JP2620428B2 - Transition Flight Attitude Control Method for Vertical Attitude Lander - Google Patents

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 transition from a cruise state in a horizontal attitude to a hover state in a vertical attitude and makes a landing.

[0002]

2. Description of the Related Art A conventional technique for controlling the flight of a flying object is disclosed in Japanese Patent Application Laid-Open No. 63-75500. In this technique, when controlling a flying object flying toward a target, an attempt is made to perform good control by adapting to flight conditions that change from moment to moment. The flying object is equipped with a flight control gain calculator, and calculates a feedback gain by solving a state equation describing how the speed and the flying attitude change depending on the operation amount of the aerodynamic control surface and the thrust deflection device. Then, the flight speed and attitude are fed back to the operation amounts of the aerodynamic control surface and the thrust deflection device.

[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 state equation is large, leading to an increase in cost.

In control surface control, a dynamic pressure schedule gain is generally used. In this method, a control gain is determined as a function of a dynamic pressure which determines a magnitude of an aerodynamic force by a state quantity representing a flight condition, based on an aerodynamic characteristic which has been clarified in advance. Then, the dynamic pressure is measured during the flight, and the control gain is changed based on the measured value. However, in the process of the vertical attitude landing aircraft transitioning from the horizontal attitude to the vertical attitude, the angle of attack and the dynamic pressure rapidly change in addition to a large angle of attack. Therefore, it is extremely difficult to measure the dynamic pressure during the 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 transition flight attitude control of a vertical attitude landing aircraft capable of optimally controlling an aerodynamic control surface and a thrust deflecting device in response to a change in flight conditions during a transition flight. The aim is to provide a method.

[0006]

SUMMARY OF THE INVENTION The present invention is directed to a vertical flight landing aircraft having an aerodynamic control surface and a thrust deflector for controlling the attitude of an airframe with respect to three axes of a pitch axis, a roll axis, and a yaw axis. In the control method, a comparator calculates a deviation between each posture angle command and an actual posture angle with respect to the three axes by a comparator, and inputs the deviation to an arithmetic unit to perform a proportional operation, an integral operation, and a differential operation. Outputting an angular acceleration command, and inputting the output angular acceleration command to a calculator for an aerodynamic rudder steering angle command system and a calculator for a thrust deflecting device operating angle command system; F _A
And thrust deflection device operating angle command system schedule gain _FT
The relative proportional constant K _T representing the magnitude of the effect of the K _A and thrust deflection device representing the magnitude of Me virtue of aerodynamic control surface, K A _· F
_A + K _T · F _T = 1 is satisfied, and the pitch posture angle or pitch posture angle command is set so that K _A · F _A becomes 1 during horizontal flight and gradually decreases as the pitch posture angle increases, and becomes 0 during the vertical hover posture. Setting as a function, multiplying the pitch attitude angle or the angular acceleration command, which is a function of the pitch attitude angle command, to output an aerodynamic rudder surface steering angle command and a thrust deflection device operating angle command, and outputting the aerodynamic rudder. Operating the aerodynamic rudder surface and the thrust deflection device based on the rudder angle command and the thrust deflection device operating angle command to perform attitude control.

[0007]

[Action] pitch axis relates the roll axis and yaw axis, the deviation is determined and the actual attitude angle and an attitude angle command, respectively, proportional operation, the angular acceleration command is determined by the integral operation and differential operation, further K _A · F _A + K _T · F _T = 1 is satisfied, and the pitch posture angle or pitch posture angle command is set so that K _A · F _A becomes 1 during horizontal flight and gradually decreases as the pitch posture angle increases, and becomes 0 during the vertical hover posture. F _a and F _T that is set as a function is multiplied by the angular acceleration command and the aerodynamic rudder Omokaji angle command and thrust deflection apparatus operating angle command obtained. The attitude control is performed by operating the aerodynamic control surface and the thrust deflection device based on this command. Thus, without using the difficult dynamic pressure measurements during the transition flight, the proportional constant K _A and K _T is a function of pitch attitude angle command obtained by easily pitch attitude angle is measured or easy calculations for using schedule gain determined by the relationship equation between the variable gain F _a and F _T, it is possible to optimize the posture control by a simple method.

[0008]

An embodiment of the present invention will be described below with reference to the drawings. First, the appearance of a vertical attitude landing aircraft to which the present embodiment is applied is shown in FIG. The left and right elevons 43 for controlling the pitch axis and the roll axis as aerodynamic control surfaces
a and 43b, and a ladder 42 for controlling the yaw axis. Further, as the thrust deflection device, thrust deflection vanes 44a to 44d are provided behind the exhaust port of the engine. Each of the thrust deflection vanes 44a to 44d has four vanes arranged in a cross shape, and each vane moves independently like a moving blade. The left and right elevons 43a to 43b and the rudder 42 as aerodynamic control surfaces are used for flight in a horizontal attitude, and the thrust deflection vanes 44a to 44d as thrust deflection devices are used for flight in a vertical hover attitude. And, during the transition flight, the left and right elevons 43a to 43a-4
3b and the rudder 42 and the thrust deflection vanes 44a to 44d are used in combination. Here, the elevons 43a and 43b take the direction of the arrow A in the positive direction, the vanes 44a and 44b take the direction of the arrow A, and the vanes 44c and 44d take the direction of the arrow B in the positive direction. 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. The attitude response characteristic when one of these axes is controlled is represented as a block diagram in 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 attitude angle θ output from the controller 12 is input to the calculator 12. The arithmetic unit 12 performs an operation using the transfer function G _C1 of the control law, and calculates the steering angle δ _{A of the} aerodynamic control surface.
And the operation angle [delta] _T of the thrust deflection device is output. here,
It is assumed that the steering characteristic and the steering angle are equal, ignoring the dynamic characteristics of the control surface and the thrust deflection device, that is, the steering angle and the delay of the operating angle with respect to the command. The portion 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 deflection device operating angle (broken line portion 17 in FIG. 3).

The steering angle δ _A and the operating angle δ _T output from the calculator 12 correspond to the amount of change from the balanced state. Steering angle [delta] _A is input to the calculator 13 is multiplied by the proportional constant K _A, K _{A · δ A} is output to the adder 15. Operating angle δ _T
Is output multiplied by the proportional constant K _T by the arithmetic unit 14 to the adder 15. Here, proportionality constant K _A represents the magnitude of the effectiveness of the aerodynamic control surface, the proportionality constant K _T represents the magnitude of the effect of the thrust vectoring device. Further, the and the operating angle [delta] _T and the steering angle [delta] _A is multiplied by a proportional constant 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 body fluctuates is smaller than the control moment generated by the control surface or the thrust deflection device and can be ignored. Therefore, the moment acting on the fuselage can be considered to occur in proportion to the operating angle [delta] _T and the steering angle [delta] _A. Thereby, the angular acceleration d ² θ / dt of the posture angle θ is obtained.
² would be considered to be proportional to the operating angle [delta] _T and the steering angle [delta] _A, is set to be multiplied by a proportional constant K _A and K _T.
The angular acceleration d ² θ / dt ² output from the adder 15 is
The integration is performed twice by the integrator 16, and the actual attitude angle θ of the airframe is output. This attitude angle θ is fed back to the comparator 11.

[0011] proportional constant K _A and K _T here, speed and posture in the transition flight varies widely depending on the change in the thrust of the engine. Therefore, if the transfer function G _C1 is an invariant function, the response characteristic of the actual command angle θ to the attitude command angle θ _C also changes greatly. Therefore, as in the control block of FIG. 4, to introduce a new variable gain F _A and F _T. Here, the variable gain F _A is aerodynamic steering Omokaji angle command based scheduling gain, a variable gain F _T is thrust deflection apparatus operating angle command based scheduling gain. Comparing the control blocks shown in FIG. 4 to that of Figure 3, the transfer function G _C2 invariant used in arithmetic unit 12, while the arithmetic unit 12 and the arithmetic units 13 and 14, variable gain F _A and calculator 21 and 22 perform multiplication using F _T is different in that it is connected in series. Thereby, the arithmetic unit 1
Output ^{d 2} .theta.c / dt ² from 2 are applied to respective arithmetic units 21 and 22, is outputted to the arithmetic unit 13 and 14 as a variable gain _{F A} and _{F T} is multiplied steering angle [delta] _A and the duration [delta] _T You. In the calculators 13 and 14, the proportional constant K is set to the steering angle δ _A and the operating angle δ _T similarly to the control block of FIG.
_A and _KT are respectively multiplied to obtain the angular acceleration d ² θ / d
t ² is output.

The following equation (1) must be established between the variable gains F _A and F _T and the proportional constants K _A and K _T.

F _A K _A + F _T K _T = 1 (1) By satisfying this relational expression, the control block shown in FIG. 4 has a relation equivalent to that of FIG. As a result, the response characteristics of the flight attitude become constant.

The proportional constants K _A and K _T have the following relationship among 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 determined in advance by the aerodynamic characteristics and the characteristics of the thrust deflection device.

[0016] From this (2) and (3), angle of attack during the flight alpha, the value of the proportional constant K _A and K _T is determined by measuring the value of the dynamic pressure Q and thrust T. Thereby, the variable gain F
It becomes possible to change the values of _A and F _T every moment during the transition flight. In this case, the thrust T can be estimated from the engine speed, temperature, and the like. However, it is difficult to measure the angle of attack α and the dynamic pressure Q because they change rapidly during the transition flight.

Therefore, the following handling is performed. In the transition flight, a certain attitude and speed during horizontal flight are caused at a predetermined speed, and the flight pattern is limited, such as flying along a predetermined route. 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 monotonously increases during the transition flight. Therefore, alpha as a function of the pitch attitude angle or pitch attitude angle command, it is possible to determine the values of Q and T, captures the value of the proportional constant K _A and K _T as a function of pitch attitude angle or 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 the control characteristics can be kept almost constant by including them in the control law as the schedule gains.

[0019] Here, the value of the variable gain _{F A} and _{F T} is
Relationship of the above equation (1) between the constants K _A and K _T are considered indefinitely as a combination. However, even in an infinite number of combinations, in an actual airframe, in general, transition flight must be able to smoothly transition from a flight state in a horizontal attitude to a flight state in a vertical hover attitude. The control surfaces of the fuselage and thrust deflecting devices also have limited capabilities. If a control that greatly depends on the aerodynamic control surface is selected in a region where the effectiveness of the aerodynamic control surface is reduced, smooth transition flight is greatly impeded by the influence of disturbance. The same can be said for the control in which the thrust is small and the control largely depends on the thrust deflection device. Therefore, the control using the more effective one of the aerodynamic control surface and the thrust deflector may be performed.However, the method of switching the aerodynamic control surface or the thrust deflector at a certain pitch attitude angle at once is based on the response of the control surface or the thrust deflector. Due to the different sexes, the control force becomes discontinuous, and a smooth transition flight cannot be achieved. In other words, for a smooth transition flight, the dependence of the aerodynamic control surface on the aerodynamic control surface is increased in the area where the effectiveness of the aerodynamic control surface is good and the effect of the thrust deflection device is poor, and conversely, the effectiveness of the thrust deflection device is poor with the effect of the aerodynamic control surface It is necessary to increase the dependence on the thrust deflecting device in an area where is good, and to gradually decrease or increase the dependence in accordance with the change in the pitch attitude angle.

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
9 shows the relationship between 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 solely by the aerodynamic control surface, except in special cases using thrust deflection devices. Therefore, in the region where the pitch attitude angle is small, the variable gain F _T for thrust deflection is used.
Is set to 0 and F _A = 1 / K _A.

Conversely, in the vertical hover attitude, the aerodynamic control surface does not affect the flight attitude, and is controlled only by the thrust deflection device. Therefore, a large area of the pitch attitude angle is set to 0 the variable gain F _A related control surface, and F _T = 1 / K _T.

A transition flight during transition from a horizontal flight to a vertical hover flight is as follows. In the case of the transition flight pattern in which the altitude is kept almost constant, the thrust increases so as to support the weight of the fuselage as the lift decreases. In the vicinity of the maximum lift coefficient attack angle corresponding to the midpoint between the horizontal flight and the vertical hover attitude during the transition flight, the dynamic pressure decreases, the effectiveness of the aerodynamic control surface is considerably reduced, and K _A indicating the effectiveness of the aerodynamic control surface Is small. However, since the lift coefficient is large in this vicinity, the decrease in lift is small, and the thrust is suppressed to a small value in order to balance the sum of the lift and the vertical component of thrust with the weight of the fuselage. _T also has a small value. When the pitch attitude angle becomes larger than the intermediate point, the dynamic pressure and the lift coefficient are greatly reduced, and the lift is drastically reduced.
_{Although A} becomes smaller, the thrust is increased so as not to lower the altitude, so that the effectiveness of the thrust deflection device is increased and _KT is increased. If the equation (1) is satisfied along with the increase in the pitch attitude in the transition flight, F _T · K _T because the thrust is still low and the _KT is small before the intermediate point.
Since the large take when it comes to setting the capacity on excessive F _T of the thrust vectoring device, it is necessary to increase the F _{A ·} K A. Because the K _A is decreases with an increase in pitch attitude,
In accordance with the rate of decrease is set so F _A increases. In the following the vicinity of the middle point, because more K _A can not be maintained the size of the F _A · K _A on a reasonable F _A longer reduced aerodynamic control surface capability, thrust deflection which has increased the K _T by a thrust increase device to keep up a smaller F _a becomes appropriate to transfer the specific gravity. Thus F _A near the midpoint takes a maximum value. Relationship between the thrust vectoring device and the pitch attitude angle is a tendency opposite to the aerodynamic control surface, F _T becomes maximum again near the midpoint. F _A
Specifically configuration steps F _T defines a pitch attitude is first expected to sum the K _A and K _T is minimized and the midpoint, decided slightly larger F _A than during level flight for this intermediate point and ,
(1) seeking the value of _{F T} by applying the formula to set the maximum value of _{F A} and _{F T.} The value of this F _T is greater than the value at the time of vertical hover position, applying the above equation (1). Then between the near horizontal flight and the intermediate point, between the vicinity of the intermediate point and the vertical hover attitude, and F _A is applied to progressively decrease to increase the value of F _A · K _A (1) Equation determine the F _T. If the value of F _A and F _T before and after the intermediate point decided in this way, as long as it greatly exceed the capabilities of the aerodynamic control surface and thrust vectoring device, the F _A and F _T midpoint re Set.

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 equation (1) is strictly maintained, a curve is obtained. However, even if approximation is made by a polygonal line as shown in FIG. 6, there is no problem in practical use, and control calculations can be simplified.

A transition flight attitude control method of the present embodiment for controlling three axes in the above manner will be described with reference to the control block diagram of FIG. For 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, respectively,
Left-wing elevon 43a, right-wing elevon 43b, rudder 4
2. Obtain an operation amount for operating the vanes 44a to 44d by the actuator. Thereby, the attitude of the aircraft can be changed as instructed.

The pitch attitude angle command θ _PC and the pitch attitude angle θ _P of the actual 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,
The proportional action P using the proportional constant K _PP and the integral constant K _IP
PID control including an integral operation I using the Laplace operator s, a differential constant _KDP and a differential operation D using the Laplace operator s is performed, and the pitch angular acceleration command d ² θ
_P / dt ² is output. Here, the content of the operation performed by the arithmetic unit 102 is the same as that of the arithmetic unit 12 in FIG. This output is output to arithmetic units 103 and 104
, And the pitch attitude angle or the pitch attitude angle command as described above is multiplied by the defined schedule gain. The result of the multiplication in the arithmetic unit 103 is calculated as a pitch elevon steering angle command δ _{PE by the} adders 401 and 40.
2 and the computing unit 104 outputs the pitch vane steering angle command δ _PV to the adders 403 and 404. Here, each computing unit 103, 104, 203, 2 in FIG.
The graphs shown at 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 on 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 arithmetic unit 2
02 is input. PID control is performed in the arithmetic unit 202, and the pitch angular acceleration command d ² θ _R / dt ² is output to the arithmetic units 203 and 204. Arithmetic units 203 and 20
At 4, the pitch attitude angle or the schedule gain as shown for the pitch attitude angle command is multiplied, respectively. The roll elevon steering angle command δ _RE output from the computing unit 203 is output to adders 401 and 402, and the roll vane steering angle command δ _RV output from the computing unit 204 is output to adders 403 to 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. PID control is performed in the arithmetic unit 302, and the yaw angular acceleration command d ² θ _Y / dt ² is calculated by the arithmetic units 303 and 30.
4 is output. The scheduler gain is multiplied by the arithmetic unit 303, and the yaw rudder steering angle command δ _YR is
It is directly output as an operation amount command to the second actuator. Multiplication result from the arithmetic unit 304 is output to the adder 405 and 406 as Yoben steering angle command [delta] _YV.

The pitch elevon steering angle command δ is supplied to the adder 401.
_PE and the roll elevon steering angle command δ _RE are input, and δ _PE −
The calculation of δ _RE is performed 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. By making the direction of the arrow A shown in FIG. 2 positive, δ _PE + δ _{RE is obtained} . An operation amount command is output to an actuator that operates the right wing elevon 43b. Thereby, the left wing elevon 43a and the right wing elevon 43b are operated.

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

In the adders 403 and 404, the pitch vane steering angle command δ _PV and the roll vane steering angle command δ _RV are respectively input, and δ _PV + δ _PV is output to the actuator for operating the vane 44a, and the actuator for operating the vane 44b. 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, and 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 each of the manipulated variable commands. Thus, the vanes 44a to 44a
44d is operated.

According to this embodiment, a pitch attitude angle which can be easily measured by a vertical gyro or the like without using a dynamic pressure which is extremely difficult to measure during a transition flight, or a pitch attitude angle command calculated by a control computer By operating the aerodynamic control surface and the thrust deflection device using a schedule gain expressed by a simple function with respect to the above, optimal flight attitude control can be performed. Therefore, a good control characteristic can be obtained without requiring a special sensor or the like for measuring the dynamic pressure and without complicating the calculation.

In this embodiment, the operation of the thrust deflection vane provided behind the exhaust port of the engine is controlled as a thrust deflection device. However, the present invention can be applied to a body using another attitude control device. Is possible. For example, when thrust deflection is performed by deflecting an exhaust nozzle of an engine, the present invention can be applied to determining a command for deflecting the nozzle. 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, according to the transition flight attitude control method of the vertical attitude landing aircraft of the present invention, the pitch attitude angle or pitch attitude that can be easily obtained without using the dynamic pressure that is difficult to measure during the transition flight. Since the control is performed using the schedule gain obtained by the relational expression between the proportionality constant, which is a function of the angle command, and the variable gain, the flight attitude can be controlled by a simple method.

[Brief description of the drawings]

FIG. 1 is a control block diagram illustrating a transition flight attitude control method for a vertical attitude landing aircraft according to an embodiment of the present invention.

FIG. 2 is a perspective view showing the appearance of a vertical attitude landing aircraft to which the transition flight attitude control method can be applied.

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

FIG. 4 is a control block diagram showing a posture response characteristic when a variable gain is introduced.

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

FIG. 6 is an explanatory diagram 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 Computing device 401-406 Adder 43a Left elevon 43b Right elevon 42 Ladder 44a-44d Vane

 ──────────────────────────────────────────────────続 き Continued on the front page (72) Inventor Naohiko Udagawa 2-23-16 Kamiochiai, Shinjuku-ku, Tokyo (72) Inventor Takashi Shinagawa 1-7-1-2 Nishi-Shinjuku, Shinjuku-ku, Tokyo Inside

Claims (2)

(57) [Claims] 1. A method for performing attitude control in a transition flight of a vertical attitude lander having an aerodynamic control surface for controlling attitude of an airframe with respect to three axes of a pitch axis, a roll axis, and a yaw axis and a thrust deflection device, Calculating a deviation between each attitude angle command and the actual attitude angle by a comparator with respect to the axis; and outputting the angular acceleration command by inputting the deviation to a calculator to perform a proportional operation, an integral operation, and a differential operation. Inputting the output angular acceleration command to an aerodynamic surface control angle command system calculator and a thrust deflection device operating angle command system calculator.
Aerodynamic steering Omokaji angle command based scheduling gain F _A and thrust deflection apparatus operating angle command based scheduling gain F _T, proportional constant K which represents the magnitude of the effect of the K _A and thrust deflection device representing the magnitude of the effectiveness of the aerodynamic control surface _{For T} , K _A · F _A + K _T ·
Together to satisfy _{F T = 1, K A ·} F A is set as a function of pitch attitude angle or pitch attitude angle commanded to the zero at the vertical hover position gradually decreases as the increase in the pitch attitude angle 1 at level flight Multiplying the angular acceleration command, which is a function of the pitch attitude angle or the pitch attitude angle command, to output an aerodynamic rudder angle command and a thrust deflecting device operation angle command, and outputting the aerodynamic rudder angle command. Operating the aerodynamic control surface and the thrust deflection device based on a thrust deflection device operating angle command to perform attitude control. 2. The step of outputting the aerodynamic rudder rudder angle command and the thrust deflector operating angle command, wherein a schedule gain that changes in a polygonal line approximating the curve gain that changes in a curve is used as the angular acceleration command. 2. The transition flight attitude control method for a vertical attitude landing aircraft according to claim 1, wherein the aerodynamic rudder rudder angle command and the thrust deflection device operation angle instruction are obtained by multiplication.