface and the thrust deflector, that is, the delay of the steering angle and the operating angle with respect to the command. The part from δ _A , δ _T to θ is a model of the response of the aircraft attitude angle to the steering angle and the thrust deflector operating angle (the broken line portion 17 in FIG. 3).

The steering angle δ _A and the operating angle δ _T output from the computing unit 12 correspond to the change amount from the balanced state. The steering angle δ _A is input to the calculator 13 and multiplied by the proportional constant K _A , and K _A · δ _A is output to the adder 15. Working angle δ _T
Is multiplied by the proportional constant K _T by the calculator 14 and output to the adder 15. Here, the steering angle δ _A and the operating angle δ _T are multiplied by the proportional constants K _A and K _T , respectively, for the following reason. It is considered that the damping moment received from the air when the attitude of the airframe changes is smaller than the control moment generated by the control surface or the thrust deflector and can be ignored. Therefore, the moment acting on the aircraft is the steering angle δ
It can be considered that it occurs in proportion to _A and the working angle δ _T. As a result, the angular acceleration of the posture angle θ is d ² θ / dt ²
Is considered to be proportional to the steering angle δ _A and the operating angle δ _T , so that the proportional constants K _A and K _T are multiplied.
The angular acceleration d ² θ / dt ² output from the adder 15 is
The integrator 16 performs integration twice and outputs the actual attitude angle θ of the machine body. This attitude angle θ is fed back to the comparator 11.

Here, the proportional constants K _A and K _T greatly change according to changes in speed, attitude and engine thrust during transition flight. Therefore, if the transfer function G _C1 is an invariant function, the response characteristic of the actual command angle θ with respect to the posture command angle θ _C also changes significantly. Therefore, variable gains F _A and F _T are newly introduced as in the control block of FIG. Comparing the control block shown in FIG. 4 with that of FIG.
The invariant transfer function G _C2 used in the computing unit 12 and the computing units 21 and 22 for performing multiplication using the variable gains F _A and F _T are respectively connected in series between the computing unit 12 and the computing units 13 and 14. The difference is that they are connected. As a result, the output d ² θC / dt ² from the arithmetic unit 12 is given to the arithmetic units 21 and 22, respectively, and the variable gains F _A and F
_The calculator 1 is multiplied by _T to obtain the steering angle δ _A and the operating angle δ _T.
It is output to 3 and 14. In the calculators 13 and 14, as in the control block of FIG. 3, the steering angle δ _A and the operating angle δ
_T is multiplied by the proportional constants K _A and K _T , respectively, and the angular acceleration d ² θ / dt ² is output.

Between the variable gains F _A and F _T and the constants of proportionality K _A and K _T , it is necessary to establish the following relationship (1).

F _A · K _A + F _T · K _T = 1 (1) When this relational expression is established, the control block shown in FIG. 4 has a relation equivalent to that of FIG. As a result, the response characteristic of the flight attitude becomes constant.

The proportional constants K _A and K _T have the following relationship with the dynamic pressure Q, the angle of attack α, and the engine thrust T.

K _A = C _A (α) · Q (2) K _T = C _T · T (3) where C _A (α) is a function of the angle of attack α, and C _T is It is a constant previously determined by the aerodynamic characteristics and the characteristics of the thrust deflector.

From the expressions (2) and (3), the values of the proportional constants K _A and K _T can be obtained by measuring the values of the attack angle α, the dynamic pressure Q and the thrust T during flight. As a result, the variable gain F
It is possible to change the values of _A and F _T moment by moment during the transition flight. In this case, the thrust T can be estimated from the engine speed, the temperature, and the like. However, the angle of attack α and the dynamic pressure Q change abruptly during transition flight, so it is difficult to measure them.

Therefore, the following handling is performed. In the transition flight, it is practically acceptable to limit the flight pattern such that the flight starts from a certain attitude and velocity during horizontal flight at a predetermined velocity and the flight follows a predetermined path. As a result, the values of α, Q, and T change in a substantially constant pattern.

Further, the pitch attitude angle or the pitch attitude angle command monotonically increases during the transition flight. Therefore, since the values of α, Q and T can be obtained as a function of the pitch attitude angle or the pitch attitude angle command, the values of the proportional constants K _A and K _T can be regarded as the function of the pitch attitude angle or the pitch attitude angle command. It becomes possible. Then, the values of the variable gains F _A and F _T are determined so as to satisfy the expression (1) and included in the control law as the schedule gain, so that the control characteristic can be kept substantially constant.

Here, the values of the variable gains F _A and F _T are
The constants K _A and K _T have the relationship of the formula (1), but there are infinite combinations. But the transition flight
It must be capable of smoothly transitioning from a horizontal attitude to a vertical hover attitude. Also,
The control surface and the thrust deflector have their respective limits. Therefore, it is necessary to determine the values of the variable gains F _A and F _T in consideration of these conditions.

FIG. 6 shows a combination example of the variable gains F _A and F _T and the proportional constants K _A and K _T. This figure is
It shows the relationship of the variable gains F _A and F _T with respect to the pitch attitude angle or the pitch attitude angle command. During level flight, the flight attitude is usually controlled only by the aerodynamic control surface, except in the special case where a thrust deflector is used. Therefore, in the region where the pitch attitude angle is small, the variable gain F _T for thrust deflection is
Is set to 0 and F _A = 1 / K _A.

On the contrary, in the vertical hover attitude, the aerodynamic control surface does not affect the flight attitude and is controlled only by the thrust deflector. Therefore, in a region where the pitch attitude angle is large, the variable gain F _A for the control surface is set to 0 and F _T = 1 / K _T.

Then, the following is a transition flight during the transition from horizontal flight to vertical hover flight. In the case of a transition flight pattern that keeps the altitude almost constant, the thrust increases as the lift decreases to support the weight of the airframe. In the vicinity of the maximum lift coefficient angle of attack, the dynamic pressure is considerably reduced, and the proportional constant K _A representing the effectiveness of the aerodynamic control surface is reduced. Although the dynamic pressure has decreased, the lift coefficient is large and the lift has not decreased much. Therefore, the thrust force is suppressed to a small value, and the proportional constant K _T, which represents the effectiveness of the thrust deflector, is also reduced. Therefore, the point where F _A = 1 / K _A and F
In the central pitch attitude angle around the point where the _T = 1 / K _T, set to both maximum value to the variable gain F _A and F _T comes. The values around the center are not determined by the above equation (1) alone. Therefore, a value is set so as not to exceed the maximum capabilities of the aerodynamic control surface and the thrust deflecting device.

The values of the variable gains F _A and F _T shown in FIG. 6 change linearly with the pitch attitude angle. If the relationship of the expression (1) is strictly maintained, it becomes a curve, but even if it is approximated by a polygonal line as shown in FIG. 6, there is no problem in practical use and the control calculation can be simplified.

The transitional flight attitude control method of this embodiment, which controls each of the three axes by the above method, will be described with reference to the control block diagram of FIG. Regarding the pitch axis, roll axis and yaw axis, based on the deviation between the attitude angle command from the guidance device and the actual attitude angle of the aircraft,
Left-wing elevon 43a, right-wing elevon 43b, rudder 4
2. Find the amount of operation for operating the vanes 44a to 44d by the actuator. This allows the attitude of the machine body to be changed according to the command.

The pitch attitude angle command θ _PC and the actual pitch attitude angle θ _P of the machine are input to the comparator 101, and the deviation θ _PC −
θ _P is input to the calculator 102. In the arithmetic unit 102,
Proportional action P using proportional constant K _PP and integral constant K _IP
And PID control consisting of an integral operation I using the Laplace operator s, a differential constant K _DP and a differential operation D using the Laplace operator s, and a pitch angular acceleration command d ² θ
_P / dt ² is output. Here, the content of the calculation performed by the arithmetic unit 102 is the same as that in the arithmetic unit 12 in FIG. This output is the arithmetic units 103 and 104.
Is input to the above-described pitch attitude angle or pitch attitude angle command, and the schedule gain defined is multiplied respectively. The multiplication result in the calculator 103 is added to the adders 401 and 40 as a pitch elevon steering angle command δ _PE.
Is output to 2, pitch vanes steering angle command [delta] _PV is output to the adder 403 and 404 from the arithmetic unit 104. Here, each of the arithmetic units 103, 104, 203, 2 in FIG.
The graphs indicated by 04, 303 and 304 illustrate the relationship between the pitch attitude angle or the gain with respect to the pitch attitude angle command. These schedule gains correspond to the variable gains FA and FT in FIG.

Similar operations are performed for the roll axis and the yaw axis. The roll attitude angle command θ _RC and the actual roll attitude angle θ _R are input to the comparator 201, and the deviation is calculated by the calculator 2.
It is input to 02. PID control is performed in the computing unit 202, and the pitch angular acceleration command d ² θ _R / dt ² is output to the computing units 203 and 204. Arithmetic units 203 and 20
At 4, the pitch attitude angles or the schedule gain as shown for the pitch attitude angle command, respectively, are multiplied. The roll elevon steering angle command δ _RE output from the calculator 203 is output to the adders 401 and 402, and the roll vane steering angle command δ _RV output from the calculator 204 is output to the adders 403-406.

The yaw attitude angle command θ _YC and the yaw attitude angle θ _Y are input to the comparator 301, and the deviation is input to the calculator 302. The PID control is performed in the calculator 302, and the yaw angular acceleration command d ² θ _Y / dt ² is calculated by the calculators 303 and 30.
4 is output. The schedule gain is multiplied in the computing unit 303, and the yaw rudder rudder angle command δ _YR becomes the rudder 4
It is directly output to the actuator for 2 as an operation amount command. The multiplication result from the computing unit 304 is output to the adders 405 and 406 as a yaw vane steering angle command δ _YV .

A pitch elevon steering angle command δ is sent to the adder 401.
_PE and the roll elevon steering angle command δ _RE are input, and δ _PE −
δ _RE is calculated and output as an operation amount command to the actuator that operates the left-wing elevon 43a. Adder 4
Similarly, the pitch elevon steering angle command δ _PE and the roll elevon steering angle command δ _RE are also input to 02, but by setting the direction of arrow A shown in FIG. 2 to be positive, δ _PE + δ _RE , It is output as an operation amount command to an actuator that operates the right wing elevon 43b. As a result, the left-wing elevon 43a and the right-wing elevon 43b are operated.

The rudder 42 is operated by the actuator to which the yaw rudder steering angle command δ _YR is given as described above.

In the adders 403 and 404, the pitch vane rudder angle command δ _PV and the roll vane rudder angle command δ _RV are respectively input, δ _PV + δ _PV is output to the actuator that operates the vane 44a, and the actuator that operates the vane 44b. Then δ _{PV −} δ _RV is output as the manipulated variable command. In the adders 405 and 406, the roll vane steering angle command δ _RV and the yaw vane steering angle command δ _YV are added, δ _YV −δ _RV is output to the actuator that operates the vane 44c, and δ is output to the actuator that operates the vane 44d. _YV +
δ _RV is output as the manipulated variable command. As a result, the vanes 44a ...
44d is operated.

According to the present embodiment, the pitch attitude angle that can be easily measured by a vertical gyro or the like without using dynamic pressure that is extremely difficult to measure during transition flight, or the pitch attitude angle command calculated by the control computer. The optimum flight attitude can be controlled by operating the aerodynamic control surface and the thrust deflecting device using the schedule gain represented by a simple function for. Therefore, it is possible to obtain good control characteristics without requiring a special sensor or the like for measuring the dynamic pressure and without complicating the contents of calculation.

In the present embodiment, the operation of the thrust deflection vane provided behind the exhaust port of the engine as the thrust deflection device is controlled, but the present invention is also applied to a machine body using another attitude control device. Is possible. For example, when thrust is deflected by deflecting the exhaust nozzle of the engine, it can be applied to determine the command of the angle for deflecting this nozzle. Further, when the attitude control moment is obtained by the reaction force obtained by ejecting the bleed air of the engine, the present invention can be applied to the determination of the opening degree command of the ejection nozzle.

[0033]

As described above, the method for controlling the transitional flight attitude of the vertical attitude lander according to the present invention can easily obtain the pitch attitude angle or pitch attitude without using the dynamic pressure which is difficult to measure during the transition flight. Since the control is performed using the control gain scheduled as a function of the angle command, the flight attitude can be controlled by a simple method.

[Brief description of drawings]

FIG. 1 is a control block diagram showing a transitional flight attitude control method for a vertical attitude lander according to an embodiment of the present invention.

FIG. 2 is a perspective view showing an appearance of a vertical attitude landing machine to which the same transitional flight attitude control method can be applied.

FIG. 3 is a control block diagram showing attitude response characteristics regarding one axis in the same transition flight attitude control method.

FIG. 4 is a control block diagram showing attitude response characteristics when a variable gain is introduced.

5 is a control block diagram equivalent to the control block shown in FIG. 4. FIG.

FIG. 6 is an explanatory view showing values of variable gains FA and FT with respect to a pitch attitude angle or a pitch attitude angle command.

[Explanation of symbols]

101, 201, 301 Comparator 102-104, 202-204, 302-304 Computational unit 401-406 Adder 43a Left wing elevon 43b Right wing elevon 42 Rudder 44a-44d Vane

 ─────────────────────────────────────────────────── ─── Continuation of the front page (72) Inventor Naohiko Udagawa 2-23-16 Kamiochiai, Shinjuku-ku, Tokyo (72) Inventor Takashi Shinagawa 1-7-2, Nishishinjuku, Shinjuku-ku, Tokyo Within Fuji Heavy Industries Ltd.

Claims (1)

[Claims] 1. A method for performing attitude control in transition flight of a vertical attitude landing machine having an aerodynamic control surface for controlling the attitude of a vehicle body about three axes of a pitch axis, a roll axis and a yaw axis, and a thrust deflector, wherein: Determining the deviation between each attitude angle command and the actual attitude angle for the axis, determining the angular acceleration command according to the control law based on the deviation obtained, and the pitch attitude angle or as a function to the pitch attitude angle command. A step of obtaining a steering angle command using a scheduled control gain and the angular acceleration command, and a step of operating the aerodynamic control surface and a thrust deflector based on the obtained steering angle command to perform attitude control And a transitional flight attitude control method for a vertical attitude lander, comprising: