JP5683224B2 - Rocket guidance calculation device, rocket guidance system, rocket, rocket guidance calculation program, and rocket guidance calculation method for rocket guidance calculation device - Google Patents

Description

  The present invention relates to a rocket guidance calculation apparatus, a rocket guidance system, a rocket, a rocket guidance calculation program, and a rocket guidance calculation method for a rocket guidance calculation apparatus for guiding a rocket.

In each country, the development of “air launch systems” that launch rockets in the air by mounting rockets on airplanes is underway.
This is because the aerial launch system has a higher degree of freedom at the launch point than the “ground launch system” that launches rockets from the ground.

  As types of rockets, there are liquid rockets using liquid fuel as a propellant (also referred to as liquid fuel rockets) and solid rockets using solid fuel as propellants (also referred to as solid fuel rockets).

Since the liquid rocket uses liquid fuel as a propellant, the amount of combustion of the liquid fuel can be adjusted by adjusting the flow rate of the liquid fuel to the combustion chamber. In other words, the liquid rocket has a “thrust cut-off function” that stops propellant combustion and eliminates thrust.
However, since the structure of the liquid rocket is complicated by pumps and piping for flowing liquid fuel to the combustion chamber, it is difficult to reduce the size of the liquid rocket.
For this reason, liquid rockets are not suitable for air-launched rockets.

On the other hand, since solid rockets use solid fuel as a propellant, they do not require complicated structures such as pumps and piping, and can be miniaturized.
For this reason, the use of solid rockets as air-launch rockets is being studied in various countries.

However, the thrust of the solid rocket cannot be stopped until the solid fuel is burned. In other words, solid rockets do not have a thrust cutoff function.
For this reason, the conventional guidance technology for guiding the liquid rocket using the thrust cutoff function cannot guide the solid rocket with high accuracy.

JP-A-62-275899 Japanese Patent Laid-Open No. 10-267597

Arthur E.M. Bryson, Jr. , Yu-Chi Ho, “Applied Optimal Control OPTIMIZATION, ESTIMATION, AND CONTROL”, Hemisphere Publishing Corporation

  An object of the present invention is to enable, for example, a solid rocket having no thrust cutoff function to be guided with high accuracy.

The rocket guidance calculation device of the present invention is
A quadratic expression of a time variable for obtaining the pitch angle of the rocket is defined as a second tangent rule, and a steering coefficient calculation unit is provided for calculating a coefficient included in the second tangent rule as a steering coefficient for obtaining the pitch angle of the rocket.

The rocket guidance calculation device further includes:
The current velocity vector of the rocket, the current position vector of the rocket, the prediction profile indicating the relationship between the elapsed time since the ignition of the rocket and the predicted value of the thrust acceleration of the rocket, and the predicted time until the end of rocket combustion A predicted state quantity calculation unit that calculates a predicted value of the rocket speed component at the end of combustion and a predicted value of the position component of the rocket at the end of combustion as a predicted state quantity,
The required value of the target value of the velocity component of the rocket at the end of combustion when the rocket flies in the target trajectory and the target value of the position component of the rocket at the end of combustion when the rocket flies in the target trajectory A request state quantity storage unit for storing,
The steering coefficient calculation unit calculates a difference between the predicted state quantity calculated by the predicted state quantity calculation unit and the requested state quantity stored in the requested state quantity storage unit as a state quantity error, and calculates the calculated state quantity error. The amount of update of the steering coefficient is calculated based on the above, and a new steering coefficient is calculated based on the calculated amount of update.

The rocket guidance system of the present invention is
The rocket guidance calculation device;
A navigation calculation device that calculates the current velocity vector of the rocket, the current position vector of the rocket, and the current non-gravity acceleration vector of the rocket by navigation calculation;
A fuselage control device for controlling the direction of thrust acceleration of the rocket according to the pitch angle of the rocket calculated from the steering coefficient calculated by the rocket guidance calculation device.

  The non-gravity acceleration vector is a thrust acceleration vector.

  The rocket of the present invention is equipped with the rocket guidance system.

The rocket guidance program of the present invention is
Define a quadratic expression of the time variable for determining the rocket pitch angle as a second tangent rule, and calculate a steering coefficient calculation process for calculating a coefficient included in the second tangent rule as a steering coefficient for determining the rocket pitch angle. Let it run.

The rocket guidance calculation program further includes:
The current velocity vector of the rocket, the current position vector of the rocket, the prediction profile indicating the relationship between the elapsed time since the ignition of the rocket and the predicted value of the thrust acceleration of the rocket, and the predicted time until the end of rocket combustion On the basis of the predicted value of the rocket velocity component at the end of combustion and the predicted value of the position component of the rocket at the end of combustion as a predicted state quantity, the computer executes a predicted state quantity calculation process,
The steering coefficient calculation process includes the target value of the rocket speed component at the end of combustion when the rocket flies on the target trajectory, and the target of the position component of the rocket at the end of combustion when the rocket flies on the target trajectory. The requested state quantity is input from a requested state quantity storage unit that stores a value as a requested state quantity, the predicted state quantity calculated by the predicted state quantity calculation unit, and the requested state quantity input from the requested state quantity storage unit, Is calculated as a state quantity error, an update amount of the steering coefficient is calculated based on the calculated state quantity error, and a new steering coefficient is calculated based on the calculated update quantity.

The rocket guidance calculation method of the rocket guidance calculation device of the present invention is:
A quadratic expression of a time variable for obtaining the rocket pitch angle is defined as a second tangent law, and a coefficient included in the second tangent law is calculated as a steering coefficient for obtaining the rocket pitch angle.

  According to the present invention, for example, a solid rocket without a thrust cutoff function can be guided with high accuracy.

1 is a configuration diagram of a rocket guidance system 200 in Embodiment 1. FIG. The figure which shows an example of the hardware resource of the rocket guidance calculation apparatus 100 in Embodiment 1. FIG. 6 is a graph showing the relationship of tan θ between the nominal profile and the second-order polynomial approximation in the first embodiment. FIG. 3 shows a dynamic coordinate system XYZ of equations of motion in the first embodiment. FIG. 3 shows a dynamic coordinate system XYZ of equations of motion in the first embodiment. 3 is a flowchart showing a rocket guidance method in the first embodiment. FIG. 3 shows a simulation apparatus 300 according to the first embodiment. 6 is a graph showing a simulation result (target trajectory: near point altitude) in the first embodiment. 6 is a graph showing a simulation result (target trajectory: far point altitude) in the first embodiment.

Embodiment 1 FIG.
A rocket guidance system for guiding a rocket to a target trajectory will be described.
This rocket guidance system is particularly effective for solid rockets that do not have a thrust cutoff function.

FIG. 1 is a configuration diagram of a rocket guidance system 200 according to the first embodiment.
The structure of the rocket guidance system 200 in Embodiment 1 is demonstrated based on FIG.

  The rocket guidance system 200 includes a rocket guidance calculation device 100, a navigation calculation device 210, and an airframe control device 220.

  The rocket guidance system 200 is mounted on a rocket (target rocket to be guided) and guides the rocket to a target trajectory.

Navigation computing device 210 calculates the position vector R of the rocket the current, the velocity vector V of the rocket the current, the navigation calculating a thrust acceleration vector a _T rocket the current.

  The airframe control device 220 controls the direction of thrust acceleration of the rocket according to the rocket pitch angle θ calculated from the steering coefficient determined by the rocket guidance calculation device 100.

The rocket guidance calculation device 100 is a device that determines the pitch angle θ of the rocket in order to guide the rocket to the target trajectory based on the second tangent law.
The quadratic tangent rule is a quadratic expression of the time variable t defined as an expression for obtaining the rocket pitch angle θ.

  The rocket guidance calculation device 100 includes a thrust acceleration prediction unit 110, a predicted state quantity calculation unit 120, a required state quantity calculation unit 130, a thrust direction determination unit 140, and a guidance calculation storage unit 190.

The guidance calculation storage unit 190 (an example of a requested state quantity storage unit) stores various data used in the rocket guidance calculation device 100.
The predicted profile 191, target trajectory information 192, predicted state quantity 193 or requested state quantity 194 is an example of data stored in the guidance calculation storage unit 190.
The prediction profile 191 is data indicating the relationship between the elapsed time since the rocket was ignited and the predicted value of thrust acceleration of the rocket.
The target trajectory information 192 is information indicating the target trajectory.
The predicted state quantity 193 is a predicted value of the rocket state quantity calculated by the predicted state quantity calculation unit 120.
The requested state quantity 194 is a target value of the rocket state quantity calculated by the requested state quantity calculating unit 130.

Thrust acceleration prediction unit 110 updates the predicted profile 191 based like thrust acceleration vector a _T rocket the current.

  The predicted state quantity calculation unit 120 calculates the predicted value of the speed component of the rocket at the end of combustion based on the current speed vector V of the rocket, the current position vector R of the rocket, and the predicted profile 191, and the end of combustion. The predicted value of the position component of the rocket at the time is calculated as a predicted state quantity 193.

  The required state quantity calculation unit 130 calculates the target value of the velocity component of the rocket at the end of combustion when the rocket flies in the target trajectory and the position component of the rocket at the end of combustion when the rocket flies in the target trajectory. The target value is calculated as the required state quantity 194 based on the prediction profile 191 and the target trajectory information 192.

The thrust direction determination unit 140 (an example of a steering coefficient calculation unit) calculates a coefficient included in the second tangent law 101 as a steering coefficient for calculating the pitch angle θ of the rocket.
In the quadratic tangent rule 101, the tangent of the pitch angle θ is expressed by a quadratic expression of a time variable t.
The steering coefficient is a coefficient A, B, or C included in the second order tangent rule 101.

Specifically, the thrust direction determination unit 140 calculates a difference between the predicted state quantity 193 and the requested state quantity 194 as a state quantity error.
The thrust direction determination unit 140 calculates the update amount of the steering coefficient based on the state quantity error.
The thrust direction determination unit 140 calculates a new steering coefficient based on the update amount of the steering coefficient.

FIG. 2 is a diagram illustrating an example of hardware resources of the rocket guidance calculation apparatus 100 according to the first embodiment.
In FIG. 2, the rocket guidance calculation apparatus 100 includes a CPU 911 (Central Processing Unit). The CPU 911 is connected to the ROM 913, the RAM 914, and the communication board 915 via the bus 912, and controls these hardware devices. The communication board 915 is connected to the network by wire or wireless.

  The ROM or RAM stores an OS (Operating System), a program group, and a file group.

  The program group includes a program that executes a function described as “unit” in the embodiment. The program is read and executed by the CPU 911. That is, the program causes the computer to function as “to part”, and causes the computer to execute the procedures and methods of “to part”.

  The file group includes various data (input, output, determination result, calculation result, processing result, etc.) used in “˜unit” described in the embodiment.

  In the embodiment, arrows included in the configuration diagrams and flowcharts mainly indicate input and output of data and signals.

  In the embodiment, what is described as “to part” may be “to circuit”, “to apparatus”, and “to device”, and “to step”, “to procedure”, and “to processing”. May be. That is, what is described as “to part” may be implemented by any of firmware, software, hardware, or a combination thereof.

  As with the rocket guidance calculation apparatus 100, the navigation calculation apparatus 210 and the airframe control apparatus 220 include hardware such as a CPU and a memory, and operate using the hardware.

  Next, the rocket tangent rule (second-order tangent rule 101) in the first embodiment will be described.

  The rocket tangent law is a relational expression between the tangent of the rocket pitch angle θ and the time variable t.

Assuming that “the sum of gravity and centrifugal force is constant”, the following bilinear tangent rule (1) is obtained.
The method for deriving the bilinear tangent rule is disclosed on pages 59-61 of Non-Patent Document 1.

  The linear tangent rule is derived by removing the distance constraint in the tangential direction from the bilinear tangent rule (1).

  Here, a second term of time is added to add a control variable in a rocket that does not have a thrust cutoff function.

In the first embodiment, the above equation (3) is referred to as “second-order tangent rule 101”.
In addition, the coefficient “A” of the zeroth order term, the coefficient “B” of the first order term, and the coefficient “C” of the second order term of the second order tangent rule (3) are called “steering coefficients”.
In particular, the coefficient “A” is referred to as “0th-order steering coefficient”, the coefficient “B” as “primary steering coefficient”, and the coefficient “C” as “secondary steering coefficient”.

FIG. 3 is a graph showing the relationship of tan θ between the nominal profile and the quadratic polynomial approximation in the first embodiment.
As shown in FIG. 3, the second-order polynomial approximation shows a value sufficiently approximated to tan θ of the nominal profile in a certain solid rocket.

  Next, the “equation of motion” which is a precondition for the rocket guidance method will be described.

4 and 5 are diagrams showing a dynamic coordinate system XYZ of the equations of motion in the first embodiment.
The dynamic coordinate system XYZ of the equation of motion in the first embodiment will be described based on FIGS.

As shown in FIG. 4, the dynamic coordinate system XYZ of the equation of motion is a coordinate system in which the point where the rocket is located is the coordinate center (origin) (the coordinate axis Y is not shown).
“A _T ” indicates the thrust acceleration vector of the rocket.
“Θ” indicates the pitch angle of the rocket.

  As shown in FIG. 5, the X axis indicates the tangential direction, the Y axis indicates the angular momentum direction, and the Z axis indicates the trajectory radial direction.

The tangential direction (X-axis) is a direction defined by “(position vector × velocity vector) × position vector” (“×” represents an outer product). However, the “position vector” and “velocity vector” are vectors expressed with the center O of the earth (see FIG. 4) as the origin.
Hereinafter, a coordinate system having the center O of the earth as the origin is referred to as an “inertial coordinate system”.

  The trajectory radial direction (Z-axis) corresponds to the direction of the position vector in the inertial coordinate system.

  The angular momentum direction (Y axis) is a direction perpendicular to the tangential direction (X axis) and the orbital radial direction (Z axis), and is defined by “position vector × velocity vector”.

  In the dynamic coordinate system XYZ shown in FIGS. 4 and 5, the following equation of motion (4) is established.

Further, the assumption of the bilinear tangent rule (1) and the linear tangent rule (2) that “the sum of gravity and centrifugal force is constant” is expressed by the following relational expression (5).
On the left side of the relational expression (5), the first term represents gravity and the second term represents centrifugal force.

FIG. 6 is a flowchart showing the rocket guidance method in the first embodiment.
A processing flow of the rocket guidance method in the first embodiment will be described with reference to FIG.

  First, the outline of the rocket guidance method will be described.

The navigation calculation device 210 calculates the rocket thrust acceleration vector a _T , velocity vector V, and position vector R (S110: navigation calculation processing).

Each “˜ unit” of the rocket guidance calculation apparatus 100 executes the following processing.
Thrust acceleration prediction unit 110, based on such a thrust acceleration vector _{a T} rocket, updates the prediction profile 191 of thrust acceleration (S120: predicting profile update process).
The predicted state quantity calculation unit 120 calculates the predicted state quantity 193 of the rocket based on the velocity vector V, the position vector R, and the prediction profile 191 (S130: predicted state quantity calculation process).
The requested state quantity calculation unit 130 calculates the requested state quantity 194 of the rocket based on the target trajectory information 192 and the predicted profile 191 (S140: requested state quantity calculation process).
The thrust direction determination unit 140 calculates the update amount of the steering coefficient based on the state amount error between the predicted state amount 193 and the requested state amount 194 (S150: state amount error calculation process, update amount calculation process).
The thrust direction determination unit 140 updates the steering coefficient based on the update amount of the steering coefficient (S151: steering coefficient calculation process).

  The airframe controller 220 changes the direction of thrust acceleration to the pitch angle θ by controlling the airframe of the rocket (S160).

  The rocket guidance method processing (S110 to S160) is repeatedly executed periodically.

  Next, details of the rocket guidance method will be described.

In S110, the navigation calculation device 210 calculates the rocket thrust acceleration vector a _T , the velocity vector V, and the position vector R at the present time.

For example, the navigation calculation apparatus 210 calculates the thrust acceleration vector a _T , the velocity vector V, and the position vector R as follows.

The navigation calculation device 210 is also called an inertial measurement unit (IMU).
The navigation calculation device 210 includes a gyro and an accelerometer. The gyro measures the angular velocity of the rocket in three axis directions (roll angle, pitch angle, yaw angle) in a dynamic coordinate system (rocket center), and the accelerometer is a rocket 3 Measure axial acceleration in a dynamic coordinate system.
Navigation computing device 210, the acceleration measured by the accelerometer coordinate conversion, and calculates the acceleration vector a _T of the inertial coordinate system (geocentric).
The navigation calculation device 210 calculates a velocity vector and a position vector of the inertial coordinate system by a general inertial navigation calculation. In the inertial navigation calculation, the navigation calculation device 210 calculates a velocity vector by integrating acceleration and angular velocity, calculates a movement vector by integrating the calculated velocity vector, and uses the calculated movement vector as the previous position vector. To calculate the current position vector.
The navigation calculation apparatus 210 performs coordinate conversion of the velocity vector and position vector calculated by inertial navigation calculation, and calculates the velocity vector V and position vector R of the inertial coordinate system.

  It progresses to S120 after S110.

In S120, the thrust acceleration prediction unit 110 of the rocket induction computing device 100, such as by entering the thrust acceleration vector a _T calculated by the navigation computing device 210, the prediction profile of thrust acceleration based like thrust acceleration vector a _T input 191 is updated.

The prediction profile 191 is history data indicating a predicted value of thrust acceleration (magnitude) of the rocket.
That is, the prediction profile 191 is data indicating the relationship between the elapsed time since the rocket was ignited and the predicted value of the thrust acceleration of the rocket.
For example, the prediction profile 191 shows the relationship between each time from when the solid fuel of the solid rocket starts to burn until the solid fuel finishes burning and the predicted value of the thrust acceleration of the rocket.

  It progresses to S130 after S120.

In S130, the predicted state quantity calculation unit 120 receives the position vector R and the velocity vector V calculated by the navigation calculation device 210, and inputs the prediction profile 191 updated in S120.
The predicted state quantity calculation unit 120 calculates the predicted state quantity 193 of the rocket at the predicted cutoff time based on the input position vector R, velocity vector V, and prediction profile 191 as follows.
The predicted cutoff time is a time indicated in the predicted profile 191 as a time when the thrust acceleration becomes zero.

The predicted state quantity calculation unit 120 performs coordinate conversion of the position vector R and the velocity vector V into a dynamic coordinate system.
The predicted state quantity calculation unit 120 integrates the above equation of motion (4) using the position vector and velocity vector obtained by coordinate conversion as initial values.
By integrating the above equation of motion (4), the predicted value of the orbital radius change rate z ′ (one point above z) at the end of combustion (the same applies hereinafter) and the predicted value of the orbital radius z A predicted value of the tangential speed x ′ can be calculated.

  The orbital radius change rate z ′ indicates the rocket velocity component (z-direction velocity), the orbit radius z indicates the rocket position component (z-direction position), and the tangential velocity x ′ indicates the rocket velocity component (x-direction). Speed).

Hereinafter, the orbit radius change rate z ′, the orbit radius z, and the tangential velocity x ′ are referred to as “rocket state quantities”.
In addition, the predicted value of the orbit radius change rate z ′ calculated in S130, the predicted value of the orbit radius z, and the predicted value of the tangential velocity x ′ are referred to as “predicted state quantity 193”.

  It progresses to S140 after S130.

In S <b> 140, the requested state quantity calculation unit 130 calculates the requested state quantity 194 based on the target trajectory information 192 and the predicted profile 191. This is because the rocket state quantity (required state quantity 194) required for guidance of the rocket is determined according to the target trajectory and the thrust acceleration.
For example, the requested state quantity calculation unit 130 calculates the requested state quantity 194 by calculating a predetermined relational expression representing the relationship among the target trajectory, the thrust acceleration, and the state quantity.
The requested state quantity 194 includes a target value for the trajectory radius change rate z ′, a target value for the trajectory radius z, and a target value for the tangential velocity x ′.
After S140, the process proceeds to S150.

In S150, the thrust direction determination unit 140 calculates a difference between the predicted state quantity 193 and the requested state quantity 194 as a state quantity error.
The thrust direction determination unit 140 calculates the following equation (6) using the calculated state quantity error and a predetermined sensitivity matrix, and calculates the update amount of the steering coefficient.

The vector on the left side of Expression (6) is a vector indicating the amount of steering coefficient update.
The matrix on the left side of the right side is a “sensitivity matrix”, and the vector on the right side of the right side is a vector indicating “state quantity error”.
Each element “∂α / ∂β” of the sensitivity matrix means a derivative obtained by partial differentiation of the relational expression of the state quantity α including the steering coefficient with the steering coefficient β. The derivative is integrated by substituting the current steering coefficients A, B, and C.

  After S150, the process proceeds to S151.

In S151, the thrust direction determination unit 140 updates the steering coefficient by adding the update amount of the steering coefficient calculated in S151 to the current steering coefficient (the previously calculated steering coefficient).
That is, the thrust direction determination unit 140 calculates the following formula (7) to calculate a new steering coefficient.
In Expression (7), “A”, “B”, and “C” indicate new steering coefficients, “A ₀ ”, “B ₀ ”, and “C ₀ ” indicate current steering coefficients, and “δA”, “δB”, “ “δC” indicates an update amount of the steering coefficient.

In S160, the body control device 220 inputs the steering coefficients “A”, “B”, and “C” calculated in S152 from the rocket guidance calculation device 100.
Aircraft control unit 220, the pitch angle (orientation of the rocket) thrust acceleration vector a _T rocket controls the rocket body such that the pitch angle θ calculated from the steering coefficients input from the rocket induced computing device 100.
After S160, the process returns to S110.

  Next, the effect of the rocket guidance method (see FIG. 6) in the first embodiment will be described.

Conventionally, “linear tangent law” or “linear sine law” has been used to guide liquid rockets using liquid fuel.
The linear tangent rule (8-1) and the linear sine rule (8-2) are shown below.
As described below, the linear tangent rule (8-1) and the linear sine rule (8-2) are linear expressions of the time variable t.

  The linear tangent rule (8-1) is an expression based on the bilinear tangent rule (1) described above.

The bilinear tangent rule (1) is based on a predetermined relationship between the boundary condition and the steering coefficient C _n with the track radius z, the track radius change rate z ′, the tangential distance x, and the tangential velocity x ′ as boundary conditions. It is a formula.
Here, the linear tangent rule (8-1) is derived by removing the tangential distance x from the boundary condition of the bilinear tangent rule (1).
A method for deriving the linear tangent rule (8-1) is disclosed on pages 60-61 of Non-Patent Document 1.

  The linear sine rule (8-2) is an expression obtained by modifying the linear tangent rule (8-1) on the assumption that “the pitch angle θ is very small”.

The liquid rocket can control the “cut-off time” at which the thrust acceleration becomes zero by stopping the combustion of the liquid fuel.
Further, the tangential speed x ′ can be controlled by controlling the cutoff time.
In order to control the trajectory radius z and the trajectory radius change rate z ′, which are the remaining boundary conditions, it is sufficient to obtain two steering coefficients C ₄ and C ₅ .

However, with the two steering coefficients C ₄ and C ₅ , the three boundary conditions “track radius z”, “track radius change rate z ′”, and “tangential velocity x ′” cannot be sufficiently controlled.

On the other hand, the solid rocket cannot control the “cut-off time” because the thrust acceleration does not become zero until the solid fuel is combusted.
That is, the tangential speed x ′ cannot be controlled by controlling the cut-off time, and the linear tangent rule (8-1) having only two steering coefficients can guide the solid rocket with high accuracy. Can not.

  Therefore, in the first embodiment, the solid rocket is guided with high accuracy by using the second-order tangent law 101 (see Expression (3)) having three steering coefficients A, B, and C.

FIG. 7 is a diagram showing a simulation apparatus 300 in the first embodiment.
FIG. 8 is a graph showing a simulation result (target trajectory: near altitude) in the first embodiment.
FIG. 9 is a graph showing a simulation result (target trajectory: far-point altitude) in the first embodiment.

  8 and 9 show the results of simulating the rocket guidance method of the first embodiment (see FIG. 6) using the simulation apparatus 300 shown in FIG.

As shown in FIG. 7, the simulation apparatus 300 includes a simple dynamics model calculation unit 310, a guidance logic calculation unit 320, and a simulation result output unit 330.
The simple dynamics model calculation unit 310 simulates the navigation calculation device 210 and the airframe control device 220 (see FIG. 1).
The guidance logic calculation unit 320 simulates the rocket guidance calculation device 100 (see FIG. 1) using the second-order tangent rule 101 (see equation (3)).
In addition, the guidance logic calculation unit 320 simulates a conventional rocket guidance calculation apparatus that uses the linear tangent rule (8-1).
The simulation result output unit 330 outputs the simulation results of the simple dynamics model calculation unit 310 and the guidance logic calculation unit 320.

  For example, the simple dynamics model calculation unit 310, the guidance logic calculation unit 320, and the simulation result output unit 330 can be implemented by a “Matlab / Simlink program”.

Similar to the rocket guidance calculation apparatus 100 (see FIG. 2), the simulation apparatus 300 includes hardware such as a CPU and a memory, and each “˜ unit” of the simulation apparatus 300 operates by hardware.
For example, the simulation result output unit 330 graphs the simulation result (FIGS. 8 and 9) and displays it on the display device.

  FIG. 8 shows the relative error (shift amount from the target trajectory) of the simulation result with respect to the near-point altitude of the target trajectory, and FIG. 9 shows the relative error of the simulation result with respect to the far-point altitude of the target trajectory. The lower the bar graph, the smaller the deviation from the target trajectory.

8 and 9, “Nominal” indicates the simulation result when no error is given to the assumed thrust and specific thrust, and “Thrust error” indicates the simulation result when the error is increased by 3% to the thrust, “Specific thrust error” indicates a simulation result when an error of 1% up is given to the specific thrust.
(L) shows a simulation result when the conventional linear tangent rule is used, and (Q) shows a simulation result when the second-order tangent rule of the first embodiment is used.
Further, (Q ′) shows a simulation result when the second tangent law is used with the thrust acceleration increased by 7%.

  As shown in FIGS. 8 and 9, the solid rocket can be guided with higher accuracy than when the second-order tangent rule is used (Q, Q ′) and when the conventional linear tangent rule is used (L). it can.

  DESCRIPTION OF SYMBOLS 100 Rocket guidance calculation apparatus, 101 Secondary tangent rule, 110 Thrust acceleration prediction part, 120 Predicted state quantity calculation part, 130 Required state quantity calculation part, 140 Thrust direction determination part, 190 Guidance calculation memory part, 191 Prediction profile, 192 Target Trajectory information, 193 predicted state quantity, 194 required state quantity, 200 rocket guidance system, 210 navigation calculation device, 220 airframe control device, 300 simulation device, 310 simple dynamics model computation unit, 320 guidance logic computation unit, 330 simulation result output unit 911 CPU, 912 bus, 913 ROM, 914 RAM, 915 communication board.

Claims (6)

A steering coefficient calculation unit for calculating a coefficient included in a second tangent law, which is a quadratic expression of a time variable for obtaining a rocket pitch angle, as a steering coefficient for obtaining a rocket pitch angle ;
The current velocity vector of the rocket, the current position vector of the rocket, the prediction profile indicating the relationship between the elapsed time since the ignition of the rocket and the predicted value of the thrust acceleration of the rocket, and the predicted time until the end of rocket combustion A predicted state quantity calculation unit that calculates a predicted value of the rocket speed component at the end of combustion and a predicted value of the position component of the rocket at the end of combustion as a predicted state quantity,
The required value of the target value of the velocity component of the rocket at the end of combustion when the rocket flies in the target trajectory and the target value of the position component of the rocket at the end of combustion when the rocket flies in the target trajectory A request state quantity storage unit for storing,
The steering coefficient calculation unit calculates a difference between the predicted state quantity calculated by the predicted state quantity calculation unit and the requested state quantity stored in the requested state quantity storage unit as a state quantity error, and calculates the calculated state quantity error. A rocket guidance calculation device, wherein an update amount of the steering coefficient is calculated based on the calculation amount, and a new steering coefficient is calculated based on the calculated update amount. Rocket induction computing device of claim 1 Symbol placement,
A navigation calculation device that calculates the current velocity vector of the rocket, the current position vector of the rocket, and the current non-gravity acceleration vector of the rocket by navigation calculation;
A rocket guidance system comprising: a body control device for controlling the direction of thrust acceleration of a rocket according to a pitch angle of the rocket calculated from a steering coefficient calculated by the rocket guidance calculation device. The rocket guidance system according to claim 2, wherein the non-gravity acceleration vector is a thrust acceleration vector. Claim 2 or claim 3 rocket equipped with rocket guidance system according. A steering coefficient calculation process for calculating a coefficient included in a second tangent law, which is a quadratic expression of a time variable for obtaining a rocket pitch angle, as a steering coefficient for obtaining a rocket pitch angle ;
The current velocity vector of the rocket, the current position vector of the rocket, the prediction profile indicating the relationship between the elapsed time since the ignition of the rocket and the predicted value of the thrust acceleration of the rocket, and the predicted time until the end of rocket combustion based on the prediction value of the velocity component of the rocket combustion end, Ru to execute the predicted state quantity calculating process to calculate a predicted state amount and a predicted value of the position components of the rocket combustion end to the computer A rocket guidance calculation program ,
The steering coefficient calculation process includes the target value of the rocket speed component at the end of combustion when the rocket flies on the target trajectory, and the target of the position component of the rocket at the end of combustion when the rocket flies on the target trajectory. The requested state quantity is input from a requested state quantity storage unit that stores a value as a requested state quantity, the predicted state quantity calculated by the predicted state quantity calculation process, and the requested state quantity input from the requested state quantity storage unit, Is calculated as a state quantity error, an update amount of the steering coefficient is calculated based on the calculated state quantity error, and a new steering coefficient is calculated based on the calculated update quantity. Rocket guidance calculation program. A steering coefficient calculation process for calculating a coefficient included in a second tangent law, which is a quadratic expression of a time variable for obtaining a rocket pitch angle, as a steering coefficient for obtaining a rocket pitch angle ;
The current velocity vector of the rocket, the current position vector of the rocket, the prediction profile indicating the relationship between the elapsed time since the ignition of the rocket and the predicted value of the thrust acceleration of the rocket, and the predicted time until the end of rocket combustion A predicted state quantity calculation process for calculating a predicted value of the velocity component of the rocket at the end of combustion and a predicted value of the position component of the rocket at the end of combustion as a predicted state quantity based on
A rocket guidance calculation method comprising:
The steering coefficient calculation process includes the target value of the rocket speed component at the end of combustion when the rocket flies on the target trajectory, and the target of the position component of the rocket at the end of combustion when the rocket flies on the target trajectory. The requested state quantity is input from a requested state quantity storage unit that stores a value as a requested state quantity, the predicted state quantity calculated by the predicted state quantity calculation process, and the requested state quantity input from the requested state quantity storage unit, The difference between the two is calculated as a state quantity error, an update amount of the steering coefficient is calculated based on the calculated state quantity error, and a new steering coefficient is calculated based on the calculated update amount. A rocket guidance calculation method for a rocket guidance calculation device.

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