JPH0958597A - Vertical landing control command device for flying object - Google Patents

Abstract

PROBLEM TO BE SOLVED: To perform automatic high-precise vertical landing even when disturbance is present by a method wherein when a flying object, such as a rocket, is vertically landed on a ground or a surface of a heavenly body from a space, a control command for engine thrust and a direction is outputted based on the present position, the speed, and acceleration of a machine body. SOLUTION: During flying of a flying object, a first command generating part 60 input an indication signal from a processing mode indication part 2 to select one of a landing mode, a space stop mode, and a transition operation mode, a set signal from a target state setting part 3, a signal for the position, the speed, and acceleration of the flying object from a movement state detecting part 20 and outputs a command signal. Further, a second command generating part 70 inputs signals from acceleration command generating parts 4 and 5 for landing control and space stop control and a signal from a machine body mass estimating part 8 and outputs a command signal. Commands from the command generating parts 60 and 70 are repeatedly executed at intervals of a given time by a switcher S9, and a thrust command for a present engine and a thrust direction command are varied.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a vertical landing control command device used for controlling vertical landing of a vehicle such as a rocket equipped with a thrust rocket engine.

[0002]

2. Description of the Related Art The conventional method for automatically landing a rocket or other flying vehicle from outer space onto the surface of the earth, moon, or planets such as planets is based on the landing control realized by the Apollo lunar lander. As can be seen, a command for the magnitude of the thrust of the rocket engine and a command for the direction in which the thrust works are generated based on the current position and speed of the air vehicle (Reference 1). (Reference 1) ARKLUMPP, APOLLO LUNAR DESCENT GUIDANC
E, Automatica vol10 No.2P133 ~ 146, May 1974 However, in rockets and other aircraft, disturbances occur due to engine mounting position and mounting angle deviations, and variations in the center of mass of the aircraft itself. .

Further, in a rocket landing vertically on the ground for the purpose of reuse, aerodynamic force acts as a disturbance. Under such a situation, it is not easy to perform vertical landing at the target point with high accuracy by the command method based only on the current position and speed.

Therefore, humans perform the operation just before the landing of the Apollo moon landing. Also, before or during the landing maneuver, due to abnormalities of the air vehicle, evacuation behavior for safety near the landing point, or other reasons, another designated point (airborne It is also necessary to quickly move the air vehicle to the specified condition (position, speed) automatically while minimizing the consumption of propellant.
However, the control device does not exist (Reference 1).

[0005]

With the conventional method of issuing a control command for the magnitude and direction of the engine thrust based on the current position and speed of the aircraft, it is easy to automatically and vertically land the aircraft. is not.

During the flight before landing, when it becomes necessary to shift the flying body from the current state to any other state due to evacuation behavior or other reasons, while automatically minimizing the propellant consumption, A command device that can be quickly realized is needed.

According to the present invention, by issuing a control command for the engine thrust and the direction based on the current position, speed and acceleration of the airframe, it is possible to automatically and accurately perform a vertical landing even in the presence of disturbance. , To quickly move the aircraft from any state during flight before landing to any state within a predetermined space by a method that automatically minimizes the consumption of engine propellant (fuel) It is an object of the present invention to provide a device capable of

[0008]

[Means for Solving the Problems]

(First Means) A vertical landing control command device for an aircraft according to the present invention is, in an aircraft having a rocket engine capable of changing the magnitude of thrust, (A) processing mode instructing section and target state setting. Unit, a first command generation unit, a second command generation unit, a body mass estimation unit, and a vertical landing control command device basic unit including a switch S, (B) the processing mode command unit is a mode According to the instruction, one of the landing mode (a), the space stop mode (b) or the transition operation mode (c) is selected, and (C) the target state setting unit corresponds to the mode selected by the processing mode instructing unit. The state of the target point is selectively set, and (D) the first command generation unit includes a landing control acceleration command generation unit, a spatial stop control acceleration command generation unit, a transition motion control thrust and a thrust direction command generation unit. Consists of the processing mode fingers And a signal from the target state setting unit, and a position, velocity and acceleration signal of the flying body from the motion state detecting unit, and (E) the second command generating unit causes the landing control acceleration. Input the signal from the command generator and the signal from the space stop control acceleration command generator, and also input the signal from the airframe mass estimator.
(F) The airframe mass estimating unit inputs signals from the transition motion control thrust of the first command generating unit, the thrust direction command generating unit, and the second command generating unit, and (G) the switching unit S makes landing. In the mode (a) and the space stop mode (b), it is turned on to the A side, and in the transition operation mode (c), it is turned on to the B side, and the target state corresponding to the mode specified before flight or during flight is set, It is characterized in that the command from the command generating unit is repeatedly executed at every designated processing time interval Δt _c to change the current thrust command and thrust direction command of the rocket engine. (Second Means) A vertical landing control command device for an aircraft according to the present invention uses the first means for vertical landing when the thrust direction of the rocket engine has a one-to-one correspondence with the attitude of the aircraft. It is characterized in that a posture command unit for converting a thrust direction command into a posture command of a flying body is added to the basic unit of the control command device. (Third Means) A vertical landing control command device for an aircraft according to the present invention is a vertical landing control command device for a vehicle in which the rocket engine can be tilted by an engine nozzle and an actuator around a mounting point on the aircraft. It is characterized in that an engine rudder angle command unit for converting a thrust direction command into an engine rudder angle command is added to the landing control command device basic unit.

(1) The vertical landing control command device basic unit 1 is installed inside an aircraft equipped with a thrust variable rocket engine (see FIG. 8 for the definition of the coordinate system). (2) The vertical landing control command device basic unit 1 is based on the rocket engine based on the current position, velocity, and acceleration signals of the flying object generated from the detection signals of various motion state detectors (sensors) 20 equipped on the flying object. Generates a thrust magnitude command and a thrust direction command.

In the usual case, the flight modes are automatically realized in the order designated in advance before the flight, but a function to allow the mode instruction from the outside is added if necessary. (3) According to the instruction of each processing mode instruction unit 2, the current mode is (a) landing mode, (b) space stop mode,
(C) Transition operation mode is selected, and the state of the target point corresponding to the mode is selected and set by the target state setting unit 3. (4) In the landing mode, the landing control acceleration command generator 4
Generates three-axis acceleration signals that the aircraft should have at the present time for realizing highly accurate vertical landing based on the current position, velocity and acceleration of the aircraft.

In the space stop mode, in order to maintain the position and speed of the target point, the space stop control acceleration command generator 5
Generates a triaxial acceleration signal that the aircraft should have at the present time based on the current position and velocity of the aircraft.

In the transition operation mode, it is designated in advance before entering the landing mode, or in the case of transitioning to the evacuation flight due to an abnormality of the flying object or safety near the landing point during flight in the landing mode, etc. In order to reach the retreat point, the transition motion control thrust and thrust direction command generation unit 6 is based on the current position and speed of the flying vehicle, and commands the magnitude of the thrust that the flying vehicle should have and the biaxial Directly generate thrust direction commands (X and Z).

The thrust direction command is represented by the thrust axis direction cosine. In the space stop mode and transition motion mode, control based on the current position and speed may be sufficient depending on the purpose. (5) In the landing mode and the space stop mode, it is necessary to convert the acceleration signals of the three axes into the thrust command and the thrust direction command of the shaft.

The body mass estimation unit 8 calculates the mass (in this case, the amount of propellant remaining in the rocket engine) from the thrust command before the present time.
The airframe mass is estimated by estimating. The thrust and thrust direction command generator 7 generates a thrust command and a biaxial (X, Z) thrust direction command based on the current estimated value of the body mass and the triaxial acceleration signal. (6) The switch S is A in the landing mode and the space stop mode.
The side is turned on, and the B side is turned on in the transition operation mode. (7) If the mode is specified before or during flight,
The target state is set. After that, the commands from the command generators 60 and 70 are repeatedly executed at the processing time interval Δt _c designated in advance to update the current thrust command and thrust direction command (see FIG. 4).

[0015]

BEST MODE FOR CARRYING OUT THE INVENTION

(First Embodiment) FIGS. 1, 4, 5, 6, 7, and 8 show a first embodiment of the device of the present invention. FIG. 1 shows the basic configuration of the device of the present invention.

FIG. 4 shows a flow chart of processing of the apparatus of the present invention. FIG. 5 is an explanatory diagram of a function f that guides the engine thrust acceleration. FIG. 6 is an explanatory diagram of Y elements of functions p and q for calculating predicted values of speed and position at the arrival time of the target point.

FIG. 7 is an explanatory diagram of the X and Z elements of the functions p and q for calculating the predicted values of the speed and the position at the arrival time of the target point. FIG.
FIG. 6 is a diagram showing the definition of a coordinate system and the like.

In the figure, (1) the processing mode instructing section 2 determines the mode related to the vertical landing motion as (a) the landing mode, based on the result of monitoring the internal condition of the aircraft or the mode instruction from the outside.
One of the (b) space stop mode and the (c) transition operation mode is selected and instructed. (2) The target state setting unit 3 determines the state of the target point (for example, landing point, one point in space, etc.) corresponding to each mode,
Set the position (￣r _T ), velocity (￣v _T ) and acceleration (￣a _t ) displayed in the coordinate system with the origin on the surface of the earth or other celestial body defined in Fig. 8 as a function of the time to the target point. .

The symbol (-) on the symbol indicates a three-dimensional vector. (-A _T ) is only for mode (a). This numerical value is stored in advance inside the generator 1.

For the sake of convenience, the acceleration (-a _T ) includes the gravity component of the celestial body. Further, in the modes (a) and (b), the initial value T ₀ of the arrival time to the target point (that is, the control end time) T is also set. (3) The current position (￣r), velocity (￣v), and acceleration (￣a) of the air vehicle (including the gravity component as in (￣a _T ) in the definition of (￣a)) are the target points. Similarly, it is given based on the data of the detector mounted on the aircraft in the coordinate system defined in FIG. (4) The landing control acceleration command generator 4 generates the engine thrust acceleration a _ti by the following iterative formula derived from the linear optimal control theory as the current triaxial acceleration signal in the landing mode.

A _ti (K) = f _i (r _i , v _i , a _i ) ·
_{Δt c + a ti (K-} 1), provided that i = X, Y, Z, f i (r i, v i, a i) indicates a value determined by a function of _{r i, v i, a i} , The details are shown in FIG.

The subscripts X, Y and Z represent vector components. K indicates the value of the current cycle, K-1 is 1
Shows the value of the previous cycle.

Δt _c is the time interval of the cycle. The iterative initial value a _ti (0) is the engine thrust speed at the start of the mode. (5) The spatial stop control acceleration command generator 5 generates the engine thrust acceleration a _ti (K) according to the following equation.

[0024] _{a ti (K) = h i} (δr i, δv i, δw i (K)), however, i = X, Y, Z where .delta.w (0) from integration of δw = ∫δr _d , Δr
Means the initial value of the integral of.

[0025] However, δ (¯r) = (¯r ) - (¯r T), δ (¯v) = (¯v) - (¯v T), δ (¯w) (K) = δ ( ￣ r) ・ Δt _c + δ (￣ w)
(K-1), and h _i is a linear function of δr _i , δv _i , δw _i (K). (Example), h _i = K _3i δv _i −K _2i δr _i −K _1i δw _i (K), where K _1i , K _2i and K _3i are constants, and (6) thrust for transition motion control and thrust direction command Generator 6
Applies the optimal control theory of minimization of propellant consumption,
A thrust magnitude command and a thrust direction command are generated as follows.

The control command for the transition motion is displayed by the vector amount u normalized by the maximum value Tmax of the thrust magnitude. Then, the absolute value of u is used as the thrust magnitude command, and the horizontal component (U _X , U _Z ) of u is given as the thrust direction command to the engine body, thereby realizing the command u as a vector amount.

In this case, the absolute value of u is obtained from the motion in the y direction on the assumption that the inclination of the thrust axis from the vertical is small (10 degrees or less) (without loss of practicality). Then, the following t ₁ and t ₂ are also obtained from the motion in the y direction.

The horizontal component (U _X , U _Z ) of U is obtained from the movements in the Y and Z directions. (6-1) A method of obtaining the absolute value of u. A thrust command is generated by the Y-axis motion control process.

At the time of mode setting, the current state and the target state are compared with each other to initialize the thrust command of the bang / bang control (the bang / bang control is a control method such as the pattern u described below). ).

Let t = 0 be the current time, t _{1 be} the switching time to u = u ₂ , and t ₁ + t ₂ be the target point arrival time. The pattern of the normalized thrust command u (0 ≦ u ≦ 1) is determined as follows.

When 0 ≦ t ≦ t ₁ , u = u ₁ t ₁ <t ≦ t ₁ + t ₂ and u = u ₂ , where u = t _c / Tmax (ratio between thrust command value and engine maximum thrust), u ₁ Is a constant, and u ₂ is a constant until the thrust command is switched, as will be described later, and thereafter becomes a variable whose value is determined by processing.

The determined thrust command u pattern and the current position,
From the velocity, the position of the target point and the velocity are displayed in the following form as an analytical solution of the equation of motion in the Y-axis direction (see FIG. 6). v _Y (t ₁ + t ₂ ) = p _Y (t ₁ , t ₂ , u ₁ , u ₂ ) r _Y (t ₁ + t ₂ ) = q _Y (t ₁ , t ₂ , u ₁ , u ₂ ) where v _Y (t ₁ + t ₂ ) and r _Y (t ₁ + t ₂ ).
Is v _Y at the undetermined target point arrival time t ₁ + t ₂ and r _Y
Is the predicted value of

P _Y and q _Y are t ₁ , t ₂ and u, respectively.
It is shown by the function of ₁ and u ₂ . In the phase before switching the thrust command (before reaching t ₁ ), u ₁ and u ₂ are fixed so that v _Y (t ₁ + t ₂ ) and r _Y (t ₁ + t ₂ ) match the target value. , That is, δv _Y = v _Y (t ₁ + t ₂ ) −v _TY = 0 δr _Y = r _Y (t ₁ + t ₂ ) −r _TY = 0, the pair of t ₁ and t ₂ can be determined. it can.

The determining method is based on the Newton method or another iterative method. At this time, the current thrust command is T _c = u ₁
It is generated as Tmax. After the thrust command is switched (after t ₁ ), both p _Y and q _Y are displayed in the functional form of t ₂ and u ₂ , so that t ₂ and u ₂ satisfying δv _Y = 0 and δr _Y = 0.
The set of can be defined by a similar iterative method.

From this, the current thrust command T _c = u ₂
Generate Tmax. (6-2) A method of obtaining the horizontal component (U _X , U _Z ) of U. By applying the optimal control theory combined with the Y-axis, the thrust command is generated based on the motion control processing of the X-axis and the Z-axis.

As a suboptimal control pattern (suboptimal means that it is derived by approximation for simplification),
Set the X and Z axis thrust direction commands using the following formula. u _i = d _i + e _i · t, where 0 ≦ t ≦ t ₁ + t ₂ , i = X, Z By performing processing in the Y-axis direction, t ₁ and t at the present time
₂ , u ₁ and u ₂ are known, and an analytical solution of the X-axis and Z-axis equations of motion can be obtained using the thrust direction command pattern set above.

This is expressed by the following equation (see FIG. 7). v _i (t ₁ + t ₂ ) = p _i (d _i , e _i ), r _i (t ₁ + t ₂ ) = q _i (d _i , e _i ), where i = X, Z where p _i (D _i , e _i ) and q _i (d _i , e _i ) are undetermined constants, which means that they are displayed by a function of d _i , e _i .

The set of constants d _i , e _i can be determined so that the analytical predicted value matches the target value. That is, a set of constants that satisfy the following can be determined. δv _i = v _i (t ₁ + t ₂ ) −v _Ti = 0 δr _i = r _i (t ₁ + t ₂ ) −r _Ti = 0, where i = X, Z The thrust direction command at the present time is t = 0. As a result, u _X = d _X and u _Z = d _Z are generated. (7) The thrust and thrust direction command generator 7 operates in the modes (a) and (b), and the thrust acceleration vector (
The thrust command based on a _t) and body mass estimate _{m (K), T c =} m (K) | ¯a t |, in the calculated result the engine output restriction range (Tmin ≦ T _c
≤ TMAX).

The || a _t | corresponding to the limited T _c is back-calculated to generate the thrust direction command u _X = a _tX / || a _t |, u _Z = a _tz / || a _t |, Let (8) The body mass estimation unit 8 calculates the current body mass m (K).
Is estimated by the previous thrust command and the functional form of the airframe mass, m (K) = M (T _c (K-1), m (K-1)).

A simple example of the function M is: M = m (K-1)-(1 / c) T _c (K-1) Δt
_c ,.

Here, C is a constant that depends on the engine. (9) FIG. 4 shows a flowchart of the processing. In FIG. 4, steps S3 to S12 are repeatedly performed at each time Δt _c when the same mode is continued.

Δt _c is from step S3 (mode branch) to
Steps S12 (thrust command, thrust direction command output) are the calculation time intervals when the calculation is repeated and all have the same constant.

However, there is a degree of freedom in the sense that the value of Δt _c can be changed depending on the mode. (Second Embodiment) FIG. 2 shows a second embodiment of the present invention.
And FIG.

The second embodiment is used when the rocket engine is fixedly attached to the flying body. In the second embodiment, since the thrust direction of the rocket engine has a one-to-one correspondence with the attitude of the airframe, the vertical landing control command device basic unit 1 of the first embodiment flies from the thrust direction command. The posture command unit 10 for converting the body posture command is added.

The output of the posture command section 10 is a biaxial posture angle command (posture control around the pitch axis and the yaw axis).
The method is to convert U _X and U _Z into attitude angles.

The relationship between the fixed coordinate system of an arbitrary flying object having the center of mass of the flying object as the origin and the fixed coordinate system of the landing point is represented by Euler angles or the like. For example, if the engine is fixedly installed on the centerline of the air vehicle, the attitude angles ψ and θ'of the thrust axis (aircraft center axis) with respect to the X axis of the fixed coordinate system at the landing point are converted as shown in FIG. To be done.

The attitude angle is changed by an auxiliary engine or a gas jet engine (not shown) for attitude control. (Third Embodiment) FIG. 3 shows a third embodiment of the present invention.
Shown in

The third embodiment is used when the rocket engine can be tilted (symbaling) around the attachment point to the flying vehicle 100 by the engine nozzle 12 and the actuator 13.

In the third embodiment, an engine rudder angle command for converting a thrust direction command into an engine rudder angle (engine nozzle tilt angle) command is given to the vertical landing control command device basic unit 1 of the first embodiment. The part 11 is added. Realization of this function requires the current attitude information of the air vehicle (see S4, S5, S6 in Fig. 4).

The output of the engine rudder angle command section 11 is a biaxial engine rudder angle command, and the attitude of the aircraft about the pitch axis and yaw axis is controlled by inclining the thrust direction of the engine (symbaling). .

[0051]

Since the present invention is constructed as described above, it has the following effects. (1) By outputting a control command based on the current position, speed, and acceleration of the flying object, automatic vertical landing on the ground or on the celestial body can be realized with high accuracy while reducing the influence of disturbance. (2) During the landing movement or before entering the landing operation, if it is necessary to evacuate or move due to an abnormality of the aircraft, safety problems near the landing point, etc., the state specified to the designated point in space , And can be automatically transitioned to minimize engine propellant consumption. (3) By selecting the three types of flight modes in any order on the ground or near the celestial body, it is possible to perform any automatic flight within the usable range of the engine propellant to be held.

[Brief description of drawings]

FIG. 1 is a diagram showing an apparatus according to a first embodiment of the present invention.

FIG. 2 is a diagram showing an apparatus according to a second embodiment of the present invention.

FIG. 3 is a diagram showing an apparatus according to a third embodiment of the present invention.

FIG. 4 is a diagram showing a flowchart of processing of the present invention.

FIG. 5 is an explanatory diagram of a function f that guides thrust acceleration of the engine.

FIG. 6 is an explanatory diagram of Y elements of functions p and q that calculate predicted values of velocity and position at a target point arrival time.

FIG. 7 is an explanatory diagram of X and Z elements of functions p and q that calculate predicted values of velocity and position at a target point arrival time.

FIG. 8 is a diagram showing a definition of a coordinate system (Part 1).

FIG. 9 is a diagram showing a definition of a coordinate system and the like (No. 2).

[Explanation of symbols]

1 ... Vertical landing control command device basic unit (generator), 2 ... Processing mode instruction unit, 3 ... Target state setting unit, 4 ... Landing control acceleration command generation unit, 5 ... Spatial stop control acceleration command generation unit, 6 ... Transitional motion control thrust and thrust direction command generator, 8 ... Airframe mass estimator, 9 ... Switch S, 10 ... Attitude command, 11 ... Engine steering angle command, 12 ... Engine nozzle, 13 ... Actuator, 20 ... Motion state detection unit (sensor), 60 ... First command generation unit, 70 ... Second command generation unit (thrust and thrust direction command generation unit), 100 ... Aircraft.

Claims (3)

[Claims] 1. In an aircraft having a rocket engine capable of changing the magnitude of thrust, (A) a processing mode instructing section (2), a target state setting section (3), and a first command generating section ( 60), a second command generation unit (70), an airframe mass estimation unit (8), and a vertical landing control command device basic unit (1) including a switch S (9), and (B) the processing The mode instruction unit (2) selects one of the landing mode (a), the space stop mode (b), or the transition operation mode (c) according to the mode instruction, and (C) the target state setting unit (3) The state of the target point corresponding to the mode selected by the processing mode instructing section (2) is selectively set, and (D) the first
The command generation section (60) includes a landing control acceleration command generation section (4), a space stop control acceleration command generation section (5),
It is composed of a thrust for thrust transition control and a thrust direction command generator (6), receives a signal from the processing mode instructor (2) and a signal from the target state setting unit (3), and also detects a motion state (20). (E) The second command generator (70) receives the signal from the landing control acceleration command generator (4) and the space stop control acceleration command generator. Input the signal from (5) and the signal from the airframe mass estimation unit (8),
(F) The airframe mass estimation unit (8) inputs signals from the transition motion control thrust of the first command generation unit (60) and the thrust direction command generation unit (6) and the second command generation unit (70). (G) The switch S (9) is turned on to the A side in the landing mode (a) and the space stop mode (b), and is turned to the B side in the transition operation mode (c) before the flight, or A target state corresponding to a designated mode is set during flight, and the command generation unit (60, 7) is set at designated processing time intervals Δt _c.
The command from 0) is repeatedly executed to change the current thrust command and thrust direction command of the rocket engine. 2. When the thrust direction of the rocket engine has a one-to-one correspondence with the attitude of the airframe, the vertical landing control command device basic unit (1) changes the thrust direction command to the attitude command of the air vehicle. The vertical landing control command device for an aircraft according to claim 1, further comprising a posture command unit (10) for conversion. 3. When the rocket engine can be tilted by the engine nozzle (12) and the actuator (13) around the attachment point to the air vehicle (100), the vertical landing control command device basic unit (1): The aircraft vertical landing control command device according to claim 1, further comprising an engine steering angle command unit (11) for converting a thrust direction command into an engine steering angle command.